Showing papers on "Spacecraft propulsion published in 1985"

Technology of closed-drift thrusters

Abstract: Introduction A CLOSED-drift thruster is defined herein as a thruster in which ions are electrostatically accelerated in essentially the thrust direction, with the accelerating electric field established by an electron current interacting with a transverse magnetic field. One component of the electron motion is counter to the ion flow. Another component is normal to that direction. The current associated with this normal component is called the Hall current. In a closed-drift accelerator there is a complete, or closed, path for the Hall current. In addition, for the ions to be accelerated in essentially a single thrust direction, the ion cyclotron radius must be much larger than the total acceleration length. Closed-drift thrusters usually employ axially symmetric electrodes and pole pieces, with the magnetic field in the radial direction and the electric field in the axial direction. The Hall current flows in a circular closed path in such a configuration. A few closed-drift thrusters without axial symmetry have also been investigated. The closed-drift thruster is particularly well suited for operation in the 1000-2000 s range of specific impulse (approximately 10,000-20,000 m/s exhaust velocity). It is difficult to operate above about 1000 s with an electrothermal thruster due to excessive excitation and ionization losses. On the other hand, the space-charge-flow limitations of gridded electrostatic thrusters will not permit practical ion current densities below about 2000 s. Within the 1000-2000 s range, the electron backflow required to establish ion acceleration can, for the most part, be used to generate ions. The generation of ions constitutes the major closed-drift thruster loss in this range of specific impulse, and this loss can be under 100 eV per beam ion. The power processing requirements are also moderate. In a properly designed closed-drift thruster only one power circuit is required for steady-state operation, with the voltage of this circuit typically in the 50-500 V range.

Antiproton Annihilation Propulsion

Abstract: : Antiproton annihilation propulsion is a new form of space propulsion, where milligrams of antimatter are used to heat tons of reaction fluid to high temperatures. The hot reaction fluid is exhausted from a nozzle to produce high thrust at high specific impulse. This study was to determine the physical, engineering, and economic feasibility of antiproton annihilation propulsion. The conclusion of the study is that antiproton propulsion is feasible, but expensive. Because the low mass of the antimatter fuel more than compensates for its high price, comparative mission studies show that antimatter fuel can be cost effective in space, where even normal chemical fuel is expensive because its mass must be lifted into orbit before it can be used. Antiproton annihilation propulsion is mission enabling, in that it allows missions to be performed that cannot be performed by any other propulsion system. Keywords: Antimatter propulsion; Antiproton; Advanced propulsion.

Effects of tripropellant engines on earth-to-orbit vehicles

Abstract: The effects of tripropellant engines on earth-to-orbit vehicles is examined in terms of their impact on engine configurations and launch capabilities. Hydrocarbon fuel with some oxygen is used in tripropellant fuels, with hydrogen as a back-up fluid for coolant and driving the pumps. Engine concepts which implement tripropellant fuels include a hydrogen-gas generator engine with staged combustion, a two-mode engine burning hydrocarbon fuel in one chamber and hydrogen in another, a dual expander engine with a central hydrocarbon nozzle and an annular hydrogen nozzle, and a dual throat engine. The use of hydrocarbons reduces the fuel weight by providing a higher specific impulse than with LOX-LH2 systems alone. Studies have shown that single-stage-to-orbit vehicles capable of lifting 13.6 Mg are possible with tripropellant engines. A dual-expander engine, is identified as offering the most fuel dry mass reduction for a given payload if cooling requirements can be satisfied. Further development of the tripropellant engines is concluded to be beneficial.

A re-examination of the laser-supported combustion wave

Abstract: Kantrowitz (1972) and Minovitch (1972) have proposed the use of laser sustained plasmas as a means to heat a rocket propellant. Recent studies of laser-powered propulsion have been directed toward the application of high-specific-impulse space propulsion systems for orbital transfer missions. Analyses of rocket performance relied heavily on the concept of the laser-supported combustion (LSC) wave. Raizer (1971) first drew the analogy between laser-sustained plasmas and combustion waves in an analysis. The Raizer model was later applied to hydrogen by Kemp and Root (1979). In connection with certain problems arising with the approach considered by Kemp and Root, the present investigation is concerned with a reexamination of the Raizer model. Attention is given to a numerical approach for the entire LSC wave in hydrogen, taking into account the incorporation of the proper boundary conditions far downstream of the wave.

