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  • Other Sources  (31)
  • SPACECRAFT PROPULSION AND POWER  (31)
  • 1990-1994  (31)
  • 1970-1974
  • 1994  (31)
  • 1
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    Performance testing of a fixed configuration microwave arcjet thruster (1994)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The microwave arcjet thruster uses microwave energy to create a free-floating plasma discharge within a microwave resonant cavity. This discharge typically absorbs 99% of the input power and converts it to thermal energy which is then transferred to the flowing propellant gas. Recent modifications have allowed the thruster to be operated in a fixed configuration where neither the cavity geometry nor the tuning mechanisms are adjusted. The prototype has demonstrated its ability to operate in this fixed configuration using a variety of propellant gases, i.e., nitrogen, helium, ammonia, and hydrogen. The current design is capable of efficient operation over a wide range of power levels (250 W to over 6000 W). Current work is focused on obtaining LIF velocimetry data of the velocity profile at the exit plane of the nozzle.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA Propulsion Engineering Research Center, Volume 2; p 75-79
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Small Satellite Propulsion Options (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Advanced chemical and low power electric propulsion offer attractive options for small satellite propulsion. Applications include orbit raising, orbit maintenance, attitude control, repositioning, and deorbit of both Earth-space and planetary spacecraft. Potential propulsion technologies for these functions include high pressure Ir/Re bipropellant engines, very low power arcjets, Hall thrusters, and pulsed plasma thrusters, all of which have been shown to operate in manners consistent with currently planned small satellites. Mission analyses show that insertion of advanced propulsion technologies enables and/or greatly enhances many planned small satellite missions. Examples of commercial, DoD, and NASA missions are provided to illustrate the potential benefits of using advanced propulsion options on small satellites.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-106701 , E-9063 , NAS 1.15:106701 , AIAA PAPER 94-2997 , Joint Propulsion Conference; Jun 27, 1994 - Jun 29, 1994; Indianapolis, IN; United States
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    NASA TECHNICAL REPORTS
  • 3
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    Power Electronics Development for the SPT-100 Thruster (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Russian electric propulsion technologies have recently become available on the world market. Of significant interest is the Stationary Plasma Thruster (SPT) which has a significant flight heritage in the former Soviet space program. The SPT has performance levels of up to 1600 seconds of specific impulse at a thrust efficiency of 0.50. Studies have shown that this level of performance is well suited for stationkeeping applications, and the SPT-100, with a 1.35 kW input power level, is presently being evaluated for use on Western commercial satellites. Under a program sponsored by the Innovative Science and Technology Division of the Ballistic Missile Defense Organization, a team of U.S. electric propulsion specialists observed the operation of the SPT-100 in Russia. Under this same program, power electronics were developed to operate the SPT-100 to characterize thruster performance and operation in the U.S. The power electronics consisted of a discharge, cathode heater, and pulse igniter power supplies to operate the thruster with manual flow control. A Russian designed matching network was incorporated in the discharge supply to ensure proper operation with the thruster. The cathode heater power supply and igniter were derived from ongoing development projects. No attempts were made to augment thruster electromagnet current in this effort. The power electronics successfully started and operated the SPT-100 thruster in performance tests at NASA Lewis, with minimal oscillations in the discharge current. The efficiency of the main discharge supply was measured at 0.92, and straightforward modifications were identified which could increase the efficiency to 0.94.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-106488 , E-8475 , NAS 1.15:106488 , IEPEC-93-044 , International Electric Propulsion Conference; Sep 13, 1993 - Sep 16, 1993; Seattle, WA; United States
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  • 4
