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At least 865 records · Page 48
All sourcesNASA NTRSDOE OSTI
NASA NTRS2018

A Micro Newton Impulse-Bit Hydrazine Thruster - Design, Test, and Mission Applications

A small thruster, that will herein be called the Hydrazine milliNewton Thruster (HmNT), has been designed and vacuum hot-fire tested, demonstrating a range of thrust and impulse far below the current state-of-the-art (SOA) hydrazine thruster. Steady-state thrust levels of 35–135 mN, and minimum impulse-bits of 25–120 μN-s were demonstrated over a typical operating range of inlet pressures. These capabilities provide new spacecraft and mission design options, including the following: Functional backup for Reaction Wheel Assemblies (RWAs), which can be used to prolong the useful life of RWAs; RWA replacement for certain missions to save mass, power, and cost; Precision delta-V that can be used for: proximity operations and docking maneuvers, formation flying, primary attitude control for SmallSats and CubeSats, and primary delta-V for CubeSats.

Scharf, Daniel P.↗
NASA NTRSDate not supplied

Performance of the 12.5-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the performance test campaign of the 12.5 kW Advanced Electric Propulsion System (AEPS) Engineering Test Unit 2 (ETU-2) thruster produced by Aerojet Rocketdyne. This test campaign had three major goals: to assess design compliance with thruster requirements, provide a comparison to the previously-tested NASA Hall Effect Rocket with Magnet Shielding (HERMeS) Technology Demonstration Units (TDUs) from which the AEPS ETU design was derived, and support AEPS application on the Power and Propulsion Element (PPE). Assessments of ETU-2 performance were conducted at discharge powers of 2.6 to 13.1 kW, which encompasses the original AEPS throttle conditions as well as those most relevant for AEPS application on PPE. The results from the test campaign indicate that ETU-2 performance meets or exceeds all AEPS thruster performance requirements and matches the values and trends previously measured with the HERMeS TDUs. In addition, similar to the HERMeS TDUs, ETU-2 performance showed minimal variation over approximately 1636 h of operating time due to the incorporation of magnetic shielding. Taken together, these results show that the AEPS design is successfully replicating the performance of the HERMeS TDUs to within the empirical uncertainty and meeting its performance requirements as the design progresses to its Critical Design Review.

Jason D Frieman↗
NASA NTRS2021

Long-Duration Wear Testing of the ASTRAEUS Thruster, Phase I: 50 kg Xe Throughput

The thruster component of the Ascendant Sub-kW Transcelestial Electric Propulsion System successfully completed a propellant throughput demonstration of 50 kg Xe. The thruster was operated for a total of 3,287 h, of which 97.9% was spent at 300 V – 1000 W, and generated 790 kN-s of total impulse. Thrust, telemetry, and temperature measurements periodically recorded at four operating conditions spanning 200 – 500 V and 200 – 1000 W revealed constant performance across the full test duration. High-fidelity plasma simulations of the erosion of the thruster’s downstream faces suggest a theoretical lifetime capability of >30 kh, corresponding to propellant throughput of >450 kg Xe and a total impulse of >7 MNs, at the 300 V – 1000 W operating condition. The test was voluntarily paused after 50 kg Xe throughput to take these erosion measurements before the test was resumed at a different operating condition.

Goebel, Dan↗
NASA NTRS2021

The Dispersion of the Modified Two-Stream and Lower Hybrid Drift Instabilities in the Near-pole Plume of a Magnetically Shielded Hall Thruster

Recent analytical and numerical investigations have shown that instabilities in the lower hybrid frequency range can grow in the plume region downstream of the front magnet pole pieces of a magnetically shielded Hall thruster. Growth of the Modified Two-Stream (MTSI) and Lower Hybrid Drift (LHDI) instabilities was found with wave propagation in the r-z (or in-) plane and azimuthal (or out-of-plane) directions, respectively. In the analysis of the LHDI electron collisions were found to be insufficient to quench the instability in a region contained by the thruster ion beam and hollow cathode plume. The implication of the findings is that ions may be heated anomalously by these instabilities, in both the out-of-plane and in-plane directions, thereby altering the velocity field and enhancing the energy of those ions that strike nearby thruster surfaces. In this article we summarize the solution of the dispersion relation for each of these instabilities and provide additional results that can be compared directly with data from future plasma diagnostics, including with direct measurements of the wave dispersion.

