Electric propulsion options for 10 kW class earth space missions

NASA Technical Reports Server (NTRS)

Patterson, M. J.; Curran, Francis M.

1989-01-01

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have been evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA II 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10(7) to 2.1x10(7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA II 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10(6) to 3.6x10(6) N-s, and approximately 662 to 1072 m/s, respectively.

Options and Risk for Qualification of Electric Propulsion System

NASA Technical Reports Server (NTRS)

Bailey, Michelle; Daniel, Charles; Cook, Steve (Technical Monitor)

2002-01-01

Electric propulsion vehicle systems envelop a wide range of propulsion alternatives including solar and nuclear, which present unique circumstances for qualification. This paper will address the alternatives for qualification of electric propulsion spacecraft systems. The approach taken will be to address the considerations for qualification at the various levels of systems definition. Additionally, for each level of qualification the system level risk implications will be developed. Also, the paper will explore the implications of analysis verses test for various levels of systems definition, while retaining the objectives of a verification program. The limitations of terrestrial testing will be explored along with the risk and implications of orbital demonstration testing. The paper will seek to develop a template for structuring of a verification program based on cost, risk and value return. A successful verification program should establish controls and define objectives of the verification compliance program. Finally the paper will seek to address the political and programmatic factors, which may impact options for system verification.

Space station propulsion-ECLSS interaction study

NASA Technical Reports Server (NTRS)

Brennan, Scott M.

1986-01-01

The benefits of the utilization of effluents of the Space Station Environmental Control and Life Support (ECLS) system are examined. Various ECLSS-propulsion system interaction options are evaluated and compared on the basis of weight, volume, and power requirements. Annual propulsive impulse to maintain station altitude during a complete solar cycle of eleven years and the effect on station resupply are considered.

High Power Electric Propulsion System for NEP: Propulsion and Trajectory Options

DOE Office of Scientific and Technical Information (OSTI.GOV)

Koppel, Christophe R.; Duchemin, Olivier; Valentian, Dominique

Recent US initiatives in Nuclear Propulsion lend themselves naturally to raising the question of the assessment of various options and particularly to propose the High Power Electric Propulsion Subsystem (HPEPS) for the Nuclear Electric Propulsion (NEP). The purpose of this paper is to present the guidelines for the HPEPS with respect to the mission to Mars, for automatic probes as well as for manned missions. Among the various options, the technological options and the trajectory options are pointed out. The consequences of the increase of the electrical power of a thruster are first an increase of the thrust itself, butmore » also, as a general rule, an increase of the thruster performance due to its higher efficiency, particularly its specific impulse increase. The drawback is as a first parameter, the increase of the thruster's size, hence the so-called 'thrust density' shall be high enough or shall be drastically increased for ions thrusters. Due to the large mass of gas needed to perform the foreseen missions, the classical xenon rare gas is no more in competition, the total world production being limited to 20 -40 tons per year. Thus, the right selection of the propellant feeding the thruster is of prime importance. When choosing a propellant with lower molecular mass, the consequences at thruster level are an increase once more of the specific impulse, but at system level the dead mass may increase too, mainly because the increase of the mass of the propellant system tanks. Other alternatives, in rupture with respect to the current technologies, are presented in order to make the whole system more attractive. The paper presents a discussion on the thruster specific impulse increase that is sometime considered an increase of the main system performances parameter, but that induces for all electric propulsion systems drawbacks in the system power and mass design that are proportional to the thruster specific power increase (kW/N). The electric thruster specific impulse shall be optimized w.r.t. the mission. The trajectories taken into account in the paper are constrained by the allowable duration of the travel and the launcher size. The multi-arcs trajectories to Mars (using an optimized combination of chemical and Electric propulsion) are presented in detail. The compatibility with NEP systems that implies orbiting a sizeable nuclear reactor and a power generation system capable of converting thermal into electric power, with minimum mass and volumes fitting in with Ariane 5 or the Space Shuttle bay, is assessed.« less

Solar Electric and Chemical Propulsion Technology Applications to a Titan Orbiter/Lander Mission

NASA Technical Reports Server (NTRS)

Cupples, Michael

2007-01-01

Several advanced propulsion technology options were assessed for a conceptual Titan Orbiter/Lander mission. For convenience of presentation, the mission was broken into two phases: interplanetary and Titan capture. The interplanetary phase of the mission was evaluated for an advanced Solar Electric Propulsion System (SEPS), while the Titan capture phase was evaluated for state-of-art chemical propulsion (NTO/Hydrazine), three advanced chemical propulsion options (LOX/Hydrazine, Fluorine/Hydrazine, high Isp mono-propellant), and advanced tank technologies. Hence, this study was referred to as a SEPS/Chemical based option. The SEPS/Chemical study results were briefly compared to a 2002 NASA study that included two general propulsion options for the same conceptual mission: an all propulsive based mission and a SEPS/Aerocapture based mission. The SEP/Chemical study assumed identical science payload as the 2002 NASA study science payload. The SEPS/Chemical study results indicated that the Titan mission was feasible for a medium launch vehicle, an interplanetary transfer time of approximately 8 years, an advanced SEPS (30 kW), and current chemical engine technology (yet with advanced tanks) for the Titan capture. The 2002 NASA study showed the feasibility of the mission based on a somewhat smaller medium launch vehicle, an interplanetary transfer time of approximately 5.9 years, an advanced SEPS (24 kW), and advanced Aerocapture based propulsion technology for the Titan capture. Further comparisons and study results were presented for the advanced chemical and advanced tank technologies.

Clustered engine study

NASA Technical Reports Server (NTRS)

Shepard, Kyle; Sager, Paul; Kusunoki, Sid; Porter, John; Campion, AL; Mouritzan, Gunnar; Glunt, George; Vegter, George; Koontz, Rob

1993-01-01

Several topics are presented in viewgraph form which together encompass the preliminary assessment of nuclear thermal rocket engine clustering. The study objectives, schedule, flow, and groundrules are covered. This is followed by the NASA groundrules mission and our interpretation of the associated operational scenario. The NASA reference vehicle is illustrated, then the four propulsion system options are examined. Each propulsion system's preliminary design, fluid systems, operating characteristics, thrust structure, dimensions, and mass properties are detailed as well as the associated key propulsion system/vehicle interfaces. A brief series of systems analysis is also covered including: thrust vector control requirements, engine out possibilities, propulsion system failure modes, surviving system requirements, and technology requirements. An assessment of vehicle/propulsion system impacts due to the lessons learned are presented.

Electric Propulsion Options for 10 kW Class Earth-Space Missions

NASA Technical Reports Server (NTRS)

Patterson, M. J.; Curran, Francis M.

1989-01-01

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment were evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA 2 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10 (exp 7) to 2.1x10 (exp 7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA 2 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10 (exp 6) to 3.6x10 (exp 6) N-s, and approximately 662 to 1072 m/s, respectively.

Propulsion Options for the Global Precipitation Measurement Core Satellite

NASA Technical Reports Server (NTRS)

Cardiff, Eric H.; Davis, Gary T.; Folta, David C.

2003-01-01

This study was conducted to evaluate several propulsion system options for the Global Precipitation Measurement (GPM) core satellite. Orbital simulations showed clear benefits for the scientific data to be obtained at a constant orbital altitude rather than with a decay/reboost approach. An orbital analysis estimated the drag force on the satellite will be 1 to 12 mN during the five-year mission. Four electric propulsion systems were identified that are able to compensate for these drag forces and maintain a circular orbit. The four systems were the UK-10/TS and the NASA 8 cm ion engines, and the ESA RMT and RITl0 EVO radio-frequency ion engines. The mass, cost, and power requirements were examined for these four systems. The systems were also evaluated for the transfer time from the initial orbit of 400 x 650 km altitude orbit to a circular 400 km orbit. The transfer times were excessive, and as a consequence a dual system concept (with a hydrazine monopropellant system for the orbit transfer and electric propulsion for drag compensation) was examined. Clear mass benefits were obtained with the dual system, but cost remains an issue because of the larger power system required for the electric propulsion system. An electrodynamic tether was also evaluated in this trade study.

Nuclear propulsion technology development - A joint NASA/Department of Energy project

NASA Technical Reports Server (NTRS)

Clark, John S.

1992-01-01

NASA-Lewis has undertaken the conceptual development of spacecraft nuclear propulsion systems with DOE support, in order to establish the bases for Space Exploration Initiative lunar and Mars missions. This conceptual evolution project encompasses nuclear thermal propulsion (NTP) and nuclear electric propulsion (NEP) systems. A technology base exists for NTP in the NERVA program files; more fundamental development efforts are entailed in the case of NEP, but this option is noted to offer greater advantages in the long term.

Hybrids of Solar Sail, Solar Electric, and Solar Thermal Propulsion for Solar-System Exploration

NASA Technical Reports Server (NTRS)

Wilcox, Brian H.

2012-01-01

Solar sails have long been known to be an attractive method of propulsion in the inner solar system if the areal density of the overall spacecraft (S/C) could be reduced to approx.10 g/sq m. It has also long been recognized that the figure (precise shape) of useful solar sails needs to be reasonably good, so that the reflected light goes mostly in the desired direction. If one could make large reflective surfaces with reasonable figure at an areal density of approx.10 g/sq m, then several other attractive options emerge. One is to use such sails as solar concentrators for solar-electric propulsion. Current flight solar arrays have a specific output of approx. 100W/kg at 1 Astronomical Unit (AU) from the sun, and near-term advances promise to significantly increase this figure. A S/C with an areal density of 10 g/sq m could accelerate up to 29 km/s per year as a solar sail at 1 AU. Using the same sail as a concentrator at 30 AU, the same spacecraft could have up to approx. 45 W of electric power per kg of total S/C mass available for electric propulsion (EP). With an EP system that is 50% power-efficient, exhausting 10% of the initial S/C mass per year as propellant, the exhaust velocity is approx. 119 km/s and the acceleration is approx. 12 km/s per year. This hybrid thus opens attractive options for missions to the outer solar system, including sample-return missions. If solar-thermal propulsion were perfected, it would offer an attractive intermediate between solar sailing in the inner solar system and solar electric propulsion for the outer solar system. In the example above, both the solar sail and solar electric systems don't have a specific impulse that is near-optimal for the mission. Solar thermal propulsion, with an exhaust velocity of the order of 10 km/s, is better matched to many solar system exploration missions. This paper derives the basic relationships between these three propulsion options and gives examples of missions that might be enabled by such hybrids.

Direct drive options for electric propulsion systems

NASA Technical Reports Server (NTRS)

Hamley, John A.

1995-01-01

Power processing units (PPU's) in an electric propulsion system provide many challenging integration issues. The PPU must provide power to the electric thruster while maintaining compatibility with all of the spacecraft power and data systems. Inefficiencies in the power processor produce heat, which must be radiated to the environment in order to ensure reliable operation. Although PPU efficiencies are generally greater than 0.9, heat loads are often substantial. This heat must be rejected by thermal control systems which generally have specific masses of 15-30 kg/kW. PPU's also represent a large fraction of the electric propulsion system dry mass. Simplification or elimination of power processing in a propulsion system would reduce the electric propulsion system specific mass and improve the overall reliability and performance. A direct drive system would eliminate all or some of the power supplies required to operate a thruster by directly connecting the various thruster loads to the solar array. The development of concentrator solar arrays has enabled power bus voltages in excess of 300 V which is high enough for direct drive applications for Hall thrusters such as the Stationary Plasma Thruster (SPT). The option of solar array direct drive for SPT's is explored to provide a comparison between conventional and direct drive system mass.

Fission Technology for Exploring and Utilizing the Solar System

NASA Technical Reports Server (NTRS)

Houts, Mike; VanDyke, Melissa; Godfroy, Tom; Pedersen, Kevin; Martin, James; Dickens, Ricky; Salvail, Pat; Hrbub, Ivana; Schmidt, George R. (Technical Monitor)

2000-01-01

Fission technology can enable rapid, affordable access to any point in the solar system. Potential fission-based transportation options include bimodal nuclear thermal rockets, high specific energy propulsion systems, and pulsed fission propulsion systems. In-space propellant re-supply enhances the effective performance of all systems, but requires significant infrastructure development. Safe, timely, affordable utilization of first-generation space fission propulsion systems will enable the development of more advanced systems. First generation space systems will build on over 45 years of US and international space fission system technology development to minimize cost,

On-Board Chemical Propulsion Technology

NASA Technical Reports Server (NTRS)

Reed, Brian D.

2004-01-01

On-board propulsion functions include orbit insertion, orbit maintenance, constellation maintenance, precision positioning, in-space maneuvering, de-orbiting, vehicle reaction control, planetary retro, and planetary descent/ascent. This paper discusses on-board chemical propulsion technology, including bipropellants, monopropellants, and micropropulsion. Bipropellant propulsion has focused on maximizing the performance of Earth storable propellants by using high-temperature, oxidation-resistant chamber materials. The performance of bipropellant systems can be increased further, by operating at elevated chamber pressures and/or using higher energy oxidizers. Both options present system level difficulties for spacecraft, however. Monopropellant research has focused on mixtures composed of an aqueous solution of hydroxl ammonium nitrate (HAN) and a fuel component. HAN-based monopropellants, unlike hydrazine, do not present a vapor hazard and do not require extraordinary procedures for storage, handling, and disposal. HAN-based monopropellants generically have higher densities and lower freezing points than the state-of-art hydrazine and can higher performance, depending on the formulation. High-performance HAN-based monopropellants, however, have aggressive, high-temperature combustion environments and require advances in catalyst materials or suitable non-catalytic ignition options. The objective of the micropropulsion technology area is to develop low-cost, high-utility propulsion systems for the range of miniature spacecraft and precision propulsion applications.

Lunar lander and return propulsion system trade study

NASA Technical Reports Server (NTRS)

Hurlbert, Eric A.; Moreland, Robert; Sanders, Gerald B.; Robertson, Edward A.; Amidei, David; Mulholland, John

1993-01-01

This trade study was initiated at NASA/JSC in May 1992 to develop and evaluate main propulsion system alternatives to the reference First Lunar Outpost (FLO) lander and return-stage transportation system concept. Thirteen alternative configurations were developed to explore the impacts of various combinations of return stage propellants, using either pressure or pump-fed propulsion systems and various staging options. Besides two-stage vehicle concepts, the merits of single-stage and stage-and-a-half options were also assessed in combination with high-performance liquid oxygen and liquid hydrogen propellants. Configurations using an integrated modular cryogenic engine were developed to assess potential improvements in packaging efficiency, mass performance, and system reliability compared to non-modular cryogenic designs. The selection process to evaluate the various designs was the analytic hierarchy process. The trade study showed that a pressure-fed MMH/N2O4 return stage and RL10-based lander stage is the best option for a 1999 launch. While results of this study are tailored to FLO needs, the design date, criteria, and selection methodology are applicable to the design of other crewed lunar landing and return vehicles.

Cooling of Electric Motors Used for Propulsion on SCEPTOR

NASA Technical Reports Server (NTRS)

Christie, Robert J.; Dubois, Arthur; Derlaga, Joseph M.

2017-01-01

NASA is developing a suite of hybrid-electric propulsion technologies for aircraft. These technologies have the benefit of lower emissions, diminished noise, increased efficiency, and reduced fuel burn. These will provide lower operating costs for aircraft operators. Replacing internal combustion engines with distributed electric propulsion is a keystone of this technology suite, but presents many new problems to aircraft system designers. One of the problems is how to cool these electric motors without adding significant aerodynamic drag, cooling system weight or fan power. This paper discusses the options evaluated for cooling the motors on SCEPTOR (Scalable Convergent Electric Propulsion Technology and Operations Research): a project that will demonstrate Distributed Electric Propulsion technology in flight. Options for external and internal cooling, inlet and exhaust locations, ducting and adjustable cowling, and axial and centrifugal fans were evaluated. The final design was based on a trade between effectiveness, simplicity, robustness, mass and performance over a range of ground and flight operation environments.

Operationally efficient propulsion system study (OEPSS) data book. Volume 7; Launch Operations Index (LOI) Design Features and Options

NASA Technical Reports Server (NTRS)

Ziese, James M.

1992-01-01

A design tool of figure of merit was developed that allows the operability of a propulsion system design to be measured. This Launch Operations Index (LOI) relates Operations Efficiency to System Complexity. The figure of Merit can be used by conceptual designers to compare different propulsion system designs based on their impact on launch operations. The LOI will improve the design process by making sure direct launch operations experience is a necessary feedback to the design process.

The NASA Electric Propulsion Program

NASA Technical Reports Server (NTRS)

Callahan, Lisa Wood; Curran, Francis M.

1996-01-01

Nearly all space missions require on-board propulsion systems and these systems typically have a major impact on spacecraft mass and cost. Electric propulsion systems offer major performance advantages over conventional chemical systems for many mission functions and the NASA Office of Space Access and Technology (OSAT) supports an extensive effort to develop the technology for high-performance, on-board electric propulsion system options to enhance and enable near- and far-term US space missions. This program includes research and development efforts on electrothermal, electrostatic, and electromagnetic propulsion system technologies to cover a wide range of potential applications. To maximize expectations of technology transfer, the program emphasizes strong interaction with the user community through a variety of cooperative and contracted approaches. This paper provides an overview of the OSAT electric propulsion program with an emphasis on recent progress and future directions.

Rocket Based Combined Cycle (RBCC) Propulsion Workshop, volume 2

NASA Technical Reports Server (NTRS)

Chojnacki, Kent T.

1992-01-01

The goal of the Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop, was to impart technology information to the propulsion community with respect to hypersonic combined cycle propulsion capabilities. The major recommendation resulting from this technology workshop was as follows: conduct a systems-level applications study to define the desired propulsion system and vehicle technology requirements for LEO launch vehicles. All SSTO and TSTO options using the various propulsion systems (airbreathing combined cycle, rocket-based combined cycle, and all rocket) must be considered. Such a study should be accomplished as soon as possible. It must be conducted with a consistent set of ground rules and assumptions. Additionally, the study should be conducted before any major expenditures on a RBCC technology development program occur.

Preliminary Assessment of Using Gelled and Hybrid Propellant Propulsion for VTOL/SSTO Launch Systems

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan; OLeary, Robert; Pelaccio, Dennis G.

1998-01-01

A novel, reusable, Vertical-Takeoff-and-Vertical-Takeoff-and-Landing, Single-Stage-to-Orbit (VTOL/SSTO) launch system concept, named AUGMENT-SSTO, is presented in this paper to help quantify the advantages of employing gelled and hybrid propellant propulsion system options for such applications. The launch vehicle system concept considered uses a highly coupled, main high performance liquid oxygen/liquid hydrogen (LO2/LH2) propulsion system, that is used only for launch, while a gelled or hybrid propellant propulsion system auxiliary propulsion system is used during final orbit insertion, major orbit maneuvering, and landing propulsive burn phases of flight. Using a gelled or hybrid propellant propulsion system for major orbit maneuver burns and landing has many advantages over conventional VTOL/SSTO concepts that use LO2/LH2 propulsion system(s) burns for all phases of flight. The applicability of three gelled propellant systems, O2/H2/Al, O2/RP-1/Al, and NTO/MMH/Al, and a state-of-the-art (SOA) hybrid propulsion system are examined in this study. Additionally, this paper addresses the applicability of a high performance gelled O2/H2 propulsion system to perform the primary, as well as the auxiliary propulsion system functions of the vehicle.

Effect of Sidewall Configurations on Hypersonic Intake Performance

NASA Astrophysics Data System (ADS)

Kim, Seihwan; Park, Ji Hyun; Jeung, In-Seuck; Lee, Hyoung Jin

For reusable space launchers and hypersonic flight vehicles, use of an air-breathing propulsion system with supersonic combustion is the most promising option in terms of cost effectiveness. At this point, only the scramjet propulsion system provides a real alternative to expensive rocket driven systems, which currently are the only way to reach a hypersonics speeds.

Utilizing Fission Technology to Enable Rapid and Affordable Access to any Point in the Solar System

NASA Technical Reports Server (NTRS)

Houts, Mike; Bonometti, Joe; Morton, Jeff; Hrbud, Ivana; Bitteker, Leo; VanDyke, Melissa; Godfroy, T.; Pedersen, K.; Dobson, C.; Patton, B.;

Beamed energy for space craft propulsion - Conceptual status and development potential

NASA Technical Reports Server (NTRS)

Sercel, Joel C.; Frisbee, Robert H.

1987-01-01

This paper outlines the results of a brief study that sought to identify and characterize beamed energy spacecraft propulsion concepts that may have positive impact on the economics of space industrialization. It is argued that the technology of beamed energy propulsion systems may significantly improve the prospects for near-term colonization of outer space. It is tentatively concluded that, for space industrialization purposes, the most attractive near-term beamed energy propulsion systems are based on microwave technology. This conclusion is reached based on consideration of the common features that exist between beamed microwave propulsion and the Solar Power Satellite (SPS) concept. Laser power beaming also continues to be an attractive option for spacecraft propulsion due to the reduced diffraction-induced beam spread afforded by laser radiation wavelengths. The conceptual status and development potential of a variety of beamed energy propulsion concepts are presented. Several alternative space transportation system concepts based on beamed energy propulsion are described.

Design and Development of a Methane Cryogenic Propulsion Stage for Human Mars Exploration

NASA Technical Reports Server (NTRS)

Percy, Thomas K.; Polsgrove, Tara; Turpin, Jason; Alexander, Leslie

2016-01-01

NASA is currently working on the Evolvabe Mars Campaign (EMC) study to outline transportation and mission options for human exploration of Mars. One of the key aspects of the EMC is leveraging current and planned near-term technology investments to build an affordable and evolvable approach to Mars exploration. This leveraging of investments includes the use of high-power Solar Electric Propulsion (SEP) systems, evolved from those currently under development in support of the Asteroid Redirect Mission (ARM), to deliver payloads to Mars. The EMC is considering several transportation options that combine solar electric and chemical propulsion technologies to deliver crew and cargo to Mars. In one primary architecture option, the SEP propulsion system is used to pre-deploy mission elements to Mars while a high-thrust chemical propulsion system is used to send crew on faster ballistic transfers between Earth and Mars. This high-thrust chemical system uses liquid oxygen - liquid methane main propulsion and reaction control systems integrated into the Methane Cryogenic Propulsion Stage (MCPS). Over the past year, there have been several studies completed to provide critical design and development information related to the MCPS. This paper is intended to provide a summary of these efforts. A summary of the current point of departure design for the MCPS is provided as well as an overview of the mission architecture and concept of operations that the MCPS is intended to support. To leverage the capabilities of solar electric propulsion to the greatest extent possible, the EMC architecture pre-deploys to Mars orbit the stages required for returning crew from Mars. While this changes the risk posture of the architecture, it can provide some mass savings by using higher-efficiency systems for interplanetary transfer. However, this does introduce significantly longer flight times to Mars which, in turn, increases the overall lifetime of the stages to as long as 2500 days. This unique aspect to the concept of operations introduces several challenges, specifically related to propellant storage and engine reliability. These challenges and some potential solutions are discussed. Specific focus is provided on two key technology areas; propulsion and cryogenic fluid management. In the area of propulsion development, the development of an integrated methane propulsion system that combines both main propulsion and reaction control is discussed. This includes an overview of potential development paths, areas where development for Mars applications are complementary to development efforts underway in other parts of the aerospace industry, and commonality between the MCPS methane propulsion applications and other Mars elements, including the Mars lander systems. This commonality is a key affordability aspect of the Evolvable Mars Campaign. A similar discussion is provided for cryogenic fluid management technologies including a discussion of how using cryo propulsion in the Mars transportation application not only provides performance benefits but also leverages decades of technology development investments made by NASA and its aerospace contractor community.

Design and Development of a Methane Cryogenic Propulsion Stage for Human Mars Exploration

NASA Technical Reports Server (NTRS)

Percy, Thomas K.; Polsgrove, Tara; Turpin, Jason; Alexander, Leslie

2016-01-01

NASA is currently working on the Evolvabe Mars Campaign (EMC) study to outline transportation and mission options for human exploration of Mars. One of the key aspects of the EMC is leveraging current and planned near-term technology investments to build an affordable and evolvable approach to Mars exploration. This leveraging of investments includes the use of high-power Solar Electric Propulsion (SEP) systems evolved from those currently under development in support of the Asteroid Redirect Mission to deliver payloads to Mars. The EMC is considering several transportation options that combine solar electric and chemical propulsion technologies to deliver crew and cargo to Mars. In one primary architecture option, the SEP propulsion system is used to pre-deploy mission elements to Mars while a high-thrust chemical propulsion system is used to send crew on faster ballistic transfers between Earth and Mars. This high-thrust chemical system uses liquid oxygen - liquid methane main propulsion and reaction control systems integrated into the Methane Cryogenic Propulsion Stage (MCPS). Over the past year, there have been several studies completed to provide critical design and development information related to the MCPS. This paper is intended to provide a summary of these efforts. A summary of the current point of departure design for the MCPS is provided as well as an overview of the mission architecture and concept of operations that the MCPS is intended to support. To leverage the capabilities of solar electric propulsion to the greatest extent possible, the EMC architecture pre-deploys the required stages for returning crew from Mars. While this changes the risk posture of the architecture, it provides mass savings by using higher-efficiency systems for interplanetary transfer. However, this does introduce significantly longer flight times to Mars which, in turn, increases the overall lifetime of the stages to as long as 3000 days. This unique aspect to the concept of operations introduces several challenges, specifically related to propellant storage and engine reliability. These challenges and some potential solutions are discussed. Specific focus is provided on two key technology areas; propulsion and cryogenic fluid management. In the area of propulsion development, the development of an integrated methane propulsion system that combines both main propulsion and reaction control is discussed. This includes an overview of potential development paths, areas where development for Mars applications are complementary to development efforts underway in other parts of the aerospace industry, and commonality between the MCPS methane propulsion applications and other Mars elements, including the Mars lander systems. This commonality is a key affordability aspect of the Evolvable Mars Campaign. A similar discussion is provided for cryogenic fluid management technologies including a discussion of how using cryo-propulsion in the Mars transportation application not only provides performance benefits but also leverages decades of technology development investments made by NASA and its aerospace contractor community.

Propulsion Risk Reduction Activities for Non-Toxic Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Fisher, Kenneth

2010-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project s primary objective is to develop propulsion system technologies for non-toxic or "green" propellants. The PCAD project focuses on the development of non-toxic propulsion technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of non-toxic propellants for space missions. Implementation of non-toxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate that non-toxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be presented.

Propulsion Risk Reduction Activities for Nontoxic Cryogenic Propulsion

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Klem, Mark D.; Fisher, Kenneth L.

2010-01-01

The Propulsion and Cryogenics Advanced Development (PCAD) Project s primary objective is to develop propulsion system technologies for nontoxic or "green" propellants. The PCAD project focuses on the development of nontoxic propulsion technologies needed to provide necessary data and relevant experience to support informed decisions on implementation of nontoxic propellants for space missions. Implementation of nontoxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate that nontoxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be presented.

Solar System Exploration Augmented by Lunar and Outer Planet Resource Utilization: Historical Perspectives and Future Possibilities

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2014-01-01

Establishing a lunar presence and creating an industrial capability on the Moon may lead to important new discoveries for all of human kind. Historical studies of lunar exploration, in-situ resource utilization (ISRU) and industrialization all point to the vast resources on the Moon and its links to future human and robotic exploration. In the historical work, a broad range of technological innovations are described and analyzed. These studies depict program planning for future human missions throughout the solar system, lunar launched nuclear rockets, and future human settlements on the Moon, respectively. Updated analyses based on the visions presented are presented. While advanced propulsion systems were proposed in these historical studies, further investigation of nuclear options using high power nuclear thermal propulsion, nuclear surface power, as well as advanced chemical propulsion can significantly enhance these scenarios. Robotic and human outer planet exploration options are described in many detailed and extensive studies. Nuclear propulsion options for fast trips to the outer planets are discussed. To refuel such vehicles, atmospheric mining in the outer solar system has also been investigated as a means of fuel production for high energy propulsion and power. Fusion fuels such as Helium 3 (3He) and hydrogen can be wrested from the atmospheres of Uranus and Neptune and either returned to Earth or used in-situ for energy production. Helium 3 and hydrogen (deuterium, etc.) were the primary gases of interest with hydrogen being the primary propellant for nuclear thermal solid core and gas core rocket-based atmospheric flight. A series of analyses have investigated resource capturing aspects of atmospheric mining in the outer solar system. These analyses included the gas capturing rate, storage options, and different methods of direct use of the captured gases. While capturing 3He, large amounts of hydrogen and 4He are produced. With these two additional gases, the potential for fueling small and large fleets of additional exploration and exploitation vehicles exists.

Solar System Exploration Augmented by Lunar and Outer Planet Resource Utilization: Historical Perspectives and Future Possibilities

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2014-01-01

Establishing a lunar presence and creating an industrial capability on the Moon may lead to important new discoveries for all of human kind. Historical studies of lunar exploration, in-situ resource utilization (ISRU) and industrialization all point to the vast resources on the Moon and its links to future human and robotic exploration. In the historical work, a broad range of technological innovations are described and analyzed. These studies depict program planning for future human missions throughout the solar system, lunar launched nuclear rockets, and future human settlements on the Moon, respectively. Updated analyses based on the visions presented are presented. While advanced propulsion systems were proposed in these historical studies, further investigation of nuclear options using high power nuclear thermal propulsion, nuclear surface power, as well as advanced chemical propulsion can significantly enhance these scenarios. Robotic and human outer planet exploration options are described in many detailed and extensive studies. Nuclear propulsion options for fast trips to the outer planets are discussed. To refuel such vehicles, atmospheric mining in the outer solar system has also been investigated as a means of fuel production for high energy propulsion and power. Fusion fuels such as helium 3 (3He) and hydrogen (H2) can be wrested from the atmospheres of Uranus and Neptune and either returned to Earth or used in-situ for energy production. Helium 3 and H2 (deuterium, etc.) were the primary gases of interest with hydrogen being the primary propellant for nuclear thermal solid core and gas core rocket-based atmospheric flight. A series of analyses have investigated resource capturing aspects of atmospheric mining in the outer solar system. These analyses included the gas capturing rate, storage options, and different methods of direct use of the captured gases. While capturing 3He, large amounts of hydrogen and 4He are produced. With these two additional gases, the potential for fueling small and large fleets of additional exploration and exploitation vehicles exists.

Main propulsion system test requirements for the two-engine Shuttle-C

NASA Technical Reports Server (NTRS)

Lynn, E. E.; Platt, G. K.

1989-01-01

The Shuttle-C is an unmanned cargo carrying derivative of the space shuttle with optional two or three space shuttle main engines (SSME's), whereas the shuttle has three SSME's. Design and operational differences between the Shuttle-C and shuttle were assessed to determine requirements for additional main propulsion system (MPS) verification testing. Also, reviews were made of the shuttle main propulsion test program objectives and test results and shuttle flight experience. It was concluded that, if significant MPS modifications are not made beyond those currently planned, then main propulsion system verification can be concluded with an on-pad flight readiness firing.

Definition of propulsion system for V/STOL research and technology aircraft

NASA Technical Reports Server (NTRS)

1977-01-01

Wind tunnel test support, aircraft contractor support, a propulsion system computer card deck, preliminary design studies, and propulsion system development plan are reported. The Propulsion system consists of two lift/cruise turbofan engines, one turboshaft engine and one lift fan connected together with shafting into a combiner gearbox. Distortion parameter levels from 40 x 80 test data were within the established XT701-AD-700 limits. The three engine-three fan system card deck calculates either vertical or conventional flight performance, installed or uninstalled. Design study results for XT701 engine modifications, bevel gear cross shaft location, fixed and tilt fan frames and propulsion system controls are described. Optional water-alcohol injection increased total net thrust 10.3% on a 90 F day. Engines have sufficient turbine life for 500 hours of the RTA duty cycle.

Pressurization, Pneumatic, and Vent Subsystems of the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

Hedayat, A.; Steadman, T. E.; Brown, T. M.; Knight, K. C.; White, C. E., Jr.; Champion, R. H., Jr.

1998-01-01

In pressurization systems, regulators and orifices are use to control the flow of the pressurant. For the X-34 Main Propulsion System, three pressurization subsystem design configuration options were considered. In the first option, regulators were used while in the other options, orifices were considered. In each design option, the vent/relief system must be capable of relieving the pressurant flow without allowing the tank pressure to rise above proof, therefore, impacts on the propellant tank vent system were investigated and a trade study of the pressurization system was conducted. The analysis indicated that design option using regulators poses least risk. Then, a detailed transient thermal/fluid analysis of the recommended pressurization system was performed. Helium usage, thermodynamic conditions, and overpressurization of each propellant tank were evaluated. The pneumatic and purge subsystem is used for pneumatic valve actuation, Inter-Propellant Seal purges, Engine Spin Start, and engine purges at the required interface pressures, A transient analysis of the pneumatic and purge subsystem provided helium usage and flow rates to Inter-Propellant Seal and engine interfaces. Fill analysis of the helium bottles of pressurization and pneumatic subsystems during ground operation was performed. The required fill time and the stored

An integral nuclear power and propulsion system concept

NASA Astrophysics Data System (ADS)

Choong, Phillip T.; Teofilo, Vincent L.; Begg, Lester L.; Dunn, Charles; Otting, William

An integral space power concept provides both the electrical power and propulsion from a common heat source and offers superior performance capabilities over conventional orbital insertion using chemical propulsion systems. This paper describes a hybrid (bimodal) system concept based on a proven, inherently safe solid fuel form for the high temperature reactor core operation and rugged planar thermionic energy converter for long-life steady state electric power production combined with NERVA-based rocket technology for propulsion. The integral system is capable of long-life power operation and multiple propulsion operations. At an optimal thrust level, the integral system can maintain the minimal delta-V requirement while minimizing the orbital transfer time. A trade study comparing the overall benefits in placing large payloads to GEO with the nuclear electric propulsion option shows superiority of nuclear thermal propulsion. The resulting savings in orbital transfer time and the substantial reduction of overall lift requirement enables the use of low-cost launchers for several near-term military satellite missions.

The PEGASUS Drive: A nuclear electric propulsion system for the space exploration initiative

NASA Astrophysics Data System (ADS)

Coomes, Edmund P.; Dagle, Jeffery E.

1991-01-01

The advantages of using electric propulsion for propulsion are well-known in the aerospace community. The high specific impulse, lower propellant requirements, and lower system mass make it a very attractive propulsion option for the Space Exploration Initiative (SEI), especially for the transport of cargo. One such propulsion system is the PEGASUS Drive (Coomes et al. 1987). In its original configuration, the PEGASUS Drive consisted of a 10-MWe power source coupled to a 6-MW magnetoplasmadynamic (MPD) thruster system. The PEGASUS Drive propelled a manned vechicle to Mars and back in 601 days. By removing the crew and their associated support systems from the space craft and by incorporating technology advances in reactor design and heat rejection systems, a second generation PEGASUS Drive can be developed with an alpha less than two. Utilizing this propulsion system, a 400-MT cargo vechicle, assembled and loaded in low Earth orbit (LEO), could deliver 262 MT of supplies and hardware to MARS 282 days after escaping Earth orbit. Upon arrival at Mars the transport vehicle would place its cargo in the desired parking orbit around Mars and then proceed to synchronous orbit above the desired landing sight. Using a laser transmitter, PEGASUS could provide 2-MW on the surface to operate automated systems deployed earlier and then provide surface power to support crew activities after their arrival. The additional supplies and hardware, coupled with the availability of megawatt levels of electric power on the Mars surface, would greatly enhance and even expand the mission options being considered under SEI.

Chemical Microthruster Options

NASA Technical Reports Server (NTRS)

DeGroot, Wim; Oleson, Steve

1996-01-01

Chemical propulsion systems with potential application to microsatellites are classified by propellant phase, i.e. gas, liquid, or solid. Four promising concepts are selected based on performance, weight, size, cost, and reliability. The selected concepts, in varying stages of development, are advanced monopropellants, tridyne(TM), electrolysis, and solid gas generator propulsion. Tridyne(TM) and electrolysis propulsion are compared vs. existing cold gas and monopropellant systems for selected microsatellite missions. Electrolysis is shown to provide a significant weight advantage over monopropellant propulsion for an orbit transfer and plane change mission. Tridyne(TM) is shown to provide a significant advantage over cold gas thrusters for orbit trimming and spacecraft separation.

Space transfer concepts and analysis for exploration missions. Implementation plan and element description document (draft final). Volume 4: Solar electric propulsion vehicle

NASA Technical Reports Server (NTRS)

1991-01-01

This document presents the solar electric propulsion (SEP) concept design developed as part of the Space Transfer Concepts and Analysis for Exploration Missions (STCAEM) study. The evolution of the SEP concept is described along with the requirements, guidelines and assumptions for the design. Operating modes and options are defined and a systems description of the vehicle is presented. Artificial gravity configuration options and space and ground support systems are discussed. Finally, an implementation plan is presented which addresses technology needs, schedules, facilities, and costs.

Space transfer concepts and analysis for exploration missions. Implementation plan and element description document (draft final). Volume 5: Nuclear electric propulsion vehicle

NASA Technical Reports Server (NTRS)

1991-01-01

The nuclear electric propulsion (NEP) concept design developed in support of the Space Transfer Concepts and Analysis for Exploration Missions (STCAEM) study is presented. The evolution of the NEP concept is described along with the requirements, guidelines, and assumptions for the design. Operating modes and options are defined and a systems description of the vehicle is presented. Artificial gravity configuration options and space and ground support systems are discussed. Finally, an implementation plan is presented which addresses technology needs, schedules, facilities and costs.

Low-thrust chemical orbit transfer propulsion

NASA Technical Reports Server (NTRS)

Pelouch, J. J., Jr.

1979-01-01

The need for large structures in high orbit is reported in terms of the many mission opportunities which require such structures. Mission and transportation options for large structures are presented, and it is shown that low-thrust propulsion is an enabling requirement for some missions and greatly enhancing to many others. Electric and low-thrust chemical propulsion are compared, and the need for an requirements of low-thrust chemical propulsion are discussed in terms of the interactions that are perceived to exist between the propulsion system and the large structure.

Small Satellite Propulsion Options

NASA Technical Reports Server (NTRS)

Myers, Roger M.; Oleson, Steven R.; Curran, Francis M.; Schneider, Steven J.

1994-01-01

Advanced chemical and low power electric propulsion offer attractive options for small satellite propulsion. Applications include orbit raising, orbit maintenance, attitude control, repositioning, and deorbit of both Earth-space and planetary spacecraft. Potential propulsion technologies for these functions include high pressure Ir/Re bipropellant engines, very low power arcjets, Hall thrusters, and pulsed plasma thrusters, all of which have been shown to operate in manners consistent with currently planned small satellites. Mission analyses show that insertion of advanced propulsion technologies enables and/or greatly enhances many planned small satellite missions. Examples of commercial, DoD, and NASA missions are provided to illustrate the potential benefits of using advanced propulsion options on small satellites.

Review of ESA Experimental Research Activities for Electric Propulsion

DTIC Science & Technology

2011-01-01

detect gravitational waves, distortions of space-time occurring when a massive body is accelerated or disturbed. To achieve that goal the relative...thrusters of Electric Propulsion systems accelerate the propellant ions to velocities of tens of kilometers per second making it a propulsion option that is...expanded through nozzle Ion electrostatically accelerated . Plasma accelerated via interaction of current and magnetic field. Concept Resistojets

Foundational Methane Propulsion Related Technology Efforts, and Challenges for Applications to Human Exploration Beyond Earth Orbit

NASA Technical Reports Server (NTRS)

Brown, Thomas; Klem, Mark; McRight, Patrick

2016-01-01

Current interest in human exploration beyond earth orbit is driving requirements for high performance, long duration space transportation capabilities. Continued advancement in photovoltaic power systems and investments in high performance electric propulsion promise to enable solar electric options for cargo delivery and pre-deployment of operational architecture elements. However, higher thrust options are required for human in-space transportation as well as planetary descent and ascent functions. While high thrust requirements for interplanetary transportation may be provided by chemical or nuclear thermal propulsion systems, planetary descent and ascent systems are limited to chemical solutions due to their higher thrust to weight and potential planetary protection concerns. Liquid hydrogen fueled systems provide high specific impulse, but pose challenges due to low propellant density and the thermal issues of long term propellant storage. Liquid methane fueled propulsion is a promising compromise with lower specific impulse, higher bulk propellant density and compatibility with proposed in-situ propellant production concepts. Additionally, some architecture studies have identified the potential for commonality between interplanetary and descent/ascent propulsion solutions using liquid methane (LCH4) and liquid oxygen (LOX) propellants. These commonalities may lead to reduced overall development costs and more affordable exploration architectures. With this increased interest, it is critical to understand the current state of LOX/LCH4 propulsion technology and the remaining challenges to its application to beyond earth orbit human exploration. This paper provides a survey of NASA's past and current methane propulsion related technology efforts, assesses the accomplishments to date, and examines the remaining risks associated with full scale development.

Advanced Space Fission Propulsion Systems

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Borowski, Stanley K.

2010-01-01

Fission has been considered for in-space propulsion since the 1940s. Nuclear Thermal Propulsion (NTP) systems underwent extensive development from 1955-1973, completing 20 full power ground tests and achieving specific impulses nearly twice that of the best chemical propulsion systems. Space fission power systems (which may eventually enable Nuclear Electric Propulsion) have been flown in space by both the United States and the Former Soviet Union. Fission is the most developed and understood of the nuclear propulsion options (e.g. fission, fusion, antimatter, etc.), and fission has enjoyed tremendous terrestrial success for nearly 7 decades. Current space nuclear research and technology efforts are focused on devising and developing first generation systems that are safe, reliable and affordable. For propulsion, the focus is on nuclear thermal rockets that build on technologies and systems developed and tested under the Rover/NERVA and related programs from the Apollo era. NTP Affordability is achieved through use of previously developed fuels and materials, modern analytical techniques and test strategies, and development of a small engine for ground and flight technology demonstration. Initial NTP systems will be capable of achieving an Isp of 900 s at a relatively high thrust-to-weight ratio. The development and use of first generation space fission power and propulsion systems will provide new, game changing capabilities for NASA. In addition, development and use of these systems will provide the foundation for developing extremely advanced power and propulsion systems capable of routinely and affordably accessing any point in the solar system. The energy density of fissile fuel (8 x 10(exp 13) Joules/kg) is more than adequate for enabling extensive exploration and utilization of the solar system. For space fission propulsion systems, the key is converting the virtually unlimited energy of fission into thrust at the desired specific impulse and thrust-to-weight ratio. This presentation will discuss potential space fission propulsion options ranging from first generation systems to highly advanced systems. Ongoing research that shows promise for enabling second generation NTP systems with Isp greater than 1000 s will be discussed, as will the potential for liquid, gas, or plasma core systems. Space fission propulsion systems could also be used in conjunction with simple (water-based) propellant depots to enable routine, affordable missions to various destinations (e.g. moon, Mars, asteroids) once in-space infrastructure is sufficiently developed. As fuel and material technologies advance, very high performance Nuclear Electric Propulsion (NEP) systems may also become viable. These systems could enable sophisticated science missions, highly efficient cargo delivery, and human missions to numerous destinations. Commonalities between NTP, fission power systems, and NEP will be discussed.

Electric Propulsion Performance from Geo-transfer to Geosynchronous Orbits

NASA Technical Reports Server (NTRS)

Dankanich, John W.; Carpenter, Christian B.

2007-01-01

For near-Earth application, solar electric propulsion advocates have focused on Low Earth Orbit (LEO) to Geosynchronous (GEO) low-thrust transfers because of the significant improvement in capability over chemical alternatives. While the performance gain attained from starting with a lower orbit is large, there are also increased transfer times and radiation exposure risk that has hindered the commercial advocacy for electric propulsion stages. An incremental step towards electric propulsion stages is the use of integrated solar electric propulsion systems (SEPS) for GTO to GEO transfer. Thorough analyses of electric propulsion systems options and performance are presented. Results are based on existing or near-term capabilities of Arcjets, Hall thrusters, and Gridded Ion engines. Parametric analyses based on "rubber" thruster and launch site metrics are also provided.

NASA electrothermal auxiliary propulsion technology

NASA Technical Reports Server (NTRS)

Stone, J. R.

1986-01-01

Electrothermal auxiliary propulsion systems provide high performance options which can have major mission benefits. There are several electrothermal concepts which offer a range of characteristics and benefits. Resistojets are the highest thrust to power option and are currently operational at mission average values of specific impulse, I sub sp approximately 295 sec. Long life, multipropellant resistojets are being developed for the space station, and resistojet technology advancements are being pursued to improve the I sub sp by more than 20 percent for resistojets used in satellite applications. Direct current arcjets have the potential of I sub sp over 400 sec with storable propellants and should provide over 1000 sec with hydrogen. Advanced concepts are being investigated to provide high power density options and possible growth to primary propulsion applications. Broad based experimental and analytical research and technology programs of NASA are summarized and recent significant advances are reviewed.

Mission and Design Sensitivities for Human Mars Landers Using Hypersonic Inflatable Aerodynamic Decelerators

NASA Technical Reports Server (NTRS)

Polsgrove, Tara P.; Thomas, Herbert D.; Dwyer Ciancio, Alicia; Collins, Tim; Samareh, Jamshid

2017-01-01

Landing humans on Mars is one of NASA's long term goals. NASA's Evolvable Mars Campaign (EMC) is focused on evaluating architectural trade options to define the capabilities and elements needed to sustain human presence on the surface of Mars. The EMC study teams have considered a variety of in-space propulsion options and surface mission options. Understanding how these choices affect the performance of the lander will allow a balanced optimization of this complex system of systems problem. This paper presents the effects of mission and vehicle design options on lander mass and performance. Beginning with Earth launch, options include fairing size assumptions, co-manifesting elements with the lander, and Earth-Moon vicinity operations. Capturing into Mars orbit using either aerocapture or propulsive capture is assessed. For entry, descent, and landing both storable as well as oxygen and methane propellant combinations are considered, engine thrust level is assessed, and sensitivity to landed payload mass is presented. This paper focuses on lander designs using the Hypersonic Inflatable Aerodynamic Decelerators, one of several entry system technologies currently considered for human missions.

Megawatt level electric propulsion perspectives

NASA Technical Reports Server (NTRS)

Jahn, Robert G.; Kelly, Arnold J.

1987-01-01

For long range space missions, deliverable payload fraction is an inverse exponential function of the propellant exhaust velocity or specific impulse of the propulsion system. The exhaust velocity of chemical systems are limited by their combustion chemistry and heat transfer to a few km/s. Nuclear rockets may achieve double this range, but are still heat transfer limited and ponderous to develop. Various electric propulsion systems can achieve exhaust velocities in the 10 km/s range, at considerably lower thrust densities, but require an external electrical power source. A general overview is provided of the currently available electric propulsion systems from the perspective of their characteristics as a terminal load for space nuclear systems. A summary of the available electric propulsion options is shown and generally characterized in the power vs. exhaust velocity plot. There are 3 general classes of electric thruster devices: neutral gas heaters, plasma devices, and space charge limited electrostatic or ion thrusters.

Planetary mission applications for space storable propulsion

NASA Technical Reports Server (NTRS)

Chase, R. L.; Cork, M. J.; Young, D. L.

1974-01-01

This paper presents the results of a study to compare space-storable with earth-storable spacecraft propulsion systems, space-storable with solid kick stages, and several space-storable development options on the basis of benefits received for cost expenditures required. The results show that, for a launch vehicle with performance less than that of Shuttle/Centaur, space-storable spacecraft propulsion offers an incremental benefit/cost ratio between 1.0 and 5.5 when compared to earth-storable systems for three of the four missions considered. In the case of VOIR 83, positive benefits were apparent only for a specific launch vehicle-spacecraft propulsion combination. A space-storable propulsion system operating at thrust of 600 lbf, 355 units of specific impulse, and with blowdown pressurization, represents the best choice for the JO 81 mission on a Titan/Centaur if only spacecraft propulsion modifications are considered. For still higher performance, a new solid-propellant kick stage with space-storable spacecraft propulsion is preferred over a system which uses space-storable propellants for both the kick stage and the spacecraft system.

Space transportation systems, launch systems, and propulsion for the Space Exploration Initiative: Results from Project Outreach

NASA Technical Reports Server (NTRS)

Garber, T.; Hiland, J.; Orletsky, D.; Augenstein, B.; Miller, M.

1991-01-01

A number of transportation and propulsion options for Mars exploration missions are analyzed. As part of Project Outreach, RAND received and evaluated 350 submissions in the launch vehicle, space transportation, and propulsion areas. After screening submissions, aggregating those that proposed identical or nearly identical concepts, and eliminating from further consideration those that violated known physical princples, we had reduced the total number of viable submissions to 213. In order to avoid comparing such disparate things as launch vehicles and electric propulsion systems, six broad technical areas were selected to categorize the submissions: space transportation systems; earth-to-orbit (ETO) launch systems; chemical propulsion; nuclear propulsion; low-thrust propulsion; and other. To provide an appropriate background for analyzing the submissions, an extensive survey was made of the various technologies relevant to the six broad areas listed above. We discuss these technologies with the intent of providing the reader with an indication of the current state of the art, as well as the advances that might be expected within the next 10 to 20 years.

Neptune Orbiters Utilizing Solar and Radioisotope Electric Propulsion

NASA Technical Reports Server (NTRS)

Fiehler, Douglas I.; Oleson, Steven R.

2004-01-01

In certain cases, Radioisotope Electric Propulsion (REP), used in conjunction with other propulsion systems, could be used to reduce the trip times for outer planetary orbiter spacecraft. It also has the potential to improve the maneuverability and power capabilities of the spacecraft when the target body is reached as compared with non-electric propulsion spacecraft. Current missions under study baseline aerocapture systems to capture into a science orbit after a Solar Electric Propulsion (SEP) stage is jettisoned. Other options under study would use all REP transfers with small payloads. Compared to the SEP stage/Aerocapture scenario, adding REP to the science spacecraft as well as a chemical capture system can replace the aerocapture system but with a trip time penalty. Eliminating both the SEP stage and the aerocapture system and utilizing a slightly larger launch vehicle, Star 48 upper stage, and a combined REP/Chemical capture system, the trip time can nearly be matched while providing over a kilowatt of science power reused from the REP maneuver. A Neptune Orbiter mission is examined utilizing single propulsion systems and combinations of SEP, REP, and chemical systems to compare concepts.

Photochemically Etched Construction Technology Developed for Digital Xenon Feed Systems

NASA Technical Reports Server (NTRS)

Otsap, Ben; Cardin, Joseph; Verhey, Timothy R.; Rawlin, Vincent K.; Mueller, Juergen; Aadlund, Randall; Kay, Robert; Andrews, Michael

2005-01-01

Electric propulsion systems are quickly emerging as attractive options for primary propulsion in low Earth orbit, in geosynchronous orbit, and on interplanetary spacecraft. The driving force behind the acceptance of these systems is the substantial reduction in the propellant mass that can be realized. Unfortunately, system designers are often forced to utilize components designed for chemical propellants in their electric systems. Although functionally acceptable, these relatively large, heavy components are designed for the higher pressures and mass flow rates required by chemical systems. To fully realize the benefits of electric propulsion, researchers must develop components that are optimized for the low flow rates, critical leakage needs, low pressures, and limited budgets of these emerging systems.

Spacecraft Impacts with Advanced Power and Electric Propulsion

NASA Technical Reports Server (NTRS)

Mason, Lee S.; Oleson, Steven R.

2000-01-01

A study was performed to assess the benefits of advanced power and electric propulsion systems for various space missions. Advanced power technologies that were considered included multiband gap and thin-film solar arrays, lithium batteries, and flywheels. Electric propulsion options included Hall effect thrusters and Ion thrusters. Several mission case studies were selected as representative of future applications for advanced power and propulsion systems. These included a low altitude Earth science satellite, a LEO communications constellation, a GEO military surveillance satellite, and a Mercury planetary mission. The study process entailed identification of overall mission performance using state-of-the-art power and propulsion technology, enhancements made possible with either power or electric propulsion advances individually, and the collective benefits realized when advanced power and electric propulsion are combined. Impacts to the overall spacecraft included increased payload, longer operational life, expanded operations and launch vehicle class step-downs.

National Institute for Rocket Propulsion Systems 1st Annual Workshop

NASA Technical Reports Server (NTRS)

Doreswamy, Rajiv; Fry, Emma; Swindell, Tina

2012-01-01

The National Institute for Rocket Propulsion Systems (NIRPS) is a Government -wide initiative that seeks to ensure the resiliency of the Nation fs rocket propulsion community in order for the enterprise to remain vibrant and capable of providing reliable and affordable propulsion systems for the nation fs defense, civil and commercial needs. Recognizing that rocket propulsion is a multi-use technology that ensures the nation fs leadership in aerospace, the Government has a vested interest in maintaining this strategic capability through coordinated and synchronized acquisition programs and continual investments in research and development. NIRPS is a resource for collaboration and integration between all sectors of the U.S. propulsion enterprise, supporting policy development options, identifying technology requirements, and offering solutions that maximize national resources while ensuring that capability exists to meet future demand. NIRPS functions as a multi ]agency organization that our nation fs decision makers can look to for comprehensive information regarding all issues concerning the propulsion enterprise.

In-Space Chemical Propulsion System Model

NASA Technical Reports Server (NTRS)

Byers, David C.; Woodcock, Gordon; Benfield, Michael P. J.

2004-01-01

Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystem. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.

In-Space Chemical Propulsion System Model

NASA Technical Reports Server (NTRS)

Byers, David C.; Woodcock, Gordon; Benfield, M. P. J.

2004-01-01

Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystems. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.

A High Power Solar Electric Propulsion - Chemical Mission for Human Exploration of Mars

NASA Technical Reports Server (NTRS)

Burke, Laura M.; Martini, Michael C.; Oleson, Steven R.

2014-01-01

Recently Solar Electric Propulsion (SEP) as a main propulsion system has been investigated as an option to support manned space missions to near-Earth destinations for the NASA Gateway spacecraft. High efficiency SEP systems are able to reduce the amount of propellant long duration chemical missions require, ultimately reducing the required mass delivered to Low Earth Orbit (LEO) by a launch vehicle. However, for long duration interplanetary Mars missions, using SEP as the sole propulsion source alone may not be feasible due to the long trip times to reach and insert into the destination orbit. By combining an SEP propulsion system with a chemical propulsion system the mission is able to utilize the high-efficiency SEP for sustained vehicle acceleration and deceleration in heliocentric space and the chemical system for orbit insertion maneuvers and trans-earth injection, eliminating the need for long duration spirals. By capturing chemically instead of with low-thrust SEP, Mars stay time increases by nearly 200 days. Additionally, the size the of chemical propulsion system can be significantly reduced from that of a standard Mars mission because the SEP system greatly decreases the Mars arrival and departure hyperbolic excess velocities (V(sub infinity)).

Mercury orbiter transport study

NASA Technical Reports Server (NTRS)

Friedlander, A. L.; Feingold, H.

1977-01-01

A data base and comparative performance analyses of alternative flight mode options for delivering a range of payload masses to Mercury orbit are provided. Launch opportunities over the period 1980-2000 are considered. Extensive data trades are developed for the ballistic flight mode option utilizing one or more swingbys of Venus. Advanced transport options studied include solar electric propulsion and solar sailing. Results show the significant performance tradeoffs among such key parameters as trip time, payload mass, propulsion system mass, orbit size, launch year sensitivity and relative cost-effectiveness. Handbook-type presentation formats, particularly in the case of ballistic mode data, provide planetary program planners with an easily used source of reference information essential in the preliminary steps of mission selection and planning.

Electric propulsion options for the SP-100 reference mission

NASA Technical Reports Server (NTRS)

Hardy, T. L.; Rawlin, V. K.; Patterson, M. J.

1987-01-01

Analyses were performed to characterize and compare electric propulsion systems for use on a space flight demonstration of the SP-100 nuclear power system. The component masses of resistojet, arcjet, and ion thruster systems were calculated using consistent assumptions and the maximum total impulse, velocity increment, and thrusting time were determined, subject to the constraint of the lift capability of a single Space Shuttle launch. From the study it was found that for most systems the propulsion system dry mass was less than 20 percent of the available mass for the propulsion system. The maximum velocity increment was found to be up to 2890 m/sec for resistojet, 3760 m/sec for arcjet, and 23 000 m/sec for ion thruster systems. The maximum thruster time was found to be 19, 47, and 853 days for resistojet, arcjet, and ion thruster systems, respectively.

Human Mars Mission: SEP Architecture Crew Taxi Propulsion Stage Study and Design and Technology for Reaction and Control System. Part 1

NASA Technical Reports Server (NTRS)

Young, Archie

1999-01-01

The Mars exploration is a candidate pathway to expand human presence and useful activities in the solar system. There are several propulsion system options being considered to place the Mars payload on its inter-planetary transfer trajectory. One propulsion option is the use of Solar Electric Propulsion (SEP) to spiral out with the Mars payload from an initial Low Earth Orbit (LEO) to an elliptical High Earth Orbit (HEO). This report, presented in annotated facing page format, describes the work completed on the design of a crew taxi propulsion stage used in conjunction with the SEP. Transportation system/mission analysis topics covered in this report include sub-system analysis, trajectory profile description, mass performance and crew taxi stage sizing, stage configuration, stage cost, and Trans-Mars Injection (TMI) launch window. The high efficiency of SEP is used to provide the major part of the TMI propulsion maneuver. Orbital energy is continuously added over a period of approximately twelve months. The SEP and Mars payload follow a spiral trajectory from an initial LEO to a final elliptical HEO. A small chemical stage is then used to provide the final part of the TMI. The now unloaded SEP returns to LEO to repeat another spiral trajectory with payload to HEO. The spiral phase of the SEP's trajectory takes several months to reach HEO, thus significantly increasing the exposure time of the crew to zero-gravity. In order to minimize the long zero-gravity effects, a high thrust chemical stage delivers the crew to the SEP's HEO. The crew rendezvous with the Mars payload in HEO. After a checkout period the Mars payload with the crew is injected onto a Trans-Mars Trajectory by a small chemical stage.

Human Mars Mission: SEP Architecture, Crew Taxi Propulsion Stage Study and Design and Technology for Reaction and Control System. Pt. 1

NASA Technical Reports Server (NTRS)

Young, Archie

1999-01-01

The Mars exploration is a candidate pathway to expand human presence and useful activities in the solar system. There are several propulsion system options being considered to place the Mars payload on its interplanetary transfer trajectory. One propulsion option is the use of Solar Electric Propulsion (SEP) to spiral out with the Mars payload from an initial Low Earth Orbit (LEO) to an elliptical High Earth Orbit (HEO). This report, presented in annotated facing page format, describes the work completed on the design of a crew taxi propulsion stage used in conjunction with the SEP. Transportation system/mission analysis topics covered in this report include sub-system analysis, trajectory profile description, mass performance and crew taxi stage sizing, stage configuration, stage cost, and Trans-Mars Injection (TMI) launch window. The high efficiency of SEP is used to provide the major part of the TMI propulsion maneuver. Orbital energy is continuously added over a period of approximately twelve months. The SEP and Mars payload follow a spiral trajectory from an initial LEO to a final elliptical HEO. A small chemical stage is then used to provide the final part of the TMI. The now unloaded SEP returns to LEO to repeat another spiral trajectory with payload to HEO. The spiral phase of the SEP's trajectory takes several months to reach HEO, thus significantly increasing the exposure time of the crew to zero-gravity. In order to minimize the long zero-gravity effects, a high thrust chemical stage delivers the crew to the SEP's HEO. The crew rendezvous with the Mars payload in HEO. After a checkout period the Mars payload with the crew is injected onto a Trans-Mars Trajectory by a small chemical stage.

Solid propulsion advanced concepts

NASA Technical Reports Server (NTRS)

Nakamura, Y.; Shafer, J. I.

1972-01-01

The feasibility and application of a solid propulsion powered spacecraft concept to implement high energy missions independent of multiplanetary swingby opportunities are assessed and recommendations offered for future work. An upper stage, solid propulsion launch vehicle augmentation system was selected as the baseline configuration in view of the established program goals of low cost and high reliability. Spacecraft and propulsion system data that characterize mission performance capabilities were generated to serve as the basis for subsequent tradeoff studies. A cost effectiveness model was used for the preliminary feasibility assessment to provide a meaningful comparative effectiveness measure of the various candidate designs. The results substantiated the feasibility of the powered spacecraft concept when used in conjunction with several intermediate-sized launch vehicles as well as the existence of energy margins by which to exploit the attainment of extended mission capabilities. Additionally, in growth option applications, the employment of advanced propulsion systems and alternate spacecraft approaches appear promising.

Comparison of Propulsion Options for Human Exploration of Mars

NASA Technical Reports Server (NTRS)

Drake, Bret G.; McGuire, Melissa L.; McCarty, Steven L.

2018-01-01

NASA continues to advance plans to extend human presence beyond low-Earth orbit leading to human exploration of Mars. The plans being laid out follow an incremental path, beginning with initial flight tests followed by deployment of a Deep Space Gateway (DSG) in cislunar space. This Gateway, will serve as the initial transportation node for departing and returning Mars spacecraft. Human exploration of Mars represents the next leap for humankind because it will require leaving Earth on a long mission with very limited return, rescue, or resupply capabilities. Although Mars missions are long, approaches and technologies are desired which can reduce the time that the crew is away from Earth. This paper builds off past analyses of NASA's exploration strategy by providing more detail on the performance of alternative in-space transportation options with an emphasis on reducing total mission duration. Key options discussed include advanced chemical, nuclear thermal, nuclear electric, solar electric, as well as an emerging hybrid propulsion system which utilizes a combination of both solar electric and chemical propulsion.

