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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Performance Measurements of Electric Solid Propellant in an Ablative Pulsed Electric Thruster

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. These propellants may also be used for electric in-space propulsion, specifically in the ablative pulsed plasma thruster. In this paper, we will investigate the performance of an electric solid propellant operating in an ablation-fed pulsed plasma device by use of an inverted pendulum micro-Newton thrust stand. Namely, the impulse-per-pulse and the specific impulse of the device using the electric solid propellant will be reported for test runs of 100 pulses and energy levels of 5, 10, 15 and 20 J. Further, the device will also be tested using the current state-of-the-art pulsed plasma thruster propellant, polytetrafluoroethylene. The performance of each propellant will be compared for each energy level using an identical setup and apparatus. This comparison of performance between propellants in a controlled setting will allow for better understanding of previous experimental observations.

Glascock, Matthew S.↗
NASA NTRS1992

The evolutionary development of high specific impulse electric thruster technology

Electric propulsion flight and technology demonstrations conducted primarily by Europe, Japan, China, the U.S., and the USSR are reviewed. Evolutionary mission applications for high specific impulse electric thruster systems are discussed, and the status of arcjet, ion, and magnetoplasmadynamic thrusters and associated power processor technologies are summarized.

Sovey, James S.↗
NASA NTRS1991

Electric thruster models for multimegawatt nuclear electric propulsion mission design

Three types of electric thrusters currently under development at JPL have potential to support future missions which utilize multimegawatt nuclear electric propulsion. These electric thrusters are the electron bombardment ion thruster, the magnetoplasmadynamic (MPD) thruster, and the electron-cyclotron-resonance (ECR) thruster. The electron bombardment ion thruster is a relatively mature technology which has been developed for operation at kilowatt power levels but will require new development for application in the multimegawatt regime. The MPD engine represents a technology which may be very well suited to steady-state multimegawatt applications but which has been limited to sub-scale (100's of kW) and pulsed (MW) testing thus far. The ECR plasma engine represents a class of very promising new concepts which are still in the basic research phase of development, but which may possess important fundamental advantages over other electric thruster technologies. Models of these thrusters are described and used to make projections of thrusters specific mass, efficiency, and power handling capacity for operation in the multimegawatt regime.

Leifer, Stephanie D.↗
NASA NTRS1992

The evolutionary development of high specific impulse electric thruster technology

Electric propulsion flight and technology demonstrations conducted in the USA, Europe, Japan, China, and USSR are reviewed with reference to the major flight qualified electric propulsion systems. These include resistojets, ion thrusters, ablative pulsed plasma thrusters, stationary plasma thrusters, pulsed magnetoplasmic thrusters, and arcjets. Evolutionary mission applications are presented for high specific impulse electric thruster systems. The current status of arcjet, ion, and magnetoplasmadynamic thrusters and their associated power processor technologies are summarized.

Sovey, James S.↗
NASA NTRS1973

On the Design and Test of a Liquid Injection Electric Thruster

A liquid injection electric thruster (LINJET) was designed and tested. The results of the tests were very encouraging with thruster performance levels well in excess of design goals. Supporting activities to the engine design and test included a five-million pulse life test on the main capacitor, a 46-million pulse test on the trigger electronics, design and fabrication of a zero resistance torque connector for use with the torsional pendulum thrust stand, design and fabrication of a logic box for control of engine firing, and a physical and chemical properties characterization of the perfluorocarbon propellant. While the results were encouraging, testing was limited, as many problems existed with the design. The most significant problem was involved with excessive propellant flow which contributed to false triggering and shorting. Low power active thermal control of the propellant storage cavity, coupled with a re-evaluation of the injection ring pore size and area exposed to the main capacitor discharge are areas that should be investigated should this design be carried forward.

