High Accuracy Positioning using Jet Thrusters for Quadcopter

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Pi ChenHuan

2018-01-01

Full Text Available A quadcopter is equipped with four additional jet thrusters on its horizontal plane and vertical to each other in order to improve the maneuverability and positioning accuracy of quadcopter. A dynamic model of the quadcopter with jet thrusters is derived and two controllers are implemented in simulation, one is a dual loop state feedback controller for pose control and another is an auxiliary jet thruster controller for accurate positioning. Step response simulations showed that the jet thruster can control the quadcopter with less overshoot compared to the conventional one. Over 10s loiter simulation with disturbance, the quadcopter with jet thruster decrease 85% of RMS error of horizontal disturbance compared to a conventional quadcopter with only a dual loop state feedback controller. The jet thruster controller shows the possibility for further accurate in the field of quadcopter positioning.

Predictive fault-tolerant control of an all-thruster satellite in 6-DOF motion via neural network model updating

Science.gov (United States)

Tavakoli, M. M.; Assadian, N.

2018-03-01

The problem of controlling an all-thruster spacecraft in the coupled translational-rotational motion in presence of actuators fault and/or failure is investigated in this paper. The nonlinear model predictive control approach is used because of its ability to predict the future behavior of the system. The fault/failure of the thrusters changes the mapping between the commanded forces to the thrusters and actual force/torque generated by the thruster system. Thus, the basic six degree-of-freedom kinetic equations are separated from this mapping and a set of neural networks are trained off-line to learn the kinetic equations. Then, two neural networks are attached to these trained networks in order to learn the thruster commands to force/torque mappings on-line. Different off-nominal conditions are modeled so that neural networks can detect any failure and fault, including scale factor and misalignment of thrusters. A simple model of the spacecraft relative motion is used in MPC to decrease the computational burden. However, a precise model by the means of orbit propagation including different types of perturbation is utilized to evaluate the usefulness of the proposed approach in actual conditions. The numerical simulation shows that this method can successfully control the all-thruster spacecraft with ON-OFF thrusters in different combinations of thruster fault and/or failure.

Multi-Axis Thrust Measurements of the EO-1 Pulsed Plasma Thruster

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Arrington, Lynn A.; Haag, Thomas W.

1999-01-01

Pulsed plasma thrusters are low thrust propulsive devices which have a high specific impulse at low power. A pulsed plasma thruster is currently scheduled to fly as an experiment on NASA's Earth Observing-1 satellite mission. The pulsed plasma thruster will be used to replace one of the reaction wheels. As part of the qualification testing of the thruster it is necessary to determine the nominal thrust as a function of charge energy. These data will be used to determine control algorithms. Testing was first completed on a breadboard pulsed plasma thruster to determine nominal or primary axis thrust and associated propellant mass consumption as a function of energy and then later to determine if any significant off-axis thrust component existed. On conclusion that there was a significant off-axis thrust component with the bread-board in the direction of the anode electrode, the test matrix was expanded on the flight hardware to include thrust measurements along all three orthogonal axes. Similar off-axis components were found with the flight unit.

Design and model experiments on thruster assisted mooring system; Futaishiki kaiyo kozobutsu no thruster ni yoru choshuki doyo seigyo

Energy Technology Data Exchange (ETDEWEB)

Nakamura, M; Koterayama, W [Kyushu Univ., Fukuoka (Japan). Research Inst. for Applied Mechanics; Kajiwara, H [Kyushu Institute of Technology, Kitakyushu (Japan). Faculty of Computer Science and System Engineering; Hyakudome, T [Kyushu University, Fukuoka (Japan)

1997-12-31

Described herein are dynamics and model experiments of the system in which positioning of a floating marine structure by mooring is combined with thruster-controlled positioning. Coefficients of dynamic forces acting on a floating structure model are determined experimentally and by the three-dimensional singularity distribution method, and the controller is designed by the PID, LQI and H{infinity} control theories. A model having a scale ratio of 1/100 was used for the experiments, where 2 thrusters were arranged in a diagonal line, one on the X-axis. It is found that the LQI and H{infinity} controllers of the thruster can control long-cycle rolling of the floating structure. They allow thruster control which is insensitive to wave cycle motion, and efficiently reduce positioning energy. The H{infinity} control regulates frequency characteristics of a closed loop more finely than the LQI control, and exhibits better controllability. 25 refs., 25 figs.

