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

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NASA NTRS1980

DOD low-thrust mission studies

The space transportation system (STS) will be the principal means of launching USAF spacecraft beginning in the 1980's. Since it is manned and reusable it provides new opportunities for unique approaches for cost effective utilization of its capabilities. The STS also places additional requirements and constraints on advanced spacecraft deployment systems that did not previously exist for expandable launch vehicles. To fully utilize these new capabilities designers must be prepared by having cost effective technologies available. Advanced propulsion technology that would provide flexibility, performance, and economic benefits to future Air Force missions was identified. Both electric and chemical propulsion systems are discussed. An LO2/LH2 stage with a torus LO2 tank and 500 lbf pumb fed engine is high on the list of propulsion technology.

Pipes, W. E.↗
NASA NTRS2019

SmallSat missions enabled by paired low-thrust hybrid rocket and low-power long-life Hall thruster

The capabilities of a SmallSat-class spacecraft targeting the outer solar system and using a combined chemical and electric propulsion system are explored. The development of compact hybrid rockets has enabled high-thrust engines to be packaged tightly enough to fit on cubesat and SmallSat spacecraft. These hybrid rockets provide 10’s-100 N of thrust depending on the propellant load & >300 s of specific impulse and have been demonstrated in both ambient and vacuum environments. Advancements in low-power long-life Hall thruster technologies have provided the potential for significantly greater propellant throughputs, enabling their use as a primary propulsion element on interplanetary spacecraft. In a recent characterization test campaign, the MaSMi-DM Hall thruster demonstrated power throttling from 150 – 1000 W with >1500 s of specific impulse available at >500 W and >40% total thrust efficiency available at >300 W; peak values of 1940 s and 53% were observed. A notional low-mass spacecraft employing a combined hybrid rocket and low-power electric propulsion system was designed and used for mission concept analysis targeting the outer solar system. Using an imposed wet mass limit of 400 kg, mission trajectories to Saturn and Uranus were generated. Orbit capture with >40% of the launch mass was shown to be possible at either target, with mission transfer times of 7.5 years and 13.5 years for Saturn and Uranus, respectively. Significant follow-on mission activities near Saturn (e.g. to Titan & Enceladus) were also possible by carrying extra propellant mass while remaining under the total wet mass limit.

Karp, Ashley C.↗
NASA NTRS1996

Inlet and Propulsion Integration of Scram Propelled Vehicles

The material to be presented in these two lectures begins with cycle considerations of the turbojet engine combined with a ramjet engine to provide thrust over the range of Mach 0 to 5. We will then examine in some detail the aerodynamic behavior that occurs in the inlet operating near the peak speed. Following that, we shall view a numerical simulation through a baseline scramjet engine, starting at the entrance to the inlet, proceeding into the combustor and through the nozzle. In the next segment, we examine a combined rocket and ramjet propulsion system. Analysis and test results will be examined with a view toward evaluation of the concept as a practical device. Two other inlets will then be reviewed: a Mach 12 inlet and a Mach 18 configuration. Finally, we close our lectures with a discussion of the Detonation Wave engine, and inspect the physical and chemical behavior obtained from numerical simulation. A few final remarks will be made regarding the application of CFD for hypersonic propulsion components.

