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A top-level feasibility study was conducted that identified and characterized promising chemical propulsion system designs which use two or more of the following propellant combinations: LOX/H2, LOX/CH4, and LOX/CO. The engine systems examined emphasized the usage of common subsystem/component hardware where possible. In support of this study, numerous mission scenarios were characterized that used various combinations of Earth, lunar, and Mars propellants to establish engine system requirements to assess the promising engine system design concept examined, and to determine overall exploration leverage of such systems compared to state-of-the-art cryogenic (LOX/H2) propulsion systems. Initially in the study, critical propulsion system technologies were assessed. Candidate expander and gas generator cycle LOX/H2/CO, LOX/H2/CH4, and LOX/CO/CH4 engine system designs were parametrically evaluated. From this evaluation baseline, tripropellant Mars Transfer Vehicle (MTV) LOX cooled and bipropellant Lunar Excursion Vehicle (LEV) and Mars Excursion Vehicle (MEV) engine systems were identified. Representative tankage designs for a MTV were also investigated. Re-evaluation of the missions using the baseline engine design showed that in general the slightly lower performance, smaller, lower weight gas generator cycle-based engines required less overall mission Mars and in situ propellant production (ISPP) infrastructure support compared to the larger, heavier, higher performing expander cycle engine systems.

Performance characterization of a flight-type 1.4 kW arcjet system were conducted at the Rocket Research Company (RRC) in Redmond, WA, and at the NASA LeRC in Cleveland, OH. The objectives of these tests were as follows: to compare low-power arcjet performance at two different test facilities; to compare arcjet performance obtained with a 2:1 mixture of gaseous hydrogen and nitrogen and hydrazine; and to quantify the effects of test cell pressure on thruster operating characteristics. Performance and thruster temperature distributions were measured at thruster input power levels and propellant mass flow rates ranging from 1274 to 1370 W and from 3.2 x 10(exp -5) to 5.1 x 10(exp -5) kg/s, respectively. Specific impulses measured at the two facilities, at comparable test cell pressures, using gaseous hydrogen-nitrogen propellant mixtures agreed to within 1 percent over the range of operating conditions tested. The specific impulses measured using hydrazine propellant were higher than that for the cold hydrogen-nitrogen mixtures. Agreement between by hydrazine and gas mixture data was good, however, when the differences in propellant enthalpies at the thruster inlet were considered. Specific impulse showed a strong dependence on test facility pressure, and was 3 to 4 percent higher below 0.1 Pa than for test cell pressures above 5 Pa.

A proposed aerospace transporter for performing earth-to-orbit/return missions as well as transatmospheric and in-space high-energy missions is argued to be an effective and feasible alternative for development. The vehicles are based on the exploitation of 'synerjet' technologies that use airbreathing/rocket combined-cycle propulsion. The vehicle is shown to be capable of terrestrial-service intercontinental transglobal flight in an 'orbital cruise' mode as well as round trips from low orbits to high orbits and to the moon. The operational flexibility is linked to the development of synerjet propulsion, so directions are proposed for technology development and validation efforts. The development of the NASP X-30 by NASA is shown to be an important initial contribution to the R&D in this direction.

The characteristics of the Flowing Gas Radiation Receiver (FGRR), a device that absorbs solar radiation volumetrically in a gas to produce high temperatures for space propulsion and power applications, are analyzed using a two-dimensional axisymmetric numerical model of the flow and radiation fields within a diffusely reflecting channel. The results show that an FGRR system is capable of generating temperatures in excess of 3000 K with collection efficiencies of approximately 75 percent for a channel with a reflectivity of 0.9. For a collinear radiation source, outflow temperatures of 3193 and 3092 K were achieved for axial and radial flow inputs, respectively, with receiver efficiencies of 0.82 and 0.76.

A space propulsion system utilizing stored thermal energy is analyzed with emphasis on the preliminary design of the heat exchanger and on the general performance of engines incorporating this concept. A highenergy propellant is developed by heat transfer from a phase-change material (Si) to a low-density gas (H2) through a concentric cylinder heat exchanger. The phase-change material is heated by a radioisotope (PoGd) over a period of time which is significantly longer than the firing time. It is shown that such an engine can most efficiently operate just above and just below a propellant flow regime defined by gradual flow transition in the heat exchanger. Two designs are developed to illustrate each distinct operating scenario. A 10-N engine utilizing laminar flow is shown to be an attractive alternative to chemical and electric systems for the maneuvering of a vehicle in the vicinity of a larger space structure. A 450-N engine utilizing primarily turbulent heat transfer is shown to be an alternative to other systems for orbit transfer. Rapid cooling of the thermal storage material in the entrance region of the heat exchanger, which can be mitigated by varying the flow area or increasing the length of the engine, is a primary limitation on an engine utilizing thermal storage.