Ground launched space rocket

Abstract: The rocket propulsion unit for space flight comprises a low altitude thrust nozzle (3) and an axially movable (9) high-altitude nozzle (4) which is initially retracted and is extendable such that its forward end connects with the rear end (3b) of the nozzle (3). Up to a height of 12-14 km, the nozzle (3) acts as over-expansion nozzle producing a thrust flow (5) exit pressure of 0.2-0.15 bar whilst the retracted nozzle (4) has its rear end (4b) a little behind the end (3b) of the nozzle (3) and, together with that nozzle (3) and its thrust flow (5), defines an annular outer thrust system (6). The system has a convergent-divergent nozzle passage (7) providing an outer thrust flow (8) subject to injector action by the flow (5). At 12-14 km height or 0.2-0.15 bar ambient pressure, the outer nozzle (4) is fully extended, this nozzle being designed such that it then acts as over-expansion nozzle with a thrust flow exit pressure of 0.03-0.02 bar.

Solar Rocket Component Study

Abstract: : A program was conducted to identify the most promising solar-thermal rocket thruster concept and to design, fabricate, and deliver to the Air Force Rocket Propulsion Laboratory, a demonstrator absorber/thruster based on the selected concept for subsequent ground test evaluation. A concept assessment, in which windowless and windowed designs involving both direct and indirect propellant heating, resulted in the selection of the windowless heat exchanger cavity configuration. The selection considered the criteria of achievable specific impulse, thrust achievable, durability, complexity, technical risk, and relative cost. Based on the concept selected, a ground test demonstration absorber/thruster was designed. Supporting detailed thermal, stress, and performance analyses were conducted together with evaluations of critical materials and fabrication processes. the projected performance of the demonstrator absorber/thruster was a specific impulse of 7930 N sec/kg (807 lbf sec/lbm), with a thrust of 3.69 N (0.83 lbf). Fabrication and assembly were completed and the thruster satisfactorily leak-checked. The absorber coil was fabricated from rhenium by the same process. Radiation shielding and insulation minimized heat losses. Keywords include: Solar Thermal Propulsion, Solar Absorber, Thruster performance, Absorber/Thruster Design, Rhenium fabrication, Hydrogen Performance, Solar Window, Window Heat Transfer, Absorber/Thruster, and Heat Transfer.

Design study of an integrated aerobraking orbital transfer vehicle

Abstract: An aerobraking orbital transfer vehicle (AOTV) concept, which has an aerobrake structure that is integrated with the propulsion stage, is discussed. The concept vehicle is to be assembled in space and is space-based. The advantages of aeroassist over an all propulsive vehicle are discussed and it is shown that the vehicle considered is very competitive with inflatable and deployable concepts from mass and performance aspects. The aerobrake geometry is an ellipsoidally blunted, raked-off, elliptical wide-angle cone with a toroidal skirt. Propellant tanks, engines, and subsystems are integrated into a closed, isogrid aerobrake structure which provides rigidity. The vehicle has two side-firing, gimbaled RL-10 type engines and carries 38,000 kg of useable propellant. The trajectory during aerobraking is determined from an adaptive guidance logic, and the heating is determined from engineering correlations as well as 3-D Navier-Stokes solutions. The AOTV is capable of placing 13,500 kg payload into geosynchronous Earth orbit (GEO) or carrying a LEO-GEO-LEO round-trip payload of 7100 kg. A two-stage version considered for lunar missions results in a lunar surface delivery capability of 18,000 kg or a round-trip capability of 6800 kg with 3860 kg delivery-only capability.

Laser propulsion for the orbital transfer mission

Abstract: America's space activities in the 1990s and beyond will partly consist of missions involving the transportation of cargo from low earth orbit (LEO) to higher orbits or to an escape trajectory. Such missions are to be performed with the aid of an orbit transfer vehicle (OTV). The operation of the OTV can be based on different propulsion concepts. A chemical OTV is characterized by a high thrust and low specific impulse. The result is a short trip time at the cost of large quantities of propellant. On the other hand, low-thrust systems such as electric propulsion units, consume very little propellant, but would have a long trip time. The present paper is concerned with a compromise between these two extremes. The employed propulsion system utilizes laser thermal propulsion, in which a ground or space-based laser is used to beam energy to a thruster on the OTV. The laser light is absorbed by a propellant. The resulting heating of the propellant causes an expansion of the propellant through a nozzle to produce thrust. Details regarding this propulsion concept are discussed, taking into account operational questions and missions.

