Showing papers on "Spacecraft propulsion published in 1996"

Analysis of Hall-effect thrusters and ion engines for orbit transfer missions

Abstract: Analytical methods were combined with actual thruster data to create a model used to predict the performance of systems based on two types of electric propulsion thrusters, Hall-effect thrusters and ion engines, for several orbit transfer missions. Two missions were trip time constrained: a LEO-GEO transfer and a LEO constellation transfer. Hall thrusters were able to deliver greater payload due to their higher overall specific power. For the power limited orbit topping mission, the choice of thruster is dependent on the user’s need. Ion engines can deliver the greatest payload due to their higher specific impulse, but they do so at the cost of higher trip time. Study of reusable electric orbit transfer vehicle systems indicates that they can offer payload mass gains over chemical systems, but that these gains are less than those offered by other electric propulsion transfer scenarios due to the necessity of carrying propellant for return trips. Additionally, solar array degradation leads to increased trip time for subsequent reusable transfers. * Research Aerospace Engineer, Member AIAA ** Group Leader, USAF Electric Propulsion Lab, Member AIAA This paper is declared a work of the US Government and is not subject to copyright protection in the United States. INTRODUCTION: The US Air Force has recently completed several studies to investigate the potential advantages of advanced space propulsion for several orbit transfer scenarios. The first study investigated advanced propulsion concepts for expendable orbit transfer vehicles and concluded that the potential launch vehicle downsizing that resulted from the use of high specific impulse thrusters provided significant cost savings over base line chemical launch vehicle/upper stage systems. The second study looked at reusable advanced upper stages and preliminary indications are that while there remains the potential for launch vehicle downsizing, it is significantly reduced compared to expendable systems. This difference was largely due to the added propellant required to perform the round trip mission from low-earth orbit to geostationary orbit. Both studies pointed out advantages for advanced electric propulsion systems based on xenon propellant. The objective of this paper is to analyze the tradeoffs between Hall-effect thrusters and ion engines as a high power propulsion system for orbit transfer missions. Both the Hall-effect thruster and the gridded ion engine are classified as electrostatic thrusters and operate on heavy noble gases, primarily xenon. These electric propulsion devices are capable of specific impulses ranging from approximately 1500 to 4000 seconds, compared to chemical systems which typically operate in the range of 300 to 400 seconds. Electric propulsion is a type of rocket propulsion for space vehicles and satellites which utilizes electric and/or magnetic processes to accelerate a propellant at a much higher specific impulse than attainable using classical chemical propulsion. The concomitant reduction in required propellant mass results in increased payload mass capability. The method of analysis used in this study is based on the model developed by Messerole. It has been modified to reflect the most current information on thruster development levels and

Evaluation of Low Power Hall Thruster Propulsion

Abstract: Hall thruster systems based on the SPT-50 and the TAL D-38 were evaluated and mission studies were performed. The 0.3 kilowatt SPT-50 operated with a specific impulse of 1160 seconds and an efficiency of 0.32. The 0.8 kilowatt D-38 provided a specific impulse above 1700 seconds at an efficiency of 0.5. The D-38 system was shown to offer a 56 kilogram propulsion system mass savings over a 101 kilogram hydrazine monopropellant system designed to perform North-South station keeping maneuvers on board a 430 kilogram geostationary satellite. The SPIT-50 system offered a greater than 50% propulsion system mass reduction in comparison to the chemical system on board a 200 kilogram low Earth orbit spacecraft performing two orbit raises and drag makeup over two years. The performance characteristics of the SPF-50 were experimentally evaluated at a number of operating conditions. The ion current density distribution of this engine was measured. The performance and system mass benefits of advanced systems based on both engines were considered.