Ortega, A. Lopez↗
NASA NTRS2021

Self-Consistent Implementation of a Zero-Equation Transport Model Into a Predictive Model for a Hall Effect Thruster

The performance of an axisymmetric multi-fluid Hall thruster code that incorporates a self-consistent, data-driven closure model for the anomalous electron transport is investigated. Five different operating conditions of the H9 magnetically shielded Hall thruster are simulated with the Jet Propulsion Laboratory’s Hall2De. In order to capture the inherent uncertainty associated with the closure model, O(100) simulations are run for each condition, each of using a coefficient set sampled randomly from a probability distribution. The results of these simulations provide probabilistic predictions of thruster performance quantities including thrust, and discharge current, as well as several component efficiencies and centerline plasma properties. The model is found to yield converged solutions at all conditions, with large 10 kHzrange oscillations and performance trends with voltage and flow rate similar to experiment. The model under-predicts the thrust by 15-25% and over-predicts the discharge current by 20% on average compared to experiments at the same discharge voltage and mass flow rate. This performance discrepancy is due to lower beam utilization, mass utilization, and divergence efficiency than experiment, resulting from high Hall parameters in the acceleration region, which lead to a protracted ion acceleration region. The physical processes underlying this result are discussed in the context of future data-driven modeling efforts.

Jorns, Benjamin A.↗
NASA NTRSDate not supplied

Deflection of Solar Energetic Protons by an Ion Thruster

While Galactic Cosmic Rays (GCR) represent a long-term cancer risk to deep space human missions, Solar Energetic Particles (SEP) represent a significant acute radiation hazard to human spaceflight. This work investigates wave-particle interactions caused by the incidence of a beam of energetic protons, representing a notional SEP event, onto a cloud of ions representing a notional ion thruster plume. The resulting electromagnetic ion/ion instability is developed to investigate the degree to which energetic protons can be scattered out of the initial beam. This work includes investigations of deflection rate dependences on incident SEP energy and flux, ion thruster power and orientation relative to SEP incidence, plume ion species, and background plasma. Also investigated are the spectrum and angular distribution of electromagnetic radiation resulting from the wave-particle interactions. The primary objective of this investigation is to develop a technique for active mitigation of SEP radiation threats to spacecraft and crew, by exploring the wave-particle interaction of a SEP event with the plasma plume expelled from an ion thruster. Since the cost in expended fuel restricts this technique to intermittent use during interplanetary flights, effective prediction of when a SEP event might impact the mission is still important, and the observation of conditions that might lead to SEP events, human-in-the-loop space weather analyses, and development of SEP prediction models remain essential to crew safety, spacecraft function, and mission success.

Interplanetary Space Flight↗
NASA NTRS1994

(abstract) Experimental Evaluation of Russian Anode Layer Thruster

This paper describes the results of a preliminary evaluation of a 1.35 kW anode layer thruster. Data on performance and plume characteristics will be presented. Data on thruster characteristics such as thrust, plume current densities, discharge voltage, discharge current, and mass flow rate along with the results of a 500-hour endurance test wear will be presented.

anode layer thrusters SPT thrusters propulsion end↗
NASA NTRS1996

A Robust Digital Autopilot for Spacecraft Equipped with Pulse-Operated Thrusters

The analysis and design of attitude control systems for spacecraft employing pulse-operated (on-off) thrusters is usually accomplished through a combination of modeling approximations and empirical techniques. In this paper a new thruster pulse-modulation scheme for pointing and tracking applications is developed from nonlinear control theory.

attitude control spacecraft thrusters control theo↗
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 NTRS2001