Space station propulsion: The advanced development program at Lewis

Abstract: A reference configuration was established for the initial operating capability (IOC) station. The reference configuration has assumed hydrazine fueled thrusters as the propulsion system. This was to establish costing and as a reference for comparison when other propulsion systems are considered. An integral part of the plan to develop the Space Station is the advanced development program. The objective of this program is to provide advanced technology alternatives for the initial and evolutionary Space Station which optimize the system's functional characteristics in terms of performance, cost, and utilization. The portion of the Advanced Development Program that is concerned with auxiliary propulsion and the research and programmatic activities conducted are discussed.

Electric propulsion produced environments and possible interactions with the SP-100 power system

Abstract: An examination of a conceptual SP-100 type power system has been conducted to assess the effect of possible electric thruster induced contamination. The operating characteristics and known exhaust plume properties of various electric thrusters are summarized along with the natural environment and its effects to provide background for the contamination discussion. Charge buildup on nonconducting surfaces, differential charging of surfaces and arcing should be moderated by the operation of an electric thruster system. Thin film deposition of eroded thruster materials is a possible contamination mechanism. Electric thruster field efflux and propellant condensation should have only a minor contaminating effect upon a nuclear power system.

Thermal arcjet technology for space propulsion

Abstract: Advanced space propulsion systems are required to meet projected Air Force needs through the year 2000 Most of these missions require a large, on-orbit impulse capability High specific impulse (I sub sp) electric engines can provide this impulse while consuming relatively little propellant An arcjet engine system, which operates in the range of 800 to 2000 s I sub sp, is a promising candidate to meet these projected Air Force mission needs This electric propulsion system is ideally suited to missions currently under consideration, such as the Space-based Radar and other space platforms, because sufficient power is already installed for other functions on the spacecraft Also, arcjet systems are attractive for NASA near-term, low-cost Mariner Mark II missions to Saturn and Uranus Development of arcjet engines was an Air Force and NASA-sponsored activity that proceeded vigorously from its inception during the late 1950's up to the mid-1960's when the programs were terminated This paper describes thermal arcjet technology as it was developed over two decades ago and points to the direction this technology development should proceed in the future In particular, operation with storable propellants such as ammonia and hydrazine are considered The performance, applicability and advantages of these systems in terms of increased payload and/or decreased trip times are discussed

Design of a nuclear electric propulsion orbital transfer vehicle

Abstract: Payload increases of three to five times that of the Shuttle/Centaur can be achieved using nuclear electric propulsion. Various nuclear power plant options being pursued by the SP-100 Program are described. These concepts can grow from 100 kWe to 1 MWe output. Spacecraft design aspects are addressed, including thermal interactions, plume interactions, and radiation fluences. A baseline configuration is described accounting for these issues. Starting the orbital transfer vehicle from an altitude of 300 km indicates no significant additional risk to the biosphere than a reactor powerplant that has run for 7 years at a 300-year orbit.

NASA electric propulsion technology

Abstract: It is pointed out that the requirements for future electric propulsion cover an extremely large range of technical and programmatic characteristics. A NASA program is to provide options for the many potential mission applications, taking into account work on electrostatic, electromagnetic, and electrothermal propulsion systems. The present paper is concerned with developments regarding the three classes of electric propulsion. Studies concerning electrostatic propulsion are concerned with ion propulsion for primary propulsion for planetary and earth-orbit transfer vehicles, stationkeeping for geosynchronous spacecraft, and ion thruster systems. In connection with investigations related to electromagnetic propulsion, attention is given to electromagnetic launchers, the Hall current thruster, and magnetoplasmadynamic thrusters. In a discussion of electrothermal developments, space station resistojets are considered along with high performance resistojets, arcjets, and a laser thruster.

An analysis of low-thrust, resistojet reboost for the Space Station

Abstract: This paper presents results of an analysis of low-thrust orbit maintenance of the Space Station. Propellant requirements and transfer times are given for reboost of the station through various altitude increments. The reboost can readily be accomplished with thrust levels that subject the station to an acceleration of less than the desired upper limit of 10 to the -5th g's. The variation in time and the probabilistic aspect of the predicted upper-atmospheric density as well as the variation in time of sun-pointing drag areas were taken into account. Estimates of the propellant requirements at different times during an 11-year solar cycle are given. It is shown that the amount of CO2 available from the station life-support system is sufficient, over most of the solar cycle, to provide the propellant for a resistojet orbit-maintenance system.