Design and model experiments on thruster assisted mooring system; Futaishiki kaiyo kozobutsu no thruster ni yoru choshuki doyo seigyo

Energy Technology Data Exchange (ETDEWEB)

Nakamura, M.; Koterayama, W. [Kyushu Univ., Fukuoka (Japan). Research Inst. for Applied Mechanics; Kajiwara, H. [Kyushu Institute of Technology, Kitakyushu (Japan). Faculty of Computer Science and System Engineering; Hyakudome, T. [Kyushu University, Fukuoka (Japan)

1996-12-31

Described herein are dynamics and model experiments of the system in which positioning of a floating marine structure by mooring is combined with thruster-controlled positioning. Coefficients of dynamic forces acting on a floating structure model are determined experimentally and by the three-dimensional singularity distribution method, and the controller is designed by the PID, LQI and H{infinity} control theories. A model having a scale ratio of 1/100 was used for the experiments, where 2 thrusters were arranged in a diagonal line, one on the X-axis. It is found that the LQI and H{infinity} controllers of the thruster can control long-cycle rolling of the floating structure. They allow thruster control which is insensitive to wave cycle motion, and efficiently reduce positioning energy. The H{infinity} control regulates frequency characteristics of a closed loop more finely than the LQI control, and exhibits better controllability. 25 refs., 25 figs.

Magnesium Hall Thruster

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Szabo, James J.

2015-01-01

This Phase II project is developing a magnesium (Mg) Hall effect thruster system that would open the door for in situ resource utilization (ISRU)-based solar system exploration. Magnesium is light and easy to ionize. For a Mars- Earth transfer, the propellant mass savings with respect to a xenon Hall effect thruster (HET) system are enormous. Magnesium also can be combusted in a rocket with carbon dioxide (CO2) or water (H2O), enabling a multimode propulsion system with propellant sharing and ISRU. In the near term, CO2 and H2O would be collected in situ on Mars or the moon. In the far term, Mg itself would be collected from Martian and lunar regolith. In Phase I, an integrated, medium-power (1- to 3-kW) Mg HET system was developed and tested. Controlled, steady operation at constant voltage and power was demonstrated. Preliminary measurements indicate a specific impulse (Isp) greater than 4,000 s was achieved at a discharge potential of 400 V. The feasibility of delivering fluidized Mg powder to a medium- or high-power thruster also was demonstrated. Phase II of the project evaluated the performance of an integrated, highpower Mg Hall thruster system in a relevant space environment. Researchers improved the medium power thruster system and characterized it in detail. Researchers also designed and built a high-power (8- to 20-kW) Mg HET. A fluidized powder feed system supporting the high-power thruster was built and delivered to Busek Company, Inc.

Cathode Effects in Cylindrical Hall Thrusters

Energy Technology Data Exchange (ETDEWEB)

Granstedt, E.M.; Raitses, Y.; Fisch, N. J.

2008-09-12

Stable operation of a cylindrical Hall thruster (CHT) has been achieved using a hot wire cathode, which functions as a controllable electron emission source. It is shown that as the electron emission from the cathode increases with wire heating, the discharge current increases, the plasma plume angle reduces, and the ion energy distribution function shifts toward higher energies. The observed effect of cathode electron emission on thruster parameters extends and clarifies performance improvements previously obtained for the overrun discharge current regime of the same type of thruster, but using a hollow cathode-neutralizer. Once thruster discharge current saturates with wire heating, further filament heating does not affect other discharge parameters. The saturated values of thruster discharge parameters can be further enhanced by optimal placement of the cathode wire with respect to the magnetic field.

Oxygen-Methane Thruster

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Pickens, Tim

2012-01-01

An oxygen-methane thruster was conceived with integrated igniter/injector capable of nominal operation on either gaseous or liquid propellants. The thruster was designed to develop 100 lbf (approximately 445 N) thrust at vacuum conditions and use oxygen and methane as propellants. This continued development included refining the design of the thruster to minimize part count and manufacturing difficulties/cost, refining the modeling tools and capabilities that support system design and analysis, demonstrating the performance of the igniter and full thruster assembly with both gaseous and liquid propellants, and acquiring data from this testing in order to verify the design and operational parameters of the thruster. Thruster testing was conducted with gaseous propellants used for the igniter and thruster. The thruster was demonstrated to work with all types of propellant conditions, and provided the desired performance. Both the thruster and igniter were tested, as well as gaseous propellants, and found to provide the desired performance using the various propellant conditions. The engine also served as an injector testbed for MSFC-designed refractory combustion chambers made of rhenium.

Control of the electric-field profile in the Hall thruster

International Nuclear Information System (INIS)

Fruchtman, A.; Fisch, N.J.; Raitses, Y.

