Showing papers on "Spacecraft propulsion published in 1988"

Performance and lifetime assessment of MPD arc thruster technology

Abstract: A summary of performance and lifetime characteristics of pulsed and steady-state magnetoplasmadynamic (MPD) thrusters is presented. The technical focus is on cargo vehicle propulsion for exploration-class missions to the Moon and Mars. Relatively high MPD thruster efficiencies of 0.43 and 0.69 have been reported at about 5000 s specific impulse using hydrogen and lithium, respectively. Efficiencies of 0.10 to 0.35 in the 1000 to 4500 s specific impulse range have been obtained with other propellants (e.g., Ar, NH3, N2). Thermal efficiency data in excess of 0.80 at MW power levels using pulsed thrusters indicate the potential of high MPD thruster performance. Extended tests of pulsed and steady-state MPD thrusters yield total impulses at least two to three orders of magnitude below that necessary for cargo vehicle propulsion. Performance tests and diagnostics for life-limiting mechanisms of megawatt-class thrusters will require high fidelity test stands which handle in excess of 10 kA and a vacuum facility whose operational pressure is less than 3 x 10 to the -4 torr.

Low pressure reaction control propulsion system for a spacecraft

Abstract: A spacecraft propulsion system which integrates the function of the apogee kick motor (AKM) and reaction control system (RCS) is disclosed. In accordance with this invention, a pump-fed AKM is employed which results in lightweight main tanks and pressurization systems. The RCS thrusters are operated by small bellows tanks which are intermittently pressurized by a gas pressurization system to provide high pressure for operation of the RCS thrusters. The system according to this invention enables use of lighter weight main propellant tanks since they do not have to withstand high internal pressures and also enables realization of the numerous advantages of a pump-fed AKM. Several embodiments describe various methods for cycling the bellows tanks.

A reusable rocket engine intelligent control

Abstract: An intelligent control system for reusable space propulsion systems for future launch vehicles is described. The system description includes a framework for the design. The framework consists of an execution level with high-speed control and diagnostics, and a coordination level which marries expert system concepts with traditional control. A comparison is made between air breathing and rocket engine control concepts to assess the relative levels of development and to determine the applicability of air breathing control concepts to future reusable rocket engine systems.

Performance of 10-kW class xenon ion thrusters

Abstract: Presented are performance data for laboratory and engineering model 30 cm-diameter ion thrusters operated with xenon propellant over a range of input power levels from approximately 2 to 20 kW. Also presented are preliminary performance results obtained from laboratory model 50 cm-diameter cusp- and divergent-field ion thrusters operating with both 30 cm- amd 50 cm-diameter ion optics up to a 20 kW input power. These data include values of discharge chamber propellant and power efficiencies, as well as values of specific impulse, thruster efficiency, thrust and power. The operation of the 30 cm- and 50 cm-diameter ion optics are also discussed.

Fluid mechanics in a magnetoplasmadynamic thruster

Abstract: Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1988.

Modular, Ion-Propelled, Orbit-Transfer Vehicle

Abstract: The design approach is presented for a modular, ion-propelled, orbit transfer vehicle (OTV). The OTV consists of a propulsion module that can be returned to Earth via the Shuttle for refueling and refurbishment, and a reusable power bus that mates to the spacecraft paylpad and remains in orbit. The technologies required to make the OTV concept both technically and economically feasible are identified. As an example of how the OTV could be applied, the NAVSTAR/Global Positioning System (GPS) Block 3 mission is examined using both conventional (expendable) chemical stages and the ion propulsion OTV. The OTV approach is shown to be particularly attractive, from a cost standpoint, for the specific application to GPS. The high specific impulse provided by ion propulsion is shown to result in a new reduction of $145 to 195 million in overall cost for the GPS Block 3 mission as compared with the cost using the Payload Assist Module (PAM) D-II chemical propulsion stage. This reusable OTV approach is believed to be equally attractive for other missions that require multiple launches.