Launch vehicle and power level impacts on electric GEO insertion

Abstract: Solar Electric Propulsion (SEP) has been shown to increase net geosynchronous spacecraft mass when used for station keeping and final orbit insertion. The impact of launch vehicle selection and power level on the benefits of this approach were examined for 20 and 25 kW systems launched using the Ariane 5, Atlas IIAR, Long March, Proton, and Sea Launch vehicles. Two advanced on-board propulsion technologies, 5 kW ion and Hall thruster systems, were used to establish the relative merits of the technologies and launch vehicles. GaAs solar arrays were assumed. The analysis identifies the optimal starting orbits for the SEP orbit raising/plane changing while considering the impacts of radiation degradation in the Van Allen belts, shading, power degradation, and oblateness. This use of SEP to provide part of the orbit insertion results in net mass increases of 15 - 38% and 18 - 46% for one to two month trip times, respectively, over just using SEP for 15 years of north/south station keeping. SEP technology was shown to have a greater impact on net masses of launch vehicles with higher launch latitudes when avoidance of solar array and payload degradation is desired. This greater impact of SEP could help reduce the plane changing disadvantage of high latitude launch sites. Comparison with results for 10 and 15 kW systems show clear benefits of incremental increases in SEP power level, suggesting that an evolutionary approach to high power SEP for geosynchronous spacecraft is possible.

Electrodynamic Tethers for Spacecraft Propulsion

Abstract: Relatively short electrodynamic tethers can use solar power to 'push' against a planetary magnetic field to achieve propulsion without the expenditure of propellant. The groundwork has been laid for this type of propulsion. NASA began developing tether technology for space applications in the 1960's. Important recent milestones include retrieval of a tether in space (TSS-1, 1992), successful deployment of a 20-km-long tether in space (SEDS-1, 1993), and operation of an electrodynamic tether with tether current driven in both directions-power and thrust modes (PMG, 1993). The planned Propulsive Small Expendable Deployer System (ProSEDS) experiment will demonstrate electrodynamic tether thrust during its flight in early 2000. ProSEDS will use the flight-proven Small Expendable Deployer System (SEDS) to deploy a 5 km bare copper tether from a Delta II upper stage to achieve approximately 0.4 N drag thrust, thus deorbiting the stage. The experiment will use a predominantly 'bare' tether for current collection in lieu of the endmass collector and insulated tether approach used on previous missions. Theory and ground-based plasma chamber testing indicate that the bare tether is a highly-efficient current collector. The flight experiment is a precursor to utilization of the technology on the International Space Station for reboost application and the more ambitious electrodynamic tether upper stage demonstration mission which will be capable of orbit raising, lowering and inclination changes - all using electrodynamic thrust. In addition, the use of this type of propulsion may be attractive for future missions at Jupiter and any other planetary body with a magnetosphere.

Apparatus and method for simulating rocket-to-ramjet transition in a propulsion system

Abstract: A test facility for realistically simulating selected operating conditions for a multi-mode aircraft propulsion system employs bypass ducting to rapidly introduce heated compressed air to a ramjet engine immediately following rocket booster operation. Heated compressed airflow from a suitable supply is initially utilized to ensure that incidental ducting, control elements, and instrumentation are stabilized in temperature, by bypassing the totality of the heated compressed airflow from a location close to the multi-mode propulsion system out to ambient atmosphere. Upon stable conditions being realized, the rocket thrust component of the propulsion system is actuated and selected physical parameters e.g., pressure, in the multi-mode propulsion system are measured. Upon certain criteria being fulfilled, inlet port covers are opened, allowing heated, compressed air to pass through to the ramjet engine, as well as continuing to bypass a reduced amount of air. A mixture of fuel and a portion of the available heated compressed air is then burned in a ramburner component of the propulsion system. The thrust generated by the propulsion system in its rocket propulsion mode, rocket-to-ramjet transition, and ramjet propulsion mode is measured and analyzed in conjunction with other measured parameters.