Povinelli, Louis A.↗
NASA NTRSDate not supplied

Space Nuclear Propulsion for Deep Space Science Missions

The use of nuclear thermal propulsion (NTP) 1 and nuclear electric propulsion (NEP) 2 systems on deep space science missions to the outer planets and into the interstellar medium 3 can yield significant spacecraft system and mission performance benefits and improvements relative to the use of conventional chemical propulsion systems. Several recent and ongoing programs are developing the technologies and systems required to realize a near-term deep space nuclear propulsion capability. NTP provides improved propulsion efficiencies compared to chemical propulsion, while also providing substantial thrust. This combination of high thrust and increased specific impulse (I_sp) provides high acceleration and extended thrusting periods, enabling greatly reduced trip-times on certain types of missions compared to various propulsive alternatives. For examples, compared to a baseline mission using chemical propulsion, NTP-powered missions to Jupiter or Uranus could deliver approximately 2.4-3.6 times more payload (in the case of Jupiter, the payload delivery is significantly larger than the Juno spacecraft). In this comparison, the higher end of the payload advantage is obtained when the trip time is held equal for the NTP-powered vehicle and a vehicle using a chemical propulsion departure stage. NTP systems are presently under development by multiple government agencies. NASA’s Space Nuclear Propulsion (SNP) project aims to demonstrate a hydrogen-fed NTP engine at 900 s specific impulse (I_sp) and approximately 10-15 klb_f of thrust. DARPA’s Demonstration Rocket for Agile Cislunar Operations (DRACO) program is targeting a demonstration of an NTP system in the cislunar space between the Earth and the Moon. An appropriately phased development plan that applies the development of the reactor technology for an NTP engine in this performance class and leverages mature, existing liquid rocket component hardware provides a path to a lower cost propulsion system that can be realized on a shorter development schedule. NEP, with high Isp in the 2,000-8,000 s range, can also provide advantages over chemical propulsion, including a much greater payload delivery mass and the flexibility for planners to trade between delivered mass and a wider window of mission trajectory options. Electric propulsion (EP) systems have demonstrated great utility, performing notably on the Dawn mission to enable rendezvous and orbital insertion at two separate bodies, Vesta and Ceres. An NEP-powered vehicle would have a similar capability to visit multiple bodies, loitering at each before moving to the next. A 10 kW_e NEP system provides a power- rich environment on the spacecraft that is simply not possible using present radioisotope power systems, giving mission planners more scientific instrument and communication hardware options. Several programs and projects are presently developing NEP systems and subsystems in the 10 kW_e power range, leveraging past reactor work and recent nuclear power generation risk-reduction demonstration activities such as the Demonstration Using Flattop Fission (DUFF) and the Kilopower Reactor Using Stirling TechnologY (KRUSTY). The goal of the Air Force Research Laboratory’s Joint Energy Technology Supplying On-Orbit Nuclear Power (JETSON) program is an in-space demonstration vehicle that has a 10 kW_e -class fission power source. These past and present efforts can be combined with the ongoing development of 10 kW_e -class electric propulsion systems (notably the NEXT-C ion thruster or the Hall-effect thrusters for Power and Propulsion Element of the Lunar Gateway) to provide a pathway to a low-cost, reliable NEP system for deep space science application.

Kurt A. Polzin↗
NASA NTRSDate not supplied

Space Nuclear Propulsion for Deep Space Science Missions

The use of nuclear thermal propulsion (NTP) 1 and nuclear electric propulsion (NEP) 2 systems on deep space science missions to the outer planets and into the interstellar medium 3 can yield significant spacecraft system and mission performance benefits and improvements relative to the use of conventional chemical propulsion systems. Several recent and ongoing programs are developing the technologies and systems required to realize a near-term deep space nuclear propulsion capability. NTP provides improved propulsion efficiencies compared to chemical propulsion, while also providing substantial thrust. This combination of high thrust and increased specific impulse (I_sp) provides high acceleration and extended thrusting periods, enabling greatly reduced trip-times on certain types of missions compared to various propulsive alternatives. For examples, compared to a baseline mission using chemical propulsion, NTP-powered missions to Jupiter or Uranus could deliver approximately 2.4-3.6 times more payload (in the case of Jupiter, the payload delivery is significantly larger than the Juno spacecraft). In this comparison, the higher end of the payload advantage is obtained when the trip time is held equal for the NTP-powered vehicle and a vehicle using a chemical propulsion departure stage. NTP systems are presently under development by multiple government agencies. NASA’s Space Nuclear Propulsion (SNP) project aims to demonstrate a hydrogen-fed NTP engine at 900 s specific impulse (I_sp) and approximately 10-15 klb_f of thrust. DARPA’s Demonstration Rocket for Agile Cislunar Operations (DRACO) program is targeting a demonstration of an NTP system in the cislunar space between the Earth and the Moon. An appropriately phased development plan that applies the development of the reactor technology for an NTP engine in this performance class and leverages mature, existing liquid rocket component hardware provides a path to a lower cost propulsion system that can be realized on a shorter development schedule. NEP, with high Isp in the 2,000-8,000 s range, can also provide advantages over chemical propulsion, including a much greater payload delivery mass and the flexibility for planners to trade between delivered mass and a wider window of mission trajectory options. Electric propulsion (EP) systems have demonstrated great utility, performing notably on the Dawn mission to enable rendezvous and orbital insertion at two separate bodies, Vesta and Ceres. An NEP-powered vehicle would have a similar capability to visit multiple bodies, loitering at each before moving to the next. A 10 kW_e NEP system provides a power- rich environment on the spacecraft that is simply not possible using present radioisotope power systems, giving mission planners more scientific instrument and communication hardware options. Several programs and projects are presently developing NEP systems and subsystems in the 10 kW_e power range, leveraging past reactor work and recent nuclear power generation risk-reduction demonstration activities such as the Demonstration Using Flattop Fission (DUFF) and the Kilopower Reactor Using Stirling TechnologY (KRUSTY). The goal of the Air Force Research Laboratory’s Joint Energy Technology Supplying On-Orbit Nuclear Power (JETSON) program is an in-space demonstration vehicle that has a 10 kW_e -class fission power source. These past and present efforts can be combined with the ongoing development of 10 kW_e -class electric propulsion systems (notably the NEXT-C ion thruster or the Hall-effect thrusters for Power and Propulsion Element of the Lunar Gateway) to provide a pathway to a low-cost, reliable NEP system for deep space science application.