Power Without Wires (POWOW)

NASA Technical Reports Server (NTRS)

Brandhorst, Henry W., Jr.; Howell, Joe (Technical Monitor)

2002-01-01

Electric propulsion has emerged as a cost-effective solution to a wide range of satellite applications. Deep Space 1 successfully demonstrated electric propulsion as the primary propulsion source for a satellite. The POWOW concept is a solar-electric propelled spacecraft capable of significant cargo and short trip times for traveling to Mars. There it would enter areosynchronous orbit (Mars GEO equivalent) and beam power to surface installations via lasers. The concept has been developed with industrial partner expertise in high efficiency solar cells, advanced concentrator modules, innovative arrays, and high power electric propulsion systems. The present baseline spacecraft design providing 898 kW using technologies expected to be available in 2003 will be described. Areal power densities approaching 350 W/sq m at 80 C operating temperatures and wing level specific powers of over 350 W/kg are projected. Details of trip times and payloads to Mars are presented. Electric propulsion options include Hall, MPD, and ion thrusters of various power levels and trade studies have been conducted to define the most advantageous options. Because the design is modular, learning curve methodology has been applied to determine expected cost reductions and is included.

Propulsion Options for Primary Thrust and Attitude Control of Microspacecraft

NASA Technical Reports Server (NTRS)

deGroot, W. A.

1998-01-01

Order of magnitude decreases in the size of scientific satellites and spacecraft could provide concurrent decreases in mission costs because of lower launch and fabrication costs. Although many subsystems are amenable to dramatic size reductions, miniaturization of the propulsion subsystems is not straightforward. There are a range of requirements for both primary and attitude control propulsion, dictated by mission requirements, satellite size, and power restrictions. Many of the established propulsion technologies can not currently be applied to microspacecraft. Because of this, micro-electromechanical systems (MEMS) fabrication technology is being explored as a path for miniaturization.

Comparison of mission design options for manned Mars missions

NASA Technical Reports Server (NTRS)

Babb, Gus R.; Stump, William R.

1986-01-01

A number of manned Mars mission types, propulsion systems, and operational techniques are compared. Conjunction and opposition class missions for cryogenic, hybrid (cryo/storable), and NERVA propulsion concepts are addressed. In addition, both Earth and Mars orbit aerobraking, direct entry of landers, hyperbolic rendezvous, and electric propulsion cases are examined. A common payload to Mars was used for all cases. The basic figure of merit used was weight in low Earth orbit (LEO) at mission initiation. This is roughly proportional to launch costs.

Solar Electric and Chemical Propulsion for a Titan Mission

NASA Technical Reports Server (NTRS)

Cupples, Michael; Green, Shaun E.; Donahue, Benjamin B.; Coverstone, Victoria L.

2005-01-01

Systems analyses were performed for a Titan Explorer Mission characterized by Earth-Saturn transfer stages using solar electric power generation and propulsion systems for primary interplanetary propulsion, and chemical propulsion for capture at Titan. An examination of a range of system factors was performed to determine their effect on the payload delivery capability to Titan. The effect of varying launch vehicle type, solar array power level, ion thruster number, specific impulse, trip time, and Titan capture stage chemical propellant choice was investigated. The major purpose of the study was to demonstrate the efficacy of applying advanced ion propulsion system technologies like NASA's Evolutionary Xenon Thruster (NEXT), coupled with state-of-the-art (SOA) and advanced chemical technologies to a Flagship class mission. This study demonstrated that a NASA Design Reference Mission (DRM) payload of 406 kg could be successfully delivered to Titan using the baseline advanced ion propulsion system in conjunction with SOA chemical propulsion for Titan capture. In addition, the SEPS/Chemical system of this study is compared to an all- chemical NASA DRM mission. Results showed that the NEXT- based SEPS/Chemical system was able to deliver the required payload to Titan in 5 years less transfer time and on a smaller launch vehicle than the SOA chemical option.

Space tug point design study. Volume 3: Design definition. Part 1: Propulsion and mechanical, avionics, thermal control and electrical power subsystems

NASA Technical Reports Server (NTRS)

1973-01-01

A study was conducted to determine the configuration and performance of a space tug. Details of the space tug systems are presented to include: (1) propulsion systems, (2) avionics, (3) thermal control, and (4) electric power subsystems. The data generated include engineering drawings, schematics, subsystem operation, and component description. Various options investigated and the rational for the point design selection are analyzed.

High- and low-thrust propulsion systems for the space station

NASA Technical Reports Server (NTRS)

Jones, R. E.

1987-01-01

The purpose of the Advanced Development program was to investigate propulsion options for the space station. Two options were investigated in detail: a high-thrust system consisting of 25 to 50 lbf gaseous oxygen/hydrogen rockets, and a low-thrust system of 0.1 lbf multipropellant resistojets. An effort is also being conducted to determine the life capability of hydrazine-fueled thrusters. During the course of this program, studies clearly identified the benefits of utilizing waste water and other fluids as propellant sources. The results of the H/O thruster test programs are presented and the plan to determine the life of hydrazine thrusters is discussed. The background required to establish a long-life resistojet is presented and the first design model is shown in detail.

Trajectory Optimization of an Interstellar Mission Using Solar Electric Propulsion

NASA Technical Reports Server (NTRS)

Kluever, Craig A.

1996-01-01

This paper presents several mission designs for heliospheric boundary exploration using spacecraft with low-thrust ion engines as the primary mode of propulsion The mission design goal is to transfer a 200-kg spacecraft to the heliospheric boundary in minimum time. The mission design is a combined trajectory and propulsion system optimization problem. Trajectory design variables include launch date, launch energy, burn and coast arc switch times, thrust steering direction, and planetary flyby conditions. Propulsion system design parameters include input power and specific impulse. Both SEP and NEP spacecraft arc considered and a wide range of launch vehicle options are investigated. Numerical results are presented and comparisons with the all chemical heliospheric missions from Ref 9 are made.

The NASA In-Space Propulsion Technology Project, Products, and Mission Applicability

NASA Technical Reports Server (NTRS)

Anderson, David J.; Pencil, Eric; Liou, Larry; Dankanich, John; Munk, Michelle M.; Kremic, Tibor

2009-01-01

The In-Space Propulsion Technology (ISPT) Project, funded by NASA s Science Mission Directorate (SMD), is continuing to invest in propulsion technologies that will enable or enhance NASA robotic science missions. This overview provides development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, advanced chemical thrusters, and systems analysis tools. Aerocapture investments improved: guidance, navigation, and control models of blunt-body rigid aeroshells; atmospheric models for Earth, Titan, Mars, and Venus; and models for aerothermal effects. Investments in electric propulsion technologies focused on completing NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6 to 7 kW throttle-able gridded ion system. The project is also concluding its High Voltage Hall Accelerator (HiVHAC) mid-term product specifically designed for a low-cost electric propulsion option. The primary chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. The project is also delivering products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. In-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations.

NASA's In-Space Propulsion Technology Project Overview, Near-term Products and Mission Applicability

NASA Technical Reports Server (NTRS)

Dankanich, John; Anderson, David J.

2008-01-01

The In-Space Propulsion Technology (ISPT) Project, funded by NASA's Science Mission Directorate (SMD), is continuing to invest in propulsion technologies that will enable or enhance NASA robotic science missions. This overview provides development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, advanced chemical thrusters, and systems analysis tools. Aerocapture investments improved (1) guidance, navigation, and control models of blunt-body rigid aeroshells, 2) atmospheric models for Earth, Titan, Mars and Venus, and 3) models for aerothermal effects. Investments in electric propulsion technologies focused on completing NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system. The project is also concluding its High Voltage Hall Accelerator (HiVHAC) mid-term product specifically designed for a low-cost electric propulsion option. The primary chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. The project is also delivering products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. In-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations.

Nuclear powered Mars cargo transport mission utilizing advanced ion propulsion

NASA Technical Reports Server (NTRS)

Galecki, Diane L.; Patterson, Michael J.

1987-01-01

Nuclear-powered ion propulsion technology was combined with detailed trajectory analysis to determine propulsion system and trajectory options for an unmanned cargo mission to Mars in support of manned Mars missions. A total of 96 mission scenarios were identified by combining two power levels, two propellants, four values of specific impulse per propellant, three starting altitudes, and two starting velocities. Sixty of these scenarios were selected for a detailed trajectory analysis; a complete propulsion system study was then conducted for 20 of these trajectories. Trip times ranged from 344 days for a xenon propulsion system operating at 300 kW total power and starting from lunar orbit with escape velocity, to 770 days for an argon propulsion system operating at 300 kW total power and starting from nuclear start orbit with circular velocity. Trip times for the 3 MW cases studied ranged from 356 to 413 days. Payload masses ranged from 5700 to 12,300 kg for the 300 kW power level, and from 72,200 to 81,500 kg for the 3 MW power level.

Status report on nuclear electric propulsion systems

NASA Technical Reports Server (NTRS)

Stearns, J. W.

1975-01-01

Progress in nuclear electric propulsion (NEP) systems for a multipayload multimission vehicle needed in both deep-space missions and a variety of geocentric missions is reviewed. The space system power level is a function of the initial launch vehicle mass, but developments in out-of-core nuclear thermionic direct conversion have broadened design options. Cost, design, and performance parameters are compared for reusable chemical space tugs and NEP reusable space tugs. Improvements in heat pipes, ion engines, and magnetoplasmadynamic arc jet thrust subsystems are discussed.

Effluent treatment for nuclear thermal propulsion ground testing

NASA Technical Reports Server (NTRS)

Shipers, Larry R.

1993-01-01

The objectives are to define treatment functions, review concept options, discuss PIPET effluent treatment system (ETS), and outline future activities. The topics covered include the following: reactor exhaust; effluent treatment functions; effluent treatment categories; effluent treatment options; concept evaluation; PIPETS ETS envelope; PIPET effluent treatment concept; and future activities.

Upper stages utilizing electric propulsion

NASA Technical Reports Server (NTRS)

Byers, D. C.

1980-01-01

The payload characteristics of geocentric missions which utilize electron bombardment ion thruster systems are discussed. A baseline LEO to GEO orbit transfer mission was selected to describe the payload capabilities. The impacts on payloads of both mission parameters and electric propulsion technology options were evaluated. The characteristics of the electric propulsion thrust system and the power requirements were specified in order to predict payload mass. This was completed by utilizing a previously developed methodology which provides a detailed thrust system description after the final mass on orbit, the thrusting time, and the specific impulse are specified. The impact on payloads of total mass in LEO, thrusting time, propellant type, specific impulse, and power source characteristics was evaluated.

The Use of Nuclear Propulsion, Power and 'In-Situ' Resources for Routine Lunar Space Transportation and Commercial Base Development

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

2003-01-01

This viewgraph presentation illustrates possible future strategies for solar system exploration supported by Nuclear Thermal Rocket (NTR) Propulsion. Topics addressed in the presentation include: lunar mining, Liquid Oxygen (LOX) augmented NTR (LANTR), 'Shuttle-Derived' Heavy Lift Vehicle (SDHLV) options for future human Lunar missions, and lunar-produced oxygen (LUNOX).

Implementation Options for the PROPEL Electrodynamic Tether Demonstration Mission

NASA Technical Reports Server (NTRS)

Bilen, Sven G.; Johnson, C. Les; Gilchrist, Brian E.; Hoyt, Robert P.; Elder, Craig H.; Fuhrhop, Keith P.; Scadera, Michael; Stone, Nobie

2014-01-01

The PROPEL ("Propulsion using Electrodynamics") flight demonstration mission concept will demonstrate the use of an electrodynamic tether (EDT) for generating thrust, which will allow the propulsion system to overcome the limitations of the rocket equation. The mission concept has been developed by a team of government, industry, and academia partners led by NASA Marshall Space Flight Center (MSFC). PROPEL is being designed for versatility of the EDT system with multiple end users in mind and to be flexible with respect to platform. Previously, we reported on a comprehensive mission design for PROPEL with a mission duration of six months or longer with multiple mission goals including demonstration of significant boost, deboost, inclination change, and drag make-up activities. To explore a range of possible configurations, primarily driven by cost considerations, other mission concept designs have been pursued. In partnership with the NASA's Office of Chief Technologist (OCT) Game Changing Program, NASA MSFC Leadership, and the MSFC Advanced Concepts Office, a mission concept design was developed for a near-term EDT propulsion flight validation mission. The Electrodynamic Tether Propulsion Study (ETPS) defined an EDT propulsion system capable of very large delta-V for use on future missions developed by NASA, DoD, and commercial customers. To demonstrate the feasibility of an ETPS, the study focused on a space demonstration mission concept design with configuration of a pair of tethered satellite busses, one of which is the Japanese H-II Transfer Vehicle (HTV). The HTV would fly its standard ISS resupply mission. When resupply mission is complete, the ISS reconfigures and releases the HTV to perform the EDT experiment at safe orbital altitudes below the ISS. Though the focus of this particular mission concept design addresses a scenario involving the HTV or a similar vehicle, the propulsion system's capability is relevant to a number of applications, as noted above. The ETPS builds on prior work on long-life, failure-resistant, conducting tethers and includes an instrument suite with demonstrated heritage capable of performing necessary diagnostics to measure performance against predictions for a given system size (to be determined) and boost rate. Mission designs in other configurations and launch vehicle options are being developed such that the system can be demonstration should a flight opportunity be identified. We will report on past and ongoing implementation options for PROPEL.

Power Processing for a Conceptual Project Prometheus Electric Propulsion System

NASA Technical Reports Server (NTRS)

Scina, Joseph E., Jr.; Aulisio, Michael; Gerber, Scott S.; Hewitt, Frank; Miller, Leonard; Elbuluk, Malik; Pinero, Luis R. (Technical Monitor)

2005-01-01

NASA has proposed a bold mission to orbit and explore the moons of Jupiter. This mission, known as the Jupiter Icy Moons Orbiter (JIMO), would significantly increase NASA s capability to explore deep space by making use of high power electric propulsion. One electric propulsion option under study for JIMO is an ion propulsion system. An early version of an ion propulsion system was successfully used on NASA's Deep Space 1 mission. One concept for an ion thruster system capable of meeting the current JIMO mission requirement would have individual thrusters that are 16 to 25 kW each and require voltages as high as 8.0 kV. The purpose of this work is to develop power processing schemes for delivering the high voltage power to the spacecraft ion thrusters based upon a three-phase AC distribution system. In addition, a proposed DC-DC converter topology is presented for an ion thruster ancillary supply based upon a DC distribution system. All specifications discussed in this paper are for design convenience and are speculative in nature.

Green micro-resistojet research at Delft University of Technology: new options for Cubesat propulsion

NASA Astrophysics Data System (ADS)

Cervone, A.; Zandbergen, B.; Guerrieri, D. C.; De Athayde Costa e Silva, M.; Krusharev, I.; van Zeijl, H.

2017-03-01

The aerospace industry is recently expressing a growing interest in green, safe and non-toxic propellants for the propulsion systems of the new generation of space vehicles, which is especially true in the case of Cubesat micro-propulsion systems. Demanding requirements are associated to the future missions and challenges offered by this class of spacecraft, where the availability of a propulsion system might open new possibilities for a wide range of applications including orbital maintenance and transfer, formation flying and attitude control. To accomplish these requirements, Delft University of Technology is currently developing two different concepts of water-propelled micro-thrusters based on MEMS technologies: a free molecular micro-resistojet operating with sublimating solid water (ice) at low plenum gas pressure of less than 600 Pa, and a more conventional micro-resistojet operating with liquid water heated and vaporized by means of a custom designed silicon heating chamber. In this status review paper, the current design and future expected developments of the two micro-propulsion concepts is presented and discussed, together with an initial analysis of the expected performance and potential operational issues. Results of numerical simulations conducted to optimize the design of the heating and expansion slots, as well as a detailed description of the manufacturing steps for the conventional micro-resistojet concept, are presented. Some intended steps for future research activities, including options for thrust intensity and direction control, are briefly introduced.

Entry, Descent, and Landing With Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2012-01-01

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet s moons atmosphere for entry, descent, and landing can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. As part of those efforts, NASA began to conduct experiments to gather the experimental data to make informed decisions on the "best" EDL options. A model of a three engine retro-propulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. The rocket engine flow was simulated with a non-combusting flow of air.

Evolutionary use of nuclear electric propulsion

NASA Technical Reports Server (NTRS)

Hack, K. J.; George, J. A.; Riehl, J. P.; Gilland, J. H.

1990-01-01

Evolving new propulsion technologies through a rational and conscious effort to minimize development costs and program risks while maximizing the performance benefits is intuitively practical. A phased approach to the evolution of nuclear electric propulsion from use on planetary probes, to lunar cargo vehicles, and finally to manned Mars missions with a concomitant growth in technology is considered. Technology levels and system component makeup are discussed for nuclear power systems and both ion and magnetoplasmadynamic thrusters. Mission scenarios are described, which include analysis of a probe to Pluto, a lunar cargo mission, Martian split, all-up, and quick-trip mission options. Evolutionary progression of the use of NEP in such missions is discussed.

From Earth to Orbit: An assessment of transportation options

NASA Technical Reports Server (NTRS)

Gavin, Joseph G., Jr.; Blond, Edmund; Brill, Yvonne C.; Budiansky, Bernard; Cooper, Robert S.; Demisch, Wolfgang H.; Hawk, Clark W.; Kerrebrock, Jack L.; Lichtenberg, Byron K.; Mager, Artur

1992-01-01

The report assesses the requirements, benefits, technological feasibility, and roles of Earth-to-Orbit transportation systems and options that could be developed in support of future national space programs. Transportation requirements, including those for Mission-to-Planet Earth, Space Station Freedom assembly and operation, human exploration of space, space science missions, and other major civil space missions are examined. These requirements are compared with existing, planned, and potential launch capabilities, including expendable launch vehicles (ELV's), the Space Shuttle, the National Launch System (NLS), and new launch options. In addition, the report examines propulsion systems in the context of various launch vehicles. These include the Advanced Solid Rocket Motor (ASRM), the Redesigned Solid Rocket Motor (RSRM), the Solid Rocket Motor Upgrade (SRMU), the Space Shuttle Main Engine (SSME), the Space Transportation Main Engine (STME), existing expendable launch vehicle engines, and liquid-oxygen/hydrocarbon engines. Consideration is given to systems that have been proposed to accomplish the national interests in relatively cost effective ways, with the recognition that safety and reliability contribute to cost-effectiveness. Related resources, including technology, propulsion test facilities, and manufacturing capabilities are also discussed.

Mission and Design Sensitivities for Human Mars Landers Using Hypersonic Inflatable Aerodynamic Decelerators

NASA Technical Reports Server (NTRS)

Polsgrove, Tara P.; Thomas, Herbert D.; Collins, Tim; Dwyer Cianciolo, Alicia; Samareh, Jamshid

2017-01-01

Landing humans on Mars is one of NASA's long term goals. The Evolvable Mars Campaign (EMC) is focused on evaluating architectural trade options to define the capabilities and elements needed for a sustainable human presence on the surface of Mars. The EMC study teams have considered a variety of in-space propulsion options and surface mission options. As we seek to better understand how these choices affect the performance of the lander, this work informs and influences requirements for transportation systems to deliver the landers to Mars and enable these missions. This paper presents the effects of mission and vehicle design options on lander mass and performance. Beginning with Earth launch, options include fairing size assumptions, co-manifesting other elements with the lander, and Earth-Moon vicinity operations. Capturing into Mars orbit using either aerocapture or propulsive capture is assessed. For entry, descent, and landing both storable as well as oxygen and methane propellant combinations are considered, engine thrust level is assessed, and sensitivity to landed payload mass is presented. This paper focuses on lander designs using the Hypersonic Inflatable Aerodynamic Decelerators (HIAD), one of several entry system technologies currently considered for human missions.

Propulsion Technology Assessment: Science and Enabling Technologies to Explore the Interstellar Medium

NASA Technical Reports Server (NTRS)

Hopkins, Randall C.; Thomas, Herbert D.; Wiegmann, Bruce M.; Heaton, Andrew F.; Johnson, Les; Baysinger, Michael F.; Beers, Benjamin R.

2016-01-01

Led by the Keck Institute for Space Studies at the California Institute of Technology, the Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study to assess what low-thrust advanced propulsion system candidates, existing and near term, could deliver a small, Voyager-like satellite to our solar system's heliopause, approximately 100 AU from the center of the sun, within 10 years and within a 2025-2035 launch window. The advanced propulsion system trade study consisted of three candidates, including a Magnetically Shielded Miniature (MaSMi) Hall thruster, a solar sail and an electric sail. Two aerial densities, and thus characteristic accelerations, 0.426 mm/sq s and 0.664 mm/sq s were analyzed for the solar sail option in order understand the impact of near and long term development of this technology. Similarly, two characteristic accelerations, 1 mm/s2 and 2 mm/sq s, were also analyzed for the electric sail option in addition to tether quantities of 10 and 20, respectively, and individual tether length of 20 km. A second analysis was conducted to determine what existing solid rocket motor kick stage(s) would be required to provide additional thrust at various points in the trajectory, assuming an earth departure characteristic energy capability provided by a Space Launch System (SLS) Block 1B vehicle architecture carrying an 8.4 meter payload fairing. Two trajectory profiles were considered, including an escape trajectory using a Jupiter gravity assist (E-Ju), and an escape trajectory first performing a Jupiter gravity assist followed by an Oberth maneuver around the sun and an optional Saturn gravity assist (E-Ju-Su-Sa). The Oberth maneuver would need to be performed very close to the sun, wherein this study assumed a perihelion distance of approximately 11 solar radii, or 0.05 AU, away from the surface. The heat shield technology required to perform this type of ambitious maneuver was assumed to be similar to that of NASA's Solar Probe Plus mission, which is slated to launch in July 2018. With respect to a SLS Block 1B earth departure characteristic energy capability of 100 km2/sq s for the E-Ju trajectory option, results indicated that compared to having no advanced propulsion system onboard, both the MaSMi Hall thruster and solar sail options subtract approximately 8 to 10 years from the total trip time while the electric sail outperforms all options by subtracting up to 20 years. With respect to an average kick stage velocity capability of 2.5 to 3.5 km/s at perihelion, the most sensitive segment of the E-Ju-Su-Sa trajectory option, results indicated that both the MaSMi Hall thrust and solar sail options only subtract 1 to 3 years from the total trip time whereas the electric sail again outperforms all other options by subtracting up to 5 years. In other words, if the Technology Readiness Level of an electric sail could be increased in time, this propulsion technology could not only enable a satellite to reach 100 AU in 10 years but it could potentially do so even faster. Completing such an ambitious mission in that short of a timespan would be very attractive to many as it would be well within the average career span of any of those involved.

Propulsion Technology Assessment: Science & Enabling Technologies to Explore the Interstellar Medium

NASA Technical Reports Server (NTRS)

Hopkins, Randall C.; Thomas, Herbert D.; Wiegmann, Bruce M.; Heaton, Andrew F.; Johnson, Les; Baysinger, Michael F.; Beers, Benjamin R.

2015-01-01

As part of a larger effort led by the Keck Institute for Space Studies at the California Institute of Technology, the Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a study to assess what low-thrust advanced propulsion system candidates, existing and near term, could deliver a small, Voyager-like satellite to our solar system's heliopause, approximately 100 AU from the center of the sun, within 10 years and within a 2025 to 2035 launch window. The advanced propulsion system trade study consisted of three candidates, including a Magnetically Shielded Miniature (MaSMi) Hall thruster, a solar sail and an electric sail. Two aerial densities, and thus characteristic accelerations, 0.426 mm/s2 and 0.664 mm/s2, were analyzed for the solar sail option in order understand the impact of near and long term development of this technology. Similarly, two characteristic accelerations, 1 mm/s2 and 2 mm/s2, were also analyzed for the electric sail option in addition to tether quantities of 10 and 20, respectively, and individual tether length of 20 km. A second analysis was conducted to determine what existing solid rocket motor kick stage(s) would be required to provide additional thrust at various points in the trajectory, assuming an earth departure characteristic energy capability provided by a Space Launch System (SLS) Block 1B vehicle architecture carrying an 8.4 meter payload fairing. Two trajectory profiles were considered, including an escape trajectory using a Jupiter gravity assist (E-Ju), and an escape trajectory first performing a Jupiter gravity assist followed by an Oberth maneuver around the sun and an optional Saturn gravity assist (E-Ju-Su-Sa). The Oberth maneuver would need to be performed very close to the sun, wherein this study assumed a perihelion distance of approximately 11 solar radii, or 0.05 AU, away from the surface. The heat shield technology required to perform this type of ambitious maneuver was assumed to be similar to that of NASA's Solar Probe Plus mission, which is slated to launch in July 2018. With respect to a SLS Block 1B earth departure characteristic energy capability of 100 sq km/s2 for the E-Ju trajectory option, results indicated that compared to having no advanced propulsion system onboard, both the MaSMi Hall thruster and solar sail options subtract approximately 8 to 10 years from the total trip time while the electric sail outperforms all options by subtracting up to 20 years. With respect to an average kick stage velocity capability of 2.5 to 3.5 km/s at perihelion, the most sensitive segment of the E-Ju-Su-Sa trajectory option, results indicated that both the MaSMi Hall thrust and solar sail options only subtract 1 to 3 years from the total trip time whereas the electric sail again outperforms all other options by subtracting up to 5 years. In other words, if the Technology Readiness Level of an electric sail could be increased in time, this propulsion technology could not only enable a satellite to reach 100 AU in 10 years but it could potentially do so even faster. Completing such an ambitious mission in that short of a timespan would be very attractive to many as it would be well within the average career span of any of those involved.

Enhancing space transportation: The NASA program to develop electric propulsion

NASA Technical Reports Server (NTRS)

Bennett, Gary L.; Watkins, Marcus A.; Byers, David C.; Barnett, John W.

1990-01-01

The NASA Office of Aeronautics, Exploration, and Technology (OAET) supports a research and technology (R and T) program in electric propulsion to provide the basis for increased performance and life of electric thruster systems which can have a major impact on space system performance, including orbital transfer, stationkeeping, and planetary exploration. The program is oriented toward providing high-performance options that will be applicable to a broad range of near-term and far-term missions and vehicles. The program, which is being conducted through the Jet Propulsion Laboratory (JPL) and Lewis Research Center (LeRC) includes research on resistojet, arcjets, ion engines, magnetoplasmadynamic (MPD) thrusters, and electrodeless thrusters. Planning is also under way for nuclear electric propulsion (NEP) as part of the Space Exploration Initiative (SEI).

A Smoothed Eclipse Model for Solar Electric Propulsion Trajectory Optimization

NASA Technical Reports Server (NTRS)

Aziz, Jonathan D.; Scheeres, Daniel J.; Parker, Jeffrey S.; Englander, Jacob A.

2017-01-01

Solar electric propulsion (SEP) is the dominant design option for employing low-thrust propulsion on a space mission. Spacecraft solar arrays power the SEP system but are subject to blackout periods during solar eclipse conditions. Discontinuity in power available to the spacecraft must be accounted for in trajectory optimization, but gradient-based methods require a differentiable power model. This work presents a power model that smooths the eclipse transition from total eclipse to total sunlight with a logistic function. Example trajectories are computed with differential dynamic programming, a second-order gradient-based method.

Thrust augmentation options for the Beta 2 two-stage-to-orbit vehicle

NASA Technical Reports Server (NTRS)

Snyder, Christopher A.

1993-01-01

NASA LeRC is continuing to study propulsion concepts for a horizontal takeoff and landing, fully reusable, two-stage-to-orbit vehicle. This will be capable of launching and returning a 10,000 pound payload to a 100 nautical mile polar orbit using low-risk technology. The vehicle, Beta 2, is a derivative of the USAF/Boeing Beta vehicle which was designed to deliver a 50,000 pound payload to a similar orbit. Beta 2 stages at Mach 6.5 and about 100,000 ft altitude. The propulsion system for the booster is an over/under turbine bypass engine/ramjet configuration. In this paper, several options for thrust augmentation were studied in order to improve the performance of this engine where there was a critical need. Options studies were turbine engine overspeed in the transonic region, water injection at a various turbine engine locations also during the transonic region, and water injection at the turbine engine face during high speed operation. The methodology, constraints, propulsion performance, and mission study results are presented.

Systems design and analysis of the microwave radiometer spacecraft

NASA Technical Reports Server (NTRS)

Garrett, L. B.

1981-01-01

Systems design and analysis data were generated for microwave radiometer spacecraft concept using the Large Advanced Space Systems (LASS) computer aided design and analysis program. Parametric analyses were conducted for perturbations off the nominal-orbital-altitude/antenna-reflector-size and for control/propulsion system options. Optimized spacecraft mass, structural element design, and on-orbit loading data are presented. Propulsion and rigid-body control systems sensitivities to current and advanced technology are established. Spacecraft-induced and environmental effects on antenna performance (surface accuracy, defocus, and boresight off-set) are quantified and structured material frequencies and modal shapes are defined.

Evaluation of solar electric propulsion technologies for discovery class missions

NASA Technical Reports Server (NTRS)

Oh, David Y.

2005-01-01

A detailed study examines the potential benefits that advanced electric propulsion (EP) technologies offer to the cost-capped missions in NASA's Discovery program. The study looks at potential cost and performance benefits provided by three EP technologies that are currently in development: NASA's Evolutionary Xenon Thruster (NEXT), an Enhanced NSTAR system, and a Low Power Hall effect thruster. These systems are analyzed on three straw man Discovery class missions and their performance is compared to a state of the art system using the NSTAR ion thruster. An electric propulsion subsystem cost model is used to conduct a cost-benefit analysis for each option. The results show that each proposed technology offers a different degree of performance and/or cost benefit for Discovery class missions.

POwer WithOut Wire (POWOW): A SEP Concept for Space Exploration

NASA Technical Reports Server (NTRS)

Brandhorst, Henry W., Jr.; ONeill, Mark

2000-01-01

Electric propulsion has emerged as a cost-effective solution to a wide range of satellite applications. Deep Space 1 demonstrated electric propulsion as a primary propulsion source for a spacecraft. The POwer WithOut Wires (POWOW) concept has been developed as a solar electric propelled spacecraft that would travel to Mars, for example, enter selenosynchronous orbit and then use lasers to beam power to surface installations. This concept has been developed with industrial expertise in high efficiency solar cells, advanced concentrator modules, innovative arrays, and high power electric propulsion systems. The paper will present the latest version of the spacecraft, the technologies involved, possible missions and trip times to Mars and laser beaming options. The POWOW spacecraft is a general purpose solar electric propulsion system that includes technologies that are directly applicable to commercial and government spacecraft with power levels ranging from 4 kW in Low Earth Orbits (LEO) to about 1 MW. The system is modular and expandable. Learning curve costing methodologies are used to demonstrate cost effectiveness of a modular system.

Status and Mission Applicability of NASA's In-Space Propulsion Technology Project

NASA Technical Reports Server (NTRS)

Anderson, David J.; Munk, Michelle M.; Dankanich, John; Pencil, Eric; Liou, Larry

2009-01-01

The In-Space Propulsion Technology (ISPT) project develops propulsion technologies that will enable or enhance NASA robotic science missions. Since 2001, the ISPT project developed and delivered products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. These in-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations. This paper provides status of the technology development, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of advanced chemical thrusters, electric propulsion, aerocapture, and systems analysis tools. The current chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. Investments in electric propulsion technologies focused on completing NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system, and the High Voltage Hall Accelerator (HiVHAC) thruster, which is a mid-term product specifically designed for a low-cost electric propulsion option. Aerocapture investments developed a family of thermal protections system materials and structures; guidance, navigation, and control models of blunt-body rigid aeroshells; atmospheric models for Earth, Titan, Mars and Venus; and models for aerothermal effects. In 2009 ISPT started the development of propulsion technologies that would enable future sample return missions. The paper describes the ISPT project's future focus on propulsion for sample return missions. The future technology development areas for ISPT is: Planetary Ascent Vehicles (PAV), with a Mars Ascent Vehicle (MAV) being the initial development focus; multi-mission technologies for Earth Entry Vehicles (MMEEV) needed for sample return missions from many different destinations; propulsion for Earth Return Vehicles (ERV), transfer stages to the destination, and Electric Propulsion for sample return and low cost missions; and Systems/Mission Analysis focused on sample return propulsion. The ISPT project is funded by NASA's Science Mission Directorate (SMD).

Nuclear Cryogenic Propulsion Stage Affordable Development Strategy

NASA Technical Reports Server (NTRS)

Doughty, Glen E.; Gerrish, H. P.; Kenny, R. J.

2014-01-01

The development of nuclear power for space use in nuclear thermal propulsion (NTP) systems will involve significant expenditures of funds and require major technology development efforts. The development effort must be economically viable yet sufficient to validate the systems designed. Efforts are underway within the National Aeronautics and Space Administration's (NASA) Nuclear Cryogenic Propulsion Stage Project (NCPS) to study what a viable program would entail. The study will produce an integrated schedule, cost estimate and technology development plan. This will include the evaluation of various options for test facilities, types of testing and use of the engine, components, and technology developed. A "Human Rating" approach will also be developed and factored into the schedule, budget and technology development approach.

Propulsion options for the HI SPOT long endurance drone airship. Final report, November 1978-August 1979

DOE Office of Scientific and Technical Information (OSTI.GOV)

Marcy, W.L.; Hookway, R.O.

1979-09-15

Airbreathing, monofueled, stored-energy, and solar-rechargeable propulsion systems have been studied for the HI SPOT Long Endurance Drone Airship, providing constant-level electrical power as well as variable aerodynamic thrust to maintain position in winds varying from 15 to 100 knots at high altitude. A hydrogen fueled airbreathing engine is optimum for mission lengths up to 30 days or more.

Solar thermal propulsion for planetary spacecraft

NASA Technical Reports Server (NTRS)

Sercel, J. C.

1985-01-01

Previous studies have shown that many desirable planetary exploration missions require large injection delta-V. Solar Thermal Rocket (STR) propulsion, under study for orbit-raising applications may enhance or enable such high-energy missions. The required technology of thermal control for liquid hydrogen propellant is available for the required storage duration. Self-deploying, inflatable solar concentrators are under study. The mass penalty for passive cryogenic thermal control, liquid hydrogen tanks and solar concentrators does not compromise the specific impulse advantage afforded by the STR as compared to chemical propulsion systems. An STR injection module is characterized and performance is evaluated by comparison to electric propulsion options for the Saturn Orbiter Titan Probe (SOTP) and Uranus Flyby Uranus Probe (UFUP) missions.

Propulsion Technology Assessment: Science and Enabling Technologies to Explore the Interstellar Medium

NASA Technical Reports Server (NTRS)

Hopkins, Randall C.; Thomas, Herbert D.; Wiegmann, Bruce M.; Heaton, Andrew F.; Johnson, Les; Baysinger, Michael F.; Beers, Benjamin R.

2016-01-01

As part of a larger effort led by the Keck Institute for Space Studies at the California Institute of Technology, the Advanced Concepts Office at NASA’s George C. Marshall Space Flight Center conducted a study to assess what low-thrust advanced propulsion system candidates, existing and near term, could deliver a small, Voyager-like satellite to our solar system’s heliopause, approximately 100 AU from the center of the sun, within 10 years and within a 2025 to 2035 launch window. The advanced propulsion system trade study consisted of three candidates, including a Magnetically Shielded Miniature (MaSMi) Hall thruster, a solar sail and an electric sail. Two aerial densities, and thus characteristic accelerations, 0.426 mm/s(exp 2) and 0.664 mm/s(exp 2), were analyzed for the solar sail option in order understand the impact of near and long term development of this technology. Similarly, two characteristic accelerations, 1 mm/s(exp 2) and 2 mm/s(exp 2), were also analyzed for the electric sail option in addition to tether quantities of 10 and 20, respectively, and individual tether length of 20 km. A second analysis was conducted to determine what existing solid rocket motor kick stage(s) would be required to provide additional thrust at various points in the trajectory, assuming an earth departure characteristic energy capability provided by a Space Launch System (SLS) Block 1B vehicle architecture carrying an 8.4 meter payload fairing. Two trajectory profiles were considered, including an escape trajectory using a Jupiter gravity assist (E-Ju), and an escape trajectory first performing a Jupiter gravity assist followed by an Oberth maneuver around the sun and an optional Saturn gravity assist (E-Ju-Su-Sa). The Oberth maneuver would need to be performed very close to the sun, wherein this study assumed a perihelion distance of approximately 11 solar radii, or 0.05 AU, away from the surface. The heat shield technology required to perform this type of ambitious maneuver was assumed to be similar to that of NASA’s Solar Probe Plus mission, which is slated to launch in July 2018. With respect to a SLS Block 1B earth departure characteristic energy capability of 100 km(exp 2)/s(exp 2) for the E-Ju trajectory option, results indicated that compared to having no advanced propulsion system onboard, both the MaSMi Hall thruster and solar sail options subtract approximately 8 to 10 years from the total trip time while the electric sail outperforms all options by subtracting up to 20 years. With respect to an average kick stage velocity capability of 2.5 to 3.5 km/s at perihelion, the most sensitive segment of the E-Ju-Su-Sa trajectory option, results indicated that both the MaSMi Hall thrust and solar sail options only subtract 1 to 3 years from the total trip time whereas the electric sail again outperforms all other options by subtracting up to 5 years. In other words, if the Technology Readiness Level of an electric sail could be increased in time, this propulsion technology could not only enable a satellite to reach 100 AU in 10 years but it could potentially do so even faster. Completing such an ambitious mission in that short of a timespan would be very attractive to many as it would be well within the average career span of any of those involved.

Strategic Technologies for Deep Space Transport

NASA Technical Reports Server (NTRS)

Litchford, Ronald J.

2016-01-01

Deep space transportation capability for science and exploration is fundamentally limited by available propulsion technologies. Traditional chemical systems are performance plateaued and require enormous Initial Mass in Low Earth Orbit (IMLEO) whereas solar electric propulsion systems are power limited and unable to execute rapid transits. Nuclear based propulsion and alternative energetic methods, on the other hand, represent potential avenues, perhaps the only viable avenues, to high specific power space transport evincing reduced trip time, reduced IMLEO, and expanded deep space reach. Here, key deep space transport mission capability objectives are reviewed in relation to STMD technology portfolio needs, and the advanced propulsion technology solution landscape is examined including open questions, technical challenges, and developmental prospects. Options for potential future investment across the full compliment of STMD programs are presented based on an informed awareness of complimentary activities in industry, academia, OGAs, and NASA mission directorates.

Mission Design for the Innovative Interstellar Explorer Vision Mission

NASA Technical Reports Server (NTRS)

Fiehler, Douglas I.; McNutt, Ralph L.

2005-01-01

The Innovative Interstellar Explorer, studied under a NASA Vision Mission grant, examined sending a probe to a heliospheric distance of 200 Astronomical Units (AU) in a "reasonable" amount of time. Previous studies looked at the use of a near-Sun propulsive maneuver, solar sails, and fission reactor powered electric propulsion systems for propulsion. The Innovative Interstellar Explorer's mission design used a combination of a high-energy launch using current launch technology, a Jupiter gravity assist, and electric propulsion powered by advanced radioisotope power systems to reach 200 AU. Many direct and gravity assist trajectories at several power levels were considered in the development of the baseline trajectory, including single and double gravity assists utilizing the outer planets (Jupiter, Saturn, Uranus, and Neptune). A detailed spacecraft design study was completed followed by trajectory analyses to examine the performance of the spacecraft design options.

Nuclear electric propulsion options for piloted Mars missions

NASA Technical Reports Server (NTRS)

George, Jeffrey A.

1993-01-01

Three nuclear electric propulsion (NEP) systems are discussed. The three systems are as follows: a system based on current SP-100 technology; a potassium Rankine-cycle based power conversion system, and an argon ion thruster system. The system will be researched for implementation in several possible vehicle configurations. The following are among the possible Mars vehicle configurations: a piloted 15 MWe multi-reactor vehicle; a piloted 10 MWe vehicle with ECCV; a piloted 10 MWe modular vehicle; piloted 10 and 15 MWe vehicles with ECCV and MEV; a piloted 5 MWe vehicle with ECCV; a 5 MWe cargo vehicle with 2 MEV's; and a 2.5 MWe vehicle with MEV.

Low Thrust Orbital Maneuvers Using Ion Propulsion

NASA Astrophysics Data System (ADS)

Ramesh, Eric

2011-10-01

Low-thrust maneuver options, such as electric propulsion, offer specific challenges within mission-level Modeling, Simulation, and Analysis (MS&A) tools. This project seeks to transition techniques for simulating low-thrust maneuvers from detailed engineering level simulations such as AGI's Satellite ToolKit (STK) Astrogator to mission level simulations such as the System Effectiveness Analysis Simulation (SEAS). Our project goals are as follows: A) Assess different low-thrust options to achieve various orbital changes; B) Compare such approaches to more conventional, high-thrust profiles; C) Compare computational cost and accuracy of various approaches to calculate and simulate low-thrust maneuvers; D) Recommend methods for implementing low-thrust maneuvers in high-level mission simulations; E) prototype recommended solutions.

MW-Class Electric Propulsion System Designs for Mars Cargo Transport

NASA Technical Reports Server (NTRS)

Gilland, James H.; LaPointe, Michael R.; Oleson, Steven; Mercer, Carolyn; Pencil, Eric; Maosn, Lee

2011-01-01

Multi-kilowatt electric propulsion systems are well developed and have been used on commercial and military satellites in Earth orbit for several years. Ion and Hall thrusters have also propelled robotic spacecraft to encounters with asteroids, the Moon, and minor planetary bodies within the solar system. High power electric propulsion systems are currently being considered to support piloted missions to near earth asteroids, as cargo transport for sustained lunar or Mars exploration, and for very high-power piloted missions to Mars and the outer planets. Using NASA Mars Design Architecture 5.0 as a reference, a preliminary parametric analysis was performed to determine the suitability of a nuclear powered, MW-class electric propulsion system for Mars cargo transport. For this initial analysis, high power 100-kW Hall thrusters and 250-kW VASIMR engines were separately evaluated to determine optimum vehicle architecture and estimated performance. The DRA 5.0 cargo mission closed for both propulsion options, delivering a 100 t payload to Mars orbit and reducing the number of heavy lift launch vehicles from five in the baseline DRA 5.0 architecture to two using electric propulsion. Under an imposed single engine-out mission success criteria, the VASIMR system took longer to reach Mars than did the Hall system, arising from the need to operate the VASIMR thrusters in pairs during the spiral out from low Earth orbit.

LOX/hydrocarbon auxiliary propulsion system study

NASA Technical Reports Server (NTRS)

Orton, G. F.; Mark, T. D.; Weber, D. D.

1982-01-01

Liquid oxygen (LOX)/hydrocarbon propulsion concepts for a "second generation' orbiter auxiliary propulsion system was evaluated. The most attractive fuel and system design approach identified, and the technology advancements that are needed to provide high confidence for a subsequent system development were determined. The fuel candidates were ethanol, methane, propane, and ammonia. Even though ammonia is not a hydrocarbon, it was included for evaluation because it is clean burning and has a good technology base. The major system design options were pump versus pressure feed, cryogenic versus ambient temperature RCS propellant feed, and the degree of OMS-RCS integration. Ethanol was determined to be the best fuel candidate. It is an earth-storable fuel with a vapor pressure slightly higher than monomethyl hydrazine. A pump-fed OMS was recommended because of its high specific impulse, enabling greater velocity change and greater payload capability than a pressure fed system.

Investigation into Hybrid Rockets and Other Cost-Effective Propulsion System Options for Small Satellites

DTIC Science & Technology

1996-05-01

8-7 COMPLETE TEXT OF THESIS ROCKET PROPULSION FUNDEMENTALS EXPERIMENTAL DATA (MICROSOFT EXCEL FILES) 4 ANALYSIS WORKSHEETS (MATHSOFT MATHCAD FILES...up and running. At ~413,000, this represents a very small investment considering it encompasses the entire program. Similar programs run at... investment would be -needed along with over two man-years of effort. However, this is for the first flight article. Subsequent flight articles of identical

Space transfer concepts and analysis for exploration missions

NASA Technical Reports Server (NTRS)

1990-01-01

The progress and results are summarized for mission/system requirements database; mission analysis; GN and C (Guidance, Navigation, and Control), aeroheating, Mars landing; radiation protection; aerobrake mass analysis; Shuttle-Z, TMIS (Trans-Mars Injection Stage); Long Duration Habitat Trade Study; evolutionary lunar and Mars options; NTR (Nuclear Thermal Rocket); NEP (Nuclear Electric Propulsion) update; SEP (Solar Electric Propulsion) update; orbital and space-based requirements; technology; piloted rover; programmatic task; and evolutionary and innovative architecture.

Guide to Flow Measurement for Electric Propulsion Systems

NASA Technical Reports Server (NTRS)

Frieman, Jason D.; Walker, Mitchell L. R.; Snyder, Steve

2013-01-01

In electric propulsion (EP) systems, accurate measurement of the propellant mass flow rate of gas or liquid to the thruster and external cathode is a key input in the calculation of thruster efficiency and specific impulse. Although such measurements are often achieved with commercial mass flow controllers and meters integrated into propellant feed systems, the variability in potential propellant options and flow requirements amongst the spectrum of EP power regimes and devices complicates meter selection, integration, and operation. At the direction of the Committee on Standards for Electric Propulsion Testing, a guide was jointly developed by members of the electric propulsion community to establish a unified document that contains the working principles, methods of implementation and analysis, and calibration techniques and recommendations on the use of mass flow meters in laboratory and spacecraft electric propulsion systems. The guide is applicable to EP devices of all types and power levels ranging from microthrusters to high-power ion engines and Hall effect thrusters. The establishment of a community standard on mass flow metering will help ensure the selection of the proper meter for each application. It will also improve the quality of system performance estimates by providing comprehensive information on the physical phenomena and systematic errors that must be accounted for during the analysis of flow measurement data. This paper will outline the standard methods and recommended practices described in the guide titled "Flow Measurement for Electric Propulsion Systems."

Design of a Mars Airplane Propulsion System for the Aerial Regional-Scale Environmental Survey (ARES) Mission Concept

NASA Technical Reports Server (NTRS)

Kuhl, Christopher A.

2008-01-01

The Aerial Regional-Scale Environmental Survey (ARES) is a Mars exploration mission concept that utilizes a rocket propelled airplane to take scientific measurements of atmospheric, surface, and subsurface phenomena. The liquid rocket propulsion system design has matured through several design cycles and trade studies since the inception of the ARES concept in 2002. This paper describes the process of selecting a bipropellant system over other propulsion system options, and provides details on the rocket system design, thrusters, propellant tank and PMD design, propellant isolation, and flow control hardware. The paper also summarizes computer model results of thruster plume interactions and simulated flight performance. The airplane has a 6.25 m wingspan with a total wet mass of 185 kg and has to ability to fly over 600 km through the atmosphere of Mars with 45 kg of MMH / MON3 propellant.

Evaluation of High-Power Solar Electric Propulsion using Advanced Ion, Hall, MPD, and PIT Thrusters for Lunar and Mars Cargo Missions

NASA Technical Reports Server (NTRS)

Frisbee, Robert H.

2006-01-01

This paper presents the results of mission analyses that expose the advantages and disadvantages of high-power (MWe-class) Solar Electric Propulsion (SEP) for Lunar and Mars Cargo missions that would support human exploration of the Moon and Mars. In these analyses, we consider SEP systems using advanced Ion thrusters (the Xenon [Xe] propellant Herakles), Hall thrusters (the Bismuth [Bi] propellant Very High Isp Thruster with Anode Layer [VHITAL], magnetoplasmadynamic (MPD) thrusters (the Lithium [Li] propellant Advanced Lithium-Fed, Applied-field Lorentz Force Accelerator (ALFA2), and pulsed inductive thruster (PIT) (the Ammonia [NH3] propellant Nuclear-PIT [NuPIT]). The analyses include comparison of the advanced-technology propulsion systems (VHITAL, ALFA2, and NuPIT) relative to state-of-theart Ion (Herakles) propulsion systems and quantify the unique benefits of the various technology options such as high power-per-thruster (and/or high power-per-thruster packaging volume), high specific impulse (Isp), high-efficiency, and tankage mass (e.g., low tankage mass due to the high density of bismuth propellant). This work is based on similar analyses for Nuclear Electric Propulsion (NEP) systems.

TAL Performance and Mission Analysis in a CDL Capacitor Powered Direct-Drive Configuration

NASA Technical Reports Server (NTRS)

Hrbud, Ivana; Rose, M. Frank; Oleson, Steve R.; Jenkins, Rhonald M.

1999-01-01

The goals of this research are (1) to prove the concept feasibility of a direct-drive electric propulsion system, and (2) to evaluate the performance and characteristics of a Russian TAL (Thruster with Anode Layer) operating in a long-pulse mode, powered by a capacitor-based power source developed at Space Power Institute. The TAL, designated D-55, is characterized by an external acceleration zone and is powered by a unique chemical double layer (CDL) capacitor bank with a capacitance of 4 F at a charge voltage of 400 V. Performance testing of this power supply on the TAL was conducted at NASA Lewis Research Center in Cleveland, OH. Direct thrust measurements of the TAL were obtained at CDL power levels ranging from 450 to 1750 W. The specific impulse encompassed a range from 1150 s to 2200 s, yielding thruster system efficiencies between 50 and 60%. Preliminary mission analysis of the CDL direct-drive concept and other electric propulsion options was performed for the ORACLE spacecraft in 6am/6pm and 12am/12pm, 300 km sun-synchronous orbits. The direct-drive option was competitive with the other systems by increasing available net mass between 5 and 42% and reducing two-year system wet mass between 18 and 63%. Overall, the electric propulsion power requirements for the satellite solar array were reduced between 57 and 91% depending oil the orbit evaluated The direct-drive, CDL capacitor-based concept in electric propulsion thus promises to be a highly-efficient, viable alternative for satellite operations in specific near-Earth missions.

Overview of Innovative Aircraft Power and Propulsion Systems and Their Applications for Planetary Exploration

NASA Technical Reports Server (NTRS)

Colozza, Anthony; Landis, Geoffrey; Lyons, Valerie

2003-01-01

Planetary exploration may be enhanced by the use of aircraft for mobility. This paper reviews the development of aircraft for planetary exploration missions at NASA and reviews the power and propulsion options for planetary aircraft. Several advanced concepts for aircraft exploration, including the use of in situ resources, the possibility of a flexible all-solid-state aircraft, the use of entomopters on Mars, and the possibility of aerostat exploration of Titan, are presented.

Nuclear Thermal Propulsion Mars Mission Systems Analysis and Requirements Definition

NASA Technical Reports Server (NTRS)

Mulqueen, Jack; Chiroux, Robert C.; Thomas, Dan; Crane, Tracie

2007-01-01

This paper describes the Mars transportation vehicle design concepts developed by the Marshall Space Flight Center (MSFC) Advanced Concepts Office. These vehicle design concepts provide an indication of the most demanding and least demanding potential requirements for nuclear thermal propulsion systems for human Mars exploration missions from years 2025 to 2035. Vehicle concept options vary from large "all-up" vehicle configurations that would transport all of the elements for a Mars mission on one vehicle. to "split" mission vehicle configurations that would consist of separate smaller vehicles that would transport cargo elements and human crew elements to Mars separately. Parametric trades and sensitivity studies show NTP stage and engine design options that provide the best balanced set of metrics based on safety, reliability, performance, cost and mission objectives. Trade studies include the sensitivity of vehicle performance to nuclear engine characteristics such as thrust, specific impulse and nuclear reactor type. Tbe associated system requirements are aligned with the NASA Exploration Systems Mission Directorate (ESMD) Reference Mars mission as described in the Explorations Systems Architecture Study (ESAS) report. The focused trade studies include a detailed analysis of nuclear engine radiation shield requirements for human missions and analysis of nuclear thermal engine design options for the ESAS reference mission.

Mars, Phobos, and Deimos Sample Return Enabled by ARRM Alternative Trade Study Spacecraft

NASA Technical Reports Server (NTRS)

Englander, Jacob A.; Vavrina, Matthew; Merrill, Raymond G.; Qu, Min; Naasz, Bo J.

2014-01-01

The Asteroid Robotic Redirect Mission (ARRM) has been the topic of many mission design studies since 2011. The reference ARRM spacecraft uses a powerful solar electric propulsion (SEP) system and a bag device to capture a small asteroid from an Earth-like orbit and redirect it to a distant retrograde orbit (DRO) around the moon. The ARRM Option B spacecraft uses the same propulsion system and multi-Degree of Freedom (DoF) manipulators device to retrieve a very large sample (thousands of kilograms) from a 100+ meter diameter farther-away Near Earth Asteroid (NEA). This study will demonstrate that the ARRM Option B spacecraft design can also be used to return samples from Mars and its moons - either by acquiring a large rock from the surface of Phobos or Deimos, and or by rendezvousing with a sample-return spacecraft launched from the surface of Mars.

Mars, Phobos, and Deimos Sample Return Enabled by ARRM Alternative Trade Study Spacecraft

NASA Technical Reports Server (NTRS)

Englander, Jacob A.; Vavrina, Matthew; Naasz, Bo; Merill, Raymond G.; Qu, Min

2014-01-01

The Asteroid Robotic Redirect Mission (ARRM) has been the topic of many mission design studies since 2011. The reference ARRM spacecraft uses a powerful solar electric propulsion (SEP) system and a bag device to capture a small asteroid from an Earth-like orbit and redirect it to a distant retrograde orbit (DRO) around the moon. The ARRM Option B spacecraft uses the same propulsion system and multi-Degree of Freedom (DoF) manipulators device to retrieve a very large sample (thousands of kilograms) from a 100+ meter diameter farther-away Near Earth Asteroid (NEA). This study will demonstrate that the ARRM Option B spacecraft design can also be used to return samples from Mars and its moons - either by acquiring a large rock from the surface of Phobos or Deimos, and/or by rendezvousing with a sample-return spacecraft launched from the surface of Mars.

Cryogenic system options for a superconducting aircraft propulsion system

NASA Astrophysics Data System (ADS)

Berg, F.; Palmer, J.; Bertola, L.; Miller, Paul; Dodds, Graham

2015-12-01

There is a perceived need in the future for a move away from traditional aircraft designs in order to meet ambitious emissions and fuel burn targets. High temperature superconducting distributed propulsion may be an enabler for aircraft designs that have better propulsive efficiency and lower drag. There has been significant work considering the electrical systems required, but less on the cryogenics to enable it. This paper discusses some of the major choices to be faced in cryocooling for aircraft. The likely need for a disposable cryogen to reduce power demand is explained. A set of cryocooling methods are considered in a sensitivity study, which shows that the feasibility of the cryogenic system will depend strongly on the superconducting technology and the aircraft platform. It is argued that all three aspects must be researched and designed in close collaboration to reach a viable solution.

Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

2013-01-01

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. A three engine retro-propulsion configuration with a 2.5 inch diameter sphere-cone aeroshell model was tested in the NASA Glenn 1x1 Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 degree Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retro-propulsion induced shock waves and retro-propulsion for Earth launched booster recovery are also addressed.

NASA's Nuclear Thermal Propulsion Project

NASA Technical Reports Server (NTRS)

Houts, Michael; Mitchell, Sonny; Kim, Tony; Borowski, Stanley; Power, Kevin; Scott, John; Belvin, Anthony; Clement, Steven

2015-01-01

Space fission power systems can provide a power rich environment anywhere in the solar system, independent of available sunlight. Space fission propulsion offers the potential for enabling rapid, affordable access to any point in the solar system. One type of space fission propulsion is Nuclear Thermal Propulsion (NTP). NTP systems operate by using a fission reactor to heat hydrogen to very high temperature (>2500 K) and expanding the hot hydrogen through a supersonic nozzle. First generation NTP systems are designed to have an Isp of approximately 900 s. The high Isp of NTP enables rapid crew transfer to destinations such as Mars, and can also help reduce mission cost, improve logistics (fewer launches), and provide other benefits. However, for NTP systems to be utilized they must be affordable and viable to develop. NASA's Advanced Exploration Systems (AES) NTP project is a technology development project that will help assess the affordability and viability of NTP. Early work has included fabrication of representative graphite composite fuel element segments, coating of representative graphite composite fuel element segments, fabrication of representative cermet fuel element segments, and testing of fuel element segments in the Compact Fuel Element Environmental Tester (CFEET). Near-term activities will include testing approximately 16" fuel element segments in the Nuclear Thermal Rocket Element Environmental Simulator (NTREES), and ongoing research into improving fuel microstructure and coatings. In addition to recapturing fuels technology, affordable development, qualification, and utilization strategies must be devised. Options such as using low-enriched uranium (LEU) instead of highly-enriched uranium (HEU) are being assessed, although that option requires development of a key technology before it can be applied to NTP in the thrust range of interest. Ground test facilities will be required, especially if NTP is to be used in conjunction with high value or crewed missions. There are potential options for either modifying existing facilities or constructing new ground test facilities. At least three potential options exist for reducing (or eliminating) the release of radioactivity into the environment during ground testing. These include fully containing the NTP exhaust during the ground test, scrubbing the exhaust, or utilizing an existing borehole at the Nevada National Security Site (NNSS) to filter the exhaust. Finally, the project is considering the potential for an early flight demonstration of an engine very similar to one that could be used to support human Mars or other ambitious missions. The flight demonstration could be an important step towards the eventual utilization of NTP.

Technology needs for lunar and Mars space transfer systems

NASA Technical Reports Server (NTRS)

Woodcock, Gordon R.; Cothran, Bradley C.; Donahue, Benjamin; Mcghee, Jerry

1991-01-01

The determination of appropriate space transportation technologies and operating modes is discussed with respect to both lunar and Mars missions. Three levels of activity are set forth to examine the sensitivity of transportation preferences including 'minimum,' 'full science,' and 'industrialization and settlement' categories. High-thrust-profile missions for lunar and Mars transportation are considered in terms of their relative advantages, and transportation options are defined in terms of propulsion and braking technologies. Costs and life-cycle cost estimates are prepared for the transportation preferences by using a parametric cost model, and a return-on-investment summary is given. Major technological needs for the programs are listed and include storable propulsion systems; cryogenic engines and fluids management; aerobraking; and nuclear thermal, nuclear electric, electric, and solar electric propulsion technologies.

Fuel Cells: A Real Option for Unmanned Aerial Vehicles Propulsion

PubMed Central

2014-01-01

The possibility of implementing fuel cell technology in Unmanned Aerial Vehicle (UAV) propulsion systems is considered. Potential advantages of the Proton Exchange Membrane or Polymer Electrolyte Membrane (PEMFC) and Direct Methanol Fuel Cells (DMFC), their fuels (hydrogen and methanol), and their storage systems are revised from technical and environmental standpoints. Some operating commercial applications are described. Main constraints for these kinds of fuel cells are analyzed in order to elucidate the viability of future developments. Since the low power density is the main problem of fuel cells, hybridization with electric batteries, necessary in most cases, is also explored. PMID:24600326

Nuclear Electric Propulsion Application: RASC Mission Robotic Exploration of Venus

NASA Technical Reports Server (NTRS)

McGuire, Melissa L.; Borowski, Stanley K.; Packard, Thomas W.

2004-01-01

The following paper documents the mission and systems analysis portion of a study in which Nuclear Electric Propulsion (NEP) is used as the in-space transportation system to send a series of robotic rovers and atmospheric science airplanes to Venus in the 2020 to 2030 timeframe. As part of the NASA RASC (Revolutionary Aerospace Systems Concepts) program, this mission analysis is meant to identify future technologies and their application to far reaching NASA missions. The NEP systems and mission analysis is based largely on current technology state of the art assumptions. This study looks specifically at the performance of the NEP transfer stage when sending a series of different payload package point design options to Venus orbit.

Low-Enriched Uranium Nuclear Thermal Propulsion Systems

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Mitchell, Doyce P.; Aschenbrenner, Ken

2017-01-01

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. For example, using NTP for human Mars missions can provide faster transit and/or round trip times for crew; larger mission payloads; off nominal mission opportunities (including wider injection windows); and crew mission abort options not available from other architectures. The use of NTP can also reduce required earth-to-orbit launches, reducing cost and improving ground logistics. In addition to enabling robust human Mars mission architectures, NTP can be used on exploration missions throughout the solar system. A first generation NTP system could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. Progress made under the NTP project could also help enable high performance fission power systems and Nuclear Electric Propulsion (NEP). Guidance, navigation, and control of NTP may have some unique but manageable characteristics.

Implementation Options for the PROPEL Electrodynamic Tether Demonstration Mission

NASA Technical Reports Server (NTRS)

Bilen, Sven G.; Johnson, Les; Gilchrist, Brian E.; Hoyt, Robert P.; Elder, Craig H.; Fuhrhop, Keith P.; Scadera, Michael P.; Stone, Nobie H.