Method and apparatus for coupling space vehicles

Abstract: A first space vehicle is attached with a second space vehicle which includes a rocket propulsion nozzle having a combustion chamber upstream from the nozzle throat. Apparatus 10 includes an elongate grappling arm 12 extending from the first space vehicle, a pair of inflatable bladders 16, 18 positioned about the grappling arm for engaging an inner surface of the combustion chamber upon inflation, and a pair of rear bladders 20, 22 positioned about the grappling arm for engaging an inner surface of the nozzle downstream from the combustion chamber upon inflation for aligning the grappling arm and the rocket propulsion nozzle. A pressurized fluid source 24 is provided on the first space vehicle for supplying fluid pressure to the inflatable bladders, and a fluid control valve manifold 30 selectively controls the release of pressurized fluid to the bladders. According to the method of the invention, the grappling arm is inserted into the rocket propulsion nozzle, and the control valves are actuated to first inflate the front bladders and thereby interconnect the grappling arm and the rocket propulsion nozzle. The rear bladders are subsequently inflated to align a central axis of the grappling arm with a central axis of the rocket propulsion nozzle. Inflation of the rear bladders provides an axial reaction load to balance the axial load provided by the front bladders. Attaching the vehicles in space may be controlled from the earth by activating the control valves to inflate the bladders. The cost of attaching space vehicles is significantly reduced by attaching a grappling arm on one vehicle with the existing rocket propulsion nozzle of another vehicle.

A scientific view of the productivity of abrasive blasting nozzles

Abstract: ABRASIVE blasting is a sensitive issue in infrastructure refurbishment. Recent environmental laws and resulting costs have brought about a powerful incentive for improvement, but the scientific basis of abrasive blasting has remained essentially the same for over a century. Everyone knows that abrasive blasting "doesn' t take a rocket scientist." However, a rocket propulsion scientist, in fact, can bring about some much needed improvements in abrasive blasting, especially in the efficiency and productivity of blasting nozzles. For instance, the exhaust of the space shuttle main engine contains "shock diamonds," which are characteristic of supersonic flow. So does a blasting nozzle, although the diamonds are seldom visible to the naked eye. Micrometer-sized alumina particles are accelerated through the nozzles of the shuttle's solid rocket boosters (Ref 1), just as abrasive particles are accelerated through a blasting nozzle. (Actually, it is easier to analyze the shuttle booster problem because the particles are so small.) This article will explain how some of the extensive knowledge and sophistication of rocket nozzle technology (Ref 2) can be used to improve the productivity of abrasive blasting nozzles. It begins with a brief review of the history and state-of-the-art of blasting nozzles, followed by a somewhat technical explanation of bow to improve productivity. Why is the so-called venturi nozzle a better choice for abrasive blasting than the earlier straight-bore nozzle? From the field of gas dynamics (Ref 4) comes an immediate answer: just as in a rocket nozzle, it is the only way to achieve high speed flow at the nozzle exit. The nozzle must contract to a minimum area for the flow to reach the speed of sound, and then it expands to produce supersonic airspeeds. Rocket nozzles like the one shown in Fig. 2 obey the same principle, although in order to be effective at high altitudes, they expand to much larger exit diameters than the earthbound blasting nozzle. However, blasting nozzles have not benefited from any twentieth-century aerospace technology developments. For example, consider the double-venturi scheme shown in Fig. 1. It is a type of ejector nozzle that entrains outside air, but there is no

Chemical Microthruster Options

Abstract: 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.