The nonnuclear subsystem technologies required for incorporating nuclear thermal propulsion (NTP) into space-exploration missions are discussed. Of particular interest to planned missions are such technologies as materials, instrumentation and controls, turbomachinery, CFD modeling, nozzle extension designs and models, and analyses of exhaust plumes. NASA studies are described and/or proposed for refractory metals and alloys, robotic NTP controls, and turbopump materials candidates. Alternative nozzle concepts such as aerospikes and truncated plugs are proposed, and numerical simulations are set forth for studying heavy molecules and the backstreaming of highly reactive free-radical hydrogen in the exhaust plume. The critical technologies described in the paper are central to the development of NTP, and NTP has the potential to facilitate a range of space exploration activities.

A new method of providing power to space vehicles consists of using ground-based lasers to beam power to photovoltaic receivers in space. This can be used as a power source for electrically propelled orbital transfer vehicles.

The Consortium for CFD applications in propulsion technology has been formed at NASA/Marshall Space Flight Center. The combustion devices technology team is one of the three teams that constitute the Consortium. While generally aiming to advance combustion devices technology for rocket propulsion, the team's efforts for the last 1 and 1/2 years have been focused on issues relating to the Space Transportation Main Engine (STME) nozzle. The nozzle design uses hydrogen-rich turbine exhaust to cool the wall in a film/dump scheme. This method of cooling presents challenges and associated risks for the nozzle designers and the engine/vehicle integrators. Within the nozzle itself, a key concern is the ability to effectively and efficiently film cool the wall. From the National Launch System vehicle base standpoint, there are concerns with dumping combustible gases at the nozzle exit and their potential adverse effects on the base thermal environment. The Combustion Team has developed and is implementing plans to use validated CFD tools to aid in risk mitigation for both areas.

Safety is the prime design requirement for nuclear thermal propulsion (NTP). It must be built in at the initiation of the design process. An understanding of safety concerns is fundamental to the development of nuclear rockets for manned missions to Mars and many other applications that will be enabled or greatly enhanced by the use of nuclear propulsion. To provide an understanding of the basic issues, a tutorial has been prepared. This tutorial covers a range of topics including safety requirements and approaches to meet these requirements, risk and safety analysis methodology, NERVA reliability and safety approach, and life cycle risk assessments.

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.

The theory and procedures for determining the characteristics of pressure oscillations in rocket engines with prescribed burning rate oscillations are presented. Analyses including radial and hub baffles and absorbers can be performed in one, two, or three dimensions. Pressure and velocity oscillations calculated using this procedure are presented for the SSME to show the influence of baffles and absorbers on the burning rate oscillations required to achieve neutral stability. Comparisons are made between the results obtained utilizing 1D, 2D, and 3D assumption with regards to capturing the physical phenomena of interest and computational requirements.

There are potential space exploration missions which may significantly benefit from the use of electric propulsion at power levels of hundreds of kilowatts. The applied magnetic field MPD thruster is potentially capable of efficient, high specific impulse operation in this power range. This paper describes current experimental and analytical efforts to further the development of such a thruster and presents the latest results. In particular, efforts to measure, simultaneously, the thrust developed by the archead and by the electromagnet, and to evaluate the effect of a diffuser on vacuum tank back pressure, are presented and discussed. It was found that with ammonia vapor as propellant, the vacuum tank pressure was reduced from 8 to 4.9 Pa at a power level of 80 kW. This pressure decrease is expected to become greater as the power and applied field are increased. Also, the development of a cathode/plasma interaction model for determining the heat loads to the cathode as functions of the various free stream plasma parameters is presented. This model is combined with a cathode thermal model in order to provide a complete and integrated picture of MPD thruster cathode operation. Several computational examples are used to illustrate the combined model.

In order to advance rocket propulsion technology, the Consortium for Computational Fluid Dynamics (CFD) Application in Propulsion Technology has been formed at Marshall Space Flight Center (MSFC). The Consortium consists of three Teams: the turbine stage team, the pump stage team (PST), and the combustion devices team. The PST has formulated and is implementing a plan for pump technology development whose end product will be validated CFD codes suitable for application to pump components, test data suitable for validating CFD codes, and advanced pump components optimized using CFD codes. The PST's work during the fall of 1991 and the winter and spring of 1992 is discussed in this paper. This work is highlighted by CFD analyses of an advanced impeller design and collection of laser two-focus velocimeter data for the Space Shuttle Main Engine High Pressure Fuel Pump impeller.