Electromagnetic ground to orbit propulsion method and operating system for high mass payloads

Abstract: A reusable and regenerative electromagnetic propulsion method and operating system is provided for propelling high mass payloads to orbital velocities which does not require a vacuum environment. The propulsion system comprises a self supporting superconducting dipole coil several kilometers in diamater that is accelerated by magnetic repulsive forces generated by a plurality of giant superconducting field coils mounted in the underground tunnels. The propulsion dipole is mounted inside a circular hypersonic wing-like structure equipped with movable aerodyanmic control surfaces for guidance. The propulsion system can accelerate a payload with any desired launch azimuth by accelerating along a line of magnetic induction generated by the field coils having the desired azimuth angle. The payload is attached to the propulsion system by a plurality of cables. After reaching orbital velocity, the payload is detached from the propulsion system and the propulsion system is decelerated back to the earth's surface by magnetic repulsive forces generated by the field coils. A large fraction of the orbital energy of the propulsion system is reconverted back into electrical energy by the inductive coupling between the magnetically decelerated propulsion coil and the field coils which is used to launch another payload.

Air Augmented Rocket Propulsion Concepts

Abstract: : This report presents the results of a study of five combined cycle propulsion systems capable of accelerating a single stage vehicle with lifting surfaces from ground launch to Earth orbit. The engines studied were airbreathing, rocket based combined cycle (RBCC) engines. 'Combined Cycle' engines integrate airbreathing and rocket propulsion systems into a single engine system. These engines transition from initial air-augmented rocket mode takeoff and initial acceleration to ramjet to scramjet and finally to rocket propulsion to orbital insertion velocity. Engine systems, engine/vehicle integration, vehicle structure, propellant storage systems and thermal protection system (TPS) design, sizing and weight estimation, overall performance on various trajectories, and ground support systems were studied. A technology assessment, subscale engine test plan, development program plan and life cycle costs estimates are presented. Keywords: Air augmented rockets, Ramjets, Scramjets, Orbit on demand, Combined cycle engines, Space vehicles.

Advanced propulsion for the Mars Rover Sample Return Mission

Abstract: The present evaluation of highly detailed advanced propulsion system design concepts for the Mars Rover Sample Return Mission proceeded by comparing a baseline chemical propulsion option with both storable and cryogenic advanced chemical propulsion alternatives and solar- and nuclear-based electric propulsion OTVs. Substantial launch mass reductions and commensurate payload mass increases were obtainable with both advanced chemical and electric propulsion cycles.

Nuclear propulsion - A vital technology for the exploration of Mars and the planets beyond

Abstract: The physics and technology issues and performance potential of various direct thrust fission and fusion propulsion concepts are examined. Next to chemical propulsion the solid core fission thermal rocket (SCR) is the olny other concept to be experimentally tested at the power (approx 1.5 to 5.0 GW) and thrust levels (approx 0.33 to 1.11 MN) required for manned Mars missions. With a specific impulse of approx 850 s, the SCR can perform various near-Earth, cislunar and interplanetary missions with lower mass and cost requirements than its chemical counterpart. The gas core fission thermal rocket, with a specific power and impulse of approx 50 kW/kg and 5000 s offers the potential for quick courier trips to Mars (of about 80 days) or longer duration exploration cargo missions (lasting about 280 days) with starting masses of about 1000 m tons. Convenient transportation to the outer Solar System will require the development of magnetic and inertial fusion rockets (IFRs). Possessing specific powers and impulses of approx 100 kW/kg and 200-300 kilosecs, IFRs will usher in the era of the true Solar System class speceship. Even Pluto will be accessible with roundtrip times of less than 2 years and starting masses of about 1500 m tons.

Evaluation of innovative rocket engines for single-stage earth-to-orbit vehicles

Abstract: Computer models of rocket engines and single-stage-to-orbit vehicles that were developed by the authors at DFVLR and NASA have been combined. The resulting code consists of engine mass, performance, trajectory and vehicle sizing models. The engine mass model includes equations for each subsystem and describes their dependences on various propulsion parameters. The engine performance model consists of multidimensional sets of theoretical propulsion properties and a complete thermodynamic analysis of the engine cycle. The vehicle analyses include an optimized trajectory analysis, mass estimation, and vehicle sizing. A vertical-takeoff, horizontal-landing, single-stage, winged, manned, fully reusable vehicle with a payload capability of 13.6 Mg (30,000 lb) to low earth orbit was selected. Hydrogen, methane, propane, and dual-fuel engines were studied with staged-combustion, gas-generator, dual bell, and the dual-expander cycles. Mixture ratio, chamber pressure, nozzle exit pressure liftoff acceleration, and dual fuel propulsive parameters were optimized.

Recent advances in low-thrust propulsion technology

Abstract: 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.

