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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Design and testing of the U.S. Space Station Freedom primary propulsion system

The primary propulsion system (PPS) for the Space Station Freedom is discussed in terms of salient design characteristics and key testing procedures. The rocket engine modules contain reboost and attitude control thrusters, and their designs are illustrated showing the mounting structures, thruster solenoid valves, and thrust chambers. The propellant tank assembly for storing gaseous N pressurant and hydrazine propellant is described as are the system avionics, thruster solenoid valves, and latching isolation valves. PPS testing conducted on the development systems includes the use of a propulsion-module development unit, a development test article, and system qualification testing. Specific test articles include functional heaters, mass/thermal simulated components, flight-quality structures, and software control operations.

Morano, Joseph S.↗
NASA NTRS2001

Design of the EO-1 Pulsed Plasma Thruster Attitude Control Experiment

The Pulsed Plasma Thruster (PPT) Experiment on the Earth Observing 1 (EO-1) spacecraft has been designed to demonstrate the capability of a new generation PPT to perform spacecraft attitude control. The PPT is a small, self-contained pulsed electromagnetic Propulsion system capable of delivering high specific impulse (900-1200 s), very small impulse bits (10-1000 micro N-s) at low average power (less than 1 to 100 W). EO-1 has a single PPT that can produce torque in either the positive or negative pitch direction. For the PPT in-flight experiment, the pitch reaction wheel will be replaced by the PPT during nominal EO-1 nadir pointing. A PPT specific proportional-integral-derivative (PID) control algorithm was developed for the experiment. High fidelity simulations of the spacecraft attitude control capability using the PPT were conducted. The simulations, which showed PPT control performance within acceptable mission limits, will be used as the benchmark for on-orbit performance. The flight validation will demonstrate the ability of the PPT to provide precision pointing resolution. response and stability as an attitude control actuator.

Zakrzwski, Charles↗
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 NTRS2008

Performance Evaluation of the Prototype Model NEXT Ion Thruster

The performance testing results of the first prototype model NEXT ion engine, PM1, are presented. The NEXT program has developed the next generation ion propulsion system to enhance and enable Discovery, New Frontiers, and Flagship-type NASA missions. The PM1 thruster exhibits operational behavior consistent with its predecessors, the engineering model thrusters, with substantial mass savings, enhanced thermal margins, and design improvements for environmental testing compliance. The dry mass of PM1 is 12.7 kg. Modifications made in the thruster design have resulted in improved performance and operating margins, as anticipated. PM1 beginning-of-life performance satisfies all of the electric propulsion thruster mission-derived technical requirements. It demonstrates a wide range of throttleability by processing input power levels from 0.5 to 6.9 kW. At 6.9 kW, the PM1 thruster demonstrates specific impulse of 4190 s, 237 mN of thrust, and a thrust efficiency of 0.71. The flat beam profile, flatness parameters vary from 0.66 at low-power to 0.88 at full-power, and advanced ion optics reduce localized accelerator grid erosion and increases margins for electron backstreaming, impingement-limited voltage, and screen grid ion transparency. The thruster throughput capability is predicted to exceed 750 kg of xenon, an equivalent of 36,500 hr of continuous operation at the full-power operating condition.

Herman, Daniel A.↗
NASA NTRS1973

An engineering model 30 cm ion thruster

Thruster development at Hughes Research Laboratories and NASA Lewis Research Center has brought the 30-cm mercury bombardment ion thruster to the state of an engineering model. This thruster has been designed to have sufficient internal strength for direct mounting on gimbals, to weigh 7.3 kg, to operate with a corrected overall efficiency of 71%, and to have 10,000 hours lifetime. Subassemblies, such as the ion optical system, isolators, etc., have been upgraded to meet launch qualification standards. This paper presents a summary of the design specifications and performance characteristics which define the interface between the thruster module and the remainder of the propulsion system.

Poeschel, R. L.↗
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.

