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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Thruster array design approaches for a solar electric propulsion Encke Flyby mission

Design approaches are described and evaluated for a mercury electron-bombardment ion thruster array. Such an array might be used on a solar electric interplanetary spacecraft that obtains electrical energy from large solar panels. Thruster array designs are described and evaluated as they would apply to an Encke Flyby mission. Besides several well known approaches, a new concept utilizing individual two-axis gimbal actuators on each thruster is described and shown to have many structural and thermal advantages.

Ross, R. G., Jr.↗
NASA NTRS2003

Minimum impulse thruster valve design and development

The design and development of a minimum impulse thruster valve was conducted, by Moog, under contract by NASA's Jet Propulsion Laboratory, California Institute of Technology, for deep space propulsion systems. The effort was focused on applying known solenoid design techniques scaled to provide a 1 -millisecond response capability for monopropellant, hydrazine ACS thruster applications. The valve has an extended operating temperature range of 20(deg)F to +350(deg)F with a total mass of less than 25 grams and nominal power draw of 7 watts. The design solution resulted in providing a solenoid valve that is one-tenth the scale of the standard product line. The valve has the capability of providing a mass flow rate of 0.0009 pounds per second hydrazine. The design life of 1,000,000 cycles was demonstrated both dry and wet. Not all design factors scaled as expected and proved to be the focus of the final development effort. These included the surface interactions, hydrodynamics and driver electronics. The resulting solution applied matured design approaches to minimize the program risk with innovative methods to address the impacts of scale.

impulse thruster valve↗
NASA NTRS1992

Long life monopropellant hydrazine thruster evaluation for Space Station Freedom application - Test results

In support of propulsion system thruster development activity for Space Station Freedom (SSF), NASA Johnson Space Center (JSC) conducted a hydrazine thruster technology demonstration program. The goal of this program was to identify impulse life capability of state-of-the-art long life hydrazine thrusters nominally rated for 50 pounds thrust at 300 psia supply pressure. The SSF propulsion system requirement for impulse life of this thruster class is 1.5 million pounds-seconds, corresponding to a throughput of approximately 6400 pounds of propellant. Long life thrusters were procured from The Marquardt Company, Hamilton Standard, and Rocket Research Company, Testing at JSC was completed on the thruster designs to quantify life while simulating expected thruster firing duty cycles and durations for SSF. This paper presents a review of the SSF propulsion system hydrazine thruster requirements, summaries of the three long life thruster designs procured by JSC and acceptance test results for each thruster, the JSC thruster life evaluation test program, and the results of the JSC test program.

Popp, Christopher G.↗
NASA NTRS1991

Long life monopropellant hydrazine thruster evaluation for Space Station Freedom application

In support of propulsion system thruster development activity for Space Station Freedom (SSF), NASA Johnson Space Center (JSC) is conducting a hydrazine thruster technology demonstration program. The goal of this program is to identify impulse life capability of state-of-the-art long life hydrazine thrusters nominally rated for 50 pounds thrust at 300 psia supply pressure. The SSF propulsion system requirement for impulse life of this thruster class is 1.5 million pound-seconds, corresponding to a throughput of approximately 6400 pounds of propellant, with a high performance (234 pound-seconds per propellant pound). Long life thrusters were procured from Hamilton Standard, The Marquardt Company, and Rocket Research Company. Testing has initiated on the thruster designs to identify life while simulating expected thruster firing duty cycles and durations for SSF using monopropellant grade hydrazine. This paper presents a review of the SSF propulsion system and requirements as applicable to hydrazine thrusters, the three long life thruster designs procured by JSC and the resultant acceptance test data for each thruster, and the JSC test plan and facility.

Popp, Christopher G.↗
NASA NTRS2007

Design of a Low-Energy FARAD Thruster

The design of an electrodeless thruster that relies on a pulsed, rf-assisted discharge and electromagnetic acceleration using an inductive coil is presented. The thruster design is optimized using known performance,scaling parameters, and experimentally-determined design rules, with design targets for discharge energy, plasma exhaust velocity; and thrust efficiency of 100 J/pulse, 25 km/s, and 50%, respectively. Propellant is injected using a high-speed gas valve and preionized by a pulsed-RF signal supplied by a vector inversion generator, allowing for current sheet formation at lower discharge voltages and energies relative to pulsed inductive accelerators that do not employ preionization. The acceleration coil is designed to possess an inductance of at least 700 nH while the target stray (non-coil) inductance in the circuit is 70 nH. A Bernardes and Merryman pulsed power train or a pulse compression power train provide current to the acceleration coil and solid-state components are used to switch both powertrains.

