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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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NASA NTRS1993

Low power pulsed MPD thruster system analysis and applications

Pulsed magnetoplasmadynamic (MPD) thruster systems were analyzed for application to solar-electric orbit transfer vehicles at power levels ranging from 10 to 40 kW. Potential system level benefits of pulsed propulsion technology include ease of power scaling without thruster performance changes, improved transportability from low power flight experiments to operational systems, and reduced ground qualification costs. Required pulsed propulsion system components include a pulsed applied-field MPD thruster, a pulse-forming network, a charge control unit, a cathode heater supply, and high speed valves. Mass estimates were obtained for each propulsion subsystem and spacecraft component using off-the-shelf technology whenever possible. Results indicate that for payloads of 1000 and 2000 kg pulsed MPD thrusters can reduce launch mass by between 1000 and 2500 kg over those achievable with hydrogen arcjets, which can be used to reduce launch vehicle class and the associated launch cost. While the achievable mass savings depends on the trip time allowed for the mission, cases are shown in which the launch vehicle required for a mission is decreased from an Atlas IIAS to an Atlas I or Delta 7920.

Myers, Roger M.↗

NASA NTRS1977

Dynamics and control simulation of the Spacelab Experiment Pointing Mount

Computer simulations were developed to evaluate the performance of four Experiment Pointing Mounts (EPM) being considered for Spacelab experiments in the 1980-1990 time frame. The system modeled compromises a multibody system consisting of the shuttle, a mechanical isolation device, the EPM, celestial and inertial sensors, bearings, gimbal torque motors and associated nonlinearities, the experiment payload, and control and estimator algorithms. Each mount was subjected to a common disturbance (shuttle vernier thruster firing and man push off) and command (stellar pointing or solar raster scan) input. The fundamental limitation common to all mounts was found to be sensor noise. System dynamics and hardware nonlinearities have secondary effects on pointing performance for sufficiently high bandwidth.

Marsh, E. L.↗

NASA NTRS1990

Multimegawatt electric propulsion system design considerations

Piloted Mars mission requirements of relatively short trip times and low initial mass in earth orbit as identified by the NASA Space Exploration Initiative, indicate the need for multimegawatt electric propulsion systems. The design considerations and results for two thruster types, the argon ion and hydrogen magnetoplasmadynamic thrusters, are addressed in terms of configuration, performance, and mass projections. Preliminary estimates of Power Management and Distribution for these systems are given. Some assessment of these systems' performance in a reference Space Exploration Initiative piloted mission are discussed. Research and development requirements of these systems are also described.

Gilland, J. H.↗

NASA NTRS1991

Multimegawatt electric propulsion system design considerations

Piloted Mars Mission Requirements of relatively short trip times and low initial mass in Earth orbit as identified by the NASA Space Exploration Initiative, indicate the need for multimegawatt electric propulsion systems. The design considerations and results for two thruster types, the argon ion, and hydrogen magnetoplasmadynamic thrusters, are addressed in terms of configuration, performance, and mass projections. Preliminary estimates of power management and distribution for these systems are given. Some assessment of these systems' performance in a reference Space Exploration Initiative piloted mission are discussed. Research and development requirements of these systems are also described.

Gilland, J. H.↗

NASA NTRS2009

Magnetospheric Multiscale (MMS) Mission Commissioning Phase Orbit Determination Error Analysis

The Magnetospheric MultiScale (MMS) mission commissioning phase starts in a 185 km altitude x 12 Earth radii (RE) injection orbit and lasts until the Phase 1 mission orbits and orientation to the Earth-Sun li ne are achieved. During a limited time period in the early part of co mmissioning, five maneuvers are performed to raise the perigee radius to 1.2 R E, with a maneuver every other apogee. The current baseline is for the Goddard Space Flight Center Flight Dynamics Facility to p rovide MMS orbit determination support during the early commissioning phase using all available two-way range and Doppler tracking from bo th the Deep Space Network and Space Network. This paper summarizes th e results from a linear covariance analysis to determine the type and amount of tracking data required to accurately estimate the spacecraf t state, plan each perigee raising maneuver, and support thruster cal ibration during this phase. The primary focus of this study is the na vigation accuracy required to plan the first and the final perigee ra ising maneuvers. Absolute and relative position and velocity error hi stories are generated for all cases and summarized in terms of the ma ximum root-sum-square consider and measurement noise error contributi ons over the definitive and predictive arcs and at discrete times inc luding the maneuver planning and execution times. Details of the meth odology, orbital characteristics, maneuver timeline, error models, and error sensitivities are provided.