HERMeS↗
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 NTRS2010

Effect of a Second, Parallel Capacitor on the Performance of a Pulse Inductive Plasma Thruster

Pulsed inductive plasma accelerators are electrodeless space propulsion devices where a capacitor is charged to an initial voltage and is then discharged through an inductive coil that couples energy into the propellant, ionizing and accelerating it to produce thrust. A model that employs a set of circuit equations (as illustrated in Fig. 1a) coupled to a one-dimensional momentum equation has been previously used by Lovberg and Dailey [1] and Polzin et al. [2-4] to model the plasma acceleration process in pulsed inductive thrusters. In this paper an extra capacitor, inductor, and resistor are added to the system in the manner illustrated in the schematic shown in Fig. 1b. If the second capacitor has a smaller value than the initially charged capacitor, it can serve to increase the current rise rate through the inductive coil. Increasing the current rise rate should serve to better ionize the propellant. The equation of motion is solved to find the effect of an increased current rise rate on the acceleration process. We examine the tradeoffs between enhancing the breakdown process (increasing current rise rate) and altering the plasma acceleration process. These results provide insight into the performance of modified circuits in an inductive thruster, revealing how this design permutation can affect an inductive thruster's performance.

Polzin, Kurt A.↗
DOE OSTI2025

Study of the Design and Characteristics of a Modified Pulsed Plasma Thruster with Graphite and Tungsten Trigger Electrodes

The paper presents experimental results for a modified pulsed plasma thruster (PPT) with solid propellant, using a coaxial anode–cathode design. Graphite from pencil leads served as propellant, and a tungsten trigger electrode was tested to reduce carbonization effects. Experiments were performed in a vacuum chamber at 0.001 Pa, employing diagnostics such as discharge current/voltage recording, power measurement, ballistic pendulum, time-of-flight (TOF) method, and a Faraday cup. Current and voltage waveforms matched an oscillatory RLC circuit with variable plasma channel resistance. Key discharge parameters were measured, including current pulse duration/amplitude and plasma channel formation/decay dynamics. Impulse bit values, obtained with a ballistic pendulum, reached up to 8.5 μN·s. Increasing trigger capacitor capacitance reduced thrust due to unstable “pre-plasma” formation and partial pre-discharge energy loss. Using TOF and Faraday cup diagnostics, plasma front velocity, ion current amplitude, current density, and ion concentration were determined. Tungsten electrodes produced lower charged particle concentrations than graphite but offered better adhesion resistance, minimal carbonization, and stable long-term performance. The findings support optimizing trigger electrode materials and PPT operating modes to extend lifetime and stabilize thrust output.

Faraday cup↗
NASA NTRSDate not supplied

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

This work presents a summary of the detailed characterization testof the 12.5 kW Advanced Electric Propulsion System Engineering Test Unit 2 (AEPS ETU-2) thruster produced by Aerojet Rocketdyne. This test campaign had two major goals: to assessthe risk of design compliance with thruster requirements and providea comparison to the previously-tested NASA Hall Effect Rocket with Magnet Shielding Technology Demonstration Units (HERMeS TDUs) from which the AEPS ETU design was derived.

AEPS↗
NASA NTRS2000

Electromagnetic Propulsion

The design and development of an Electromagnetic Propulsion is discussed. Specific Electromagnetic Propulsion Topics discussed include: (1) Technology for Pulse Inductive Thruster (PIT), to design, develop, and test of a multirepetition rate pulsed inductive thruster, Solid-State Switch Technology, and Pulse Driver Network and Architecture; (2) Flight Weight Magnet Survey, to determine/develop light weight high performance magnetic materials for potential application Advanced Space Flight Systems as these systems develop; and (3) Magnetic Flux Compression, to enable rapid/robust/reliable omni-planetary space transportation within realistic development and operational costs constraints.

Schafer, Charles↗
NASA NTRS1982

Developing a scalable inert gas ion thruster

Analytical studies to identify and then design a high performance scalable ion thruster operating with either argon or xenon for use in large space systems are presented. The magnetoelectrostatic containment concept is selected for its efficient ion generation capabilities. The iterative nature of the bounding magnetic fields allows the designer to scale both the diameter and length, so that the thruster can be adapted to spacecraft growth over time. Three different thruster assemblies (conical, hexagonal and hemispherical) are evaluated for a 12 cm diameter thruster and performance mapping of the various thruster configurations shows that conical discharge chambers produce the most efficient discharge operation, achieving argon efficiencies of 50-80% mass utilization at 240-310 eV/ion and xenon efficiencies of 60-97% at 240-280 eV/ion. Preliminary testing of the large 30 cm thruster, using argon propellant, indicates a 35% improvement over the 12 cm thruster in mass utilization efficiency. Since initial performance is found to be better than projected, a larger 50 cm thruster is already in the development stage.

