Augmenting Thrust With Waste Heat
According to NASA report, energy rejected by nuclear reactor adds to rocket thrust. With augmentation, specific impulse increased by as much as 23 percent over that of conventional engine.
Frisbee, R. H.↗
SEARCH · Engineering Papers
Results for “Specific impulse”
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Development of the Engineering Test Satellite-3 (ETS-3) ion engine system
The ion engine system onboard the ETS-3 is discussed. The system consists of two electron bombardment type mercury ion engines with 2 mN thrust and 2,000 sec specific impulse and a power conditioner with automatic control functions. The research and development of the system, development of its EM, PM and FM, the system test and the technical achievements leading up to final launch are discussed.
Kitamura, S.↗
Ion thruster project
The mercury bombardment electrostatic ion thruster is the most successful electric thruster available today. A 5 cm diameter ion thruster with 3,000 specific impulse and 5mN thrust is described. The advantages of electric propulsion and the tests that will be performed are also presented.
Perche, G. E.↗
SSME lifetime prediction and verification, integrating environments, structures, materials: The challenge
The planned missions for the space shuttle dictated a unique and technology-extending rocket engine. The high specific impulse requirements in conjunction with a 55-mission lifetime, plus volume and weight constraints, produced unique structural design, manufacturing, and verification requirements. Operations from Earth to orbit produce severe dynamic environments, which couple with the extreme pressure and thermal environments associated with the high performance, creating large low cycle loads and high alternating stresses above endurance limit which result in high sensitivity to alternating stresses. Combining all of these effects resulted in the requirements for exotic materials, which are more susceptible to manufacturing problems, and the use of an all-welded structure. The challenge of integrating environments, dynamics, structures, and materials into a verified SSME structure is discussed. The verification program and developmental flight results are included. The first six shuttle flights had engine performance as predicted with no failures. The engine system has met the basic design challenges.
Ryan, R. S.↗
Advanced OTV engine concepts
The results and status of engine technology efforts to date and related company funded activities are presented. Advanced concepts in combustors and injectors, high speed turbomachinery, controls, and high-area-ratio nozzles that package within a short length result is engines with specific impulse values 35 to 46 seconds higher than those now realized by operational systems. The improvement in life, reliability, and maintainability of OTV engines are important.
Zachary, A. T.↗
SSME thrust chamber simulation using Navier-Stokes equations
The capability of the PHOENICS fluid dynamics code in predicting two-dimensional, compressible, and reacting flow in the combustion chamber and nozzle of the space shuttle main engine (SSME) was evaluated. A non-orthogonal body fitted coordinate system was used to represent the nozzle geometry. The Navier-Stokes equations were solved for the entire nozzle with a turbulence model. The wall boundary conditions were calculated based on the wall functions which account for pressure gradients. Results of the demonstration test case reveal all expected features of the transonic nozzle flows. Of particular interest are the locations of normal and barrel shocks, and regions of highest temperature gradients. Calculated performance (global) parameters such as thrust chamber flow rate, thrust, and specific impulse are also in good agreement with available data.
Przekwas, A. J.↗
Ring-cusp ion thruster with shell anode
An improved ion thruster for low specific impulse operation in the 1500 sec to 6000 sec range has a multicusp boundary field provided by high strength magnets on an iron anode shell which lengthens the paths of electrons from a hollow cathode assembly. A downstream anode pole piece in the form of an iron ring supports a ring of magnets to provide a more uniform beam profile. A cylindrical cathode magnet can be moved selectively in an axial direction along a feed tube to produce the desired magnetic field at the cathode tip.
