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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Laboratory-Model Integrated-System FARAD Thruster

Pulsed inductive plasma accelerators are spacecraft propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is electrodeless, inducing a plasma current sheet in propellant located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity (order of 10 km/s) through the interaction of the plasma current with an induced magnetic field. The Faraday Accelerator with RF-Assisted Discharge (FARAD) thruster [1,2] is a type of pulsed inductive plasma accelerator in which the plasma is preionized by a mechanism separate from that used to form the current sheet and accelerate the gas. Employing a separate preionization mechanism in this manner allows for the formation of an inductive current sheet at much lower discharge energies and voltages than those found in previous pulsed inductive accelerators like the Pulsed Inductive Thruster (PIT). In a previous paper [3], the authors presented a basic design for a 100 J/pulse FARAD laboratory-version thruster. The design was based upon guidelines and performance scaling parameters presented in Refs. [4, 5]. In this paper, we expand upon the design presented in Ref. [3] by presenting a fully-assembled and operational FARAD laboratory-model thruster and addressing system and subsystem-integration issues (concerning mass injection, preionization, and acceleration) that arose during assembly. Experimental data quantifying the operation of this thruster, including detailed internal plasma measurements, are presented by the authors in a companion paper [6]. The thruster operates by first injecting neutral gas over the face of a flat, inductive acceleration coil and at some later time preionizing the gas. Once the gas is preionized current is passed through the acceleration coil, inducing a plasma current sheet in the propellant that is accelerated away from the coil through electromagnetic interaction with the time-varying magnetic field. Neutral gas is injected over the face of the acceleration coil through a fast-acting valve that feeds a central distribution manifold. The thruster is designed to preionize the gas using an RF-frequency ringing signal produced by a discharging Vector Inversion Generator (VIG). The acceleration stage consists of a multiple-turn, multiple-strand spiral induction coil (see Fig. 1, left panel) and is designed for operation at discharge energies on the order of 100 J/pulse. Several different pulsed power train modules can be used to drive current through the acceleration coil. One such power train is based upon the Bernardes and Merryman circuit topology, which restricts voltage reversal on the capacitor banks and can be clamped to eliminate current reversal in the coil. A second option is a pulse-compression-ring power train (see Fig. 1, right panel), which takesa temporally broad, low current pulse and transforms it into a short, high current pulse.

Polzin, K.A.↗
NASA NTRS2007

Space Technology 7 : Micropropulsion and Mass Distribution

The NASA New Millennium Program Space Technology 7 (ST7) project will validate technology for precision spacecraft control. The ST7 disturbance reduction system (DRS) will contain new micropropulsion technology to be flown as part of the European Space Agency's LISA (laser interferometer space antenna) Pathfinder project. After launch into a low Earth orbit in early 2010, the LISA Pathfinder spacecraft will be maneuvered to a halo orbit about the Earth-Sun LI Lagrange point for operations. The DRS will control the position of the spacecraft relative to a reference to an accuracy of one nanometer over time scales of several thousand seconds. To perform the control the spacecraft will use a new colloid thruster technology. The thrusters will operate over the range of 5 to 30 micro-Newtons with precision of 0.1 micro-Newton. The thrust will be generated by using a high electric field to extract charged droplets of a conducting colloid fluid and accelerating them with a precisely adjustable voltage. The control position reference will be provided by the European LISA Technology Package, which will include two nearly free-floating test masses. The test mass position and attitude will be sensed and adjusted using electrostatic capacitance bridges. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass. The dynamic control system will cover eighteen degrees of freedom, six for each of the test masses and six for the spacecraft. In the absence of other disturbances, the test masses will slowly gravitate toward local concentrations of spacecraft mass. The test mass acceleration must be minimized to maintain the acceleration of the enclosing drag-free spacecraft within the control authority of the micropropulsion system. Therefore, test mass acceleration must be predicted by accurate measurement of mass distribution, then offset by the placement of specially shaped balance masses near each test mass. The - acceleration is characterized by calculating the gravitational effect of over ten million modeled points of a nearly 500-kg spacecraft. This paper provides an overview of the mission technology and the process of precision mass modeling of the DRS equipment.