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    A Low-Erosion Starting Technique for High-Performance Arcjets (1994)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The NASA arcjet program is currently sponsoring development of high specific impulse thrusters for next generation geosynchronous communications satellites (2 kW-class) and low-power arcjets for power limited spacecraft (approx. 0.5 kW-class). Performance goals in both of these efforts will require up to 1000 starts at propellant mass flow rates significantly below those used in state-of-the-art arcjet thruster systems (i.e., high specific power levels). Reductions in mass flow rate can lead to damaging modes of operation, particularly at thruster ignition. During the starting sequence, the gas dynamic force due to low propellant flow is often insufficient to rapidly push the arc anode attachment to its steady-state position in the diverging section of the nozzle. This paper describes the development and demonstration of a technique which provides for non-damaging starts at low steady-state flow rates. The technique employs a brief propellant pressure pulse at ignition to increase gas dynamic forces during the critical ignition/transition phase of operation. Starting characteristics obtained using both pressure-pulsed and conventional starting techniques were compared across a wide range of propellant flow rates. The pressure-pulsed starting technique provided reliable starts at mass flow rates down to 21 mg/s, typically required for 700 s specific impulse level operation of 2 kW thrusters. Following the comparison, a 600 start test was performed across a wide flow rate range. Post-test inspection showed minimal erosion of critical arcjet anode/nozzle surfaces.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-106627 , E-8921 , NAS 1.15:106627 , Joint Propulsion Conference; Jun 27, 1994 - Jun 29, 1994; Indianapolis, IN; United States
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  • 5
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    SOSPAC- SOLAR SPACE POWER ANALYSIS CODE (1994)
    In:  Other Sources
    Details
    Publication Date: 2011-08-24
    Description: The Solar Space Power Analysis Code, SOSPAC, was developed to examine the solar thermal and photovoltaic power generation options available for a satellite or spacecraft in low earth orbit. SOSPAC is a preliminary systems analysis tool and enables the engineer to compare the areas, weights, and costs of several candidate electric and thermal power systems. The configurations studied include photovoltaic arrays and parabolic dish systems to produce electricity only, and in various combinations to provide both thermal and electric power. SOSPAC has been used for comparison and parametric studies of proposed power systems for the NASA Space Station. The initial requirements are projected to be about 40 kW of electrical power, and a similar amount of thermal power with temperatures above 1000 degrees Centigrade. For objects in low earth orbit, the aerodynamic drag caused by suitably large photovoltaic arrays is very substantial. Smaller parabolic dishes can provide thermal energy at a collection efficiency of about 80%, but at increased cost. SOSPAC allows an analysis of cost and performance factors of five hybrid power generating systems. Input includes electrical and thermal power requirements, sun and shade durations for the satellite, and unit weight and cost for subsystems and components. Performance equations of the five configurations are derived, and the output tabulates total weights of the power plant assemblies, area of the arrays, efficiencies, and costs. SOSPAC is written in FORTRAN IV for batch execution and has been implemented on an IBM PC computer operating under DOS with a central memory requirement of approximately 60K of 8 bit bytes. This program was developed in 1985.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NPO-16855
    Format: text
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  • 6
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    Development of CFD model for augmented core tripropellant rocket engine (1994)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The Space Shuttle era has made major advances in technology and vehicle design to the point that the concept of a single-stage-to-orbit (SSTO) vehicle appears more feasible. NASA presently is conducting studies into the feasibility of certain advanced concept rocket engines that could be utilized in a SSTO vehicle. One such concept is a tripropellant system which burns kerosene and hydrogen initially and at altitude switches to hydrogen. This system will attain a larger mass fraction because LOX-kerosene engines have a greater average propellant density and greater thrust-to-weight ratio. This report describes the investigation to model the tripropellant augmented core engine. The physical aspects of the engine, the CFD code employed, and results of the numerical model for a single modular thruster are discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Alabama Univ., Research Reports: 1994 NASA(ASEE Summer Faculty Fellowship Program; 6 p
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  • 7