Jones, T. A.↗
NASA NTRS2015

Development of Long-Lifetime Pulsed Gas Valves for Pulsed Electric Thrusters

It is advantageous for gas-fed pulsed electric thrusters to employ pulsed valves so propellant is only flowing to the device during operation. The propellant utilization of the thruster will be maximized when all the gas injected into the thruster is acted upon by the fields produced by the electrical pulse. Gas that is injected too early will diffuse away from the thruster before the electrical pulse can act to accelerate the propellant. Gas that is injected too late will miss being accelerated by the already-completed electrical pulse. As a consequence, the valve must open quickly and close equally quickly, only remaining open for a short duration. In addition, the valve must have only a small amount of volume between the sealing body and the thruster so the front and back ends of the pulse are as coincident as possible with the valve cycling, with very little latent propellant remaining in the feed lines after the valve is closed. For a real mission of interest, a pulsed thruster can be expected to pulse at least 10(exp 10) - 10(exp 11) times, setting the range for the number of times a valve must open and close. The valves described in this paper have been fabricated and tested for operation in an inductive pulsed plasma thruster (IPPT) for in-space propulsion. In general, an IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged, producing a high-current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed, it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. The valve characteristics needed for the IPPT application require a fast-acting valve capable of a minimum of 10(exp 10) valve actuation cycles. Since even 10(exp 9) cycles is well above anything demonstrated, this lower value was selected as the design point for the present work. The valve seal must remain leak-tight throughout operation, and the body must maintain a low internal leakage at relatively high operating temperatures. The full set of design requirements used for this program are summarized in Table 1. In this work, we describe two pulsed gas valves that have been fabricated to have long lifetime and demonstrate the characteristics listed above. The first is a miniaturized, conventional electromagnet-based valve while the second is a piezoelectric-based valve design. The conventional valve, shown in Fig. 1, is opened by use of a solenoid electromagnetic actuator. When current is applied to the solenoid coil, magnetic forces pull the plunger away from the valve seat, allowing fluid to flow through the valve. Removal of electrical current permits the spring and fluid pressure to seat the plunger, halting the flow of fluid. The valve body is fabricated from 304L corrosion resistant steel (CRES) and while the parts that form the magnetic circuit are fabricated from 430 CRES. This material does not have optimum magnetic properties, but its corrosion resistance permits incorporation into a design without requiring an additional plating process. A viton O-ring compound (Parker V0884-75), selected for its mechanical strength at elevated temperatures, was used for the valve seat seal. The design was based solely on the use of analytical sizing calculations, as opposed to a more rigorous finite element analysis. While this valve is small and relatively lightweight, it does not represent a design that is optimized for mass and/or a given volume envelope. The piezoelectric valve is a "puller" valve design. Applying a voltage to the piezo crystal causes it to elongate and pull a pintle off the seat, opening the valve. The valve seal consists of the pintle with an external, spherically-formed tip fabricated from Inconel 625 and the seat fabricated from Inconel 718 and possessing a spherical surface that permits the pintle to align with the valve body centerline after each cycle, maintaining the circumferential metal-on-metal seal. The contact stress on the valve pintle is expected in the worst case to be 96.5 MPa (14 ksi), which is well below the work hardening value of 413.7 MPa (60 ksi) for Inconel 615. Results presented in this work will compare the performance of each valve, both to each other and to the valve design requirements given in Table 1. In general, both valves meet all the performance envelope requirements, but the piezoelectric valve consumes considerably less power and has a faster response time over the conventional solenoid-driven valves.

Burkhardt, Wendel M.↗
NASA NTRS2005

Design, Fabrication, and Testing of Emissive Probes to Determine the Plasma Potential of the Plumes of Various Electric Thrusters

A significant problem in the use of electric thrusters in spacecraft is the formation of low-energy ions in the thruster plume. Low-energy ions are formed in the plume via random collisions between high-velocity ions ejected from the thruster and slow-moving neutral atoms of propellant effusing from the engine. The sputtering of spacecraft materials due to interactions with low-energy ions may result in erosion or contamination of the spacecraft. The trajectory of these ions is determined primarily by the plasma potential of the plume. Thus, accurate characterization of the plasma potential is essential to predicting low-energy ion contamination. Emissive probes were utilized to determine the plasma potential. When the ion and electron currents to the probe are balanced, the potential of such probes float to the plasma potential. Two emissive probes were fabricated; one utilizing a DC power supply, another utilizing a rectified AC power source. Labview programs were written to coordinate and automate probe motion in the thruster plume. Employing handshaking interaction, these motion programs were synchronized to various data acquisition programs to ensure precision and accuracy of the measurements. Comparing these experimental values to values from theoretical models will allow for a more accurate prediction of low-energy ion interaction.

