Showing papers on "Spacecraft propulsion published in 1980"

Time-optimal orbit transfer trajectory for solar sail spacecraft

Abstract: This paper considers the problem of optimally controlling the sail steering angle of a solar sail spacecraft (a spaceship moving under the influence of solar radiation pressure) so as to execute a minimum-time coplanar orbit transfer from the mean orbital distance of Earth to the mean Martian orbital distance around the sun. This problem has been modeled as a free-terminal time-optimal control problem with an unbounded control variable and with state variable equality constraints at the final time. It has been solved by the penalty function approach using the conjugate gradient algorithm. Detailed computational results regarding the convergence of the optimal control, the state trajectories, and the norm of the gradient trajectory, along with the minimization of the performance functional, are presented. In general, the optimal solution is found to be compatible with that obtained by some of the earlier investigators of this problem. In conclusion, the optimized orbit transfer time of the solar sail spacecraft is compared with that of an ionic propulsion system.

Optimization of Rocket Propulsion Systems for Advanced Earth-to-Orbit Shuttles

Abstract: A computerized preliminary design system is used to evaluate potential main liquid-rocket propulsion systems for advanced technology winged single-state-to-orbit launch vehicles. Evaluated are tradeoffs between ascent flight trajectory performance and flight vehicle sizing driven by engine mass and propellant requirements. Numerous mission, flight, and vehicle-related requirements and constraints are satisfied in the design process. With the design system, five dual-mode propulsion system concepts are compared to a baseline hydrogen and oxygen system in terms of the changes in vehicle dry mass and gross mass.

Model rocket propulsion system

Abstract: The invention relates generally to a small model rocket motor. The motor includes a plastic composite propellant, and plastic ablative nozzle, and provides means for producing thrust, a time delay, and a charge to activate external devices of desirable design.

Advanced oxygen-hydrocarbon rocket engine study

Abstract: This study identifies and evaluates promising LO2/HC rocket engine cycles, produces a consistent and reliable data base for vehicle optimization and design studies, demonstrates the significance of propulsion system improvements, and selects the critical technology areas necessary to realize an improved surface to orbit transportation system. Parametric LO2/HC engine data were generated over a range of thrust levels from 890 to 6672 kN (200K to 1.5M 1bF) and chamber pressures from 6890 to 34500 kN (1000 to 5000 psia). Engine coolants included RP-1, refined RP-1, LCH4, LC3H8, LO2, and LH2. LO2/RP-1 G.G. cycles were found to be not acceptable for advanced engines. The highest performing LO2/RP-1 staged combustion engine cycle utilizes LO2 as the coolant and incorporates an oxidizer rich preburner. The highest performing cycle for LO2/LCH4 and LO2/LC3H8 utilizes fuel cooling and incorporates both fuel and oxidizer rich preburners. LO2/HC engine cycles permitting the use of a third fluid LH2 coolant and an LH2 rich gas generator provide higher performance at significantly lower pump discharge pressures. The LO2/HC dual throat engine, because of its high altitude performance, delivers the highest payload for the vehicle configuration that was investigated.

A prelude to interstellar flight

Abstract: A 20 to 50 year interstellar precursor mission extending 400 to 1000 AU from the solar system is outlined as a means of bringing out and solving engineering problems inherent in a star mission, and of studying the heliopause, the interstellar medium, and cosmic rays outside the heliosphere. Solar or laser sailing combined with a 500 kWe nuclear-electric propulsion system using fission would achieve a heliocentric excess velocity of 100km/s for the 32,000 kg spacecraft having a Shuttle derivative as a launch vehicle, and containing a Pluto flyby or separate orbiter powered by radioiosotope thermoelectric generators. X-band transmission using 40 w of power, a 15 m diameter spacecraft antenna and a 100 m receiving antenna on earth and providing 100 b/s is proposed, but a rate of 2 to 4 kb/s via 500 to 1000 w of power using the K-band and a 300 m diameter receiving antenna located on an Orbiting Deep Space Relay Station is also considered.

Electric propulsion for near-Earth space missions

Abstract: A set of missions was postulated that was considered to be representative of those likely to be desirable/feasible over the next three decades. The characteristics of these missions, and their payloads, that most impact the choice/design of the requisite propulsion system were determined. A system-level model of the near-Earth transportation process was constructed, which incorporated these mission/system characteristics, as well as the fundamental parameters describing the technology/performance of an ion bombardment based electric propulsion system. The model was used for sensitivity studies to determine the interactions between the technology descriptors and program costs, and to establish the most cost-effective directions for technology advancement. The most important factor was seen to be the costs associated with the duration of the mission, and this in turn makes the development of advanced electric propulsion systems having moderate to high efficiencies ( 50 percent) at intermediate ranges of specific impulse (approximately 1000 seconds) very desirable.