A theoretical analysis of vacuum arc thruster performance

In vacuum arc discharges the current is conducted through vapor evaporated from the cathode surface. In these devices very dense, highly ionized plasmas can be created from any metallic or conducting solid used as the cathode. This paper describes theoretical models of performance for several thruster configurations which use vacuum arc plasma sources. This analysis suggests that thrusters using vacuum arc sources can be operated efficiently with a range of propellant options that gives great flexibility in specific impulse. In addition, the efficiency of plasma production in these devices appears to be largely independent of scale because the metal vapor is ionized within a few microns of the cathode electron emission sites, so this approach is well-suited for micropropulsion.

thruster↗
NASA NTRS2004

Deep Space Mission Applications for NEXT: NASA's Evolutionary Xenon Thruster

NASA's Evolutionary Xenon Thruster (NEXT) is designed to address a need for advanced ion propulsion systems on certain future NASA deep space missions. This paper surveys seven potential missions that have been identified as being able to take advantage of the unique capabilities of NEXT. Two conceptual missions to Titan and Neptune are analyzed, and it is shown that ion thrusters could decrease launch mass and shorten trip time, to Titan compared to chemical propulsion. A potential Mars Sample return mission is described, and compassion made between a chemical mission and a NEXT based mission. Four possible near term applications to New Frontiers and Discovery class missions are described, and comparisons are made to chemical systems or existing NSTAR ion propulsion system performance. The results show that NEXT has potential performance and cost benefits for missions in the Discovery, New Frontiers, and larger mission classes.

low thrust trajectories↗
NASA NTRS2014

High Input Voltage Discharge Supply for High Power Hall Thrusters Using Silicon Carbide Devices

A power processing unit for a 15 kW Hall thruster is under development at NASA Glenn Research Center. The unit produces up to 400 VDC with two parallel 7.5 kW discharge modules that operate from a 300 VDC nominal input voltage. Silicon carbide MOSFETs and diodes were used in this design because they were the best choice to handle the high voltage stress while delivering high efficiency and low specific mass. Efficiencies in excess of 97 percent were demonstrated during integration testing with the NASA-300M 20 kW Hall thruster. Electromagnet, cathode keeper, and heater supplies were also developed and will be integrated with the discharge supply into a vacuum-rated brassboard power processing unit with full flight functionality. This design could be evolved into a flight unit for future missions that requires high power electric propulsion.

Electric Propulsion↗
NASA NTRS2014

Effect of Background Pressure on the Plasma Oscillation Characteristics of the HiVHAc Hall Thruster

During a component compatibility test of the NASA HiVHAc Hall thruster, a high-speed camera and a set of high-speed Langmuir probes were implemented to study the effect of varying facility background pressure on thruster operation. The results show a rise in the oscillation frequency of the breathing mode with rising background pressure, which is hypothesized to be due to a shortening accelerationionization zone. An attempt is made to apply a simplified ingestion model to the data. The combined results are used to estimate the maximum acceptable background pressure for performance and wear testing.

Hall Thrusters↗
NASA NTRS2016

Laser Induced Fluorescence Measurements in a Hall Thruster Plume as a Function of Background Pressure

A set of Laser Induced Fluorescence (LIF) measurements in the near-field region of the NASA- 173M Hall thruster plume is presented at four background pressure conditions varying from 9.4 x 10(exp -6) torr to 3.3 x 10(exp -5) torr. The xenon ion velocity distribution function was measured simultaneously along the axial and radial directions. An ultimate exhaust velocity of 19.6+/-0.25 km/s achieved at a distance of 20 mm was measured, and that value was not sensitive to pressure. On the other hand, the ion axial velocity at the thruster exit was strongly influenced by pressure, indicating that the accelerating electric field moved inward with increased pressure. The shift in electric field corresponded to an increase in measured thrust. Pressure had a minor effect on the radial component of ion velocity, mainly affecting ions exiting close to the channel inner wall. At that radial location the radial component of ion velocity was approximately 1000 m/s greater at the lowest pressure than at the highest pressure. A reduction of the inner magnet coil current by 0.6 A resulted in a lower axial ion velocity at the channel exit while the radial component of ion velocity at the channel inner wall location increased by 1300 m/s, and at the channel outer wall location the radial ion velocity remained unaffected. The ultimate exhaust velocity was not significantly affected by the inner magnet current.