Polzin, K. A.↗
NASA NTRS1982

J series thruster thermal test results

Test experience with J series ion thrusters have indicated that the present thruster design may result in excessive temperatures in areas which utilize organic materials such as wire insulation, with the resultant outgassing and potential contamination of insulating materials. Further, it appears that thermal data obtained with earlier thruster designs, such as the 700 series thruster, may not be directly applicable to the J series design. Two J series thrusters were fitted with thermocouples and critical temperatures measured for a variety of configurations and operating parameters. Completely enclosing the thruster to reduce facility contamination significantly increased temperatures prompting the selection of a compromise geometry for life testing. The operating parameter having the largest effect on temperatures was discharge power, while beam power affected little else than extraction system temperatures. Several off-normal operating modes were also investigated. Data believed to be sufficient to effectively modify existing thermal models were obtained from the tests.

Bechtel, R. T.↗
NASA NTRS1978

Prediction of plasma properties in mercury ion thrusters

A simplified theoretical model was developed which obtains to first order the plasma properties in the discharge chamber of a mercury ion thruster from basic thruster design and controllable operating parameters. The basic operation and design of ion thrusters is discussed, and the important processes which influence the plasma properties are described in terms of the design and control parameters. The conservation for mass, charge and energy were applied to the ion production region, which was defined as the region of the discharge chamber having as its outer boundary the surface of revolution of the innermost field line to intersect the anode. Mass conservation and the equations describing the various processes involved with mass addition and removal from the ion production region are satisfied by a Maxwellian electron density spatial distribution in that region.

Longhurst, G. R.↗
NASA NTRS1974

Status of 30 cm mercury ion thruster development

Two engineering model 30-cm ion thrusters were 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

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 NTRS1995

The 2.3 kW Ion Thruster Wear Test

A 30-cm diameter xenon ion thruster is under development at NASA to provide an ion propulsion option for auxiliary and primary propulsion on missions of national interest. Specific efforts include thruster design optimizations, component life testing and validation, and performance characterizations. Under this program, the ion thruster will be brought to engineering model development status. This paper describes the results of a 2.3-kW 2000-hour wear test performed to identify life limiting phenomena, measure the performance and characterize the operation of the thruster, and obtain wear, erosion, and surface contamination data. These data are being using as a data base for proceeding with additional life validation tests, and to provide input to flight thruster design requirements.

Parkes, James↗
NASA NTRS2001

NASA's Hall Thruster Program

NASA's Hall thruster program has base research and focused development efforts in support of the Advanced Space Transportation Program, Space-Based Program, and various other programs. The objective of the base research is to gain an improved understanding of the physical processes and engineering constraints of Hall thrusters to enable development of advanced Hall thruster designs. Specific technical questions that are current priorities of the base effort are: (1) How does thruster life vary with operating point? (2) How can thruster lifetime and wear rate be most efficiently evaluated? (3) What are the practical limitations for discharge voltage as it pertains to high specific impulse operation (high discharge voltage) and high thrust operation (low discharge voltage)? (4) What are the practical limits for extending Hall thrusters to very high input powers? and (5) What can be done during thruster design to reduce cost and integration concerns? The objective of the focused development effort is to develop a 50 kW-class Hall propulsion system, with a milestone of a 50 kW engineering model thruster/system by the end of program year 2006. Specific program wear 2001 efforts, along with the corporate and academic participation, are described.

Jankovsky, Robert S.↗
NASA NTRS1974

Study of monopropellants for electrothermal thrusters

A 333 mN electrothermal thruster designed to use MIL-grade hydrazine was demonstrated to be suitable for operation with low freezing point monopropellants containing hydrazine azide, monomethylhydrazine, unsymmetrical-dimethylhydrazine and ammonia. The steady-state specific impulse was greater than 200 sec for all propellants. The pulsed-mode specific impulse for an azide blend exceeded 175 sec for pulse widths greater than 50 msec; propellants containing carbonaceous species delivered 175 sec pulsed-mode specific impulses for pulse widths greater than 100 msec. Longer thrust chamber residence times were required for the carbonaceous propellants; the original thruster design was modified by increasing the characteristic chamber length and screen packing density. Specific recommendations were made for the work required to design and develop flight worthy thrusters, including methods to increase propellant dispersal at injection, thruster geometry changes to reduce holding power levels and methods to initiate the rapid decomposition of the carbonaceous propellants.