Prospects for a breakthrough in field-dependent 'propulsion'

Abstract: If sufficient justification exists to start the development of a field physics laboratory, attention might well be given to the development of two advanced types of field-independent propulsion systems These are gravimagnetic systems and field resonance systems The first are multipurpose propulsion systems employing the 'gravitational' effects of coherent electromagnetic energy configurations The second are deep-space propulsion systems which artificially generate an energy pattern that matches precisely, or resonates with, a virtual pattern associated with a distant space-time point A new model of space-time/field interactions is used in the present paper to describe the potential characteristics of electromagnetic/gravitational field interactions and the performance capabilities of these two propulsion systems

Laser propulsion - 1980

Abstract: An overview of the current government sponsored work in laser propulsion is presented and the NASA program is discussed Attention is given to the overall NASA plan in laser propulsion and the laser rocket engine technology program Some results of an analytical effort at Physical Sciences Inc are presented, as well as results of the NASA/Army Missile Command experimental effort Finally, future plans are briefly summarized

Upper stages utilizing electric propulsion

Abstract: The payload characteristics of geocentric missions which utilize electron bombardment ion thruster systems are discussed. A baseline LEO to GEO orbit transfer mission was selected to describe the payload capabilities. The impacts on payloads of both mission parameters and electric propulsion technology options were evaluated. The characteristics of the electric propulsion thrust system and the power requirements were specified in order to predict payload mass. This was completed by utilizing a previously developed methodology which provides a detailed thrust system description after the final mass on orbit, the thrusting time, and the specific impulse are specified. The impact on payloads of total mass in LEO, thrusting time, propellant type, specific impulse, and power source characteristics was evaluated.

Nuclear electric propulsion system utilization for earth orbit transfer of large spacecraft structures

Abstract: The paper discusses a potential application of electric propulsion to perform orbit transfer of a large spacecraft structure to geosynchronous orbit (GEO) from LEO, utilizing a nuclear reactor space power source in the spacecraft on a shared basis. The discussions include spacecraft, thrust system, and nuclear reactor space power system concepts. Emphasis is placed on orbiter payload arrangements, spacecraft launch constraints, and spacecraft LEO assembly and deployment sequences.

A long-life bipropellant system demonstration: Viking Orbiter Propulsion System - 4 years in space and operating

Abstract: On April 23, 1980, Viking Orbiter One (VO-1), operating in the blowdown mode, completed a ten second Mars orbit trim maneuver to position the spacecraft for its final science sequence in May-June 1980. This brought the number of propulsive maneuvers for VO-1 to 23. Total accumulated operating time for the rocket engine was 2896 seconds, representing a total impulse of 3.93 x 10 to the 6th N-sec. The estimated propellant remaining was sufficient to operate the rocket engine for an additional 30 seconds. VO-1 has completed more than 1700 days in space, 1400 days in orbit around Mars and more than two years of attitude control system operation with helium gas transferred from the propulsion system pressurant tank. The mass of helium remaining is expected to be sufficient for attitude control through June 1980. This paper presents the flight history of the Viking 75 Orbiter Propulsion Systems and summarizes the design and test philosophy which have contributed to their success.

Technology Status of a Fluorine-Hydrazine Propulsion System for Planetary Spacecraft

Abstract: The basic technology exists and a system integration program is well underway to allow incorporation of a fluorine-hydrazine propulsion system into future spacecraft required for unmanned planetary missions. These spacecraft would be inserted in earth orbit using the Space Transportation System Shuttle and given its initial sendoff by the Inertial Upper Stage (IUS). The design of a typical propulsion system, assessment of thermal and structural impacts on a selected spacecraft and comparative studies with conventional propulsion systems have been completed. A major part of the current JPL Program involves assembly of a 3650 N thrust demonstration system using titanium tanks, flight weight components and structure. This system will be used to demonstrate the state-of-the-art throughout a representative flight system's qualification.

