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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2010

NEXT Ion Thruster Thermal Model

As the NEXT ion thruster progresses towards higher technology readiness, it is necessary to develop the tools that will support its implementation into flight programs. An ion thruster thermal model has been developed for the latest prototype model design to aid in predicting thruster temperatures for various missions. This model is comprised of two parts. The first part predicts the heating from the discharge plasma for various throttling points based on a discharge chamber plasma model. This model shows, as expected, that the internal heating is strongly correlated with the discharge power. Typically, the internal plasma heating increases with beam current and decreases slightly with beam voltage. The second is a model based on a finite difference thermal code used to predict the thruster temperatures. Both parts of the model will be described in this paper. This model has been correlated with a thermal development test on the NEXT Prototype Model 1 thruster with most predicted component temperatures within 5 to 10 C of test temperatures. The model indicates that heating, and hence current collection, is not based purely on the footprint of the magnet rings, but follows a 0.1:1:2:1 ratio for the cathode-to-conical-to-cylindrical-to-front magnet rings. This thermal model has also been used to predict the temperatures during the worst case mission profile that is anticipated for the thruster. The model predicts ample thermal margin for all of its components except the external cable harness under the hottest anticipated mission scenario. The external cable harness will be re-rated or replaced to meet the predicted environment.

VanNoord, Jonathan L.↗
NASA NTRS2007

Environmental Testing of the NEXT PM1R Ion Engine

The NEXT propulsion system is an advanced ion propulsion system presently under development that is oriented towards robotic exploration of the solar system using solar electric power. The subsystem includes an ion engine, power processing unit, feed system components, and thruster gimbal. The Prototype Model engine PM1 was subjected to qualification-level environmental testing in 2006 to demonstrate compatibility with environments representative of anticipated mission requirements. Although the testing was largely successful, several issues were identified including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration which led to abbreviated thermal testing, and generation of particulate contamination from manufacturing processes and engine materials. The engine was reworked to address most of these findings, renamed PM1R, and the environmental test sequence was repeated. Thruster functional testing was performed before and after the vibration and thermal-vacuum tests. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 Grms for 2 min in each of three axes. Thermal-vacuum testing included three thermal cycles from 120 to 215 C with hot engine re-starts. Thruster performance was nominal throughout the test program, with minor variations in a few engine operating parameters likely caused by facility effects. There were no significant changes in engine performance as characterized by engine operating parameters, ion optics performance measurements, and beam current density measurements, indicating no significant changes to the hardware as a result of the environmental testing. The NEXT PM1R engine and the breadboard gimbal were found to be well-designed against environmental requirements based on the results reported herein. The redesigned cathode heater terminations successfully survived the vibration environments. Based on the results of this test program and confidence in the engineering solutions available for the remaining findings of the first test program, specifically the particulate contamination, the hardware environmental qualification program can proceed with confidence

Snyder, John S.↗
NASA NTRS2007

NEXT Thruster Component Verification Testing

Component testing is a critical part of thruster life validation activities under NASA s Evolutionary Xenon Thruster (NEXT) project testing. The high voltage propellant isolators were selected for design verification testing. Even though they are based on a heritage design, design changes were made because the isolators will be operated under different environmental conditions including temperature, voltage, and pressure. The life test of two NEXT isolators was therefore initiated and has accumulated more than 10,000 hr of operation. Measurements to date indicate only a negligibly small increase in leakage current. The cathode heaters were also selected for verification testing. The technology to fabricate these heaters, developed for the International Space Station plasma contactor hollow cathode assembly, was transferred to Aerojet for the fabrication of the NEXT prototype model ion thrusters. Testing the contractor-fabricated heaters is necessary to validate fabrication processes for high reliability heaters. This paper documents the status of the propellant isolator and cathode heater tests.