2014-01-01

The PROPEL flight mission concept will demonstrate the safe use of an electrodynamic tether for generating thrust. PROPEL is being designed to be a versatile electrodynamic-tether system for multiple end users and to be flexible with respect to platform. As such, several implementation options are being explored, including a comprehensive mission design for PROPEL with a mission duration of six months; a space demonstration mission concept design with configuration of a pair of tethered satellites, one of which is the Japanese H-II Transfer Vehicle; and an ESPA-based system. We report here on these possible implementation options for PROPEL. electrodynamic tether; PROPEL demonstration mission; propellantless propulsion

World Ships: The Solar-Photon Sail Option

NASA Astrophysics Data System (ADS)

Matloff, G. L.

The World Ship, a spacecraft large enough to simulate a small-scale terrestrial internal environment, may be the best feasible option to transfer members of a technological civilization between neighboring stars. Because of the projected size of these spacecraft, journey durations of ~1,000 years seem likely. One of the propulsion options for World Ships is the hyper-thin, likely space-manufactured solar-photon sail, unfurled as close to the migrating civilization's home star as possible. Because the sail and associated structure can be wound around the habitat while not in use, it represents the only known ultimately feasible interstellar propulsion system that can be applied for en route galactic-cosmic ray shielding as well as acceleration/ deceleration. This paper reviews the three suggested sail configurations that can be applied to world ship propulsion: parachute, hollow-body and hoop sails. Possible existing and advanced sail and structure materials and the predicted effects on the sail of the near-Sun space environment are reviewed. Consideration of solar-photon-sail World Ships also affects SETI (the Search for Extraterrestrial Intelligence). Can we detect such craft in flight? When in a star's lifetime is migration using such craft likely? What classes of stars are good candidates for solar-sail World-Ship searches?

Electric propulsion for geostationary orbit insertion

NASA Technical Reports Server (NTRS)

Oleson, Steven R.; Curran, Francis M.; Myers, Roger M.

1995-01-01

Solar electric propulsion (SEP) technology is already being used for geostationary satellite stationkeeping to increase payload mass. By using this same technology to perform part of the orbit transfer additional increases in payload mass can be achieved. Advanced chemical and N2H4 arcjet systems are used to increase the payload mass by performing stationkeeping and part of the orbit transfer. Four mission options are analyzed which show the impact of either sharing the orbit transfer between chemical and SEP systems or having either complete the transfer alone. Results show that for an Atlas 2AS payload increases in net mass (geostationary satellite mass less wet propulsion system mass) of up to 100 kg can be achieved using advanced chemical for the transfer and advanced N2H4 arcjets for stationkeeping. An additional 100 kg can be added using advanced N2H4 arcjets for part of a 40 day orbit transfer.

Space-based laser-powered orbital transfer vehicle (Project SLICK)

NASA Technical Reports Server (NTRS)

1988-01-01

A conceptual design study of a laser-powered orbital transfer vehicle (LOTV) is presented. The LOTV, nicknamed SLICK (Space Laser Interorbital Cargo Kite), will be utilized for the transfer of 16000 kg of cargo between Low Earth Orbit (LEO) and either Geosynchronous Earth Orbit (GEO) or Low Lunar Orbit (LLO). This design concentrates primarily on the LEO/GEO scenario, which will have typical LEO-to-GEO trip time of 6 days and two return versions. One version uses an all propulsive return while the other utilizes a ballute aerobrake for the return trip. Furthermore, three return cargo options of 16000 kg, 5000 kg (standard option), and 1600 kg are considered for this scenario. The LEO/LLO scenario uses only a standard, aerobraked version. The basic concept behind the LOTV is that the power for the propulsion system is supplied by a source separate from the LOTV itself. For the LEO/GEO scenario the LOTV utilizes a direct solar-pumped iodide laser and possibly two relay stations, all orbiting at an altitude of one Earth radius and zero inclination. An additional nuclear-powered laser is placed on the Moon for the LEO/LLO scenario. The propulsion system of the LOTV consists of a single engine fueled with liquid hydrogen. The laser beam is captured and directed by a four mirror optical system through a window in the thrust chamber of the engine. There, seven plasmas are created to convert the laser beam energy into thermal energy at an efficiency of at least 50 percent. For the LEO/LLO scenario the laser propulsion is supplemented by LH2/LOX chemical thrusters.

Solar Electric Propulsion Vehicle Design Study for Cargo Transfer to Earth-moon L1

NASA Technical Reports Server (NTRS)

Sarver-Verhey, Timothy R.; Kerslake, Thomas W.; Rawlin, Vincent K.; Falck, Robert D.; Dudzinski, Leonard J.; Oleson, Steven R.

2002-01-01

A design study for a cargo transfer vehicle using solar electric propulsion was performed for NASA's Revolutionary Aerospace Systems Concepts program. Targeted for 2016, the solar electric propulsion (SEP) transfer vehicle is required to deliver a propellant supply module with a mass of approximately 36 metric tons from Low Earth Orbit to the first Earth-Moon libration point (LL1) within 270 days. Following an examination of propulsion and power technology options, a SEP transfer vehicle design was selected that incorporated large-area (approx. 2700 sq m) thin film solar arrays and a clustered engine configuration of eight 50 kW gridded ion thrusters mounted on an articulated boom. Refinement of the SEP vehicle design was performed iteratively to properly estimate the required xenon propellant load for the out-bound orbit transfer. The SEP vehicle performance, including the xenon propellant estimation, was verified via the SNAP trajectory code. Further efforts are underway to extend this system model to other orbit transfer missions.

Materials Requirements for Advanced Propulsion Systems

NASA Technical Reports Server (NTRS)

Whitaker, Ann F.; Cook, Mary Beth; Clinton, R. G., Jr.

2005-01-01

NASA's mission to "reach the Moon and Mars" will be obtained only if research begins now to develop materials with expanded capabilities to reduce mass, cost and risk to the program. Current materials cannot function satisfactorily in the deep space environments and do not meet the requirements of long term space propulsion concepts for manned missions. Directed research is needed to better understand materials behavior for optimizing their processing. This research, generating a deeper understanding of material behavior, can lead to enhanced implementation of materials for future exploration vehicles. materials providing new approaches for manufacture and new options for In response to this need for more robust materials, NASA's Exploration Systems Mission Directorate (ESMD) has established a strategic research initiative dedicated to materials development supporting NASA's space propulsion needs. The Advanced Materials for Exploration (AME) element directs basic and applied research to understand material behavior and develop improved materials allowing propulsion systems to operate beyond their current limitations. This paper will discuss the approach used to direct the path of strategic research for advanced materials to ensure that the research is indeed supportive of NASA's future missions to the moon, Mars, and beyond.

Lightweight Radiator for in Space Nuclear Electric Propulsion

NASA Technical Reports Server (NTRS)

Craven, Paul; Tomboulian, Briana; SanSoucie, Michael

2014-01-01

Nuclear electric propulsion (NEP) is a promising option for high-speed in-space travel due to the high energy density of nuclear fission power sources and efficient electric thrusters. Advanced power conversion technologies may require high operating temperatures and would benefit from lightweight radiator materials. Radiator performance dictates power output for nuclear electric propulsion systems. Game-changing propulsion systems are often enabled by novel designs using advanced materials. Pitch-based carbon fiber materials have the potential to offer significant improvements in operating temperature, thermal conductivity, and mass. These properties combine to allow advances in operational efficiency and high temperature feasibility. An effort at the NASA Marshall Space Flight Center to show that woven high thermal conductivity carbon fiber mats can be used to replace standard metal and composite radiator fins to dissipate waste heat from NEP systems is ongoing. The goals of this effort are to demonstrate a proof of concept, to show that a significant improvement of specific power (power/mass) can be achieved, and to develop a thermal model with predictive capabilities making use of constrained input parameter space. A description of this effort is presented.

NEP processing, operations, and disposal

NASA Technical Reports Server (NTRS)

Stancati, Mike

1993-01-01

Several recent studies by ASAO/NPO staff members at LeRC and by other organizations have highlighted the potential benefits of using Nuclear Electric Propulsion (NEP) as the primary transportation means for some of the proposed missions of the Space Exploration Initiative. These include the potential to reduce initial mass in orbit and Mars transit time. Modular NEP configurations also introduce fully redundant main propulsion to Mars flight systems adding several abort or fall back options not otherwise available. Recent studies have also identified mission operations, such as on orbital assembly, refurbishment, and reactor disposal, as important discriminators for propulsion system evaluation. This study is intended to identify and assess 'end-to-end' operational issues associated with using NEP for transporting crews and cargo between Earth and Mars. We also include some consideration of lunar cargo transfer as well.

NASA Exploration Team (NExT) In-Space Transportation Overview

NASA Technical Reports Server (NTRS)

Drake, Bret G.; Cooke, Douglas R.; Kos, Larry D.; Brady, Hugh J. (Technical Monitor)

2002-01-01

This presentation provides an overview of NASA Exploration Team's (NEXT) vision of in-space transportation in the future. Hurdles facing in-space transportation include affordable power sources, crew health and safety, optimized robotic and human operations and space systems performance. Topics covered include: exploration of Earth's neighborhood, Earth's neighborhood architecture and elements, Mars mission trajectory options, delta-v variations, Mars mission duration options, Mars mission architecture, nuclear electric propulsion advantages and miscellaneous technology needs.

GASP- General Aviation Synthesis Program. Volume 5: Weight

NASA Technical Reports Server (NTRS)

Hague, D.

1978-01-01

Subroutines for determining the weights of propulsion system related components and the airframe components of an aircraft configuration are presented. Subroutines that deal with design load conditions, aircraft balance, and tail sizing are included. Options for turbine and internal combustion engines are provided.

Personal Access Satellite System (PASS) study. Fiscal year 1989 results

NASA Technical Reports Server (NTRS)

Sue, Miles K. (Editor)

1990-01-01

The Jet Propulsion Laboratory is exploring the potential and feasibility of a personal access satellite system (PASS) that will offer the user greater freedom and mobility than existing or currently planned communications systems. Studies performed in prior years resulted in a strawman design and the identification of technologies that are critical to the successful implementation of PASS. The study efforts in FY-89 were directed towards alternative design options with the objective of either improving the system performance or alleviating the constraints on the user terminal. The various design options and system issues studied this year and the results of the study are presented.

Review of Nuclear Thermal Propulsion Ground Test Options

NASA Technical Reports Server (NTRS)

Coote, David J.; Power, Kevin P.; Gerrish, Harold P.; Doughty, Glen

2015-01-01

High efficiency rocket propulsion systems are essential for humanity to venture beyond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rockets with relatively high thrust and twice the efficiency of highest performing chemical propellant engines. NTP utilizes the coolant of a nuclear reactor to produce propulsive thrust. An NTP engine produces thrust by flowing hydrogen through a nuclear reactor to cool the reactor, heating the hydrogen and expelling it through a rocket nozzle. The hot gaseous hydrogen is nominally expected to be free of radioactive byproducts from the nuclear reactor; however, it has the potential to be contaminated due to off-nominal engine reactor performance. NTP ground testing is more difficult than chemical engine testing since current environmental regulations do not allow/permit open air testing of NTP as was done in the 1960's and 1970's for the Rover/NERVA program. A new and innovative approach to rocket engine ground test is required to mitigate the unique health and safety risks associated with the potential entrainment of radioactive waste from the NTP engine reactor core into the engine exhaust. Several studies have been conducted since the ROVER/NERVA program in the 1970's investigating NTP engine ground test options to understand the technical feasibility, identify technical challenges and associated risks and provide rough order of magnitude cost estimates for facility development and test operations. The options can be divided into two distinct schemes; (1) real-time filtering of the engine exhaust and its release to the environment or (2) capture and storage of engine exhaust for subsequent processing.

Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion

NASA Astrophysics Data System (ADS)

Borowski, Stanley K.; Dudzinski, Leonard A.; McGuire, Melissa L.

2002-12-01

The nuclear thermal rocket (NTR) is one of the leading propulsion options for future human missions to Mars because of its high specific impulse (1sp is approximately 850-1000 s) capability and its attractive engine thrust-to-weight ratio (approximately 3-10). To stay within the available mass and payload volume limits of a "Magnum" heavy lift vehicle, a high performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by three 15 thousand pounds force (klbf) NTR engines is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule burnup of enriched uranium-235 during the Earth departure phase (approximately 10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of this high performance stage. By reconfiguring the engines for both propulsive thrust and modest power generation (referred to as "bimodal" operation), a robust, multiple burn, "power-rich" stage with propulsive Mars capture and reuse capability is possible. A family of modular bimodal NTR (BNTR) vehicles are described which utilize a common "core" stage powered by three 15 klbf BNTRs that produce 50 kWe of total electrical power for crew life support, an active refrigeration / reliquification system for long term, zero-boiloff liquid hydrogen (LH2) storage, and high data rate communications. An innovative, spine-like "saddle truss" design connects the core stage and payload element and is open underneath to allow supplemental "in-line" propellant tanks and contingency crew consumables to be easily jettisoned to improve vehicle performance. A "modified" DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies space operations. By taking the next logical step--use of the BNTR for propulsive capture of all payload elements into Mars orbit--the power available in Mars orbit grows to 150 kWe compared to 30 kWe for the DRM. Propulsive capture also eliminates the complex, higher risk aerobraking and capture maneuver which is replaced by a simpler reentry using a standardized, lower mass "aerodescent" shell. The attractiveness of the "all BNTR" option is further increased by the substitution of the lightweight, inflatable "TransHab" module in place of the heavier, hard-shell hab module. Use of TransHab introduces the potential for propulsive recovery and reuse of the BNTR / Earth return vehicle (ERV). It also allows the crew to travel to and from Mars on the same BNTR transfer vehicle thereby cutting the duration of the ERV mission in half--from approximately 4.7 to 2.5 years. Finally, for difficult Mars options, such as Phobos rendezvous and sample return missions, volume (not mass) constraints limit the performance of the "all LH2" BNTR stage. The use of "LOX-augmented" NTR (LANTR) engines, operating at a modest oxygen-to-hydrogen mixture ratio (MR) of 0.5, helps to increase "bulk" propellant density and total thrust during the TMI burn. On all subsequent burns, the bimodal LANTR engines operate on LH2 only (MR=0) to maximize vehicle performance while staying within the mass limits of two Magnum launches.

Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion. Revised Dec. 1998

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Dudzinski, Leonard A.; McGuire, Melissa L.

2002-01-01

The nuclear thermal rocket (NTR) is one of the leading propulsion options for future human missions to Mars because of its high specific impulse (1sp is approximately 850-1000 s) capability and its attractive engine thrust-to-weight ratio (approximately 3-10). To stay within the available mass and payload volume limits of a "Magnum" heavy lift vehicle, a high performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by three 15 thousand pounds force (klbf) NTR engines is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule burnup of enriched uranium-235 during the Earth departure phase (approximately 10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of this high performance stage. By reconfiguring the engines for both propulsive thrust and modest power generation (referred to as "bimodal" operation), a robust, multiple burn, "power-rich" stage with propulsive Mars capture and reuse capability is possible. A family of modular bimodal NTR (BNTR) vehicles are described which utilize a common "core" stage powered by three 15 klbf BNTRs that produce 50 kWe of total electrical power for crew life support, an active refrigeration / reliquification system for long term, zero-boiloff liquid hydrogen (LH2) storage, and high data rate communications. An innovative, spine-like "saddle truss" design connects the core stage and payload element and is open underneath to allow supplemental "in-line" propellant tanks and contingency crew consumables to be easily jettisoned to improve vehicle performance. A "modified" DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies space operations. By taking the next logical step--use of the BNTR for propulsive capture of all payload elements into Mars orbit--the power available in Mars orbit grows to 150 kWe compared to 30 kWe for the DRM. Propulsive capture also eliminates the complex, higher risk aerobraking and capture maneuver which is replaced by a simpler reentry using a standardized, lower mass "aerodescent" shell. The attractiveness of the "all BNTR" option is further increased by the substitution of the lightweight, inflatable "TransHab" module in place of the heavier, hard-shell hab module. Use of TransHab introduces the potential for propulsive recovery and reuse of the BNTR / Earth return vehicle (ERV). It also allows the crew to travel to and from Mars on the same BNTR transfer vehicle thereby cutting the duration of the ERV mission in half--from approximately 4.7 to 2.5 years. Finally, for difficult Mars options, such as Phobos rendezvous and sample return missions, volume (not mass) constraints limit the performance of the "all LH2" BNTR stage. The use of "LOX-augmented" NTR (LANTR) engines, operating at a modest oxygen-to-hydrogen mixture ratio (MR) of 0.5, helps to increase "bulk" propellant density and total thrust during the TMI burn. On all subsequent burns, the bimodal LANTR engines operate on LH2 only (MR=0) to maximize vehicle performance while staying within the mass limits of two Magnum launches.

Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Dudzinski, Leonard A.; McGuire, Melissa L.

1998-01-01

The nuclear thermal rocket (NTR) is one of the leading propulsion options for future human missions to Mars because of its high specific impulse (Isp-850-1000 s) capability and its attractive engine thrust-to-weight ratio (approximately equal 3-10). To stay within the available mass and payload volume limits of a "Magnum" heavy lift vehicle, a high performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by three 15 thousand pounds force (klbf) NTR engines is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule burnup of enriched uranium-235 during the Earth departure phase (approximately 10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of this high performance stage. By reconfiguring the engines for both propulsive thrust and modest power generation (referred to as "bimodal" operation), a robust, multiple burn, "power-rich" stage with propulsive Mars capture and reuse capability is possible, A family of modular "bimodal" NTR (BNTR) vehicles are described which utilize a common "core" stage powered by three 15 klbf BNTRs that produce 50 kWe of total electrical power for crew life support, an active refrigeration / reliquification system for long term, "zero-boiloff" liquid hydrogen (LH2) storage, and high data rate communications. An innovative, spine-like "saddle truss" design connects the core stage and payload element and is open underneath to allow supplemental "in-line" propellant tanks and contingency crew consumables to be easily jettisoned to improve vehicle performance. A "modified" DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies space operations. By taking the next logical step--use of the BNTR for propulsive capture of all payload elements into Mars orbit--the power available in Mars orbit grows to 150 kWe compared to 30 kWe for the DRM. Propulsive capture also eliminates the complex, higher risk aerobraking and capture maneuver which is replaced by a simpler reentry using a standardized, lower mass "aerodescent" shell. The attractiveness of the "all BNTR" option is further increased by the substitution of the lightweight, inflatable "TransHab" module in place of the heavier, hard-shell hab module. Use of TransHab introduces the potential for propulsive recovery and reuse of the BNTR/ERV. It also allows the crew to travel to and from Mar on the same BNTR transfer vehicle thereby cutting the duration of the ERV mission in half--from approximately 4.7 to 2.5 years. Finally, for difficult Mars options, such as Phobos rendezvous and sample return missions, volume (not mass) constraints limit the performance of the "all LH2" BNTR stage. The use of "LOX-augmented" NTR (LANTR) engines, operating at a modest oxygen-to-hydrogen mixutre ratio (MR) of 0.5, helps to increase "bulk" propellant density and total thrust during the TMI burn. On all subsequent burns, the bimodal LANTR engines operate on LH2 only (MR=0) to maximize vehicle performance while staying within the lift capability of two Magnum launches.

A potassium Rankine multimegawatt nuclear electric propulsion concept

NASA Technical Reports Server (NTRS)

Baumeister, E.; Rovang, R.; Mills, J.; Sercel, J.; Frisbee, R.

1990-01-01

Multimegawatt nuclear electric propulsion (NEP) has been identified as a potentially attractive option for future space exploratory missions. A liquid-metal-cooled reactor, potassium Rankine power system that is being developed is suited to fulfill this application. The key features of the nuclear power system are described, and system characteristics are provided for various potential NEP power ranges and operational lifetimes. The results of recent mission studies are presented to illustrate some of the potential benefits to future space exploration to be gained from high-power NEP. Specifically, mission analyses have been performed to assess the mass and trip time performance of advanced NEP for both cargo and piloted missions to Mars.

Electrical power systems for Mars

NASA Technical Reports Server (NTRS)

Giudici, Robert J.

1986-01-01

Electrical power system options for Mars Manned Modules and Mars Surface Bases were evaluated for both near-term and advanced performance potential. The power system options investigated for the Mission Modules include photovoltaics, solar thermal, nuclear reactor, and isotope power systems. Options discussed for Mars Bases include the above options with the addition of a brief discussion of open loop energy conversion of Mars resources, including utilization of wind, subsurface thermal gradients, and super oxides. Electrical power requirements for Mission Modules were estimated for three basic approaches: as a function of crew size; as a function of electric propulsion; and as a function of transmission of power from an orbiter to the surface of Mars via laser or radio frequency. Mars Base power requirements were assumed to be determined by production facilities that make resources available for follow-on missions leading to the establishment of a permanently manned Base. Requirements include the production of buffer gas and propellant production plants.

Electrical power systems for Mars

NASA Astrophysics Data System (ADS)

Giudici, Robert J.

1986-05-01

Electrical power system options for Mars Manned Modules and Mars Surface Bases were evaluated for both near-term and advanced performance potential. The power system options investigated for the Mission Modules include photovoltaics, solar thermal, nuclear reactor, and isotope power systems. Options discussed for Mars Bases include the above options with the addition of a brief discussion of open loop energy conversion of Mars resources, including utilization of wind, subsurface thermal gradients, and super oxides. Electrical power requirements for Mission Modules were estimated for three basic approaches: as a function of crew size; as a function of electric propulsion; and as a function of transmission of power from an orbiter to the surface of Mars via laser or radio frequency. Mars Base power requirements were assumed to be determined by production facilities that make resources available for follow-on missions leading to the establishment of a permanently manned Base. Requirements include the production of buffer gas and propellant production plants.

NEXT Ion Propulsion System Configurations and Performance for Saturn System Exploration

NASA Technical Reports Server (NTRS)

Benson, Scott W.; Riehl, John P.; Oleson, Steven R.

2007-01-01

The successes of the Cassini/Huygens mission have heightened interest to return to the Saturn system with focused robotic missions. The desire for a sustained presence at Titan, through a dedicated orbiter and in-situ vehicle, either a lander or aerobot, has resulted in definition of a Titan Explorer flagship mission as a high priority in the Solar System Exploration Roadmap. The discovery of active water vapor plumes erupting from the tiger stripes on the moon Enceladus has drawn the attention of the space science community. The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system is well suited to future missions to the Saturn system. NEXT is used within the inner solar system, in combination with a Venus or Earth gravity assist, to establish a fast transfer to the Saturn system. The NEXT system elements are accommodated in a separable Solar Electric Propulsion (SEP) module, or are integrated into the main spacecraft bus, depending on the mission architecture and performance requirements. This paper defines a range of NEXT system configurations, from two to four thrusters, and the Saturn system performance capability provided. Delivered mass is assessed parametrically over total trip time to Saturn. Launch vehicle options, gravity assist options, and input power level are addressed to determine performance sensitivities. A simple two-thruster NEXT system, launched on an Atlas 551, can deliver a spacecraft mass of over 2400 kg on a transfer to Saturn. Similarly, a four-thruster system, launched on a Delta 4050 Heavy, delivers more than 4000 kg spacecraft mass. A SEP module conceptual design, for a two thruster string, 17 kW solar array, configuration is characterized.

Solar System Exploration Augmented by In-Situ Resource Utilization: Human Planetary Base Issues for Mercury and Saturn

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan A.

2017-01-01

Establishing a lunar presence and creating an industrial capability on the Moon may lead to important new discoveries for all of human kind. Historical studies of lunar exploration, in-situ resource utilization (ISRU) and industrialization all point to the vast resources on the Moon and its links to future human and robotic exploration. In references 1 through 9, a broad range of technological innovations are described and analyzed. Figures 1 depicts program planning for future human missions throughout the solar system which included lunar launched nuclear rockets, and future human settlements on the Moon. Figures 2 and 3 present the results for human Mercury missions, including LEO departure masses and round trip Mercury lander masses. Using in-situ resources, the missions become less burdensome to the LEO launch infrastructure. In one example using Mercury derived hydrogen, the LEO mass of the human Mercury missions can be reduced from 2,800 MT to 1,140 MT (Ref. 15). Additional analyses of staging options for human Mercury missions will be presented. Figures 4 shows an option for thermal control for long term in-space cryogenic storage and Figure 5 depicts the potentially deleterious elements emanating from Mercury that must be addressed, respectively. Updated analyses based on the visions presented will be presented. While advanced propulsion systems were proposed in these historical studies, further investigation of nuclear options using high power nuclear thermal and nuclear electric propulsion as well as advanced chemical propulsion can significantly enhance these scenarios. Human bases at Mercury may have to be resupplied from resources from regolith and water resources in permanently shadowed craters at its northern pole.

Mars 2024/2026 Pathfinder Mission: Mars Architectures, Systems, and Technologies for Exploration and Resources Project

NASA Technical Reports Server (NTRS)

Zeitlin, Nancy; Mueller, Robert; Muscatello, Anthony

2015-01-01

Integrate In Situ Resource Utilization (ISRU) sub-systems and examine advanced capabilities and technologies to verify Mars 2024 Forward architecture precursor pathfinder options: Integrated spacecraft/surface infrastructure fluid architecture: propulsion, power, life support center dot Power system feed and propellant scavenging from propulsion system center dot High quality oxygen for life support and EVA Fluid/cryogenic zero-loss transfer and long-term storage center dot Rapid depot-to-rover/spacecraft center dot Slow ISRU plant-to-ascent vehicle Integration of ISRU consumable production center dot Oxygen only from Mars atmosphere carbon dioxide center dot Oxygen, fuel, water, from extraterrestrial soil/regolith Test bed to evaluate long duration life, operations, maintenance on hardware, sensors, and autonomy

Extended applications study of AMOOS and AMRS

NASA Technical Reports Server (NTRS)

White, J.

1977-01-01

The potential advantages of the Aeromaneuvering Orbit-to-Orbit Shuttle (AMOOS) over the all-propulsive Orbit Transfer Vehicle (OTV) are shown. In particular, the kit concept studies and the dual fueled AMOOS studies show its versatility and option potential over the all-propulsive vehicle. All of this potential of AMOOS and the Aeromaneuvering Recovery System (AMRS) depends upon the ability to control the trajectory during atmospheric flight and so use an ablative TPS. In turn, this TPS must be light weight, which can be attained by spraying a lightweight ablator directly onto the load bearing skin.

Space nuclear power system and the design of the nuclear electric propulsion OTV

NASA Technical Reports Server (NTRS)

Buden, D.; Garrison, P. W.

1984-01-01

Payload increases of three to five times that of the Shuttle/Centaur can be achieved using nuclear electric propulsion. Various nuclear power plant options being pursued by the SP-100 Program are described. These concepts can grow from 100 kWe to 1 MWe output. Spacecraft design aspects are addressed, including thermal interactions, plume interactions, and radiation fluences. A baseline configuration is described accounting for these issues. Safety aspects of starting the OTV transfer from an altitude of 300 km indicate no significant additional risk to the biosphere.

Material Challenges and Opportunities for Commercial Electric Aircraft

NASA Technical Reports Server (NTRS)

Misra, Ajay

2014-01-01

Significant reduction in carbon dioxide emission for future air transportation system will require adoption of electric propulsion system and more electric architectures. Various options for aircraft electric propulsion include hybrid electric, turboelectric, and full electric system. Realization of electric propulsion system for commercial aircraft applications will require significant increases in power density of electric motors and energy density of energy storage system, such as the batteries and fuel cells. In addition, transmission of MW of power in the aircraft will require high voltage power transmission system to reduce the weight of the power transmission system. Finally, there will be significant thermal management challenges. Significant advances in material technologies will be required to meet these challenges. Technologies of interest include materials with higher electrical conductivity than Cu, high thermal conductivity materials, and lightweight electrically insulating materials with high breakdown voltage, high temperature magnets, advanced battery and fuel cell materials, and multifunctional materials. The presentation will include various challenges for commercial electric aircraft and provide an overview of material improvements that will be required to meet these challenges.

Evaluation of High-Performance Space Nuclear Electric Generators for Electric Propulsion Application

NASA Technical Reports Server (NTRS)

Woodcock, Gordon; Kross, Dennis A. (Technical Monitor)

2002-01-01

Electric propulsion applications are enhanced by high power-to-mass ratios for their electric power sources. At multi-megawatt levels, we can expect thrust production systems to be less than 5 kg/kWe. Application of nuclear electric propulsion to human Mars missions becomes an attractive alternative to nuclear thermal propulsion if the propulsion system is less than about 10 kg/kWe. Recent references have projected megawatt-plus nuclear electric sources at specific mass values from less than 1 kg/kWe to about 5 kg/kWe. Various assumptions are made regarding power generation cycle (turbogenerator; MHD (magnetohydrodynamics)) and reactor heat source design. The present paper compares heat source and power generation options on the basis of a parametric model that emphasizes heat transfer design and realizable hardware concept. Pressure drop (important!) is included in the power cycle analysis, and MHD and turbogenerator cycles are compared. Results indicate that power source specific mass less than 5 kg/kWe is attainable, even if peak temperatures achievable are limited to 1500 K. Projections of specific mass less than 1 kg/kWe are unrealistic, even at the highest peak temperatures considered.

Concept Design of High Power Solar Electric Propulsion Vehicles for Human Exploration

NASA Technical Reports Server (NTRS)

Hoffman, David J.; Kerslake, Thomas W.; Hojnicki, Jeffrey S.; Manzella, David H.; Falck, Robert D.; Cikanek, Harry A., III; Klem, Mark D.; Free, James M.

2011-01-01

Human exploration beyond low Earth orbit will require enabling capabilities that are efficient, affordable and reliable. Solar electric propulsion (SEP) has been proposed by NASA s Human Exploration Framework Team as one option to achieve human exploration missions beyond Earth orbit because of its favorable mass efficiency compared to traditional chemical propulsion systems. This paper describes the unique challenges associated with developing a large-scale high-power (300-kWe class) SEP vehicle and design concepts that have potential to meet those challenges. An assessment of factors at the subsystem level that must be considered in developing an SEP vehicle for future exploration missions is presented. Overall concepts, design tradeoffs and pathways to achieve development readiness are discussed.

Mars Earth Return Vehicle (MERV) Propulsion Options

NASA Technical Reports Server (NTRS)

Oleson, Steven R.; McGuire, Melissa L.; Burke, Laura; Fincannon, James; Warner, Joe; Williams, Glenn; Parkey, Thomas; Colozza, Tony; Fittje, Jim; Martini, Mike;

Electric Propulsion Apparatus

NASA Technical Reports Server (NTRS)

Patterson, Michael J. (Inventor)

2013-01-01

An electric propulsion machine includes an ion thruster having an annular discharge chamber housing an anode having a large surface area. The ion thruster includes flat annular ion optics with a small span to gap ratio. Optionally, a second electric propulsion thruster may be disposed in a cylindrical space disposed within an interior of the annulus.

Instellar Exploration: Propulsion Options for Precursors and Beyond

NASA Technical Reports Server (NTRS)

Johnson, Charles Les; Leifer, Stephanie

1999-01-01

NASA is considering a mission to explore near-interstellar space early in the next decade as the first step toward a vigorous interstellar exploration program. A key enabling technology for such an ambitious science and exploration effort is the development of propulsion systems capable of providing fast trip times; mission duration should not exceed the professional lifetime of the investigative team. Advanced propulsion technologies that might support an interstellar precursor mission early in the next century include some combination of solar sails, nuclear electric propulsion systems, and aerogravity assists. Follow-on missions to far beyond the heliopause will require the development of propulsion technologies that are only at the conceptual stage today. These include 1) matter-antimatter annihilation, 2) beamed-energy sails, and 3) fusion systems. For years, the scientific community has been interested in the development of solar sail technology to support exploration of the inner and outer planets. Progress in thin-film technology and the development of technologies that may enable the remote assembly of large sails in space are only now maturing to the point where ambitious interstellar precursor missions can be considered. Electric propulsion is now being demonstrated for planetary exploration by the Deep Space 1 mission. The primary issues for it's adaptation to interstellar precursor applications include the nuclear reactor that would be required and the engine lifetime. For further term interstellar missions, matter-antimatter annihilation propulsion system concepts have the highest energy density of any propulsion systems using onboard propellants. However, there are numerous challenges to production and storage of antimatter that must be overcome before it can be seriously considered for interstellar flight. Off-board energy systems (laser sails) are candidates for long-distance interstellar flight but development of component technologies and necessary infrastructure have not begun.. Fusion propulsion has been studied extensively. However, fusion technology is still considered immature, even after many decades of well-funded research. Furthermore, fusion alone does not offer high enough energy density to make it a viable candidate for interstellar propulsion unless propellant can be collected in situ, as was considered by R. Bussard for his interstellar ramjet concept. The current research in investigating these propulsion systems will be described, and the range of application of each technology will be explored.

Rotorcraft convertible engine study

NASA Technical Reports Server (NTRS)

Gill, J. C.; Earle, R. V.; Mar, H. M.

1982-01-01

The objective of the Rotorcraft Convertible Engine Study was to define future research and technology effort required for commercial development by 1988 of convertible fan/shaft gas turbine engines for unconventional rotorcraft transports. Two rotorcraft and their respective missions were defined: a Fold Tilt Rotor aircraft and an Advancing Blade Concept (ABC) rotorcraft. Sensitivity studies were conducted with these rotorcraft to determine parametrically the influence of propulsion characteristics on aircraft size, mission fuel requirements, and direct operating costs (DOC). The two rotorcraft were flown with conventional propulsion systems (separate lift/cruise engines) and with convertible propulsion systems to determine the benefits to be derived from convertible engines. Trade-off studies were conducted to determine the optimum engine cycle and staging arrangement for a convertible engine. Advanced technology options applicable to convertible engines were studied. Research and technology programs were identified which would ensure technology readiness for commercial development of convertible engines by 1988.

Incident Energy Focused Design and Validation for the Floating Potential Probe

NASA Technical Reports Server (NTRS)

Fincannon, James

2002-01-01

Utilizing the spacecraft shadowing and incident energy analysis capabilities of the NASA Glenn Research Center Power and Propulsion Office's SPACE System Power Analysis for Capability Evaluation) computer code, this paper documents the analyses for various International Space Station (ISS) Floating Potential Probe (EPP) preliminary design options. These options include various solar panel orientations and configurations as well as deployment locations on the ISS. The incident energy for the final selected option is characterized. A good correlation between the predicted data and on-orbit operational telemetry is demonstrated. Minor deviations are postulated to be induced by degradation or sensor drift.

Australia’s Submarine Design Capabilities and Capacities: Challenges and Options for the Future Submarine

DTIC Science & Technology

2011-01-01

stealth features requiring specialised noise and vibra- tion skills and propulsion plants requiring other unique skill sets. Personnel with these...analysis Acoustic, wake , thermal, electromagnetic, and other signature analysis Combat systems and ship control Combat system integration, combat system...to-diagnose flow-induced radiated noise Own-sensor performance degradation Note: Risks can be reduced for given designs using scale models

Investigation of propulsion system for large LNG ships

NASA Astrophysics Data System (ADS)

Sinha, R. P.; Nik, Wan Mohd Norsani Wan

2012-09-01

Requirements to move away from coal for power generation has made LNG as the most sought after fuel source, raising steep demands on its supply and production. Added to this scenario is the gradual depletion of the offshore oil and gas fields which is pushing future explorations and production activities far away into the hostile environment of deep sea. Production of gas in such environment has great technical and commercial impacts on gas business. For instance, laying gas pipes from deep sea to distant receiving terminals will be technically and economically challenging. Alternative to laying gas pipes will require installing re-liquefaction unit on board FPSOs to convert gas into liquid for transportation by sea. But, then because of increased distance between gas source and receiving terminals the current medium size LNG ships will no longer remain economical to operate. Recognizing this business scenario shipowners are making huge investments in the acquisition of large LNG ships. As power need of large LNG ships is very different from the current small ones, a variety of propulsion derivatives such as UST, DFDE, 2-Stroke DRL and Combined cycle GT have been proposed by leading engine manufacturers. Since, propulsion system constitutes major element of the ship's capital and life cycle cost, which of these options is most suited for large LNG ships is currently a major concern of the shipping industry and must be thoroughly assessed. In this paper the authors investigate relative merits of these propulsion options against the benchmark performance criteria of BOG disposal, fuel consumption, gas emissions, plant availability and overall life cycle cost.

ISS Propulsion Module Crew Systems Interface Analysis in the Intelligent Synthesis Environment

NASA Technical Reports Server (NTRS)

Chen, Di-Wen

1999-01-01

ERGO, a human modeling software for ergonomic assessment and task analysis, was used for the crew systems interface analysis of the International Space Station (ISS) Propulsion Module (PM). The objective of analysis was to alleviate passageway size concerns. Three basic passageway configuration concepts: (1) 45" clear passageway without centerline offset (2) 50" clear passageway, 12" centerline offset, (3) 50" clear passageway, no centerline offset, and were reviewed. 95 percentile male and female models which were provided by the software performed crew system analysis from an anthropometric point of view. Four scenarios in which the crew floats in microgravity through a 50" no-offset passageway as they carry a 16" x 20" x 30" avionics box were simulated in the 10-weeks of intensive study. From the results of the analysis, concept (3) was the preferred option. A full scale, three-dimensional virtual model of the ISS Propulsion Module was created to experience the sense of the Intelligent Synthesis Environment and to evaluate the usability and applicability of the software.

Nuclear Engine System Simulation (NESS). Version 2.0: Program user's guide

NASA Technical Reports Server (NTRS)

Pelaccio, Dennis G.; Scheil, Christine M.; Petrosky, Lyman

1993-01-01

This Program User's Guide discusses the Nuclear Thermal Propulsion (NTP) engine system design features and capabilities modeled in the Nuclear Engine System Simulation (NESS): Version 2.0 program (referred to as NESS throughout the remainder of this document), as well as its operation. NESS was upgraded to include many new modeling capabilities not available in the original version delivered to NASA LeRC in Dec. 1991, NESS's new features include the following: (1) an improved input format; (2) an advanced solid-core NERVA-type reactor system model (ENABLER 2); (3) a bleed-cycle engine system option; (4) an axial-turbopump design option; (5) an automated pump-out turbopump assembly sizing option; (6) an off-design gas generator engine cycle design option; (7) updated hydrogen properties; (8) an improved output format; and (9) personal computer operation capability. Sample design cases are presented in the user's guide that demonstrate many of the new features associated with this upgraded version of NESS, as well as design modeling features associated with the original version of NESS.

Space transfer concepts and analysis for exploration missions. Implementation plan and element description document (draft final). Volume 2: Cryo/aerobrake vehicle

NASA Technical Reports Server (NTRS)

1991-01-01

The cryogenic/aerobrake (CAB) and the cryogenic all-propulsive (CAP) concept designs developed in support of the Space Transfer Concepts and Analysis for Exploration Missions (STCAEM) study are presented. The evolution of the CAB and CAP concepts is described along with the requirements, guidelines and assumptions for the designs. Operating modes and options are defined and systems descriptions of the vehicles are presented. Artificial gravity configuration options and space and ground support systems are discussed. Finally, an implementation plan is presented which addresses technology needs, schedules, facilities, and costs.

Hot Water Propulsion for Horizontal Rocket Assisted Take-Off Systems for Future Reusable Launch Vehicles

NASA Astrophysics Data System (ADS)

Pilz, N.; Adirim, H.; Lo, R.; Schildknecht, A.

2004-10-01

Among other concepts, reusable space transportation systems that comprise winged reusable launch vehicles (RLV) with horizontal take-off and horizontal landing (HTHL) are under worldwide investigation, e.g. the respective concepts within ESA's FESTIP-Study (Future European Space Transportation Integration Program) or the HOPPER concept by EADS-ST. The payload of these RLVs could be significantly increased by means of a ground-based take-off assistance system that would accelerate the vehicle along a horizontal track until it reaches the desired speed to ignite its onboard engines for leaving the ground and launching into orbit. This paper illustrates the advantages of horizontal take-off for winged RLVs and provides an overview of launch-assist options for HTHL RLVs. It presents hot water propulsion for ground-based take-off assistance systems for future RLVs as an attractive choice besides magnetic levitation and acceleration (maglev) technology. Finally, preliminary design concepts are presented for a rocket assisted take-off system (RATOS) with hot water propulsion followed by an analysis of its improvement potential.

Affordable Development and Demonstrationof a Small NTR Engine and Stage: A Preliminary NASA, DOE and Industry Assessment

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Sefcik, Robert J.; Qualls, A. L.; Schnitzler, B.G.; Joyner, C. R.

2014-01-01

Formulation of Affordable and Sustainable NTP Development Strategy is Underway Involving NASA, DOE and Industry. In FY11, Nuclear Thermal Propulsion (NTP) was identified as a key propulsion option under the Advanced In-Space Propulsion (AISP) component of NASA's Exploration Technology Development and Demonstration (ETDD) program.

Transportation analyses for the lunar-Mars initiative

NASA Technical Reports Server (NTRS)

Woodcock, Gordon R.; Buddington, Patricia A.

1991-01-01

This paper focuses on certain results of an ongoing NASA-sponsored study by Boeing, including (1) a series of representative space exploration scenarios; (2) the levels of effort required to accomplish each; and (3) a range of candidate transportation system as partial implementations of the scenarios. This effort predated release of the Synthesis report; the three levels of activity described are not responses to the Synthesis architectures. These three levels (minimum, median and ambitious), do envelope the range of scope described in the four Synthesis architecture models. The level of analysis detail was to the current known level of detail of transportation hardware systems and mission scenarios. The study did not include detailed analysis of earth-to-orbit transportation, surface systems, or tracking and communications systems. The influence of earth-to-orbit systems was considered in terms of delivery capacity and cost. Aspects of additional options, such as in situ resource utilization are explored as needed to indicate potential benefits. Results favored cryogenic chemical propulsion for low activity levels and undemanding missions (such as cargo and some lunar missions), nuclear thermal propulsion for median activity levels similar to the Synthesis architectures, and nuclear thermal propulsion with aerobraking or nuclear electric propulsion for high activity levels. Solar electric propulsion was seen as having an important role if the present high unit cost (i.e., dollars per watt) of space photovoltaics could be reduced by a factor of five or more at production rates of megawatts per year.

Small Reactor Designs Suitable for Direct Nuclear Thermal Propulsion: Interim Report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bruce G. Schnitzler

Advancement of U.S. scientific, security, and economic interests requires high performance propulsion systems to support missions beyond low Earth orbit. A robust space exploration program will include robotic outer planet and crewed missions to a variety of destinations including the moon, near Earth objects, and eventually Mars. Past studies, in particular those in support of both the Strategic Defense Initiative (SDI) and the Space Exploration Initiative (SEI), have shown nuclear thermal propulsion systems provide superior performance for high mass high propulsive delta-V missions. In NASA's recent Mars Design Reference Architecture (DRA) 5.0 study, nuclear thermal propulsion (NTP) was again selectedmore » over chemical propulsion as the preferred in-space transportation system option for the human exploration of Mars because of its high thrust and high specific impulse ({approx}900 s) capability, increased tolerance to payload mass growth and architecture changes, and lower total initial mass in low Earth orbit. The recently announced national space policy2 supports the development and use of space nuclear power systems where such systems safely enable or significantly enhance space exploration or operational capabilities. An extensive nuclear thermal rocket technology development effort was conducted under the Rover/NERVA, GE-710 and ANL nuclear rocket programs (1955-1973). Both graphite and refractory metal alloy fuel types were pursued. The primary and significantly larger Rover/NERVA program focused on graphite type fuels. Research, development, and testing of high temperature graphite fuels was conducted. Reactors and engines employing these fuels were designed, built, and ground tested. The GE-710 and ANL programs focused on an alternative ceramic-metallic 'cermet' fuel type consisting of UO2 (or UN) fuel embedded in a refractory metal matrix such as tungsten. The General Electric program examined closed loop concepts for space or terrestrial applications as well as open loop systems for direct nuclear thermal propulsion. Although a number of fast spectrum reactor and engine designs suitable for direct nuclear thermal propulsion were proposed and designed, none were built. This report summarizes status results of evaluations of small nuclear reactor designs suitable for direct nuclear thermal propulsion.« less

Development of a unified guidance system for geocentric transfer. [for solar electric propulsion spacecraft

NASA Technical Reports Server (NTRS)

Cake, J. E.; Regetz, J. D., Jr.

1975-01-01

A method is presented for open loop guidance of a solar electric propulsion spacecraft to geosynchronous orbit. The method consists of determining the thrust vector profiles on the ground with an optimization computer program, and performing updates based on the difference between the actual trajectory and that predicted with a precision simulation computer program. The motivation for performing the guidance analysis during the mission planning phase is discussed, and a spacecraft design option that employs attitude orientation constraints is presented. The improvements required in both the optimization program and simulation program are set forth, together with the efforts to integrate the programs into the ground support software for the guidance system.

Development of a unified guidance system for geocentric transfer. [solar electric propulsion spacecraft

NASA Technical Reports Server (NTRS)

Cake, J. E.; Regetz, J. D., Jr.

1975-01-01

A method is presented for open loop guidance of a solar electric propulsion spacecraft to geosynchronsus orbit. The method consists of determining the thrust vector profiles on the ground with an optimization computer program, and performing updates based on the difference between the actual trajectory and that predicted with a precision simulation computer program. The motivation for performing the guidance analysis during the mission planning phase is discussed, and a spacecraft design option that employs attitude orientation constraints is presented. The improvements required in both the optimization program and simulation program are set forth, together with the efforts to integrate the programs into the ground support software for the guidance system.

Electric Propulsion

NASA Astrophysics Data System (ADS)

Baggett, R.

2004-11-01

Next Generation Electric Propulsion (NGEP) technology development tasks are working towards advancing solar-powered electric propulsion systems and components to levels ready for transition to flight systems. Current tasks within NGEP include NASA's Evolutionary Xenon Thruster (NEXT), Carbon Based Ion Optics (CBIO), NSTAR Extended Life Test (ELT) and low-power Hall Effect thrusters. The growing number of solar electric propulsion options provides reduced cost and flexibility to capture a wide range of Solar System exploration missions. Benefits of electric propulsion systems over state-of-the-art chemical systems include increased launch windows, which reduce mission risk; increased deliverable payload mass for more science; and a reduction in launch vehicle size-- all of which increase the opportunities for New Frontiers and Discovery class missions. The Dawn Discovery mission makes use of electric propulsion for sequential rendezvous with two large asteroids (Vesta then Ceres), something not possible using chemical propulsion. NEXT components and thruster system under development have NSTAR heritage with significant increases in maximum power and Isp along with deep throttling capability to accommodate changes in input power over the mission trajectory. NEXT will produce engineering model system components that will be validated (through qualification-level and integrated system testing) and ready for transition to flight system development. NEXT offers Discovery, New Frontiers, Mars Exploration and outer-planet missions a larger deliverable payload mass and a smaller launch vehicle size. CBIO addresses the need to further extend ion thruster lifetime by using low erosion carbon-based materials. Testing of 30-cm Carbon-Carbon and Pyrolytic graphite grids using a lab model NSTAR thruster are complete. In addition, JPL completed a 1000 hr. life test on 30-cm Carbon-Carbon grids. The NSTAR ELT was a life time qualification test started in 1999 with a goal of 88 kg throughput of Xenon propellant. The test was intentionally terminated in 2003 after accumulating 233 kg throughput. The thruster has been completely disassembled and the conditions of all components documented. Because most of the NSTAR design features have been used in the NEXT thruster, the success of the ELT goes a long way toward qualifying NEXT by similarity Recent mission analyses for Discovery and New Frontiers class missions have also identified potential benefits of low-power, high thrust Hall Effect thrusters. Estimated to be ready for mission implementation by 2008, low-power Hall systems could increase mission capture for electric propulsion by greatly reducing propulsion cost, mass and complexity.

Sunmaster: An SEP cargo vehicle for Mars missions

NASA Technical Reports Server (NTRS)

Chiles, Aleasa; Fraser, Jennifer; Halsey, Andy; Honeycutt, David; Madden, Michael; Mcgough, Brian; Paulsen, David; Spear, Becky; Tarkenton, Lynne; Westley, Kevin

1991-01-01

Options are examined for an unmanned solar powered electric propulsion cargo vehicle for Mars missions. The 6 prime areas of study include: trajectory, propulsion system, power system, supporting structure, control system, and launch consideration. Optimization of the low thrust trajectory resulted in a total round trip mission time just under 4 years. The argon propelled electrostatic ion thruster system consists of seventeen 5 N engines and uses a specific impulse of 10,300 secs. At Earth, the system uses 13 engines to produce 60 N of thrust; at Mars, five engines are used, producing 25 N thrust. The thrust of the craft is varied between 60 N at Earth and 24 N at Mars due to reduced solar power available. Solar power is collected by a Fresnel lens concentrator system using a multistacked cell. This system provides 3.5 MW to the propulsion system after losses. Control and positioning to the craft are provided by a system of three double gimballed control moment gyros. Four shuttle 'C' launches will be used to transport the unassembled vehicle in modular units to low Earth orbit where it will be assembled using the Mobile Transporter of the Space Station Freedom.

Propulsion options for very low Earth orbit microsatellites

NASA Astrophysics Data System (ADS)

Leomanni, Mirko; Garulli, Andrea; Giannitrapani, Antonio; Scortecci, Fabrizio

2017-04-01

The growing competitiveness in the commercial space market has raised the interest in operating small spacecraft at very low altitudes. To make this feasible, the space industry has started developing propulsion options tailored specifically to these platforms. This paper presents a review of emerging micropropulsion technologies and evaluates their applicability to microsatellite missions in the altitude range 250-500 km. The results of the proposed analysis are demonstrated on two different remote sensing applications.

V/STOL Systems Research Aircraft: A Tool for Cockpit Integration

NASA Technical Reports Server (NTRS)

Stortz, Michael W.; ODonoghue, Dennis P.; Tiffany, Geary (Technical Monitor)

1995-01-01

The next generation ASTOVL aircraft will have a complicated propulsion System. The configuration choices include Direct Lift, Lift-Fan and Lift+Lift /Cruise but the aircraft must also have supersonic performance and low-observable characteristics. The propulsion system may have features such as flow blockers, vectoring nozzles and flow transfer schemes. The flight control system will necessarily fully integrate the aerodynamic surfaces and the propulsive elements. With a fully integrated, fly-by-wire flight/propulsion control system, the options for cockpit integration are interesting and varied. It is possible to decouple longitudinal and vertical responses allowing the pilot to close the loop on flight path and flight path acceleration directly. In the hover, the pilot can control the translational rate directly without having to stabilize the inner rate and attitude loops. The benefit of this approach, reduced workload and increased precision. has previously been demonstrated through several motion-based simulations. In order to prove the results in flight, the V/STOL System Research Aircraft (VSRA) was developed at the NASA Ames Research Center. The VSRA is the YAV-8B Prototype modified with a research flight control system using a series-parallel servo configuration in all the longitudinal degrees of freedom (including thrust and thrust vector angle) to provide an integrated flight and propulsion control system in a limited envelope. Development of the system has been completed and flight evaluations of the response types have been performed. In this paper we will discuss the development of the VSRA, the evolution of the flight path command and translational rate command response types and the Guest Pilot evaluations of the system. Pilot evaluation results will be used to draw conclusions regarding the suitability of the system to satisfy V/STOL requirements.

V/STOL systems research aircraft: A tool for cockpit integration

NASA Technical Reports Server (NTRS)

Stortz, Michael W.; ODonoghue, Dennis P.

1995-01-01

The next generation ASTOVL aircraft will have a complicated propulsion system. The configuration choices include Direct Lift, Lift-Fan and Lift + Lift/Cruise but the aircraft must also have supersonic performance and low-observable characteristics. The propulsion system may have features such as flow blockers, vectoring nozzles and flow transfer schemes. The flight control system will necessarily fully integrate the aerodynamic surfaces and the propulsive elements. With a fully integrated, fly-by-wire flight/propulsion control system, the options for cockpit integration are interesting and varied. It is possible to de-couple longitudinal and vertical responses allowing the pilot to close the loop on flightpath and flightpath acceleration directly. In the hover, the pilot can control the translational rate directly without having to stabilize the inner rate and attitude loops. The benefit of this approach, reduced workload and increased precision, has previously been demonstrated through several motion-based simulations. In order to prove the results in flight, the V/STOL System Research Aircraft (VSRA) was developed at the NASA Ames Research Center. The VSRA is the YAV-8B Prototype modified with a research flight control system using a series-parallel servo configuration in all the longitudinal degrees of freedom (including thrust and thrust vector angle) to provide an integrated flight and propulsion control system in a limited envelope. Development of the system has been completed and flight evaluations of the response types have been performed. In this paper we will discuss the development of the VSRA, the evolution of the flightpath command and translational rate command response types and the Guest Pilot evaluations of the system. Pilot evaluation results are used to draw conclusions regarding the suitability of the system to satisfy V/STOL requirements.

Comparisons of selected laser beam power missions to conventionally powered missions

NASA Technical Reports Server (NTRS)

Bozek, John M.; Oleson, Steven R.; Landis, Geoffrey A.; Stavnes, Mark W.

1993-01-01

Earth-based laser sites beaming laser power to space assets have shown benefits over competing power system concepts for specific missions. Missions analyzed in this report that show benefits of laser beam power are low Earth orbit (LEO) to geosynchronous Earth orbit (GEO) transfer, LEO to low lunar orbit (LLO) cargo missions, and lunar-base power. Both laser- and solar-powered orbit-transfer vehicles (OTV's) make a 'tug' concept viable, which substantially reduces cumulative initial mass to LEO in comparison to chemical propulsion concepts. Lunar cargo missions utilizing laser electric propulsion from Earth-orbit to LLO show substantial mass saving to LEO over chemical propulsion systems. Lunar-base power system options were compared on a landed-mass basis. Photovoltaics with regenerative fuel cells, reactor-based systems, and laser-based systems were sized to meet a generic lunar-base power profile. A laser-based system begins to show landed mass benefits over reactor-based systems when proposed production facilities on the Moon require power levels greater than approximately 300 kWe. Benefit/cost ratios of laser power systems for an OTV, both to GEO and LLO, and for a lunar base were calculated to be greater than 1.

Propulsion engineering study for small-scale Mars missions

DOE Office of Scientific and Technical Information (OSTI.GOV)

Whitehead, J.

1995-09-12

Rocket propulsion options for small-scale Mars missions are presented and compared, particularly for the terminal landing maneuver and for sample return. Mars landing has a low propulsive {Delta}v requirement on a {approximately}1-minute time scale, but at a high acceleration. High thrust/weight liquid rocket technologies, or advanced pulse-capable solids, developed during the past decade for missile defense, are therefore more appropriate for small Mars landers than are conventional space propulsion technologies. The advanced liquid systems are characterize by compact lightweight thrusters having high chamber pressures and short lifetimes. Blowdown or regulated pressure-fed operation can satisfy the Mars landing requirement, but hardwaremore » mass can be reduced by using pumps. Aggressive terminal landing propulsion designs can enable post-landing hop maneuvers for some surface mobility. The Mars sample return mission requires a small high performance launcher having either solid motors or miniature pump-fed engines. Terminal propulsion for 100 kg Mars landers is within the realm of flight-proven thruster designs, but custom tankage is desirable. Landers on a 10 kg scale also are feasible, using technology that has been demonstrated but not previously flown in space. The number of sources and the selection of components are extremely limited on this smallest scale, so some customized hardware is required. A key characteristic of kilogram-scale propulsion is that gas jets are much lighter than liquid thrusters for reaction control. The mass and volume of tanks for inert gas can be eliminated by systems which generate gas as needed from a liquid or a solid, but these have virtually no space flight history. Mars return propulsion is a major engineering challenge; earth launch is the only previously-solved propulsion problem requiring similar or greater performance.« less

Recent advances in low-thrust propulsion technology

NASA Technical Reports Server (NTRS)

Stone, James R.

1988-01-01

The NASA low-thrust propulsion technology program is aimed at providing high performance options to a broad class of near-term and future missions. Major emphases of the program are on storable and hydrogen/oxygen low-thrust chemical, low-power (auxiliary) electrothermal, and high-power electric propulsion. This paper represents the major accomplishments of the program and discusses their impact.

Propulsion issues, options and trades

NASA Technical Reports Server (NTRS)

Forsythe, Doug J.

1986-01-01

Several different types of propulsion concepts are discussed: pulsed fission; continuous nuclear fission; chemical; and chemical boost with advanced nuclear fission. Some of the key characteristics of each type are provided, and typical concepts of each are shown.

Asteroid (Flora and Eros) sample-return missions via solar electric propulsion

NASA Technical Reports Server (NTRS)

Friedlander, A. L.

1971-01-01

The characteristics and capabilities of solar electric propulsion for performing sample-return missions to the asteroids Flora and Eros are considered. Trajectory/payload analysis and mission design tradeoff options are emphasized.

Status of Liquid Oxygen/Liquid Methane Injector Study for a Mars Ascent Engine

NASA Technical Reports Server (NTRS)

Trinh, Huu Ogyic; Cramer, John M.

1998-01-01

Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for non-toxic chemical rockets only a cryogenic propulsion system would provide high enough performance to be considered for a Mars ascent vehicle. Although the mission is possible with Earth-supplied propellants for this vehicle, utilization of in-situ propellants is highly attractive. This option would significantly reduce the overall mass of the return vehicle. Consequently, the cost of the mission would be greatly reduced because the number and size of the Earth launch vehicle(s) needed for the mission decrease. NASA/Johnson Space Center has initiated several concept studies (2) of in-situ propellant production plants. Liquid oxygen (LOX) is the primary candidate for an in-situ oxidizer. In-situ fuel candidates include methane (CH4), ethylene (C2H4), and methanol (CH3OH). MSFC initiated a technology development program for a cryogenic propulsion system for the Mars human exploration mission in 1998. One part of this technology program is the effort described here: an evaluation of propellant injection concepts for a LOX/liquid methane Mars Ascent Engine (MAE) with an emphasis on light-weight, high efficiency, reliability, and thermal compatibility. In addition to the main objective, hot-fire tests of the subject injectors will be used to test other key technologies including light-weight combustion chamber materials and advanced ignition concepts. This state-of-the-art technology will then be applied to the development of a cryogenic propulsion system that will meet the requirements of the planned Mars sample return (MSR) mission. The current baseline propulsion system for the MSR mission uses a storable propellant combination [monomethyl hydrazine/mixed oxides of nitrogen-25(MMH/MON-25)]. However, a mission option that incorporates in-situ propellant production and utilization for the ascent stage is being carefully considered as a subscale precursor to a future human mission to Mars.

Advancement of a 30K W Solar Electric Propulsion System Capability for NASA Human and Robotic Exploration Missions

NASA Technical Reports Server (NTRS)

Smith, Bryan K.; Nazario, Margaret L.; Manzella, David H.

2012-01-01

Solar Electric Propulsion has evolved into a demonstrated operational capability performing station keeping for geosynchronous satellites, enabling challenging deep-space science missions, and assisting in the transfer of satellites from an elliptical orbit Geostationary Transfer Orbit (GTO) to a Geostationary Earth Orbit (GEO). Advancing higher power SEP systems will enable numerous future applications for human, robotic, and commercial missions. These missions are enabled by either the increased performance of the SEP system or by the cost reductions when compared to conventional chemical propulsion systems. Higher power SEP systems that provide very high payload for robotic missions also trade favorably for the advancement of human exploration beyond low Earth orbit. Demonstrated reliable systems are required for human space flight and due to their successful present day widespread use and inherent high reliability, SEP systems have progressively become a viable entrant into these future human exploration architectures. NASA studies have identified a 30 kW-class SEP capability as the next appropriate evolutionary step, applicable to wide range of both human and robotic missions. This paper describes the planning options, mission applications, and technology investments for representative 30kW-class SEP mission concepts under consideration by NASA

Iodine Propulsion Advantages for Low Cost Mission Applications and the Iodine Satellite (ISAT) Technology Demonstration

NASA Technical Reports Server (NTRS)

Dankanich, John W.; Schumacher, Daniel M.

2015-01-01

The NASA Marshall Space Flight Center Science and Technology Office is continuously exploring technology options to increase performance or reduce cost and risk to future NASA missions including science and exploration. Electric propulsion is a prevalent technology known to reduce mission costs by reduction in launch costs and spacecraft mass through increased post launch propulsion performance. The exploration of alternative propellants for electric propulsion continues to be of interest to the community. Iodine testing has demonstrated comparable performance to xenon. However, iodine has a higher storage density resulting in higher ?V capability for volume constrained systems. Iodine's unique properties also allow for unpressurized storage yet sublimation with minimal power requirements to produce required gas flow rates. These characteristics make iodine an ideal propellant for secondary spacecraft. A range of mission have been evaluated with a focus on low-cost applications. Results highlight the potential for significant cost reduction over state of the art. Based on the potential, NASA has been developing the iodine Satellite for a near-term iodine Hall propulsion technology demonstration. Mission applications and progress of the iodine Satellite project are presented.

Nuclear electric power for multimegawatt orbit transfer vehicles

NASA Astrophysics Data System (ADS)

Casagrande, R. D.

Multimegawatt nuclear propulsion is an attractive option for orbit transfer vehicles. The masses of these platforms are expected to exceed the capability of a single launch from Earth necessitating assembly in space in a parking orbit. The OTV would transfer the platform from the parking orbit to the operational orbit and then return for the next mission. Electric propulsion is advantageous because of the high specific impulse achieved by the technology, 1000 to 5000 s and beyond, to reduce the propellant required. Nuclear power is attractive as the power system because of the weight savings over solar systems in the multimegawatt regime, and multimegawatts of power are required. A conceptual diagram is shown of an OTV with a command control module using electric thrusters powered from an SP-100 class nuclear reactor power system.