Arcjet Propulsion System for an SP-100 Flight Experiment

Abstract: The design and performance of a arcjet nuclear electric propulsion spacecraft, suitable for use in a space reactor power system (SRPS) flight experiment, are outlined. The vehicle design is based on a 92-kW ammonia arcjet system operating at a specific impulse of 1050 s and an efficiency of 45 percent. The arcjet/gimbal system, power processing unit, and propellant feed system are described. A 100-kW SRPS is assumed and the spacecraft mass is baselined at 5250 kg, excluding the propellant and propellant feed system. A radiation/arcjet efflux diagnostics package is included in the performance analysis. This spacecraft, assuming a Shuttle launch from Kennedy Space Center, can perform a 35-deg inclination change and reach a final orbit of 35,860 km with a 120-day trip time, thus providing a four-month active load for the SRPS. Alternatively, a Titan IV launch could provide a mass margin of 120 kg to a 1000km, 58-deg final orbit in 74 days.

Future spacecraft propulsion

Abstract: Propulsion requirements for launch vehicles, upper stages, satellites and platforms, and planetary spacecraft are described from a functional perspective and compared on an energy basis. Mission velocity requirements for a range of missions are presented. A simple model relating optimum exhaust velocity and maximum system delta-V as a function of system-specific energy is developed, which provides insight into the relationship between system performance and various power and propulsion subsystem characteristics. Based on this model, various advanced propulsion options, e.g., the solid-core nuclear rocket and nuclear electric propulsion, are evaluated, and the implications of this analysis for propulsion and power system technology development programs are discussed. The objective of this paper is to provide an overview of future propulsion requirements for the nonspecialist.

Electrothermal propulsion of spacecraft with millimeter and submillimeter electromagnetic energy.

Abstract: The concept of millimeter and submillimeter wave electrothermal propulsion is considered. State-of-the-art radiation sources from 30-1000 GHz are examined to determine their applicability to electrothermal propulsion systems. The problem of energy conversion and power conditioning in this frequency range is also addressed. The potential advantage of utilizing power beaming with millimeter and submillimeter systems is examined. Finally, areas of future research and development are indicated.

A tandem mirror hybrid plume plasma propulsion facility

Abstract: A concept in electrodeless plasma propulsion, which is also capable of delivering a variable Isp, is presented. The concept involves a three-stage system of plasma injection, heating, and subsequent ejection through a magnetic nozzle. The nozzle produces the hybrid plume by the coaxial injection of hypersonic neutral gas. The gas layer, thus formed, protects the material walls from the hot plasma and, through increased collisions, helps detach it from the diverging magnetic field. The physics of this concept is evaluated numerically through full spatial and temporal simulations; these explore the operating characteristics of such a device over a wide region of parameter space. An experimental facility to study the plasma dynamics in the hybrid plume was built. The device consists of a tandem mirror operating in an asymmetric mode. A later upgrade of this system will incorporate a cold plasma injector at one end of the machine. Initial experiments involve the full characterization of the operating envelope, as well as extensive measurements of plasma properties at the exhaust. The results of the numerical simulations are described.

Heat transfer in aerospace propulsion

Abstract: 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

Use of high temperature superconductors in magnetoplasmadynamic systems

Abstract: The use of Tesla-class high-temperature superconducting magnets may have an extremely large impact on critical development issues (erosion, heat transfer, and performance) related to magnetoplasmadynamic (MPD) thrusters and also may provide significant benefits in reducing the mass of magnetics used in the power processing system. These potential performance improvements, coupled with additional benefits of high-temperature superconductivity, provide a very strong motivation to develop high-temperature superconductivity (HTS) applied-field MPD thruster propulsion systems. The application of HTS to MPD thruster propulsion systems may produce an enabling technology for these electric propulsion systems. This paper summarizes the impact that HTS may have upon MPD propulsion systems.

Parametric studies of electric propulsion systems for orbit transfer vehicles

Abstract: The present parametric tradeoff study for OTV electric propulsion systems encompasses ammonia and hydrogen arcjets as well as Xe-ion propulsion systems with performance characteristics currently being projected for 1993 operation. In all cases, the power source is a nuclear-electric system with 30 kg/kW(e) specific mass, and the mission involves the movement of payloads from lower orbits to GEO. Attention is given to payload capabilities and associated propellant requirements. Mission trip time is identified as the key parameter for selection; while arcjets are preferable for shorter trip times, ion propulsion is more advantageous for longer trip times due to reduced propellant mass fraction.

Energy Conversion Measurements in Laser-Sustained Argon Plasmas for Application to Rocket Propulsion.