Pinero, Luis R.↗
NASA NTRS1996

Solar and Heliospheric Observatory (SOHO) Flight Dynamics Simulations Using MATLAB (R)

This paper describes a study to verify onboard attitude control laws in the coarse Sun-pointing (CSP) mode by simulation and to develop procedures for operational support for the Solar and Heliospheric Observatory (SOHO) mission. SOHO was launched on December 2, 1995, and the predictions of the simulation were verified with the flight data. This study used a commercial off the shelf product MATLAB(tm) to do the following: Develop procedures for computing the parasitic torques for orbital maneuvers; Simulate onboard attitude control of roll, pitch, and yaw during orbital maneuvers; Develop procedures for predicting firing time for both on- and off-modulated thrusters during orbital maneuvers; Investigate the use of feed forward or pre-bias torques to reduce the attitude handoff during orbit maneuvers - in particular, determine how to use the flight data to improve the feed forward torque estimates for use on future maneuvers. The study verified the stability of the attitude control during orbital maneuvers and the proposed use of feed forward torques to compensate for the attitude handoff. Comparison of the simulations with flight data showed: Parasitic torques provided a good estimate of the on- and off-modulation for attitude control; The feed forward torque compensation scheme worked well to reduce attitude handoff during the orbital maneuvers. The work has been extended to prototype calibration of thrusters from observed firing time and observed reaction wheel speed changes.

Headrick, R. D.↗
NASA NTRS2001

MEMS-Based Satellite Micropropulsion Via Catalyzed Hydrogen Peroxide Decomposition

Micro-electromechanical systems (MEMS) techniques offer great potential in satisfying the mission requirements for the next generation of "micro-scale" satellites being designed by NASA and Department of Defense agencies. More commonly referred to as "nanosats", these miniature satellites feature masses in the range of 10-100 kg and therefore have unique propulsion requirements. The propulsion systems must be capable of providing extremely low levels of thrust and impulse while also satisfying stringent demands on size, mass, power consumption and cost. We begin with an overview of micropropulsion requirements and some current MEMS-based strategies being developed to meet these needs. The remainder of the article focuses the progress being made at NASA Goddard Space Flight Center towards the development of a prototype monopropellant MEMS thruster which uses the catalyzed chemical decomposition of high concentration hydrogen peroxide as a propulsion mechanism. The products of decomposition are delivered to a micro-scale converging/diverging supersonic nozzle which produces the thrust vector; the targeted thrust level approximately 500 N with a specific impulse of 140-180 seconds. Macro-scale hydrogen peroxide thrusters have been used for satellite propulsion for decades; however, the implementation of traditional thruster designs on a MEMS scale has uncovered new challenges in fabrication, materials compatibility, and combustion and hydrodynamic modeling. A summary of the achievements of the project to date is given, as is a discussion of remaining challenges and future prospects.

Hitt, Darren L.↗
NASA NTRS2007

Liquid Bismuth Propellant Management System for the Very High Specific Impulse Thruster with Anode Layer

Two prototype bismuth propellant feed systems were constructed and operated in conjunction with a propellant vaporizer. One system provided bismuth to a vaporizer using gas pressurization but did not include a means to measure the flow rate. The second system incorporated an electromagnetic pump to provide fine control of the hydrostatic pressure and a new type of in-line flow sensor that was developed for accurate, real-time measurement of the mass flow rate. High-temperature material compatibility was a driving design requirement for the pump and flow sensor, leading to the selection of Macor for the main body of both components. Posttest inspections of both components revealed no degradation of the material. The gas pressurization system demonstrated continuous pressure control over a range from zero to 200 torr. In separate proof-of-concept experiments, the electromagnetic pump produced a linear pressure rise as a function of current that compared favorably with theoretical pump pressure predictions, producing a pressure rise of 10 kPa at 30 A. Preliminary flow sensor operation indicated a bismuth flow rate of 6 mg/s with an uncertainty of plus or minus 6%. An electronics suite containing a real-time controller was successfully used to control the entire system, simultaneously monitoring all power supplies and performing data acquisition duties.