Development of the Cassini Spacecraft Propulsion Subsystem

Abstract: The Cassini Spacecraft will be launched on an expedition to Saturn in October 1997. The mission is an eleven-year operation; the first seven years traveling to Saturn via a combination of propulsive burns and VenusVenus-Earth-Jupiter gravity-assist, and the remaining four years orbiting Saturn while exploring the planet, its moons, rings, and nearby icy satellites. The propulsion module subsystem provides thrust and torque to the spacecraft. Larger Delta-V maneuvers are conducted with a primary (with redundant backup) pressure-regulated 445-N engine, which burns nitrogen tetroxide (NTO) and monomethylhydrazine (MMH). Total propellant capacity is 3000 kg. Saturn Orbit Insertion (SOI) is accomplished with a 170minute (maximum duration) continuous firing of the bipropellant engine. Attitude control of the spacecraft is maintained by 1-N thrusters (arranged in two redundant pairs per each of four clusters), which operate in a blowdown mode, with the monopropellant tank (containing 132 kg of hydrazine) recharged once from a dedicated helium pressurant tank. A significant development program was implemented to provide design verification of major assemblies, and to extend the performance capabilities of heritage components. The flight hardware is currently in the final integration and test phase prior to shipment to JPL for integration with the spacecraft. Introduction Historical Background Cassini has its roots in the Saturnian system exploration studies that began in 1989 with the CRAF/Cassini Program. That Program involved two separate spacecraft — the CRAF (Comet Rendezvous and Asteroid Flyby), originally targeted for a 1995 launch to flyby an asteroid and eventually rendezvous with a comet; and Cassini, slated for a 1996 launch to explore Saturn and its many satellites and rings. Although going through major changes since its inception, Cassini remains an international cooperative effort of NASA, which is producing the main orbiter spacecraft (S/C); the European Space Agency (ESA), which is providing the Huygens Probe; and the Italian Space Agency (ASI), responsible for the S/C radio antenna and portions of three scientific experiments. The mission is being managed by NASA's Jet Propulsion Laboratory (JPL), where the orbiter is being designed, built and tested. In keeping with the international flavor of CRAF/Cassini from its inception, the earliest S/C Propulsion Subsystem was to be built by the Federal Republic of Germany (FRG). Technically, the Germans were contributing one propulsion subsystem for the CRAF Mission and a spare subsystem that was * Member AIAA * Senior Member AIAA Copyright © 1996 Lockheed Martin Astronautics. Published by The American Institute of Aeronautics and Astronautics, Inc. with Permission

`Galileo Galilei' flight experiment on the equivalence principle with field emission electric propulsion

Abstract: An experiment to test the equivalence of inertial to gravitational (passive) mass in space offers two main advantages: a signal about a factor of a thousand larger than on Earth and the possibility of exploiting the absence of weight. `Galileo Galilei' (GG) is a small satellite mission currently under study in Italy with the financial support of ASI (Agenzia Spaziale Italiana). The mission concerns a small, low Earth satellite ( total mass, altitude) with two objectives. One is scientific, in the field of fundamental physics, and the other technological within the framework of spacecraft propulsion and drag compensation. The scientific goal is to test the equivalence principle to one part in , four orders of magnitude better than the best ground results. The technological goal is a full, comprehensive test of FEEP (field emission electric propulsion) thrusters for accurate drag compensation, a technology developed in Europe by the ESA (European Space Agency) which is likely to become an essential component of all space experiments which require measurement of small forces. The GG experiment is based on novel concepts and does not require low temperatures.

Electrodynamic Tethers for Reboost of the International Space Station and Spacecraft Propulsion.

Abstract: The International Space Station (ISS) will require periodic reboost due to atmospheric aerodynamic drag. This is nominally achieved through the use of thruster firings by the attached Progress M spacecraft. Many Progress flights to the ISS are required annually. Electrodynamic tethers provide an attractive alternative in that they can provide periodic reboost or continuous drag cancellation using no consumables, propellant, nor conventional propulsion elements. The system could also serve as an emergency backup reboost system used only in the event resupply and reboost are delayed for some reason. The system also has direct application to spacecraft and upper stage propulsion. Electrodynamic tethers have been demonstrated in space previously with the Plasma Motor Generator (PMG) experiment and the Tethered Satellite System (TSS-1R). The advanced electrodynamic tether proposed for this application has significant advantages over previous systems in that higher thrust is achievable with significantly shorter tethers and without the need for an active current collection device, hence making the system simpler and much less expensive. Propellantless Reboost for the ISS: An Electrodvnamic Tether Thruster The need for an alternative to chemical thruster reboost of the ISS has become increasingly apparent as the station nears completion. We propose a system to utilize ISS electrical power to generate thrust by means of a new type of electrodynamic tether attached to the station (Fig. 1). A flexible system could be developed to generate an average thrust of 0.5 to 0.8 N for 5 to 10 kW of electrical power. By comparison, aerodynamic drag on 755 is expected to average from 0.3 to 1.1 N (depending upon the year). The proposed system uses a tether with a kilometers-long uninsulated (bare) segment capable of collecting currents greater than 10 A from the ionosphere. The new design exhibits a remarkable insensitivity to electron density variations, allowing it