Sovey, J. S.↗
Ignition characterization of LOX/hydrocarbon propellants
The results of an evaluation of the ignition characteristics of the gaseous oxygen (Gox)/Ethanol propellant combination are presented. Ignition characterization was accomplished through the analysis, design, fabrication and testing of a spark initiated torch igniter and prototype 620 lbF thruster/igniter assembly. The igniter was tested over a chamber pressure range of 74 to 197 psia and mixture ratio range of 0.778 to 3.29. Cold (-92 to -165 F) and ambient (44 to 80 F) propellant temperatures were used. Spark igniter ignition limits and thruster steady state and pulse mode, performance, cooling and stability data are presented. Spark igniter ignition limits are presented in terms of cold flow pressure, ignition chamber diameter and mixture ratio. Thruster performance is presented in terms of vacuum specific impulse versus engine mixture ratio. Gox/Ethanol propellants were shown to be ignitable over a wide range of mixture ratios. Cold propellants were shown to have a minor effect on igniter ignition limits. Thruster pulse mode capability was demonstrated with multiple pulses of 0.08 sec duration and less.
Lawver, B. R.↗
Ignition characterization of LOX/hydrocarbon propellants
The results of an evaluation of the ignition characteristics of the gaseous oxygen (Gox)/Ethanol propellant combination are summarized. Ignition characterization was accomplished through the analysis, fabrication and testing of a spark initiated torch igniter and prototype 620 1bF thruster/igniter assembly. The igniter was tested over a chamber pressure range of 3.3 to 262 psia and a mixture ratio range of 0.4 to 40. The prototype thruster/igniter assembly was tested over the chamber pressure range of 74 to 197 psia and mixture ratio range of 0.778 to 3.29. Cold (-92 to -165 F) and ambient (44 to 80 F) propellant temperatures were used. Spark igniter ignition limits and thruster steady state and pulse mode, performance, cooling and stability data are presented. Spark igniter ignition limits are presented in terms of cold flow pressure, ignition chamber diameter and mixture ratio. Thruster performance is presented in terms of vacuum specific impulse versus engine mixture ratio. Gox/Ethanol propellants were shown to be ignitable over a wide range of mixture ratios. Cold propellants were shown to have a minor effect on igniter ignition limits. Thruster pulse mode capability was demonstrated with multiple pulses of 0.08 sec duration and less.
Lawver, B. R.↗
Performance capabilities of the 12-centimeter Xenon ion thruster
The 8- and 12-cm mercury ion thruster systems were developed primarily to provide N-S station keeping of satellites with masses up to about 1800 to 3600 kg respectively. The on-orbit propulsion requirements of recently proposed Large Space Systems (LSS) are beyond the thrust capabilities of the baseline 8- and 12-cm thruster systems. This paper presents a characterization of the performance capabilities of the 12-cm Xenon ion thruster to enable an evaluation of its application to LSS auxiliary propulsion requirements. With minor thruster modifications and simplifications the thrust was increased to 64 mN, a factor of six over the baseline 12-cm mercury thruster performance. The thruster was operated over a range of specific impulse of about 2000 to 4000 seconds and at total efficiencies up to 68.0 percent. The operating levels reached in this study were found to be close to the operating limits of the thruster design in terms of perveance, grid breakdown voltage and thruster component temperatures such as those of the magnets and cathode baffle.
Mantenieks, M.↗
Electrothermal thruster diagnostics
A flight-qualified electrothermal thruster demonstrated its adaptability to a variety of propellants. Originally qualified for operation with hydrazine propellant, it was operated with nitrogen, hydrogen, and ammonia propellants, demonstrating 73, 61, and 52 percent overall efficiency with these propellants, respectively, when tested over a wide range of operating conditions. By introducing a preheater to admit hot, rather than cold, propellant inlet gases to the thruster's augmentation heat exchanger, delivered specific impulse closer to theoretical performance limits should be achieved.
Zafran, S.↗
Experiments with a microwave electrothermal thruster concept
The design and test of two microwave electrothermal thrusters is described. The coaxial and cylindrical cavity concepts are tested in nitrogen gas with 200-2,000 W of 2.45 GHz input power. Experimental measurements and calculations of thrust, specific impulse and energy efficiency are presented for different gas flow rates and discharge pressures. Measured energy efficiencies varied between 10-60 percent and the performance compared favorably with other electrothermal thrusters operating in nitrogen gas. The experimental performance demonstrated the feasibility of the concept.