Technologies involved in configuring an advanced earth-to-orbit transport for low structural mass

Abstract: The current space shuttle is expected to adequately meet Government and industry needs for the transport of cargo to and from orbit well into the 1990's. However, continual study of potential follow-on shuttle systems is necessary and desirable in order to complement ongoing research in materials, structures, propulsion, aerodynamics, and other related areas. By studying alternate systems well in advance, it will be possible to explore the various technologies and develop those for which there is the greatest apparent payoff. In this paper a single-stage Earth-to-orbit transport designed for delivery of approximately 29,500 kg (65,000 lb) payload will be described. The vehicle, which takes off vertically and lands horizontally, is 60 m (197 feet) long and weighs approximately 1.8 Gg (4 M lb) at liftoff. In the interest of weight reduction, a simple body of revolution is utilized for the main body shell. In this design the main propulsion tanks serve as a primary load-carrying structure. Further, in order to minimize structural mass, the cargo bay is located between two of the main propellant tanks. The cargo volume, at 396 cu m (14,000 cu feet), exceeds that provided by the shuttle; but the bay itself is nonconforming in shape - being approximately 10 m (32 feet) in diameter by 5 m (17 feet) long. Dual-fuel propulsion is employed, since a number of studies have shown that (though lowering performance) the operation of hydrocarbon (RP) engines in parallel with LOX/LH2 engines results in a net reduction in the vehicle's physical size and structural mass. Other weight-saving features entail the extensive use of honeycomb sandwiches, advanced materials, and advanced fabrication techniques. The vehicle presented is utilized only as a means to study and identify various technologies needed in order to develop a low mass Earth-to-orbit transportation system for the future. The conclusion of this study is that vehicle geometry and structural/materials technology are critical to the development of efficient single-stage Earth-to-orbit transports.

Low-thrust chemical orbit to orbit propulsion system propellant management study

Abstract: Low thrust chemical propulsion systems were sized for transfer of large space systems from LEO to GEO. The influence of propellant combination, tankage and insulation requirements, and propellant management techniques on the LTPS mass and volume were studied. Liquid oxygen combined with hydrogen, methane or kerosene were the propellant combinations. Thrust levels of 445, 2230, and 4450 N were combined with 1, 4 and 8 perigee burn strategies. This matrix of systems was evaluated using multilayer insulation and spray-on-foam insulation systems. Various combinations of toroidal, cylindrical with ellipsoidal domes, and ellipsoidal tank shapes were investigated. Results indicate that low thrust (445 N) and single perigee burn approaches are considerably less efficient than the higher thrust level and multiple burn strategies. A modified propellant settling approach minimized propellant residuals and decreased system complexity, in addition, the toroid/ellipsoidal tank combination was predicted to be shortest.

Looking toward single-stage launch systems

Abstract: The prospects for utilization of single-stage-to-orbit launch vehicles to meet requirements for improved space transportation economics near the turn of the century are addressed. Vehicle concepts based on incremental or derivative advancements in technology are described. Comparative sizing and economic features of single and two-stage concepts are shown.

U. S. Air Force approach to plume contamination

Abstract: Exhaust products from rocket engine firings can produce undesirable effects on sensitivesatellite surfaces, such as optical systems, solar cells, and thermal control surfaces. TheAir Force has an objective of minimizing the effect of rocket plume contamination on space-craft mission effectiveness. Plume contamination can result from solid rocket motors, liq-uid propellant engines, and electric thrusters. To solve the plume contamination problem,the Air Force Rocket Propulsion Laboratory (AFRPL) has developed a plume contamination com-puter model which predicts the production, transport, and deposition of rocket exhaust pro- ducts. In addition, an experimental data base is being obtained through ground -based vacu-um chamber experiments and in- flight measurements with which to compare the analytical re- sults. Finally, the experimental data is being used to verify and improve the analyticalmodel. The plume contamination model, known as CONTAM, has been used to make contaminationpredictions for various engines. The experimental programs have yielded quantitative data,such as species concentrations and temperatures, in all regions of the plume. The resultof the modelling and experimental programs will ultimately be computer models which can beused by the satellite designer to analyze and to minimize the effect plume contaminationwill have on a particular spacecraft system.IntroductionSpacecraft contamination can not be attributed to one or even a few sources, but rathermust be attributed to all sources of migratory material which are foreign to the specificsystem of interest. Examples of potential contamination sources include outgassing of sol-id materials, the exhaust effluents of onboard propulsion systems, lubricants used formechanical devices, etc. The spacecraft designer must therefore be careful in choosing anymaterial for use on the spacecraft, basing his decision on the material properties and be-havior under vacuum conditions. In addition, proper care prior to launch must be consider-

Magnetic Propulsion for a Hypervelocity Launcher

Abstract: The concept of using electromagnetic forces to launch projectiles to high-velocity has been pursued since the laws of electro-motion were first derived by Oersted and Ampere in the early 1800’s. During the 1960’s, there were a number of attempts to develop practical hypervelocity accelerators using electro-magnetic forces. These investigations did not meet with notable success [1,2]. There has been greatly increased interest in electromagnetic propulsion in recent years for a variety of applications including hypervelocity weapons, meteoroid simulation at impact velocities above 10 km/s, high pressure shock physics, and space propulsion. This paper describes an electromagnetic launcher which was developed at the Australian National University (ANU) in the early 1970’s, [3] and was successfully used to accelerate projectiles to hypervelocities [4].