Space power technology

Abstract: The trends today are toward smaller satellites with increased capability. These two trends place enormous demands on the technology used to power such space systems. Typical power sources are nuclear, photovoltaic, and chemical. As the power available on a typical satellite decreases, the use of long pulsed power sources becomes more attractive. Technologies such as electrochemical capacitors, pulsed batteries and flywheels offer the possibility of employing high power subsystems for limited times without undue burden on the spacecraft power train. The state of the art in each of the pulsed storage devices is discussed and preliminary experiments described which use an electrochemical power source to power a spacecraft propulsion unit such as an arc-jet or Hall effect thruster.

Commercially-driven human interplanetary propulsion systems: Rationale, concept, technology, and performance requirements

Abstract: Previous studies of human interplanetary missions are largely characterized by long trip times, limited performance capabilities, and enormous costs. Until these missions become dramatically more ‘‘commercial‐friendly’’, their funding source and rationale will be restricted to national governments and their political/scientific interests respectively. A rationale is discussed for human interplanetary space exploration predicated on the private sector. Space propulsion system requirements are identified for interplanetary transfer times of no more than a few weeks/months to and between the major outer planets. Nuclear fusion is identified as the minimum requisite space propulsion technology. A conceptual design is described and evolutionary catalyzed‐DD to DHe3 fuel cycles are proposed. Magnetic nozzles for direct thrust generation and quantifying the operational aspects of the energy exchange mechanisms between high energy reaction products and neutral propellants are identified as two of the many key supporting technologies essential to satisfying system performance requirements. Government support of focused, breakthrough technologies is recommended at funding levels appropriate to other ongoing federal research.

Repositioning of geostationary spacecraft - Chemical and electric propulsion options

Abstract: This paper explores the application of electric propulsion to the three major on-board propulsion tasks needed for geosynchronous satellites: north-south stationkeeping, high altitude orbit raising, and station relocation. It is a general survey of practical repositioning in geostationary orbit, making use of on-board electric propulsion and power only. The paper covers two chemical and three electric propulsion options, four levels of longitudinal displ acement, and three spacecraft models across a range of satellite power/mass ratio. In addition, the merit of 'jump-starting' the maneuver with stored power is briefly examined. Chemical propulsion options are hydrazine monopropellant and storable bi-propellant systems. The full range of electric propulsion is explored using the hydrazine arcjet, the SPT-100 gridless ion thruster, and the conventional gridded xenon ion thruster. Velocity increments corresponding to longitude changes of 30, 60, 1200, and 180 deg are considered. Three S/C models are chosen across a range of satellite power/mass ratios. (Author)

Electrodynamic tethers for reboost of the International Space Station and spacecraft propulsion