ST-7↗
NASA NTRS2010

Reliability of High-Voltage Tantalum Capacitors. Parts 3 and 4)

Weibull grading test is a powerful technique that allows selection and reliability rating of solid tantalum capacitors for military and space applications. However, inaccuracies in the existing method and non-adequate acceleration factors can result in significant, up to three orders of magnitude, errors in the calculated failure rate of capacitors. This paper analyzes deficiencies of the existing technique and recommends more accurate method of calculations. A physical model presenting failures of tantalum capacitors as time-dependent-dielectric-breakdown is used to determine voltage and temperature acceleration factors and select adequate Weibull grading test conditions. This model is verified by highly accelerated life testing (HALT) at different temperature and voltage conditions for three types of solid chip tantalum capacitors. It is shown that parameters of the model and acceleration factors can be calculated using a general log-linear relationship for the characteristic life with two stress levels.

Teverovsky, Alexander↗
NASA NTRS2007

Plasma Measurements in an Integrated-System FARAD Thruster

Pulsed inductive plasma accelerators are spacecraft propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is electrodeless, inducing a current sheet in a plasma located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity (order of 10 km/s) through the interaction of the plasma current and the induced magnetic field. The Faraday Accelerator with RF-Assisted Discharge (FARAD) thruster[1,2] is a type of pulsed inductive plasma accelerator in which the plasma is preionized by a mechanism separate from that used to form the current sheet and accelerate the gas. Employing a separate preionization mechanism allows for the formation of an inductive current sheet at much lower discharge energies and voltages than those used in previous pulsed inductive accelerators like the Pulsed Inductive Thruster (PIT). A benchtop FARAD thruster was designed following guidelines and similarity performance parameters presented in Refs. [3,4]. This design is described in detail in Ref. [5]. In this paper, we present the temporally and spatially resolved measurements of the preionized plasma and inductively-accelerated current sheet in the FARAD thruster operating with a Vector Inversion Generator (VIG) to preionize the gas and a Bernardes and Merryman circuit topology to provide inductive acceleration. The acceleration stage operates on the order of 100 J/pulse. Fast-framing photography will be used to produce a time-resolved, global view of the evolving current sheet. Local diagnostics used include a fast ionization gauge capable of mapping the gas distribution prior to plasma initiation; direct measurement of the induced magnetic field using B-dot probes, induced azimuthal current measurement using a mini-Rogowski coil, and direct probing of the number density and electron temperature using triple probes.

Polzin, K. A.↗
NASA NTRS1966

Rectilinear accelerometer possesses self- calibration feature

Rectilinear accelerometer operates from an ac source with a phase-sensitive ac voltage output proportional to the applied accelerations. The unit includes an independent circuit for self-test which provides a sensor output simulating an acceleration applied to the sensitive axis of the accelerometer.

Henderson, R. B.↗
NASA NTRS1977

Baseline tests of the EVA contractor electric passenger vehicle

The EVA Contactor four door sedan, an electric passenger vehicle, was tested to characterize the state-of-the-art of electric vehicles. It is a four passenger sedan that was converted to an electric vehicle. It is powered by 16 series connected 6 volt electric vehicle batteries through a four step contactor controller actuated by a foot accelerator pedal. The controller changes the voltage applied to the separately excited DC motor. The braking system is a vacuum assisted hydraulic braking system. Regenerative braking was also provided.

Bozek, J. M.↗
NASA NTRS2002

Evaluation of Cycle Life and Characterization of YTP 45 Ah Li-Ion Battery for EMU

Li-ion batteries, with longer cycle life and higher energy density features, are now more and more attractive and applied in multiple fields. The YTP 45 Ah Li-ion battery has been evaluated here and may be employed in EMU in the future. Evaluations were on: (1) Cycle life tests - 500 cycles total (completed 40 cycles in simulated shuttle use mode and 460 cycles in an accelerated use mode, and recorded differential voltage of individual cell in battery); (2) Characterization test - discharge capacity measurement in environment temperature of -10, 25, 50 C before and after 500 cycles; and (3) Thermal testing - charge and discharge at 50 C and -10 C before and after 500 cycles. The battery showed less than a 9% drop of initial discharge capacity and energy within 500 cycles with 475 cycles 59% DOD plus 25 cycles 100% DOD. The EOD voltage ranged from 16.0 to 18.0 V, which fits the requirement for operating the EMU.