Kurt A. Polzin↗
NASA NTRS2016

Affordable Development and Demonstration of a Small Nuclear Thermal Rocket (NTR) Engine and Stage: How Small Is Big Enough?

The Nuclear Thermal Rocket (NTR) derives its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core. It generates high thrust and has a specific impulse potential of approximately 900 specific impulse - a 100 percent increase over today's best chemical rockets. The Nuclear Thermal Propulsion (NTP) project, funded by NASA's Advanced Exploration Systems (AES) program, includes five key task activities: (1) Recapture, demonstration, and validation of heritage graphite composite (GC) fuel (selected as the Lead Fuel option); (2) Engine Conceptual Design; (3) Operating Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable Development Strategy. During fiscal year (FY) 2014, a preliminary Design Development Test and Evaluation (DDT&E) plan and schedule for NTP development was outlined by the NASA Glenn Research Center (GRC), Department of Energy (DOE) and industry that involved significant system-level demonstration projects that included Ground Technology Demonstration (GTD) tests at the Nevada National Security Site (NNSS), followed by a Flight Technology Demonstration (FTD) mission. To reduce cost for the GTD tests and FTD mission, small NTR engines, in either the 7.5 or 16.5 kilopound-force thrust class, were considered. Both engine options used GC fuel and a common fuel element (FE) design. The small approximately 7.5 kilopound-force criticality-limited engine produces approximately157 thermal megawatts and its core is configured with parallel rows of hexagonal-shaped FEs and tie tubes (TTs) with a FE to TT ratio of approximately 1:1. The larger approximately 16.5 kilopound-force Small Nuclear Rocket Engine (SNRE), developed by Los Alamos National Laboratory (LANL) at the end of the Rover program, produces approximately 367 thermal megawatts and has a FE to TT ratio of approximately 2:1. Although both engines use a common 35-inch (approximately 89-centimeters) -long FE, the SNRE's larger diameter core contains approximately 300 more FEs needed to produce an additional 210 thermal megawatts of power. To reduce the cost of the FTD mission, a simple one-burn lunar flyby mission was considered to reduce the liquid hydrogen (LH2) propellant loading, the stage size and complexity. Use of existing and flight proven liquid rocket and stage hardware (e.g., from the RL10B-2 engine and Delta Cryogenic Second Stage) was also maximized to further aid affordability. This paper examines the pros and cons of using these two small engine options, including their potential to support future human exploration missions to the Moon, near Earth asteroids (NEA), and Mars, and recommends a preferred size. It also provides a preliminary assessment of the key activities, development options, and schedule required to affordably build, ground test and fly a small NTR engine and stage within a 10-year timeframe.