Chen, Erinna M.↗
NASA NTRS2011

Non-Contact Thrust Stand Calibration Method for Repetitively-Pulsed Electric Thrusters

A thrust stand calibration technique for use in testing repetitively-pulsed electric thrusters for in-space propulsion has been developed and tested using a modified hanging pendulum thrust stand. In the implementation of this technique, current pulses are applied to a solenoidal coil to produce a pulsed magnetic field that acts against the magnetic field produced by a permanent magnet mounted to the thrust stand pendulum arm. The force on the magnet is applied in this non-contact manner, with the entire pulsed force transferred to the pendulum arm through a piezoelectric force transducer to provide a time-accurate force measurement. Modeling of the pendulum arm dynamics reveals that after an initial transient in thrust stand motion the quasisteady average deflection of the thrust stand arm away from the unforced or zero position can be related to the average applied force through a simple linear Hooke s law relationship. Modeling demonstrates that this technique is universally applicable except when the pulsing period is increased to the point where it approaches the period of natural thrust stand motion. Calibration data were obtained using a modified hanging pendulum thrust stand previously used for steady-state thrust measurements. Data were obtained for varying impulse bit at constant pulse frequency and for varying pulse frequency. The two data sets exhibit excellent quantitative agreement with each other as the constant relating average deflection and average thrust match within the errors on the linear regression curve fit of the data. Quantitatively, the error on the calibration coefficient is roughly 1% of the coefficient value.

Wong, Andrea R.↗
NASA NTRS1996

Hybrid Electric Thruster For Spacecraft

Report describes proposed hybrid electric spacecraft thruster combining best features of resistojet and arcjet, both of which proven spacecraft thrusters.

Pivirotto, Thomas J.↗
NASA NTRS1992

Characterization of electric thruster plumes using multiplexed laser induced fluorescence measurements

The use of laser-induced fluorescence to obtain spatially resolved measurements of propellant velocities and temperatures in electric thruster plumes is discussed, with emphasis on two innovations of the technique, namely simultaneous recording of the optogalvanic signal in a hollow cathode lamp for the purpose of calibrating Doppler shifts, and two-beam multiplexing to allow the measurement of two velocity components at once. It is also shown how information on plume fluctuations can be obtained from the multiplxed LIF data. The techniques are demonstrated on the plume from a low power arcjet, operated on argon, and its extension to the measurement of ion velocities in electrostatic ion thrusters and stationary plasma thrusters is discussed.

Ruyten, W. M.↗
NASA NTRS2015

Development of Long-Lifetime Pulsed Gas Valves for Pulsed Electric Thrusters

The design and test results for two types of pulsed gas valves are presented. The valves, a piezo valve and a solenoid actuated valve, must have exceedingly long lifetime to support gas-fed pulsed electric thruster operation for missions of interest. The performance of both valves was tested, with both demonstrating the capability to throttle the gas flow rate while maintaining low leakage levels below 10(exp -3) sccs of He at the beginning of valve lifetime. The piezo valve varies the flow rate by changing the amount that the valve is open, which is a function of applied voltage. This valve demonstrated continuous throttlability from 0-10 mL/s, with opening and closing times of 100 microsecond or less. The solenoid actuated valve flow rate changes as a function of the inlet gas pressure, with demonstrated flow rates in these tests from 2.7-11 mL per second. The valve response time is slower than the piezo valve, opening in 1-2 ms and closing in several ms. The solenoid actuated valve was tested to one million cycles, with the valve performance remaining relatively unchanged throughout the test. Galling of the sliding plunger caused the valve to bind and fail just after one million cycles, but at this point in the test the valve sealing surface leak rate still appeared to be well below the maximum target leak rake of 1×10(exp -3) sccs of He.

Burkhardt, Wendel M.↗
NASA NTRS1975

Factors in the design of spacecraft utilizing multiple electric thrusters

An analysis has been performed which describes the electrical interactions between simultaneously operating ion thrusters and between the ion thrusters and a spacecraft bus. The problem is reduced to a relatively simple equivalent circuit including all leakage paths. The influence of the various circuits elements on a spacecraft potential is studied. A thruster grounding scheme is introduced which is designed to minimize the electrical interactions and includes protection for both thrusters and spacecraft from possible damage due to neutralizer failure. The paper includes the results of various experiments which have been conducted to quantitatively define the elements of the equivalent circuit.

Fitzgerald, D. J.↗