Space transfer concepts and analyses for exploration missions, phase 4

NASA Technical Reports Server (NTRS)

Woodcock, Gordon R.

1993-01-01

Earlier studies carried out under this contract covered a wide range of lunar and Mars transportation options, and lunar rove concepts and technology needs. The current report discusses the activities conducted under Technical Directives 16 and 17. Mars transportation was addressed as well as a review and update of architectures and propulsion systems.

An Overview of Facilities and Capabilities to Support the Development of Nuclear Thermal Propulsion

DOE Office of Scientific and Technical Information (OSTI.GOV)

James Werner; Sam Bhattacharyya; Mike Houts

Abstract. The future of American space exploration depends on the ability to rapidly and economically access locations of interest throughout the solar system. There is a large body of work (both in the US and the Former Soviet Union) that show that Nuclear Thermal Propulsion (NTP) is the most technically mature, advanced propulsion system that can enable this rapid and economical access by its ability to provide a step increase above what is a feasible using a traditional chemical rocket system. For an NTP system to be deployed, the earlier measurements and recent predictions of the performance of the fuelmore » and the reactor system need to be confirmed experimentally prior to launch. Major fuel and reactor system issues to be addressed include fuel performance at temperature, hydrogen compatibility, fission product retention, and restart capability. The prime issue to be addressed for reactor system performance testing involves finding an affordable and environmentally acceptable method to test a range of engine sizes using a combination of nuclear and non-nuclear test facilities. This paper provides an assessment of some of the capabilities and facilities that are available or will be needed to develop and test the nuclear fuel, and reactor components. It will also address briefly options to take advantage of the greatly improvement in computation/simulation and materials processing capabilities that would contribute to making the development of an NTP system more affordable. Keywords: Nuclear Thermal Propulsion (NTP), Fuel fabrication, nuclear testing, test facilities.« less

Nuclear Engine System Simulation (NESS). Volume 1: Program user's guide

NASA Astrophysics Data System (ADS)

Pelaccio, Dennis G.; Scheil, Christine M.; Petrosky, Lyman J.

1993-03-01

A Nuclear Thermal Propulsion (NTP) engine system design analysis tool is required to support current and future Space Exploration Initiative (SEI) propulsion and vehicle design studies. Currently available NTP engine design models are those developed during the NERVA program in the 1960's and early 1970's and are highly unique to that design or are modifications of current liquid propulsion system design models. To date, NTP engine-based liquid design models lack integrated design of key NTP engine design features in the areas of reactor, shielding, multi-propellant capability, and multi-redundant pump feed fuel systems. Additionally, since the SEI effort is in the initial development stage, a robust, verified NTP analysis design tool could be of great use to the community. This effort developed an NTP engine system design analysis program (tool), known as the Nuclear Engine System Simulation (NESS) program, to support ongoing and future engine system and stage design study efforts. In this effort, Science Applications International Corporation's (SAIC) NTP version of the Expanded Liquid Engine Simulation (ELES) program was modified extensively to include Westinghouse Electric Corporation's near-term solid-core reactor design model. The ELES program has extensive capability to conduct preliminary system design analysis of liquid rocket systems and vehicles. The program is modular in nature and is versatile in terms of modeling state-of-the-art component and system options as discussed. The Westinghouse reactor design model, which was integrated in the NESS program, is based on the near-term solid-core ENABLER NTP reactor design concept. This program is now capable of accurately modeling (characterizing) a complete near-term solid-core NTP engine system in great detail, for a number of design options, in an efficient manner. The following discussion summarizes the overall analysis methodology, key assumptions, and capabilities associated with the NESS presents an example problem, and compares the results to related NTP engine system designs. Initial installation instructions and program disks are in Volume 2 of the NESS Program User's Guide.

Nuclear Engine System Simulation (NESS). Volume 1: Program user's guide

NASA Technical Reports Server (NTRS)

Pelaccio, Dennis G.; Scheil, Christine M.; Petrosky, Lyman J.

1993-01-01

A Nuclear Thermal Propulsion (NTP) engine system design analysis tool is required to support current and future Space Exploration Initiative (SEI) propulsion and vehicle design studies. Currently available NTP engine design models are those developed during the NERVA program in the 1960's and early 1970's and are highly unique to that design or are modifications of current liquid propulsion system design models. To date, NTP engine-based liquid design models lack integrated design of key NTP engine design features in the areas of reactor, shielding, multi-propellant capability, and multi-redundant pump feed fuel systems. Additionally, since the SEI effort is in the initial development stage, a robust, verified NTP analysis design tool could be of great use to the community. This effort developed an NTP engine system design analysis program (tool), known as the Nuclear Engine System Simulation (NESS) program, to support ongoing and future engine system and stage design study efforts. In this effort, Science Applications International Corporation's (SAIC) NTP version of the Expanded Liquid Engine Simulation (ELES) program was modified extensively to include Westinghouse Electric Corporation's near-term solid-core reactor design model. The ELES program has extensive capability to conduct preliminary system design analysis of liquid rocket systems and vehicles. The program is modular in nature and is versatile in terms of modeling state-of-the-art component and system options as discussed. The Westinghouse reactor design model, which was integrated in the NESS program, is based on the near-term solid-core ENABLER NTP reactor design concept. This program is now capable of accurately modeling (characterizing) a complete near-term solid-core NTP engine system in great detail, for a number of design options, in an efficient manner. The following discussion summarizes the overall analysis methodology, key assumptions, and capabilities associated with the NESS presents an example problem, and compares the results to related NTP engine system designs. Initial installation instructions and program disks are in Volume 2 of the NESS Program User's Guide.

Solar power satellite. System definition study. Part 1, volume 5: SPS transportation. Representative system descriptions

NASA Technical Reports Server (NTRS)

1977-01-01

Both LEO transportation (earth to low earth orbit) and GEO transportation (low earth orbit to geosynchronous orbit) segments were addressed. LEO options include both a 2 stage winged space freighter vehicle and a 2 stage ballistic recoverable vehicle. Both incorporate LO(2)/RP-1/LH(2) engines on the booster and standard SSME's on the upper stage. The orbit transfer vehicle options included chemical for geosynchronous satellite assembly and self powered electric propulsion for low earth orbit satellite assembly. An exhaust products analysis was conducted for the earth to LEO vehicle since atmospheric pollution could be a concern.

Interplanetary Electric Propulsion Uranus Mission Trades Supporting the Decadal Survey

NASA Technical Reports Server (NTRS)

Dankanich, John W.; McAdams, James

2011-01-01

The Decadal Survey Committee was tasked to develop a comprehensive science and mission strategy for planetary science that updates and extends the National Academies Space Studies Board s current solar system exploration decadal survey. A Uranus orbiter mission has been evaluated as a part of this 2013-2022 Planetary Science Decadal Survey. A comprehensive Uranus orbiter mission design was completed, including a broad search of interplanetary electric propulsion transfer options. The scope of interplanetary trades was limited to electric propulsion concepts, both solar and radioisotope powered. Solar electric propulsion offers significant payloads to Uranus. Inserted mass into the initial science orbit due is highly sensitive to transfer time due to arrival velocities. The recommended baseline trajectory is a 13 year transfer with an Atlas 551, a 1+1 NEXT stage with 15 kW of power using an EEJU trajectory and a 1,000km EGA flyby altitude constraint. This baseline delivers over 2,000kg into the initial science orbit. Interplanetary trajectory trades and sensitivity analyses are presented herein.

Natural Gas Propulsion Options for Short Sea Shipping Routes

DTIC Science & Technology

2010-08-01

that are involved with gas and the relevant systems, along with personal protection issues that come into effect when handling both compressed and...a compressed air system for engine starting, which is stored in compressed air storage cylinders . The system leads compressed air through a valve...directly into the cylinder heads at 30bar to begin the rotation of the engine. After this rotation occurs, the engine is supplied with diesel fuel

Nuclear Thermal Rocket (NTR) Propulsion and Power Systems for Outer Planetary Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, S. K.; Cataldo, R. L.

2001-01-01

The high specific impulse (I (sub sp)) and engine thrust generated using liquid hydrogen (LH2)-cooled Nuclear Thermal Rocket (NTR) propulsion makes them attractive for upper stage applications for difficult robotic science missions to the outer planets. Besides high (I (sub sp)) and thrust, NTR engines can also be designed for "bimodal" operation allowing substantial amounts of electrical power (10's of kWe ) to be generated for onboard spacecraft systems and high data rate communications with Earth during the course of the mission. Two possible options for using the NTR are examined here. A high performance injection stage utilizing a single 15 klbf thrust engine can inject large payloads to the outer planets using a 20 t-class launch vehicle when operated in an "expendable mode". A smaller bimodal NTR stage generating approx. 1 klbf of thrust and 20 to 40 kWe for electric propulsion can deliver approx. 100 kg using lower cost launch vehicles. Additional information is contained in the original extended abstract.

Lightweight Damage Tolerant Radiators for In-Space Nuclear Electric Power and Propulsion

NASA Technical Reports Server (NTRS)

Craven, Paul; SanSoucie, Michael P.; Tomboulian, Briana; Rogers, Jan; Hyers, Robert

2014-01-01

Nuclear electric propulsion (NEP) is a promising option for high-speed in-space travel due to the high energy density of nuclear power sources and efficient electric thrusters. Advanced power conversion technologies for converting thermal energy from the reactor to electrical energy at high operating temperatures would benefit from lightweight, high temperature radiator materials. Radiator performance dictates power output for nuclear electric propulsion systems. Pitch-based carbon fiber materials have the potential to offer significant improvements in operating temperature and mass. An effort at the NASA Marshall Space Flight Center to show that woven high thermal conductivity carbon fiber mats can be used to replace standard metal and composite radiator fins to dissipate waste heat from NEP systems is ongoing. The goals of this effort are to demonstrate a proof of concept, to show that a significant improvement of specific power (power/mass) can be achieved, and to develop a thermal model with predictive capabilities. A description of this effort is presented.

A unique nuclear thermal rocket engine using a particle bed reactor

NASA Astrophysics Data System (ADS)

Culver, Donald W.; Dahl, Wayne B.; McIlwain, Melvin C.

1992-01-01

Aerojet Propulsion Division (APD) studied 75-klb thrust Nuclear Thermal Rocket Engines (NTRE) with particle bed reactors (PBR) for application to NASA's manned Mars mission and prepared a conceptual design description of a unique engine that best satisfied mission-defined propulsion requirements and customer criteria. This paper describes the selection of a sprint-type Mars transfer mission and its impact on propulsion system design and operation. It shows how our NTRE concept was developed from this information. The resulting, unusual engine design is short, lightweight, and capable of high specific impulse operation, all factors that decrease Earth to orbit launch costs. Many unusual features of the NTRE are discussed, including nozzle area ratio variation and nozzle closure for closed loop after cooling. Mission performance calculations reveal that other well known engine options do not support this mission.

Solar Electric Propulsion Vehicle Demonstration to Support Future Space Exploration Missions

NASA Technical Reports Server (NTRS)

Smith, Bryan K.; Nazario, Margaret L.; Cunningham, Cameron C.

2012-01-01

Human and robotic exploration beyond Low Earth Orbit (LEO) will require enabling capabilities that are efficient, affordable, and reliable. Solar Electric Propulsion (SEP) is highly advantageous because of its favorable in-space mass transfer efficiency compared to traditional chemical propulsion systems. The NASA studies have demonstrated that this advantage becomes highly significant as missions progress beyond Earth orbit. Recent studies of human exploration missions and architectures evaluated the capabilities needed to perform a variety of human exploration missions including missions to Near Earth Objects (NEOs). The studies demonstrated that SEP stages have potential to be the most cost effective solution to perform beyond LEO transfers of high mass cargoes for human missions. Recognizing that these missions require power levels more than 10X greater than current electric propulsion systems, NASA embarked upon a progressive pathway to identify critical technologies needed and a plan for an incremental demonstration mission. The NASA studies identified a 30kW class demonstration mission that can serve as a meaningful demonstration of the technologies, operational challenges, and provide the appropriate scaling and modularity required. This paper describes the planning options for a representative demonstration 30kW class SEP mission.

Hydrogen Fuel System Design Trades for High-Altitude Long-Endurance Remotely- Operated Aircraft

NASA Technical Reports Server (NTRS)

Millis, Marc G.; Tornabene, Robert T.; Jurns, John M.; Guynn, Mark D.; Tomsik, Thomas M.; VanOverbeke, Thomas J.

2009-01-01

Preliminary design trades are presented for liquid hydrogen fuel systems for remotely-operated, high-altitude aircraft that accommodate three different propulsion options: internal combustion engines, and electric motors powered by either polymer electrolyte membrane fuel cells or solid oxide fuel cells. Mission goal is sustained cruise at 60,000 ft altitude, with duration-aloft a key parameter. The subject aircraft specifies an engine power of 143 to 148 hp, gross liftoff weight of 9270 to 9450 lb, payload of 440 lb, and a hydrogen fuel capacity of 2650 to 2755 lb stored in two spherical tanks (8.5 ft inside diameter), each with a dry mass goal of 316 lb. Hydrogen schematics for all three propulsion options are provided. Each employs vacuum-jacketed tanks with multilayer insulation, augmented with a helium pressurant system, and using electric motor driven hydrogen pumps. The most significant schematic differences involve the heat exchangers and hydrogen reclamation equipment. Heat balances indicate that mission durations of 10 to 16 days appear achievable. The dry mass for the hydrogen system is estimated to be 1900 lb, including 645 lb for each tank. This tank mass is roughly twice that of the advanced tanks assumed in the initial conceptual vehicle. Control strategies are not addressed, nor are procedures for filling and draining the tanks.

Nuclear electric propulsion: A better, safer, cheaper transportation system for human exploration of Mars

NASA Technical Reports Server (NTRS)

Clark, John S.; George, Jeffrey A.; Gefert, Leon P.; Doherty, Michael P.; Sefcik, Robert J.

1994-01-01

NASA has completed a preliminary mission and systems study of nuclear electric propulsion (NEP) systems for 'split-sprint' human exploration and related robotic cargo missions to Mars. This paper describes the study, the mission architecture selected, the NEP system and technology development needs, proposed development schedules, and estimated development costs. Since current administration policy makers have delayed funding for key technology development activities that could make Mars exploration missions a reality in the near future, NASA will have time to evaluate various alternate mission options, and it appears prudent to ensure that Mars mission plans focus on astronaut and mission safety, while reducing costs to acceptable levels. The split-sprint nuclear electric propulsion system offers trip times comparable to nuclear thermal propulsion (NTP) systems, while providing mission abort opportunities that are not possible with 'reference' mission architectures. Thus, NEP systems offer short transit times for the astronauts, reducing the exposure of the crew to intergalactic cosmic radiation. The high specific impulse of the NEP system, which leads to very low propellant requirements, results in significantly lower 'initial mass in low earth orbit' (IMLEO). Launch vehicle packaging studies show that the NEP system can be launched, assembled, and deployed, with about one less 240-metric-ton heavy lift launch vehicle (HLLV) per mission opportunity - a very Technology development cost of the nuclear reactor for an NEP system would be shared with the proposed nuclear surface power systems, since nuclear systems will be required to provide substantial electrical power on the surface of Mars. The NEP development project plan proposed includes evolutionary technology development for nuclear electric propulsion systems that expands upon SP-100 (Space Power - 100 kw(e)) technology that has been developed for lunar and Mars surface nuclear power, and small NEP systems for interplanetary probes. System upgrades are expected to evolve that will result in even shorter trip times, improved payload capabilities, and enhanced safety and reliability.

Missions to the Outer Solar System and Beyond - Concept Study for a Kuiper Belt Sample-Return

NASA Astrophysics Data System (ADS)

Ganapathy, Rohan M.

The exploration of Kuiper belt objects (KBOs) might deliver crucial data for answering questions about the evolution of the solar system and the origin of life. Whereas the current New Horizons mission performs a flyby at KBOs, an in-depth exploration of the Kuiper belt requires an orbiter, lander or even a sample return. In this paper, we present a range of potential mission architectures for a Kuiper belt sample return mission. We use the Systems Modeling Language (SysML) for the necessary modeling and the systems engineering tool MagicDraw. A process similar to the NASA Rapid Mission Architecture approach was used. We start with a rationale a KBO sample return, dene science objectives, high-level requirements and select a strawman payload. From a key trade-matrix, mission architecture options are generated. Finally, necessary technologies and prerequisites for the mission are identied. We conclude that one of the dwarf planets Pluto, Haumea, Orcus or Quaoar and their moons should be considered as a target for the mission. The samples should be collected from the dwarf planet of choice or from its moon(s), which omits the rather high velocity requirements for a landing and departure from the dwarf planet itself. Attractive mission architectures include radioisotopic electric propulsion-based missions, missions with a combination of a solar electric propulsion stage and radioisotopic electric propulsion, or missions using nuclear electric propulsion.

The Rationale/Benefits of Nuclear Thermal Rocket Propulsion for NASA's Lunar Space Transportation System

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

1994-01-01

The solid core nuclear thermal rocket (NTR) represents the next major evolutionary step in propulsion technology. With its attractive operating characteristics, which include high specific impulse (approximately 850-1000 s) and engine thrust-to-weight (approximately 4-20), the NTR can form the basis for an efficient lunar space transportation system (LTS) capable of supporting both piloted and cargo missions. Studies conducted at the NASA Lewis Research Center indicate that an NTR-based LTS could transport a fully-fueled, cargo-laden, lunar excursion vehicle to the Moon, and return it to low Earth orbit (LEO) after mission completion, for less initial mass in LEO than an aerobraked chemical system of the type studied by NASA during its '90-Day Study.' The all-propulsive NTR-powered LTS would also be 'fully reusable' and would have a 'return payload' mass fraction of approximately 23 percent--twice that of the 'partially reusable' aerobraked chemical system. Two NTR technology options are examined--one derived from the graphite-moderated reactor concept developed by NASA and the AEC under the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) programs, and a second concept, the Particle Bed Reactor (PBR). The paper also summarizes NASA's lunar outpost scenario, compares relative performance provided by different LTS concepts, and discusses important operational issues (e.g., reusability, engine 'end-of life' disposal, etc.) associated with using this important propulsion technology.

Main-belt asteroid exploration - Mission options for the 1990s

NASA Technical Reports Server (NTRS)

Yen, C.-W. L.

1982-01-01

Mission configurations, propulsion systems, and target bodies for possible NASA asteroid exploration projects are examined. Noting that an announced delay in the development of a solar electric propulsion system has led to a consideration of chemical rocket systems, asteroid missions are grouped in terms of five potential areas for investigation, each successively further from the sun. The Shuttle-launched IUS is suggested as the prime candidate for boosting probes into trajectories for asteroid rendezvous with a number of the 3000 known asteroids. Planetary swingbys are mentioned as the only suitable method for satisfying the large energy requirements of the asteroid missions. Performance analyses are presented of the IUS 2-stage/Star-48 and Centaur vehicles, and sample missions to Fortuna, Anahita, and Urania in 1990 and further missions to the middle, outer, and Trojans asteroids are outlined.

Small Fast Spectrum Reactor Designs Suitable for Direct Nuclear Thermal Propulsion

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bruce G. Schnitzler; Stanley K. Borowski

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. Past studies, in particular those in support of both the Strategic Defense Initiative (SDI) and Space Exploration Initiative (SEI), have shown nuclear thermal propulsion systems provide superior performance for high mass high propulsive delta-V missions. The recent NASA Design Reference Architecture (DRA) 5.0 Study re-examined mission, payload, and transportation system requirements for a human Mars landing mission in the post-2030 timeframe. Nuclear thermal propulsion was again identified asmore » the preferred in-space transportation system. A common nuclear thermal propulsion stage with three 25,000-lbf thrust engines was used for all primary mission maneuvers. Moderately lower thrust engines may also have important roles. In particular, lower thrust engine designs demonstrating the critical technologies that are directly extensible to other thrust levels are attractive from a ground testing perspective. An extensive nuclear thermal rocket technology development effort was conducted from 1955-1973 under the Rover/NERVA Program. Both graphite and refractory metal alloy fuel types were pursued. Reactors and engines employing graphite based fuels were designed, built and ground tested. A number of fast spectrum reactor and engine designs employing refractory metal alloy fuel types were proposed and designed, but none were built. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art graphite based fuel design incorporating lessons learned from the very successful technology development program. The SNRE was a nominal 16,000-lbf thrust engine originally intended for unmanned applications with relatively short engine operations and the engine and stage design were constrained to fit within the payload volume of the then planned space shuttle. The SNRE core design utilized hexagonal fuel elements and hexagonal structural support elements. The total number of elements can be varied to achieve engine designs of higher or lower thrust levels. Some variation in the ratio of fuel elements to structural elements is also possible. Options for SNRE-based engine designs in the 25,000-lbf thrust range were described in a recent (2010) Joint Propulsion Conference paper. The reported designs met or exceeded the performance characteristics baselined in the DRA 5.0 Study. Lower thrust SNRE-based designs were also described in a recent (2011) Joint Propulsion Conference paper. Recent activities have included parallel evaluation and design efforts on fast spectrum engines employing refractory metal alloy fuels. These efforts include evaluation of both heritage designs from the Argonne National Laboratory (ANL) and General Electric Company GE-710 Programs as well as more recent designs. Results are presented for a number of not-yet optimized fast spectrum engine options.« less

Small Fast Spectrum Reactor Designs Suitable for Direct Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Schnitzler, Bruce G.; Borowski, Stanley K.

2012-01-01

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. Past studies, in particular those in support of the Space Exploration Initiative (SEI), have shown nuclear thermal propulsion systems provide superior performance for high mass high propulsive delta-V missions. The recent NASA Design Reference Architecture (DRA) 5.0 Study re-examined mission, payload, and transportation system requirements for a human Mars landing mission in the post-2030 timeframe. Nuclear thermal propulsion was again identified as the preferred in-space transportation system. A common nuclear thermal propulsion stage with three 25,000-lbf thrust engines was used for all primary mission maneuvers. Moderately lower thrust engines may also have important roles. In particular, lower thrust engine designs demonstrating the critical technologies that are directly extensible to other thrust levels are attractive from a ground testing perspective. An extensive nuclear thermal rocket technology development effort was conducted from 1955-1973 under the Rover/NERVA Program. Both graphite and refractory metal alloy fuel types were pursued. Reactors and engines employing graphite based fuels were designed, built and ground tested. A number of fast spectrum reactor and engine designs employing refractory metal alloy fuel types were proposed and designed, but none were built. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art graphite based fuel design incorporating lessons learned from the very successful technology development program. The SNRE was a nominal 16,000-lbf thrust engine originally intended for unmanned applications with relatively short engine operations and the engine and stage design were constrained to fit within the payload volume of the then planned space shuttle. The SNRE core design utilized hexagonal fuel elements and hexagonal structural support elements. The total number of elements can be varied to achieve engine designs of higher or lower thrust levels. Some variation in the ratio of fuel elements to structural elements is also possible. Options for SNRE-based engine designs in the 25,000-lbf thrust range were described in a recent (2010) Joint Propulsion Conference paper. The reported designs met or exceeded the performance characteristics baselined in the DRA 5.0 Study. Lower thrust SNRE-based designs were also described in a recent (2011) Joint Propulsion Conference paper. Recent activities have included parallel evaluation and design efforts on fast spectrum engines employing refractory metal alloy fuels. These efforts include evaluation of both heritage designs from the Argonne National Laboratory (ANL) and General Electric Company GE-710 Programs as well as more recent designs. Results are presented for a number of not-yet optimized fast spectrum engine options.

Development of Modeling Approaches for Nuclear Thermal Propulsion Test Facilities

NASA Technical Reports Server (NTRS)

Jones, Daniel R.; Allgood, Daniel C.; Nguyen, Ke

2014-01-01

High efficiency of rocket propul-sion systems is essential for humanity to venture be-yond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rock-ets with relatively high thrust and twice the efficiency of the Space Shuttle Main Engine. NASA is in the pro-cess of developing a new NTP engine, and is evaluat-ing ground test facility concepts that allow for the thor-ough testing of NTP devices. NTP engine exhaust, hot gaseous hydrogen, is nominally expected to be free of radioactive byproducts from the nuclear reactor; how-ever, it has the potential to be contaminated due to off-nominal engine reactor performance. Several options are being investigated to mitigate this hazard potential with one option in particular that completely contains the engine exhaust during engine test operations. The exhaust products are subsequently disposed of between engine tests. For this concept (see Figure 1), oxygen is injected into the high-temperature hydrogen exhaust that reacts to produce steam, excess oxygen and any trace amounts of radioactive noble gases released by off-nominal NTP engine reactor performance. Water is injected to condense the potentially contaminated steam into water. This water and the gaseous oxygen (GO2) are subsequently passed to a containment area where the water and GO2 are separated into separate containment tanks.

Viability of a Reusable In-Space Transportation System

NASA Technical Reports Server (NTRS)

Jefferies, Sharon A.; McCleskey, Carey M.; Nufer, Brian M.; Lepsch, Roger A.; Merrill, Raymond G.; North, David D.; Martin, John G.; Komar, David R.

2015-01-01

The National Aeronautics and Space Administration (NASA) is currently developing options for an Evolvable Mars Campaign (EMC) that expands human presence from Low Earth Orbit (LEO) into the solar system and to the surface of Mars. The Hybrid in-space transportation architecture is one option being investigated within the EMC. The architecture enables return of the entire in-space propulsion stage and habitat to cis-lunar space after a round trip to Mars. This concept of operations opens the door for a fully reusable Mars transportation system from cis-lunar space to a Mars parking orbit and back. This paper explores the reuse of in-space transportation systems, with a focus on the propulsion systems. It begins by examining why reusability should be pursued and defines reusability in space-flight context. A range of functions and enablers associated with preparing a system for reuse are identified and a vision for reusability is proposed that can be advanced and implemented as new capabilities are developed. Following this, past reusable spacecraft and servicing capabilities, as well as those currently in development are discussed. Using the Hybrid transportation architecture as an example, an assessment of the degree of reusability that can be incorporated into the architecture with current capabilities is provided and areas for development are identified that will enable greater levels of reuse in the future. Implications and implementation challenges specific to the architecture are also presented.

NASA Hypersonic Propulsion: Overview of Progress from 1995 to 2005

NASA Technical Reports Server (NTRS)

Cikanek, Harry A., III; Bartolotta, Paul A.; Klem, Mark D.; Rausch, Vince L.

2007-01-01

Hypersonic propulsion work supported by the United States National Aeronautics and Space Administration had a primary focus on Space Transportation during the period from 1995 to 2005. The framework for these advances was established by policy and pursued with substantial funding. Many noteworthy advances were made, highlighted by the pinnacle flights of the X-43. This paper reviews and summarizes the programs and accomplishments of this era. The accomplishments are compared to the goals and objectives to lend an overarching perspective to what was achieved. At least dating back to the early days of the Space Shuttle program, NASA has had the objective of reducing the cost of access to space and concurrently improving safety and reliability. National Space Transportation Policy in 1994 coupled with a base of prior programs such as the National Aerospace Plane and the need to look beyond the Space Shuttle program set the stage for NASA to pursue Space Transportation Advances. Programs defined to pursue the advances represented a broad approach addressing classical rocket propulsion as well as airbreathing propulsion in various combinations and forms. The resulting portfolio of activities included systems analysis and design studies, discipline research and technology, component technology development, propulsion system ground test demonstration and flight demonstration. The types of propulsion systems that were pursued by these programs included classical rocket engines, "aerospike" rocket engines, high performance rocket engines, scram jets, rocket based combined cycles, and turbine based combined cycles. Vehicle architectures included single and two stage vehicles. Either single types of propulsion systems or combinations of the basic propulsion types were applied to both single and two stage vehicle design concepts. Some of the propulsion system design concepts were built and tested at full scale, large scale and small scale. Many flight demonstrators were conceptually defined, fewer designed and some built and one flown to demonstrate several technical advancements including propulsion. The X-43 flights were a culmination of these efforts for airbreathing propulsion. During the course of that period, there was a balance of funding and emphasis toward rocket propulsion but still very substantial airbreathing propulsion effort. The broad objectives of these programs were to both advance and test the state of the art so as to provide a basis for options to be pursued for broad space transportation needs, most importantly focused on crew carrying capability. NASA cooperated with the Department of Defense in planning and implementation of these programs to make efficient use of objectives and capabilities where appropriate. Much of the work was conducted in industry and academia as well as Government laboratories. Many test articles and data-bases now exist as a result of this work. At the conclusion of the period, the body of work made it clear that continued research and technology development was warranted, because although not ready for a NASA system development decision, results continued to support the promise of air-breathing propulsion for access to space.

The Hybrid Propellant Module (HPM): A New Concept for Space Transfer in the Earth's Neighborhood and Beyond

NASA Technical Reports Server (NTRS)

Mankins, John C.; Mazanek, Daniel D.

2001-01-01

The safe, affordable and effective transfer of ever-larger payloads and eventually personnel beyond Low Earth Orbit (LEO) is a major challenge facing future commercial development and human exploration of space. Without reusable systems, sustained exploration or large scale development beyond LEO appears to be economically non-viable. However, reusable systems must be capable of both good fuel efficiency and "high utilization of capacity", or else economic costs will remain unacceptably high. Various options exist that can provide high fuel efficiency - for example, Solar Electric Propulsion Systems (SEPS) - but only at the cost of low thrust and concomitant long transit times. Chemical propulsion systems offer the potential for high thrust and short transit times - including both cryogenic and non-cryogenic options - but only at the cost of relatively low specific impulse (Isp). Nuclear thermal propulsion systems offer relatively good thrust-to-weight and Isp - but involve public concerns that may be insurmountable for all except the most-critical of public purposes. Fixed infrastructures have been suggested as one approach to solving this challenge; for example, rotating tether approaches. However, these systems tend to suffer from high initial costs or unacceptable operational constraints. A new concept has been identified - the Hybrid Propellant Module (HPM) - that integrates the best features of both chemical and solar electric transportation architectures. The HPM approach appears to hold promise of solving the issues associated with other approaches, opening a new family of capabilities for future space exploration and development of near-Earth space and beyond. This paper provides a summary overview of the challenge of Earth neighborhood transportation and discusses how various systems concepts might be applied to meet the needs of these architectures. The paper describes a new approach, the HPM, and illustrates the application of the concept for a typical mission concept. The paper concludes with a discussion of needed technologies and a possible timeline for the development and evolution of this class of systems concepts.

A Power Conversion Concept for the Jupiter Icy Moons Orbiter

NASA Technical Reports Server (NTRS)

Mason, Lee S.

2003-01-01

The Jupiter Icy Moons Orbiter (JIMO) mission is currently under study by the Office of Space Science under the Project Prometheus Program. JIMO is examining the use of Nuclear Electric Propulsion (NEP) to carry scientific payloads to three Jovian moons. A potential power system concept includes dual 100 kWe Brayton converters, a deployable pumped loop heat rejection subsystem, and a 400 Vac Power Management and Distribution (PMAD) bus. Many trades were performed in aniving at this candidate power system concept. System-level studies examined design and off-design operating modes, determined startup requirements, evaluated subsystem redundancy options, and quantified the mass and radiator area of reactor power systems from 20 to 200 kWe. In the Brayton converter subsystem, studies were performed to investigate converter packaging options, and assess the induced torque effects on spacecraft dynamics due to rotating machinery. In the heat rejection subsystem, design trades were conducted on heat transport approaches, material and fluid options, and deployed radiator geometries. In the PMAD subsystem, the overall electrical architecture was defined and trade studies examined distribution approaches, voltage levels, and cabling options.

Mission and system optimization of nuclear electric propulsion vehicles for lunar and Mars missions

NASA Technical Reports Server (NTRS)

Gilland, James H.

1991-01-01

The detailed mission and system optimization of low thrust electric propulsion missions is a complex, iterative process involving interaction between orbital mechanics and system performance. Through the use of appropriate approximations, initial system optimization and analysis can be performed for a range of missions. The intent of these calculations is to provide system and mission designers with simple methods to assess system design without requiring access or detailed knowledge of numerical calculus of variations optimizations codes and methods. Approximations for the mission/system optimization of Earth orbital transfer and Mars mission have been derived. Analyses include the variation of thruster efficiency with specific impulse. Optimum specific impulse, payload fraction, and power/payload ratios are calculated. The accuracy of these methods is tested and found to be reasonable for initial scoping studies. Results of optimization for Space Exploration Initiative lunar cargo and Mars missions are presented for a range of power system and thruster options.

Highlights of Nanosatellite Development Program at NASA-Goddard Space Flight Center

NASA Technical Reports Server (NTRS)

Rhee, Michael S.; Zakrzwski, Chuck M.; Thomas, Mike A.; Bauer, Frank H. (Technical Monitor)

2000-01-01

Currently the GN&C's Propulsion Branch of the NASA's Goddard Space Flight Center (GSFC) is conducting a broad technology development program for propulsion devices that are ideally suited for nanosatellite missions. The goal of our program is to develop nanosatellite propulsion systems that can be flight qualified in a few years and flown in support of nanosatellite missions. The miniature cold gas thruster technology, the first product from the GSFC's propulsion component technology development program, will be flown on the upcoming ST-5 mission in 2003. The ST-5 mission is designed to validate various nanosatellite technologies in all major subsystem areas. It is a precursor mission to more ambitious nanosatellite missions such as the Magnetospheric Constellation mission. By teaming with the industry and government partners, the GSFC propulsion component technology development program is aimed at pursuing a multitude of nanosatellite propulsion options simultaneously, ranging from miniaturized thrusters based on traditional chemical engines to MEMS based thruster systems. After a conceptual study phase to determine the feasibility and the applicability to nanosatellite missions, flight like prototypes of selected technology are fabricated for testing. The development program will further narrow down the effort to those technologies that are considered "mission-enabling" for future nanosatellite missions. These technologies will be flight qualified to be flown on upcoming nanosatellite missions. This paper will report on the status of our development program and provide details on the following technologies: Low power miniature cold gas thruster Nanosatellite solid rocket motor. Solid propellant gas generator system for cold gas thruster. Low temperature hydrazine blends for miniature hydrazine thruster. MEMS mono propellant thruster using hydrogen peroxide.

Multi-Thruster Propulsion Apparatus

NASA Technical Reports Server (NTRS)

Patterson, Michael J. (Inventor)

2016-01-01

An electric propulsion machine includes an ion thruster having a discharge chamber housing a large surface area anode. The ion thruster includes flat annular ion optics with a small span to gap ratio. Optionally, at least a second thruster may be disposed radially offset from the ion thruster.

Dual-fuel propulsion - Why it works, possible engines, and results of vehicle studies. [on earth-to-orbit Space Shuttle flights

NASA Technical Reports Server (NTRS)

Martin, J. A.; Wilhite, A. W.

1979-01-01

The reasons why dual-fuel propulsion works are discussed. Various engine options are discussed, and vehicle mass and cost results are presented for earth-to-orbit vehicles. The results indicate that dual-fuel propulsion is attractive, particularly with the dual-expander engine. A unique orbit-transfer vehicle is described which uses dual-fuel propulsion. One Space Shuttle flight and one flight of a heavy-lift Shuttle derivative are used for each orbit-transfer vehicle flight, and the payload capability is quite attractive.

Joint Radioisotope Electric Propulsion Studies - Neptune System Explorer

NASA Technical Reports Server (NTRS)

Khan, M. Omair; Amini, Rashied; Ervin, Joan; Lang, Jared; Landau, Damon; Oleson, Steven; Spilker, Thomas; Strange, Nathan

2011-01-01

The Neptune System Explorer (NSE) mission concept study assessed opportunities to conduct Cassini-like science at Neptune with a radioisotope electric propulsion (REP) based spacecraft. REP is based on powering an electric propulsion (EP) engine with a radioisotope power source (RPS). The NSE study was commissioned under the Joint Radioisotope Electric Propulsion Studies (JREPS) project, which sought to determine the technical feasibility of flagship class REP applications. Within JREPS, special emphasis was given toward identifying tall technology tent poles, as well as recommending any new RPS technology developments that would be required for complicated REP missions. Based on the goals of JREPS, multiple RPS (e.g. thermoelectric and Stirling based RPS) and EP (e.g. Hall and ion engines) technology combinations were traded during the NSE study to determine the most favorable REP design architecture. Among the findings from the study was the need for >400We RPS systems, which was driven by EP operating powers and the requirement for a long-lived mission in the deep solar system. Additionally multiple development and implementation risks were identified for the NSE concept, as well as REP missions in general. Among the strengths of the NSE mission would be the benefits associated with RPS and EP use, such as long-term power (approx. 2-3kW) at Neptune and flexible trajectory options for achieving orbit or tours of the Neptune system. Although there are still multiple issues to mitigate, the NSE concept demonstrated distinct advantages associated with using REP for deep space flagship-class missions.

Current Development of Nuclear Thermal Propulsion technologies at the Center for Space Nuclear Research

DOE Office of Scientific and Technical Information (OSTI.GOV)

Robert C. O'Brien; Steven K. Cook; Nathan D. Jerred

Nuclear power and propulsion has been considered for space applications since the 1950s. Between 1955 and 1972 the US built and tested over twenty nuclear reactors / rocket engines in the Rover/NERVA programs1. The Aerojet Corporation was the prime contractor for the NERVA program. Modern changes in environmental laws present challenges for the redevelopment of the nuclear rocket. Recent advances in fuel fabrication and testing options indicate that a nuclear rocket with a fuel composition that is significantly different from those of the NERVA project can be engineered; this may be needed to ensure public support and compliance with safetymore » requirements. The Center for Space Nuclear Research (CSNR) is pursuing a number of technologies, modeling and testing processes to further the development of safe, practical and affordable nuclear thermal propulsion systems.« less

Manned Mars Explorer project: Guidelines for a manned mission to the vicinity of Mars using Phobos as a staging outpost; schematic vehicle designs considering chemical and nuclear electric propulsion

NASA Technical Reports Server (NTRS)

Nolan, Sean; Neubek, Deb; Baxmann, C. J.

1988-01-01

The Manned Mars Explorer (MME) project responds to the fundamental problems of sending human beings to Mars in a mission scenario and schematic vehicle designs. The mission scenario targets an opposition class Venus inbound swingby for its trajectory with concentration on Phobos and/or Deimos as a staging base for initial and future Mars vicinity operations. Optional vehicles are presented as a comparison using nuclear electric power/propulsion technology. A Manned Planetary Vehicle and Crew Command Vehicle are used to accomplish the targeted mission. The Manned Planetary Vehicle utilizes the mature technology of chemical propulsion combined with an advanced aerobrake, tether and pressurized environment system. The Crew Command Vehicle is the workhorse of the mission performing many different functions including a manned Mars landing, and Phobos rendezvous.

Trade Study of Multiple Thruster Options for the Mars Airplane Concept

NASA Technical Reports Server (NTRS)

Kuhl, Christopher A.; Gayle, Steven W.; Hunter, Craig A.; Kenney, Patrick S.; Scola, Salvatore; Paddock, David A.; Wright, Henry S.; Gasbarre, Joseph F.

2009-01-01

A trade study was performed at NASA Langley Research Center under the Planetary Airplane Risk Reduction (PARR) project (2004-2005) to examine the option of using multiple, smaller thrusters in place of a single large thruster on the Mars airplane concept with the goal to reduce overall cost, schedule, and technical risk. The 5-lbf (22N) thruster is a common reaction control thruster on many satellites. Thousands of these types of thrusters have been built and flown on numerous programs, including MILSTAR and Intelsat VI. This study has examined the use of three 22N thrusters for the Mars airplane propulsion system and compared the results to those of the baseline single thruster system.

Well-to-Wheels Analysis of Advanced Fuel/Vehicle Systems: A North American Study of Energy Use, Greenhouse Gas Emissions, and Criteria Pollutant Emissions

DOE Office of Scientific and Technical Information (OSTI.GOV)

Brinkman, Norman; Wang, Michael; Weber, Trudy

An accurate assessment of future fuel/propulsion system options requires a complete vehicle fuel-cycle analysis, commonly called a well-to-wheels (WTW) analysis. This WTW study analyzes energy use and emissions associated with fuel production (or well-to-tank [WTT]) activities and energy use and emissions associated with vehicle operation (or tank-to-wheels [TTW]) activities.

Assessing the Relative Risk of Aerocapture Using Probabalistic Risk Assessment

NASA Technical Reports Server (NTRS)

Percy, Thomas K.; Bright, Ellanee; Torres, Abel O.

2005-01-01

A recent study performed for the Aerocapture Technology Area in the In-Space Propulsion Technology Projects Office at the Marshall Space Flight Center investigated the relative risk of various capture techniques for Mars missions. Aerocapture has been proposed as a possible capture technique for future Mars missions but has been perceived by many in the community as a higher risk option as compared to aerobraking and propulsive capture. By performing a probabilistic risk assessment on aerocapture, aerobraking and propulsive capture, a comparison was made to uncover the projected relative risks of these three maneuvers. For mission planners, this knowledge will allow them to decide if the mass savings provided by aerocapture warrant any incremental risk exposure. The study focuses on a Mars Sample Return mission currently under investigation at the Jet Propulsion Laboratory (JPL). In each case (propulsive, aerobraking and aerocapture), the Earth return vehicle is inserted into Martian orbit by one of the three techniques being investigated. A baseline spacecraft was established through initial sizing exercises performed by JPL's Team X. While Team X design results provided the baseline and common thread between the spacecraft, in each case the Team X results were supplemented by historical data as needed. Propulsion, thermal protection, guidance, navigation and control, software, solar arrays, navigation and targeting and atmospheric prediction were investigated. A qualitative assessment of human reliability was also included. Results show that different risk drivers contribute significantly to each capture technique. For aerocapture, the significant drivers include propulsion system failures and atmospheric prediction errors. Software and guidance hardware contribute the most to aerobraking risk. Propulsive capture risk is mainly driven by anomalous solar array degradation and propulsion system failures. While each subsystem contributes differently to the risk of each technique, results show that there exists little relative difference in the reliability of these capture techniques although uncertainty for the aerocapture estimates remains high given the lack of in-space demonstration.

Propulsion System for Very High Altitude Subsonic Unmanned Aircraft

NASA Technical Reports Server (NTRS)

Bents, David J.; Mockler, Ted; Maldonado, Jaime; Harp, James L., Jr.; King, Joseph F.; Schmitz, Paul C.

1998-01-01

This paper explains why a spark ignited gasoline engine, intake pressurized with three cascaded stages of turbocharging, was selected to power NASA's contemplated next generation of high altitude atmospheric science aircraft. Beginning with the most urgent science needs (the atmospheric sampling mission) and tracing through the mission requirements which dictate the unique flight regime in which this aircraft has to operate (subsonic flight at greater then 80 kft) we briefly explore the physical problems and constraints, the available technology options and the cost drivers associated with developing a viable propulsion system for this highly specialized aircraft. The paper presents the two available options (the turbojet and the turbocharged spark ignited engine) which are discussed and compared in the context of the flight regime. We then show how the unique nature of the sampling mission, coupled with the economic considerations pursuant to aero engine development, point to the spark ignited engine as the only cost effective solution available. Surprisingly, this solution compares favorably with the turbojet in the flight regime of interest. Finally, some remarks are made about NASA's present state of development, and future plans to flight demonstrate the three stage turbocharged powerplant.

Solar Electric Propulsion for Mars Exploration

NASA Technical Reports Server (NTRS)

Hack, Kurt J.

1998-01-01

Highly propellant-efficient electric propulsion is being combined with advanced solar power technology to provide a non-nuclear transportation option for the human exploration of Mars. By virtue of its high specific impulse, electric propulsion offers a greater change in spacecraft velocity for each pound of propellant than do conventional chemical rockets. As a result, a mission to Mars based on solar electric propulsion (SEP) would require fewer heavy-lift launches than a traditional all-chemical space propulsion scenario would. Performance, as measured by mass to orbit and trip time, would be comparable to the NASA design reference mission for human Mars exploration, which utilizes nuclear thermal propulsion; but it would avoid the issues surrounding the use of nuclear reactors in space.

Low Carbon Propulsion Strategic Thrust Overview

NASA Technical Reports Server (NTRS)

Dryer, Jay

2014-01-01

NASA is taking a leadership role with regard to developing new options for low-carbon propulsion. Work related to the characterization of alternative fuels is coordinated with our partners in government and industry, and NASA is close to concluding a TC in this area. Research on alternate propulsion concepts continues to grow and is an important aspect of the ARMD portfolio. Strong partnerships have been a key enabling factor for research on this strategic thrust.

NIRPS - Solutions Facilitator Team Overview and Accomplishments

NASA Technical Reports Server (NTRS)

Brown, Thomas M., III; Childress, Rhonda

2013-01-01

National Institute for Rocket Propulsion Systems (NIRPS) purpose is to help preserve and align government and private rocket propulsion capabilities to meet present and future US commercial, civil, and defense needs, while providing authoritative insight and recommendations to National decisional authorities. Stewardship: Monitor and analyze the state of the industry in order to formulate and recommend National Policy options and strategies that promote a healthy industrial base and ensure best-value for the American taxpayer. Technology: Identify technology needs and recommend technology insertions by leading roadmap assessments and actively participating in program formulation activities. Solutions Facilitator/Provider: Maintain relationships and awareness across the Government, industry and academia, to align available capacity with emerging demand.

Recovery of Lunar Surface Access Module Residual and Reserve Propellants

NASA Technical Reports Server (NTRS)

Notardonato, William U.

2007-01-01

The Vision for Space Exploration calls for human exploration of the lunar surface in the 2020 timeframe. Sustained human exploration of the lunar surface will require supply, storage, and distribution of consumables for a variety of mission elements. These elements include propulsion systems for ascent and descent stages, life support for habitats and extra-vehicular activity, and reactants for power systems. NASA KSC has been tasked to develop technologies and strategies for consumables transfer for lunar exploration as part of the Exploration Technology Development Program. This paper will investigate details of operational concepts to scavenge residual propellants from the lunar descent propulsion system. Predictions on the mass of residuals and reserves are made. Estimates of heat transfer and boiloff rates are calculated and transient tank thermodynamic issues post-engine cutoff are modeled. Recovery and storage options including cryogenic liquid, vapor and water are discussed, and possible reuse of LSAM assets is presented.

Enrichment Zoning Options for the Small Nuclear Rocket Engine (SNRE)

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bruce G. Schnitzler; Stanley K. Borowski

2010-07-01

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. In NASA’s recent Mars Design Reference Architecture (DRA) 5.0 study (NASA-SP-2009-566, July 2009), nuclear thermal propulsion (NTP) was again selected over chemical propulsion as the preferred in-space transportation system option because of its high thrust and high specific impulse (-900 s) capability, increased tolerance to payload mass growth and architecture changes, and lower total initial mass in low Earth orbit. An extensive nuclear thermal rocket technology development effortmore » was conducted from 1955-1973 under the Rover/NERVA Program. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art design incorporating lessons learned from the very successful technology development program. Past activities at the NASA Glenn Research Center have included development of highly detailed MCNP Monte Carlo transport models of the SNRE and other small engine designs. Preliminary core configurations typically employ fuel elements with fixed fuel composition and fissile material enrichment. Uniform fuel loadings result in undesirable radial power and temperature profiles in the engines. Engine performance can be improved by some combination of propellant flow control at the fuel element level and by varying the fuel composition. Enrichment zoning at the fuel element level with lower enrichments in the higher power elements at the core center and on the core periphery is particularly effective. Power flattening by enrichment zoning typically results in more uniform propellant exit temperatures and improved engine performance. For the SNRE, element enrichment zoning provided very flat radial power profiles with 551 of the 564 fuel elements within 1% of the average element power. Results for this and alternate enrichment zoning options for the SNRE are compared.« less

Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to an Asteroid

NASA Technical Reports Server (NTRS)

Mercer, Carolyn R.; Oleson, Steven R.; Pencil, Eric J.; Piszczor, Michael F.; Mason, Lee S.; Bury, Kristen M.; Manzella, David H.; Kerslake, Thomas W.; Hojinicki, Jeffrey S.; Brophy, John P.

2012-01-01

NASA s goal for human spaceflight is to expand permanent human presence beyond low Earth orbit (LEO). NASA is identifying potential missions and technologies needed to achieve this goal. Mission options include crewed destinations to LEO and the International Space Station; high Earth orbit and geosynchronous orbit; cis-lunar space, lunar orbit, and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference missions to drive out required functions and capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One conclusion from this exercise was that a solar electric propulsion stage could reduce mission cost by reducing the required number of heavy lift launches and could increase mission reliability by providing a robust architecture for the long-duration crewed mission. Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its surface, and visiting its moons. This paper describes the parameterized assessment of power and propulsion technologies for a piloted solar electric vehicle to a near-Earth asteroid. The objective of the assessment was to determine technology drivers to advance the state of the art of electric propulsion systems for human exploration. Sensitivity analyses on the performance characteristics of the propulsion and power systems were done to determine potential system-level impacts of improved technology. Starting with a "reasonable vehicle configuration" bounded by an assumed launch date, we introduced technology improvements to determine the system-level benefits (if any) that those technologies might provide. The results of this assessment are discussed and recommendations for future work are described.

Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to an Asteroid

NASA Technical Reports Server (NTRS)

Mercer, Carolyn R.; Oleson, Steven R.; Pencil, Eric J.; Piszczor, Michael F.; Mason, Lee S.; Bury, Kristen M.; Manzella, David H.; Kerslake, Thomas W.; Hojinicki, Jeffrey S.; Brophy, John P.

2011-01-01

NASA's goal for human spaceflight is to expand permanent human presence beyond low Earth orbit (LEO). NASA is identifying potential missions and technologies needed to achieve this goal. Mission options include crewed destinations to LEO and the International Space Station; high Earth orbit and geosynchronous orbit; cis-lunar space, lunar orbit, and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference missions to drive out required functions and capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One conclusion from this exercise was that a solar electric propulsion stage could reduce mission cost by reducing the required number of heavy lift launches and could increase mission reliability by providing a robust architecture for the long-duration crewed mission. Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its surface, and visiting its moons. This paper describes the parameterized assessment of power and propulsion technologies for a piloted solar electric vehicle to a near-Earth asteroid. The objective of the assessment was to determine technology drivers to advance the state of the art of electric propulsion systems for human exploration. Sensitivity analyses on the performance characteristics of the propulsion and power systems were done to determine potential system-level impacts of improved technology. Starting with a "reasonable vehicle configuration" bounded by an assumed launch date, we introduced technology improvements to determine the system-level benefits (if any) that those technologies might provide. The results of this assessment are discussed and recommendations for future work are described.

Mars Mission Analysis Trades Based on Legacy and Future Nuclear Propulsion Options

NASA Astrophysics Data System (ADS)

Joyner, Russell; Lentati, Andrea; Cichon, Jaclyn

2007-01-01

The purpose of this paper is to discuss the results of mission-based system trades when using a nuclear thermal propulsion (NTP) system for Solar System exploration. The results are based on comparing reactor designs that use a ceramic-metallic (CERMET), graphite matrix, graphite composite matrix, or carbide matrix fuel element designs. The composite graphite matrix and CERMET designs have been examined for providing power as well as propulsion. Approaches to the design of the NTP to be discussed will include an examination of graphite, composite, carbide, and CERMET core designs and the attributes of each in regards to performance and power generation capability. The focus is on NTP approaches based on tested fuel materials within a prismatic fuel form per the Argonne National Laboratory testing and the ROVER/NERVA program. NTP concepts have been examined for several years at Pratt & Whitney Rocketdyne for use as the primary propulsion for human missions beyond earth. Recently, an approach was taken to examine the design trades between specific NTP concepts; NERVA-based (UC)C-Graphite, (UC,ZrC)C-Composite, (U,Zr)C-Solid Carbide and UO2-W CERMET. Using Pratt & Whitney Rocketdyne's multidisciplinary design analysis capability, a detailed mission and vehicle model has been used to examine how several of these NTP designs impact a human Mars mission. Trends for the propulsion system mass as a function of power level (i.e. thrust size) for the graphite-carbide and CERMET designs were established and correlated against data created over the past forty years. These were used for the mission trade study. The resulting mission trades presented in this paper used a comprehensive modeling approach that captures the mission, vehicle subsystems, and NTP sizing.

Space Shuttle propulsion parameter estimation using optimal estimation techniques, volume 1

NASA Technical Reports Server (NTRS)

1983-01-01

The mathematical developments and their computer program implementation for the Space Shuttle propulsion parameter estimation project are summarized. The estimation approach chosen is the extended Kalman filtering with a modified Bryson-Frazier smoother. Its use here is motivated by the objective of obtaining better estimates than those available from filtering and to eliminate the lag associated with filtering. The estimation technique uses as the dynamical process the six degree equations-of-motion resulting in twelve state vector elements. In addition to these are mass and solid propellant burn depth as the ""system'' state elements. The ""parameter'' state elements can include aerodynamic coefficient, inertia, center-of-gravity, atmospheric wind, etc. deviations from referenced values. Propulsion parameter state elements have been included not as options just discussed but as the main parameter states to be estimated. The mathematical developments were completed for all these parameters. Since the systems dynamics and measurement processes are non-linear functions of the states, the mathematical developments are taken up almost entirely by the linearization of these equations as required by the estimation algorithms.

Spacecraft attitude control for a solar electric geosynchronous transfer mission

NASA Technical Reports Server (NTRS)

Leroy, B. E.; Regetz, J. D., Jr.

1975-01-01

A study of the Attitude Control System (ACS) is made for a solar electric propulsion geosynchronous transfer mission. The basic mission considered is spacecraft injection into a low altitude, inclined orbit followed by low thrust orbit changing to achieve geosynchronous orbit. Because of the extended thrusting time, the mission performance is a strong function of the attitude control system. Two attitude control system design options for an example mission evolve from consideration of the spacecraft configuration, the environmental disturbances, and the probable ACS modes of operation. The impact of these design options on other spacecraft subsystems is discussed. The factors which must be considered in determining the ACS actuation and sensing subsystems are discussed. The effects of the actuation and sensing subsystems on the mission performance are also considered.

Mars Sample Return Orbiter Architecture Options — Results of the ESA Mars Sample Return Architecture Assessment Study

NASA Astrophysics Data System (ADS)

Derz, U.; Joffre, E.; Perkinson, M.-C.; Huesing, J.; Beyer, F.; Sanchez Perez, J. M.

2018-04-01

This paper presents the identified most promising chemical and electric propulsion architecture options of the Mars Sample Return (MSR) orbiter identified during the recent ESA MSR Architecture Assessment Study.

Rapid Response R&D for the Propulsion Directorate. Delivery Order 0019: Advanced Alternative Energy Technologies, Subtask: Life Cycle Greenhouse Gas Analysis of Advanced Jet Propulsion Fuels: Fischer-Tropsch Based SPK-1 Case Study

DTIC Science & Technology

2011-09-01

carry finished jet fuel from the CBTL facility. The pipeline connects the CBTL facility to a petroleum refinery located in Wood River, Illinois...Under Option 1, all the blended jet fuel is transported via pipeline from the refinery in Wood River to Chicago’s O’Hare airport. Under Option 2...shipping F-T jet fuel to a refinery in Wood River, Illinois (near St. Louis, Missouri) for blending and final transport of the blended jet fuel to

US Decadal Survey Outer Solar System Missions: Trajectory Options

NASA Astrophysics Data System (ADS)

Spilker, T. R.; Atkinson, D. H.; Strange, N. J.; Landau, D.

2012-04-01

The report of the US Planetary Science Decadal Survey (PSDS), released in draft form March 7, 2011, identifies several mission concepts involving travel to high-priority outer solar system (OSS) destinations. These include missions to Europa and Jupiter, Saturn and two of its satellites, and Uranus. Because travel to the OSS involves much larger distances and larger excursions out of the sun's gravitational potential well than inner solar system (ISS) missions, transfer trajectories for OSS missions are stronger drivers of mission schedule and resource requirements than for ISS missions. Various characteristics of each planet system, such as obliquity, radiation belts, rings, deep gravity wells, etc., carry ramifications for approach trajectories or trajectories within the systems. The maturity of trajectory studies for each of these destinations varies significantly. Europa has been the focus of studies for well over a decade. Transfer trajectory options from Earth to Jupiter are well understood. Current studies focus on trajectories within the Jovian system that could reduce the total mission cost of a Europa orbiter mission. Three missions to the Saturn system received high priority ratings in the PSDS report: two flagship orbital missions, one to Titan and one to Enceladus, and a Saturn atmospheric entry probe mission for NASA's New Frontiers Program. The Titan Saturn System Mission (TSSM) studies of 2007-2009 advanced our understanding of trajectory options for transfers to Saturn, including solar electric propulsion (SEP) trajectories. But SEP trajectories depend more on details of spacecraft and propulsion system characteristics than chemical trajectories, and the maturity of SEP trajectory search tools has not yet caught up with chemical trajectory tools, so there is still more useful research to be done on Saturn transfers. The TSSM studies revealed much about Saturn-orbiting trajectories that yield efficient and timely delivery to Titan or Enceladus orbits. PSDS studies suggest there is significant flexibility in trajectory options for a Saturn entry probe. Many different trajectory approaches and a wide range of trajectory parameters yield probe-to-carrier data relay communications of sufficient duration and at data rates surpassing those of the 1995 Galileo probe at Jupiter. Research into trajectories to Uranus and within the Uranian system is the least mature. SEP and chemical transfer trajectories have been examined but the searches are not yet comprehensive, though the fundamentals of such transfers are fairly well understood. Study of orbital tours within the Uranian system suggests they can be analogous to such tours in the Jovian system, though in the time frame of launches within the next couple of decades arrivals at Uranus would be from high declinations, so the transition from an initial orbit to an equatorial satellite tour could be lengthy. This paper will present results from the most recent work on these trajectories. This research was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under contract with NASA. Copyright 2012 California Institute of Technology. Government sponsorship acknowledged.

Self-rescue strategies for EVA crewmembers equipped with the SAFER backpack

NASA Technical Reports Server (NTRS)

Williams, Trevor; Baughman, David

1994-01-01

An extravehicular astronaut who becomes separated from a space station has three options available: grappling the station immediately by means of a 'shepherd's crook' device; rescue by either a second crewmember flying an MMU or a robotic-controlled MMU; or self-rescue by means of a propulsive system. The first option requires very fast response by a tumbling astronaut; the second requires constant availability of an MMU, as well as a rendezvous procedure thousands of feet from the station. This paper will consider the third option, propulsive self-rescue. In particular, the capability of the new Simplified Aid for EVA Rescue (SAFER) propulsive backpack, which is to be tested on STS-64 in Sep. 1994, will be studied. This system possesses an attitude hold function that can automatically detumble an astronaut after separation. On-orbit tests of candidate self-rescue systems have demonstrated the need for such a feature. SAFER has a total delta(v) capability of about 10 fps, to cover both rotations and translations, compared with a possible separation rate of 2.5 fps. But the delta(v) required for self-rescue is critically dependent on the delay before return can be initiated, as a consequence of orbital effects. A very important practical question is then whether the total delta(v) of SAFER is adequate to perform self-rescue for worst case values of separation speed, time to detumble, and time for the astronaut to visually acquire the station. This paper shows that SAFER does indeed have sufficient propellant to carry out self-rescue in all realistic separation cases, as well as in cases which are considerably more severe than anything likely to be encountered in practice. The return trajectories and total delta(v)'s discussed are obtained by means of an 'inertial line-of-sight targeting' scheme, derived in the paper, which allows orbital effects to be corrected by making use of the visual information available to the pilot, namely the line-of-sight direction to the station relative to the stars.

Radioisotope Electric Propulsion for Fast Outer Planetary Orbiters

NASA Technical Reports Server (NTRS)

Oleson, Steven; Benson, Scott; Gefert, Leon; Patterson, Michael; Schreiber, Jeffrey

2002-01-01

Recent interest in outer planetary targets by the Office of Space Science has spurred the search for technology options to enable relatively quick missions to outer planetary targets. Several options are being explored including solar electric propelled stages combined with aerocapture at the target and nuclear electric propulsion. Another option uses radioisotope powered electric thrusters to reach the outer planets. Past work looked at using this technology to provide faster flybys. A better use for this technology is for outer planet orbiters. Combined with medium class launch vehicles and a new direct trajectory these small, sub-kilowatt ion thrusters and Stirling radioisotope generators were found to allow missions as fast as 5 to 12 years for objects from Saturn to Pluto, respectively. Key to the development is light spacecraft and science payload technologies.