James, E.↗
NASA NTRS1987

A theoretical evaluation of laser-sustained plasma thruster performance

An extensive numerical experiment has been conducted to evaluate rocket thruster performance using a laser-sustained hydrogen plasma as the propellant. The plasma was sustained using a 30 kW CO2 laser beam operated at 10.6 microns focused inside the thruster. The steady-state Navier-Stokes equations coupled with the laser power absorption process have been solved numerically. A pressure based Navier-Stokes solver using body-fitted coordinate was used to calculate the laser-supported rocket flow which included both recirculating and transonic flow regions. The local thermodynamic equilibrium (LTE) assumption was used for the plasma thermophysical and optical properties. Geometric ray tracing was adopted to describe the laser beam. Several different throat size thrusters operated at 150 and 300 kPa chamber stagnation pressure were studied. It was found that the thruster performance (vacuum specific impulse) was highly dependent on the operating conditions, and a properly designed laser supported thruster can attain a specific impulse around 1500 secs. The heat loading on the thruster wall was also estimated and was in the range of that for a conventional chemical rocket.

Jeng, San-Mou↗
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 NTRS1973

Testing and evaluation of the LES-6 pulsed plasma thruster by means of a torsion pendulum system

Performance characteristics of the LES-6 pulsed plasma thruster over a range of input conditions were investigated by means of a torsion pendulum system. Parameters of particular interest included the impulse bit and time average thrust (and their repeatability), specific impulse, mass ablated per discharge, specific thrust, energy per unit area, efficiency, and variation of performance with ignition command rate. Intermittency of the thruster as affected by input energy and igniter resistance were also investigated. Comparative experimental data correlation with the data presented. The results of these tests indicate that the LES-6 thruster, with some identifiable design improvements, represents an attractive reaction control thruster for attitude contol applications on long-life spacecraft requiring small metered impulse bits for precise pointing control of science instruments.

Hamidian, J. P.↗
NASA NTRS1980

Inert gas ion thruster development

Two 12 cm magneto-electrostatic containment (MESC) ion thrusters were performance mapped with argon and xenon. The first, hexagonal, thruster produced optimized performance of 48.5to 79 percent argon mass utilization efficiencies at discharge energies of 240 to 425 eV/ion, respectively, Xenon mass utilization efficiencies of 78 to 95 percent were observed at discharge energies of 220 to 290 eV/ion with the same optimized hexagonal thruster. Changes to the cathode baffle reduced the discharge anode potential during xenon operation from approximately 40 volts to about 30 volts. Preliminary tests conducted with the second, hemispherical, MESC thruster showed a nonuniform anode magnetic field adversely affected thruster performance. This performance degradation was partially overcome by changes in the boundary anode placement. Conclusions drawn the hemispherical thruster tests gave insights into the plasma processes in the MESC discharge that will aid in the design of future thrusters.

Ramsey, W. D.↗
NASA NTRS2014

Implementation and Initial Validation of a 100-Kilowatt Class Nested-Channel Hall Thruster

The X3 is a 100-kilowatt class nested-channel Hall thruster developed by the Plasmadynamics and Electric Propulsion Laboratory at the University of Michigan in collaboration with the Air Force Research Laboratory and NASA. The cathode, magnetic circuit, boron nitride channel rings, and anodes all required specific design considerations during thruster development, and thermal modeling was used to properly account for thermal growth in material selection and component design. A number of facility upgrades were required at the University of Michigan to facilitate operation of the X3. These upgrades included a re-worked propellant feed system, a completely redesigned power and telemetry break-out box, and numerous updates to thruster handling equipment. The X3 was tested on xenon propellant at two current densities, 37% and 73% of the nominal design value. It was operated to a maximum steady-state discharge power of 60.8 kilowatts. The tests presented here served as an initial validation of thruster operation. Thruster behavior was monitored with telemetry, photography and high-speed current probes. The photography showed a uniform plume throughout testing. At constant current density, reductions in mass flow rate of 18% and 26% were observed in the three-channel operating configuration as compared to the superposition of each channel running individually. The high-speed current probes showed that the thruster was stable at all operating points and that the channels influence each other when more than one is operating simultaneously. Additionally, the ratio of peak-to-peak AC-coupled discharge current oscillations to mean discharge current did not exceed 51% for any operating points reported here, and did not exceed 17% at the higher current density.

High Power↗