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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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At least 559 records · Page 31
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Performance of a 15 cm ion thruster with reliable restart capability

Experimental evaluation of a modified SERT II thruster with dished grids capable of producing a 600 ma ion beam at 2690 sec specific impulse and achieving rapid, reliable startup is described. In addition to steady-state performance curves, thruster performance is examined over the duration of rapid startup and one hour operating sequence. Rapid startup utilizing high voltage spark ignition of the main and neutralizer cathodes is described. Minimal deterioration of the cathodes and the ignition system during 1000 neutralizer cathode startups and 2000 main cathode startups is demonstrated. The need for heaters and a vaporizer control system to facilitate more rapid startup is identified.

Wilbur, P. J.↗
NASA NTRS1973

Detonation propulsion for high pressure environments

One limitation encountered by chemical rocket propulsion in high pressure planetary atmospheres is illustrated by the conflict between the dependence of specific impulse on the ratio of chamber pressure to ambient pressure and the dependence of the motor structural design on the difference between the pressures. This work proposes to resolve the difficulty by employing detonating propellant in which the high pressures necessary for efficient propulsion are developed over a short time and need not be contained statically. Experimental results are presented to substantiate this claim of relatively high performance, together with an analytical development that approximately describes the flow dynamics.

Varsi, G.↗
NASA NTRS1973

Space Shuttle bipropellant RCS engine

A 'Reaction Control System' rocket engine is described which meets the Space Shuttle requirements. These include high performance/reliability and minimum weight with 100-mission life, with emphasis on reusability and minimum maintenance/servicing. The columbium fuel-vortex-cooled flight-type engine has a performance of 295 seconds vacuum specific impulse at 600 lbf thrust and a chamber pressure of 200 psia with maximum insulated-wall temperature below 2100 F. The engine has successfully demonstrated 9900 seconds operation including 6300 firing cycles without the need for maintenance.

Chazen, M. L.↗
NASA NTRS1973

New horizons in chemical propulsion

After a review of the work of the late-Fifties on free radicals for propulsion, it is concluded that atomic hydrogen would provide a potentially large increase in specific impulse. Work conducted to find an approach for isolating atomic hydrogen is considered. Other possibilities for obtaining propellants of greatly increased capability might be connected with the technology for the generation of activated states of gases, metallic hydrogen, fuels obtained from other planets, and laser transfer of energy.

Cohen, W.↗
NASA NTRS1974

Manufacture and deflagration of an atomic hydrogen propellant

It is observed that the use of very low temperatures (in the range from 0.1 to 1.5 K) produced by advanced cryogenic apparatus and the use of very strong magnetic fields (in the range from 50 to 100 kG) produced by superconducting magnets can yield a significant improvement in the atomic hydrogen trapping effectiveness of an H2 matrix. The use of a radioactive beta-ray emiter isotope may yield H-H2 propellants (with a specific impulse of about 740 sec) by secondary electron impact dissociations of H2 in an impregnated matrix maintained below 1 K in a strong magnetic field. Another method for manufacturing an H-H2 propellant involves bombardment of supercooled solid H2 with a cyclotron-produced beam of 10-MeV hydrogen atoms. The matrix-isolated atomic hydrogen must be used directly without prior melting as a solid propellant, and an analysis of the steady deflagration is presented.

Rosen, G.↗
NASA NTRS1971

Space shuttle main engine definition (phase B). Volume 2: Avionics

The advent of the space shuttle engine with its requirements for high specific impulse, long life, and low cost have dictated a combustion cycle and a closed loop control system to allow the engine components to run close to operating limits. These performance requirements, combined with the necessity for low operational costs, have placed new demands on rocket engine control, system checkout, and diagnosis technology. Based on considerations of precision environment, and compatibility with vehicle interface commands, an electronic control, makes available many functions that logically provide the information required for engine system checkout and diagnosis.

Source record↗
NASA NTRS1974

Platelet injectors for Space Shuttle orbit maneuvering engine

The Space Shuttle Orbit Maneuvering Subsystem Rocket Engine employs a platelet element injector concept. This injector has demonstrated 316-sec vacuum specific impulse performance under simulated altitude conditions when tested with a milled slot/electroformed nickel close-out regenerative chamber and a full 71 area ratio nozzle. To date, over 300 altitude engine tests and 300 stability bomb tests have demonstrated stable, erosion free operation with this concept to test durations of 150 seconds. The injector and chamber also meet the reusable requirements of the shuttle with a cycle life capability in excess of 1000 cycles. An extensive altitude restart program has also demonstrated OMS-engine operation over large variations in the burn and coast times with helium saturated propellants.