Primary propulsion of electrothermal, ion and chemical systems for space-based radar orbit transfer

Abstract: An orbit transfer mission concept has been studied for a Space-Based Radar (SBR) where 40 kW required for radar operation is assumed available for orbit transfer propulsion. Arcjet, pulsed electrothermal (PET), ion, and storable chemical systems are considered for the primary propulsion. Transferring two SBR per shuttle flight to 1112 km/60 deg using electrical propulsion systems offers an increased payload at the expense of increased trip time, up to 2000 kg each, which may be critical for survivability. Trade offs between payload mass, transfer time, launch site, inclination, and height of parking orbits are presented.

Space station propulsion options

Abstract: The selection of the propulsion system for the Space Station represents a complex issue. The present paper provides a summary of the Station design factors which dictate the propulsion requirements, taking into account approaches for meeting these requirements. Factors which affect propulsion system selection are related to thrusting strategy, volume and mass limitations, safety and contamination, electrical power, time phasing, synergistic opportunities, propellant scavenging, water electrolysis, and free-flyers. In a discussion of propulsion systems, attention is given to monopropellant options, bipropellant options, and resistojets.

Radiation Energy Receiver for Solar Propulsion Systems

Abstract: The concept of remotely heating a rocket propellant with a high-intensity radiant energy flux is especially attractive due to its high specific impulse and large payload mass capabilities. In this paper, a radiation receiverthruster which is especially suited to the particular thermodynamic and spectral characteristics of highly concentrated solar energy is proposed. In this receiver, radiant energy is volumetrically absorbed within a hydrogen gas seeded with alkali metal vapors. The alkali atoms and molecules absorb the radiant flux and subsequently transfer their internal excitation to hydrogen molecules through collisional quenching. It is shown that such a radiation receiver would outperform a blackbody cavity type receiver in both efficiency and maximum operating temperatures. A solar rocket equipped with such a receiver-thruster would deliver thrusts of several hundred Newtons at a specific impulse of 1000 seconds.

Station keeping of geostationary satellites by electric propulsion

Abstract: As various types of perturbations tend to drive a geostationary satellite away from its prescribed position, occasional orbit corrections have to be carried out by means of a suitable propulsion system. In future geostationary missions, low thrust electric propulsion is likely to be applied for station keeping because of considerable mass savings. In this paper a station keeping strategy for electric propulsion systems is developed. Both the unconstrained case and the case where thrust operation constraints are present are considered and tested by computer simulation of a realistic example.

Solar thermal propulsion for planetary spacecraft

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

Integratable propulsion systems for the Space Station

Abstract: Oxygen/hydrogen propulsion system options for space station orbit maintenance and attitude control were developed and evaluated relative to monopropellant and storable bipropellant propulsion systems. Space station propulsion requirements were analyzed with reference to such considerations as station size, altitude, power, crew size, and orbit transfer vehicle and orbital maneuvering vehicle servicing requirements. The evolutionary growth of oxygen/hydrogen bipropellant propulsion as an integral part of several interrelated space station functions, e.g., life support, power, and thermal management was considered. Propellant resupply evolves from resupply based on transport of liquid oxygen and liquid hydrogen to water. The advantages of the operation of the space station based on an oxygen/hydrogen economy are presented and discussed.

Reactor power system deployment and startup

Abstract: This paper addresses issues that should receive further examination in the near-term as concept selection for development of a U.S. space reactor power system is approached. The issues include: the economics, practicality and system reliability associated with transfer of nuclear spacecraft from low earth shuttle orbits to operational orbits, via chemical propulsion versus nuclear electric propulsion; possible astronaut supervised reactor and nuclear electric propulsion startup in low altitude Shuttle orbit; potential deployment methods for nuclear powered spacecraft from Shuttle; the general public safety of low altitude startup and nuclear safe and disposal orbits; the question of preferred reactor power level; and the question of frozen versus molten alkali metal coolant during launch and deployment. These issues must be considered now because they impact the SP-100 concept selection, power level selection, weight and size limits, use of deployable radiators, reliability requirements, and economics, as well as the degree of need for and the urgency of developing space reactor power systems.