Asmussen, J.↗
The DE-AC accelerator system for high beam current density
A radically different ion accelerator system design has been developed to increase the extracted beam current density of ion thrusters. By independently controlling plasma ion extraction and acceleration, the Decoupled Extraction and Amplified Current (DE-AC) accelerator system gives higher beam currents at lower electric field stress levels. Results are presented for DE-AC grid system operation with the Test Bed Ion Engine. Using argon propellant, feasibility of a 60 mA/sq cm beam current density is shown. A discussion of power supply requirements, specific impulse operation and grid system geometrical design is presented also.
Aston, G.↗
Progress report - Advanced cryogenic OTV engine technology
New technologies for space-based, reusable, throttleable, cryogenic orbit transfer propulsion are being evaluated. A variable-thrust (200 to 3000 lbF), 2000 psi chamber pressure, LO2/LH2 engine has been selected to demonstrate the 20-hour, 500-restart life goal, and a specific impulse in excess of 480 lbF-sec/lbM. The results of recent vehicle-engine integration analyses and the progress in design, fabrication, and testing are provided. Emphasis is placed on the following technology areas being investigated in support of the advanced engine design: LOX hydrostatic bearings; burn-resistant materials for high-pressure GOX turbines and valves; high surface-low flux annular combustion chambers for the dual propellant expander cycle; improved cooling approaches for high-pressure combustion chambers, new concepts in integrated controls; and engine health diagnostics.
Schoenman, L.↗
Laser propulsion for the orbital transfer mission
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.
Frisbee, R. H.↗
An experimental data base for material selection and design of high-speed, high-pressure, oxygen turbomachinery
New technologies for space-based, reusable, throttleable, cryogenic, orbit transfer propulsion are being evaluated. A variable-thrust (200 to 3000 lbF), 2000 psi chamber pressure, LO2/LH2 engine has been selected to demonstrate the 20-hour, 500-restart life goal, and a specific impulse of 480 lbF-sec/lbM. The advanced design uses warm oxygen to power a fully integrated turbopump that delivers 4500 psi LO2. The selected engine cycle provides the following advantages over conventional expander cycle designs which use H2 to drive the LH2 and LO2 pumps; higher engine operating pressures and performance in a smaller envelope, lower turbine operating temperatures, elimination of interpropellant seals and purges, and an extended throttling range. The design approach and results of testing to characterize materials for use in high-pressure gaseous oxygen are discussed. Test methods include particle impact testing in a sonic-flow, hot GO2 streams and forced, high-speed friction rubbing testing. Materials are exposed to oxygen and nitrogen atmospheres at pressures of 100, 1000, and 3000 psi to identify separately the gas cooling effects from the metal oxidation effects at the rubbing surface. The selection of candidate design materials is based on an analytical parameter defined as the burn factor. Typical materials tested include carbon steel, stainless steel, nickel, copper, and monel alloys.
Schoenman, L.↗
Solar thermal propulsion for planetary spacecraft
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.
Sercel, J. C.↗
NASA electrothermal auxiliary propulsion technology
Electrothermal auxiliary propulsion systems provide high performance options which can have major mission benefits. There are several electrothermal concepts which offer a range of characteristics and benefits. Resistojets are the highest thrust to power option and are currently operational at mission average values of specific impulse, I sub sp approximately 295 sec. Long life, multipropellant resistojets are being developed for the space station, and resistojet technology advancements are being pursued to improve the I sub sp by more than 20 percent for resistojets used in satellite applications. Direct current arcjets have the potential of I sub sp over 400 sec with storable propellants and should provide over 1000 sec with hydrogen. Advanced concepts are being investigated to provide high power density options and possible growth to primary propulsion applications. Broad based experimental and analytical research and technology programs of NASA are summarized and recent significant advances are reviewed.
Stone, J. R.↗