Polzin, K. A.↗
NASA NTRS2016

Electrospray Thruster Performance with Microfabricated Emitter Arrays and Indium Propellant

The feasibility of a microfabricated indium-fueled electrospray thruster with excellent performance was demonstrated. High efficiency electrospray thrusters with microfabricated components are under development for very compact, distributable propulsion systems that can be employed on both very small and large spacecraft with 10X improvement over SOA in mass, volume and specific impulse The critical components of this technology are the microfabricated emitter arrays and the capillary force driven indium feed system. Grey scale electron-beam lithography patterning and reactive ion dry etching provided the required micron-scale etch precision and uniformity across an array of 400 emitters in 1 cm2. Initial tests of single microfabricated silicon emitters demonstrated better performance than industry standard single emitters and the performance required to achieve 200 micronewtons when scaled up to a thruster with 400 emitters in 1 cm2. Arrays of emitters tested in a preliminary prototype Microfluidic Electrospray Propulsion (MEP) thruster assembly demonstrated stable performance at estimated thrust levels of 5, 50, and 120 µN at extraction voltages less than 4 kV. Current stability was within 0.15 % at 120 µN with only 1.5% of the emitter current collected by the extractor electrode. Post-test inspections revealed that more than 99% of the 400 emitters were electrospraying during the test. Specific impulse was estimated to be >3100 s from measurements of total charge and consumed indium mass at an emitter voltage of 1470 V. The results of this investigation suggest that microfabricated indium electrospray thruster technology is both feasible and capable of excellent performance as a highly compact microthruster.

Marrese-Reading, Colleen M.↗
NASA NTRS2010

More About N2O-Based Propulsion and Breathable-Gas Systems

A concept was evaluated of using nitrous oxide as (1) a monopropellant in thrusters for space suits and spacecraft and (2) a source of breathable gas inside space suits and spacecraft, both by exploiting the controlled decomposition of N2O into N2 and O2. Relative to one prior monopropellant hydrazine, N2O is much less toxic, yet offers comparable performance. N2O can be stored safely as a liquid at room temperature and unlike another prior monopropellant hydrogen peroxide does not decompose spontaneously. A prototype N2O-based thruster has been demonstrated. It has also been proposed to harness N2O-based thrusters for generating electric power and to use the N2 + O2 decomposition product as a breathable gas. Because of the high performance, safety, and ease of handling of N2O, it can be expected to be economically attractive to equip future spacecraft and space suits with N2O-based thrusters and breathable-gas systems.

Zubrin, Robert↗
NASA NTRS2019

NEXT Single String Integration Tests in Support of the Double Asteroid Redirection Test Mission

A system integration test has been performed utilizing a prototype model NEXT ion thruster, an engineering model power processing unit, and a laboratory model command and data handling system. The objectives of the test were to: a) verify that the integrated system meets performance requirements, b) demonstrate that the integrated system is functional across the anticipated thermal, power processor, and Xe propellant ranges for the DART mission, and to c) evaluate fault detection and operation of the command and data handling system. Measurements made during this test included: thruster performance, PPU input voltages, PPU electrical and thermal telemetry, software states, and fault flags. Additionally, a far-field electrostatic probe diagnostic was used to infer relative changes in the thrust vector across the various propellant flow splits. This manuscript presents the results of these tests, which include integrated ion propulsion system demonstrations of performance, details on the execution of DART flight algorithms, and software fault handling.

Thomas, Robert E.↗
NASA NTRS2019

NEXT Single String Integration Tests in Support of the Double Asteroid Redirection Test Mission

A system integration test has been performed utilizing a prototype model NEXT ion thruster, an engineering model power processing unit, and a laboratory model command and data handling system. The objectives of the test were to: a) verify that the integrated system meets performance requirements, b) demonstrate that the integrated system is functional across the anticipated thermal, power processor, and Xe propellant ranges for the DART mission, and to c) evaluate fault detection and operation of the command and data handling system. Measurements made during this test included: thruster performance, PPU input voltages, PPU electrical and thermal telemetry, software states, and fault flags. Additionally, a far-field electrostatic probe diagnostic was used to infer relative changes in the thrust vector across the various propellant flow splits. This manuscript presents the results of these tests, which include integrated ion propulsion system demonstrations of performance, details on the execution of DART flight algorithms, and software fault handling.