2001-01-01

Control of the electric-field profile in the Hall thruster through the positioning of an additional electrode along the channel is shown theoretically to enhance the efficiency. The reduction of the potential drop near the anode by use of the additional electrode increases the plasma density there, through the increase of the electron and ion transit times, causing the ionization in the vicinity of the anode to increase. The resulting separation of the ionization and acceleration regions increases the propellant and energy utilizations. An abrupt sonic transition is forced to occur at the axial location of the additional electrode, accompanied by the generation of a large (theoretically infinite) electric field. This ability to generate a large electric field at a specific location along the channel, in addition to the ability to specify the electric potential there, allows us further control of the electric-field profile in the thruster. In particular, when the electron temperature is high, a large abrupt voltage drop is induced at the vicinity of the additional electrode, a voltage drop that can comprise a significant part of the applied voltage

Thrusting maneuver control of a small spacecraft via only gimbaled-thruster scheme

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Kabganian, Mansour; Kouhi, Hamed; Shahravi, Morteza; Fani Saberi, Farhad

2018-05-01

The thrust vector control (TVC) scheme is a powerful method in spacecraft attitude control. Since the control of a small spacecraft is being studied here, a solid rocket motor (SRM) should be used instead of a liquid propellant motor. Among the TVC methods, gimbaled-TVC as an efficient method is employed in this paper. The spacecraft structure is composed of a body and a gimbaled-SRM where common attitude control systems such as reaction control system (RCS) and spin-stabilization are not presented. A nonlinear two-body model is considered for the characterization of the gimbaled-thruster spacecraft where, the only control input is provided by a gimbal actuator. The attitude of the spacecraft is affected by a large exogenous disturbance torque which is generated by a thrust vector misalignment from the center of mass (C.M). A linear control law is designed to stabilize the spacecraft attitude while rejecting the mentioned disturbance torque. A semi-analytical formulation of the region of attraction (RoA) is developed to ensure the local stability and fast convergence of the nonlinear closed-loop system. Simulation results of the 3D maneuvers are included to show the applicability of this method for use in a small spacecraft.

The effects of 1 kW class arcjet thruster plumes on spacecraft charging and spacecraft thermal control materials

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Bogorad, A.; Lichtin, D. A.; Bowman, C.; Armenti, J.; Pencil, E.; Sarmiento, C.

1992-01-01

Arcjet thrusters are soon to be used for north/south stationkeeping on commercial communications satellites. A series of tests was performed to evaluate the possible effects of these thrusters on spacecraft charging and the degradation of thermal control material. During the tests the interaction between arcjet plumes and both charged and uncharged surfaces did not cause any significant material degradation. In addition, firing an arcjet thruster benignly reduced the potential of charged surfaces to near zero.

Power processing systems for ion thrusters.

Science.gov (United States)

Herron, B. G.; Garth, D. R.; Finke, R. C.; Shumaker, H. A.

1972-01-01

The proposed use of ion thrusters to fulfill various communication satellite propulsion functions such as east-west and north-south stationkeeping, attitude control, station relocation and orbit raising, naturally leads to the requirement for lightweight, efficient and reliable thruster power processing systems. Collectively, the propulsion requirements dictate a wide range of thruster power levels and operational lifetimes, which must be matched by the power processing. This paper will discuss the status of such power processing systems, present system design alternatives and project expected near future power system performance.

Aerodynamic Interactions of Propulsive Deceleration and Reaction Control System Jets on Mars-Entry Aeroshells

Science.gov (United States)

Alkandry, Hicham

Future missions to Mars, including sample-return and human-exploration missions, may require alternative entry, descent, and landing technologies in order to perform pinpoint landing of heavy vehicles. Two such alternatives are propulsive deceleration (PD) and reaction control systems (RCS). PD can slow the vehicle during Mars atmospheric descent by directing thrusters into the incoming freestream. RCS can provide vehicle control and steering by inducing moments using thrusters on the hack of the entry capsule. The use of these PD and RCS jets, however, involves complex flow interactions that are still not well understood. The fluid interactions induced by PD and RCS jets for Mars-entry vehicles in hypersonic freestream conditions are investigated using computational fluid dynamics (CFD). The effects of central and peripheral PD configurations using both sonic and supersonic jets at various thrust conditions are examined in this dissertation. The RCS jet is directed either parallel or transverse to the freestream flow at different thrust conditions in order to examine the effects of the thruster orientation with respect to the center of gravity of the aeroshell. The physical accuracy of the computational method is also assessed by comparing the numerical results with available experimental data. The central PD configuration decreases the drag force acting on the entry capsule due to a shielding effect that prevents mass and momentum in the hypersonic freestream from reaching the aeroshell. The peripheral PD configuration also decreases the drag force by obstructing the flow around the aeroshell and creating low surface pressure regions downstream of the PD nozzles. The Mach number of the PD jets, however, does not have a significant effect on the induced fluid interactions. The reaction control system also alters the flowfield, surface, and aerodynamic properties of the aeroshell, while the jet orientation can have a significant effect on the control effectiveness

Development of a green bipropellant hydrogen peroxide thruster for attitude control on satellites

Science.gov (United States)

Woschnak, A.; Krejci, D.; Schiebl, M.; Scharlemann, C.