Spacecraft design↗
NASA NTRS1981

Low thrust chemical propulsion for orbit transfer of large space structures

For transporting Large Space Structures (LSS) from Low Earth Orbit (LEO) to Geosynchronous Equatorial Orbit (GEO), a variety of low thrust engine cycles have been evaluated for oxygen/hydrogen and oxygen/hydrocarbon propellants. The engine cycles included conventional propellant turbine drives (gas generator, expander, and staged-combustion cycles), turboalternator/electric motor pump drive, and fuel-cell/electric motor pump drive, as well as pressure-fed engines. The thrust chamber cooling limits and the engine cycle limits were established for a range of thrust levels. The candidate engine cycles were analyzed, screened, rated, and two engine cycle/configurations were selected for preliminary engine design. Preliminary engine designs for these two engines were formulated and engine design layouts prepared and parametric engine data generated.

Shoji, J. M.↗
NASA NTRS2012

Modular Growth NTR Space Transportation System for Future NASA Human Lunar, NEA and Mars Exploration Missions

The nuclear thermal rocket (NTR) is a proven, high thrust propulsion technology that has twice the specific impulse (I(sub sp) approx.900 s) 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 - everything required for affordable human missions beyond LEO. 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 IMLEO, versatile vehicle design, and growth potential. Furthermore, the NTR requires no large technology scale-ups since the smallest engine tested during the Rover program - the 25 klb(sub f) "Pewee" engine is sufficient for human Mars missions when used in a clustered engine configuration. The "Copernicus" crewed Mars transfer vehicle developed for DRA 5.0 was an expendable design sized for fast-conjunction, long surface stay Mars missions. It therefore has significant propellant capacity allowing a reusable "1-year" round trip human mission to a large, high energy near Earth asteroid (NEA) like Apophis in 2028. Using a "split mission" approach, Copernicus and its two key elements - a common propulsion stage and integrated "saddle truss" and LH2 drop tank assembly - configured as an Earth Return Vehicle / propellant tanker, can also support a short round trip (approx.18 month) / short orbital stay (60 days) Mars reconnaissance mission in the early 2030's before a landing is attempted. The same short stay orbital mission can be performed with an "all-up" vehicle by adding an "in-line" LH2 tank to Copernicus to supply the extra propellant needed for this higher energy, opposition-class mission. To transition to a reusable Mars architecture, Copernicus' saddle truss / drop tank assembly is replaced by an in-line tank and "star truss" assembly with paired modular drop tanks to further increase the vehicle's propellant capacity. Shorter "1-way" transit time fast-conjunction Mars missions are another possibility using this vehicle configuration but, as with reusability, increased launch mass is required. "Scaled down" versions of Copernicus (sized to a SLS lift capability of approx.70 t - 100 t) can be developed initially allowing reusable lunar cargo delivery and crewed landing missions, easy NEA missions (e.g., 2000 SG344 also in 2028) or an expendable mission to Apophis. Mission scenario descriptions, key vehicle features and operational characteristics are provided along with a brief discussion of NASA's current activities and its "pre-decisional" plans for future NTR development.

Borowski, Stanley K.↗
NASA NTRS1993

NASA Propulsion Engineering Research Center, volume 2

On 8-9 Sep. 1993, the Propulsion Engineering Research Center (PERC) at The Pennsylvania State University held its Fifth Annual Symposium. PERC was initiated in 1988 by a grant from the NASA Office of Aeronautics and Space Technology as a part of the University Space Engineering Research Center (USERC) program; the purpose of the USERC program is to replenish and enhance the capabilities of our Nation's engineering community to meet its future space technology needs. The Centers are designed to advance the state-of-the-art in key space-related engineering disciplines and to promote and support engineering education for the next generation of engineers for the national space program and related commercial space endeavors. Research on the following areas was initiated: liquid, solid, and hybrid chemical propulsion, nuclear propulsion, electrical propulsion, and advanced propulsion concepts.

Source record↗
NASA NTRS1998

Light-Weight Injector Technology for Cryogenic Mars Ascent Engines

Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for 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 launch vehicles. 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 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. 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.