Thomas, Robert↗
NASA NTRS2014

Testing of Diode-Clamping in an Inductive Pulsed Plasma Thruster Circuit

Testing of a 5.5 kV silicon (Si) diode and 5.8 kV prototype silicon carbide (SiC) diode in an inductive pulsed plasma thruster (IPPT) circuit was performed to obtain a comparison of the resulting circuit recapture efficiency,eta(sub r), defined as the percentage of the initial charge energy remaining on the capacitor bank after the diode interrupts the current. The diode was placed in a pulsed circuit in series with a silicon controlled rectifier (SCR) switch, and the voltages across different components and current waveforms were collected over a range of capacitor charge voltages. Reverse recovery parameters, including turn-off time and peak reverse recovery current, were measured and capacitor voltage waveforms were used to determine the recapture efficiency for each case. The Si fast recovery diode in the circuit was shown to yield a recapture efficiency of up to 20% for the conditions tested, while the SiC diode further increased recapture efficiency to nearly 30%. The data presented show that fast recovery diodes operate on a timescale that permits them to clamp the discharge quickly after the first half cycle, supporting the idea that diode-clamping in IPPT circuit reduces energy dissipation that occurs after the first half cycle

Toftul, Alexandra↗
NASA NTRS2014

Design and Preliminary Testing Plan of Electronegative Ion Thruster

Electronegative ion thrusters are a new iteration of existing gridded ion thruster technology differentiated by their ability to produce and accelerate both positive and negative ions. The primary motivations for electronegative ion thruster development include the elimination of lifetime-limiting cathodes from a thruster system and the ability to generate appreciable thrust through the acceleration of both positive or negative-charged ions. Proof-of-concept testing of the PEGASES (Plasma Propulsion with Electronegative GASES) thruster demonstrated the production of positively and negatively-charged ions (argon and sulfur hexafluoride, respectively) in an RF discharge and the subsequent acceleration of each charge species through the application of a time-varying electric field to a pair of metallic grids similar to those found in gridded ion thrusters. Leveraging the knowledge gained through experiments with the PEGASES I and II prototypes, the MINT (Marshall's Ion-ioN Thruster) is being developed to provide a platform for additional electronegative thruster proof-of-concept validation testing including direct thrust measurements. The design criteria used in designing the MINT are outlined and the planned tests that will be used to characterize the performance of the prototype are described.

Schloeder, Natalie R.↗
NASA NTRS2016

Status of the Development of Flight Power Processing Units for the NASAs Evolutionary Xenon Thruster - Commercial (NEXT-C) Project

A pathfinder prototype unit and two flight power processing units (PPUs) are being developed by the Aerojet Rocketdyne Corporation in Redmond, Washington and ZIN Technologies in Cleveland, Ohio, in support of the NEXT-C Project. This project is being led by the NASA Glenn Research Center in Cleveland, Ohio, and will also yield two flight thrusters. This hardware is being considered to be provided as Government Furnished Equipment for the New Frontiers Program, and is applicable to a variety of planetary science missions and astrophysics science missions. The design of the NEXT-C PPU evolves from the hardware fabricated under the NEXT technology development project. The power processing unit operates from two sources: a wide input 80 to 160 V high-power bus and a nominal 28 V low-power bus. The unit includes six power supplies. Four power supplies (beam, accelerator, discharge, and neutralizer keeper) are needed for steady state operation, while two cathode heater power supplies (neutralizer and discharge) are utilized during thruster startup. The unit in total delivers up to 7 kW of regulated power to a single gridded-ion thruster. Significant modifications to the initial design include: high-power adaptive-delay control, upgrade of design to EEE-INST-002 compliance, telemetry accuracy improvements, incorporation of telemetry to detect plume-mode operation, and simplification of the design in select areas to improve manufacturability and commercialization potential. The project is presently in the prototype phase and preparing for qualification level environmental testing.

NASAaEuro(TM)s Evolutionary Xenon Thruster↗
NASA NTRS2015

Low-Mass, Low-Power Hall Thruster System

NASA is developing an electric propulsion system capable of producing 20 mN thrust with input power up to 1,000 W and specific impulse ranging from 1,600 to 3,500 seconds. The key technical challenge is the target mass of 1 kg for the thruster and 2 kg for the power processing unit (PPU). In Phase I, Busek Company, Inc., developed an overall subsystem design for the thruster/cathode, PPU, and xenon feed system. This project demonstrated the feasibility of a low-mass power processing architecture that replaces four of the DC-DC converters of a typical PPU with a single multifunctional converter and a low-mass Hall thruster design employing permanent magnets. In Phase II, the team developed an engineering prototype model of its low-mass BHT-600 Hall thruster system, with the primary focus on the low-mass PPU and thruster. The goal was to develop an electric propulsion thruster with the appropriate specific impulse and propellant throughput to enable radioisotope electric propulsion (REP). This is important because REP offers the benefits of nuclear electric propulsion without the need for an excessively large spacecraft and power system.