Kahl, R. C.↗
NASA NTRS1974

Status of 30 cm mercury ion thruster development

Two engineering model 30-cm ion thrusters have been assembled, calibrated, and qualification tested. This paper discusses the thruster design, performance, and power system. Test results include documentation of thrust losses due to doubly charged mercury ions and beam divergence by both direct thrust measurements and beam probes. Diagnostic vibration tests have led to improved designs of the thruster backplate structure, feed system, and harness. Thruster durability is being demonstrated over a thrust range of 97 to 113 mN at a specific impulse of about 2900 seconds. As of August 15, 1974, the thruster has successfully operated for over 4000 hours.

Sovey, J. S.↗
NASA NTRS1974

Planetary mission applications for space storable propulsion

This paper presents the results of a study to compare space-storable with earth-storable spacecraft propulsion systems, space-storable with solid kick stages, and several space-storable development options on the basis of benefits received for cost expenditures required. The results show that, for a launch vehicle with performance less than that of Shuttle/Centaur, space-storable spacecraft propulsion offers an incremental benefit/cost ratio between 1.0 and 5.5 when compared to earth-storable systems for three of the four missions considered. In the case of VOIR 83, positive benefits were apparent only for a specific launch vehicle-spacecraft propulsion combination. A space-storable propulsion system operating at thrust of 600 lbf, 355 units of specific impulse, and with blowdown pressurization, represents the best choice for the JO 81 mission on a Titan/Centaur if only spacecraft propulsion modifications are considered. For still higher performance, a new solid-propellant kick stage with space-storable spacecraft propulsion is preferred over a system which uses space-storable propellants for both the kick stage and the spacecraft system.

Chase, R. L.↗
NASA NTRS1975

Analytic solution for low thrust out of ecliptic mission design

Description of an analytic trajectory solution for low thrust out of ecliptic mission design based on the two-variable asymptotic expansion technique. A practical steering program utilizing normal thrusting symmetrical about the nodes is assumed. A two-term expansion yields an essentially exact solution of the problem which is in a form suitable for evaluation on small desk-top computers. It is shown that for a circular base orbit at 1 AU, a 15-kw solar electric system with a specific impulse of 3000 sec is capable of delivering an 1100-kg net payload to a heliographic latitude of about 63 deg.

Flandro, G. A.↗
NASA NTRS1975

Flow field characteristics and performance limitations of quasi-steady magnetoplasmadynamic accelerators

Measurements of terminal voltage, exhaust velocity, current and potential distributions in a quasi-steady MPD arc operating over a range of currents from 3 to 25 kA countervene previous indications of basic limitations on MPD performance. Terminal voltage fluctuations, previously used as an indicator of a unique limiting value of the parameter current squared over mass flow, are shown to be related to cathode surface area rather than to interelectrode plasma phenomena. With improved electrodes and insulators, satisfactory arc operation at specific impulses somewhat beyond previously proposed limits has been demonstrated, and the significant thrust producing mechanisms in such arcs have been identified.

Boyle, M. J.↗
NASA NTRS1975

One-millipound mercury ion thruster

A mercury ion thruster has been developed for efficient operation at the nominal 1-mlb thrust level with a specific impulse of about 3,000 sec and a total power consumption of about 120 W. At a beam voltage of 1,200 V and beam current of 72 mA, the discharge chamber operates with a propellant efficiency of 93.8% at an ion-generation energy of 276 eV/ion. The 8-cm diameter thruster advances proven component technology to assure the capability for thruster operation over an accumulated beam-on time in excess of 20,000 hours with a capability for 10,000 on-off duty cycles. Discharge chamber optimization has combined stable current-voltage characteristics with high performance efficiency by careful placement of the discharge cathode near the location of a magnetic-field zero just upstream of the thruster endplate.