Nuclear Thermal Rocket (Ntr) Propulsion: A Proven Game-Changing Technology for Future Human Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2012-01-01

The NTR represents the next evolutionary step in high performance rocket propulsion. It generates high thrust and has a specific impulse (Isp) of approx.900 seconds (s) or more V twice that of today s best chemical rockets. The technology is also proven. During the previous Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) nuclear rocket programs, 20 rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability V all the requirements needed for a human mission to Mars. Ceramic metal cermet fuel was also pursued, as a backup option. The NTR also has significant growth and evolution potential. Configured as a bimodal system, it can generate electrical power for the spacecraft. Adding an oxygen afterburner nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, simple assembly and mission operations. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program V the 25,000 lbf (25 klbf) Pewee engine is sufficient for human Mars missions when used in a clustered engine arrangement. The Copernicus crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth asteroid (NEA) and Mars orbital missions prior to a Mars landing mission. Initially, the basic Copernicus vehicle can enable reusable 1-year round trip human missions to candidate NEAs like 1991 JW and Apophis in the late 2020 s to check out vehicle systems. Afterwards, the Copernicus spacecraft and its 2 key components, now configured as an Earth Return Vehicle / propellant tanker, would be used for a short round trip (approx.18 - 20 months)/short orbital stay (60 days) Mars / Phobos survey mission in 2033 using a split mission approach. The paper also discusses NASA s current Foundational Technology Development activities and its pre-decisional plans for future system-level Technology Demonstrations that include ground testing a small (approx.7.5 klbf) scalable NTR before the decade is out with a flight test shortly thereafter.

Implementation of a Low-Thrust Trajectory Optimization Algorithm for Preliminary Design

NASA Technical Reports Server (NTRS)

Sims, Jon A.; Finlayson, Paul A.; Rinderle, Edward A.; Vavrina, Matthew A.; Kowalkowski, Theresa D.

2006-01-01

A tool developed for the preliminary design of low-thrust trajectories is described. The trajectory is discretized into segments and a nonlinear programming method is used for optimization. The tool is easy to use, has robust convergence, and can handle many intermediate encounters. In addition, the tool has a wide variety of features, including several options for objective function and different low-thrust propulsion models (e.g., solar electric propulsion, nuclear electric propulsion, and solar sail). High-thrust, impulsive trajectories can also be optimized.

Lunar Cube Transfer Trajectory Options

NASA Technical Reports Server (NTRS)

Folta, David; Dichmann, Donald James; Clark, Pamela E.; Haapala, Amanda; Howell, Kathleen

2015-01-01

Numerous Earth-Moon trajectory and lunar orbit options are available for Cubesat missions. Given the limited Cubesat injection infrastructure, transfer trajectories are contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these transfers can be restricted by the selection or designs of Cubesat subsystems such as propulsion or communication. Nonetheless, many trajectory options can b e considered which have a wide range of transfer duration, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several options including deployment from low Earth orbit (LEO) geostationary transfer orbits (GTO) and higher energy direct lunar transfer and the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory.

Lunar Cube Transfer Trajectory Options

NASA Technical Reports Server (NTRS)

Folta, David; Dichmann, Donald J.; Clark, Pamela; Haapala, Amanda; Howell, Kathleen

2015-01-01

Numerous Earth-Moon trajectory and lunar orbit options are available for Cubesat missions. Given the limited Cubesat injection infrastructure, transfer trajectories are contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these transfers can be restricted by the selection or designs of Cubesat subsystems such as propulsion or communication. Nonetheless, many trajectory options can be considered which have a wide range of transfer durations, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several options including deployment from low Earth orbit (LEO), geostationary transfer orbits (GTO), and higher energy direct lunar transfers and the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory.

The effect of interplanetary trajectory options on a manned Mars aerobrake configuration

NASA Technical Reports Server (NTRS)

Braun, Robert D.; Powell, Richard W.; Hartung, Lin C.

1990-01-01

Manned Mars missions originating in low Earth orbit (LEO) in the time frame 2010 to 2025 were analyzed to identify preferred mission opportunities and their associated vehicle and trajectory characteristics. Interplanetary and Mars atmospheric trajectory options were examined under the constraints of an initial manned exploration scenario. Two chemically propelled vehicle options were considered: (1) an all propulsive configuration, and (2) a configuration which employs aerobraking at Earth and Mars with low lift/drag (L/D) shapes. Both the interplanetary trajectory options as well as the Mars atmospheric passage are addressed to provide a coupled trajectory simulation. Direct and Venus swingby interplanetary transfers with a 60 day Mars stopover are considered. The range and variation in both Earth and Mars entry velocity are also defined. Two promising mission strategies emerged from the study: (1) a 1.0 to 2.0 year Venus swingby mission, and (2) a 2.0 to 2.5 year direct mission. Through careful trajectory selection, 11 mission opportunities are identified in which the Mars entry velocity is between 6 and 10 km/sec and Earth entry velocity ranges from 11.5 to 12.5 km/sec. Simulation of the Earth return aerobraking maneuver is not performed. It is shown that a low L/D configuration is not feasible for Mars aerobraking without substantial improvements in the interplanetary navigation system. However, even with an advanced navigation system, entry corridor and aerothermal requirements restrict the number of potential mission opportunities. It is also shown that for a large blunt Mars aerobrake configuration, the effects of radiative heating can be significant at entry velocities as low as 6.2 km/sec and will grow to dominate the aerothermal environment at entry velocities above 8.5 km/sec. Despite the additional system complexity associated with an aerobraking vehicle, the use of aerobraking was shown to significantly lower the required initial LEO weight. In comparison with an all propulsive mission, savings between 19 and 59 percent were obtained depending upon launch date.

Updated Mars Mission Architectures Featuring Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Rodriguez, Mitchell A.; Percy, Thomas K.

2017-01-01

Nuclear thermal propulsion (NTP) can potentially enable routine human exploration of Mars and the solar system. By using nuclear fission instead of a chemical combustion process, and using hydrogen as the propellant, NTP systems promise rocket efficiencies roughly twice that of the best chemical rocket engines currently available. The most recent major Mars architecture study featuring NTP was the Design Reference Architecture 5.0 (DRA 5.0), performed in 2009. Currently, the predominant transportation options being considered are solar electric propulsion (SEP) and chemical propulsion; however, given NTP's capabilities, an updated architectural analysis is needed. This paper provides a top-level overview of several different architectures featuring updated NTP performance data. New architectures presented include a proposed update to the DRA 5.0 as well as an investigation of architectures based on the current Evolvable Mars Campaign, which is the focus of NASA's current analyses for the Journey to Mars. Architectures investigated leverage the latest information relating to NTP performance and design considerations and address new support elements not available at the time of DRA 5.0, most notably the Orion crew module and the Space Launch System (SLS). The paper provides a top level quantitative comparison of key performance metrics as well as a qualitative discussion of improvements and key challenges still to be addressed. Preliminary results indicate that the updated NTP architectures can significantly reduce the campaign mass and subsequently the costs for assembly and number of launches.

Preliminary Design of Low-Thrust Interplanetary Missions

NASA Technical Reports Server (NTRS)

Sims, Jon A.; Flanagan, Steve N.

1997-01-01

For interplanetary missions, highly efficient electric propulsion systems can be used to increase the mass delivered to the destination and/or reduce the trip time over typical chemical propulsion systems. This technology is being demonstrated on the Deep Space 1 mission - part of NASA's New Millennium Program validating technologies which can lower the cost and risk and enhance the performance of future missions. With the successful demonstration on Deep Space 1, future missions can consider electric propulsion as a viable propulsion option. Electric propulsion systems, while highly efficient, produce only a small amount of thrust. As a result, the engines operate during a significant fraction of the trajectory. This characteristic makes it much more difficult to find optimal trajectories. The methods for optimizing low-thrust trajectories are typically categorized as either indirect, or direct. Indirect methods are based on calculus of variations, resulting in a two-point boundary value problem that is solved by satisfying terminal constraints and targeting conditions. These methods are subject to extreme sensitivity to the initial guess of the variables - some of which are not physically intuitive. Adding a gravity assist to the trajectory compounds the sensitivity. Direct methods parameterize the problem and use nonlinear programming techniques to optimize an objective function by adjusting a set of variables. A variety of methods of this type have been examined with varying results. These methods are subject to the limitations of the nonlinear programming techniques. In this paper we present a direct method intended to be used primarily for preliminary design of low-thrust interplanetary trajectories, including those with multiple gravity assists. Preliminary design implies a willingness to accept limited accuracy to achieve an efficient algorithm that executes quickly.

Assessment of Options for Enhancing Surface Ship Acquisition.

DTIC Science & Technology

1996-03-01

propulsion control, cargo handling, and the ship’s self defense capability. Cost growth was 50 percent even with a reduction from 9 ships to 5. For...technology, high - value weapon and sensor systems Cargo Arrangements Every area of ship densely packed with equipment, cables, pipes, ventilation...acquisition reform has a high potential to improve the outcomes of ship acquisition programs. Progress has been made in the reduction of military

Replacement Capability Options for the United States Space Shuttle

DTIC Science & Technology

2013-09-01

extended periods, and to expand our knowledge of solar astronomy well beyond Earth-based observations.” During the Skylab missions, both the man...determined Skylab’s orbit was no longer stable due to higher than predicted solar activity. Therefore, Skylab had to be de-orbited earlier than...Module houses the oxygen, life support, power, communications, thermal control, and propulsions systems. The solar arrays for the Soyuz are also

A survey of propulsion options for cargo and piloted missions to Mars.

PubMed

Sankaran, K; Cassady, L; Kodys, A D; Choueiri, E Y

2004-05-01

In this paper, high-power electric propulsion options are surveyed in the context of cargo and piloted missions to Mars. A low-thrust trajectory optimization program (raptor) is utilized to analyze this mission. Candidate thrusters are chosen based upon demonstrated performance in the laboratory. Hall, self-field magnetoplasmadynamic (MPDT), self-field lithium Lorentz force accelerator (LiLFA), arcjet, and applied-field LiLFA systems are considered for this mission. In this first phase of the study, all thrusters are assumed to operate at a single power level (regardless of the efficiency-power curve), and the thruster specific mass and power plant specific mass are taken to be the same for all systems. Under these assumptions, for a 7.5 MW, 60 mT payload, piloted mission, the self-field LiLFA results in the shortest trip time (340 days) with a reasonable propellant mass fraction of 57% (129 mT). For a 150 kW, 9 mT payload, cargo mission, both the applied-field LiLFA and the Hall thruster seem reasonable choices with propellant mass fractions of 42 to 45%(7 to 8 mT). The Hall thrusters provide better trip times (530-570 days) compared to the applied-field LiLFA (710 days) for the relatively less demanding mission.

Near-Earth Asteroid Retrieval Mission (ARM) Study

NASA Technical Reports Server (NTRS)

Brophy, John R.; Muirhead, Brian

2013-01-01

The Asteroid Redirect Mission (ARM) concept brings together the capabilities of the science, technology, and the human exploration communities on a grand challenge combining robotic and human space exploration beyond low Earth orbit. This paper addresses the key aspects of this concept and the options studied to assess its technical feasibility. Included are evaluations of the expected number of potential targets, their expected discovery rate, the necessity to adequately characterize candidate mission targets, the process to capture a non-cooperative asteroid in deep space, and the power and propulsion technology required for transportation back to the Earth-Moon system. Viable options for spacecraft and mission designs are developed. Orbits for storing the retrieved asteroid that are stable for more than a hundred years, yet allow for human exploration and commercial utilization of a redirected asteroid, are identified. The study concludes that the key aspects of finding, capturing and redirecting an entire small, near-Earth asteroid to the Earth-Moon system by the first half of the next decade are technically feasible. The study was conducted from January 2013 through March 2013 by the Jet Propulsion Laboratory (JPL) in collaboration with Glenn Research Center (GRC), Johnson Space Center (JSC), Langley Research Center (LaRC), and Marshall Space Flight Center (MSFC).

Construction of Power Receiving Rectenna Using Mars- In-Situ Materials; A Low Energy Materials Processing Approach

NASA Technical Reports Server (NTRS)

Curreri, Peter A.; Rose, M. Franklin (Technical Monitor)

2001-01-01

It is highly desirable to have a non-nuclear power rich option for the human exploration of Mars. Utilizing a Solar Electric Propulsion, SEP, / Power Beaming architecture for a non-nuclear power option for a human Mars base potentially avoids the weather and dust sensitivities of the surface photovoltaic option. Further from Mars areosynchronous orbit near year round power can be provided. Mission analysis, however, concludes that ultra high (245 GHz) frequencies or laser transmission technologies are required for Mars landed mass competitiveness with the surface photovoltaic option if the receiving rectifying antenna "rectenna" is transported from Earth. It is suggested in this paper that producing rectenna in situ on Mars surface might make a more conventional 5.8 GHz system competitive with surface PV. The premium of a competitive, robust, continuous base power might make the development of a 10 plus MWe class SEP for human Mars mission a more attractive non-nuclear option.

Advanced Solar-propelled Cargo Spacecraft for Mars Missions

NASA Technical Reports Server (NTRS)

Auziasdeturenne, Jacqueline; Beall, Mark; Burianek, Joseph; Cinniger, Anna; Dunmire, Barbrina; Haberman, Eric; Iwamoto, James; Johnson, Stephen; Mccracken, Shawn; Miller, Melanie

1989-01-01

Three concepts for an unmanned, solar powered, cargo spacecraft for Mars support missions were investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: A Solar Radiation Absorption (SRA) system, a Solar-Pumped Laser (SPL) system and a solar powered magnetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sunsynchronous Earth orbit converts solar energy to laser energy. The MPD system used indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft and the SPL powered spacecraft return to Earth for subsequent missions. The MPD propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

From Earth to orbit. [assessment of transportation options

NASA Technical Reports Server (NTRS)

Gavin, Joseph G., Jr.; Blond, Edmund; Brill, Yvonne C.; Budiansky, Bernard; Cooper, Robert S.; Demisch, Wolfgang H.; Hawk, Clark W.; Kerrebrock, Jack L.; Lichtenberg, Byron K.; Mager, Arthur

1992-01-01

Within this document, the National Research Council (NRC) assesses the requirements, benefits, technological feasibility, and roles of Earth-to-orbit transportation options that could be developed in support of the national space program. Among the topics covered are launch vehicles and infrastructure, propulsion, and technology.

Lightweight Damage Tolerant, High-Temperature Radiators for Nuclear Power and Propulsion

NASA Technical Reports Server (NTRS)

Craven, Paul D.; SanSoucie, Michael P.

2015-01-01

NASA is increasingly emphasizing exploration to bodies beyond near-Earth orbit. New propulsion systems and new spacecraft are being built for these missions. As the target bodies get further out from Earth, high energy density systems, e.g., nuclear fusion, for propulsion and power will be advantageous. The mass and size of these systems, including supporting systems such as the heat exchange system, including thermal radiators, will need to be as small as possible. Conventional heat exchange systems are a significant portion of the total thermal management mass and size. Nuclear electric propulsion (NEP) is a promising option for high-speed, in-space travel due to the high energy density of nuclear fission power sources and efficient electric thrusters. Heat from the reactor is converted to power for use in propulsion or for system power. The heat not used in the power conversion is then radiated to space as shown in figure 1. Advanced power conversion technologies will require high operating temperatures and would benefit from lightweight radiator materials. Radiator performance dictates power output for nuclear electric propulsion systems. Pitch-based carbon fiber materials have the potential to offer significant improvements in operating temperature, thermal conductivity, and mass. These properties combine to allow significant decreases in the total mass of the radiators and significant increases in the operating temperature of the fins. A Center-funded project at NASA Marshall Space Flight Center has shown that high thermal conductivity, woven carbon fiber fins with no matrix material, can be used to dissipate waste heat from NEP systems and because of high specific power (kW/kg), will require less mass and possibly less total area than standard metal and composite radiator fins for radiating the same amount of heat. This project uses an innovative approach to reduce the mass and size required for the thermal radiators to the point that in-space NEP and power is enabled. High thermal conductivity carbon fibers are lightweight, damage tolerant, and can be heated to high temperature. Areal densities in the NASA set target range of 2 to 4 kg/m2 (for enabling NEP) are achieved and with specific powers (kW/kg) a factor of about 7 greater than conventional metal fins and about 1.5 greater than carbon composite fins. Figure 2 shows one fin under test. All tests were done under vacuum conditions.

Propulsion Technology Assessment: Science and Enabling Technologies to Explore the Interstellar Medium

NASA Technical Reports Server (NTRS)

Hopkins, Randall C.; Thomas, Herbert D.; Wiegmann, Bruce M.; Heaton, Andrew F.; Johnson, Les; Beers, Benjamin R.

2015-01-01

The Advanced Concepts Office at NASA’s George C. Marshall Space Flight Center conducted a study to assess what low-thrust advanced propulsion system candidates, existing and near term, could deliver a small, Voyager-like satellite to our solar system’s heliopause, approximately 100 AU from the sun, within 10 years. The advanced propulsion system trade study consisted of three candidates, including a Magnetically Shielded Miniature Hall thruster, a solar sail and an electric sail. A second analysis was conducted to determine which solid rocket motor kick stage(s) would be required to provide additional thrust at various points in the trajectory, assuming a characteristic energy capability provided by a Space Launch System Block 1B vehicle architecture carrying an 8.4 meter payload fairing. Two trajectory profiles were considered, including an escape trajectory using a Jupiter gravity assist and an escape trajectory first performing a Jupiter gravity assist followed by an Oberth maneuver around the sun and an optional Saturn gravity assist. Results indicated that if the Technology Readiness Level of an electric sail could be increased in time, this technology could not only enable a satellite to reach 100 AU in 10 years but it could potentially do so in even less time.

Advanced Filter Technology For Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Castillon, Erick

2015-01-01

The Scrubber System focuses on using HEPA filters and carbon filtration to purify the exhaust of a Nuclear Thermal Propulsion engine of its aerosols and radioactive particles; however, new technology may lend itself to alternate filtration options, which may lead to reduction in cost while at the same time have the same filtering, if not greater, filtering capabilities, as its predecessors. Extensive research on various types of filtration methods was conducted with only four showing real promise: ionization, cyclonic separation, classic filtration, and host molecules. With the four methods defined, more research was needed to find the devices suitable for each method. Each filtration option was matched with a device: cyclonic separators for the method of the same name, electrostatic separators for ionization, HEGA filters, and carcerands for the host molecule method. Through many hours of research, the best alternative for aerosol filtration was determined to be the electrostatic precipitator because of its high durability against flow rate and its ability to cleanse up to 99.99% of contaminants as small as 0.001 micron. Carcerands, which are the only alternative to filtering radioactive particles, were found to be non-existent commercially because of their status as a "work in progress" at research institutions. Nevertheless, the conclusions after the research were that HEPA filters is recommended as the best option for filtering aerosols and carbon filtration is best for filtering radioactive particles.

100-Lb(f) LO2/LCH4 Reaction Control Engine Technology Development for Future Space Vehicles

NASA Technical Reports Server (NTRS)

Robinson, Philip J.; Veith, Eric M.; Hurlbert, Eric A.; Jimenez, Rafael; Smith, Timothy D.

2008-01-01

The National Aeronautics and Space Administration (NASA) has identified liquid oxygen (LO2)/liquid methane (LCH4) propulsion systems as promising options for some future space vehicles. NASA issued a contract to Aerojet to develop a 100-lbf (445 N) LO2/LCH4 Reaction Control Engine (RCE) aimed at reducing the risk of utilizing a cryogenic reaction control system (RCS) on a space vehicle. Aerojet utilized innovative design solutions to develop an RCE that can ignite reliably over a broad range of inlet temperatures, perform short minimum impulse bits (MIB) at small electrical pulse widths (EPW), and produce excellent specific impulse (Isp) across a range of engine mixture ratios (MR). These design innovations also provide a start transient with a benign MR, ensuring good thrust chamber compatibility and long life. In addition, this RCE can successfully operate at MRs associated with main engines, enabling the RCE to provide emergency backup propulsion to minimize vehicle propellant load and overall system mass.

100-LBF LO2/LCH4 - Reaction Control Engine Technology Development for Future Space Vehicles

NASA Technical Reports Server (NTRS)

Robinson, Philip J.; Veith, Eric M.; Hurlbert, Eric A.; Jimenez, Rafael; Smith, Timothy D.

2008-01-01

The National Aeronautics and Space Administration (NASA) has identified liquid oxygen (LO2)/liquid methane (LCH4) propulsion systems as promising options for some future space vehicles. NASA issued a contract to Aerojet to develop a 100-lbf (445 N) LO2/LCH4 Reaction Control Engine (RCE) aimed at reducing the risk of utilizing a cryogenic reaction control system (RCS) on a space vehicle. Aerojet utilized innovative design solutions to develop an RCE that can ignite reliably over a broad range of inlet temperatures, perform short minimum impulse bits (MIB) at small electrical pulse widths (EPW), and produce excellent specific impulse (Isp) across a range of engine mixture ratios (MR). These design innovations also provide a start transient with a benign MR, ensuring good thrust chamber compatibility and long life. In addition, this RCE can successfully operate at MRs associated with main engines, enabling the RCE to provide emergency backup propulsion to minimize vehicle propellant load and overall system mass.

Mars Ascent Vehicle Gross Lift-off Mass Sensitivities for Robotic Mars Sample Return

NASA Technical Reports Server (NTRS)

Dux, Ian J.; Huwaldt, Joseph A.; McKamey, R. Steve; Dankanich, John W.

2011-01-01

The Mars ascent vehicle is a critical element of the robotic Mars Sample Return (MSR) mission. The Mars ascent vehicle must be developed to survive a variety of conditions including the trans-Mars journey, descent through the Martian atmosphere and the harsh Martian surface environments while maintaining the ability to deliver its payload to a low Mars orbit. The primary technology challenge of developing the Mars ascent vehicle system is designing for all conditions while ensuring the mass limitations of the entry descent and landing system are not exceeded. The NASA In-Space Propulsion technology project has initiated the development of Mars ascent vehicle technologies with propulsion system performance and launch environments yet to be defined. To support the project s evaluation and development of various technology options the sensitivity of the Mars ascent vehicle gross lift-off mass to engine performance, inert mass, target orbits, and launch conditions has been completed with the results presented herein.

Future earth orbit transportation systems/technology implications

NASA Technical Reports Server (NTRS)

Henry, B. Z.; Decker, J. P.

1976-01-01

Assuming Space Shuttle technology to be state-of-the-art, projected technological advances to improve the capabilities of single-stage-to-orbit (SSTO) derivatives are examined. An increase of about 30% in payload performance can be expected from upgrading the present Shuttle system through weight and drag reductions and improvements in the propellants and engines. The ODINEX (Optimal Design Integration Executive Computer Program) program has been used to explore design options. An advanced technology SSTO baseline system derived from ODINEX analysis has a conventional wing-body configuration using LOX/LH engines, three with two-position nozzles with expansion ratios of 40 and 200 and four with fixed nozzles with an expansion ratio of 40. Two assisted-takeoff approaches are under consideration in addition to a concept in which the orbital vehicle takes off empty using airbreathing propulsion and carries out a rendezvous with two large cryogenic tankers carrying propellant at an altitude of 6100 m. Further approaches under examination for propulsion, aerothermodynamic design, and design integration are described.

TOPEX satellite concept. TOPEX option study report

NASA Technical Reports Server (NTRS)

Meyer, D. P.; Case, C. M.

1982-01-01

Candidate bus equipment from the Viking, Applications Explorer Mission, and Small Scientific Satellite programs for application to the TOPEX mission options is assessed. Propulsion module equipment and subsystem candidates from the Applications Explorer Mission satellites and the Small Scientific Satellite spacecraft are evaluated for those TOPEX options. Several subsystem concepts appropriate to the TOPEX options are described. These descriptions consider performance characteristics of the subsystems. Cost and availability information on the candidate equipment and subsystems are also provided.

MisTec - A software application for supporting space exploration scenario options and technology development analysis and planning

NASA Technical Reports Server (NTRS)

Horsham, Gary A. P.

1992-01-01

This structure and composition of a new, emerging software application, which models and analyzes space exploration scenario options for feasibility based on technology development projections is presented. The software application consists of four main components: a scenario generator for designing and inputting scenario options and constraints; a processor which performs algorithmic coupling and options analyses of mission activity requirements and technology capabilities; a results display which graphically and textually shows coupling and options analysis results; and a data/knowledge base which contains information on a variety of mission activities and (power and propulsion) technology system capabilities. The general long-range study process used by NASA to support recent studies is briefly introduced to provide the primary basis for comparison for discussing the potential advantages to be gained from developing and applying this kind of application. A hypothetical example of a scenario option to facilitate the best conceptual understanding of what the application is, how it works, or the operating methodology, and when it might be applied is presented.

MisTec: A software application for supporting space exploration scenario options and technology development analysis and planning

NASA Technical Reports Server (NTRS)

Horsham, Gary A. P.

1991-01-01

The structure and composition of a new, emerging software application, which models and analyzes space exploration scenario options for feasibility based on technology development projections is presented. The software application consists of four main components: a scenario generator for designing and inputting scenario options and constraints; a processor which performs algorithmic coupling and options analyses of mission activity requirements and technology capabilities; a results display which graphically and textually shows coupling and options analysis results; and a data/knowledge base which contains information on a variety of mission activities and (power and propulsion) technology system capabilities. The general long-range study process used by NASA to support recent studies is briefly introduced to provide the primary basis for comparison for discussing the potential advantages to be gained from developing and applying this king of application. A hypothetical example of a scenario option to facilitate the best conceptual understanding of what the application is, how it works, or the operating methodology, and when it might be applied is presented.

CVX: Propulsion System Decision. Industrial Base Implications of Nuclear and Non-Nuclear Options

DTIC Science & Technology

1998-01-01

industrial base may be affected by the demand for either a nuclear or non-nuclear CVX. If CVX were conventional, the cost of components for other...air- craft carrier, designated CVX. Because of the high cost of designing a new carrier and because of competing demands within its budget, the Navy...reactor re- fuelings. The organizations also supplied workforce requirements associated with different nuclear workload demands and the costs of

Feasibility of Large High-Powered Solar Electric Propulsion Vehicles: Issues and Solutions

NASA Technical Reports Server (NTRS)

Capadona, Lynn A.; Woytach, Jeffrey M.; Kerslake, Thomas W.; Manzella, David H.; Christie, Robert J.; Hickman, Tyler A.; Schneidegger, Robert J.; Hoffman, David J.; Klem, Mark D.

2012-01-01

Human exploration beyond low Earth orbit will require the use of enabling technologies that are efficient, affordable, and reliable. Solar electric propulsion (SEP) has been proposed by NASA s Human Exploration Framework Team as an option to achieve human exploration missions to near Earth objects (NEOs) because of its favorable mass efficiency as compared to traditional chemical systems. This paper describes the unique challenges and technology hurdles associated with developing a large high-power SEP vehicle. A subsystem level breakdown of factors contributing to the feasibility of SEP as a platform for future exploration missions to NEOs is presented including overall mission feasibility, trip time variables, propellant management issues, solar array power generation, array structure issues, and other areas that warrant investment in additional technology or engineering development.

Nuclear power systems for lunar and Mars exploration

NASA Technical Reports Server (NTRS)

Sovie, R. J.; Bozek, J. M.

1990-01-01

Initial studies of a variety of mission scenarios for the new Space Exploration Initiative, and the technologies necessary to enable or significantly enhance them, have identified the development of advanced space power systems whether solar, chemical or nuclear to be of prime importance. Lightweight, compact, reliable power systems for planetary rovers and a variety of surface vehicles, utility surface power, and power for advanced propulsion systems have been identified as critical needs for these missions. These mission scenarios, the concomitant power system requirements, and power system options considered are discussed. The significant potential benefits of nuclear power are identified for meeting the power needs of the above applications.

Architectural Options for a Future Deep Space Optical Communications Network

NASA Technical Reports Server (NTRS)

Edwards, B. L.; Benjamin, T.; Scozzafava, J.; Khatri, F.; Sharma, J.; Parvin, B.; Liebrecht, P. E.; Fitzgerald, R. J.

2004-01-01

This paper provides an overview of different options at Earth to provide Deep Space optical communication services. It is based mainly on work done for the Mars Laser Communications Demonstration (MLCD) Project, a joint project between NASA's Goddard Space Flight Center (GSFC), the Jet Propulsion Laboratory, California Institute of Technology (JPL), and the Massachusetts Institute of Technology Lincoln Laboratory (MIT/LL). It also reports preliminary conclusions from the Tracking and Data Relay Satellite System Continuation Study at GSFC. A lasercom flight terminal will be flown on the Mars Telecommunications Orbiter (MTO) to be launched by NASA in 2009, and will be the first high rate deep space demonstration of this revolutionary technology.

Space Nuclear Power and Propulsion: Materials Challenges for the 21st Century

NASA Technical Reports Server (NTRS)

Houts, Mike

2008-01-01

The current focus of NASA s space fission effort is Fission Surface Power (FSP). FSP systems could be used to provide power anytime, anywhere on the surface of the Moon or Mars. FSP systems could be used at locations away from the lunar poles or in permanently shaded regions, with no performance penalty. A potential reference 40 kWe option has been devised that is cost-competitive with alternatives while providing more power for less mass. The potential reference system is readily extensible for use on Mars. At Mars the system could be capable of operating through global dust storms and providing year-round power at any Martian latitude. To ensure affordability, the potential near-term, 40 kWe reference concept is designed to use only well established materials and fuels. However, if various materials challenges could be overcome, extremely high performance fission systems could be devised. These include high power, low mass fission surface power systems; in-space systems with high specific power; and high performance nuclear thermal propulsion systems. This tutorial will provide a brief overview of space fission systems and will focus on materials challenges that, if overcome, could help enable advanced exploration and utilization of the solar system.

Nuclear Thermal Propulsion (NTP): A Proven Growth Technology for Human NEO/Mars Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2012-01-01

The nuclear thermal rocket (NTR) represents the next "evolutionary step" in high performance rocket propulsion. Unlike conventional chemical rockets that produce their energy through combustion, the NTR derives its energy from fission of Uranium-235 atoms contained within fuel elements that comprise the engine s reactor core. Using an "expander" cycle for turbopump drive power, hydrogen propellant is raised to a high pressure and pumped through coolant channels in the fuel elements where it is superheated then expanded out a supersonic nozzle to generate high thrust. By using hydrogen for both the reactor coolant and propellant, the NTR can achieve specific impulse (Isp) values of 900 seconds (s) or more - twice that of today s best chemical rockets. From 1955 - 1972, twenty rocket reactors were designed, built and ground tested in the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs. These programs demonstrated: (1) high temperature carbide-based nuclear fuels; (2) a wide range of thrust levels; (3) sustained engine operation; (4) accumulated lifetime at full power; and (5) restart capability - all the requirements needed for a human Mars mission. Ceramic metal "cermet" fuel was pursued as well, as a backup option. The NTR also has significant "evolution and growth" capability. Configured as a "bimodal" system, it can generate its own electrical power to support spacecraft operational needs. Adding an oxygen "afterburner" nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, no large technology scale-ups are required for NTP either. In fact, the smallest engine tested during the Rover program - the 25,000 lbf (25 klbf) "Pewee" engine is sufficient when used in a clustered engine arrangement. The "Copernicus" crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth object (NEO) and Mars orbital missions prior to a Mars landing mission. The paper also discusses NASA s current activities and future plans for NTP development that include system-level Technology Demonstrations - specifically ground testing a small, scalable NTR by 2020, with a flight test shortly thereafter.

Nuclear thermal propulsion engine system design analysis code development

NASA Astrophysics Data System (ADS)

Pelaccio, Dennis G.; Scheil, Christine M.; Petrosky, Lyman J.; Ivanenok, Joseph F.

1992-01-01

A Nuclear Thermal Propulsion (NTP) Engine System Design Analyis Code has recently been developed to characterize key NTP engine system design features. Such a versatile, standalone NTP system performance and engine design code is required to support ongoing and future engine system and vehicle design efforts associated with proposed Space Exploration Initiative (SEI) missions of interest. Key areas of interest in the engine system modeling effort were the reactor, shielding, and inclusion of an engine multi-redundant propellant pump feed system design option. A solid-core nuclear thermal reactor and internal shielding code model was developed to estimate the reactor's thermal-hydraulic and physical parameters based on a prescribed thermal output which was integrated into a state-of-the-art engine system design model. The reactor code module has the capability to model graphite, composite, or carbide fuels. Key output from the model consists of reactor parameters such as thermal power, pressure drop, thermal profile, and heat generation in cooled structures (reflector, shield, and core supports), as well as the engine system parameters such as weight, dimensions, pressures, temperatures, mass flows, and performance. The model's overall analysis methodology and its key assumptions and capabilities are summarized in this paper.

Radioisotope Stirling Generator Options for Pluto Fast Flyby Mission

DOE Office of Scientific and Technical Information (OSTI.GOV)

Schock, Alfred

The preceding paper described conceptual designs and analytical results for five Radioisotope Thermoelectric Generator (RTG) options for the Pluto Fast Flyby (PFF) mission, and the present paper describes three Radioisotope Stirling Generator (RSG) options for the same mission. The RSG options are based on essentially the same radioisotope heat source modules used in previously flown RTGs and on designs and analyses of a 75-watt free-piston Stirling engine produced by Mechanical Technology Incorporated (MTI) for NASA's Lewis Research Center. The integrated system design options presented were generated in a Fairchild Space study sponsored by the Department of Energy's Office of Specialmore » Applications, in support of ongoing PFF mission and spacecraft studies that the Jet Propulsion Laboratory (JPL) is conducting for the National Aeronautics and Space Administration (NASA). That study's NASA-directed goal is to reduce the spacecraft mass from its baseline value of 166 kg to ~110 kg, which implies a mass goal of less than 10 kg for a power source able to deliver 69 watts(e) at the end of the 9.2-year mission. In general, the Stirling options were found to be lighter than the thermoelectric options described in the preceding paper. But they are less mature, requiring more development, and entailing greater programmatic risk. The Stirling power system mass ranged from 7.3 kg (well below the 10-kg goal) for a non-redundant system to 11.3 kg for a redundant system able to maintain full power if one of its engines fails. In fact, the latter system could deliver as much as 115 watts(e) if desired by the mission planners. There are 2 copies in the file.« less

A candidate architecture for monitoring and control in chemical transfer propulsion systems

NASA Technical Reports Server (NTRS)

Binder, Michael P.; Millis, Marc G.

1990-01-01

To support the exploration of space, a reusable space-based rocket engine must be developed. This engine must sustain superior operability and man-rated levels of reliability over several missions with limited maintenance or inspection between flights. To meet these requirements, an expander cycle engine incorporating a highly capable control and health monitoring system is planned. Alternatives for the functional organization and the implementation architecture of the engine's monitoring and control system are discussed. On the basis of this discussion, a decentralized architecture is favored. The trade-offs between several implementation options are outlined and future work is proposed.

In-Space Transportation Propulsion Architecture Assessment

NASA Technical Reports Server (NTRS)

Woodcock, Gordon

2000-01-01

Almost all space propulsion development and application has been chemical. Aerobraking has been used at Venus and Mars, and for entry at Jupiter. One electric propulsion mission has been flown (DS-1) and electric propulsion is in general use by commercial communications satellites for stationkeeping. Gravity assist has been widely used for high-energy missions (Voyager, Galileo, Cassini, etc.). It has served as a substitute for high-energy propulsion but is limited in energy gain, and adds mission complexity as well as launch opportunity restrictions. It has very limited value for round trip missions such as humans to Mars and return. High-energy space propulsion has been researched for many years, and some major developments, such as nuclear thermal propulsion (NTP), undertaken. With the exception of solar electric propulsion at a scale of a few kilowatts, high-energy space propulsion has never been used on a mission. Most mission studies have adopted TRL 6 technology because most have looked for a near-term start. The current activity is technology planning aimed at broadening the options available to mission planners. Many of the illustrations used in this report came from various NASA sources; their use is gratefully acknowledged.

Heatpipe power system and heatpipe bimodal system design and development options

NASA Technical Reports Server (NTRS)

Houts, M. G.; Poston, D. I.; Emrich, W. J., Jr.

1997-01-01

The Heatpipe Power System (HPS) is a potential, near-term, low-cost space fission power system. The Heatpipe Bimodal System (HBS) is a potential, near-term, low-cost space fission power and/or propulsion system. Both systems will be composed of independent modules, and all components operate within the existing databases. The HPS and HBS have relatively few system integration issues; thus, the successful development of a module is a significant step toward verifying system feasibility and performance estimates. A prototypic HPS module is being fabricated, and testing is scheduled to begin in November 1996. A successful test will provide high confidence that the HPS can achieve its predicted performance.

Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Mitchell, Sonny; Houts, Michael G.; Kim, Tony

2015-01-01

Development efforts in the United States for nuclear thermal propulsion (NTP) systems began with Project Rover (1955-1973) which completed 22 high-power rocket reactor tests. Results indicated that an NTP system with a high thrust-to-weight ratio and a specific impulse greater than 900 s would be feasible. John F. Kennedy, in his historic special address to Congress on the importance of Space on May 25, 1961, said, "First, I believe that this nation should commit itself to achieving the goal, before this decade is out, of landing a man on the Moon and returning him safely to the Earth..." This was accomplished. He also said, "Secondly ... accelerate development of the Rover nuclear rocket. This gives promise of someday providing a means for even more exciting and ambitious exploration of space... to the very end of the solar system itself." The current NTP project focuses on demonstrating the affordability and viability of a fully integrated NTP system with emphasis on fuel fabrication and testing and an affordable development and qualification strategy. The goal is to enable NTP to be considered a mainstream option for supporting human Mars and other missions beyond Earth orbit.

Design of a High Temperature Radiator for the Variable Specific Impulse Magnetoplasma Rocket

NASA Technical Reports Server (NTRS)

Sheth, Rubik B.; Ungar, Eugene K.; Chambliss, Joe P.

2012-01-01

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR), currently under development by Ad Astra Rocket Company (Webster, TX), is a unique propulsion system that could change the way space propulsion is performed. VASIMR's efficiency, when compared to that of a conventional chemical rocket, reduces the propellant needed for exploration missions by a factor of 10. Currently plans include flight tests of a 200 kW VASIMR system, titled VF-200, on the International Space Station (ISS). The VF-200 will consist of two 100 kW thruster units packaged together in one engine bus. Each thruster core generates 27 kW of waste heat during its 15 minute firing time. The rocket core will be maintained between 283 and 573 K by a pumped thermal control loop. The design of a high temperature radiator is a unique challenge for the vehicle design. This paper will discuss the path taken to develop a steady state and transient-based radiator design. The paper will describe the radiator design option selected for the VASIMR thermal control system for use on ISS, and how the system relates to future exploration vehicles.

Initial Test Firing Results for Solid CO/GOX Cryogenic Hybrid Rocket Engine for Mars ISRU Propulsion Applications

NASA Technical Reports Server (NTRS)

Rice, Eric E.; St. Clair, Christopher P.; Chiaverini, Martin J.; Knuth, William H.; Gustafson, Robert J.; Gramer, Daniel J.

1999-01-01

ORBITEC is developing methods for producing, testing, and utilizing Mars-based ISRU fuel/oxidizer combinations to support low cost, planetary surface and flight propulsion and power systems. When humans explore Mars we will need to use in situ resources that are available, such as: energy (solar); gases or liquids for life support, ground transportation, and flight to and from other surface locations and Earth; and materials for shielding and building habitats and infrastructure. Probably the easiest use of Martian resources to reduce the cost of human exploration activities is the use of the carbon and oxygen readily available from the CO2 in the Mars atmosphere. ORBITEC has conducted preliminary R&D that will eventually allow us to reliably use these resources. ORBITEC is focusing on the innovative use of solid CO as a fuel. A new advanced cryogenic hybrid rocket propulsion system is suggested that will offer advantages over LCO/LOX propulsion, making it the best option for a Mars sample return vehicle and other flight vehicles. This technology could also greatly support logistics and base operations by providing a reliable and simple way to store solar or nuclear generated energy in the form of chemical energy that can be used for ground transportation (rovers/land vehicles) and planetary surface power generators. This paper describes the overall concept and the test results of the first ever solid carbon monoxide/oxygen rocket engine firing.

Propulsion Options for the HI SPOT Long Endurance Drone Airship

DTIC Science & Technology

1979-09-15

Carnot Stirling Ranklne Entropy Entropy Entropy Temper- ature ’iconst P ’Ccost V Coat P 01 1 ___PConsit V Cnt V C:)nst P Diesel Otto Brayton Entropy...ignition) and Brayton (gas turbine) systems. All of these are within a few percentage points of efficiency, though the Brayton engine is generally less...fuel consumption. The ultimate lightweight engine is the gas turbine, or Brayton cycle engine. However, while good specific fuel consumption can be

Propulsive Small Expendable Deployer System (ProSEDS)

NASA Technical Reports Server (NTRS)

Curtis, Leslie; Johnson, Les; Brown, Norman S. (Technical Monitor)

2002-01-01

The Propulsive Small Expendable Deployer System (ProSEDS) space experiment will demonstrate the use of an electrodynamic tether propulsion system to generate thrust in space by decreasing the orbital altitude of a Delta 11 Expendable Launch Vehicle second stage. ProSEDS, which is planned on an Air Force GPS Satellite replacement mission in June 2002, will use the flight proven Small Expendable Deployer System (SEDS) to deploy a tether (5 km bare wire plus 10 km non-conducting Dyneema) from a Delta 11 second stage to achieve approx. 0.4N drag thrust. ProSEDS will utilize the tether-generated current to provide limited spacecraft power. The ProSEDS instrumentation includes Langmuir probes and Differential Ion Flux Probes, which will determine the characteristics of the ambient ionospheric plasma. Two Global Positioning System (GPS) receivers will be used (one on the Delta and one on the endmass) to help determine tether dynamics and to limit transmitter operations to occasions when the spacecraft is over selected ground stations. The flight experiment is a precursor to the more ambitious electrodynamic tether upper stage demonstration mission, which will be capable of orbit raising, lowering and inclination changes-all using electrodynamic thrust. An immediate application of ProSEDS technology is for the removal of spent satellites for orbital debris mitigation. In addition to the use of this technology to provide orbit transfer and debris mitigation it may also be an attractive option for future missions to Jupiter and any other planetary body with a magnetosphere.

Propulsive Small Expendable Deployer System (ProSEDS)

NASA Technical Reports Server (NTRS)

Ballance, Judy; Johnson, Les; Rogacki, John R. (Technical Monitor)

2000-01-01

The Propulsive Small Expendable Deployer System (ProSEDS) space experiment will demonstrate the use of an electrodynamic tether propulsion system to generate thrust in space by decreasing the orbital altitude of a Delta II Expendable Launch Vehicle (ELV) second stage. ProSEDS, which is planned to fly in 2001, will use the flight proven Small Expendable Deployer System (SEDS) to deploy a tether (5km bare wire plus 10 km spectra or dyneema) from a Delta II second stage to achieve approximately 0.4N drag thrust. ProSEDS will utilize the tether-generated current to provide limited spacecraft power. The ProSEDs instrumentation includes a Langmuir probe and Differential Ion Flux Probe, which will determine the characteristics of the ambient ionospheric plasma. Two Global Positioning System (GPS) receivers will be used (one on the Delta and one on the endmass) to help determine tether dynamics and to limit transmitter operations to occasions when the spacecraft is over selected ground stations, The flight experiment is a precursor to the more ambitious electrodynamic tether upper stage demonstration mission, which will be capable of orbit raising, lowering and inclination changes-all using electrodynamic thrust. An immediate application of ProSEDS technology is for the deorbit of spent satellites for orbital debris mitigation. In addition to the use of this technology to provide orbit transfer and debris mitigation it may also be an attractive option for future missions to Jupiter and any other planetary body with a magnetosphere.

Lunar He-3, fusion propulsion, and space development

NASA Technical Reports Server (NTRS)

Santarius, John F.

1992-01-01

The recent identification of a substantial lunar resource of the fusion energy fuel He-3 may provide the first terrestrial market for a lunar commodity and, therefore, a major impetus to lunar development. The impact of this resource-when burned in D-He-3 fusion reactors for space power and propulsion-may be even more significant as an enabling technology for safe, efficient exploration and development of space. One possible reactor configuration among several options, the tandem mirror, illustrates the potential advantages of fusion propulsion. The most important advantage is the ability to provide either fast, piloted vessels or high-payload-fraction cargo vessels due to a range of specific impulses from 50 sec to 1,000,000 sec at thrust-to-weight ratios from 0.1 to 5x10(exp -5). Fusion power research has made steady, impressive progress. It is plausible, and even probable, that fusion rockets similar to the designs presented here will be available in the early part of the twenty-first century, enabling a major expansion of human presence into the solar system.

Isolated Open Rotor Noise Prediction Assessment Using the F31A31 Historical Blade Set

NASA Technical Reports Server (NTRS)

Nark, Douglas M.; Jones, William T.; Boyd, D. Douglas, Jr.; Zawodny, Nikolas S.

2016-01-01

In an effort to mitigate next-generation fuel efficiency and environmental impact concerns for aviation, open rotor propulsion systems have received renewed interest. However, maintaining the high propulsive efficiency while simultaneously meeting noise goals has been one of the challenges in making open rotor propulsion a viable option. Improvements in prediction tools and design methodologies have opened the design space for next generation open rotor designs that satisfy these challenging objectives. As such, validation of aerodynamic and acoustic prediction tools has been an important aspect of open rotor research efforts. This paper describes validation efforts of a combined computational fluid dynamics and Ffowcs Williams and Hawkings equation methodology for open rotor aeroacoustic modeling. Performance and acoustic predictions were made for a benchmark open rotor blade set and compared with measurements over a range of rotor speeds and observer angles. Overall, the results indicate that the computational approach is acceptable for assessing low-noise open rotor designs. Additionally, this approach may be used to provide realistic incident source fields for acoustic shielding/scattering studies on various aircraft configurations.

Using Gravity Assists in the Earth-moon System as a Gateway to the Solar System

NASA Technical Reports Server (NTRS)

McElrath, Timothy P.; Lantoine, Gregory; Landau, Damon; Grebow, Dan; Strange, Nathan; Wilson, Roby; Sims, Jon

2012-01-01

For spacecraft departing the Earth - Moon system, lunar flybys can significantly increase the hyperbolic escape energy (C3, in km (exp 2) /sec (exp 2) ) for a modest increase in flight time. Within approx 2 months, lunar flybys can produce a C3 of approx 2. Over 4 - 6 months, lunar flybys alone can increase the C3 to approx 4.5, or they can provide for additional periapsis burns to increase the C3 from approx 2 -3 to 10 or more, suitable for planetary missions. A lunar flyby departure can be followed by additional delta -V (such as that efficiently provided by a low thrust system, eg. Solar Electric Propulsion (SEP)) to raise the Earth - relative velocity (at a ratio of more than 2:1) before a subsequent Earth flyby, which redirects that velocity to a more distant target, all within not more than a year. This paper describes the applicability of lunar flybys for different flight times and propulsion systems, and illustrates this with instances of past usage and future possibilities. Examples discussed include ISEE-3, Nozomi, STEREO, 2018 Mars studies (which showed an 8% payload increase), and missions to Near Earth Objects (NEOs). In addition, the options for the achieving the initial lunar flyby are systematically discussed, with a view towards their practical use within a compact launch period. In particular, we show that launches to geosynchronous transfer orbit (GTO) as a secondary payload provide a feasible means of obtaining a lunar flyby for an acceptable cost, even for SEP systems that cannot easily deliver large delta-Vs at periapsis. Taken together, these results comprise a myriad of options for increasing the mission performance, by the efficient use of lunar flybys within an acceptable extension of the flight time.

Mission analysis for the Martian Moons Explorer (MMX) mission

NASA Astrophysics Data System (ADS)

Campagnola, Stefano; Yam, Chit Hong; Tsuda, Yuichi; Ogawa, Naoko; Kawakatsu, Yasuhiro

2018-05-01

Mars Moon eXplorer (MMX) is JAXA's next candidate flagship mission to be launched in the early 2020s. MMX will explore the Martian moons and return a sample from Phobos. This paper presents the mission analysis work, focusing on the transfer legs and comparing several architectures, such as hybrid options with chemical and electric propulsion modules. The selected baseline is a chemical-propulsion Phobos sample return, which is discussed in detail with the launch- and return-window analysis. The trajectories are optimized with the jTOP software, using planetary ephemerides for Mars and the Earth; Earth re-entry constraints are modeled with simple analytical equations. Finally, we introduce an analytical approximation of the three-burn capture strategy used in the Mars system. The approximation can be used together with a Lambert solver to quickly determine the transfer Δ v costs.

Mars transit vehicle thermal protection system: Issues, options, and trades

NASA Technical Reports Server (NTRS)

Brown, Norman

1986-01-01

A Mars mission is characterized by different mission phases. The thermal control of cryogenic propellant in a propulsive vehicle must withstand the different mission environments. Long term cryogenic storage may be achieved by passive or active systems. Passive cryo boiloff management features will include multilayer insulation, vapor cooled shield, and low conductance structural supports and penetrations. Active boiloff management incorporates the use of a refrigeration system. Key system trade areas include active verses passive system boiloff management (with respect to safety, reliability, and cost) and propellant tank insulation optimizations. Technology requirements include refrigeration technology advancements, insulation performance during long exposure, and cryogenic fluid transfer system for mission vehicle propellant tanking during vehicle buildip in LEO.

Places Only Sails Can Go

NASA Technical Reports Server (NTRS)

Montgomery, Edward E., IV; Heaton, Andrew F.; Garbe, Gregory P.

2003-01-01

Solar sails are a near term, low thrust, propellantless propulsion technology suitable for orbital maneuvering, station keeping, and attitude control applications for small payloads. Furthermore, these functions can be highly integrated, reducing mass, cost and complexity. The solar sail concept is based on momentum exchange with solar flux reflected from a large, deployed thin membrane. Thrust performance increases as the square of the distance to the sun. In comparison to conventional chemical systems, there are missions where solar sails are vastly more and less economical. The less attractive applications involve large payloads, outer solar system transfers, and short trip times. However, for inclination changes and station keeping at locations requiring constant thrust, the solar sail is the only economical option for missions of more than a few weeks duration. We compare the location and energies required for these applications between solar sails, advanced electric propulsion, and conventional rockets. We address the effect on mass fraction to understand solar sail mission cost and capability. Finally, the benefit of potential applications to near term science missions is reported.

Next generation: In-space transportation system(s)

NASA Technical Reports Server (NTRS)

Huffaker, Fredrick; Redus, Jerry; Kelley, David L.

1991-01-01

The development of the next generation In-Space Transportation System presents a unique challenge to the design of a propulsion system for the Space Exploration Initiative (SEI). Never before have the requirements for long-life, multiple mission use, space basing, high reliability, man-rating, and minimum maintenance come together with performance in one system that must protect the lives of space travelers, support the mission logistics needs, and do so at an acceptable cost. The challenge that is presented is to quantify the bounds of these requirements. The issue is one of degree. The length of acceptable life in space, the time it takes for reuse to pay off, and the degree to which space basing is practical (full, partial, or expended) are the issues that determine the reusable bounds of a design and include dependability, contingency capabilities, resilency, and minimum dependence on a maintenance node in preparation for and during a mission. Missions to planet earth, other non-NASA missions, and planetary missions will provide important but less demanding requirements for the transportation systems of the future. The mission proposed for the SEI require a family of transportation vehicles to meet the requirements for establishing a permanent human presence on the Moon and eventually on Mars. Specialized vehicles are needed to accomplish the different phases of each mission. These large scale missions require assembly in space and will provide the greatest usage of the planned integrated transportation system. The current approach to defining the In-Space Transportation System for the SEI Moon missions with later Mars mission applications is presented. Several system development options, propulsion concepts, current/proposed activities are reviewed, and key propulsion design criteria, issues, and technology challenges for the next generation In-Space Transportation System(s) are outlined.

Design of a power management and distribution system for a thermionic-diode powered spacecraft

NASA Technical Reports Server (NTRS)

Kimnach, Greg L.

1996-01-01

The Electrical Systems Development Branch of the Power Technology Division at the NASA Lewis Research Center in Cleveland, Ohio is designing a Power Management and Distribution (PMAD) System for the Air Force's Integrated Solar Upper Stage (ISUS) Engine Ground Test Demonstration (EGD). The ISUS program uses solar-thermal propulsion to perform orbit transfers from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO) and from LEO to Molnya. The ISUS uses the same energy conversion receiver to perform the LEO to High Earth Orbit (HEO) transfer and to generate on-orbit electric power for the payloads. On-orbit power generation is accomplished via two solar concentrators heating a dual-cavity graphite-core which has Thermionic Diodes (TMD's) encircling each cavity. The graphite core and concentrators together are called the Receiver and Concentrator (RAC). The TDM-emitters reach peak temperatures of approximately 2200K, and the TID-collectors are run at approximately 1000K. Because of the high Specific Impulse (I(sup sp)) of solar thermal propulsion relative to chemical propulsion, and because a common bus is used for communications, GN&C, power, etc., a substantial increase in payload weight is possible. This potentially allows for a stepdown in the required launch vehicle size or class for similar payload weight using conventional chemical propulsion and a separate spacecraft bus. The ISUS power system is to provide 1000W(sub e) at 28+/-6V(sub dc) to the payload/spacecraft from a maximum TID generation capability of 1070W(sub e) at 2200K. Producing power with this quality, protecting the spacecraft from electrical faults and accommodating operational constraints of the TID's are the responsibilities of the PMAD system. The design strategy and system options examined along with the proposed designs for the Flight and EGD configurations are discussed herein.

Electric Propulsion Upper-Stage for Launch Vehicle Capability Enhancement

NASA Technical Reports Server (NTRS)

Kemp, Gregory E.; Dankanich, John W.; Woodcock, Gordon R.; Wingo, Dennis R.

2007-01-01

The NASA In-Space Propulsion Technology Project Office initiated a preliminary study to evaluate the performance benefits of a solar electric propulsion (SEP) upper-stage with existing and near-term small launch vehicles. The analysis included circular and elliptical Low Earth Orbit (LEO) to Geosynchronous Earth Orbit (GEO) transfers, and LEO to Low Lunar Orbit (LLO) applications. SEP subsystem options included state-of-the-art and near-term solar arrays and electric thrusters. In-depth evaluations of the Aerojet BPT-4000 Hall thruster and NEXT gridded ion engine were conducted to compare performance, cost and revenue potential. Preliminary results indicate that Hall thruster technology is favored for low-cost, low power SEP stages, while gridded-ion engines are favored for higher power SEP systems unfettered by transfer time constraints. A low-cost point design is presented that details one possible stage configuration and outlines system limitations, in particular fairing volume constraints. The results demonstrate mission enhancements to large and medium class launch vehicles, and mission enabling performance when SEP system upper stages are mounted to low-cost launchers such as the Minotaur and Falcon 1. Study results indicate the potential use of SEP upper stages to double GEO payload mass capability and to possibly enable launch on demand capability for GEO assets. Transition from government to commercial applications, with associated cost/benefit analysis, has also been assessed. The sensitivity of system performance to specific impulse, array power, thruster size, and component costs are also discussed.

Advanced Chemical Propulsion Study

NASA Technical Reports Server (NTRS)

Woodcock, Gordon; Byers, Dave; Alexander, Leslie A.; Krebsbach, Al

2004-01-01

A study was performed of advanced chemical propulsion technology application to space science (Code S) missions. The purpose was to begin the process of selecting chemical propulsion technology advancement activities that would provide greatest benefits to Code S missions. Several missions were selected from Code S planning data, and a range of advanced chemical propulsion options was analyzed to assess capabilities and benefits re these missions. Selected beneficial applications were found for higher-performing bipropellants, gelled propellants, and cryogenic propellants. Technology advancement recommendations included cryocoolers and small turbopump engines for cryogenic propellants; space storable propellants such as LOX-hydrazine; and advanced monopropellants. It was noted that fluorine-bearing oxidizers offer performance gains over more benign oxidizers. Potential benefits were observed for gelled propellants that could be allowed to freeze, then thawed for use.

Turbulent Mixing and Combustion for High-Speed Air-Breathing Propulsion Application

DTIC Science & Technology

2007-08-12

deficit (the velocity of the wake relative to the free-stream velocity), decays rapidly with downstream distance, so that the streamwise velocity is...switched laser with double-pulse option) and a new imaging system (high-resolution: 4008x2672 pix2, low- noise (cooled) Cooke PCO-4000 CCD camera). The...was designed in-house for high-speed low- noise image acquisition. The KFS CCD image sensor was designed by Mark Wadsworth of JPL and has a resolution

Low Cost Space Demonstration for a Single-Person Spacecraft

NASA Technical Reports Server (NTRS)

Griffin, Brand N.; Dischinger, Charles

2011-01-01

This paper introduces a concept for a single-person spacecraft and presents plans for flying a low-cost, robotic demonstration mission. Called FlexCraft, the vehicle integrates propulsion and robotics into a small spacecraft that enables rapid, shirt-sleeve access to space. It can be flown by astronauts or tele-operated and is equipped with interchangeable manipulators used for maintaining the International Space Station (ISS), exploring asteroids, and servicing telescopes or satellites. Most FlexCraft systems are verified using ground facilities; however, a test in the weightless environment is needed to assess propulsion and manipulator performance. For this, a simplified, unmanned, version of FlexCraft is flown on a low-cost launch vehicle to a 350 km circular orbit. After separation from the upper stage, the vehicle returns to a target box mounted on the stage testing the propulsion and control capability. The box is equipped with manipulator test items that are representative of tasks performed on ISS, asteroid missions, or for satellites servicing. Nominal and off-nominal operations are conducted over 3 days then the vehicle re-enters the atmosphere without becoming a debris hazard. From concept to management to operations, the FlexCraft demonstration is designed to be low cost project that is launched within three years. This is possible using a simplified test configuration that eliminates nine systems unique to the operational version and by designing-to-availability. For example, the propulsion system is the same as the Manned Maneuvering Unit because it capable, simple, human-rated and all components or equivalent parts are available. A description of the launch vehicle options, mission operations, configuration, and demonstrator subsystems is presented.

Second Beamed Space-Power Workshop

NASA Technical Reports Server (NTRS)

Deyoung, Russell J. (Editor)

1989-01-01

Potential missions for microwave and laser power beaming in space are discussed. Power beaming options, millimeter wave technology, laser technology, lunar bases, spacecraft propulsion, and near-Earth applications are covered.

Earth reencounter probabilities for aborted space disposal of hazardous nuclear waste

NASA Technical Reports Server (NTRS)

Friedlander, A. L.; Feingold, H.

1977-01-01

A quantitative assessment is made of the long-term risk of earth reencounter and reentry associated with aborted disposal of hazardous material in the space environment. Numerical results are presented for 10 candidate disposal options covering a broad spectrum of disposal destinations and deployment propulsion systems. Based on representative models of system failure, the probability that a single payload will return and collide with earth within a period of 250,000 years is found to lie in the range .0002-.006. Proportionately smaller risk attaches to shorter time intervals. Risk-critical factors related to trajectory geometry and system reliability are identified as possible mechanisms of hazard reduction.

TOPEX satellite option study

NASA Technical Reports Server (NTRS)

1982-01-01

The basic design of the fleet satellite communication spacecraft (FLTSATCOM) can easily accommodate any of the three payload options for the ocean dynamic topography experiment (TOPEX). The principal mission requirements as well as the payload accommodations and communications systems needed for launching this payload are reviewed. The existing FLTSATCOM satellite design is identified and the approaches for the proposed propulsion system are described in addition to subsystems for mechanical; power; attitude and velocity control; and telemetry, tracking and control are described. The compatability of FLTSATCOM with the launch vehicle is examined and its capabilities vs TOPEX requirements are summarized. Undetermined changes needed to meet data storage, thermal control, and area to mass ratio requirements are discussed. Cost estimates are included for budgetary and planning purposes. The availability of the described design is assessed based on the continuing production of FLTSATCOM spacecraft during the schedule span planned for TOPEX.

Propulsion Systems Panel deliberations

NASA Technical Reports Server (NTRS)

Bianca, Carmelo J.; Miner, Robert; Johnston, Lawrence M.; Bruce, R.; Dennies, Daniel P.; Dickenson, W.; Dreshfield, Robert; Karakulko, Walt; Mcgaw, Mike; Munafo, Paul M.

1993-01-01

The Propulsion Systems Panel was established because of the specialized nature of many of the materials and structures technology issues related to propulsion systems. This panel was co-chaired by Carmelo Bianca, MSFC, and Bob Miner, LeRC. Because of the diverse range of missions anticipated for the Space Transportation program, three distinct propulsion system types were identified in the workshop planning process: liquid propulsion systems, solid propulsion systems and nuclear electric/nuclear thermal propulsion systems.

Propulsion Systems Panel deliberations

NASA Astrophysics Data System (ADS)

Bianca, Carmelo J.; Miner, Robert; Johnston, Lawrence M.; Bruce, R.; Dennies, Daniel P.; Dickenson, W.; Dreshfield, Robert; Karakulko, Walt; McGaw, Mike; Munafo, Paul M.

1993-02-01

The Propulsion Systems Panel was established because of the specialized nature of many of the materials and structures technology issues related to propulsion systems. This panel was co-chaired by Carmelo Bianca, MSFC, and Bob Miner, LeRC. Because of the diverse range of missions anticipated for the Space Transportation program, three distinct propulsion system types were identified in the workshop planning process: liquid propulsion systems, solid propulsion systems and nuclear electric/nuclear thermal propulsion systems.

Study of a Tricarbide Grooved Ring Fuel Element for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Taylor, Brian; Emrich, Bill; Tucker, Dennis; Barnes, Marvin; Donders, Nicolas; Benensky, Kelsa

2018-01-01

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nuclear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the operating temperature of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment. Results of experiments and associated analysis are discussed. The authors demonstrated success in reaching desired densities with some success in material distribution and reaching a solid solution. Future work is needed to improve distribution of material, minimize oxidation during the milling process, and define a fabrication process that will serve for constructing grooved ring fuel rods for large system tests.