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    A preliminary analysis of low frequency pressure oscillations in hybrid rocket motors (1994)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Past research with hybrid rockets has suggested that certain motor operating conditions are conducive to the formation of pressure oscillations, or flow instabilities, within the motor combustion chamber. These combustion-related vibrations or pressure oscillations may be encountered in virtually any type of rocket motor and typically fall into three frequency ranges: low frequency oscillations (0-300 Hz); intermediate frequency oscillations (400-1000 Hz); and high frequency oscillations (greater than 1000 Hz). In general, combustion instability is characterized by organized pressure oscillations occurring at well-defined intervals with pressure peaks that may maintain themselves, grow, or die out. Usually, such peaks exceed +/- 5% of the mean chamber pressure. For hybrid motors, these oscillations have been observed to grow to a limiting amplitude which may be dependent on factors such as fuel characteristics, oxidizer injector characteristics, average chamber pressure, oxidizer mass flux, combustion chamber length, and grain geometry. The approach taken in the present analysis is to develop a modified chamber length, L, instability theory which accounts for the relationship between pressure and oxidizer to fuel concentration ratio in the motor.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Alabama Univ., Research Reports: 1994 NASA(ASEE Summer Faculty Fellowship Program; 6 p
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  • 8
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    Interface issues for on-board propulsion systems (1994)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Integration issues associated with the use of new chemical and electric propulsion technologies are a primary concern to the user community. Experience indicates that integration impacts must be addressed to the satisfaction of both spacecraft builders and operators prior to the acceptance of new propulsion systems. The NASA Lewis Research Center (LeRC) conducts an aggressive program to develop and transfer new propulsion technologies and this includes a major effort to identify and address integration issues associated with their use. This paper provides an overview of integration issues followed by a brief description of the spacecraft integration program at LeRC.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Pennsylvania State Univ., NASA Propulsion Engineering Research Center, Volume 2; p 88-92
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  • 9
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    A hydrogen-oxygen rocket engine coolant passage design program (RECOP) for fluid-cooled thrust chambers and nozzles (1994)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The design of coolant passages in regeneratively cooled thrust chambers is critical to the operation and safety of a rocket engine system. Designing a coolant passage is a complex thermal and hydraulic problem requiring an accurate understanding of the heat transfer between the combustion gas and the coolant. Every major rocket engine company has invested in the development of thrust chamber computer design and analysis tools; two examples are Rocketdyne's REGEN code and Aerojet's ELES program. In an effort to augment current design capabilities for government and industry, the NASA Lewis Research Center is developing a computer model to design coolant passages for advanced regeneratively cooled thrust chambers. The RECOP code incorporates state-of-the-art correlations, numerical techniques and design methods, certainly minimum requirements for generating optimum designs of future space chemical engines. A preliminary version of the RECOP model was recently completed and code validation work is in progress. This paper introduces major features of RECOP and compares the analysis to design points for the first test case engine; the Pratt & Whitney RL10A-3-3A thrust chamber.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Pennsylvania State Univ., NASA Propulsion Engineering Research Center, Volume 2; p 168-172
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  • 10
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    Active chlorine and nitric oxide formation from chemical rocket plume afterburning (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Chlorine and oxides of nitrogen (NO(x)) released into the atmosphere contribute to acid rain (ground level or low-altitude sources) and ozone depletion from the stratosphere (high-altitude sources). Rocket engines have the potential for forming or activating these pollutants in the rocket plume. For instance, H2/O2 rockets can produce thermal NO(x) in their plumes. Emphasis, in the past, has been placed on determining the impact of chlorine release on the stratosphere. To date, very little, if any, information is available to understand what contribution NO(x) emissions from ground-based engine testing and actual rocket launches have on the atmosphere. The goal of this work is to estimate the afterburning emissions from chemical rocket plumes and determine their local stratospheric impact. Our study focuses on the space shuttle rocket motors, which include both the solid rocket boosters (SRB's) and the liquid propellant main engines (SSME's). Rocket plume afterburning is modeled employing a one-dimensional model incorporating two chemical kinetic systems: chemical and thermal equilibria with overlayed nitric oxide chemical kinetics (semi equilibrium) and full finite-rate chemical kinetics. Additionally, the local atmospheric impact immediately following a launch is modeled as the emissions diffuse and chemically react in the stratosphere.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-197503 , NAS 1.26:197503
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