Trihn, Huu Phuoc↗
NASA NTRS1999

Liquid Methane/Liquid Oxygen Injectors for Potential Future Mars Ascent Engines

Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for 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 launch vehicles. 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 would decrease. NASA/Johnson Space Center has initiated several concept studies 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 paper will address the results of the liquid methane/LOX injector study conducted at MSFC. A total of four impinging injector configurations were tested under combustion conditions in a modular combustor test article (MCTA), equipped with optically accessible windows. A series of forty hot-fire tests, which covered a wide range of engine operating conditions with the chamber pressure varied from 320 to 510 and the mixture ratio from 1.5 to 3.5, were performed. The test matrix also included a variation in the combustion chamber length for the purpose of investigating its effects on the combustion performance and stability.

Trinh, Huu Phuoc↗
NASA NTRS1994

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

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.

Borowski, Stanley K.↗
NASA NTRS2001

Applications of the Method of Space-Time Conservation Element and the Solution Element to Unsteady Chemically Reactive Flows

This document reports the conclusion and findings of our research activities for this grant. The goal of the project is the development and application of the method of Space-Time Conservation Element and Solution Element, or the CE/SE method, to simulate chemically reacting flows. The product of this project will be a high-fidelity, time-accurate flow solver analyzing unsteady flow fields advanced propulsion concepts, including the low-emission turbojet engine combustion and flow fields of the Pulse Detonation Engines (PDE). Based on the documents and computer software of the CE/SE method that we have received from the CE/SE working group at NASA Lewis, we have focused our research effort on addressing outstanding technical issues related to the extension of the CE/SE method for unsteady, chemically reactive flows. In particular, we have made progresses in the following three aspects: (1) Derivation of the governing equations for reacting flows; (2) Numerical treatments of stiff source terms; and (3) Detailed simulations of ZND detonation waves.

Yu, Sheng-Tao↗
NASA NTRS2012

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

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.

Borowski, Stanley↗
NASA NTRS1973

Chemical propulsion - The old and the new challenges

The historical background concerning the application of liquid propellant rockets is considered. Progress to date in chemical liquid propellant rocket engines can be summarized as an increase in performance through the use of more energetic propellant combinations and increased combustion pressure. New advances regarding liquid propellant rocket engines are related to the requirement for reusability in connection with the development of the Space Shuttle.

Mccarty, J. P.↗
NASA NTRS1998

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

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.

Trinh, Huu Ogyic↗
NASA NTRS1990

The Centauri project: Manned interstellar travel

The development of antimatter engines for spacecraft propulsion will allow man to expand to the nearest stellar neighbors such as the Alpha Centuri system. Compared to chemically powered rockets like the Apollo mission class which would take 50,000 years to reach the Centauri system, antimatter propulsion would reduce one way trip time to 30 years or less. The challenges encountered by manned interstellar travel are formidable. The spacecraft must be a combination of sublight speed transportation system and a traveling microplanet serving an expanding population. As the population expands from the initial 100 people to approximately 300, the terraformed asteroid, enclosed by a man-made shell will allow for expansion over its surface in the fashion of a small terrestrial town. All aspects of human life - birth; death; physical, emotional, and educational needs; and government and law must be met by the structure, systems, and institutions on-board.

Ciesla, Thomas M.↗
NASA NTRS1997

On-Board Chemical Propulsion

NASA Lewis Research Center's On-Board Propulsion program (OBP) is developing low-thrust chemical propulsion technologies for both satellite and vehicle reaction control applications. There is a vigorous international competition to develop new, highperformance bipropellant engines. High-leverage bipropellant systems are critical to both commercial competitiveness in the international communications market and to cost-effective mission design in government sectors. To significantly improve bipropellant engine performance, we must increase the thermal margin of the chamber materials. Iridium-coated rhenium (Ir/Re) engines, developed and demonstrated under OBP programs, can operate at temperatures well above the constraints of state-of-practice systems, providing a sufficient margin to maximize performance with the hypergolic propellants used in most satellite propulsion systems.

Schneider, Steven J.↗