Study of a Tricarbide Grooved Ring Fuel Element for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Taylor, Brian; Emrich, Bill; Tucker, Dennis; Barnes, Marvin; Donders, Nicolas; Benensky, Kelsa

2018-01-01

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nu- clear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the operating temperature of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment. Results of experiments and associated analysis are desired densities with some success in material distribution and reaching a solid solution. Future work is needed to improve distribution of material, minimize oxidation during the milling process, and de ne a fabrication process that will serve for constructing grooved ring fuel rods for large system tests.

Forest fire advanced system technology (FFAST) conceptual design study

NASA Technical Reports Server (NTRS)

Nichols, J. David; Warren, John R.

1987-01-01

The National Aeronautics and Space Administration's Jet Propulsion Laboratory (JPL) and the U.S. Department of Agriculture (USDA) Forest Service completed a conceptual design study that defined an integrated forest fire detection and mapping system that will be based upon technology available in the 1990s. Potential system configuration options in emerging and advanced technologies related to the conceptual design were identified and recommended for inclusion as preferred system components. System component technologies identified for an end-to-end system include airborne mounted, thermal infrared (IR) linear array detectors, automatic onboard georeferencing and signal processing, geosynchronous satellite communications links, and advanced data integration and display. Potential system configuration options were developed and examined for possible inclusion in the preferred system configuration. The preferred system configuration will provide increased performance and be cost effective over the system currently in use. Forest fire management user requirements and the system component emerging technologies were the basis for the system configuration design. The conceptual design study defined the preferred system configuration that warrants continued refinement and development, examined economic aspects of the current and preferred system, and provided preliminary cost estimates for follow-on system prototype development.

In-situ propellant rocket engines for Mars missions ascent vehicle

NASA Technical Reports Server (NTRS)

Roncace, Elizabeth A.

1991-01-01

When contemplating the human exploration of Mars, many scenarios using various propulsion systems have been considered. One propulsion option among them is a vehicle stage with multiple, pump fed rocket engines capable of operating on propellants available on Mars. This reduces the earth launch mass requirements, resulting in economic and payload benefits. No plentiful sources of hydrogen on Mars have been identified on the surface of Mars, so most commonly used high performance liquid fuels, such as hydrogen and hydrocarbons, can be eliminated as possible in situ propellants. But 95 pct of the Martian atmosphere consists of carbon dioxide, which can be converted into carbon monoxide and oxygen. The carbon monoxide oxygen propellant combination is a candidate for a Martian in situ propellant rocket engine. The feasibility is analyzed of a pump fed engine cycle using the propellant combination of carbon monoxide and oxygen.

In-situ propellant rocket engines for Mars mission ascent vehicle

NASA Technical Reports Server (NTRS)

Roncace, Elizabeth A.

1991-01-01

When comtemplating the human exploration of Mars, many scenarios using various propulsion systems have been considered. One propulsion option among them is a vehicle stage with multiple, pump fed rocket engines capable of operating on propellants available on Mars. This reduces the Earth launch mass requirements, resulting in economic and payload benefits. No plentiful sources of hydrogen on Mars have been identified on the surface of Mars, so most commonly used high performance liquid fuels, such as hydrogen and hydrocarbons, can be eliminated as possible in-situ propellants. But 95 pct. of the Martian atmosphere consists of carbon dioxide, which can be converted into carbon monoxide and oxygen. The carbon monoxide oxygen propellant conbination is a candidate for a Martian in-situ propellant rocket engine. The feasibility is analyzed of a pump fed engine cycle using the propellant combination of carbon monoxide and oxygen.

The Iodine Satellite (iSat) Project Development Towards Critical Design Review

NASA Technical Reports Server (NTRS)

Dankanich, John W.; Calvert, Derek; Kamhawi, Hani; Hickman, Tyler; Szabo, James; Byrne, Lawrence

2015-01-01

Despite the prevalence of small satellites in recent years, the systems flown to date have very limited propulsion capability. SmallSats are typically secondary payloads and have significant constraints for volume, mass, and power in addition to limitations on the use of hazardous propellants or stored energy. These constraints limit the options for SmallSat maneuverability. NASA's Space Technology Mission Directorate approved the iodine Satellite flight project for a rapid demonstration of iodine Hall thruster technology in a 12U (cubesat units) configuration under the Small Spacecraft Technology Program. The mission is a partnership between NASA MSFC, NASA GRC, and Busek Co, Inc., with the Air Force supporting the propulsion technology maturation. The team is working towards the critical design review in the final design and fabrication phase of the project. The current design shows positive technical performance margins in all areas. The iSat project is planned for launch readiness in the spring of 2017.

Advanced Single-Aisle Transport Propulsion Design Options Revisited

NASA Technical Reports Server (NTRS)

Guynn, Mark D.; Berton, Jeffrey J.; Tong, Michael T.; Haller, William J.

2013-01-01

Future propulsion options for advanced single-aisle transports have been investigated in a number of previous studies by the authors. These studies have examined the system level characteristics of aircraft incorporating ultra-high bypass ratio (UHB) turbofans (direct drive and geared) and open rotor engines. During the course of these prior studies, a number of potential refinements and enhancements to the analysis methodology and assumptions were identified. This paper revisits a previously conducted UHB turbofan fan pressure ratio trade study using updated analysis methodology and assumptions. The changes incorporated have decreased the optimum fan pressure ratio for minimum fuel consumption and reduced the engine design trade-offs between minimizing noise and minimizing fuel consumption. Nacelle drag and engine weight are found to be key drivers in determining the optimum fan pressure ratio from a fuel efficiency perspective. The revised noise analysis results in the study aircraft being 2 to 4 EPNdB (cumulative) quieter due to a variety of reasons explained in the paper. With equal core technology assumed, the geared engine architecture is found to be as good as or better than the direct drive architecture for most parameters investigated. However, the engine ultimately selected for a future advanced single-aisle aircraft will depend on factors beyond those considered here.

Advanced planning activity. [for interplanetary flight and space exploration

NASA Technical Reports Server (NTRS)

1974-01-01

Selected mission concepts for interplanetary exploration through 1985 were examined, including: (1) Jupiter orbiter performance characteristics; (2) solar electric propulsion missions to Mercury, Venus, Neptune, and Uranus; (3) space shuttle planetary missions; (4) Pioneer entry probes to Saturn and Uranus; (5) rendezvous with Comet Kohoutek and Comet Encke; (6) space tug capabilities; and (7) a Pioneer mission to Mars in 1979. Mission options, limitations, and performance predictions are assessed, along with probable configurational, boost, and propulsion requirements.

Magnetic Nozzle Simulation Studies for Electric Propulsion

NASA Astrophysics Data System (ADS)

Tarditi, Alfonso

2010-11-01

Electric Propulsion has recently re-gained interest as one of the key technologies to enable NASA's long-range space missions. Options are being considered also in the field of aneutronic fusion propulsion for high-power electric thrusters. To support these goals the study of the exhaust jet in a plasma thruster acquires a critical importance because the need of high-efficiency generation of thrust. A model of the plasma exhaust has been developed with the 3D magneto-fluid NIMROD code [1] to study the physics of the plasma detachment in correlation with experimentally relevant configurations. The simulations show the role of the plasma diamagnetism and of the magnetic reconnection process in the formation of a detached plasma. Furthermore, in direct fusion-propulsion concepts high-energy (MeV range) fusion products have to be efficiently converted into a slower and denser plasma jet (with specific impulse down to few 1000's seconds, for realistic missions in the Solar System). For this purpose, a two-stage conversion process is being modeled where high-energy ions are non-adiabatically injected and confined into a magnetic duct leading to the magnetic nozzle, transferring most of their energy into their gyro-motion and drifting at slower speed along with the plasma propellant. The propellant acquires then thermal energy that gets converted into the direction of thrust by the magnetic nozzle. [1] C. R. Sovinec et al., J. Comput. Phys. 195, 355 (2004).

Orbit Transfer Systems with Emphasis on Shuttle Applications, 1986-1991

NASA Technical Reports Server (NTRS)

1977-01-01

A systems study is presented for a transportation system which will follow the interim upper stage and spinning solid upper stage. Included are concepts, concept comparisons, trends, parametric data, etc. associated with the future system. Relevant technical and programmatic information is developed. This information is intended to focus future activity to identify attractive options and to summarize the major issues associated with the future development of the system. To establish a common basis for identifying current transportation concepts, an orbit transfer vehicle (OTV) is defined as a propulsive (velocity producing) rocket or stage. When used with a crew transfer module, a manned sortie module or other payloads, the combination becomes an orbit transfer system (OTS). Standardization of OTV's and OTS's is required.

Waste Management Options for Long-Duration Space Missions: When to Reject, Reuse, or Recycle

NASA Technical Reports Server (NTRS)

Linne, Diane L.; Palaszewski, Bryan A.; Gokoglu, Suleyman; Gallo, Christopher A.; Balasubramaniam, Ramaswamy; Hegde, Uday G.

2014-01-01

The amount of waste generated on long-duration space missions away from Earth orbit creates the daunting challenge of how to manage the waste through reuse, rejection, or recycle. The option to merely dispose of the solid waste through an airlock to space was studied for both Earth-moon libration point missions and crewed Mars missions. Although the unique dynamic characteristics of an orbit around L2 might allow some discarded waste to intersect the lunar surface before re-impacting the spacecraft, the large amount of waste needed to be managed and potential hazards associated with volatiles recondensing on the spacecraft surfaces make this option problematic. A second option evaluated is to process the waste into useful gases to be either vented to space or used in various propulsion systems. These propellants could then be used to provide the yearly station-keeping needs at an L2 orbit, or if processed into oxygen and methane propellants, could be used to augment science exploration by enabling lunar mini landers to the far side of the moon.

ISS Contingency Attitude Control Recovery Method for Loss of Automatic Thruster Control

NASA Technical Reports Server (NTRS)

Bedrossian, Nazareth; Bhatt, Sagar; Alaniz, Abran; McCants, Edward; Nguyen, Louis; Chamitoff, Greg

2008-01-01

In this paper, the attitude control issues associated with International Space Station (ISS) loss of automatic thruster control capability are discussed and methods for attitude control recovery are presented. This scenario was experienced recently during Shuttle mission STS-117 and ISS Stage 13A in June 2007 when the Russian GN&C computers, which command the ISS thrusters, failed. Without automatic propulsive attitude control, the ISS would not be able to regain attitude control after the Orbiter undocked. The core issues associated with recovering long-term attitude control using CMGs are described as well as the systems engineering analysis to identify recovery options. It is shown that the recovery method can be separated into a procedure for rate damping to a safe harbor gravity gradient stable orientation and a capability to maneuver the vehicle to the necessary initial conditions for long term attitude hold. A manual control option using Soyuz and Progress vehicle thrusters is investigated for rate damping and maneuvers. The issues with implementing such an option are presented and the key issue of closed-loop stability is addressed. A new non-propulsive alternative to thruster control, Zero Propellant Maneuver (ZPM) attitude control method is introduced and its rate damping and maneuver performance evaluated. It is shown that ZPM can meet the tight attitude and rate error tolerances needed for long term attitude control. A combination of manual thruster rate damping to a safe harbor attitude followed by a ZPM to Stage long term attitude control orientation was selected by the Anomaly Resolution Team as the alternate attitude control method for such a contingency.

Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Corban, Robert R.; Mcguire, Melissa L.; Beke, Erik G.

1995-01-01

The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 3-10), the NTR can also be configured as a 'dual mode' system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split cargo-piloted sprint' Mars mission architecture. Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a reference 240 t-class heavy lift launch vehicle (HLLV) and smaller 120 t HLLV option. Attractive performance characteristics and high-leverage technologies associated with both the engine and stage are identified, and supporting parametric sensitivity data is provided. The potential for commonality of engine and stage components to satisfy a broad range of lunar and Mars missions is also discussed.

Spacecraft Chemical Propulsion Systems at NASA's Marshall Space Flight Center: Heritage and Capabilities

NASA Technical Reports Server (NTRS)

McRight, Patrick S.; Sheehy, Jeffrey A.; Blevins, John A.

2005-01-01

NASA Marshall Space Flight Center (MSFC) is well known for its contributions to large ascent propulsion systems such as the Saturn V and the Space Shuttle. This paper highlights a lesser known but equally rich side of MSFC - its heritage in spacecraft chemical propulsion systems and its current capabilities for in-space propulsion system development and chemical propulsion research. The historical narrative describes the efforts associated with developing upper-stage main propulsion systems such as the Saturn S-IVB as well as orbital maneuvering and reaction control systems such as the S-IVB auxiliary propulsion system, the Skylab thruster attitude control system, and many more recent activities such as Chandra, the Demonstration of Automated Rendezvous Technology, X-37, the X-38 de-orbit propulsion system, the Interim Control Module, the US Propulsion Module, and several technology development activities. Also discussed are MSFC chemical propulsion research capabilities, along with near- and long-term technology challenges to which MSFC research and system development competencies are relevant.

National Aero-Space Plane (NASP) program

NASA Technical Reports Server (NTRS)

Tank, Ming H.

1991-01-01

A program to develop the technology for reusable airbreathing hypersonic/transatmospheric vehicles is addressed. Information on the following topics is presented in viewgraph form: (1) the National Aerospace Plane (NASP) program schedule; (2) the NASP program organization; (3) competitive strategy; (4) propulsion options; (5) wind tunnel data available for NASP; (6) ground track of envelope expansion; and (7) altitude vs. Mach number. A NASP/Space Shuttle comparison, NASP configuration matrix, and the propulsion concept of a high speed scramjet are also briefly addressed.

Numerical Propulsion System Simulation (NPSS): An Award Winning Propulsion System Simulation Tool

NASA Technical Reports Server (NTRS)

Stauber, Laurel J.; Naiman, Cynthia G.

2002-01-01

The Numerical Propulsion System Simulation (NPSS) is a full propulsion system simulation tool used by aerospace engineers to predict and analyze the aerothermodynamic behavior of commercial jet aircraft, military applications, and space transportation. The NPSS framework was developed to support aerospace, but other applications are already leveraging the initial capabilities, such as aviation safety, ground-based power, and alternative energy conversion devices such as fuel cells. By using the framework and developing the necessary components, future applications that NPSS could support include nuclear power, water treatment, biomedicine, chemical processing, and marine propulsion. NPSS will dramatically reduce the time, effort, and expense necessary to design and test jet engines. It accomplishes that by generating sophisticated computer simulations of an aerospace object or system, thus enabling engineers to "test" various design options without having to conduct costly, time-consuming real-life tests. The ultimate goal of NPSS is to create a numerical "test cell" that enables engineers to create complete engine simulations overnight on cost-effective computing platforms. Using NPSS, engine designers will be able to analyze different parts of the engine simultaneously, perform different types of analysis simultaneously (e.g., aerodynamic and structural), and perform analysis in a more efficient and less costly manner. NPSS will cut the development time of a new engine in half, from 10 years to 5 years. And NPSS will have a similar effect on the cost of development: new jet engines will cost about a billion dollars to develop rather than two billion. NPSS is also being applied to the development of space transportation technologies, and it is expected that similar efficiencies and cost savings will result. Advancements of NPSS in fiscal year 2001 included enhancing the NPSS Developer's Kit to easily integrate external components of varying fidelities, providing the initial Visual-Based Syntax (VBS) capability, and developing additional capabilities to support space transportation. NPSS was supported under NASA's High Performance Computing and Communications Program. Through the NASA/Industry Cooperative Effort agreement, NASA Glenn and its industry and Government partners are developing NPSS. The NPSS team consists of propulsion experts and software engineers from GE Aircraft Engines, Pratt & Whitney, The Boeing Company, Honeywell, Rolls-Royce Corporation, Williams International, Teledyne Continental Motors, Arnold Engineering Development Center, Wright Patterson Air Force Base, and the NASA Glenn Research Center. Glenn is leading the way in developing NPSS--a method for solving complex design problems that's faster, better, and cheaper.

U.S. Space Station Freedom waste fluid disposal system with consideration of hydrazine waste gas injection thrusters

NASA Technical Reports Server (NTRS)

Winters, Brian A.

1990-01-01

The results are reported of a study of various methods for propulsively disposing of waste gases. The options considered include hydrazine waste gas injection, resistojets, and eutectic salt phase change heat beds. An overview is given of the waste gas disposal system and how hydrozine waste gas injector thruster is implemented within it. Thruster performance for various gases are given and comparisons with currently available thruster models are made. The impact of disposal on station propellant requirements and electrical power usage are addressed. Contamination effects, reliability and maintainability assessments, safety issues, and operational scenarios of the waste gas thruster and disposal system are considered.

Photovoltaic options for solar electric propulsion

NASA Technical Reports Server (NTRS)

Stella, Paul M.; Flood, Dennis J.

1990-01-01

This paper discusses both state-of-the-art and advanced development cell and array technology. Present technology includes rigid, roll-out, and foldout flexible substrate designs, with silicon and GaAs solar cells. The use of concentrator array systems is discussed based on both DOD efforts and NASA work. The benefits of advanced lightweight array technology, for both near term and far term utilization, and of advanced high efficiency thin radiation resistant cells is examined. This includes gallium arsenide/germanium, indium phosphide, and thin film devices such as copper indium disclenide.

Study of Power Options for Jupiter and Outer Planet Missions

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Fincannon, James

2015-01-01

Power for missions to Jupiter and beyond presents a challenging goal for photovoltaic power systems, but NASA missions including Juno and the upcoming Europa Clipper mission have shown that it is possible to operate solar arrays at Jupiter. This work analyzes photovoltaic technologies for use in Jupiter and outer planet missions, including both conventional arrays, as well as analyzing the advantages of advanced solar cells, concentrator arrays, and thin film technologies. Index Terms - space exploration, spacecraft solar arrays, solar electric propulsion, photovoltaic cells, concentrator, Fresnel lens, Jupiter missions, outer planets.

A feasibility assessment of installation, operation and disposal options for nuclear reactor power system concepts for a NASA growth space station

NASA Technical Reports Server (NTRS)

Bloomfield, Harvey S.; Heller, Jack A.

1987-01-01

A preliminary feasibility assessment of the integration of reactor power system concepts with a projected growth space station architecture was conducted to address a variety of installation, operational disposition, and safety issues. A previous NASA sponsored study, which showed the advantages of space station - attached concepts, served as the basis for this study. A study methodology was defined and implemented to assess compatible combinations of reactor power installation concepts, disposal destinations, and propulsion methods. Three installation concepts that met a set of integration criteria were characterized from a configuration and operational viewpoint, with end-of-life disposal mass identified. Disposal destinations that met current aerospace nuclear safety criteria were identified and characterized from an operational and energy requirements viewpoint, with delta-V energy requirement as a key parameter. Chemical propulsion methods that met current and near-term application criteria were identified and payload mass and delta-V capabilities were characterized. These capabilities were matched against concept disposal mass and destination delta-V requirements to provide the feasibility of each combination.

Aero-Assisted Pre-Stage for Ballistic and Aero-Assisted Launch Vehicles

NASA Technical Reports Server (NTRS)

Ustinov, Eugene A.

2012-01-01

A concept of an aero-assisted pre-stage is proposed, which enables launch of both ballistic and aero-assisted launch vehicles from conventional runways. The pre-stage can be implemented as a delta-wing with a suitable undercarriage, which is mated with the launch vehicle, so that their flight directions are coaligned. The ample wing area of the pre-stage combined with the thrust of the launch vehicle ensure prompt roll-out and take-off of the stack at airspeeds typical for a conventional jet airliner. The launch vehicle is separated from the pre-stage as soon as safe altitude is achieved, and the desired ascent trajectory is reached. Nominally, the pre-stage is non-powered. As an option, to save the propellant of the launch vehicle, the pre-stage may have its own short-burn propulsion system, whereas the propulsion system of the launch vehicle is activated at the separation point. A general non-dimensional analysis of performance of the pre-stage from roll-out to separation is carried out and applications to existing ballistic launch vehicle and hypothetical aero-assisted vehicles (spaceplanes) are considered.

MEMS Reaction Control and Maneuvering for Picosat Beyond LEO

NASA Technical Reports Server (NTRS)

Alexeenko, Alina

2016-01-01

The MEMS Reaction Control and Maneuvering for Picosat Beyond LEO project will further develop a multi-functional small satellite technology for low-power attitude control, or orientation, of picosatellites beyond low Earth orbit (LEO). The Film-Evaporation MEMS Tunable Array (FEMTA) concept initially developed in 2013, is a thermal valving system which utilizes capillary forces in a microchannel to offset internal pressures in a bulk fluid. The local vapor pressure is increased by resistive film heating until it exceeds meniscus strength in a nozzle which induces vacuum boiling and provides a stagnation pressure equal to vapor pressure at that point which is used for propulsion. Interplanetary CubeSats can utilize FEMTA for high slew rate attitude corrections in addition to desaturating reaction wheels. The FEMTA in cooling mode can be used for thermal control during high-power communication events, which are likely to accompany the attitude correction. Current small satellite propulsion options are limited to orbit correction whereas picosatellites are lacking attitude control thrusters. The available attitude control systems are either quickly saturated reaction wheels or movable high drag surfaces with long response times.

Solar-Powered Electric Propulsion Systems: Engineering and Applications

NASA Technical Reports Server (NTRS)

Stearns, J. W.; Kerrisk, D. J.

1966-01-01

Lightweight, multikilowatt solar power arrays in conjunction with electric propulsion offer potential improvements to space exploration, extending the usefulness of existing launch vehicles to higher-energy missions. Characteristics of solar-powered electric propulsion missions are outlined, and preliminary performance estimates are shown. Spacecraft system engineering is discussed with respect to parametric trade-offs in power and propulsion system design. Relationships between mission performance and propulsion system performance are illustrated. The present state of the art of electric propulsion systems is reviewed and related to the mission requirements identified earlier. The propulsion system design and test requirements for a mission spacecraft are identified and discussed. Although only ion engine systems are currently available, certain plasma propulsion systems offer some advantages in over-all system design. These are identified, and goals are set for plasma-thrustor systems to make them competitive with ion-engine systems for mission applications.

Introduction: The challenge of optimum integration of propulsion systems and large space structures

NASA Technical Reports Server (NTRS)

Carlisle, R. F.

1980-01-01

A functional matrix of possible propulsion system characteristics for a spacecraft for deployable and assembled spacecraft structures shows that either electric propulsion or low thrust chemical propulsion systems could provide the propulsion required. The trade-off considerations of a single propulsion engine or multiengines are outlined and it is shown that a single point engine is bounded by some upper limit of thrust for assembled spacecraft. The matrix also shows several additional functions that can be provided to the spacecraft if a propulsion system is an integral part of the spacecraft. A review of all of the functions that can be provided for a spacecraft by an integral propulsion system may result in the inclusion of the propulsion for several functions even if no single function were mandatory. Propulsion interface issues for each combination of engines are identified.

46 CFR 62.35-5 - Remote propulsion-control systems.

Code of Federal Regulations, 2014 CFR

2014-10-01

... 46 Shipping 2 2014-10-01 2014-10-01 false Remote propulsion-control systems. 62.35-5 Section 62.35... AUTOMATION Requirements for Specific Types of Automated Vital Systems § 62.35-5 Remote propulsion-control systems. (a) Manual propulsion control. All vessels having remote propulsion control from the navigating...

46 CFR 62.35-5 - Remote propulsion-control systems.

Code of Federal Regulations, 2013 CFR

2013-10-01

... 46 Shipping 2 2013-10-01 2013-10-01 false Remote propulsion-control systems. 62.35-5 Section 62.35... AUTOMATION Requirements for Specific Types of Automated Vital Systems § 62.35-5 Remote propulsion-control systems. (a) Manual propulsion control. All vessels having remote propulsion control from the navigating...

46 CFR 62.35-5 - Remote propulsion-control systems.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 46 Shipping 2 2011-10-01 2011-10-01 false Remote propulsion-control systems. 62.35-5 Section 62.35... AUTOMATION Requirements for Specific Types of Automated Vital Systems § 62.35-5 Remote propulsion-control systems. (a) Manual propulsion control. All vessels having remote propulsion control from the navigating...

46 CFR 62.35-5 - Remote propulsion-control systems.

Code of Federal Regulations, 2010 CFR

2010-10-01

... 46 Shipping 2 2010-10-01 2010-10-01 false Remote propulsion-control systems. 62.35-5 Section 62.35... AUTOMATION Requirements for Specific Types of Automated Vital Systems § 62.35-5 Remote propulsion-control systems. (a) Manual propulsion control. All vessels having remote propulsion control from the navigating...

46 CFR 62.35-5 - Remote propulsion-control systems.

Code of Federal Regulations, 2012 CFR

2012-10-01

... 46 Shipping 2 2012-10-01 2012-10-01 false Remote propulsion-control systems. 62.35-5 Section 62.35... AUTOMATION Requirements for Specific Types of Automated Vital Systems § 62.35-5 Remote propulsion-control systems. (a) Manual propulsion control. All vessels having remote propulsion control from the navigating...

Design Analysis of a High Temperature Radiator for the Variable Specific Impulse Magnetoplasma Rocket (VASIMR)

NASA Technical Reports Server (NTRS)

Sheth, Rubik B.; Ungar, Eugene K.; Chambliss, Joe P.; Cassady, Leonard D.

2011-01-01

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR), currently under development by Ad Astra Rocket Company, is a unique propulsion system that can potentially change the way space propulsion is performed. VASIMR's efficiency, when compared to that of a conventional chemical rocket, reduce propellant needed for exploration missions by a factor of 10. Currently plans include flight tests of a 200 kW VASIMR system, titled VF-200, on the International Space Station. The VF-200 will consist of two 100 kW thruster units packaged together in one engine bus. Each thruster unit has a unique heat rejection requirement of about 27 kW over a firing time of 15 minutes. In order to control rocket core temperatures, peak operating temperatures of about 300 C are expected within the thermal control loop. Design of a high temperature radiator is a unique challenge for the vehicle design. This paper will discuss the path taken to develop a steady state and transient based radiator design. The paper will describe radiator design options for the VASIMR thermal control system for use on ISS as well as future exploration vehicles.

Electric Propulsion Applications and Impacts

NASA Technical Reports Server (NTRS)

Curran, Frank M.; Wickenheiser, Timothy J.

1996-01-01

Most space missions require on-board propulsion systems and these systems are often dominant spacecraft mass drivers. Presently, on-board systems account for more than half the injected mass for commercial communications systems and even greater mass fractions for ambitious planetary missions. Anticipated trends toward the use of both smaller spacecraft and launch vehicles will likely increase pressure on the performance of on-board propulsion systems. The acceptance of arcjet thrusters for operational use on commercial communications satellites ushered in a new era in on-board propulsion and exponential growth of electric propulsion across a broad spectrum of missions is anticipated. NASA recognizes the benefits of advanced propulsion and NASA's Office of Space Access and Technology supports an aggressive On-Board Propulsion program, including a strong electric propulsion element, to assure the availability of high performance propulsion systems to meet the goals of the ambitious missions envisioned in the next two decades. The program scope ranges from fundamental research for future generation systems through specific insertion efforts aimed at near term technology transfer. The On-Board propulsion program is committed to carrying technologies to levels required for customer acceptance and emphasizes direct interactions with the user community and the development of commercial sources. This paper provides a discussion of anticipated missions, propulsion functions, and electric propulsion impacts followed by an overview of the electric propulsion element of the NASA On-Board Propulsion program.

High-speed flight propulsion systems. Progress in Astronautics and Aeronautics. Vol. 137

DOE Office of Scientific and Technical Information (OSTI.GOV)

Murthy, S.N.B.; Curran, E.T.

1991-01-01

Various papers on high-speed flight propulsion systems are presented. The topics addressed are: propulsion systems from takeoff to high-speed flight, propulsion system performance and integration for high Mach air-breathing flight, energy analysis of high-speed flight systems, waves and thermodynamics in high Mach number propulsive ducts, turbulent free shear layer mixing and combustion, turbulent mixing in supersonic combustion systems, mixing and mixing enhancement in supersonic reacting flowfields, study of combustion and heat-exchange processes in high-enthalpy short-duration facilities, and facility requirements for hypersonic propulsion system testing.

Lunar Cube Transfer Trajectory Options

NASA Technical Reports Server (NTRS)

Folta, David C.; Dichman, Don; Clark, Pamela; Haapala, Amanda; Howell, Kathleen

2014-01-01

Contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these designs can be restricted by the selection of the Cubesat subsystem design such as propulsion or communication. Nonetheless, many trajectory options can be designed with have a wide range of transfer durations, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several design options including deployment into low Earth orbit (LEO), geostationary transfer orbits (GTO), and higher energy direct lunar transfer orbits. In addition to direct transfer options from these initial orbits, we also investigate the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory. In this article we examine several design options that meet the above limited deployment and subsystem drivers. We study ways that both impulsive and low-thrust Solar Electric Propulsion (SEP) engines can be used to place the Cubesat first into a highly eccentric Earth orbit, enter the Moon's Sphere of Influence, and finally achieve a highly eccentric lunar orbit. We show that such low-thrust transfers are feasible with a realistic micro-thruster model, assuming that the Cubesat can generate sufficient power for the SEP. Two examples are shown here: (1) A Cubestat injected by Exploration Mission 1 (EM-1) then employing low thrust; and (2) a CubSat deployed in a GTO, then employing impulsive maneuvers. For the EM-1 injected initial design, we increase the EM-1 targeted lunar flyby distance to reduce the energy of the lunar flyby to match that of a typical lMoon system heteroclinic manifold. Figure 1 presents an option that encompasses the similar dynamics as that of the ARTEMIS mission design. Low-thrust maneuvers are used along the manifold trajectory to raise perigee to that of a lunar orbit, adjust the timing with respect to the Moon, rotate the line of apsides, and target a ballistic lunar encounter. In this design a second flyby decreases the orbital energy with respect to the Moon, so that C3 -0.1 km2s2. Another design, shown in Figure 2 emanates from a GTO then uses impulsive maneuvers to phase onto a local Earth-Moon manifold, which then transfers the CubeSat to a lunar encounter.

Space Launch System Upper Stage Technology Assessment

NASA Technical Reports Server (NTRS)

Holladay, Jon; Hampton, Bryan; Monk, Timothy

2014-01-01

The Space Launch System (SLS) is envisioned as a heavy-lift vehicle that will provide the foundation for future beyond low-Earth orbit (LEO) exploration missions. Previous studies have been performed to determine the optimal configuration for the SLS and the applicability of commercial off-the-shelf in-space stages for Earth departure. Currently NASA is analyzing the concept of a Dual Use Upper Stage (DUUS) that will provide LEO insertion and Earth departure burns. This paper will explore candidate in-space stages based on the DUUS design for a wide range of beyond LEO missions. Mission payloads will range from small robotic systems up to human systems with deep space habitats and landers. Mission destinations will include cislunar space, Mars, Jupiter, and Saturn. Given these wide-ranging mission objectives, a vehicle-sizing tool has been developed to determine the size of an Earth departure stage based on the mission objectives. The tool calculates masses for all the major subsystems of the vehicle including propellant loads, avionics, power, engines, main propulsion system components, tanks, pressurization system and gases, primary structural elements, and secondary structural elements. The tool uses an iterative sizing algorithm to determine the resulting mass of the stage. Any input into one of the subsystem sizing routines or the mission parameters can be treated as a parametric sweep or as a distribution for use in Monte Carlo analysis. Taking these factors together allows for multi-variable, coupled analysis runs. To increase confidence in the tool, the results have been verified against two point-of-departure designs of the DUUS. The tool has also been verified against Apollo moon mission elements and other manned space systems. This paper will focus on trading key propulsion technologies including chemical, Nuclear Thermal Propulsion (NTP), and Solar Electric Propulsion (SEP). All of the key performance inputs and relationships will be presented and discussed in light of the various missions. For each mission there are several trajectory options and each will be discussed in terms of delta-v required and transit duration. Each propulsion system will be modeled, sized, and judged based on their applicability to the whole range of beyond LEO missions. Criteria for scoring will include the resulting dry mass of the stage, resulting propellant required, time to destination, and an assessment of key enabling technologies. In addition to the larger metrics, this paper will present the results of several coupled sensitivity studies. The ultimate goals of these tools and studies are to provide NASA with the most mass-, technology-, and cost-effective in-space stage for its future exploration missions.

Trajectory options for the DART mission

NASA Astrophysics Data System (ADS)

Atchison, Justin A.; Ozimek, Martin T.; Kantsiper, Brian L.; Cheng, Andrew F.

2016-06-01

This study presents interplanetary trajectory options for the Double Asteroid Redirection Test (DART) spacecraft to reach the near Earth object, Didymos binary system, during its 2022 Earth conjunction. DART represents a component of a joint NASA-ESA mission to study near Earth object kinetic impact deflection. The DART trajectory must satisfy mission objectives for arrival timing, geometry, and lighting while minimizing launch vehicle and spacecraft propellant requirements. Chemical propulsion trajectories are feasible from two candidate launch windows in late 2020 and 2021. The 2020 trajectories are highly perturbed by Earth's orbit, requiring post-launch deep space maneuvers to retarget the Didymos system. Within these windows, opportunities exist for flybys of additional near Earth objects: Orpheus in 2021 or 2007 YJ in 2022. A second impact attempt, in the event that the first impact is unsuccessful, can be added at the expense of a shorter launch window and increased (∼3x) spacecraft ΔV . However, the second impact arrival geometry has poor lighting, high Earth ranges, and would require additional degrees of freedom for solar panel and/or antenna gimbals. A low-thrust spacecraft configuration increases the trajectory flexibility. A solar electric propulsion spacecraft could be affordably launched as a secondary spacecraft in an Earth orbit and spiral out to target the requisite interplanetary departure condition. A sample solar electric trajectory was constructed from an Earth geostationary transfer using a representative 1.5 kW thruster. The trajectory requires 9 months to depart Earth's sphere of influence, after which its interplanetary trajectory includes a flyby of Orpheus and a second Didymos impact attempt. The solar electric spacecraft implementation would impose additional bus design constraints, including large solar arrays that could pose challenges for terminal guidance. On the basis of this study, there are many feasible options for DART to meet its mission design objectives and enable this unique kinetic impact experiment.

Cost-effective technology advancement directions for electric propulsion transportation systems in earth-orbital missions

NASA Technical Reports Server (NTRS)

Regetz, J. D., Jr.; Terwilliger, C. H.

1979-01-01

The directions that electric propulsion technology should take to meet the primary propulsion requirements for earth-orbital missions in the most cost effective manner are determined. The mission set requirements, state of the art electric propulsion technology and the baseline system characterized by it, adequacy of the baseline system to meet the mission set requirements, cost optimum electric propulsion system characteristics for the mission set, and sensitivities of mission costs and design points to system level electric propulsion parameters are discussed. The impact on overall costs than specific masses or costs of propulsion and power systems is evaluated.

Status of the NASA YF-12 Propulsion Research Program

NASA Technical Reports Server (NTRS)

Albers, J. A.

1976-01-01

The YF-12 research program was initiated to establish a technology base for the design of an efficient propulsion system for supersonic cruise aircraft. The major technology areas under investigation in this program are inlet design analysis, propulsion system steady-state performance, propulsion system dynamic performance, inlet and engine control systems, and airframe/propulsion system interactions. The objectives, technical approach, and status of the YF-12 propulsion program are discussed. Also discussed are the results obtained to date by the NASA Ames, Lewis, and Dryden research centers. The expected technical results and proposed future programs are also given. Propulsion system configurations are shown.

Controlling air pollution from passenger ferries: cost-effectiveness of seven technological options.

PubMed

Farrell, Alexander E; Corbett, James J; Winebrake, James J

2002-12-01

Continued interest in improving air quality in the United States along with renewed interest in the expansion of urban passenger ferry service has created concern about air pollution from ferry vessels. This paper presents a methodology for estimating the air pollution emissions from passenger ferries and the costs of emissions control strategies. The methodology is used to estimate the emissions and costs of retrofitting or re-powering ferries with seven technological options (combinations of propulsion and emission control systems) onto three vessels currently in service in San Francisco Bay. The technologies include improved engine design, cleaner fuels (including natural gas), and exhaust gas cleanup devices. The three vessels span a range of ages and technologies, from a 25-year-old monohull to a modern, high-speed catamaran built only four years ago. By looking at a range of technologies, vessel designs, and service conditions, a sense of the broader implications of controlling emissions from passenger ferries across a range of vessels and service profiles is provided. Tier 2-certified engines are the most cost-effective choice, but all options are cost-effective relative to other emission control strategies already in place in the transportation system.

Identification of propulsion systems

NASA Technical Reports Server (NTRS)

Merrill, Walter; Guo, Ten-Huei; Duyar, Ahmet

1991-01-01

This paper presents a tutorial on the use of model identification techniques for the identification of propulsion system models. These models are important for control design, simulation, parameter estimation, and fault detection. Propulsion system identification is defined in the context of the classical description of identification as a four step process that is unique because of special considerations of data and error sources. Propulsion system models are described along with the dependence of system operation on the environment. Propulsion system simulation approaches are discussed as well as approaches to propulsion system identification with examples for both air breathing and rocket systems.

Comparison of Mars Aircraft Propulsion Systems

NASA Technical Reports Server (NTRS)

Colozza, Anthony J.

2003-01-01

The propulsion system is a critical aspect of the performance and feasibility of a Mars aircraft. Propulsion system mass and performance greatly influence the aircraft s design and mission capabilities. Various propulsion systems were analyzed to estimate the system mass necessary for producing 35N of thrust within the Mars environment. Three main categories of propulsion systems were considered: electric systems, combustion engine systems and rocket systems. Also, the system masses were compared for mission durations of 1, 2, and 4 h.

Comparing Trash Disposal and Reuse Options for Deep Space Gateway and Mars Missions

NASA Technical Reports Server (NTRS)

Ewert, Michael; Broyan, James; Goodliff, Kandyce; Clowdsley, Martha; Singleterry, Robert

2017-01-01

Taking out the trash at NASA's newly proposed Deep Space Gateway (DSG) will not be a trivial task. While not the most important aspect of planning this cislunar outpost, there are several options that should be carefully considered since they may affect the crew as well as mission mass and volume. This study extends an earlier one, which focused on waste disposal options for a Mars Transit Vehicle. In that study, gasifying and venting trash along the way was found to noticeably reduce propellant needs and launch mass, whereas keeping processed trash on board in the form of radiation shielding tiles would significantly lower the crew's radiation dose during a solar particle event. Another favorable strategy was packing trash in a used logistics module for disposal. Since the DSG does not need much propulsion to maintain its orbit and Orion will be present with its own radiation storm shelter at the Gateway, the driving factors of the waste disposal trade study are different than for the Mars mission. Besides reviewing the propulsion and radiation shielding factors, potential drivers such as mass, power, volume, crew time, and human factors (e.g. smell) were studied. Disposal options for DSG include jettison of a used logistics module containing waste after every human stay, jettison of the same logistics module after several missions once it is full, regular disposal of trash via an airlock, or gasifying waste products for easier disposal or reuse. Conversely, a heat melt compactor device could be used to remove water and stabilize trash into tiles which could be more compactly stored on board and used as radiation shielding. Equivalent system mass analysis is used to tally the benefits and costs (mass, volume, power, crew time) of each case on an equivalent mass basis. Other more subjective factors are also discussed. Recommendations are made for DSG and Mars mission waste disposal.

Near Earth Asteroid Human Mission Possibilities Using Nuclear Thermal Rocket (NTR) Propulsion

NASA Technical Reports Server (NTRS)

Borowski, Stanley; McCurdy, David R.; Packard, Thomas W.

2012-01-01

The NTR is a proven technology that generates high thrust and has a specific impulse (Isp (is) approximately 900 s) twice that of today's best chemical rockets. During the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs, twenty rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability - all the requirements needed for a human mission to Mars. Ceramic metal fuel was also evaluated as a backup option. In NASA's recent Mars Design reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program - the 25 klbf 'Pewee' engine is sufficient for a human Mars mission when used in a clustered engine configuration. The 'Copernicus crewed NTR Mars transfer vehicle design developed for DRA 5.0 has significant capability that can enable reusable '1-year' round trip human missions to candidate near Earth asteroids (NEAs) like 1991 JW in 2027, or 2000 SG344 and Apophis in 2028. A robotic precursor mission to 2000 SG344 in late 2023 could provide an attractive Flight Technology Demonstration of a small NTR engine that is scalable to the 25 klbf-class engine used for human missions 5 years later. In addition to the detailed scientific data gathered from on-site inspection, human NEA missions would also provide a valuable 'check out' function for key elements of the NTR transfer vehicle (its propulsion module, TransHab and life support systems, etc.) in a 'deep space' environment prior to undertaking the longer duration Mars orbital and landing missions that would follow. The initial mass in low Earth orbit required for a mission to Apophis is approximately 323 t consisting of the NTR propulsion module ((is) approximately 138 t), the integrated saddle truss and LH2 drop tank assembly ((is) approximately 123 t), and the 6-crew payload element ((is) approximately 62 t). The later includes a multi-mission Space Excursion Vehicle (MMSEV) used for close-up examination and sample gathering. The total burn time and required restarts on the three 25 klbf 'Pewee-class' engines operating at Isp (is) approximately 906 s, are approximately 76.2 minutes and 4, respectively, well below the 2 hours and 27 restarts demonstrated on the NERVA eXperimental Engine, the NRX-XE. The paper examines the benefits, requirements and characteristics of using NTP for the above NEA missions. The impacts on vehicle design of HLV payload volume and lift capability, crew size, and reusability are also quantified.

Advanced NSTS propulsion system verification study

NASA Technical Reports Server (NTRS)

Wood, Charles

1989-01-01

The merits of propulsion system development testing are discussed. The existing data base of technical reports and specialists is utilized in this investigation. The study encompassed a review of all available test reports of propulsion system development testing for the Saturn stages, the Titan stages, and the Space Shuttle main propulsion system. The knowledge on propulsion system development and system testing available from specialists and managers was also 'tapped' for inclusion.

INSPACE CHEMICAL PROPULSION SYSTEMS AT NASA's MARSHALL SPACE FLIGHT CENTER: HERITAGE AND CAPABILITIES

NASA Technical Reports Server (NTRS)

McRight, P. S.; Sheehy, J. A.; Blevins, J. A.

2005-01-01

NASA s Marshall Space Flight Center (MSFC) is well known for its contributions to large ascent propulsion systems such as the Saturn V rocket and the Space Shuttle external tank, solid rocket boosters, and main engines. This paper highlights a lesser known but very rich side of MSFC-its heritage in the development of in-space chemical propulsion systems and its current capabilities for spacecraft propulsion system development and chemical propulsion research. The historical narrative describes the flight development activities associated with upper stage main propulsion systems such as the Saturn S-IVB as well as orbital maneuvering and reaction control systems such as the S-IVB auxiliary propulsion system, the Skylab thruster attitude control system, and many more recent activities such as Chandra, the Demonstration of Automated Rendezvous Technology (DART), X-37, the X-38 de-orbit propulsion system, the Interim Control Module, the US Propulsion Module, and multiple technology development activities. This paper also highlights MSFC s advanced chemical propulsion research capabilities, including an overview of the center s Propulsion Systems Department and ongoing activities. The authors highlight near-term and long-term technology challenges to which MSFC research and system development competencies are relevant. This paper concludes by assessing the value of the full range of aforementioned activities, strengths, and capabilities in light of NASA s exploration missions.

X-37 Storable Propulsion System Design and Operations

NASA Technical Reports Server (NTRS)

Rodriguez, Henry; Popp, Chris; Rehagen, Ronald J.

2005-01-01

In a response to NASA's X-37 TA-10 Cycle-1 contract, Boeing assessed nitrogen tetroxide (N2O4) and monomethyl hydrazine (MMH) Storable Propellant Propulsion Systems to select a low risk X-37 propulsion development approach. Space Shuttle lessons learned, planetary spacecraft, and Boeing Satellite HS-601 systems were reviewed to arrive at a low risk and reliable storable propulsion system. This paper describes the requirements, trade studies, design solutions, flight and ground operational issues which drove X-37 toward the selection of a storable propulsion system. The design of storable propulsion systems offers the leveraging of hardware experience that can accelerate progress toward critical design. It also involves the experience gained from launching systems using MMH and N2O4 propellants. Leveraging of previously flight-qualified hardware may offer economic benefits and may reduce risk in cost and schedule. This paper summarizes recommendations based on experience gained from Space Shuttle and similar propulsion systems utilizing MMH and N2O4 propellants. System design insights gained from flying storable propulsion are presented and addressed in the context of the design approach of the X-37 propulsion system.

X-37 Storable Propulsion System Design and Operations

NASA Technical Reports Server (NTRS)

Rodriguez, Henry; Popp, Chris; Rehegan, Ronald J.

2006-01-01

In a response to NASA's X-37 TA-10 Cycle-1 contract, Boeing assessed nitrogen tetroxide (N2O4) and monomethyl hydrazine (MMH) Storable Propellant Propulsion Systems to select a low risk X-37 propulsion development approach. Space Shuttle lessons learned, planetary spacecraft, and Boeing Satellite HS-601 systems were reviewed to arrive at a low risk and reliable storable propulsion system. This paper describes the requirements, trade studies, design solutions, flight and ground operational issues which drove X-37 toward the selection of a storable propulsion system. The design of storable propulsion systems offers the leveraging of hardware experience that can accelerate progress toward critical design. It also involves the experience gained from launching systems using MMH and N2O4 propellants. Leveraging of previously flight-qualified hardware may offer economic benefits and may reduce risk in cost and schedule. This paper summarizes recommendations based on experience gained from Space Shuttle and similar propulsion systems utilizing MMH and N2O4 propellants. System design insights gained from flying storable propulsion are presented and addressed in the context of the design approach of the X-37 propulsion system.

Fusion energy for space missions in the 21st Century

NASA Technical Reports Server (NTRS)

Schulze, Norman R.

1991-01-01

Future space missions were hypothesized and analyzed and the energy source for their accomplishment investigated. The mission included manned Mars, scientific outposts to and robotic sample return missions from the outer planets and asteroids, as well as fly-by and rendezvous mission with the Oort Cloud and the nearest star, Alpha Centauri. Space system parametric requirements and operational features were established. The energy means for accomplishing the High Energy Space Mission were investigated. Potential energy options which could provide the propulsion and electric power system and operational requirements were reviewed and evaluated. Fusion energy was considered to be the preferred option and was analyzed in depth. Candidate fusion fuels were evaluated based upon the energy output and neutron flux. Reactors exhibiting a highly efficient use of magnetic fields for space use while at the same time offering efficient coupling to an exhaust propellant or to a direct energy convertor for efficient electrical production were examined. Near term approaches were identified.

Multi-objective decision-making under uncertainty: Fuzzy logic methods

NASA Technical Reports Server (NTRS)

Hardy, Terry L.

1994-01-01

Selecting the best option among alternatives is often a difficult process. This process becomes even more difficult when the evaluation criteria are vague or qualitative, and when the objectives vary in importance and scope. Fuzzy logic allows for quantitative representation of vague or fuzzy objectives, and therefore is well-suited for multi-objective decision-making. This paper presents methods employing fuzzy logic concepts to assist in the decision-making process. In addition, this paper describes software developed at NASA Lewis Research Center for assisting in the decision-making process. Two diverse examples are used to illustrate the use of fuzzy logic in choosing an alternative among many options and objectives. One example is the selection of a lunar lander ascent propulsion system, and the other example is the selection of an aeration system for improving the water quality of the Cuyahoga River in Cleveland, Ohio. The fuzzy logic techniques provided here are powerful tools which complement existing approaches, and therefore should be considered in future decision-making activities.

Radioisotope Thermoelectric Generator Options for Pluto Fast Flyby Mission

NASA Astrophysics Data System (ADS)

Schock, Alfred

1994-07-01

A small spacecraft design for the Pluto Fast Flyby (PFF) mission is under study by the Jet Propulsion Laboratory (PL) for the National Aeronautics and Space Administration (NASA), for a possible launch as early as 1998. JPL's 1992 baseline design calls for a power source able to furnish an energy output of 3963 kWh and a power output of 69 Watts(e) at the end of the 9.2-year mission. Satisfying those demands is made difficult because NASA management has set a goal of reducing the spacecraft mass from a baseline value of 166 kg to ~110 kg, which implies a mass goal of less than 10 kg for the power source. To support the ongoing NASA/JPL studies, the Department of Energy's Office of Special Applications (DOE/OSA) commissioned Fairchild Space to prepare and analyze conceptual designs of radioisotope power systems for the PFF mission. Thus far, a total of eight options employing essentially the same radioisotope heat source modules were designed and subjected to thermal, electrical, structural, and mass analyses by Fairchild. Five of these - employing thermoelectric converters - are described in the present paper, and three - employing free-piston Stirling converters - are described in the companion paper presented next. The system masses of the thermoelectric options ranged from 19.3 kg to 10.2 kg. In general, the options requiring least development are the heaviest, and the lighter options require more development with greater programmatic risk.

Advanced space power and propulsion based on lasers

NASA Astrophysics Data System (ADS)

Roth, M.; Logan, B. G.

2015-10-01

One of the key components for future space exploration, manned or unmanned, is the availability of propulsion systems beyond the state of the art. The rapid development in conventional propulsion systems since the middle of the 20th century has already reached the limits of chemical propulsion technology. To enhance mission radius, shorten the transit time and also extend the lifetime of a spacecraft more efficient, but still powerful propulsion system must be developed. Apart from the propulsion system a major weight contribution arises from the required energy source. Envisioning rapid development of future high average power laser systems and especially the ICAN project we review the prospect of advanced space propulsion based on laser systems.

NASA/USRA advanced design program, 1990 - 1991

NASA Technical Reports Server (NTRS)

1991-01-01

Twenty-four UCLA students, in groups of four or five, participated in a mission design for a manned expedition to Mars that was based on the concept of mid-course refueling from electrically propelled tankers launched ahead of the manned mission. The study was conducted during the 1991 spring term. Some of the student groups opted for non-nuclear propulsion of the manned ship, based on LOX and LH2; others opted for one based on nuclear thermal propulsion. By way of example, the first option is presented here.

Advanced Development of a Compact 5-15 lbf Lox/Methane Thruster for an Integrated Reaction Control and Main Engine Propulsion System

NASA Technical Reports Server (NTRS)

Hurlbert, Eric A.; McManamen, John Patrick; Sooknanen, Josh; Studak, Joseph W.

2011-01-01

This paper describes the advanced development and testing of a compact 5 to 15 lbf LOX/LCH4 thruster for a pressure-fed integrated main engine and RCS propulsion system to be used on a spacecraft "vertical" test bed (VTB). The ability of the RCS thruster and the main engine to operate off the same propellant supply in zero-g reduces mass and improves mission flexibility. This compact RCS engine incorporates several features to dramatically reduce mass and parts count, to ease manufacturing, and to maintain acceptable performance given that specific impulse (Isp) is not the driver. For example, radial injection holes placed on the chamber body for easier drilling, and high temperature Haynes 230 were selected for the chamber over other more expensive options. The valve inlets are rotatable before welding allowing different orientations for vehicle integration. In addition, the engine design effort selected a coil-on-plug ignition system which integrates a relay and coil with the plug electrode, and moves some exciter electronics to avionics driver board. The engine injector design has small dribble volumes to target minimum pulse widths of 20 msec. and an efficient minimum impulse bit of less than 0.05 lbf-sec. The propellants, oxygen and methane, were chosen because together they are a non-toxic, Mars-forward, high density, space storable, and high performance propellant combination that is capable of pressure-fed and pump-fed configurations and integration with life support and power subsystems. This paper will present the results of the advanced development testing to date of the RCS thruster and the integration with a vehicle propulsion system.

Lunar surface base propulsion system study, volume 1

NASA Technical Reports Server (NTRS)

1987-01-01

The efficiency, capability, and evolution of a lunar base will be largely dependent on the transportation system that supports it. Beyond Space Station in low Earth orbit (LEO), a Lunar-derived propellant supply could provide the most important resource for the transportation infrastructure. The key to an efficient Lunar base propulsion system is the degree of Lunar self-sufficiency (from Earth supply) and reasonable propulsion system performance. Lunar surface propellant production requirements must be accounted in the measurement of efficiency of the entire space transportation system. Of all chemical propellant/propulsion systems considered, hydrogen/oxygen (H/O) OTVs appear most desirable, while both H/O and aluminum/oxygen propulsion systems may be considered for the lander. Aluminized-hydrogen/oxygen and Silane/oxygen propulsion systems are also promising candidates. Lunar propellant availability and processing techniques, chemical propulsion/vehicle design characteristics, and the associated performance of the total transportation infrastructure are reviewed, conceptual propulsion system designs and vehicle/basing concepts, and technology requirements are assessed in context of a Lunar Base mission scenario.

Next Generation Heavy-Lift Launch Vehicle: Large Diameter, Hydrocarbon-Fueled Concepts

NASA Technical Reports Server (NTRS)

Holliday, Jon; Monk, Timothy; Adams, Charles; Campbell, Ricky

2012-01-01

With the passage of the 2010 NASA Authorization Act, NASA was directed to begin the development of the Space Launch System (SLS) as a follow-on to the Space Shuttle Program. The SLS is envisioned as a heavy lift launch vehicle that will provide the foundation for future large-scale, beyond low Earth orbit (LEO) missions. Supporting the Mission Concept Review (MCR) milestone, several teams were formed to conduct an initial Requirements Analysis Cycle (RAC). These teams identified several vehicle concept candidates capable of meeting the preliminary system requirements. One such team, dubbed RAC Team 2, was tasked with identifying launch vehicles that are based on large stage diameters (up to the Saturn V S-IC and S-II stage diameters of 33 ft) and utilize high-thrust liquid oxygen (LOX)/RP engines as a First Stage propulsion system. While the trade space for this class of LOX/RP vehicles is relatively large, recent NASA activities (namely the Heavy Lift Launch Vehicle Study in late 2009 and the Heavy Lift Propulsion Technology Study of 2010) examined specific families within this trade space. Although the findings from these studies were incorporated in the Team 2 activity, additional branches of the trade space were examined and alternative approaches to vehicle development were considered. Furthermore, Team 2 set out to define a highly functional, flexible, and cost-effective launch vehicle concept. Utilizing this approach, a versatile two-stage launch vehicle concept was chosen as a preferred option. The preferred vehicle option has the capability to fly in several different configurations (e.g. engine arrangements) that gives this concept an inherent operational flexibility which allows the vehicle to meet a wide range of performance requirements without the need for costly block upgrades. Even still, this concept preserves the option for evolvability should the need arise in future mission scenarios. The foundation of this conceptual design is a focus on low cost and effectiveness rather than efficiency or cutting-edge technology. This paper details the approach and process, as well as the trade space analysis, leading to the preferred vehicle concept.

Deriving the Generalized Power and Efficiency Equations for Jet Propulsion Systems

NASA Astrophysics Data System (ADS)

Lee, Hsing-Juin; Chang, Chih-Luong

The kinetic power and efficiency equations for general jet propulsion systems are classically given in a much cursory, incomplete, and ununified format. This situation prohibits the propulsion designer from seeing the panorama of interrelated propulsion parameters and effects. And in some cases, it may lead to an energy-inefficient propulsion system design, or induce significant offset in propulsion performance as demonstrated in this study. Thus, herein we attempt to clarify some related concepts and to rigorously derive the associated generalized equations with a complete spectrum of physical parameters to be manipulated in quest of better performance. By a highly efficient interweaved transport scheme, we have derived the following equations for general jet propulsion systems: i.e., generalized total kinetic power, generalized kinetic power delivered to the jet propulsion system, generalized thrust power, generalized available propulsion power, and relevant generalized propulsive, thermal, and overall efficiency equations. Further, the variants of these equations under special conditions are also considered. For taking advantage of the above propulsion theories, we also illustrate some novel propulsion strategies in the final discussion, such as the dive-before-climb launch of rocket from highland mountain on eastbound rail, with perhaps minisatellites as the payloads.

Mars exploration advances: Missions to Mars - Mars base

NASA Technical Reports Server (NTRS)

Dejarnette, Fred R.; Mckay, Christopher P.

1992-01-01

An overview is presented of Mars missions and related planning with attention given to four mission architectures in the light of significant limitations. Planned unpiloted missions are discussed including the Mars Orbital Mapping Mission, the Mars Rover Sample Return, the Mars Aeronomy Orbiter, and the Mars Environmental Survey. General features relevant to the missions are mentioned including launch opportunities, manned-mission phases, and propulsion options. The four mission architectures are set forth and are made up of: (1) the Mars-exploration infrastructures; (2) science emphasis for the moon and Mars; (3) the moon to stay and Mars exploration; and (4) space resource utilization. The possibility of robotic missions to the moon and Mars is touched upon and are concluded to be possible by the end of the century. The ramifications of a Mars base are discussed with specific reference to habitability and base activities, and the human missions are shown to require a heavy-lift launcher and either chemical/aerobrake or nuclear-thermal propulsion system.

Space Debris Removal Using Multi-Mission Modular Spacecraft

NASA Astrophysics Data System (ADS)

Savioli, L.; Francesconi, A.; Maggi, F.; Olivieri, L.; Lorenzini, E.; Pardini, C.

2013-08-01

The study and development of ADR missions in LEO have become an issue of topical interest to the attention of the space community since the future space flight activities could be threatened by collisional cascade events. This paper presents the analysis of an ADR mission scenario where modular remover kits are employed to de-orbit some selected debris in SSO, while a distinct space tug performs the orbital transfers and rendezvous manoeuvres, and installs the remover kits on the client debris. Electro-dynamic tether and electric propulsion are considered as de-orbiting alternatives, while chemical propulsion is employed for the space tug. The total remover mass and de-orbiting time are identified as key parameters to compare the performances of the two de-orbiting options, while an optimization of the ΔV required to move between five selected objects is performed for a preliminary design at system level of the space tug. Final controlled re-entry is also considered and performed by means of a hybrid engine.

The Iodine Satellite (iSat) Project Development Towards Critical Design Review (CDR)

NASA Technical Reports Server (NTRS)

Dankanich, John W.; Selby, Michael; Polzin, Kurt A.; Kamhawi, Hani; Hickman, Tyler; Byrne, Larry

2016-01-01

Despite the prevalence of Small Satellites in recent years, the systems flown to date have very limited propulsion capability. SmallSats are typically secondary payloads and have significant constraints for volume, mass, and power in addition to limitations on the use of hazardous propellants or stored energy (i.e. high pressure vessels). These constraints limit the options for SmallSat maneuverability. NASA's Space Technology Mission Directorate approved the iodine Satellite flight project for a rapid demonstration of iodine Hall thruster technology in a 12U configuration under the Small Spacecraft Technology Program. The project formally began in FY15 as a partnership between NASA MSFC, NASA GRC, and Busek Co, Inc., with the Air Force supporting the propulsion technology maturation. The team is in final preparation of the Critical Design Review prior to initiating the fabrication and integration phase of the project. The iSat project is on schedule for a launch opportunity in November 2017.

Development of the electric vehicle analyzer

NASA Astrophysics Data System (ADS)

Dickey, Michael R.; Klucz, Raymond S.; Ennix, Kimberly A.; Matuszak, Leo M.

1990-06-01

The increasing technological maturity of high power (greater than 20 kW) electric propulsion devices has led to renewed interest in their use as a means of efficiently transferring payloads between earth orbits. Several systems and architecture studies have identified the potential cost benefits of high performance Electric Orbital Transfer Vehicles (EOTVs). These studies led to the initiation of the Electric Insertion Transfer Experiment (ELITE) in 1988. Managed by the Astronautics Laboratory, ELITE is a flight experiment designed to sufficiently demonstrate key technologies and options to pave the way for the full-scale development of an operational EOTV. An important consideration in the development of the ELITE program is the capability of available analytical tools to simulate the orbital mechanics of a low thrust, electric propulsion transfer vehicle. These tools are necessary not only for ELITE mission planning exercises but also for continued, efficient, accurate evaluation of DoD space transportation architectures which include EOTVs. This paper presents such a tool: the Electric Vehicle Analyzer (EVA).

Space Propulsion Technology Program Overview

NASA Technical Reports Server (NTRS)

Escher, William J. D.

1991-01-01

The topics presented are covered in viewgraph form. Focused program elements are: (1) transportation systems, which include earth-to-orbit propulsion, commercial vehicle propulsion, auxiliary propulsion, advanced cryogenic engines, cryogenic fluid systems, nuclear thermal propulsion, and nuclear electric propulsion; (2) space platforms, which include spacecraft on-board propulsion, and station keeping propulsion; and (3) technology flight experiments, which include cryogenic orbital N2 experiment (CONE), SEPS flight experiment, and cryogenic orbital H2 experiment (COHE).

System Analysis and Performance Benefits of an Optimized Rotorcraft Propulsion System

NASA Technical Reports Server (NTRS)

Bruckner, Robert J.

2007-01-01

The propulsion system of rotorcraft vehicles is the most critical system to the vehicle in terms of safety and performance. The propulsion system must provide both vertical lift and forward flight propulsion during the entire mission. Whereas propulsion is a critical element for all flight vehicles, it is particularly critical for rotorcraft due to their limited safe, un-powered landing capability. This unparalleled reliability requirement has led rotorcraft power plants down a certain evolutionary path in which the system looks and performs quite similarly to those of the 1960 s. By and large the advancements in rotorcraft propulsion have come in terms of safety and reliability and not in terms of performance. The concept of the optimized propulsion system is a means by which both reliability and performance can be improved for rotorcraft vehicles. The optimized rotorcraft propulsion system which couples an oil-free turboshaft engine to a highly loaded gearbox that provides axial load support for the power turbine can be designed with current laboratory proven technology. Such a system can provide up to 60% weight reduction of the propulsion system of rotorcraft vehicles. Several technical challenges are apparent at the conceptual design level and should be addressed with current research.