Pote, Bruce↗
NASA NTRS1987

Status of xenon ion propulsion technology

This paper describes a working-model xenon ion propulsion subsystem (XIPS) designed for north-south stationkeeping (NSSK) of 2500-kg-class geosynchronous communication satellites. The XIPS consists of a 25-cm-diameter laboratory-model thruster, a breadboard-model power supply, and a flight-prototype pressure regulator (the critical component of the pressure-regulated xenon feed system). With a thrust of 63.5 mN, specific impulse of 2800 sec, and thruster efficiency of 65 percent, the XIPS performance is believed to be the highest ever reported for an ion thruster operated at 1.3-kW input power. The XIPS power supply accepts an input power of about 1.4 kW from a 28- to 35-V bus and converts it into the seven outputs required for startup and operation of the thruster. The simplified power supply contains only about 500 parts and has demonstrated an unprecedented efficiency of 90 percent and a specific mass of about 8 kg/kW. The results of a highly successful wear-mechanism test in which the working-model XIPS was operated for 4350 hours and 3850 ON/OFF cycles are presented. These hours and cycles are equivalent to over ten years of NSSK on large communication satellites.

Beattie, J. R.↗
NASA NTRS2004

Ultra High Voltage Propellant Isolators and Insulators for JIMO Ion Thrusters

Within NASA's Project Prometheus, high specific impulse ion thrusters for electric propulsion of spacecraft for the proposed Jupiter Icy Moon Orbiter (JIMO) mission to three of Jupiter's moons: Callisto, Ganymede and Europa will require high voltage operation to meet mission propulsion. The anticipated approx.6,500 volt net ion energy will require electrical insulation and propellant isolation which must exceed that used successfully by the NASA Solar Electric Propulsion Technology Readiness (NSTAR) Deep Space 1 mission thruster by a factor of approx.6. Xenon propellant isolator prototypes that operate at near one atmosphere and prototypes that operate at low pressures (<100 Torr) have been designed and are being tested for suitability to the JIMO mission requirements. Propellant isolators must be durable to Paschen breakdown, sputter contamination, high temperature, and high voltage while operating for factors longer duration than for the Deep Space 1 Mission. Insulators used to mount the thrusters as well as those needed to support the ion optics have also been designed and are under evaluation. Isolator and insulator concepts, design issues, design guidelines, fabrication considerations and performance issues are presented. The objective of the investigation was to identify candidate isolators and insulators that are sufficiently robust to perform durably and reliably during the proposed JIMO mission.

Banks, Bruce A.↗
NASA NTRS1999

Nanosatellite Propulsion Development Program

Earth-orbiting nanosatellite constellations are a unique and exciting means toward fulfilling part of the mission of the Goddard Space Flight Center (GSFC). These constellations, which may consist of several hundred 10-kg spacecraft, present unique challenges in the area of propulsion. Many mission concepts require significant delta-v and attitude control capability to reside in the nanosatellites. In response to requirements from mission feasibility studies, such as the Magnetospheric Constellation study, the GSFC has initiated industry and government partnerships to develop enabling propulsion technologies. The largest challenge has been to meet the power constraints of nanosatellites. These power issues, combined with the high thrust required by many of the missions studied, have led the GSFC to concentrate its efforts on chemical propulsion technology. Electric propulsion technologies capable of performing efficiently at very low power are also of interest to the GSFC as potential candidates for nanosatellite formation flying missions. This paper provides the status of specific industrial or government partnerships undertaken by the GSFC to develop nano/micro propulsion components. Three specific technologies are described in detail: 1) Nanosatellite Solid Rocket Motor Prototype 2) Ultra-Low-Power Cold Gas Thruster for Spin-Axis Precession 3) Micro-Machined Solid-Propellant Gas Generators.

Gagosian, J. S.↗