Deng, Yi↗
NASA NTRS2005

Space Technology 7 Disturbance Reduction System - precision control flight Validation

The NASA New Millennium Program Space Technology 7 (ST7) project will validate technology for precision spacecraft control. The Disturbance Reduction System (DRS) will be part of the European Space Agency's LISA Pathfinder project. The DRS will control the position of the spacecraft relative to a reference to an accuracy of one nanometer over time scales of several thousand seconds. To perform the control, the spacecraft will use a new colloid thruster technology. The thrusters will operate over the range of 5 to 30 micro-Newtons with precision of 0.1 micro- Newton. The thrust will be generated by using a high electric field to extract charged droplets of a conducting colloid fluid and accelerating them with a precisely adjustable voltage. The control reference will be provided by the European LISA Technology Package, which will include two nearly freefloating test masses. The test mass positions and orientations will be measured using a capacitance bridge. The test mass position and attitude will be adjustable using electrostatically applied forces and torques. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass. The dynamic control system will cover eighteen degrees of freedom: six for each of the test masses and six for the spacecraft. After launch in late 2009 to a low Earth orbit, the LISA Pathfinder spacecraft will be maneuvered to a halo orbit about the Earth-Sun L1 Lagrange point for operations.

microthrusters↗
NASA NTRS2007

Design of an Integrated-System FARAD Thruster

Pulsed inductive plasma accelerators are spacecraft propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is electrodeless, inducing a current s heet in a plasma located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity (order of 10 km/s) through the interaction of the plasma current and the induced magne tic field, The Faraday Accelerator with RF-Assisted Discharge (FARAD) thruster is a type of pulsed inductive plasma accelerator in which t he plasma is preionized by a mechanism separate from that used to for m the current sheet and accelerate the gas. Employing a separate preionization mechanism allows for the formation of an inductive current s heet at much lower discharge energies and voltages than those used in previous pulsed inductive accelerators like the Pulsed Inductive Thr uster (PIT). In this paper, we present the design of a benchtop FARAD thruster with all the subsystems (mass injection, preionization, and acceleration) integrated into a single unit. Design of the thruster follows the guidelines and similarity performance parameters presented elsewhere. The system is designed to use the ringing, RF-frequency s ignal produced by a discharging Vector Inversion Generator (VIG) to p reionize the gas. The acceleration stage operates on the order of 100 J/pulse and can be driven by several different pulsed powertrains. These include a simple capacitor coupled to the system, a Bernardes and Merryman configuration, and a pulsecompression circuit that takes a temporally broad, low current pulse and transforms it into a short, h igh current pulse. A set of applied magnetic field coils are integrated into the system to guide the preionized propellant as it spreads ov er the face of the inductive acceleration coil. The coils are operate d in a pulsed mode, and the thruster can be operated without using the coils to determine if there is a performance improvement gain realiz ed when an applied field is present.

Polzin, K.A.↗
NASA NTRS2014

A Reliability Model for Ni-BaTiO3-Based (BME) Ceramic Capacitors

The evaluation of multilayer ceramic capacitors (MLCCs) with base-metal electrodes (BMEs) for potential NASA space project applications requires an in-depth understanding of their reliability. The reliability of an MLCC is defined as the ability of the dielectric material to retain its insulating properties under stated environmental and operational conditions for a specified period of time t. In this presentation, a general mathematic expression of a reliability model for a BME MLCC is developed and discussed. The reliability model consists of three parts: (1) a statistical distribution that describes the individual variation of properties in a test group of samples (Weibull, log normal, normal, etc.), (2) an acceleration function that describes how a capacitors reliability responds to external stresses such as applied voltage and temperature (All units in the test group should follow the same acceleration function if they share the same failure mode, independent of individual units), and (3) the effect and contribution of the structural and constructional characteristics of a multilayer capacitor device, such as the number of dielectric layers N, dielectric thickness d, average grain size r, and capacitor chip size S. In general, a two-parameter Weibull statistical distribution model is used in the description of a BME capacitors reliability as a function of time. The acceleration function that relates a capacitors reliability to external stresses is dependent on the failure mode. Two failure modes have been identified in BME MLCCs: catastrophic and slow degradation. A catastrophic failure is characterized by a time-accelerating increase in leakage current that is mainly due to existing processing defects (voids, cracks, delamination, etc.), or the extrinsic defects. A slow degradation failure is characterized by a near-linear increase in leakage current against the stress time; this is caused by the electromigration of oxygen vacancies (intrinsic defects). The two identified failure modes follow different acceleration functions. Catastrophic failures follow the traditional power-law relationship to the applied voltage. Slow degradation failures fit well to an exponential law relationship to the applied electrical field. Finally, the impact of capacitor structure on the reliability of BME capacitors is discussed with respect to the number of dielectric layers in an MLCC unit, the number of BaTiO3 grains per dielectric layer, and the chip size of the capacitor device.