Abstract: The International Space Station (ISS) will require periodic reboost due to atmospheric aerodynamic drag. This is nominally achieved through the use of thruster firings by the attached Progress M spacecraft. Many Progress flights to the ISS are required annually. Electrodynamic tethers provide an attractive alternative in that they can provide periodic reboost or continuous drag cancellation using no consumables, propellant, nor conventional propulsion elements. The system could also serve as an emergency backup reboost system used only in the event resupply and reboost are delayed for some reason. The system also has direct application to spacecraft and upper stage propulsion. Electrodynamic tethers have been demonstrated in space previously with the Plasma Motor Generator (PMG) experiment and the Tethered Satellite System (TSS-1R). The advanced electrodynamic tether proposed for this application has significant advantages over previous systems in that higher thrust is achievable with significantly shorter tethers and without the need for an active current collection device, hence making the system simpler and much less expensive. Propellantless Reboost for the ISS: An Electrodvnamic Tether Thruster The need for an alternative to chemical thruster reboost of the ISS has become increasingly apparent as the station nears completion. We propose a system to utilize ISS electrical power to generate thrust by means of a new type of electrodynamic tether attached to the station (Fig. 1). A flexible system could be developed to generate an average thrust of 0.5 to 0.8 N for 5 to 10 kW of electrical power. By comparison, aerodynamic drag on 755 is expected to average from 0.3 to 1.1 N (depending upon the year). The proposed system uses a tether with a kilometers-long uninsulated (bare) segment capable of collecting currents greater than 10 A from the ionosphere. The new design exhibits a remarkable insensitivity to electron density variations, allowing it

Investigation Into Hybrid Rockets and Other Cost-Effective Propulsion System Options For Small Satellites.

Abstract: Affordable small satellites need affordable propulsion systems. The primary objective of this research was to investigate cost-effective propulsion system options for small satellites. The research comprised four interrelated goals: Identify key cost drivers for spacecraft hardware. Characterise propulsion technology costs. Characterise propulsion system costs. Evaluate and compare the cost-effectiveness of system options. Each of these goals was attained. Important results and conclusions emerged. To begin with, key spacecraft hardware cost drivers were shown to occur during each phase of a mission—definition, design and acquisition—and a process for resolving them was advanced. Furthermore, propulsion system costs were shown to include far more than performance and price. A new paradigm for understanding the total cost of propulsion systems was defined that encompasses nine dimensions— mass, volume, time, power, system price, integration, logistics, safety and technical risk. This paradigm was used to characterise propulsion technology options. From this effort, hybrid rockets emerged as a promising but underdeveloped technology with great potential for cost-effective application. To evaluate this potential, a dedicated research program was completed during which a hybrid motor was designed, built and tested using 85% hydrogen peroxide as oxidiser and polythene as fuel. The basic concept for a hybrid upper stage was proven. Excellent combustion performance was measured and characterised. Real total costs for future small satellite applications were assessed. This research demonstrated that hybrid rockets offer a safe, reliable upper stage option that is a versatile, cost-effective alternative to solid rocket motors. In addition, despite negative industry bias, hydrogen peroxide proved itself as a safe, effective oxidiser for hybrid and mono-propellant applications. The characterisation of propulsion system costs led to a complete design case study for a minisatellite aimed at the most cost-effective solution. It was shown that by focusing on the key cost drivers and trading among the cost dimensions, a truly versatile, cost-effective system design can be achieved. Finally, an innovative technique was derived to parametrically combine the diverse cost dimensions into a useful, quantifiable figure of merit for mission and research planning. Overall, it was shown that the most cost-effective solution is found by weighing all options along the nine dimensions of the cost paradigm within the context of a specific mission. Overall, the research advances the state of the art of hybrid rocket technology specifically, and satellite engineering in general.