An initial investigation of the design considerations in clustering nuclear rocket engines for space transfer vehicles has been made. The clustering of both propulsion modules (which include start tanks) and nuclear rocket engines installed directly to a vehicle core tank appears to be feasible. Special provisions to shield opposite run tanks and the opposite side of a core tank - in the case of the boost pump concept - are required; the installation of a circumferential reactor side shield sector appears to provide an effective solution to this problem. While the time response to an engine-out event does not appear to be critical, the gimbal displacement required appears to be important. Since an installation of three engines offers a substantial reduction in gimbal requirements for engine-out and it may be possible to further enhance mission reliability with the greater number of engines, it is recommended that a cluster of four engines be considered.

A NASA workshop on rocket-based combined-cycle propulsion technologies is described emphasizing the development of a starting point for earth-to-orbit (ETO) rocket technologies. The tutorial is designed with attention given to the combined development of aeronautical airbreathing propulsion and space rocket propulsion. The format, agenda, and group deliberations for the tutorial are described, and group deliberations include: (1) mission and space transportation infrastructure; (2) vehicle-integrated propulsion systems; (3) development operations, facilities, and human resource needs; and (4) spaceflight fleet applications and operations. Although incomplete the workshop elevates the subject of combined-cycle hypersonic propulsion and develops a common set of priniciples regarding the development of these technologies.

Research aimed at developing a fundamental understanding of the plasma processes as applied to spacecraft propulsion is presented. Calorimetric, photographic, and spectrophotometric measurements based on the TM011 and TM012 modes in the resonance cavity have been performed. The efficiency of a thruster has been calculated using a theoretical model for predicting temperature, velocity, and species density within the propellant. It is concluded that the microwave electrothermal thruster is a viable alternative to electrode thrusters.

A 25-year historical review of the Space Electric Rocket Test 2 (SERT 2) mission is presented. The Agena launch vehicle; the SERT 2 spacecraft; and mission-peculiar spacecraft hardware, including two ion thruster systems, are described. The 3 1/2-year development period, from 1966 to 1970, that was needed to design, fabricate, and qualify the ion thruster system and the supporting spacecraft components, is documented. Major testing of two ion thruster systems and related auxiliary experiments that were conducted in space after the 3 Feb. 1970, launch are reviewed. Extended ion thruster restarts from 1973 to 1981 are reported, in addition to cross-neutralization tests. Tests of a reflector erosion experiment were continued in 1989 to 1991. The continuing performance of spacecraft subsystems, including the solar arrays, over the 1970-1991 period is summarized. Finally, the knowledge of thruster-spacecraft interactions learned from SERT 2 is listed.

The indigenous propellants of oxygen and carbon monoxide were studied for use in a Mars Ascent Vehicle (MAV). Both the oxygen and carbon monoxide were evaluated as turbine working fluids for a full expander engine cycle used in the MAV. Two oxygen working fluid engines and four carbon monoxide engines were investigated. The maximum hot-gas-side wall temperature was limited to either 445 K (800R) or 556 K (1000 R) over a range of working fluid mass flows. An engine thrust of 44.5 kN (1,000 lb(sub f)) was assumed. The engine characteristics of coolant inlet pressure, total engine length, specific impulse, pump efficiencies, and turbine power required were optimized. Of the six engines evaluated, the study showed that either type of working fluid was feasible for a MAV engine. Using carbon monoxide as the working fluid resulted in the best engine characteristics.

A small contingent of engineers at NASA LeRC pioneered the basic research on liquid propellants for rockets shortly after World War 2. Carried on through the 1950s, this work influenced the important early decisions made by Abe Silverstein when he took charge of the Office of Space Flight Programs for NASA. He strongly supported the development of liquid hydrogen as a propulsion fuel in the face of resistance from Wernher von Braun. Members of the LeRC staff played an important role in bringing liquid hydrogen technology to the point of reliability through their management of the Centaur Program. This paper demonstrates how the personality and engineering intuition of Abe Silverstein shaped the Centaur program and left a lasting imprint on the laboratory research tradition. Many of the current leaders of LeRC received their first hands-on engineering experience when they worked on the Centaur program in the 1960s.

This paper discusses alternative methods for enabling human explorers to achieve long range mobility on the surface of Mars. Vehicle types considered include ground rovers, winged and lighter-than air atmospheric vehicles, and suborbital ballistic vehicles making use of rocket propulsion. Energy sources considered for such vehicles include batteries, fuel cells, radioisotope generators, solar photovoltaics, chemical combustion engines, and nuclear reactors. It is found that the production of fuels and propellants out of indigenous Martian resources greatly enhances the potential for long range mobility, so much so that it may be considered the essential central element of any manned Mars exploration strategy.