Chung, Lauren R.↗

NASA NTRS2011

MW-Class Electric Propulsion System Designs for Mars Cargo Transport

Multi-kilowatt electric propulsion systems are well developed and have been used on commercial and military satellites in Earth orbit for several years. Ion and Hall thrusters have also propelled robotic spacecraft to encounters with asteroids, the Moon, and minor planetary bodies within the solar system. High power electric propulsion systems are currently being considered to support piloted missions to near earth asteroids, as cargo transport for sustained lunar or Mars exploration, and for very high-power piloted missions to Mars and the outer planets. Using NASA Mars Design Architecture 5.0 as a reference, a preliminary parametric analysis was performed to determine the suitability of a nuclear powered, MW-class electric propulsion system for Mars cargo transport. For this initial analysis, high power 100-kW Hall thrusters and 250-kW VASIMR engines were separately evaluated to determine optimum vehicle architecture and estimated performance. The DRA 5.0 cargo mission closed for both propulsion options, delivering a 100 t payload to Mars orbit and reducing the number of heavy lift launch vehicles from five in the baseline DRA 5.0 architecture to two using electric propulsion. Under an imposed single engine-out mission success criteria, the VASIMR system took longer to reach Mars than did the Hall system, arising from the need to operate the VASIMR thrusters in pairs during the spiral out from low Earth orbit.

Gilland, James H.↗

NASA NTRS1988

The identification of a distributed parameter model for a flexible structure

A computational method is developed for the estimation of parameters in a distributed model for a flexible structure. The structure we consider (part of the RPL experiment) consists of a cantilevered beam with a thruster and linear accelerometer at the free end. The thruster is fed by a pressurized hose whose horizontal motion effects the transverse vibration of the beam. The Euler-Bernoulli theory is used to model the vibration of the beam and treat the hose thruster assembly as a lumped or point mass dashpot spring system at the tip. Measurements of linear acceleration at the tip are used to estimate the hose parameters (mass, stiffness, damping) and a Voigt-Kelvin viscoelastic structural damping parameter for the beam using a least squares fit to the data. Spline based approximations are considered to the hybrid (coupled ordinary and partial differential equations) systems; theoretical convergence results and numerical studies with both simulation and actual experimental data obtained from the structure are presented and discussed.

Banks, H. T.↗

NASA NTRS1986

The identification of a distributed parameter model for a flexible structure

A computational method is developed for the estimation of parameters in a distributed model for a flexible structure. The structure we consider (part of the RPL experiment) consists of a cantilevered beam with a thruster and linear accelerometer at the free end. The thruster is fed by a pressurized hose whose horizontal motion effects the transverse vibration of the beam. The Euler-Bernoulli theory is used to model the vibration of the beam and treat the hose-thruster assembly as a lumped or point mass-dashpot-spring system at the tip. Using measurements of linear acceleration at the tip, it is estimated that the parameters (mass, stiffness, damping) and a Voight-Kelvin viscoelastic structural damping parameter for the beam using a least squares fit to the data. Spline based approximations to the hybrid (coupled ordinary and partial differential equations) system are considered; theoretical convergence results and numerical studies with both simulation and actual experimental data obtained from the structure are presented and discussed.

Banks, H. T.↗

NASA NTRS1996

Proportional and scale change models to project failures of mechanical components with applications to space station

In this paper we develop the mathematical theory of proportional and scale change models to perform reliability analysis. The results obtained will be applied for the Reaction Control System (RCS) thruster valves on an orbiter. With the advent of extended EVA's associated with PROX OPS (ISSA & MIR), and docking, the loss of a thruster valve now takes on an expanded safety significance. Previous studies assume a homogeneous population of components with each component having the same failure rate. However, as various components experience different stresses and are exposed to different environments, their failure rates change with time. In this paper we model the reliability of a thruster valves by treating these valves as a censored repairable system. The model for each valve will take the form of a nonhomogeneous process with the intensity function that is either treated as a proportional hazard model, or a scale change random effects hazard model. Each component has an associated z, an independent realization of the random variable Z from a distribution G(z). This unobserved quantity z can be used to describe heterogeneity systematically. For various models methods for estimating the model parameters using censored data will be developed. Available field data (from previously flown flights) is from non-renewable systems. The estimated failure rate using such data will need to be modified for renewable systems such as thruster valve.