Advanced Space Propulsion

Abstract: This presentation describes a number of advanced space propulsion technologies with the potential for meeting the need for dramatic reductions in the cost of access to space, and the need for new propulsion capabilities to enable bold new space exploration (and, ultimately, space exploitation) missions of the 21st century. For example, current Earth-to-orbit (e.g., low Earth orbit, LEO) launch costs are extremely high (ca. $10,000/kg); a factor 25 reduction (to ca. $400/kg) will be needed to produce the dramatic increases in space activities in both the civilian and government sectors identified in the Commercial Space Transportation Study (CSTS). Similarly, in the area of space exploration, all of the relatively 'easy' missions (e.g., robotic flybys, inner solar system orbiters and landers; and piloted short-duration Lunar missions) have been done. Ambitious missions of the next century (e.g., robotic outer-planet orbiters/probes, landers, rovers, sample returns; and piloted long-duration Lunar and Mars missions) will require major improvements in propulsion capability. In some cases, advanced propulsion can enable a mission by making it faster or more affordable, and in some cases, by directly enabling the mission (e.g., interstellar missions). As a general rule, advanced propulsion systems are attractive because of their low operating costs (e.g., higher specific impulse, ISD) and typically show the most benefit for relatively 'big' missions (i.e., missions with large payloads or AV, or a large overall mission model). In part, this is due to the intrinsic size of the advanced systems as compared to state-of-the-art (SOTA) chemical propulsion systems. Also, advanced systems often have a large 'infrastructure' cost, either in the form of initial RD and directly supported independent research at universities and industry. The cooperative program consists of mission studies, research and development of ion engine technology using C60 (Buckminsterfullerene) propellant, and research and development of lithium-propellant Lorentz-force accelerator (LFA) engine technology. The university/industry-supported research includes modeling and proof-of-concept experiments in advanced, high-lsp, long-life electric propulsion, and in fusion propulsion.

Economical rocket flight

Abstract: A mathematical model of a rocket travelling in the earth's atmosphere was developed. Numerical solutions of the model were used to demonstrate that efficient operation of a rocket is critically dep...

Space Propulsion and Power: Operational Effectiveness and Cost Study (OECS).

Abstract: : The performance of chemical upper stage rocket propulsion cannot be improved much in the near future. However, promising alternative solar and nuclear powered upper stages could be operational within ten years. This report provides a detailed discussion of these alternative upper stage concepts and their likely applications and performance. Included are descriptions of the methodology and the results of a cost effectiveness analysis. The report also evaluates satellite electrical power generation because many of the concepts have propulsion and electrical power subsystems integrated with the satellite. We found frequent, significant performance and cost advantages for the alternative technologies compared with today's baseline. These technologies would make it possible to place the same payload mass into operational orbit with a smaller, less costly booster or to place more payload mass into orbit with the original booster. Off-setting these advantages are low thrust levels and, consequently, longer orbital transfer times.

NSTAR Ion Thruster and Breadboard Power Processor Functional Integration Test Results

Abstract: A 23 kW Breadboard Power Processing Unit (BBPPU) was developed as part of the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) Program The NSTAR program will deliver an electric propulsion system based on a 30 cm xenon ion thruster to the New Millennium (NM) program for use as the primary propulsion system for the initial NM flight The final development test for the BBPPU, the Functional Integration Test, was carried out to demonstrate all aspects of BBPPU operation with an Engineering Model Thruster Test objectives included: (1) demonstration and validation of automated thruster start procedures, (2) demonstration of stable closed loop control of the thruster beam current, (3) successful response and recovery to thruster faults, and (4) successful safing of the system during simulated spacecraft faults These objectives were met over the specified 80-120 VDC input voltage range and 05-23 output power capability of the BBPPU Two minor anomalies were noted in discharge and neutralizer keeper current These anomalies did not affect the stability of the system and were successfully corrected