Options for human ``return to the moon'' using tomorrow's SSTO, ISRU, and LOX-augmented NTR technologies

NASA Astrophysics Data System (ADS)

Borowski, Stanley K.

1996-03-01

The feasibility of conducting human missions to the Moon is examined assuming the use of three ``high leverage'' technologies: (1) a single-stage-to-orbit (SSTO) launch vehicle, (2) ``in-situ'' resource utilization (ISRU)—specifically ``lunar-derived'' liquid oxygen (LUNOX), and (3) LOX-augmented nuclear thermal rocket (LANTR) propulsion. Lunar transportation system elements consisting of a LANTR-powered lunar transfer vehicle (LTV) and a chemical propulsion lunar landing/Earth return vehicle (LERV) are configured to fit within the ``compact'' dimensions of the SSTO cargo bay (diameter: 4.6 m/length: 9.0 m) while satisfying an initial mass in low Earth orbit (IMLEO) limit of ˜60 t (3 SSTO launches). Using ˜8 t of LUNOX to ``reoxidize'' the LERV for a ``direct return'' flight to Earth reduces its size and mass allowing delivery to LEO on a single 20 t SSTO launch. Similarly, the LANTR engine's ability to operate at any oxygen/hydrogen mixture ratio from 0 to 7 with high specific impulse (˜940 to 515 s) is exploited to reduce hydrogen tank volume, thereby improving packaging of the LANTR LTV's ``propulsion'' and ``propellant modules''. Expendable and reusable, piloted and cargo missions and vehicle designs are presented along with estimates of LUNOX production required to support the different mission modes.

Investigation of a Tricarbide Grooved Ring Fuel Element for a Nuclear Thermal Rocket

NASA Technical Reports Server (NTRS)

Taylor, Brian D.; Emrich, Bill; Tucker, Dennis; Barnes, Marvin; Donders, Nicolas; Benensky, Kelsa

2017-01-01

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nuclear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the temperature limitations of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment.

Apollo Command and Service Module Propulsion Systems Overview

NASA Technical Reports Server (NTRS)

Interbartolo, Michael A.

2009-01-01

An overview of the Apollo Command and Service Module (CSM) propulsion systems is provided. The systems for CSM propulsion and control are defined, the times during the mission when each system is used are listed, and, the basic components and operation of the service propulsion system, SM reaction control system and CM reaction control system are described.

A real time Pegasus propulsion system model for VSTOL piloted simulation evaluation

NASA Technical Reports Server (NTRS)

Mihaloew, J. R.; Roth, S. P.; Creekmore, R.

1981-01-01

A real time propulsion system modeling technique suitable for use in man-in-the-loop simulator studies was developd. This technique provides the system accuracy, stability, and transient response required for integrated aircraft and propulsion control system studies. A Pegasus-Harrier propulsion system was selected as a baseline for developing mathematical modeling and simulation techniques for VSTOL. Initially, static and dynamic propulsion system characteristics were modeled in detail to form a nonlinear aerothermodynamic digital computer simulation of a Pegasus engine. From this high fidelity simulation, a real time propulsion model was formulated by applying a piece-wise linear state variable methodology. A hydromechanical and water injection control system was also simulated. The real time dynamic model includes the detail and flexibility required for the evaluation of critical control parameters and propulsion component limits over a limited flight envelope. The model was programmed for interfacing with a Harrier aircraft simulation. Typical propulsion system simulation results are presented.

NASA's Evolutionary Xenon Thruster: The NEXT Ion Propulsion System for Solar System Exploration

NASA Technical Reports Server (NTRS)

Pencil, Eric J.; Benson, Scott W.

2008-01-01

This viewgraph presentation reviews NASA s Evolutionary Xenon Thruster (NEXT) Ion Propulsion system. The NEXT project is developing a solar electric ion propulsion system. The NEXT project is advancing the capability of ion propulsion to meet NASA robotic science mission needs. The NEXT system is planned to significantly improve performance over the state of the art electric propulsion systems, such as NASA Solar Electric Propulsion Technology Application Readiness (NSTAR). The status of NEXT development is reviewed, including information on the NEXT Thruster, the power processing unit, the propellant management system (PMS), the digital control interface unit, and the gimbal. Block diagrams NEXT system are presented. Also a review of the lessons learned from the Dawn and NSTAR systems is provided. In summary the NEXT project activities through 2007 have brought next-generation ion propulsion technology to a sufficient maturity level.

Raising Nuclear Thermal Propulsion (NTP) Technology Readiness Above 3

NASA Technical Reports Server (NTRS)

Gerrish, Harold P., Jr.

2014-01-01

NTP development is currently supported by the NASA program office "Advanced Exploration Systems". The concept is a main propulsion option being considered for human missions to Mars in the 2030's. Major NTP development took place in the 1960's and 1970's under the Rover/NERVA program. The technology had matured to TRL 6 and was preparing to go to TRL 7 with a prototype flight engine before the program was cancelled. Over the last 40 years, a variety of continuations started, but only lasted a few years each. The Rover/NERVA infrastructure is almost all gone. The only remains are a few pieces of hardware, final reports and a few who worked the Rover/NERVA. Two types of nuclear fuel are being investigated to meet the current engine design specific impulse of 900 seconds compared to approximately 850 seconds demonstrated during Rover/NERVA. One is a continuation of composite fuel with new coatings to better control mid-band corrosion. The other type is a CERMET fuel made of Tungsten and UO2. Both fuels are being made from Rover/NERVA lessons learned, but with slightly different recipes to increase fuel endurance at higher operating temperatures. The technology readiness level (TRL) of these current modified reactor fuels is approximately TRL 3. To keep the development cost low and help mature the TRL level past 4 quickly, a few special non-nuclear test facilities have been made to test surrogate fuel, with depleted uranium, as coupons and full length elements. Both facilities utilize inductive heating and are licensed to handle depleted uranium. TRL 5 requires exposing the fuel to a nuclear environment and TRL 6 requires a prototype ground or flight engine system test. Currently, three different NTP ground test facility options are being investigated: exhaust scrubber, bore hole, and total exhaust containment. In parallel, a prototype flight demonstration test is also being studied. The first human mission to Mars in the 2030's is currently 2033. For an advanced propulsion concept to be seriously considered for use, the engine development plans need to show it is feasible and affordable to reach TRL 8 by 2027 and can be qualified for human mission use.

The MPD thruster program at JPL

NASA Technical Reports Server (NTRS)

Barnett, John; Goodfellow, Keith; Polk, James; Pivirotto, Thomas

1991-01-01

The main topics covered include: (1) the Space Exploration Initiative (SEI) context; (2) critical issues of MPD Thruster design; and (3) the Magnetoplasmadynamic (MPD) Thruster Program at JPL. Under the section on the SEI context the nuclear electric propulsion system and some electric thruster options are addressed. The critical issues of MPD Thruster development deal with the requirements, status, and approach taken. The following areas are covered with respect to the MPD Thruster Program at JPL: (1) the radiation-cooled MPD thruster; (2) the High-Current Cathode Test Facility; (3) thruster component thermal modeling; and (4) alkali metal propellant studies.

Software For Calibration Of Polarimetric SAR Data

NASA Technical Reports Server (NTRS)

Van Zyl, Jakob; Zebker, Howard; Freeman, Anthony; Holt, John; Dubois, Pascale; Chapman, Bruce

1994-01-01

POLCAL (Polarimetric Radar Calibration) software tool intended to assist in calibration of synthetic-aperture radar (SAR) systems. In particular, calibrates Stokes-matrix-format data produced as standard product by NASA/Jet Propulsion Laboratory (JPL) airborne imaging synthetic aperture radar (AIRSAR). Version 4.0 of POLCAL is upgrade of version 2.0. New options include automatic absolute calibration of 89/90 data, distributed-target analysis, calibration of nearby scenes with corner reflectors, altitude or roll-angle corrections, and calibration of errors introduced by known topography. Reduces crosstalk and corrects phase calibration without use of ground calibration equipment. Written in FORTRAN 77.

46 CFR 111.33-11 - Propulsion systems.

Code of Federal Regulations, 2010 CFR

2010-10-01

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46 CFR 111.33-11 - Propulsion systems.

Code of Federal Regulations, 2013 CFR

2013-10-01

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46 CFR 111.33-11 - Propulsion systems.

Code of Federal Regulations, 2014 CFR

2014-10-01

... 46 Shipping 4 2014-10-01 2014-10-01 false Propulsion systems. 111.33-11 Section 111.33-11 Shipping... REQUIREMENTS Power Semiconductor Rectifier Systems § 111.33-11 Propulsion systems. Each power semiconductor rectifier system in a propulsion system must meet sections 4-8-5/5.17.9 and 4-8-5/5.17.10 of ABS Steel...

46 CFR 111.33-11 - Propulsion systems.

Code of Federal Regulations, 2012 CFR

2012-10-01

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46 CFR 111.33-11 - Propulsion systems.

Code of Federal Regulations, 2011 CFR

2011-10-01

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Propulsion System and Orbit Maneuver Integration in CubeSats: Trajectory Control Strategies Using Micro Ion Propulsion

NASA Technical Reports Server (NTRS)

Hudson, Jennifer; Martinez, Andres; Petro, Andrew

2015-01-01

The Propulsion System and Orbit Maneuver Integration in CubeSats project aims to solve the challenges of integrating a micro electric propulsion system on a CubeSat in order to perform orbital maneuvers and control attitude. This represents a fundamentally new capability for CubeSats, which typically do not contain propulsion systems and cannot maneuver far beyond their initial orbits.

An overview of the NASA Advanced Propulsion Concepts program

NASA Technical Reports Server (NTRS)

Curran, Francis M.; Bennett, Gary L.; Frisbee, Robert H.; Sercel, Joel C.; Lapointe, Michael R.

1992-01-01

NASA Advanced Propulsion Concepts (APC) program for the development of long-term space propulsion system schemes is managed by both NASA-Lewis and the JPL and is tasked with the identification and conceptual development of high-risk/high-payoff configurations. Both theoretical and experimental investigations have been undertaken in technology areas deemed essential to the implementation of candidate concepts. These APC candidates encompass very high energy density chemical propulsion systems, advanced electric propulsion systems, and an antiproton-catalyzed nuclear propulsion concept. A development status evaluation is presented for these systems.

Hazardous Environment Robotics

NASA Technical Reports Server (NTRS)

1996-01-01

Jet Propulsion Laboratory (JPL) developed video overlay calibration and demonstration techniques for ground-based telerobotics. Through a technology sharing agreement with JPL, Deneb Robotics added this as an option to its robotics software, TELEGRIP. The software is used for remotely operating robots in nuclear and hazardous environments in industries including automotive and medical. The option allows the operator to utilize video to calibrate 3-D computer models with the actual environment, and thus plan and optimize robot trajectories before the program is automatically generated.

Development of Laser Propulsion and Tracking System for Laser-Driven Micro-Airplane

NASA Astrophysics Data System (ADS)

Ishikawa, Hiroyasu; Kajiwara, Itsuro; Hoshino, Kentaro; Yabe, Takashi; Uchida, Shigeaki; Shimane, Yoshichika

2004-03-01

The purposes of this paper are to improve the control performance of the developed laser tracking system and to develop an integrated laser propulsion/tracking system for realizing a continuous flight and control of the micro-airplane. The laser propulsion is significantly effective to achieve the miniaturization and lightening of the micro-airplane. The laser-driven micro-airplane has been studied with a paper-craft airplane and YAG laser, resulting in a successful glide of the airplane. In the next stage of the laser-driven micro-airplane development, the laser tracking is expected as key technologies to achieve continuous propulsion. Furthermore, the laser propulsion system should be combined with the laser tracking system to supply continuous propulsion. Experiments are carried out to evaluate the performance of the developed laser tracking system and integrated laser propulsion/tracking system.

Status of Propulsion Technology Development Under the NASA In-space Propulsion Technology Program

NASA Technical Reports Server (NTRS)

Anderson, David; Kamhawi, Hani; Patterson, Mike; Dankanich, John; Pencil, Eric; Pinero, Luis

2014-01-01

Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing and delivering in-space propulsion technologies for NASA's Science Mission Directorate (SMD). These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems Mission Analysis. ISPT's propulsion technologies include: 1) the 0.6-7 kW NASA's Evolutionary Xenon Thruster (NEXT) gridded ion propulsion system; 2) a 0.3-3.9kW Hall-effect electric propulsion (HEP) system for low cost and sample return missions; 3) the Xenon Flow Control Module (XFCM); 4) ultra-lightweight propellant tank technologies (ULTT); and 5) propulsion technologies for a Mars Ascent Vehicle (MAV). The HEP system is composed of the High Voltage Hall Accelerator (HiVHAc) thruster, a power processing unit (PPU), and the XFCM. NEXT and the HiVHAc are throttle-able electric propulsion systems for planetary science missions. The XFCM and ULTT are two component technologies which being developed with nearer-term flight infusion in mind. Several of the ISPT technologies are related to sample return missions needs like: MAV propulsion and electric propulsion. And finally, one focus of the SystemsMission Analysis area is developing tools that aid the application or operation of these technologies on wide variety of mission concepts. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

A Brief Review of the Need for Robust Smart Wireless Sensor Systems for Future Propulsion Systems, Distributed Engine Controls, and Propulsion Health Management

NASA Technical Reports Server (NTRS)

Hunter, Gary W.; Behbahani, Alireza

2012-01-01

Smart Sensor Systems with wireless capability operational in high temperature, harsh environments are a significant component in enabling future propulsion systems to meet a range of increasingly demanding requirements. These propulsion systems must incorporate technology that will monitor engine component conditions, analyze the incoming data, and modify operating parameters to optimize propulsion system operations. This paper discusses the motivation towards the development of high temperature, smart wireless sensor systems that include sensors, electronics, wireless communication, and power. The challenges associated with the use of traditional wired sensor systems will be reviewed and potential advantages of Smart Sensor Systems will be discussed. A brief review of potential applications for wireless smart sensor networks and their potential impact on propulsion system operation, with emphasis on Distributed Engine Control and Propulsion Health Management, will be given. A specific example related to the development of high temperature Smart Sensor Systems based on silicon carbide electronics will be discussed. It is concluded that the development of a range of robust smart wireless sensor systems are a foundation for future development of intelligent propulsion systems with enhanced capabilities.

46 CFR 184.620 - Propulsion engine control systems.

Code of Federal Regulations, 2014 CFR

2014-10-01

... 46 Shipping 7 2014-10-01 2014-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

46 CFR 184.620 - Propulsion engine control systems.

Code of Federal Regulations, 2013 CFR

2013-10-01

... 46 Shipping 7 2013-10-01 2013-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

46 CFR 184.620 - Propulsion engine control systems.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 46 Shipping 7 2011-10-01 2011-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

46 CFR 184.620 - Propulsion engine control systems.

Code of Federal Regulations, 2012 CFR

2012-10-01

... 46 Shipping 7 2012-10-01 2012-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

46 CFR 184.620 - Propulsion engine control systems.

Code of Federal Regulations, 2010 CFR

2010-10-01

... 46 Shipping 7 2010-10-01 2010-10-01 false Propulsion engine control systems. 184.620 Section 184... Communications Systems § 184.620 Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of...

Fuzzy Logic Approaches to Multi-Objective Decision-Making in Aerospace Applications

NASA Technical Reports Server (NTRS)

Hardy, Terry L.

1994-01-01

Fuzzy logic allows for the quantitative representation of multi-objective decision-making problems which have vague or fuzzy objectives and parameters. As such, fuzzy logic approaches are well-suited to situations where alternatives must be assessed by using criteria that are subjective and of unequal importance. This paper presents an overview of fuzzy logic and provides sample applications from the aerospace industry. Applications include an evaluation of vendor proposals, an analysis of future space vehicle options, and the selection of a future space propulsion system. On the basis of the results provided in this study, fuzzy logic provides a unique perspective on the decision-making process, allowing the evaluator to assess the degree to which each option meets the evaluation criteria. Future decision-making should take full advantage of fuzzy logic methods to complement existing approaches in the selection of alternatives.

Design options for advanced manned launch systems

NASA Astrophysics Data System (ADS)

Freeman, Delma C.; Talay, Theodore A.; Stanley, Douglas O.; Lepsch, Roger A.; Wilhite, Alan W.

1995-03-01

Various concepts for advanced manned launch systems are examined for delivery missions to space station and polar orbit. Included are single-and two-stage winged systems with rocket and/or air-breathing propulsion systems. For near-term technologies, two-stage reusable rocket systems are favored over single-stage rocket or two-stage air-breathing/rocket systems. Advanced technologies enable viable single-stage-to-orbit (SSTO) concepts. Although two-stage rocket systems continue to be lighter in dry weight than SSTO vehicles, advantages in simpler operations may make SSTO vehicles more cost-effective over the life cycle. Generally, rocket systems maintain a dry-weight advantage over air-breathing systems at the advanced technology levels, but to a lesser degree than when near-term technologies are used. More detailed understanding of vehicle systems and associated ground and flight operations requirements and procedures is essential in determining quantitative discrimination between these latter concepts.

Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

Chojnacki, Kent T.

1992-01-01

The goal of the Rocket-Based Combined Cycle (RBCC) Propulsion Technology Workshop was to assess the RBCC propulsion system's viability for Earth-to-Orbit (ETO) transportation systems. This was accomplished by creating a forum (workshop) in which past work in the field of RBCC propulsion systems was reviewed, current technology status was evaluated, and future technology programs in the field of RBCC propulsion systems were postulated, discussed, and recommended.

Electric propulsion for lunar exploration and lunar base development

NASA Technical Reports Server (NTRS)

Palaszewski, Bryan

1992-01-01

Using electric propulsion to deliver materials to lunar orbit for the development and construction of a lunar base was investigated. Because the mass of the base and its life-cycle resupply mass are large, high specific impulse propulsion systems may significantly reduce the transportation system mass and cost. Three electric propulsion technologies (arcjet, ion, and magnetoplasmadynamic (MPD) propulsion) were compared with oxygen/hydrogen propulsion for a lunar base development scenario. Detailed estimates of the orbital transfer vehicles' (OTV's) masses and their propellant masses are presented. The fleet sizes for the chemical and electric propulsion systems are estimated. Ion and MPD propulsion systems enable significant launch mass savings over O2/H2 propulsion. Because of the longer trip time required for the low-thrust OTV's, more of them are required to perform the mission model. By offloading the lunar cargo from the manned O2/H2 OTV missions onto the electric propulsion OTV's, a significant reduction of the low Earth orbit (LEO) launch mass is possible over the 19-year base development period.

Integrated propulsion/energy transfer control systems for lift-fan V/STOL aircraft. [reduction of total propulsion system and control system installation requirements

NASA Technical Reports Server (NTRS)

Deckert, W. H.; Rolls, L. S.

1974-01-01

An integrated propulsion/control system for lift-fan transport aircraft is described. System behavior from full-scale experimental and piloted simulator investigations are reported. The lift-fan transport is a promising concept for short-to-medium haul civil transportation and for other missions. The lift-fan transport concept features high cruise airspeed, favorable ride qualities, small perceived noise footprints, high utilization, transportation system flexibility, and adaptability to VTOL, V/STOL, or STOL configurations. The lift-fan transport has high direct operating costs in comparison to conventional aircraft, primarily because of propulsion system and aircraft low-speed control system installation requirements. An integrated lift-fan propulsion system/aircraft low-speed control system that reduces total propulsion system and control system installation requirements is discussed.

Space propulsion systems. Present performance limits and application and development trends

NASA Technical Reports Server (NTRS)

Buehler, R. D.; Lo, R. E.

1981-01-01

Typical spaceflight programs and their propulsion requirements as a comparison for possible propulsion systems are summarized. Chemical propulsion systems, solar, nuclear, or even laser propelled rockets with electrical or direct thermal fuel acceleration, nonrockets with air breathing devices and solar cells are considered. The chemical launch vehicles have similar technical characteristics and transportation costs. A possible improvement of payload by using air breathing lower stages is discussed. The electrical energy supply installations which give performance limits of electrical propulsion and the electrostatic ion propulsion systems are described. The development possibilities of thermal, magnetic, and electrostatic rocket engines and the state of development of the nuclear thermal rocket and propulsion concepts are addressed.

Propulsion for CubeSats

NASA Astrophysics Data System (ADS)

Lemmer, Kristina

2017-05-01

At present, very few CubeSats have flown in space featuring propulsion systems. Of those that have, the literature is scattered, published in a variety of formats (conference proceedings, contractor websites, technical notes, and journal articles), and often not available for public release. This paper seeks to collect the relevant publically releasable information in one location. To date, only two missions have featured propulsion systems as part of the technology demonstration. The IMPACT mission from the Aerospace Corporation launched several electrospray thrusters from Massachusetts Institute of Technology, and BricSAT-P from the United States Naval Academy had four micro-Cathode Arc Thrusters from George Washington University. Other than these two missions, propulsion on CubeSats has been used only for attitude control and reaction wheel desaturation via cold gas propulsion systems. As the desired capability of CubeSats increases, and more complex missions are planned, propulsion is required to accomplish the science and engineering objectives. This survey includes propulsion systems that have been designed specifically for the CubeSat platform and systems that fit within CubeSat constraints but were developed for other platforms. Throughout the survey, discussion of flight heritage and results of the mission are included where publicly released information and data have been made available. Major categories of propulsion systems that are in this survey are solar sails, cold gas propulsion, electric propulsion, and chemical propulsion systems. Only systems that have been tested in a laboratory or with some flight history are included.

Evaluation of propfan propulsion applied to general aviation

NASA Technical Reports Server (NTRS)

Awker, R. W.

1986-01-01

Propfan propulsion on business aircraft was evaluated. Comparisons, in terms of cost and performance, were made between propfan propulsion systems and conventional turbofan propulsion systems on a typical business aircraft. In addition, configuration and cost sensitivity studies were conducted to further assess the potential of propfan propulsion.

Study of electrical and chemical propulsion systems for auxiliary propulsion of large space systems, volume 2

NASA Technical Reports Server (NTRS)

Smith, W. W.

1981-01-01

The five major tasks of the program are reported. Task 1 is a literature search followed by selection and definition of seven generic spacecraft classes. Task 2 covers the determination and description of important disturbance effects. Task 3 applies the disturbances to the generic spacecraft and adds maneuver and stationkeeping functions to define total auxiliary propulsion systems requirements for control. The important auxiliary propulsion system characteristics are identified and sensitivities to control functions and large space system characteristics determined. In Task 4, these sensitivities are quantified and the optimum auxiliary propulsion system characteristics determined. Task 5 compares the desired characteristics with those available for both electrical and chemical auxiliary propulsion systems to identify the directions technology advances should take.

Propulsion System Models for Rotorcraft Conceptual Design

NASA Technical Reports Server (NTRS)

Johnson, Wayne

2014-01-01

The conceptual design code NDARC (NASA Design and Analysis of Rotorcraft) was initially implemented to model conventional rotorcraft propulsion systems, consisting of turboshaft engines burning jet fuel, connected to one or more rotors through a mechanical transmission. The NDARC propulsion system representation has been extended to cover additional propulsion concepts, including electric motors and generators, rotor reaction drive, turbojet and turbofan engines, fuel cells and solar cells, batteries, and fuel (energy) used without weight change. The paper describes these propulsion system components, the architecture of their implementation in NDARC, and the form of the models for performance and weight. Requirements are defined for improved performance and weight models of the new propulsion system components. With these new propulsion models, NDARC can be used to develop environmentally-friendly rotorcraft designs.

NASA Electric Propulsion System Studies

NASA Technical Reports Server (NTRS)

Felder, James L.

2015-01-01

An overview of NASA efforts in the area of hybrid electric and turboelectric propulsion in large transport. This overview includes a list of reasons why we are looking at transmitting some or all of the propulsive power for the aircraft electrically, a list of the different types of hybrid-turbo electric propulsion systems, and the results of 4 aircraft studies that examined different types of hybrid-turbo electric propulsion systems.

System controls challenges of hypersonic combined-cycle engine powered vehicles

NASA Technical Reports Server (NTRS)

Morrison, Russell H.; Ianculescu, George D.

1992-01-01

Hypersonic aircraft with air-breathing engines have been described as the most complex and challenging air/space vehicle designs ever attempted. This is particularly true for aircraft designed to accelerate to orbital velocities. The propulsion system for the National Aerospace Plane will be an active factor in maintaining the aircraft on course. Typically addressed are the difficulties with the aerodynamic vehicle design and development, materials limitations and propulsion performance. The propulsion control system requires equal materials limitations and propulsion performance. The propulsion control system requires equal concern. Far more important than merely a subset of propulsion performance, the propulsion control system resides at the crossroads of trajectory optimization, engine static performance, and vehicle-engine configuration optimization. To date, solutions at these crossroads are multidisciplinary and generally lag behind the broader performance issues. Just how daunting these demands will be is suggested. A somewhat simplified treatment of the behavioral characteristics of hypersonic aircraft and the issues associated with their air-breathing propulsion control system design are presented.

Beamed energy propulsion

NASA Technical Reports Server (NTRS)

Shoji, James M.

1992-01-01

Beamed energy concepts offer an alternative for an advanced propulsion system. The use of a remote power source reduces the weight of the propulsion system in flight and this, combined with the high performance, provides significant payload gains. Within the context of this study's baseline scenario, two beamed energy propulsion concepts are potentially attractive: solar thermal propulsion and laser thermal propulsion. The conceived beamed energy propulsion devices generally provide low thrust (tens of pounds to hundreds of pounds); therefore, they are typically suggested for cargo transportation. For the baseline scenario, these propulsion system can provide propulsion between the following nodes: (1) low Earth orbit to geosynchronous Earth orbit; (2) low Earth orbit to low lunar orbit; (3) low lunar orbit to low Mars orbit--only solar thermal; and (4) lunar surface to low lunar orbit--only laser thermal.

Strategies for the sustained human exploration of Mars

NASA Astrophysics Data System (ADS)

Landau, Damon Frederick

A variety of mission scenarios are compared in this thesis to assess the strengths and weaknesses of options for Mars exploration. The mission design space is modeled along two dimensions: trajectory architectures and propulsion system technologies. Direct, semi-direct, stop-over, semi-cycler, and cycler architectures are examined, and electric propulsion, nuclear thermal rockets, methane and oxygen production on Mars, Mars water excavation, aerocapture, and reusable propulsion systems are included in the technology assessment. The mission sensitivity to crew size, vehicle masses, and crew travel time is also examined. The primary figure of merit for a mission scenario is the injected mass to low-Earth orbit (IMLEO), though technology readiness levels (TRL) are also included. Several elements in the architecture dimension are explored in more detail. The Earth-Mars semi-cycler architecture is introduced and five families of Earth-Mars semi-cycler trajectories are presented along with optimized itineraries. Optimized cycler trajectories are also presented. In addition to Earth-Mars semi-cycler and cycler trajectories, conjunction-class, free-return, Mars-Earth semi-cycler, and low-thrust trajectories are calculated. Design parameters for optimal DeltaV trajectories are provided over a range of flight times (from 120 to 270 days) and launch years (between 2009 and 2022). Unlike impulsive transfers, the mass-optimal low-thrust trajectory depends strongly on the thrust and specific impulse of the propulsion system. A low-thrust version of the rocket equation is provided where the initial mass or thrust may be minimized by varying the initial acceleration and specific impulse. Planet-centered operations are also examined. A method to rotate a parking orbit about the line of apsides to achieve the proper orientation at departure is discussed, thus coupling the effects of parking-orbit orientation with the interplanetary trajectories. Also, a guidance algorithm for rendezvous during flybys in semi-cycler and cycler missions is presented with a control law for final approach. A forty-year plan to establish a permanent base on Mars is detailed and methods to expand the base are discussed. Once a large base is established, one-, two-, or three-vehicle systems may sustain the colonization of Mars.

Space shuttle propulsion systems on-board checkout and monitoring system development study

NASA Technical Reports Server (NTRS)

1971-01-01

Investigations on the fundamental space shuttle propulsion systems program are reported, with emphasis on in-depth reviews of preliminary drafts of the guidelines. The guidelines will be used to incorporate the onboard checkout and monitoring function into the basic design of the propulsion systems and associated interfacing systems. The analysis of checkout and monitoring requirements of the Titan 3 L expandable booster propulsion systems was completed, and the techniques for accomplishing the checkout and monitoring functions were determined. Updating results of the basic study of propulsion system checkout and monitoring is continuing.

Status of Propulsion Technology Development Under the NASA In-Space Propulsion Technology Program

NASA Technical Reports Server (NTRS)

Anderson, David; Kamhawi, Hani; Patterson, Mike; Pencil, Eric; Pinero, Luis; Falck, Robert; Dankanich, John

2014-01-01

Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing and delivering in-space propulsion technologies for NASA's Science Mission Directorate (SMD). These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems/Mission Analysis. ISPT's propulsion technologies include: 1) the 0.6-7 kW NASA's Evolutionary Xenon Thruster (NEXT) gridded ion propulsion system; 2) a 0.3-3.9kW Halleffect electric propulsion (HEP) system for low cost and sample return missions; 3) the Xenon Flow Control Module (XFCM); 4) ultra-lightweight propellant tank technologies (ULTT); and 5) propulsion technologies for a Mars Ascent Vehicle (MAV). The NEXT Long Duration Test (LDT) recently exceeded 50,000 hours of operation and 900 kg throughput, corresponding to 34.8 MN-s of total impulse delivered. The HEP system is composed of the High Voltage Hall Accelerator (HIVHAC) thruster, a power processing unit (PPU), and the XFCM. NEXT and the HIVHAC are throttle-able electric propulsion systems for planetary science missions. The XFCM and ULTT are two component technologies which being developed with nearer-term flight infusion in mind. Several of the ISPT technologies are related to sample return missions needs: MAV propulsion and electric propulsion. And finally, one focus of the Systems/Mission Analysis area is developing tools that aid the application or operation of these technologies on wide variety of mission concepts. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

High-Power Hall Propulsion Development at NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Manzella, David H.; Smith, Timothy D.; Schmidt, George R.

2014-01-01

The NASA Office of the Chief Technologist Game Changing Division is sponsoring the development and testing of enabling technologies to achieve efficient and reliable human space exploration. High-power solar electric propulsion has been proposed by NASA's Human Exploration Framework Team as an option to achieve these ambitious missions to near Earth objects. NASA Glenn Research Center (NASA Glenn) is leading the development of mission concepts for a solar electric propulsion Technical Demonstration Mission. The mission concepts are highlighted in this paper but are detailed in a companion paper. There are also multiple projects that are developing technologies to support a demonstration mission and are also extensible to NASA's goals of human space exploration. Specifically, the In-Space Propulsion technology development project at NASA Glenn has a number of tasks related to high-power Hall thrusters including performance evaluation of existing Hall thrusters; performing detailed internal discharge chamber, near-field, and far-field plasma measurements; performing detailed physics-based modeling with the NASA Jet Propulsion Laboratory's Hall2De code; performing thermal and structural modeling; and developing high-power efficient discharge modules for power processing. This paper summarizes the various technology development tasks and progress made to date

High-Power Hall Propulsion Development at NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Manzella, David H.; Smith, Timothy D.; Schmidt, George R.

2012-01-01

The NASA Office of the Chief Technologist Game Changing Division is sponsoring the development and testing of enabling technologies to achieve efficient and reliable human space exploration. High-power solar electric propulsion has been proposed by NASA's Human Exploration Framework Team as an option to achieve these ambitious missions to near Earth objects. NASA Glenn Research Center is leading the development of mission concepts for a solar electric propulsion Technical Demonstration Mission. The mission concepts are highlighted in this paper but are detailed in a companion paper. There are also multiple projects that are developing technologies to support a demonstration mission and are also extensible to NASA's goals of human space exploration. Specifically, the In-Space Propulsion technology development project at the NASA Glenn has a number of tasks related to high-power Hall thrusters including performance evaluation of existing Hall thrusters; performing detailed internal discharge chamber, near-field, and far-field plasma measurements; performing detailed physics-based modeling with the NASA Jet Propulsion Laboratory's Hall2De code; performing thermal and structural modeling; and developing high-power efficient discharge modules for power processing. This paper summarizes the various technology development tasks and progress made to date.

Mars lander survey

NASA Technical Reports Server (NTRS)

Stump, William R.; Babb, Gus R.; Davis, Hubert P.

1986-01-01

The requirements, issues, and design options are reviewed for manned Mars landers. Issues such as high 1/d versus low 1/d shape, parking orbit, and use of a small Mars orbit transfer vehicle to move the lander from orbit to orbit are addressed. Plots of lander mass as a function of Isp, destination orbit, and cargo up and down, plots of initial stack mass in low Earth orbit as a function of lander mass and parking orbit, detailed weight statements, and delta V tables for a variety of options are included. Lander options include a range from minimum landers up to a single stage reusable design. Mission options include conjunction and Venus flyby trajectories using all-cryogenic, hybrid, NERVA, and Mars orbit aerobraking propulsion concepts.

Status of Solar Sail Propulsion: Moving Toward an Interstellar Probe

NASA Technical Reports Server (NTRS)

Johnson, Les; Young, Roy M.; Montgomery, Edward E., IV

2006-01-01

NASA's In-Space Propulsion Technology Program has developed the first-generation of solar sail propulsion systems sufficient to accomplish inner solar system science and exploration missions. These first-generation solar sails, when operational, will range in size from 40 meters to well over 100 meters in diameter and have an areal density of less than 13 grams-per-square meter. A rigorous, multiyear technology development effort culminated last year in the testing of two different 20-meter solar sail systems under thermal vacuum conditions. This effort provided a number of significant insights into the optimal design and expected performance of solar sails as well as an understanding of the methods and costs of building and using them. In a separate effort, solar sail orbital analysis tools for mission design were developed and tested. Laboratory simulations of the effects of long-term space radiation exposure were also conducted on two candidate solar sail materials. Detailed radiation and charging environments were defined for mission trajectories outside the protection of the earth's magnetosphere, in the solar wind environment. These were used in other analytical tools to prove the adequacy of sail design features for accommodating the harsh space environment. Preceding, and in conjunction with these technology efforts, NASA sponsored several mission application studies for solar sails, including one that would use an evolved sail capability to support humanity's first mission into nearby interstellar space. The proposed mission is called the Interstellar Probe. The Interstellar Probe might be accomplished in several ways. A 200-meter sail, with an areal density approaching 1 gram-per-square meter, could accelerate a robotic probe to the very edge of the solar system in just under 20 years from launch. A sail using the technology just demonstrated could make the same mission, but take significantly longer. Conventional chemical propulsion systems would require even longer flight times. Spinner sails of the type being explored by the Japanese may also be a good option, but the level of maturity in that technology is not clear. While the technology to support a 200-meter, ultralightweight sail mission is not yet in hand, the recent NASA investments in solar sail technology are an essential first step toward making it a reality. This paper will describe the status of solar sail propulsion within NASA, near-term solar sail mission applications, and the plan to advance the technology to the point where the Interstellar Probe mission can be flown.

Advanced Electric Propulsion for Space Solar Power Satellites

NASA Technical Reports Server (NTRS)

Oleson, Steve

1999-01-01

The sun tower concept of collecting solar energy in space and beaming it down for commercial use will require very affordable in-space as well as earth-to-orbit transportation. Advanced electric propulsion using a 200 kW power and propulsion system added to the sun tower nodes can provide a factor of two reduction in the required number of launch vehicles when compared to in-space cryogenic chemical systems. In addition, the total time required to launch and deliver the complete sun tower system is of the same order of magnitude using high power electric propulsion or cryogenic chemical propulsion: around one year. Advanced electric propulsion can also be used to minimize the stationkeeping propulsion system mass for this unique space platform. 50 to 100 kW class Hall, ion, magnetoplasmadynamic, and pulsed inductive thrusters are compared. High power Hall thruster technology provides the best mix of launches saved and shortest ground to Geosynchronous Earth Orbital Environment (GEO) delivery time of all the systems, including chemical. More detailed studies comparing launch vehicle costs, transfer operations costs, and propulsion system costs and complexities must be made to down-select a technology. The concept of adding electric propulsion to the sun tower nodes was compared to a concept using re-useable electric propulsion tugs for Low Earth Orbital Environment (LEO) to GEO transfer. While the tug concept would reduce the total number of required propulsion systems, more launchers and notably longer LEO to GEO and complete sun tower ground to GEO times would be required. The tugs would also need more complex, longer life propulsion systems and the ability to dock with sun tower nodes.

Certification Testing Approach for Propulsion System Design

NASA Technical Reports Server (NTRS)

Rodriguez, Henry; Popp, Chris

2005-01-01

The Certification of Propulsion Systems is costly and complex which involves development and qualification testing. The desire of the certification process is to assure all requirements can be demonstrated to be compliant. The purpose of this paper is to address the technical design concerns of certifying a system for flight. The authors of this paper have experience the lessons learned from supporting the Shuttle Program for Main Propulsion and On Orbit Propulsions Systems. They have collaborated design concerns for certifying propulsion systems. Presented are Pressurization, Tankage, Feed System and Combustion Instability concerns. Propulsion System Engineers are challenged with the dilemma for testing new systems to specific levels to reduce risk yet maintain budgetary targets. A methodical approach is presented to define the types of test suitable to address the technical issues for qualifying systems for retiring the risk levels.

Centralized versus distributed propulsion

NASA Technical Reports Server (NTRS)

Clark, J. P.

1982-01-01

The functions and requirements of auxiliary propulsion systems are reviewed. None of the three major tasks (attitude control, stationkeeping, and shape control) can be performed by a collection of thrusters at a single central location. If a centralized system is defined as a collection of separated clusters, made up of the minimum number of propulsion units, then such a system can provide attitude control and stationkeeping for most vehicles. A distributed propulsion system is characterized by more numerous propulsion units in a regularly distributed arrangement. Various proposed large space systems are reviewed and it is concluded that centralized auxiliary propulsion is best suited to vehicles with a relatively rigid core. These vehicles may carry a number of flexible or movable appendages. A second group, consisting of one or more large flexible flat plates, may need distributed propulsion for shape control. There is a third group, consisting of vehicles built up from multiple shuttle launches, which may be forced into a distributed system because of the need to add additional propulsion units as the vehicles grow. The effects of distributed propulsion on a beam-like structure were examined. The deflection of the structure under both translational and rotational thrusts is shown as a function of the number of equally spaced thrusters. When two thrusters only are used it is shown that location is an important parameter. The possibility of using distributed propulsion to achieve minimum overall system weight is also examined. Finally, an examination of the active damping by distributed propulsion is described.

Electric and hybrid vehicle system R/D

NASA Technical Reports Server (NTRS)

Schwartz, H. J.

1980-01-01

The work being done to characterize the level of current propulsion technology through component testing is described. Important interactions between the battery and the propulsion system will be discussed. Component development work, involving traction motors, motor controllers and transmissions are described and current results are presented. Studies of advanced electric and hybrid propulsion system studies are summarized and the status of propulsion system development work supported by the project is described. A strategy for fostering joint industry/government projects for commercialization of propulsion components and systems is described briefly.

Derivation of Failure Rates and Probability of Failures for the International Space Station Probabilistic Risk Assessment Study

NASA Technical Reports Server (NTRS)

Vitali, Roberto; Lutomski, Michael G.

2004-01-01

National Aeronautics and Space Administration s (NASA) International Space Station (ISS) Program uses Probabilistic Risk Assessment (PRA) as part of its Continuous Risk Management Process. It is used as a decision and management support tool to not only quantify risk for specific conditions, but more importantly comparing different operational and management options to determine the lowest risk option and provide rationale for management decisions. This paper presents the derivation of the probability distributions used to quantify the failure rates and the probability of failures of the basic events employed in the PRA model of the ISS. The paper will show how a Bayesian approach was used with different sources of data including the actual ISS on orbit failures to enhance the confidence in results of the PRA. As time progresses and more meaningful data is gathered from on orbit failures, an increasingly accurate failure rate probability distribution for the basic events of the ISS PRA model can be obtained. The ISS PRA has been developed by mapping the ISS critical systems such as propulsion, thermal control, or power generation into event sequences diagrams and fault trees. The lowest level of indenture of the fault trees was the orbital replacement units (ORU). The ORU level was chosen consistently with the level of statistically meaningful data that could be obtained from the aerospace industry and from the experts in the field. For example, data was gathered for the solenoid valves present in the propulsion system of the ISS. However valves themselves are composed of parts and the individual failure of these parts was not accounted for in the PRA model. In other words the failure of a spring within a valve was considered a failure of the valve itself.

A Densified Liquid Methane Delivery System for the Altair Ascent Stage

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.; Johnson, Wesley L.; Smudde, Todd D.; Femminineo, Mark F.; Schnell, Andrew R.

2010-01-01

The Altair Lunar Lander is currently carrying options for both cryogenic and hypergolic ascent stage propulsion modules. The cryogenic option uses liquid methane and liquid oxygen to propel Altair from the lunar surface back to rendezvous with the Orion command module. Recent studies have determined that the liquid methane should be densified by subcooling it to 93 K in order to prevent over-pressurization of the propellant tanks during the 210 day stay on the lunar surface. A trade study has been conducted to determine the preferred method of producing; loading, and maintaining the subcooled, densified liquid methane onboard Altair from a ground operations perspective. The trade study took into account the limitations in mass for the launch vehicle and the mobile launch platform as well as the historical reliability of various components and their thermal efficiencies. Several unique problems were encountered, namely delivering a small amount of a cryogenic propellant to a flight tank that is positioned over 350 ft above the launch pad as well as generating the desired delivery temperature of the methane at 93 K which is only 2.3 K above the methane triple point of 90.7 K. Over 20 methods of subcooled liquid methane production and delivery along with the associated system architectures were investigated to determine the best solutions to the problem. The top four cryogenic processing solutions were selected for further evaluation and detailed thermal modeling. This paper describes the results of the preliminary trade analysis of the 20 plus methane densification methods considered. The results of the detailed analysis will be briefed to the Altair Project Office and their propulsion team as well as the Ground Operations Project Office before the down-select is made between cryogenic and hypergolic ascent stages in August 2010.

Mission and space vehicle sizing data for a chemical propulsion/aerobraking option

NASA Technical Reports Server (NTRS)

Butler, John; Brothers, Bobby

1986-01-01

Sizing data is presented for various combinations of Mars missions and chemical-propulsion/aerobraking vehicles. Data is compared for vehicles utilizing opposition (2-year mission) and conjunction (3-year mission) trajectories for 1999 and 2001 opportunities, for various sizes of vehicles. Payload capabilities for manned and unmanned missions vehicles and for propulsive-braking and aerobraking cases are shown. The effect of scaling up a reference vehicle is compared to the case of utilizing two identical vehicles, for growth in payload capability. The rate of cumulative build up of weight on the surface of Mars is examined for various mission/vehicle combinations, and is compared to the landed-weight requirements for sortie missions, moving-base missions, and fixed-base missions. Also, the required buildup of weight in low Earth orbit (LEO) for various mission/vehicle combinations is presented and discussed.

YF-12 propulsion research program and results

NASA Technical Reports Server (NTRS)

Albers, J. A.; Olinger, F. V.

1976-01-01

The objectives and status of the propulsion program, along with the results acquired in the various technology areas, are discussed. The instrumentation requirements for and experience with flight testing the propulsion systems at high supersonic cruise are reported. Propulsion system performance differences between wind tunnel and flight are given. The effects of high frequency flow fluctuations (transients) on the stability of the propulsion system are described, and shock position control is evaluated.

Operationally efficient propulsion system study (OEPSS) data book. Volume 6; Space Transfer Propulsion Operational Efficiency Study Task of OEPSS

NASA Technical Reports Server (NTRS)

Harmon, Timothy J.

1992-01-01

This document is the final report for the Space Transfer Propulsion Operational Efficiency Study Task of the Operationally Efficient Propulsion System Study (OEPSS) conducted by the Rocketdyne Division of Rockwell International. This Study task studied, evaluated and identified design concepts and technologies which minimized launch and in-space operations and optimized in-space vehicle propulsion system operability.

Development of unified propulsion system for geostationary satellite

NASA Astrophysics Data System (ADS)

Murayama, S.; Kobayashi, H.; Masuda, I.; Kameishi, M.; Miyoshi, K.; Takahashi, M.

Japan's first Liquid Apogee Propulsion System (LAPS) has been developed for ETS-VI (Engineering Test Satellite - VI) 2-ton class geostationary satellite. The next largest (2-ton class) geostationary satellite, COMETS (Communication and Broadcasting Engineering Test Satellite), requires a more compact apogee propulsion system in order to increase the space for mission instruments. The study for such a propulsion system concluded with a Unified Propulsion System (UPS), which uses a common N2H4 propellant tank for both bipropellant apogee engines and monopropellant Reaction Control System (RCS) thrusters. This type of propulsion system has several significant advantages compared with popular nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) bipropellant satellite propulsion systems: The NTO/N2H4 apogee engine has a high specific impulse, and N2H4 thrusters have high reliability. Residual of N2H4 caused by propellant utilization of apogee engine firing (AEF) can be consumed by N2H4 monopropellant thrusters; that means a considerably prolonged satellite life.

Propulsion System Dynamic Modeling for the NASA Supersonic Concept Vehicle: AeroPropulsoServoElasticity

NASA Technical Reports Server (NTRS)

Kopasakis, George; Connolly, Joseph; Seidel, Jonathan

2014-01-01

A summary of the propulsion system modeling under NASA's High Speed Project (HSP) AeroPropulsoServoElasticity (APSE) task is provided with a focus on the propulsion system for the low-boom supersonic configuration developed by Lockheed Martin and referred to as the N+2 configuration. This summary includes details on the effort to date to develop computational models for the various propulsion system components. The objective of this paper is to summarize the model development effort in this task, while providing more detail in the modeling areas that have not been previously published. The purpose of the propulsion system modeling and the overall APSE effort is to develop an integrated dynamic vehicle model to conduct appropriate unsteady analysis of supersonic vehicle performance. This integrated APSE system model concept includes the propulsion system model, and the vehicle structural-aerodynamics model. The development to date of such a preliminary integrated model will also be summarized in this report.propulsion system dynamics, the structural dynamics, and aerodynamics.

An Object Oriented Extensible Architecture for Affordable Aerospace Propulsion Systems

NASA Technical Reports Server (NTRS)

Follen, Gregory J.

2003-01-01

Driven by a need to explore and develop propulsion systems that exceeded current computing capabilities, NASA Glenn embarked on a novel strategy leading to the development of an architecture that enables propulsion simulations never thought possible before. Full engine 3 Dimensional Computational Fluid Dynamic propulsion system simulations were deemed impossible due to the impracticality of the hardware and software computing systems required. However, with a software paradigm shift and an embracing of parallel and distributed processing, an architecture was designed to meet the needs of future propulsion system modeling. The author suggests that the architecture designed at the NASA Glenn Research Center for propulsion system modeling has potential for impacting the direction of development of affordable weapons systems currently under consideration by the Applied Vehicle Technology Panel (AVT).

46 CFR 58.01-35 - Main propulsion auxiliary machinery.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 46 Shipping 2 2011-10-01 2011-10-01 false Main propulsion auxiliary machinery. 58.01-35 Section 58... AUXILIARY MACHINERY AND RELATED SYSTEMS General Requirements § 58.01-35 Main propulsion auxiliary machinery. Auxiliary machinery vital to the main propulsion system must be provided in duplicate unless the system...

46 CFR 58.01-35 - Main propulsion auxiliary machinery.

Code of Federal Regulations, 2010 CFR

2010-10-01

... 46 Shipping 2 2010-10-01 2010-10-01 false Main propulsion auxiliary machinery. 58.01-35 Section 58... AUXILIARY MACHINERY AND RELATED SYSTEMS General Requirements § 58.01-35 Main propulsion auxiliary machinery. Auxiliary machinery vital to the main propulsion system must be provided in duplicate unless the system...

46 CFR 121.620 - Propulsion engine control systems.

Code of Federal Regulations, 2011 CFR

2011-10-01

... 46 Shipping 4 2011-10-01 2011-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and engine...

46 CFR 58.01-35 - Main propulsion auxiliary machinery.

Code of Federal Regulations, 2012 CFR

2012-10-01

... 46 Shipping 2 2012-10-01 2012-10-01 false Main propulsion auxiliary machinery. 58.01-35 Section 58... AUXILIARY MACHINERY AND RELATED SYSTEMS General Requirements § 58.01-35 Main propulsion auxiliary machinery. Auxiliary machinery vital to the main propulsion system must be provided in duplicate unless the system...

46 CFR 121.620 - Propulsion engine control systems.

Code of Federal Regulations, 2012 CFR

2012-10-01

... 46 Shipping 4 2012-10-01 2012-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and engine...

46 CFR 58.01-35 - Main propulsion auxiliary machinery.

Code of Federal Regulations, 2013 CFR

2013-10-01

... 46 Shipping 2 2013-10-01 2013-10-01 false Main propulsion auxiliary machinery. 58.01-35 Section 58... AUXILIARY MACHINERY AND RELATED SYSTEMS General Requirements § 58.01-35 Main propulsion auxiliary machinery. Auxiliary machinery vital to the main propulsion system must be provided in duplicate unless the system...

46 CFR 121.620 - Propulsion engine control systems.

Code of Federal Regulations, 2014 CFR

2014-10-01

... 46 Shipping 4 2014-10-01 2014-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and engine...

46 CFR 121.620 - Propulsion engine control systems.

Code of Federal Regulations, 2010 CFR

2010-10-01

... 46 Shipping 4 2010-10-01 2010-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and engine...

46 CFR 58.01-35 - Main propulsion auxiliary machinery.

Code of Federal Regulations, 2014 CFR

2014-10-01

... 46 Shipping 2 2014-10-01 2014-10-01 false Main propulsion auxiliary machinery. 58.01-35 Section 58... AUXILIARY MACHINERY AND RELATED SYSTEMS General Requirements § 58.01-35 Main propulsion auxiliary machinery. Auxiliary machinery vital to the main propulsion system must be provided in duplicate unless the system...

46 CFR 121.620 - Propulsion engine control systems.

Code of Federal Regulations, 2013 CFR

2013-10-01

... 46 Shipping 4 2013-10-01 2013-10-01 false Propulsion engine control systems. 121.620 Section 121... Propulsion engine control systems. (a) A vessel must have two independent means of controlling each propulsion engine. Control must be provided for the engine speed, direction of shaft rotation, and engine...

Preliminary study of propulsion systems and airplane wing parameters for a US Navy subsonic V/STOL aircraft

NASA Technical Reports Server (NTRS)

Zola, C. L.; Fishbach, L. H.; Allen, J. L.

1978-01-01

Two V/STOL propulsion concepts were evaluated in a common aircraft configuration. One propulsion system consists of cross coupled turboshaft engines driving variable pitch fans. The other system is a gas coupled combination of turbojet gas generators and tip turbine fixed pitch fans. Evaluations were made of endurance at low altitude, low speed loiter with equal takeoff fuel loads. Effects of propulsion system sizing, bypass ratio, and aircraft wing planform parameters were investigated and compared. Shaft driven propulsion systems appear to result in better overall performance, although at higher installed weight, than gas systems.

Performance analysis of the ascent propulsion system of the Apollo spacecraft

NASA Technical Reports Server (NTRS)

Hooper, J. C., III

1973-01-01

Activities involved in the performance analysis of the Apollo lunar module ascent propulsion system are discussed. A description of the ascent propulsion system, including hardware, instrumentation, and system characteristics, is included. The methods used to predict the inflight performance and to establish performance uncertainties of the ascent propulsion system are discussed. The techniques of processing the telemetered flight data and performing postflight performance reconstruction to determine actual inflight performance are discussed. Problems that have been encountered and results from the analysis of the ascent propulsion system performance during the Apollo 9, 10, and 11 missions are presented.

Advanced transportation system studies. Alternate propulsion subsystem concepts: Propulsion database

NASA Technical Reports Server (NTRS)

Levack, Daniel

1993-01-01

The Advanced Transportation System Studies alternate propulsion subsystem concepts propulsion database interim report is presented. The objective of the database development task is to produce a propulsion database which is easy to use and modify while also being comprehensive in the level of detail available. The database is to be available on the Macintosh computer system. The task is to extend across all three years of the contract. Consequently, a significant fraction of the effort in this first year of the task was devoted to the development of the database structure to ensure a robust base for the following years' efforts. Nonetheless, significant point design propulsion system descriptions and parametric models were also produced. Each of the two propulsion databases, parametric propulsion database and propulsion system database, are described. The descriptions include a user's guide to each code, write-ups for models used, and sample output. The parametric database has models for LOX/H2 and LOX/RP liquid engines, solid rocket boosters using three different propellants, a hybrid rocket booster, and a NERVA derived nuclear thermal rocket engine.

Automated screening of propulsion system test data by neural networks, phase 1

NASA Technical Reports Server (NTRS)

Hoyt, W. Andes; Whitehead, Bruce A.

1992-01-01

The evaluation of propulsion system test and flight performance data involves reviewing an extremely large volume of sensor data generated by each test. An automated system that screens large volumes of data and identifies propulsion system parameters which appear unusual or anomalous will increase the productivity of data analysis. Data analysts may then focus on a smaller subset of anomalous data for further evaluation of propulsion system tests. Such an automated data screening system would give NASA the benefit of a reduction in the manpower and time required to complete a propulsion system data evaluation. A phase 1 effort to develop a prototype data screening system is reported. Neural networks will detect anomalies based on nominal propulsion system data only. It appears that a reasonable goal for an operational system would be to screen out 95 pct. of the nominal data, leaving less than 5 pct. needing further analysis by human experts.

Nuclear power technology requirements for NASA exploration missions

NASA Technical Reports Server (NTRS)

Bloomfield, Harvey S.

1990-01-01

It is pointed out that future exploration of the moon and Mars will mandate developments in many areas of technology. In particular, major advances will be required in planet surface power systems. Critical nuclear technology challenges that can enable strategic self-sufficiency, acceptable operational costs, and cost-effective space transportation goals for NASA exploration missions have been identified. Critical technologies for surface power systems include stationary and mobile nuclear reactor and radioisotope heat sources coupled to static and dynamic power conversion devices. These technologies can provide dramatic reductions in mass, leading to operational and transportation cost savings. Critical technologies for space transportation systems include nuclear thermal rocket and nuclear electric propulsion options, which present compelling concepts for significantly reducing mass, cost, or travel time required for Earth-Mars transport.

Maintenance of contamination sensitive surfaces on board long-term space vehicles

NASA Technical Reports Server (NTRS)

Phillips, A.; Maag, C.

1984-01-01

In the current age, highly sensitive instruments are being flown on spacecraft, and questions of contamination have become important. The present investigation is concerned with the available approaches which can provide long-term protection for contamination sensitive surfaces. Aspects and sources of spacecraft contamination are examined, taking into account materials outgassing, particulates, propulsion system interaction, overboard venting, man-made and cosmic debris, and atomic oxygen/ambient atmosphere interaction. Suitable protection approaches provided by current technology are discussed, giving attention to aperture covers, a possibility for a retractable cover design, gaseous purges, options for prolonging the lifetime of the thermal control system, and plume shields. Some new possibilities considered are related to an early warning system for excessive amounts of contamination, a molecular/wake shield, and the use of atomic oxygen.

Applications of plasma core reactors to terrestrial energy systems

NASA Technical Reports Server (NTRS)

Latham, T. S.; Biancardi, F. R.; Rodgers, R. J.

1974-01-01

Plasma core reactors offer several new options for future energy needs in addition to space power and propulsion applications. Power extraction from plasma core reactors with gaseous nuclear fuel allows operation at temperatures higher than conventional reactors. Highly efficient thermodynamic cycles and applications employing direct coupling of radiant energy are possible. Conceptual configurations of plasma core reactors for terrestrial applications are described. Closed-cycle gas turbines, MHD systems, photo- and thermo-chemical hydrogen production processes, and laser systems using plasma core reactors as prime energy sources are considered. Cycle efficiencies in the range of 50 to 65 percent are calculated for closed-cycle gas turbine and MHD electrical generators. Reactor advantages include continuous fuel reprocessing which limits inventory of radioactive by-products and thorium-U-233 breeder configurations with about 5-year doubling times.-

Technology Area Roadmap for In Space Propulsion Technologies

NASA Technical Reports Server (NTRS)

Johnson, Les; Meyer, Mike; Coote, David; Goebel, Dan; Palaszewski, Bryan; White, Sonny

2010-01-01

This slide presentation reviews the technology area (TA) roadmap to develop propulsion technologies that will be used to enable further exploration of the solar system, and beyond. It is hoped that development of the technologies within this TA will result in technical solutions that will improve thrust levels, specific impulse, power, specific mass, volume, system mass, system complexity, operational complexity, commonality with other spacecraft systems, manufacturability and durability. Some of the propulsion technologies that are reviewed include: chemical and non-chemical propulsion, and advanced propulsion (i.e., those with a Technology Readiness level of less than 3). Examples of these advanced technologies include: Beamed Energy, Electric Sail, Fusion, High Energy Density Materials, Antimatter, Advanced Fission and Breakthrough propulsion technologies. Timeframes for development of some of these propulsion technologies are reviewed, and top technical challenges are reviewed. This roadmap describes a portfolio of in-space propulsion technologies that can meet future space science and exploration needs.

Space station integrated propulsion and fluid systems study

NASA Technical Reports Server (NTRS)

Bicknell, B.; Wilson, S.; Dennis, M.; Shepard, D.; Rossier, R.

1988-01-01

The program study was performed in two tasks: Task 1 addressed propulsion systems and Task 2 addressed all fluid systems associated with the Space Station elements, which also included propulsion and pressurant systems. Program results indicated a substantial reduction in life cycle costs through integrating the oxygen/hydrogen propulsion system with the environmental control and life support system, and through supplying nitrogen in a cryogenic gaseous supercritical or subcritical liquid state. A water sensitivity analysis showed that increasing the food water content would substantially increase the amount of water available for propulsion use and in all cases, the implementation of the BOSCH CO2 reduction process would reduce overall life cycle costs to the station and minimize risk. An investigation of fluid systems and associated requirements revealed a delicate balance between the individual propulsion and fluid systems across work packages and a strong interdependence between all other fluid systems.

Engineering of the Magnetized Target Fusion Propulsion System

NASA Technical Reports Server (NTRS)

Statham, G.; White, S.; Adams, R. B.; Thio, Y. C. F.; Santarius, J.; Alexander, R.; Chapman, J.; Fincher, S.; Philips, A.; Polsgrove, T.

2003-01-01

Engineering details are presented for a magnetized target fusion (MTF) propulsion system designed to support crewed missions to the outer solar system. Basic operation of an MTF propulsion system is introduced. Structural, thermal, radiation-management and electrical design details are presented. The propellant storage and supply system design is also presented. A propulsion system mass estimate and associated performance figures are given. The advantages of helium-3 as a fusion fuel for an advanced MTF system are discussed.

Solar Electric Propulsion System Integration Technology (SEPSIT). Volume 2: Encke rendezvous mission and space vehicle functional description

NASA Technical Reports Server (NTRS)

Gardner, J. A.

1972-01-01

A solar electric propulsion system integration technology study is discussed. Detailed analyses in support of the solar electric propulsion module were performed. The thrust subsystem functional description is presented. The space vehicle and the space mission to which the propulsion system is applied are analyzed.

Clean Cities Guide to Alternative Fuel and Advanced Medium- and Heavy-Duty Vehicles (Book)

DOE Office of Scientific and Technical Information (OSTI.GOV)

Not Available

2013-08-01

Today's fleets are increasingly interested in medium-duty and heavy-duty vehicles that use alternative fuels or advanced technologies that can help reduce operating costs, meet emissions requirements, improve fleet sustainability, and support U.S. energy independence. Vehicle and engine manufacturers are responding to this interest with a wide range of options across a steadily growing number of vehicle applications. This guide provides an overview of alternative fuel power systems?including engines, microturbines, electric motors, and fuel cells?and hybrid propulsion systems. The guide also offers a list of individual medium- and heavy-duty vehicle models listed by application, along with associated manufacturer contact information, fuelmore » type(s), power source(s), and related information.« less

Clean Cities Guide to Alternative Fuel and Advanced Medium- and Heavy-Duty Vehicles

DOE Office of Scientific and Technical Information (OSTI.GOV)

None

2013-08-01

Today's fleets are increasingly interested in medium-duty and heavy-duty vehicles that use alternative fuels or advanced technologies that can help reduce operating costs, meet emissions requirements, improve fleet sustainability, and support U.S. energy independence. Vehicle and engine manufacturers are responding to this interest with a wide range of options across a steadily growing number of vehicle applications. This guide provides an overview of alternative fuel power systems--including engines, microturbines, electric motors, and fuel cells--and hybrid propulsion systems. The guide also offers a list of individual medium- and heavy-duty vehicle models listed by application, along with associated manufacturer contact information, fuelmore » type(s), power source(s), and related information.« less

Applying the OTV to lunar logistics

NASA Technical Reports Server (NTRS)

Willcockson, W. H.