Kuenzly, J. D.↗
NASA NTRS1973

Performance of a modified downstream-cathode MPD thruster.

Discussion of the improvements incorporated since 1969 in a low-power MPD thruster design which is considered a potential candidate for satellite station keeping and attitude control. The improvements include a new cathode design, and changes in thruster geometry, with xenon used as the propellant. A better thermal design is found to be necessary for further improvement of the thruster.

Burkhart, J. A.↗
NASA NTRS1982

Operation of the J-series thruster using inert gas

Electron bombardment ion thrusters using inert gases are candidates for large space systems. The J-Series 30 cm diameter thruster, designed for operation up to 3 k-W with mercury, is at a state of technology readiness. The characteristics of operation with xenon, krypton, and argon propellants in a J-Series thruster with that obtained with mercury are compared. The performance of the discharge chamber, ion optics, and neutralizer and the overall efficiency as functions of input power and specific impulse and thruster lifetime were evaluated. As expected, the discharge chamber performance with inert gases decreased with decreasing atomic mass. Aspects of the J-Series thruster design which would require modification to provide operation at high power with insert gases were identified.

Rawlin, V. K.↗
NASA NTRS2006

Testing Fundamental Properties of Ionic Liquids for Colloid Microthruster Applications

NASA's New Millennium Program is scheduled to test a Disturbance Reduction System (DRS) on Space Technology 7 (ST7) as part of the European Space Agency's (ESA's) LISA Pathfinder Mission in late 2009. Colloid Micronewton Thrusters (CMNTs) will be used to counteract forces, mainly solar photon pressure, that could disturb gravitational reference sensors as part of the DRS. The micronewton thrusters use an ionic liquid, a room temperature molten salt, as propellant. The ionic liquid has a number of unusual properties that have a direct impact on thruster design. One of the most important issues is bubble formation before and during operation, especially during rapid pressure transitions from atmospheric to vacuum conditions. Bubbles have been observed in the feed system causing variations in propellant flow rate that can adversely affect thruster control. Bubbles in the feed system can also increase the likelihood that propellant will spray onto surfaces that can eventually lead to shorting high voltage electrodes. Two approaches, reducing the probability of bubble formation and removing bubbles with a new bubble eliminator device in the flow system, were investigated at Busek Co., Inc. and the Jet Propulsion Laboratory (JPL) to determine the effectiveness of both approaches. Results show that bubble formation is mainly caused by operation at low pressure and volatile contaminants in the propellant coming out of solution. A specification for the maximum tolerable level of contamination has been developed, and procedures for providing system cleanliness have been tested and implemented. The bubble eliminator device has also been tested successfully and has been implemented in recent thruster designs at Busek. This paper focuses on the propellant testing work at JPL, including testing of a breadboard level bubble eliminator device.

bubble formation↗
NASA NTRS1984

Ion thruster performance model

A model of ion thruster performance is developed for high flux density, cusped magnetic field thruster designs. This model is formulated in terms of the average energy required to produce an ion in the discharge chamber plasma and the fraction of these ions that are extracted to form the beam. The direct loss of high energy (primary) electrons from the plasma to the anode is shown to have a major effect on thruster performance. The model provides simple algebraic equations enabling one to calculate the beam ion energy cost, the average discharge chamber plasma ion energy cost, the primary electron density, the primary-to-Maxwellian electron density ratio and the Maxwellian electron temperature. Experiments indicate that the model correctly predicts the variation in plasma ion energy cost for changes in propellant gas (Ar, Kr and Xe), grid transparency to neutral atoms, beam extraction area, discharge voltage, and discharge chamber wall temperature. The model and experiments indicate that thruster performance may be described in terms of only four thruster configuration dependent parameters and two operating parameters. The model also suggests that improved performance should be exhibited by thruster designs which extract a large fraction of the ions produced in the discharge chamber, which have good primary electron and neutral atom containment and which operate at high propellant flow rates.

Brophy, J. R.↗