Hyman, J., Jr.↗
NASA NTRS1975

Trajectory design of a solar electric propulsion Encke slow flyby mission

A slow flyby of the periodic comet, Encke, prior to its perihelion passage in 1980 has been shown to be a feasible mission for a solar electric propulsion spacecraft using a moderate power level of around 15 kw. The mission is accomplished in slightly less than two years with a planned miss distance and flyby speed of 1000 km and 4 km/s respectively. The results of a performance tradeoff study are presented together with a description of the nominal trajectory. The trajectories used in these mission studies were calculated using an updated model of the 30 cm mercury bombardment thruster's performance. The trajectory simulations include optimum staging of operating thrusters as a function of input power level constraints and limits on the variation of thruster specific impulse and efficiency as a function of thruster input power.

Sauer, C. G., Jr.↗
NASA NTRS1975

Gaseous fuel nuclear reactor research

Gaseous-fuel nuclear reactors are described; their distinguishing feature is the use of fissile fuels in a gaseous or plasma state, thereby breaking the barrier of temperature imposed by solid-fuel elements. This property creates a reactor heat source that may be able to heat the propellant of a rocket engine to 10,000 or 20,000 K. At this temperature level, gas-core reactors would provide the breakthrough in propulsion needed to open the entire solar system to manned and unmanned spacecraft. The possibility of fuel recycling makes possible efficiencies of up to 65% and nuclear safety at reduced cost, as well as high-thrust propulsion capabilities with specific impulse up to 5000 sec.

Schwenk, F. C.↗
NASA NTRS1975

Application of blast wave theory to explosive propulsion

An analysis was carried out by using blast wave theory to delineate the important aspects of detonating explosives in nozzles, such as flow and wave phenomena, characteristic length and time scales, and the parameters on which the specific impulse is dependent. The propulsive system utilizes the momentum of the ambient gas set into motion in the nozzle by the explosion. A somewhat simplified model was considered for the situation where the mass of ambient gas in the nozzle is much greater than the mass of gas produced in the explosion, a condition of interest for dense atmospheres, e.g., near the surface of Venus. Instantaneous detonation and energy release was presumed to occur at the apex of a conical nozzle, and the shock wave generated by the explosion was taken to propagate as a spherical wave, thereby setting the ambient gas in the nozzle into one-dimensional radially outward motion.

Back, L. H.↗
NASA NTRS1975

Improved planetary orbit mission capabilities in the Shuttle era by use of space-storable propulsion systems

Planetary-orbit mission capabilities will be greatly improved by the advent of space-storable retropropulsion systems with liquid fluorine and hydrazine as bipropellants even when the Shuttle/Tug is used as launch vehicle. Having a specific impulse as large as 375 sec, a multimission propulsion module using space-storable propellants with Pioneer or Mariner spacecraft as payload can perform Mercury, Saturn, and Uranus orbiter missions, and even some comet rendezvous missions, more cost effectively and flexibly than one using earth-storable propellants. It also can reduce trip time to the outer planets significantly. This paper presents mission requirements, technology status, system design and performance, development schedules, and costs based on data derived in a recent design and feasibility study.

Meissinger, H. F.↗
NASA NTRS1975

Comparative assessment of out-of-core nuclear thermionic power systems

The hardware selections available for fabrication of a nuclear electric propulsion stage for planetary exploration were explored. The investigation was centered around a heat-pipe-cooled, fast-spectrum nuclear reactor for an out-of-core power conversion system with sufficient detail for comparison with the in-core system studies completed previously. A survey of competing power conversion systems still indicated that the modular reliability of thermionic converters makes them the desirable choice to provide the 240-kWe end-of-life power for at least 20,000 full power hours. The electrical energy will be used to operate a number of mercury ion bombardment thrusters with a specific impulse in the range of about 4,000-5,000 seconds.

Estabrook, W. C.↗
NASA NTRS1976

Investigation of applications for high-power, self-critical fissioning uranium plasma reactors

Analytical studies were conducted to investigate potentially attractive applications for gaseous nuclear cavity reactors fueled by uranium hexafluoride and its decomposition products at temperatures of 2000 to 6000 K and total pressures of a few hundred atmospheres. Approximate operating conditions and performance levels for a class of nuclear reactors in which fission energy removal is accomplished principally by radiant heat transfer from the high temperature gaseous nuclear fuel to surrounding absorbing media were determined. The results show the radiant energy deposited in the absorbing media may be efficiently utilized in energy conversion system applications which include (1) a primary energy source for high thrust, high specific impulse space propulsion, (2) an energy source for highly efficient generation of electricity, and (3) a source of high intensity photon flux for heating working fluid gases for hydrogen production or MHD power extraction.

Rodgers, R. J.↗