Facility Effects↗
NASA NTRS2016

Post-Test Inspection of NASA's Evolutionary Xenon Thruster Long Duration Test Hardware: Ion Optics

A Long Duration Test (LDT) was initiated in June 2005 as a part of NASA's Evolutionary Xenon Thruster (NEXT) service life validation approach. Testing was voluntarily terminated in February 2014, with the thruster accumulating 51,184 hours of operation, processing 918 kg of xenon propellant, and delivering 35.5 MN-s of total impulse. The post-test inspection objectives for the ion optics were derived from the original NEXT LDT test objectives, such as service life model validation, and expanded to encompass other goals that included verification of in situ measurements, test issue root causes, and past design changes. The ion optics cold grid gap had decreased only by an average of 7% of pretest center grid gap, so efforts to stabilize NEXT grid gap were largely successful. The upstream screen grid surface exhibited a chamfered erosion pattern. Screen grid thicknesses were ≥ 86% of the estimated pretest thickness, indicating that the screen grid has substantial service life remaining. Deposition was found on the screen aperture walls and downstream surfaces that was primarily composed of grid material and back-sputtered carbon, and this deposition likely caused the minor decreases in screen grid ion transparency during the test. Groove depths had eroded through up to 35% of the accelerator grid thickness. Minimum accelerator aperture diameters increased only by about 5-7% of the pretest values and downstream surface diameters increased by about 24-33% of the pretest diameters. These results suggest that increasing the accelerator aperture diameters, improving manufacturing tolerances, and masking down the perforated diameter to 36 cm were successful in reducing the degree of accelerator aperture erosion at larger radii.

Ion Thruster↗
NASA NTRS2016

Post-Test Inspection of NASA's Evolutionary Xenon Thruster Long-Duration Test Hardware: Discharge Chamber

The NEXT Long-Duration Test is part of a comprehensive thruster service life assessment intended to demonstrate overall throughput capability, validate service life models, quantify wear rates as a function of time and operating condition, and identify any unknown life-limiting mechanisms. The test was voluntarily terminated in February 2014 after demonstrating 51,184 hours of high-voltage operation, 918 kg of propellant throughput, and 35.5 MN-s of total impulse. The post-test inspection of the thruster hardware began shortly afterwards with a combination of non-destructive and destructive analysis techniques, and is presently nearing completion. This paper presents relevant results of the post-test inspection for the discharge chamber as well as other miscellaneous components such as the high-voltage propellant isolators and electrical cabling. Comparison of magnetic field measurements taken during pretest and post-test inspections indicate that the field strength did not degrade, consistent with performance data obtained during the test. Inspection of discharge chamber mesh samples show a deposition coating primarily composed of grid material that is approximately 15 micrometers in thickness. This thickness is well within the retention capability of the mesh and is therefore not expected to present any issues. Approximately 3.1 grams of deposition flakes were found at the bottom of the discharge chamber, composed primarily of grid material and carbon. Calculated size histograms of these flakes indicate that 99% have a maximum dimension of 200 micrometers or smaller, which is significantly less than the ion optics grid gap. Larger flakes that are capable of causing a grid-to-grid short will be analyzed to determine if their formation will occur in flight or is a facility effect. The high-voltage propellant isolators as well as numerous other electrical insulators were inspected and no evidence of arcing or any other issues were found.

Electric Propulsion↗
NASA NTRS2016

Post-Test Inspection of NASA's Evolutionary Xenon Thruster Long-Duration Test Hardware: Discharge and Neutralizer Cathodes

The NEXT Long-Duration Test is part of a comprehensive thruster service life assessment intended to demonstrate overall throughput capability, validate service life models, quantify wear rates as a function of time and operating condition, and identify any unknown life-limiting mechanisms. The test was voluntarily terminated in April 2014 after demonstrating 51,184 hours of high-voltage operation, 918 kg of propellant throughput, and 35.5 MN-s of total impulse. The post-test inspection of the thruster hardware began shortly afterwards with a combination of non-destructive and destructive analysis techniques, and is presently nearing completion. This presentation presents relevant results of the post-test inspection for both discharge and neutralizer cathodes.

Lifetime↗