reliability↗
NASA NTRS2014

Evaluation of Commercial Automotive-Grade BME Capacitors

Three Ni-BaTiO3 ceramic capacitor lots with the same specification (chip size, capacitance, and rated voltage) and the same reliability level, made by three different manufacturers, were degraded using highly accelerated life stress testing (HALST) with the same temperature and applied voltage conditions. The reliability, as characterized by mean time to failure (MTTF), differed by more than one order of magnitude among the capacitor lots. A theoretical model based on the existence of depletion layers at grain boundaries and the entrapment of oxygen vacancies has been proposed to explain the MTTF difference among these BME capacitors. It is the conclusion of this model that reliability will not be improved simply by increasing the insulation resistance of a BME capacitor. Indeed, Ni-BaTiO3 ceramic capacitors with a smaller degradation rate constant K will always give rise to a longer reliability life

capacitor↗
NASA NTRS2014

Evaluation of Commercial Automotive-Grade BME Capacitors

Three Ni-BaTiO3 ceramic capacitor lots with the same specification (chip size, capacitance, and rated voltage) and the same reliability level, made by three different manufacturers, were degraded using highly accelerated life stress testing (HALST) with the same temperature and applied voltage conditions. The reliability, as characterized by mean time to failure (MTTF), differed by more than one order of magnitude among the capacitor lots. A theoretical model based on the existence of depletion layers at grain boundaries and the entrapment of oxygen vacancies has been proposed to explain the MTTF difference among these BME capacitors. It is the conclusion of this model that reliability will not be improved simply by increasing the insulation resistance of a BME capacitor. Indeed, Ni-BaTiO3 ceramic capacitors with a smaller degradation rate constant K will always give rise to a longer reliability life.

reliability↗
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 NTRS2010

Design of a Microwave Assisted Discharge Inductive Plasma Accelerator

A new plasma accelerator concept that employs electrodeless plasma preionization and pulsed inductive acceleration is presented. Preionization is achieved through an electron cyclotron resonance discharge that produces a weakly-ionized plasma at the face of a conical theta pinch-shaped inductive coil. The presence of the preionized plasma allows for current sheet formation at lower discharge voltages than those found in other pulsed inductive accelerators. The location of an electron cyclotron resonance discharge can be controlled through the design of the applied magnetic field in the thruster. A finite-element model of the magnetic field was used as a design tool, allowing for the implementation of an arrangement of permanent magnets that yields a small volume of preionized propellant at the coil face. This allows for current sheet formation at the face of the inductive coil, minimizing the initial inductance of the pulse circuit and maximizing the potential efficiency of the new accelerator.

Hallock, Ashley K.↗
NASA NTRS2010

Design of a Microwave Assisted Discharge Inductive Plasma Accelerator

The design and construction of a thruster that employs electrodeless plasma preionization and pulsed inductive acceleration is described. Preionization is achieved through an electron cyclotron resonance discharge that produces a weakly-ionized plasma at the face of a conical theta pinch-shaped inductive coil. The presence of the preionized plasma allows for current sheet formation at lower discharge voltages than those employed in other pulsed inductive accelerators that do not employ preionization. The location of the electron cyclotron resonance discharge is controlled through the design of the applied magnetic field in the thruster. Finite element analysis shows that there is an arrangement of permanent magnets that yields a small volume of resonant magnetic field at the coil face. Preionization in the resonant zone leads to current sheet formation at the coil face, which minimizes the initial inductance of the pulse circuit and maximizes the potential electrical efficiency of the accelerator. A magnet assembly was constructed around an inductive coil to provide structural support to the selected arrangement of neodymium magnets. Measured values of the resulting magnetic field compare favorably with the finite element model.

Hallock, Ashley K.↗
NASA NTRS1991

Power Supply

Maxwell Laboratories capacitor charging power supply is the first commercial spinoff from the NASA CCDS program - a consortia of industries and government establishments to accelerate development of ground and space based commercial applications of NASA technology. The power supply transforms and conditions large voltages to charge capacitors used in x-ray sources, medical accelerators, etc. It is lighter, more reliable, more compact and efficient. Originally developed for space lasers, its commercial potential was soon recognized.

Source record↗
NASA NTRS1973

The 20 kilovolt rocket borne electron accelerator

The accelerator system is a preprogrammed multi-voltage system capable of operating at a current level of 1/2 ampere at the 20 kilovolt level. The five major functional areas which comprise this system are: (1) Silver zinc battery packs; (2) the electron gun assembly; (3) gun control and opening circuits; (4) the telemetry conditioning section; and (5) the power conversion section.

Harrison, R.↗