Effects of fuel formulation on regression performance in hybrid motors

Abstract: Hybrid rocket propulsion—derived from a solid fuel burned with a liquid or gaseous oxidizer—has the potential to be safer, more flexible, and less expensive (compared to solids) as no energetic materials are involved; it can be throttled, stopped, and restarted; and the fuel can be formulated from materials already in volume production. Advances in fuel and combustion technology will make an important contribution toward enabling hybrid rocket propulsion to achieve its potential and compete effectively with solids and liquids. McDonnell Douglas Aerospace-Huntsville (MDA-HSV) has assembled data on many hybrid motor firings from a variety of sources, analyzed the data, noted trends, and established comparison criteria. In addition, MDA-HSV has formulated and tested new hybrid rocket fuel compositions en route to development of a higher performance solid hybrid rocket fuel. This paper presents hybrid motor performance data from several sources including test and analysis procedures and how these affect the usefulness of the data, and reports test results on new fuel compositions. Introduction Although hybrid rockets (solid fuel with liquid oxidizer) have been flown intermittentiy since the 1930s, the technology of solid fuel development, testing, and utilization is still in its infancy. There are no standardized tests nor test equipment, no specific aspects of technique to be reported, and no standard data reduction and analysis techniques. As a result, it is often difficult to compare the performance of different fuels and/or fuel/oxidizer combinations based on what is reported in the literature However, trends as well as limitations of usefulness of incomplete data have been noted and will be discussed. The most important general correlation and predictor of fuel regression behavior in hybrid rocket motors is

Integrated Solar Upper Stage (ISUS) mission trades

Abstract: Solar thermal propulsion and propulsion/power systems were identified as key technologies in the operational effectiveness and cost comparison study (OECS) sponsored by Phillips Laboratory (PL). These technologies were found to be pervasively cost effective with short transfer times and very good performance across a wide range of missions. The on-going Integrated Solar Upper Stage (ISUS) Program sponsored by PL represents development of a solar thermal propulsion/power (bimodal) system. As part of this effort, mission trades are being conducted to further define the ISUS system for geosynchronous equatorial orbit (GEO), high Earth orbit (HEO-Molniya class), and mid Earth orbit (MEO-GPS class) missions. These trades will consider launch vehicles ranging in size from a LLV3 to an Atlas IIAS that insert the ISUS into low Earth orbit (LEO). These trades will be used to define the ISUS system for the planned Engine Ground Demonstration, a space demonstration mission, and as a future operational system.

Jupiter explorer using microspacecraft and bimodal reactor system

Abstract: Using second generation microspacecraft and space reactor bimodal systems enables having a fleet of microspacecraft throughout the Jovian Planetary System simultaneously obtaining scientific data of Jupiter and its satellites. The microspacecraft uses new micro-technology and each spacecraft has a dry mass of 10 to 20 kg as described by Collins, et al. (1995). The space reactor bimodal system, defined by an Air Force study for Earth orbital missions and reported by Weitzberg, et al. (1995), provides 10 kWe power, 1000 N thrust, 850 s Isp, with a 1500 kg system mass. Using this bimodal system, trajectories to Jupiter were examined and an optimal direct and gravity assisted trajectory selected as described by Zubrin and Mondt (1996). A conceptual design for a spacecraft using the space reactor bimodal system for propulsion and power, that is capable of performing the Jupiter mission of interest, is defined. An end-to-end example mission is defined for Jupiter and it satellites with 11 microspacecraft. An Atlas 2AS launch vehicle is used to launch the bimodal spacecraft and its payload into Earth orbit. The space reactor propulsion system is then used with Earth gravity assists to place the spacecraft and eleven second generation microspacecraft from Earth orbit into a Jupiter orbit. The space reactor bimodal spacecraft acts as a carrier and communication spacecraft for a fleet of microspacecraft. This fleet of microspacecraft act in a coordinated manner and gather science data from several scientific target in the Jovian planetary system.

Preparing for the next millennium: Report on the Sixth Annual Advanced Space Propulsion System Workshop

Abstract: On 21–23 March 1995, JPL hosted the NASA Office of Space Access and Technology’s (OSATs) Sixth Advanced Space Propulsion Workshop. These workshops provide an informal means for researchers to exchange their research information in a timely manner. Presentations spanned the range from advanced chemical propulsion and advanced fission propulsion to plasma propulsion, fusion propulsion, antiproton propulsion to some intriguing uncategorized advanced concepts. This paper will provide a short summary of these informal presentations.