Taneja, Vidya S.↗

NASA NTRS2008

Dynamic Method for Identifying Collected Sample Mass

G-Sample is designed for sample collection missions to identify the presence and quantity of sample material gathered by spacecraft equipped with end effectors. The software method uses a maximum-likelihood estimator to identify the collected sample's mass based on onboard force-sensor measurements, thruster firings, and a dynamics model of the spacecraft. This makes sample mass identification a computation rather than a process requiring additional hardware. Simulation examples of G-Sample are provided for spacecraft model configurations with a sample collection device mounted on the end of an extended boom. In the absence of thrust knowledge errors, the results indicate that G-Sample can identify the amount of collected sample mass to within 10 grams (with 95-percent confidence) by using a force sensor with a noise and quantization floor of 50 micrometers. These results hold even in the presence of realistic parametric uncertainty in actual spacecraft inertia, center-of-mass offset, and first flexibility modes. Thrust profile knowledge is shown to be a dominant sensitivity for G-Sample, entering in a nearly one-to-one relationship with the final mass estimation error. This means thrust profiles should be well characterized with onboard accelerometers prior to sample collection. An overall sample-mass estimation error budget has been developed to approximate the effect of model uncertainty, sensor noise, data rate, and thrust profile error on the expected estimate of collected sample mass.

Carson, John↗

NASA NTRS1976

Similarity solution of the boundary-layer equations for laser heated flows

The laminar boundary-layer equations with local similarity approximation are solved in order to estimate the heat transfer rate to the wall in a laser-heated rocket thruster using pure hydrogen, where the average temperature of the hot plasma core is about 14,000 K and the core Reynolds number based on nozzle throat diameter is about 2000. Under these conditions, the density-viscosity product at the wall can be 10 times the freestream value, and the hydrogen is completely dissociated. Hence the boundary layer equations with variable transport properties are solved by a quasi-linearization technique, and the equilibrium properties of hydrogen are used in the calculations. Velocity profiles are obtained, and the wall shear and wall stagnation enthalpy gradient are plotted against the pressure gradient parameter.

Wu, P. K.↗

NASA NTRS1987

MPD thruster plasma instability studies

Results of an ongoing experimental/theoretical effort to understand the influence of plasma waves and instabilities upon coaxial plasma thruster performance and longevity are discussed. An experiment which permits measurement of the plasma dispersion relation plus a suite of plasma parameters for each one millisecond quasi-steady discharge is described. This investigation is guided by a linearized Boltzmann-Poisson kinetic numerical model which can accommodate the effects of collisions, inhomogeneities and the magnetic field. Preliminary tests reveal the presence of the low-frequency branch of an ion cyclotron wave. Although the stability of these waves still has to be established through spatial growth rate measurements, it is well known that they are prone to 'current-driven' instabilities. 'Current-driven' instabilities in the collisional tensor conducting plasma of the MPD thruster are analytically discussed and their relevance to the MPD plasma is illustrated by estimating a critical total current at which their onset occurs.

Kelly, A. J.↗

NASA NTRS1991

Derated ion thruster design issues

Preliminary activities to develop and refine a lightweight 30 cm engineering model ion thruster are discussed. The approach is to develop a 'derated' ion thruster capable of performing both auxiliary and primary propulsion roles over an input power range of at least 0.5 to 5.0 kilo-W. Design modifications to a baseline thruster to reduce mass and volume are discussed. Performance data over an order of magnitude input power range are presented, with emphasis on the performance impact of engine throttling. Thruster design modifications to optimize performance over specific power envelopes are discussed. Additionally, lifetime estimates based on wear test measurements are made for the operation envelope of the engine.

Patterson, Michael J.↗

NASA NTRS2003

The Propulsive Small Expendable Deployer System (ProSEDS)

This Final Report covers the following main topics: 1) Brief Description of ProSEDS; 2) Mission Analysis; 3) Dynamics Reference Mission; 4) Dynamics Stability; 5) Deployment Control; 6) Updated System Performance; 7) Updated Mission Analysis; 8) Updated Dynamics Reference Mission; 9) Updated Deployment Control Profiles and Simulations; 10) Updated Reference Mission; 11) Evaluation of Power Delivered by the Tether; 12) Deployment Control Profile Ref. #78 and Simulations; 13) Kalman Filters for Mission Estimation; 14) Analysis/Estimation of Deployment Flight Data; 15) Comparison of ED Tethers and Electrical Thrusters; 16) Dynamics Analysis for Mission Starting at a Lower Altitude; 17) Deployment Performance at a Lower Altitude; 18) Satellite Orbit after a Tether Cut; 19) Deployment with Shorter Dyneema Tether Length; 20) Interactive Software for ED Tethers.