1988-01-01

The Orbit Transfer Vehicle (OTV), representing the next generation of upper stages, has recently been studied in a Phase A concept definition study managed by NASA's Marshall Space Flight Center. The vehicle has been previously defined as strictly an orbit-to-orbit type transfer device. Recently its application to the task of lunar surface logistics was investigated. Transfer options to the surface were considered which included direct transfer, and transfer via lunar orbit as well as the L1 libration point. The subsystem modifications required to enable lunar landings were established for the following elements: aerobrake, main propulsion system, landing legs, primary structure, and avionics. It is concluded that the majority of the basic systems required for efficient transfer to the lunar surface are already contained in the OTV.

Feasibility of a responsive, hybrid propulsion augmented, Vertical-Takeoff-and-Landing, Single-Stage-to-Orbit launch system

NASA Astrophysics Data System (ADS)

Pelaccio, Dennis G.

1996-03-01

A novel, reusable, Vertical-Takeoff-and-Landing, Single-Stage-to-Orbit (VTOL/SSTO) launch system concept, named HYP-SSTO, is presented in this paper. This launch vehicle system concept uses a highly coupled, main high performance liquid oxygen/liquid hydrogen (LOX/LH2) propulsion system, that is used only for launch, with a hybrid auxiliary propulsion system which is used during final orbit insertion, major orbit maneuvering, and landing propulsive burn phases of flight. By using a hybrid propulsion system for major orbit maneuver burns and landing, this launch system concept has many advantages over conventional VTOL/SSTO concepts that use LOX/LH2 propulsion system(s) burns for all phases of flight. Because hybrid propulsion systems are relatively simple and inert by their nature, this concept has the potential to support short turnaround times between launches, be economical to develop, and be competitive in terms of overall system life-cycle cost. This paper provides a technical description of the novel, reusable HYP-SSTO launch system concept. Launch capability performance, as well as major design and operational system attributes, are identified and discussed.

Ares V Utilization in Support of a Human Mission to Mars

NASA Technical Reports Server (NTRS)

Holladay, J. B.; Jaap, J. P.; Pinson, R. M.; Creech, S. D.; Ryan, R. M.; Monk, T. S.; Baggett. K. E.; Runager, M. D.; Dux, I. J.; Hack, K. J.;

Using Additive Manufacturing to Print a CubeSat Propulsion System

NASA Technical Reports Server (NTRS)

Marshall, William M.

2015-01-01

CubeSats are increasingly being utilized for missions traditionally ascribed to larger satellites CubeSat unit (1U) defined as 10 cm x 10 cm x 11 cm. Have been built up to 6U sizes. CubeSats are typically built up from commercially available off-the-shelf components, but have limited capabilities. By using additive manufacturing, mission specific capabilities (such as propulsion), can be built into a system. This effort is part of STMD Small Satellite program Printing the Complete CubeSat. Interest in propulsion concepts for CubeSats is rapidly gaining interest-Numerous concepts exist for CubeSat scale propulsion concepts. The focus of this effort is how to incorporate into structure using additive manufacturing. End-use of propulsion system dictates which type of system to develop-Pulse-mode RCS would require different system than a delta-V orbital maneuvering system. Team chose an RCS system based on available propulsion systems and feasibility of printing using a materials extrusion process. Initially investigated a cold-gas propulsion system for RCS applications-Materials extrusion process did not permit adequate sealing of part to make this a functional approach.

Rocket-Based Combined-Cycle (RBCC) Propulsion Technology Workshop. Tutorial session

NASA Technical Reports Server (NTRS)

1992-01-01

The goal of this workshop was to illuminate the nation's space transportation and propulsion engineering community on the potential of hypersonic combined cycle (airbreathing/rocket) propulsion systems for future space transportation applications. Four general topics were examined: (1) selections from the expansive advanced propulsion archival resource; (2) related propulsion systems technical backgrounds; (3) RBCC engine multimode operations related subsystem background; and (4) focused review of propulsion aspects of current related programs.

Advanced Chemical Propulsion

NASA Technical Reports Server (NTRS)

Alexander, Leslie, Jr.

2006-01-01

Advanced Chemical Propulsion (ACP) provides near-term incremental improvements in propulsion system performance and/or cost. It is an evolutionary approach to technology development that produces useful products along the way to meet increasingly more demanding mission requirements while focusing on improving payload mass fraction to yield greater science capability. Current activities are focused on two areas: chemical propulsion component, subsystem, and manufacturing technologies that offer measurable system level benefits; and the evaluation of high-energy storable propellants with enhanced performance for in-space application. To prioritize candidate propulsion technology alternatives, a variety of propulsion/mission analyses and trades have been conducted for SMD missions to yield sufficient data for investment planning. They include: the Advanced Chemical Propulsion Assessment; an Advanced Chemical Propulsion System Model; a LOx-LH2 small pumps conceptual design; a space storables propellant study; a spacecraft cryogenic propulsion study; an advanced pressurization and mixture ratio control study; and a pump-fed vs. pressure-fed study.

Opening the Solar System: An Advanced Nuclear Spacecraft for Human Exploration

NASA Technical Reports Server (NTRS)

Werka, R. O.; Percy, T. K.

2014-01-01

Human exploration of the solar system is limited by our technology, not our imagination. We dream of a time when we can freely travel among the planets and truly become a spacefaring people. However, the current state of our technology limits our options for architecting missions to other planets. Instead of sailing the seas of space in the way that we cruise the seas of Earth, our limited propulsion technology requires us to depart Earth on a giant cluster of gas tanks and return in a lifeboat. This inefficient approach to exploration is evident in many of today's leading mission plans for human flights to Mars, asteroids, and other destinations. The cost and complexity of this approach to mission architecting makes it extremely difficult to realize our dreams of exploration beyond Low Earth Orbit (LEO). This does not need to be the case. Researchers at NASA's Marshall Space Flight Center (MSFC) have been investigating the feasibility of a new take on nuclear propulsion with the performance to enable a paradigm shift in human space exploration. During the fall of 2013, engineers at MSFC's Advanced Concepts Office developed a spacecraft concept (pictured below) around this new propulsion technology and redefined the human Mars mission to show its full potential. This spacecraft, which can be launched with a fleet of soon-to-be available SLS launch vehicles, is fueled primarily with hydrogen, and is fully reusable with no staging required. The reusable nature of this design enables a host of alternative mission architectures that more closely resemble an ocean voyage than our current piecemeal approach to exploration.

GPIM AF-M315E Propulsion System

NASA Technical Reports Server (NTRS)

Spores, Ronald A.; Masse, Robert; Kimbrel, Scott; McLean, Chris

2014-01-01

The NASA Space Technology mission Directorate's (STMD) Green Propellant Infusion Mission (GPIM) Technology Demonstration Mission (TDM) will demonstrate an operational AF-M315E green propellant propulsion system. Aerojet-Rocketdyne is responsible for the development of the propulsion system payload. This paper statuses the propulsion system module development, including thruster design and system design; Initial test results for the 1N engineering model thruster are presented. The culmination of this program will be high-performance, green AF-M315E propulsion system technology at TRL 7+, with components demonstrated to TRL 9, ready for direct infusion to a wide range of applications for the space user community.

In-Space Transportation for NASA's Evolvable Mars Campaign

NASA Technical Reports Server (NTRS)

Percy, Thomas K.; McGuire, Melissa; Polsgrove, Tara

2015-01-01

As the nation embarks on a new and bold journey to Mars, significant work is being done to determine what that mission and those architectural elements will look like. The Evolvable Mars Campaign, or EMC, is being evaluated as a potential approach to getting humans to Mars. Built on the premise of leveraging current technology investments and maximizing element commonality to reduce cost and development schedule, the EMC transportation architecture is focused on developing the elements required to move crew and equipment to Mars as efficiently and effectively as possible both from a performance and a programmatic standpoint. Over the last 18 months the team has been evaluating potential options for those transportation elements. One of the key aspects of the EMC is leveraging investments being made today in missions like the Asteroid Redirect Mission (ARM) mission using derived versions of the Solar Electric Propulsion (SEP) propulsion systems and coupling them with other chemical propulsion elements that maximize commonality across the architecture between both transportation and Mars operations elements. This paper outlines the broad trade space being evaluated including the different technologies being assessed for transportation elements and how those elements are assembled into an architecture. Impacts to potential operational scenarios at Mars are also investigated. Trades are being made on the size and power level of the SEP vehicle for delivering cargo as well as the size of the chemical propulsion systems and various mission aspects including Inspace assembly and sequencing. Maximizing payload delivery to Mars with the SEP vehicle will better support the operational scenarios at Mars by enabling the delivery of landers and habitation elements that are appropriately sized for the mission. The purpose of this investigation is not to find the solution but rather a suite of solutions with potential application to the challenge of sending cargo and crew to Mars. The goal is that, by building an architecture intelligently with all aspects considered, the sustainable Mars program wisely invests limited resources enabling a long-term human Mars exploration program.

Nuclear Propulsion Technical Interchange Meeting, volume 2

NASA Technical Reports Server (NTRS)

1993-01-01

The purpose of the meeting was to review the work performed in fiscal year 1992 in the areas of nuclear thermal and nuclear electric propulsion technology development. These proceedings are an accumulation of the presentations provided at the meeting along with annotations provided by authors. The proceedings cover system concepts, technology development, and system modeling for nuclear thermal propulsion (NTP) and nuclear electric propulsion (NEP). The test facilities required for the development of the nuclear propulsion systems are also discussed.

National Institute for Rocket Propulsion Systems (NIRPS): Solutions Facilitator

NASA Technical Reports Server (NTRS)

Brown, Tom

2011-01-01

National Institute for Rocket Propulsion Systems (NIRPS) "Solutions" plans to enable our nation's future in rocket propulsion systems by leveraging existing skills and capabilities to support industry's future needs

Hybrid rocket propulsion

NASA Technical Reports Server (NTRS)

Holzman, Allen L.

1993-01-01

Topics addressed are: (1) comparison of the theoretical impulses; (2) comparison of the density-specific impulses; (3) general propulsion system features comparison; (4) hybrid systems, booster applications; and (5) hybrid systems, upper stage propulsion applications.

Modeling of Spacecraft Advanced Chemical Propulsion Systems

NASA Technical Reports Server (NTRS)

Benfield, Michael P. J.; Belcher, Jeremy A.

2004-01-01

This paper outlines the development of the Advanced Chemical Propulsion System (ACPS) model for Earth and Space Storable propellants. This model was developed by the System Technology Operation of SAIC-Huntsville for the NASA MSFC In-Space Propulsion Project Office. Each subsystem of the model is described. Selected model results will also be shown to demonstrate the model's ability to evaluate technology changes in chemical propulsion systems.

Solar electric propulsion system technology

NASA Technical Reports Server (NTRS)

Masek, T. D.; Macie, T. W.

1971-01-01

Achievements in the solar electric propulsion system technology program (SEPST 3) are reported and certain propulsion system-spacecraft interaction problems are discussed. The basic solar electric propulsion system concept and elements are reviewed. Hardware is discussed only briefly, relying on detailed fabrication or assembly descriptions reported elsewhere. Emphasis is placed on recent performance data, which are presented to show the relationship between spacecraft requirements and present technology.

Mirror fusion propulsion system: A performance comparison with alternate propulsion systems for the manned Mars Mission

NASA Technical Reports Server (NTRS)

Schulze, Norman R.; Carpenter, Scott A.; Deveny, Marc E.; Oconnell, T.

1993-01-01

The performance characteristics of several propulsion technologies applied to piloted Mars missions are compared. The characteristics that are compared are Initial Mass in Low Earth Orbit (IMLEO), mission flexibility, and flight times. The propulsion systems being compared are both demonstrated and envisioned: Chemical (or Cryogenic), Nuclear Thermal Rocket (NTR) solid core, NTR gas core, Nuclear Electric Propulsion (NEP), and a mirror fusion space propulsion system. The proposed magnetic mirror fusion reactor, known as the Mirror Fusion Propulsion System (MFPS), is described. The description is an overview of a design study that was conducted to convert a mirror reactor experiment at Lawrence Livermore National Lab (LLNL) into a viable space propulsion system. Design principles geared towards minimizing mass and maximizing power available for thrust are identified and applied to the LLNL reactor design, resulting in the MFPS. The MFPS' design evolution, reactor and fuel choices, and system configuration are described. Results of the performance comparison shows that the MFPS minimizes flight time to 60 to 90 days for flights to Mars while allowing continuous return-home capability while at Mars. Total MFPS IMLEO including propellant and payloads is kept to about 1,000 metric tons.

Mirror fusion propulsion system - A performance comparison with alternate propulsion systems for the manned Mars mission

NASA Technical Reports Server (NTRS)

Deveny, M.; Carpenter, S.; O'Connell, T.; Schulze, N.

1993-01-01

The performance characteristics of several propulsion technologies applied to piloted Mars missions are compared. The characteristics that are compared are Initial Mass in Low Earth Orbit (IMLEO), mission flexibility, and flight times. The propulsion systems being compared are both demonstrated and envisioned: Chemical (or Cryogenic), Nuclear Thermal Rocket (NTR) solid core, NTR gas core, Nuclear Electric Propulsion (NEP), and a mirror fusion space propulsion system. The proposed magnetic mirror fusion reactor, known as the Mirror Fusion Propulsion System (MFPS), is described. The description is an overview of a design study that was conducted to convert a mirror reactor experiment at Lawrence Livermore National Lab (LLNL) into a viable space propulsion system. Design principles geared towards minimizing mass and maximizing power available for thrust are identified and applied to the LLNL reactor design, resulting in the MFPS. The MFPS' design evolution, reactor and fuel choices, and system configuration are described. Results of the performance comparison shows that the MFPS minimizes flight time to 60 to 90 days for flights to Mars while allowing continuous return-home capability while at Mars. Total MFPS IMLEO including propellant and payloads is kept to about 1,000 metric tons.

Operationally Efficient Propulsion System Study (OEPSS) data book. Volume 4: OEPSS design concepts

NASA Technical Reports Server (NTRS)

Wong, George S.; Ziese, James M.; Farhangi, Shahram

1990-01-01

This study was initiated to identify operations problems and cost drivers for current propulsion systems and to identify technology and design approaches to increase the operational efficiency and reduce operations costs for future propulsion systems. To provide readily usable data for the Advanced Launch System (ALS) program, the results of the OEPSS study have been organized into a series of OEPSS Data Books. This volume describes three propulsion concepts that will simplify the propulsion system design and significantly reduce operational requirements. The concepts include: (1) a fully integrated, booster propulsion module concept for the ALS that avoids the complex system created by using autonomous engines with numerous artificial interfaces; (2) an LOX tank aft concept which avoids potentially dangerous geysering in long LOX propellant lines; and (3) an air augmented, rocket engine nozzle afterburning propulsion concept that will significantly reduce LOX propellant requirements, reduce vehicle size and simplify ground operations and ground support equipment and facilities.

Visions of the Future: Hybrid Electric Aircraft Propulsion

NASA Technical Reports Server (NTRS)

Bowman, Cheryl L.

2016-01-01

The National Aeronautics and Space Administration (NASA) is investing continually in improving civil aviation. Hybridization of aircraft propulsion is one aspect of a technology suite which will transform future aircraft. In this context, hybrid propulsion is considered a combination of traditional gas turbine propulsion and electric drive enabled propulsion. This technology suite includes elements of propulsion and airframe integration, parallel hybrid shaft power, turbo-electric generation, electric drive systems, component development, materials development and system integration at multiple levels.

Electric Propulsion Requirements and Mission Analysis Under NASA's In-Space Propulsion Technology Project

NASA Technical Reports Server (NTRS)

Dudzinski, Leonard a.; Pencil, Eric J.; Dankanich, John W.

2007-01-01

The In-Space Propulsion Technology Project (ISPT) is currently NASA's sole investment in electric propulsion technologies. This project is managed at NASA Glenn Research Center (GRC) for the NASA Headquarters Science Mission Directorate (SMD). The objective of the electric propulsion project area is to develop near-term and midterm electric propulsion technologies to enhance or enable future NASA science missions while minimizing risk and cost to the end user. Systems analysis activities sponsored by ISPT seek to identify future mission applications in order to quantify mission requirements, as well as develop analytical capability in order to facilitate greater understanding and application of electric propulsion and other propulsion technologies in the ISPT portfolio. These analyses guide technology investments by informing decisions and defining metrics for technology development to meet identified mission requirements. This paper discusses the missions currently being studied for electric propulsion by the ISPT project, and presents the results of recent electric propulsion (EP) mission trades. Recent ISPT systems analysis activities include: an initiative to standardize life qualification methods for various electric propulsion systems in order to retire perceived risk to proposed EP missions; mission analysis to identify EP requirements from Discovery, New Frontiers, and Flagship classes of missions; and an evaluation of system requirements for radioisotope-powered electric propulsion. Progress and early results of these activities is discussed where available.

Communications satellite business ventures - Measuring the impact of technology programmes and related policies

NASA Technical Reports Server (NTRS)

Greenberg, J. S.

1986-01-01

An economic evaluation and planning procedure which assesses the effects of various policies on fixed satellite business ventures is described. The procedure is based on a stochastic financial simulation model, the Domsat II, which evaluates spacecraft reliability, market performance, and cost uncertainties. The application of the Domsat II model to the assessment of NASA's ion thrusters for on-orbit propulsion and GaAs solar cell technology is discussed. The effects of insurance rates and the self-insurance option on the financial performance of communication satellite business ventures are investigated. The selection of a transportation system for placing the satellites into GEO is analyzed.

JT9D engine diagnostics. Task 2: Feasibility study of measuring in-service flight loads. [747 aircraft performance

NASA Technical Reports Server (NTRS)

Kafka, P. G.; Skibo, M. A.; White, J. L.

1977-01-01

The feasibility of measuring JT9D propulsion system flight inertia loads on a 747 airplane is studied. Flight loads background is discussed including the current status of 747/JT9D loads knowledge. An instrumentation and test plan is formulated for an airline-owned in-service airplane and the Boeing-owned RA001 test airplane. Technical and cost comparisons are made between these two options. An overall technical feasibility evaluation is made and a cost summary presented. Conclusions and recommendations are presented in regard to using existing inertia loads data versus conducting a flight test to measure inertia loads.

The Ion Propulsion System for the Solar Electric Propulsion Technology Demonstration Mission

NASA Technical Reports Server (NTRS)

Herman, Daniel A.; Santiago, Walter; Kamhawi, Hani; Polk, James E.; Snyder, John Steven; Hofer, Richard R.; Parker, J. Morgan

2015-01-01

The Asteroid Redirect Robotic Mission is a candidate Solar Electric Propulsion Technology Demonstration Mission whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a companion human-crewed mission. The ion propulsion system must be capable of operating over an 8-year time period and processing up to 10,000 kg of xenon propellant. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as a critical part of an affordable, beyond-low-Earth-orbit, manned-exploration architecture. Under the NASA Space Technology Mission Directorate the critical electric propulsion and solar array technologies are being developed. The ion propulsion system being co-developed by the NASA Glenn Research Center and the Jet Propulsion Laboratory for the Asteroid Redirect Vehicle is based on the NASA-developed 12.5 kW Hall Effect Rocket with Magnetic Shielding (HERMeS0 thruster and power processing technologies. This paper presents the conceptual design for the ion propulsion system, the status of the NASA in-house thruster and power processing activity, and an update on flight hardware.

Simulation model of the integrated flight/propulsion control system, displays, and propulsion system for ASTOVL lift-fan aircraft

NASA Technical Reports Server (NTRS)

Chung, W. Y. William; Borchers, Paul F.; Franklin, James A.

1995-01-01

A simulation model has been developed for use in piloted evaluations of takeoff, transition, hover, and landing characteristics of an advanced, short takeoff, vertical landing lift fan fighter aircraft. The flight/propulsion control system includes modes for several response types which are coupled to the aircraft's aerodynamic and propulsion system effectors through a control selector tailored to the lift fan propulsion system. Head-up display modes for approach and hover, tailored to their corresponding control modes are provided in the simulation. Propulsion system components modeled include a remote lift and a lift/cruise engine. Their static performance and dynamic response are represented by the model. A separate report describes the subsonic, power-off aerodynamics and jet induced aerodynamics in hover and forward flight, including ground effects.

Development Status of High-Thrust Density Electrostatic Engines

NASA Technical Reports Server (NTRS)

Patterson, Michael J.; Haag, Thomas W.; Foster, John E.; Young, Jason A.; Crofton, Mark W.

2017-01-01

Ion thruster technology offers the highest performance and efficiency of any mature electric propulsion thruster. It has by far the highest demonstrated total impulse of any technology option, demonstrated at input power levels appropriate for primary propulsion. It has also been successfully implemented for primary propulsion in both geocentric and heliocentric environments, with excellent ground/in-space correlation of both its performance and life. Based on these attributes there is compelling reasoning to continue the development of this technology: it is a leading candidate for high power applications; and it provides risk reduction for as-yet unproven alternatives. As such it is important that the operational limitations of ion thruster technology be critically examined and in particular for its application to primary propulsion its capabilities relative to thrust the density and thrust-to-power ratio be understood. This publication briefly addresses some of the considerations relative to achieving high thrust density and maximizing thrust-to-power ratio with ion thruster technology, and discusses the status of development work in this area being executed under a collaborative effort among NASA Glenn Research Center, the Aerospace Corporation, and the University of Michigan.

A review of electric propulsion systems and mission applications

NASA Technical Reports Server (NTRS)

Vondra, R.; Nock, K.; Jones, R.

1984-01-01

The satisfaction of growing demands for access to space resources will require new developments related to advanced propulsion and power technologies. A key technology in this context is concerned with the utilization of electric propulsion. A brief review of the current state of development of electric propulsion systems on an international basis is provided, taking into account advances in the USSR, the U.S., Japan, West Germany, China and Brazil. The present investigation, however, is mainly concerned with the U.S. program. The three basic types of electric thrusters are considered along with the intrinsic differences between chemical and electric propulsion, the resistojet, the augmented hydrazine thruster, the arcjet, the ion auxiliary propulsion system flight test, the pulsed plasma thruster, magnetoplasmadynamic propulsion, a pulsed inductive thruster, and rail accelerators. Attention is also given to the applications of electric propulsion.

New Propulsion Technologies For Exploration of the Solar System and Beyond

NASA Technical Reports Server (NTRS)

Johnson, Les; Cook, Stephen (Technical Monitor)

2001-01-01

In order to implement the ambitious science and exploration missions planned over the next several decades, improvements in in-space transportation and propulsion technologies must be achieved. For robotic exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs. Future missions will require 2 to 3 times more total change in velocity over their mission lives than the NASA Solar Electric Technology Application Readiness (NSTAR) demonstration on the Deep Space 1 mission. Rendezvous and return missions will require similar investments in in-space propulsion systems. New opportunities to explore beyond the outer planets and to the stars will require unparalleled technology advancement and innovation. The Advanced Space Transportation Program (ASTP) is investing in technologies to achieve a factor of 10 reduction in the cost of Earth orbital transportation and a factor of 2 reduction in propulsion system mass and travel time for planetary missions within the next 15 years. Since more than 70% of projected launches over the next 10 years will require propulsion systems capable of attaining destinations beyond Low Earth Orbit, investment in in-space technologies will benefit a large percentage of future missions. The ASTP technology portfolio includes many advanced propulsion systems. From the next generation ion propulsion system operating in the 5 - 10 kW range, to fission-powered multi-kilowatt systems, substantial advances in spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals use the environment of space itself for energy and propulsion and are generically called, "propellantless" because they do not require on-board fuel to achieve thrust. An overview of the state-of-the-art in propellantless propulsion technologies such as solar and plasma sails, electrodynamic and momentum transfer tethers, and aeroassist and aerocapture will also be described. Results of recent earth-based technology demonstrations and space tests for many of these new propulsion technologies will be discussed.

NASA In-Space Propulsion Technologies and Their Infusion Potential

NASA Technical Reports Server (NTRS)

Anderson, David J.; Pencil,Eric J.; Peterson, Todd; Vento, Daniel; Munk, Michelle M.; Glaab, Louis J.; Dankanich, John W.

2012-01-01

The In-Space Propulsion Technology (ISPT) program has been developing in-space propulsion technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (Electric and Chemical), Entry Vehicle Technologies (Aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance; 2) NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells; and aerothermal effect models. Two component technologies that will be ready for flight infusion in the near future will be Advanced Xenon Flow Control System, and ultra-lightweight propellant tank technologies. Future focuses for ISPT are sample return missions and other spacecraft bus technologies like: 1) Mars Ascent Vehicles (MAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) for sample return missions; and 3) electric propulsion for sample return and low cost missions. These technologies are more vehicle-focused, and present a different set of technology infusion challenges. While the Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion technologies to a wide variety of mission concepts. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.

NASA's In-Space Propulsion Technology Program: Overview and Status

NASA Technical Reports Server (NTRS)

Johnson, Les; Alexander, Leslie; Baggett, Randy; Bonometti, Joe; Herrmann, Melody; James, Bonnie; Montgomery, Sandy

2004-01-01

NASA's In-Space Propulsion Technology Program is investing in technologies that have the potential to revolutionize the robotic exploration of deep space. For robotic exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs and, in some cases, enable missions previously considered impossible. Continued reliance on conventional chemical propulsion alone will not enable the robust exploration of deep space - the maximum theoretical efficiencies have almost been reached and they are insufficient to meet needs for many ambitious science missions currently being considered. The In-Space Propulsion Technology Program s technology portfolio includes many advanced propulsion systems. From the next generation ion propulsion system operating in the 5 - 10 kW range, to advanced cryogenic propulsion, substantial advances in spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals use the environment of space itself for energy and propulsion and are generically called, 'propellantless' because they do not require onboard fuel to achieve thrust. Propellantless propulsion technologies include scientific innovations such as solar sails, electrodynamic and momentum transfer tethers, aeroassist, and aerocapture. This paper will provide an overview of both propellantless and propellant-based advanced propulsion technologies, and NASA s plans for advancing them as part of the $60M per year In-Space Propulsion Technology Program.

Additive Manufacturing a Liquid Hydrogen Rocket Engine

NASA Technical Reports Server (NTRS)

Jones, Carl P.; Robertson, Elizabeth H.; Koelbl, Mary Beth; Singer, Chris

2016-01-01

Space Propulsion is a 5 day event being held from 2nd May to the 6th May 2016 at the Rome Marriott Park Hotel in Rome, Italy. This event showcases products like Propulsion sub-systems and components, Production and manufacturing issues, Liquid, Solid, Hybrid and Air-breathing Propulsion Systems for Launcher and Upper Stages, Overview of current programmes, AIV issues and tools, Flight testing and experience, Technology building blocks for Future Space Transportation Propulsion Systems : Launchers, Exploration platforms & Space Tourism, Green Propulsion for Space Transportation, New propellants, Rocket propulsion & global environment, Cost related aspects of Space Transportation propulsion, Modelling, Pressure-Thrust oscillations issues, Impact of new requirements and regulations on design etc. in the Automotive, Manufacturing, Fabrication, Repair & Maintenance industries.

Economic effects of propulsion system technology on existing and future transport aircraft

NASA Technical Reports Server (NTRS)

Sallee, G. P.

1974-01-01

The results of an airline study of the economic effects of propulsion system technology on current and future transport aircraft are presented. This report represents the results of a detailed study of propulsion system operating economics. The study has four major parts: (1) a detailed analysis of current propulsion system maintenance with respect to the material and labor costs encountered versus years in service and the design characteristics of the major elements of the propulsion system of the B707, b727, and B747. (2) an analysis of the economic impact of a future representative 1979 propulsion system is presented with emphasis on depreciation of investment, fuel costs and maintenance costs developed on the basis of the analysis of the historical trends observed. (3) recommendations concerning improved methods of forecasting the maintenance cost of future propulsion systems are presented. A detailed method based on the summation of the projected labor and material repair costs for each major engine module and its installation along with a shorter form suitable for quick, less detailed analysis are presented, and (4) recommendations concerning areas where additional technology is needed to improve the economics of future commercial propulsion systems are presented along with the suggested economic benefits available from such advanced technology efforts.

Real-time fault diagnosis for propulsion systems

NASA Technical Reports Server (NTRS)

Merrill, Walter C.; Guo, Ten-Huei; Delaat, John C.; Duyar, Ahmet

1991-01-01

Current research toward real time fault diagnosis for propulsion systems at NASA-Lewis is described. The research is being applied to both air breathing and rocket propulsion systems. Topics include fault detection methods including neural networks, system modeling, and real time implementations.

Operationally Efficient Propulsion System Study (OEPSS) data book. Volume 3: Operations technology

NASA Technical Reports Server (NTRS)

Vilja, John O.

1990-01-01

The study was initiated to identify operational problems and cost drivers for current propulsion systems and to identify technology and design approaches to increase the operational efficiency and reduce operations costs for future propulsion systems. To provide readily usable data for the Advanced Launch System (ALS) program, the results of the OEPSS study were organized into a series of OEPSS Data Books. This volume describes operations technologies that will enhance operational efficiency of propulsion systems. A total of 15 operations technologies were identified that will eliminate or mitigate operations problems described in Volume 2. A recommended development plan is presented for eight promising technologies that will simplify the propulsion system and reduce operational requirements.

National space transportation systems planning

NASA Technical Reports Server (NTRS)

Lucas, W. R.

1985-01-01

In the fall of 1984, the DOD and NASA had been asked to identify launch vehicle technologies which could be made available for use in 1995 to 2010. The results of the studies of the two groups were integrated, and a consumer report, dated December 1984, was forwarded to the President. Aspects of mission planning and analysis are discussed along with a combined mission model, future launch system requirements, a launch vehicle planning background, Shuttle derivative vehicle program options, payload modularization, launch vehicle technology implications, a new engine program for the mid-1990's. Future launch systems goals are to achieve an order of magnitude reduction in future launch cost and meet the lift requirements and launch rates. Attention is given to an advanced cryogenic engine, advanced LOX/hydrocarbon engine, advanced power systems, aerodynamics/flight mechanics, reentry/recovery systems, avionics/software, advanced manufacturing techniques, autonomous ground and mission operations, advanced structures/materials, and air breathing propulsion.

LO2/LH2 propulsion for outer planet orbiter spacecraft

NASA Technical Reports Server (NTRS)

Garrison, P. W.; Sigurdson, K. B.

1983-01-01

Galileo class orbiter missions (750-1500 kg) to the outer planets require a large postinjection delta-V for improved propulsion performance. The present investigation shows that a pump-fed low thrust LO2/LH2 propulsion system can provide a significantly larger net on-orbit mass for a given delta-V than a state-of-the-art earth storable, N2O4/monomethylhydrazine pressure-fed propulsion system. A description is given of a conceptual design for a LO2/LH2 pump-fed propulsion system developed for a Galileo class mission to the outer planets. Attention is given to spacecraft configuration, details regarding the propulsion system, the thermal control of the cryogenic propellants, and aspects of mission performance.

A revolutionary lunar space transportation system architecture using extraterrestrial LOX-augmented NTR propulsion

NASA Astrophysics Data System (ADS)

Borowski, Stanley K.; Corban, Robert R.; Culver, Donald W.; Bulman, Melvin J.; McIlwain, Mel C.

1994-08-01

The concept of a liquid oxygen (LOX)-augmented nuclear thermal rocket (NTR) engine is introduced, and its potential for revolutionizing lunar space transportation system (LTS) performance using extraterrestrial 'lunar-derived' liquid oxygen (LUNOX) is outlined. The LOX-augmented NTR (LANTR) represents the marriage of conventional liquid hydrogen (LH2)-cooled NTR and airbreathing engine technologies. The large divergent section of the NTR nozzle functions as an 'afterburner' into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the NTR's choked sonic throat: 'scramjet propulsion in reverse.' By varying the oxygen-to-fuel mixture ratio (MR), the LANTR concept can provide variable thrust and specific impulse (Isp) capability with a LH2-cooled NTR operating at relatively constant power output. For example, at a MR = 3, the thrust per engine can be increased by a factor of 2.75 while the Isp decreases by only 30 percent. With this thrust augmentation option, smaller, 'easier to develop' NTR's become more acceptable from a mission performance standpoint (e.g., earth escape gravity losses are reduced and perigee propulsion requirements are eliminated). Hydrogen mass and volume is also reduced resulting in smaller space vehicles. An evolutionary NTR-based lunar architecture requiring only Shuttle C and/or 'in-line' shuttle-derived launch vehicles (SDV's) would operate initially in an 'expandable mode' with NTR lunar transfer vehicles (LTV's) delivering 80 percent more payload on piloted missions than their LOX/LH2 chemical propulsion counterparts. With the establishment of LUNOX production facilities on the lunar surface and 'fuel/oxidizer' depot in low lunar orbit (LLO), monopropellant NTR's would be outfitted with an oxygen propellant module, feed system, and afterburner nozzle for 'bipropellant' operation. The LANTR cislunar LTV now transitions to a reusable mode with smaller vehicle and payload doubling benefits on each piloted round trip mission. As the initial lunar outposts grow to centralized bases and settlements with a substantial permanent human presence, a LANTR-powered shuttle capable of 36 to 24 hour 'one-way' trip times to the moon and back becomes possible with initial mass in low earth orbit (IMLEO) requirements of approximately 160 to 240 metric tons, respectively.

A Revolutionary Lunar Space Transportation System Architecture Using Extraterrestrial Lox-augmented NTR Propulsion

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Corban, Robert R.; Culver, Donald W.; Bulman, Melvin J.; Mcilwain, Mel C.

1994-01-01

The concept of a liquid oxygen (LOX)-augmented nuclear thermal rocket (NTR) engine is introduced, and its potential for revolutionizing lunar space transportation system (LTS) performance using extraterrestrial 'lunar-derived' liquid oxygen (LUNOX) is outlined. The LOX-augmented NTR (LANTR) represents the marriage of conventional liquid hydrogen (LH2)-cooled NTR and airbreathing engine technologies. The large divergent section of the NTR nozzle functions as an 'afterburner' into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the NTR's choked sonic throat: 'scramjet propulsion in reverse.' By varying the oxygen-to-fuel mixture ratio (MR), the LANTR concept can provide variable thrust and specific impulse (Isp) capability with a LH2-cooled NTR operating at relatively constant power output. For example, at a MR = 3, the thrust per engine can be increased by a factor of 2.75 while the Isp decreases by only 30 percent. With this thrust augmentation option, smaller, 'easier to develop' NTR's become more acceptable from a mission performance standpoint (e.g., earth escape gravity losses are reduced and perigee propulsion requirements are eliminated). Hydrogen mass and volume is also reduced resulting in smaller space vehicles. An evolutionary NTR-based lunar architecture requiring only Shuttle C and/or 'in-line' shuttle-derived launch vehicles (SDV's) would operate initially in an 'expandable mode' with NTR lunar transfer vehicles (LTV's) delivering 80 percent more payload on piloted missions than their LOX/LH2 chemical propulsion counterparts. With the establishment of LUNOX production facilities on the lunar surface and 'fuel/oxidizer' depot in low lunar orbit (LLO), monopropellant NTR's would be outfitted with an oxygen propellant module, feed system, and afterburner nozzle for 'bipropellant' operation. The LANTR cislunar LTV now transitions to a reusable mode with smaller vehicle and payload doubling benefits on each piloted round trip mission. As the initial lunar outposts grow to centralized bases and settlements with a substantial permanent human presence, a LANTR-powered shuttle capable of 36 to 24 hour 'one-way' trip times to the moon and back becomes possible with initial mass in low earth orbit (IMLEO) requirements of approximately 160 to 240 metric tons, respectively.

Comparative performance evaluation of advanced AC and DC EV propulsion systems

NASA Astrophysics Data System (ADS)

MacDowall, R. D.; Crumley, R. L.

Idaho National Engineering Laboratory (INEL) evaluates EV propulsion systems and components for the U.S. Department of Energy (DOE) Electric and Hybrid Vehicle (EHV) Program. In this study, experimental data were used to evaluate the relative performances of the benchmark Chrysler/GE ETV-1 DC and the Ford/GE First Generation Single-Shaft AC (ETX-I) propulsion systems. Tests were conducted on the INEL's chassis dynamometer using identical aerodynamic and rolling resistance road-load coefficients and vehicle test weights. The results allowed a direct comparison of selected efficiency and performance characteristics for the two propulsion system technologies. The ETX-I AC system exhibited slightly lower system efficiency during constant speed testing than the ETV-1 DC propulsion system.

The Nuclear Cryogenic Propulsion Stage

NASA Technical Reports Server (NTRS)

Houts, Michael G.; Kim, Tony; Emrich, William J.; Hickman, Robert R.; Broadway, Jeramie W.; Gerrish, Harold P.; Doughty, Glen; Belvin, Anthony; Borowski, Stanley K.; Scott, John

2014-01-01

The fundamental capability of Nuclear Thermal Propulsion (NTP) is game changing for space exploration. A first generation Nuclear Cryogenic Propulsion Stage (NCPS) based on NTP could provide high thrust at a specific impulse above 900 s, roughly double that of state of the art chemical engines. Characteristics of fission and NTP indicate that useful first generation systems will provide a foundation for future systems with extremely high performance. The role of the NCPS in the development of advanced nuclear propulsion systems could be analogous to the role of the DC-3 in the development of advanced aviation. Progress made under the NCPS project could help enable both advanced NTP and advanced Nuclear Electric Propulsion (NEP). Nuclear propulsion can be affordable and viable compared to other propulsion systems and must overcome a biased public fear due to hyper-environmentalism and a false perception of radiation and explosion risk.

Implementation of an Online Database for Chemical Propulsion Systems

NASA Technical Reports Server (NTRS)

David B. Owen, II; McRight, Patrick S.; Cardiff, Eric H.

2009-01-01

The Johns Hopkins University, Chemical Propulsion Information Analysis Center (CPIAC) has been working closely with NASA Goddard Space Flight Center (GSFC); NASA Marshall Space Flight Center (MSFC); the University of Alabama at Huntsville (UAH); The Johns Hopkins University, Applied Physics Laboratory (APL); and NASA Jet Propulsion Laboratory (JPL) to capture satellite and spacecraft propulsion system information for an online database tool. The Spacecraft Chemical Propulsion Database (SCPD) is a new online central repository containing general and detailed system and component information on a variety of spacecraft propulsion systems. This paper only uses data that have been approved for public release with unlimited distribution. The data, supporting documentation, and ability to produce reports on demand, enable a researcher using SCPD to compare spacecraft easily, generate information for trade studies and mass estimates, and learn from the experiences of others through what has already been done. This paper outlines the layout and advantages of SCPD, including a simple example application with a few chemical propulsion systems from various NASA spacecraft.

The Propulsion Center at MSFC

NASA Technical Reports Server (NTRS)

Gerrish, Harold; Schmidt, George R. (Technical Monitor)

2000-01-01

The Propulsion Research Center at MSFC serves as a national resource for research of advanced, revolutionary propulsion technologies. Our mission is to move the nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft like access to earth-orbit, rapid travel throughout the solar system, and exploration of interstellar space. Current efforts cover a wide range of exciting areas, including high-energy plasma thrusters, advanced fission and fusion engines, antimatter propulsion systems, beamed energy rockets and sails, and fundamental motive physics. Activities involve concept investigation, proof-of-concept demonstration, and breadboard validation of new propulsion systems. The Propulsion Research Center at MSFC provides an environment where NASA, national laboratories, universities, and industry researchers can pool their skills together to perform landmark propulsion achievements. We offer excellent educational opportunities to students and young researchers-fostering a wellspring of innovation that will revolutionize space transportation.

MW-Class Electric Propulsion System Designs

NASA Technical Reports Server (NTRS)

LaPointe, Michael R.; Oleson, Steven; Pencil, Eric; Mercer, Carolyn; Distefano, Salvador

2011-01-01

Electric propulsion systems are well developed and have been in commercial use for several years. Ion and Hall thrusters have propelled robotic spacecraft to encounters with asteroids, the Moon, and minor planetary bodies within the solar system, while higher power systems are being considered to support even more demanding future space science and exploration missions. Such missions may include orbit raising and station-keeping for large platforms, robotic and human missions to near earth asteroids, cargo transport for sustained lunar or Mars exploration, and at very high-power, fast piloted missions to Mars and the outer planets. The Advanced In-Space Propulsion Project, High Efficiency Space Power Systems Project, and High Power Electric Propulsion Demonstration Project were established within the NASA Exploration Technology Development and Demonstration Program to develop and advance the fundamental technologies required for these long-range, future exploration missions. Under the auspices of the High Efficiency Space Power Systems Project, and supported by the Advanced In-Space Propulsion and High Power Electric Propulsion Projects, the COMPASS design team at the NASA Glenn Research Center performed multiple parametric design analyses to determine solar and nuclear electric power technology requirements for representative 300-kW class and pulsed and steady-state MW-class electric propulsion systems. This paper describes the results of the MW-class electric power and propulsion design analysis. Starting with the representative MW-class vehicle configurations, and using design reference missions bounded by launch dates, several power system technology improvements were introduced into the parametric COMPASS simulations to determine the potential system level benefits such technologies might provide. Those technologies providing quantitative system level benefits were then assessed for technical feasibility, cost, and time to develop. Key assumptions and primary results of the COMPASS MW-class electric propulsion power system study are reported, and discussion is provided on how the analysis might be used to guide future technology investments as NASA moves to more capable high power in-space propulsion systems.

State-of-the-Art for Small Satellite Propulsion Systems

NASA Technical Reports Server (NTRS)

Parker, Khary I.

2016-01-01

SmallSats are a low cost access to space with an increasing need for propulsion systems. NASA, and other organizations, will be using SmallSats that require propulsion systems to: a) Conduct high quality near and far reaching on-orbit research and b) Perform technology demonstrations. Increasing call for high reliability and high performing for SmallSat components. Many SmallSat propulsion technologies are currently under development: a) Systems at various levels of maturity and b) Wide variety of systems for many mission applications.

Mars Conjunction Crewed Missions With a Reusable Hybrid Architecture

NASA Technical Reports Server (NTRS)

Merrill, Raymond G.; Strange, Nathan J.; Qu, Min; Hatten, Noble

2015-01-01

A new crew Mars architecture has been developed that provides many potential benefits for NASA-led human Mars moons and surface missions beginning in the 2030s or 2040s. By using both chemical and electric propulsion systems where they are most beneficial and maintaining as much orbital energy as possible, the Hybrid spaceship that carries crew round trip to Mars is pre-integrated before launch and can be delivered to orbit by a single launch. After check-out on the way to cis-lunar space, it is refueled and can travel round trip to Mars in less than 1100 days, with a minimum of 300 days in Mars vicinity (opportunity dependent). The entire spaceship is recaptured into cis-lunar space and can be reused. The spaceship consists of a habitat for 4 crew attached to the Hybrid propulsion stage which uses long duration electric and chemical in-space propulsion technologies that are in use today. The hybrid architecture's con-ops has no in-space assembly of the crew transfer vehicle and requires only rendezvous of crew in a highly elliptical Earth orbit for arrival at and departure from the spaceship. The crew transfer vehicle does not travel to Mars so it only needs be able to last in space for weeks and re-enter at lunar velocities. The spaceship can be refueled and resupplied for multiple trips to Mars (every other opportunity). The hybrid propulsion stage for crewed transits can also be utilized for cargo delivery to Mars every other opportunity in a reusable manner to pre-deploy infrastructure required for Mars vicinity operations. Finally, the Hybrid architecture provides evolution options for mitigating key long-duration space exploration risks, including crew microgravity and radiation exposure.

Development of the European Small Geostationary Satellite SGEO

NASA Astrophysics Data System (ADS)

Lübberstedt, H.; Schneider, A.; Schuff, H.; Miesner, Th.; Winkler, A.

2008-08-01

The SGEO product portfolio, ranging from Satellite platform delivery up to in-orbit delivery of a turnkey system including satellite and ground control station, is designed for applications ranging from TV Broadcast to multimedia applications, Internet access, mobile or fixed services in a wide range of frequency bands. Furthermore, Data Relay missions such as the European Data Relay Satellite (EDRS) as well as other institutional missions are targeted. Key design features of the SGEO platform are high flexibility and modularity in order to accommodate a very wide range of future missions, a short development time below two years and the objective to build the system based on ITAR free subsystems and components. The system will provide a long lifetime of up to 15 years in orbit operations with high reliability. SGEO is the first European satellite to perform all orbit control tasks solely by electrical propulsion (EP). This design provides high mass efficiency and the capability for direct injection into geostationary orbit without chemical propulsion (CP). Optionally, an Apogee Engine Module based on CP will provide the perigee raising manoeuvres in case of a launch into geostationary transfer orbit (GTO). This approach allows an ideal choice out of a wide range of launcher candidates in dependence of the required payload capacity. SGEO will offer to the market a versatile and high performance satellite system with low investment risk for the customer and a short development time. This paper provides an overview of the SGEO system key features and the current status of the SGEO programme.

In situ monitoring of atmospheric nitrous acid based on multi-pumping flow system and liquid waveguide capillary cell.

PubMed

Liu, Yuhan; Lu, Keding; Dong, Huabin; Li, Xin; Cheng, Peng; Zou, Qi; Wu, Yusheng; Liu, Xingang; Zhang, Yuanhang

2016-05-01

In the last four decades, various techniques including spectroscopic, wet chemical and mass spectrometric methods, have been developed and applied for the detection of ambient nitrous acid (HONO). We developed a HONO detection system based on long path photometry which consists of three independent modules i.e., sampling module, fluid propulsion module and detection module. In the propulsion module, solenoid pumps are applied. With solenoid pumps the pulsed flow can be computer controlled both in terms of pump stroke volume and pulse frequency, which enables the attainment of a very stable flow rate. In the detection module, a customized Liquid Waveguide Capillary Cell (LWCC) is used. The customized LWCC pre-sets the optical fiber in-coupling with the liquid wave guide, providing the option of fast startup and easy maintenance of the absorption photometry. In summer 2014, our system was deployed in a comprehensive campaign at a rural site in the North China Plain. More than one month of high quality HONO data spanning from the limit of detection to 5ppb were collected. Intercomparison of our system with another established system from Forschungszentrum Juelich is presented and discussed. In conclusion, our instrument achieved a detection limit of 10pptV within 2min and a measurement uncertainty of 7%, which is well suited for investigation of the HONO budget from urban to rural conditions in China. Copyright © 2016. Published by Elsevier B.V.

Directions in propulsion control

NASA Technical Reports Server (NTRS)

Lorenzo, Carl F.

1990-01-01

Discussed here is research at NASA Lewis in the area of propulsion controls as driven by trends in advanced aircraft. The objective of the Lewis program is to develop the technology for advanced reliable propulsion control systems and to integrate the propulsion control with the flight control for optimal full-system control.

Heat transfer in aerospace propulsion

NASA Technical Reports Server (NTRS)

Simoneau, Robert J.; Hendricks, Robert C.; Gladden, Herbert J.

1988-01-01

Presented is an overview of heat transfer related research in support of aerospace propulsion, particularly as seen from the perspective of the NASA Lewis Research Center. Aerospace propulsion is defined to cover the full spectrum from conventional aircraft power plants through the Aerospace Plane to space propulsion. The conventional subsonic/supersonic aircraft arena, whether commercial or military, relies on the turbine engine. A key characteristic of turbine engines is that they involve fundamentally unsteady flows which must be properly treated. Space propulsion is characterized by very demanding performance requirements which frequently push systems to their limits and demand tailored designs. The hypersonic flight propulsion systems are subject to severe heat loads and the engine and airframe are truly one entity. The impact of the special demands of each of these aerospace propulsion systems on heat transfer is explored.

Selection and Prioritization of Advanced Propulsion Technologies for Future Space Missions

NASA Technical Reports Server (NTRS)

Eberle, Bill; Farris, Bob; Johnson, Les; Jones, Jonathan; Kos, Larry; Woodcock, Gordon; Brady, Hugh J. (Technical Monitor)

2002-01-01

The exploration of our solar system will require spacecraft with much greater capability than spacecraft which have been launched in the past. This is particularly true for exploration of the outer planets. Outer planet exploration requires shorter trip times, increased payload mass, and ability to orbit or land on outer planets. Increased capability requires better propulsion systems, including increased specific impulse. Chemical propulsion systems are not capable of delivering the performance required for exploration of the solar system. Future propulsion systems will be applied to a wide variety of missions with a diverse set of mission requirements. Many candidate propulsion technologies have been proposed but NASA resources do not permit development of a] of them. Therefore, we need to rationally select a few propulsion technologies for advancement, for application to future space missions. An effort was initiated to select and prioritize candidate propulsion technologies for development investment. The results of the study identified Aerocapture, 5 - 10 KW Solar Electric Ion, and Nuclear Electric Propulsion as high priority technologies. Solar Sails, 100 Kw Solar Electric Hall Thrusters, Electric Propulsion, and Advanced Chemical were identified as medium priority technologies. Plasma sails, momentum exchange tethers, and low density solar sails were identified as high risk/high payoff technologies.

Powersail High Power Propulsion System Design Study

NASA Astrophysics Data System (ADS)

Gulczinski, Frank S., III

2000-11-01

A desire by the United States Air Force to exploit the space environment has led to a need for increased on-orbit electrical power availability. To enable this, the Air Force Research Laboratory Space Vehicles Directorate (AFRL/ VS) is developing Powersail: a two-phased program to demonstrate high power (100 kW to 1 MW) capability in space using a deployable, flexible solar array connected to the host spacecraft using a slack umbilical. The first phase will be a proof-of-concept demonstration at 50 kW, followed by the second phase, an operational system at full power. In support of this program, the AFRL propulsion Directorate's Spacecraft Propulsion Branch (AFRL/PRS ) at Edwards AFB has commissioned a design study of the Powersail High Power Propulsion System. The purpose of this study, the results of which are summarized in this paper, is to perform mission and design trades to identify potential full-power applications (both near-Earth and interplanetary) and the corresponding propulsion system requirements and design. The design study shall farther identify a suitable low power demonstration flight that maximizes risk reduction for the fully operational system. This propulsion system is expected to be threefold: (1) primary propulsion for moving the entire vehicle, (2) a propulsion unit that maintains the solar array position relative to the host spacecraft, and (3) control propulsion for maintaining proper orientation for the flexible solar array.

NASA In-Space Propulsion Technology Program: Overview and Update

NASA Technical Reports Server (NTRS)

Johnson, Les; Alexander, Leslie; Baggett, Randy M.; Bonometti, Joseph A.; Herrmann, Melody; James, Bonnie F.; Montgomery, Sandy E.

2004-01-01

NASA's In-Space Propulsion Technology Program is investing in technologies that have the potential to revolutionize the robotic exploration of deep space. For robotic exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs and, in some cases, enable missions previously considered impossible. Continued reliance on conventional chemical propulsion alone will not enable the robust exploration of deep space - the maximum theoretical efficiencies have almost been reached and they are insufficient to meet needs for many ambitious science missions currently being considered. The In-Space Propulsion Technology Program's technology portfolio includes many advanced propulsion systems. From the next-generation ion propulsion system operating in the 5- to 10-kW range to aerocapture and solar sails, substantial advances in - spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals use the environment of space itself for energy and propulsion and are generically called 'propellantless' because they do not require onboard fuel to achieve thrust. Propellantless propulsion technologies include scientific innovations such as solar sails, electrodynamic and momentum transfer.tethers, aeroassist and aerocapture. This paper will provide an overview of both propellantless and propellant-based advanced propulsion technologies, as well as NASA's plans for advancing them as part of the In-Space Propulsion Technology Program.

NASA's In-Space Propulsion Technology Program: Overview and Update

NASA Technical Reports Server (NTRS)

Johnson, Les; Alexander, Leslie; Baggett, Randy M.; Bonometti, Joseph A.; Herrmann, Melody; James, Bonnie F.; Montgomery, Sandy E.

2004-01-01

NASA's In-Space Propulsion Technology Program is investing in technologies that have the potential to revolutionize the robotic exploration of deep space. For robotic exploration and science missions, increased efficiencies of future propulsion systems are critical to reduce overall life-cycle costs and, in some cases, enable missions previously considered impossible. Continued reliance on conventional chemical propulsion alone will not enable the robust exploration of deep space - the maximum theoretical efficiencies have almost been reached and they are insufficient to meet needs for many ambitious science missions currently being considered. The In-Space Propulsion Technology Program s technology portfolio includes many advanced propulsion systems. From the next-generation ion propulsion system operating in the 5- to 10-kW range to aerocapture and solar sails, substantial advances in spacecraft propulsion performance are anticipated. Some of the most promising technologies for achieving these goals ase the environment of space itself for energy and propulsion and are generically called 'propellantless' because they do not require onboard fuel to achieve thrust. Propellantless propulsion technologies include scientific innovations such as solar sails, electrodynamic and momentum transfer tethers, aeroassist, and aerocapture. This paper will provide an overview of both propellantless and propellant-based advanced propulsion technologies, as well as NASA s plans for advancing them as part of the In-Space Propulsion Technology Program.

The need for expanded exploration of matter-antimatter annihilation for propulsion application

NASA Technical Reports Server (NTRS)

Massier, P. F.

1982-01-01

The use of matter-antimatter annihilation as a propulsion application for interstellar travel is discussed. The physical basis for the superior energy release in such a system is summarized, and the problems associated with antimatter production, collection and storage are assessed. Advances in devising a workable propulsion system are reported, and the parameters of an antimatter propulsion system are described.

Survey of Advanced Booster Options for Potential Shuttle Derivative Vehicles

NASA Technical Reports Server (NTRS)

Sackheim, Robert L.; Ryan, Richard; Threet, Ed; Kennedy, James W. (Technical Monitor)

2001-01-01

A never-ending major goal for the Space Shuttle program is to continually improve flight safety, as long as this launch system remains in operational service. One of the options to improve system safety and to enhance vehicle performance as well, that has been seriously studied over the past several decades, is to replace the existing strap-on four segment solid rocket boosters (SRB's) with more capable units. A number of booster upgrade options have been studied in some detail, ranging from five segment solids through hybrids and a wide variety of liquid strap-ons (both pressure and pump fed with various propellants); all the way to a completely reusable liquid fly back booster (complete with air breathing engines for controlled landing and return). All of these possibilities appear to offer improvements in varying degrees; and each has their strengths and weaknesses from both programmatic and technical points of view. The most beneficial booster upgrade/design, if the shuttle program were to continue long enough to justify the required investment, would be an approach that greatly increased both vehicle and crew safety. This would be accomplished by increasing the minimum range/minimum altitude envelope that would readily allow abort to orbit (ATO), possibly even to zero/zero, and possibly reduce or eliminate the Return to Launch Site (RTLS) and even the Trans Atlantic Landing (TAL) abort mode requirements. This paper will briefly survey and discuss all of the various booster'upgrade options studied previously, and compare their relative attributes. The survey will explicitly discuss, in summary comparative form, options that include: five segment solids; several hybrid possibilities; pressure and/or pump-fed liquids using either LO2/kerosene, H2O/kerosene and LO2/J2, any of which could be either fully expendable, partly or fully reusable; and finally a fully reusable liquid fly back booster system, with a number of propellant and propulsion system options. Performance and configuration comparison illustrations and tables will be included to provide a comprehensive survey for the paper.

Hybrid rocket propulsion systems for outer planet exploration missions

NASA Astrophysics Data System (ADS)

Jens, Elizabeth T.; Cantwell, Brian J.; Hubbard, G. Scott

2016-11-01

Outer planet exploration missions require significant propulsive capability, particularly to achieve orbit insertion. Missions to explore the moons of outer planets place even more demanding requirements on propulsion systems, since they involve multiple large ΔV maneuvers. Hybrid rockets present a favorable alternative to conventional propulsion systems for many of these missions. They typically enjoy higher specific impulse than solids, can be throttled, stopped/restarted, and have more flexibility in their packaging configuration. Hybrids are more compact and easier to throttle than liquids and have similar performance levels. In order to investigate the suitability of these propulsion systems for exploration missions, this paper presents novel hybrid motor designs for two interplanetary missions. Hybrid propulsion systems for missions to Europa and Uranus are presented and compared to conventional in-space propulsion systems. The hybrid motor design for each of these missions is optimized across a range of parameters, including propellant selection, O/F ratio, nozzle area ratio, and chamber pressure. Details of the design process are described in order to provide guidance for researchers wishing to evaluate hybrid rocket motor designs for other missions and applications.

Overview of the Development of the Advanced Electric Propulsion System (AEPS)

NASA Technical Reports Server (NTRS)

Herman, Daniel; Tofil, Todd; Santiago, Walter; Kamhawi, Hani; Polk, James; Snyder, John Steven; Hofer, Richard; Picha, Frank; Schmidt, George

2017-01-01

NASA is committed to the demonstration and application of high-power solar electric propulsion to meet its future mission needs. It is continuing to develop the 14 kW Advanced Electric Propulsion System (AEPS) under a project that recently completed an Early Integrated System Test (EIST) and System Preliminary Design Review (PDR). In addition, NASA is pursuing external partnerships in order to demonstrate Solar Electric Propulsion (SEP) technology and the advantages of high-power electric propulsion-based spacecraft. The recent announcement of a Power and Propulsion Element (PPE) as the first major piece of an evolvable human architecture to Mars has replaced the Asteroid Redirect Robotic Mission (ARRM) as the most likely first application of the AEPS Hall thruster system. This high-power SEP capability, or an extensible derivative of it, has been recognized as a critical part of a new, affordable human exploration architecture for missions beyond-low-Earth-orbit. This paper presents the status of AEPS development activities, and describes how AEPS hardware will be integrated into the PPE ion propulsion system.

Factors Influencing Solar Electric Propulsion Vehicle Payload Delivery for Outer Planet Missions

NASA Technical Reports Server (NTRS)

Cupples, Michael; Green, Shaun; Coverstone, Victoria

2003-01-01

Systems analyses were performed for missions utilizing solar electric propulsion systems to deliver payloads to outer-planet destinations. A range of mission and systems factors and their affect on the delivery capability of the solar electric propulsion system was examined. The effect of varying the destination, the trip time, the launch vehicle, and gravity-assist boundary conditions was investigated. In addition, the affects of selecting propulsion system and power systems characteristics (including primary array power variation, number of thrusters, thruster throttling mode, and thruster Isp) on delivered payload was examined.

Development of technology for creating intelligent control systems for power plants and propulsion systems for marine robotic systems

NASA Astrophysics Data System (ADS)

Iakovleva, E. V.; Momot, B. A.

2017-10-01

The object of this study is to develop a power plant and an electric propulsion control system for autonomous remotely controlled vessels. The tasks of the study are as follows: to assess remotely controlled vessels usage reasonability, to define the requirements for this type of vessel navigation. In addition, the paper presents the analysis of technical diagnostics systems. The developed electric propulsion control systems for vessels should provide improved reliability and efficiency of the propulsion complex to ensure the profitability of remotely controlled vessels.

A Titan Explorer Mission Utilizing Solar Electric Propulsion and Chemical Propulsion Systems

NASA Technical Reports Server (NTRS)

Cupples, Michael; Coverstone, Vicki

2003-01-01

Mission and Systems analyses were performed for a Titan Explorer Mission scenario utilizing medium class launch vehicles, solar electric propulsion system (SEPS) for primary interplanetary propulsion, and chemical propulsion for capture at Titan. An examination of a range of system factors was performed to determine their affect on the payload delivery capability to Titan. The effect of varying the launch vehicle, solar array power, associated number of SEPS thrusters, chemical propellant combinations, tank liner thickness, and tank composite overwrap stress factor was investigated. This paper provides a parametric survey of the aforementioned set of system factors, delineating their affect on Titan payload delivery, as well as discussing aspects of planetary capture methodology.

Design of Electrical Systems for Rocket Propulsion Test Facilities at the John C. Stennis Space Center

NASA Technical Reports Server (NTRS)

Hughes, Mark S.; Davis, Dawn M.; Bakker, Henry J.; Jensen, Scott L.

2007-01-01

This viewgraph presentation reviews the design of the electrical systems that are required for the testing of rockets at the Rocket Propulsion Facility at NASA Stennis Space Center (NASA SSC). NASA/SSC s Mission in Rocket Propulsion Testing Is to Acquire Test Performance Data for Verification, Validation and Qualification of Propulsion Systems Hardware. These must be accurate reliable comprehensive and timely. Data acquisition in a rocket propulsion test environment is challenging: severe temporal transient dynamic environments, large thermal gradients, vacuum to 15 ksi pressure regimes SSC has developed and employs DAS, control systems and control systems and robust instrumentation that effectively satisfies these challenges.

The Altitude Wind Tunnel (AWT): A unique facility for propulsion system and adverse weather testing

NASA Technical Reports Server (NTRS)

Chamberlin, R.

1985-01-01

A need has arisen for a new wind tunnel facility with unique capabilities for testing propulsion systems and for conducting research in adverse weather conditions. New propulsion system concepts, new aircraft configurations with an unprecedented degree of propulsion system/aircraft integration, and requirements for aircraft operation in adverse weather dictate the need for a new test facility. Required capabilities include simulation of both altitude pressure and temperature, large size, full subsonic speed range, propulsion system operation, and weather simulation (i.e., icing, heavy rain). A cost effective rehabilitation of the NASA Lewis Research Center's Altitude Wind Tunnel (AWT) will provide a facility with all these capabilities.