Heavy lift launch vehicles for 1995 and beyond

Abstract: A Heavy Lift Launch Vehicle (HLLV) designed to deliver 300,000 lb to a 540 n mi circular polar orbit may be required to meet national needs for 1995 and beyond The vehicle described herein can accommodate payload envelopes up to 50 ft diameter by 200 ft in length Design requirements include reusability for the more expensive components such as avionics and propulsion systems, rapid launch turnaround time, minimum hardware inventory, stage and component flexibility and commonality, and low operational costs All ascent propulsion systems utilize liquid propellants, and overall launch vehicle stack height is minimized while maintaining a reasonable vehicle diameter The ascent propulsion systems are based on the development of a new liquid oxygen/hydrocarbon booster engine and liquid oxygen/liquid hydrogen upper stage engine derived from today's SSME technology Wherever possible, propulsion and avionics systems are contained in reusable propulsion/avionics modules that are recovered after each launch

Analytical investigation of solid rocket nozzle failure

Abstract: On April 5, 1983, an Inertial Upper Stage (IUS) spacecraft experienced loss of control during the burn of the second of two solid rocket motors. The anomaly investigation showed the cause to be a malfunction of the solid rocket motor. This paper presents a description of the IUS system, a failure analysis summary, an account of the thermal testing and computer modeling done at Marshall Space Flight Center, a comparison of analysis results with thermal data obtained from motor static tests, and describes some of the design enhancement incorporated to prevent recurrence of the anomaly.

Advanced Energy Storage Systems.

Abstract: : This report describes the selection, sizing, and optimization of energy storage/power conditioning systems (ES/PC) for the electromagnetic propulsion systems. An ES/PC is the interface between an electromagnetic thruster and a primary power source. The ES/PC accepts the relatively low, continuous power output from the primary power source, and converts it to a series of high power pulses and delivers it to the thruster. THe ES/PC minimizes the primary power requirements while maximizing the thruster performance. A selection, sizing, and optimization can be achieved if an optimization criterion exists, and if the ES/PC is examined within the context of the integrated propulsion system. We selected mass minimization as the optimization criterion. To develop the selection methodology, we identified the requirements imposed om the ES/PC and parametrically related the ES/PC performance to the propulsion system outputs. Originator-supplied keywords include: Electromagnetic propulsion, Orbit transfer, Stationkeeping, Optimization, Energy Storage/Power Conditioning (ES/PC), PUlsed Inductive Thruster (PIT), Teflon Pulsed Plasma Thruster (TPP), Magnetoplasmadynamic Thruster (MPD), Electric Rail Gun THruster (ERG), Inductors, Capacitors, and Compulsators.

Microwave electric propulsion for orbit transfer applications

Abstract: A Microwave Electric Propulsion (MEP) concept is developed for an unmanned Orbit Transfer Vehicle (OTV). The concept is based on the use of beamed microwave power and on an extrapolation of ion thruster technology. Beamed microwave power transmission is discussed in terms of its relationship to spacecraft propulsion. Characteristics of an MEP OTV are determined and performance is evaluated by comparison to a variety of alternative propulsion systems for the completion of a ten-year mission model.

An atomic hydrogen propulsion system

Abstract: P ERFORMANCE characteristics of a space propulsion system utilizing atomic hydrogen as a propellant are calculated. This thruster could be utilized for missions which require reliable long life operation, such as attitude correction, stationkeeping, and orbit change. A schematic diagram of the atomic hydrogen space propulsion thruster is shown in Fig. 1. An important feature of the engine is that atomic H is used as soon as it is produced by the discharge, thus eliminating the difficult problem of storage.' A mathematical model which describes the reaction chamber is presented; however, details of the microwave discharge, connecting regions, and other parts of the propulsion system are not discussed. The H2 is assumed to be completely dissociated by the microwave discharge with no impurities. The mathematical model is formulated in terms of a derived set of coupled nonlinear, first-order, differential equations governing the rate of formation and collisional dissociation of H2 and various energy loss mechanisms. These, in addition to the rate equation governing the density of atomic hydrogen in the chamber, are solved using iterative procedures. Based upon a specified flow rate of atomic hydrogen into the reaction chamber, values are given for the thrust, specific impulse, reaction chamber pressure and gas temperature, and the densities of H and H2 inside the reaction chamber for a spherical reaction chamber geometry.