Hydromagnetics and future propulsion systems

Abstract: Futuristic hydromagnetic propulsion systems for spacecraft are examined with emphasis on systems that use regular coherent patterns of magnetic and electric fields of very high strength to interact with the structure of space-time to effect a quick translation from one space-time point to another. A particular type of this system is discussed: namely, the field resonance propulsion concept which will utilize superconducting magnets and a configuration of tunable free-electron lasers.

To encounter a star - the Solar Probe mission.

Abstract: Trajectory and technology requirements for the Solar Probe mission, whereby a spacecraft will attain a perihelion of a few solar radii, are discussed. Planetary gravity assist trajectories are considered that use Jupiter and possibly the earth, and that are needed for large payloads to be carried. Hardware technology development includes a lightweight thermal shield, a high energy propulsion module, and new power generation capability. Also discussed are a drag compensation system making the spacecraft drag free to a level of 10 to the -10th g, and a telecommunications system with doppler tracking accuracy of 0.1 mm/sec.

Electric propulsion comes of age

Abstract: The article examines several aspects of electric propulsion. It is shown that because the process and rate of energy transfer in electric systems are quite different from those of chemical propulsion systems, performance and orbit mechanics are intimately related. Attention is given to the several discrete subelements of an electric primary propulsion system, showing that while they have straightforward operating principles, they present challenges in implementation as working hardware. Finally, discussion covers applications which include a range of interplanetary and near-earth missions.

Low-thrust chemical rocket engine study

Abstract: Parametric data and preliminary designs on liquid rocket engines for low thrust cargo orbit-transfer-vehicles are described and those items where technology is required to enhance the designs are identified. The results of film cooling studies to establish the upper chamber pressure limit are given. The study showed that regen cooling with RP-1 was not feasible over the entire thrust and chamber pressure ranges. The thermal data showed that the RP-1 bulk temperature exceeded the study coking temperature limit of 1010 R. Based upon the results presented, O2/H2 and O2/CH4 regen engine systems and O2/H2 film cooled engines were selected for further study in the system analysis. Six engine design concepts are examined.

Electric propulsion technology

Abstract: The advanced electric propulsion program is directed towards lowering the specific impulse and increasing the thrust per unit of ion thruster systems In addition, electrothermal and electromagnetic propulsion technologies are being developed to attempt to fill the gap between the conventional ion thruster and chemical rocket systems Most of these new concepts are exagenous and are represented by rail accelerators, ablative Teflon thrusters, MPD arcs, Free Radicals, etc Endogenous systems such as metallic hydrogen offer great promise and are also being pursued

SPS attitude control and stationkeeping: Requirements and tradeoffs

Abstract: The dominant control requirements of solar power satellites change appreciably relative to small contemporary spacecraft. Trade studies and analyses illustrated preferred control approaches. It was found that the geosynchronous equatorial orbit is preferred over the alternative orbits considered, that the solar pressure orbit perturbation dominates stationkeeping propulsion requirements and that a combined AC and SK system using ion electric propulsion can satisfy the attitude control requirements. It was also found that control system/structural dynamic interaction stability can be obtained through frequency separation with reasonable structural dynamic requirements and simplify spacecraft design.

Mission Integration Study for Solid Teflon Pulsed Plasma Millipound Propulsion System.

Abstract: : The objective of this program is to identify, validate and assess advantages to the Air Force of using high performance pulsed plasma electric propulsion technology for satellite stationkeeping, attitude control, drag make-up, and solar pressure compensation and to develop pulsed plasma propulsion system/spacecraft integration and design criteria

Electric propulsion and power

Abstract: The development of electric propulsion systems is discussed and the benefits of these systems to various space mission requirements are outlined. The characteristics and development status of 8 and 30 cm mercury ion thrusters and solar electric propulsion systems are reported. In addition the advantages of an inert gas thruster for Earth orbital missions are examined and include its capability for operation at higher values of specific impulse, the ease at which it can be integrated with space systems, and it's low pollution potential.