Lorenzini, Enrico C.↗

NASA NTRS2016

Pattern Recognition Control Design

Spacecraft control algorithms must know the expected spacecraft response to any command to the available control effectors, such as reaction thrusters or torque devices. Spacecraft control system design approaches have traditionally relied on the estimated vehicle mass properties to determine the desired force and moment, as well as knowledge of the effector performance to efficiently control the spacecraft. A pattern recognition approach can be used to investigate the relationship between the control effector commands and the spacecraft responses. Instead of supplying the approximated vehicle properties and the effector performance characteristics, a database of information relating the effector commands and the desired vehicle response can be used for closed-loop control. A Monte Carlo simulation data set of the spacecraft dynamic response to effector commands can be analyzed to establish the influence a command has on the behavior of the spacecraft. A tool developed at NASA Johnson Space Center (Ref. 1) to analyze flight dynamics Monte Carlo data sets through pattern recognition methods can be used to perform this analysis. Once a comprehensive data set relating spacecraft responses with commands is established, it can be used in place of traditional control laws and gains set. This pattern recognition approach can be compared with traditional control algorithms to determine the potential benefits and uses.

Gambone, Elisabeth↗

NASA NTRS2020

Electric Propulsion for the Psyche Mission: Development Activities and Status

NASA’s Psyche mission will launch in 2022 and begin a 3.6-year cruise to the metallic asteroid Psyche, where it will examine this unique body. The baseline spacecraft design is a hybrid of JPL’s deep-space heritage subsystems with commercial partner Maxar’s electric propulsion, power, and structure subsystems. All primary propulsion will be done with SPT-140 thrusters, which will be the first use of Hall thrusters for a NASA mission. The electric propulsion subsystem and its implementation for the Psyche mission are described here. Major testing activities have included the successful completion of subsystem integrated testing with the design modifications required for Psyche, and a series of low-power thrust repeatability tests that were performed in support of navigation analyses. Thruster performance models have been further validated with new SPT-140 flight data, and new analyses of thruster swirl torque have been performed that result in much higher values than previously estimated. Analysis of recent Maxar flight data has also provided a new understanding of in-flight propellant usage uncertainties. Subsystem integration and test activities are now underway and the status and plans are discussed.

Johnson, Ian↗

NASA NTRS1983

Endurance test of a 30-CM-diameter engineering model ion thruster. Task 12: Investigation of thin-film erosion monitors for ion thrusters

An investigation of short term measurement techniques for predicting the wearout of ion thrusters resulting from sputter erosion damage is described. The previously established laminar thin film techniques to provide high precision erosion rate data. However, the erosion rates obtained using this technique are generally substantially higher than those obtained during long term endurance tests (by virtue of the as deposited nature of the thin films), so that the results must be interpreted in a relative sense. Absolute measurements can be performed using a new masked substrate arrangement which was developed during this study. This new technique provides a means for estimating the lifetimes of critical discharge chamber components based on direct measurements of sputter erosion depths obtained during short duration (10 hour) tests. The method enables the effects on lifetime of thruster design and operating parameters to be inferred without the investment of the time and capital required to conduct long term (1000 hour) endurance tests. Results obtained using the direct measurement technique are shown to agree with sputter erosion depths calculated for the plasma conditions of the test and also with lifetest results. The direct measurement approach is shown to be applicable to both mercury and argon discharge plasma environments and should be useful in estimating the lifetimes of inert gas and extended performance mercury ion thrusters presently under development.

Beattie, J. R.↗

NASA NTRS2018

In-situ Diagnostic for Assessing Hall Thruster Wear

The design of a new diagnostic to measure the net erosion of Hall thruster surfaces is presented. This diagnostic consists of a pair of optical noncontact profilometer pens mounted to a set of motion stages, which can interrogate the surface features of multiple components of interest including the hollow cathode assembly, magnet front pole covers, and discharge channel. By comparing scans of these surfaces to reference features, estimates of the component erosion rates can be acquired throughout long-duration lifetime tests without venting and removing the thruster from the vacuum facility for external profilometry. This work presents a detailed overview of the diagnostic design including the precision positioning system. In addition, preliminary data are shown which verify diagnostic operation and establish a baseline that will be used to track the erosion of the Hall Effect Rocket with Magnetic Shielding (HERMeS) Technology Demonstration Unit 3 (TDU-3) during an ongoing long-duration wear test.

Ahern, Drew M.↗