Abstract: : Laser Propulsion is the production of high specific impulse rocket thrust using a high power laser as a remote energy source Specific impulses in excess of 1000 seconds are achievable because propellant temperatures are very high and low molecular weight gases can be used This report focuses on the energy conversion mechanisms of laser-sustained plasmas in flowing argon The status of AFOSR sponsored experiments to determine thermal efficiency and global absorption is detailed An improved testing facility has allowed plasma operating conditions never before possible The results indicate that nearly all of the input laser power can be absorbed by a plasma Plasmas at elevated gas pressure have been tested, and preliminary results presented Optimal operating conditions have yet to be determined for the available laser powers and gas pressures Further experimentation at very high argon gas velocities ( 20 m/s) must be performed in order to completely characterize plasma behavior Keywords: Beamed energy propulsion, Laser plasma formation

Ion propulsion system design and throttling strategies for planetary missions

Abstract: This paper describes recent advances in ion propulsion system design which promise to increase the propulsion system reliability by reducing the overall system complexity. The greatest simplification in the overall propulsion system operation is accomplished through a change in the ion engine throttling strategy. By using three grid optics it is possible to effect engine throttling at a constant beam curent over at least a 3.8 to 1 variation in input power. Throttling at a constant beam current results in a single discharge chamber operating point and eliminates the need for active propellant flow controllers and complex engine throttling software. Detailed mission analysis calculations for a CNSR mission performed using this constant beam current throttling strategy indicate only a small reduction in delivered payload and increase in required propellant relative to a conventional throttling profile based on varying the beam curent.

Radiation augmented energy storage system

Abstract: This invention relates to an improved method of storing solar radiation energy in a spacecraft and using it with high efficiency for space propulsion.

Radiation effect on rocket engine performance

Abstract: The effects of radiation on the performance of modern rocket propulsion systems operating at high pressure and temperature were recognized as a key issue in the design and operation of various liquid rocket engines of the current and future generations. Critical problem areas of radiation coupled with combustion of bipropellants are assessed and accounted for in the formulation of a universal scaling law incorporated with a radiation-enhanced vaporization combustion model. Numerical algorithms are developed and the pertaining data of the Variable Thrust Engine (VTE) and Space Shuttle Main Engine (SSME) are used to conduct parametric sensitivity studies to predict the principal intercoupling effects of radiation. The analysis reveals that low enthalpy engines, such as the VTE, are vulnerable to a substantial performance set back by the radiative loss, whereas the performance of high enthalpy engines such as the SSME, are hardly affected over a broad range of engine operation. Additionally, combustion enhancement by the radiative heating of the propellant has a significant impact in those propellants with high absorptivity. Finally, the areas of research related with radiation phenomena in bipropellant engines are identified.

Status of high power electric propulsion technology

Abstract: The growing emphasis on very challenging missions and the anticipated availability of high power levels in space have led to renewed interest in high power electric propulsion. The status of high power electric propulsion technology and its applicability to various missions are reviewed. The major thruster and system technology issues are identified which must be addressed in a focussed program in order to assure technology readiness for these missions.

An expert system for probabilistic description of loads on space propulsion system structural components

Abstract: LDEXPT, an expert system that generates probabilistic characterizations of the loads spectra borne by spacecraft propulsion systems' structural components, is found by recent experience at NASA-Lewis to be useful in the cases of components representative of the Space Shuttle Main Engine's turbopumps and fluid transfer ducting. LDEXPT is composed of a knowledge base management system and a rule base management system. The ANLOAD load-modeling module of LDEXPT encompasses three independent probabilistic analysis techniques.

Fission fragment rockets: A potential breakthrough

Abstract: A new reactor concept which has the potential of enabling extremely energetic and ambitious space propulsion missions is described Fission fragments are directly utilized as the propellant by guiding them out of a very low density core using magnetic fields The very high fission fragment exhaust velocities yield specific impulses of approximately a million seconds while maintaining respectable thrust levels Specific impulses of this magnitude allow acceleration of significant payload masses to several percent of the velocity of light and enable a variety of interesting missions, eg, payloads to the nearest star, Alpha Centauri, in about a hundred years for very rapid solar system transport The parameters reported in this paper are based on a very preliminary analysis Considerable trade-off studies will be required to find the optimum system We hope the optimum system proves to be as attractive as our preliminary analysis indicates, although we must admit that our limited effort is insufficient to guarantee any specific level of performance