2013-03-01

This document describes the selection assessment of propellants for a 1-newton green bipropellant thruster for attitude control on satellites. The development of this thruster was conducted as a part of the project GRASP (Green Advanced Space Propellants) within the European FP7 research program. The green propellant combinations hydrogen peroxide (highly concentrated with 87.5 %(wt.)) with kerosene or hydrogen peroxide (87.5 %(wt.)) with ethanol were identified as interesting candidates and were investigated in detail with the help of an experimental combustion chamber in the chemical propulsion laboratory at the Forschungsund Technologietransfer GmbH ― Fotec. Based on the test results, a final selection of propellants was performed.

Study on Endurance and Performance of Impregnated Ruthenium Catalyst for Thruster System.

Science.gov (United States)

Kim, Jincheol; Kim, Taegyu

2018-02-01

Performance and endurance of the Ru catalyst were studied for nitrous oxide monopropellant thruster system. The thermal decomposition of N2O requires a considerably high temperature, which make it difficult to be utilized as a thruster propellant, while the propellant decomposition temperature can be reduced by using the catalyst through the decomposition reaction with the propellant. However, the catalyst used for the thruster was frequently exposed to high temperature and high-pressure environment. Therefore, the state change of the catalyst according to the thruster operation was analyzed. Characterization of catalyst used in the operation condition of the thruster was performed using FE-SEM and EDS. As a result, performance degradation was occurred due to the volatilization of Ru catalyst and reduction of the specific surface area according to the phase change of Al2O3.

Development of HAN-based Liquid Propellant Thruster

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Hisatsune, K.; Izumi, J.; Tsutaya, H.; Furukawa, K.

2004-10-01

Many of propellants that are applied to the conventional spacecraft propulsion system are toxic propellants. Because of its toxicity, considering the environmental pollution or safety on handling, it will be necessary to apply the "green" propellant to the spacecraft propulsion system. The purpose of this study is to apply HAN based liquid propellant (LP1846) to mono propellant thruster. Compared to the hydrazine that is used in conventional mono propellant thruster, HAN based propellant is not only lower toxic but also can obtain higher specific impulse. Moreover, HAN based propellant can be decomposed by the catalyst. It means there are the possibility of applying to the mono propellant thruster that can leads to the high reliability of the propulsion system.[1],[2] However, there are two technical subjects, to apply HAN based propellant to the mono propellant thruster. One is the high combustion temperature. The catalyst will be damaged under high temperature condition. The other is the low catalytic activity. It is the serious problem on application of HAN based propellant to the mono propellant thruster that is used for attitude control of spacecraft. To improve the catalytic activity of HAN based propellant, it is necessary to screen the best catalyst for HAN based propellant. The adsorption analysis is conducted by Monte Carlo Simulation to screen the catalyst metal for HAN and TEAN. The result of analysis shows the Iridium is the best catalyst metal for HAN and TEAN. Iridium is the catalyst metal that is used at conventional mono propellant thruster catalyst Shell405. Then, to confirm the result of analysis, the reaction test about catalyst is conducted. The result of this test is the same as the result of adsorption analysis. That means the adsorption analysis is effective in screening the catalyst metal. At the evaluating test, the various types of carrier of catalyst are also compared to Shell 405 to improve catalytic activity. The test result shows the

Thermal-environmental testing of a 30-cm engineering model thruster

Science.gov (United States)

Mirtich, M. J.

1976-01-01

An experimental test program was carried out to document all 30-cm electron bombardment Hg ion bombardment thruster functions and characteristics over the thermal environment of several proposed missions. An engineering model thruster was placed in a thermal test facility equipped with -196 C walls and solar simulation. The thruster was cold soaked and exposed to simulated eclipses lasting in duration from 17 to 72 minutes. The thruster was operated at quarter, to full beam power in various thermal configurations which simulated multiple thruster operation, and was also exposed to 1 and 2 suns solar simulation. Thruster control characteristics and constraints; performance, including thrust magnitude and direction; and structural integrity were evaluated over the range of thermal environments tested.

Thrust performance, propellant ionization, and thruster erosion of an external discharge plasma thruster

Science.gov (United States)

Karadag, Burak; Cho, Shinatora; Funaki, Ikkoh

2018-04-01

It is quite a challenge to design low power Hall thrusters with a long lifetime and high efficiency because of the large surface area to volume ratio and physical limits to the magnetic circuit miniaturization. As a potential solution to this problem, we experimentally investigated the external discharge plasma thruster (XPT). The XPT produces and sustains a plasma discharge completely in the open space outside of the thruster structure through a magnetic mirror configuration. It eliminates the very fundamental component of Hall thrusters, discharge channel side walls, and its magnetic circuit consists solely of a pair of hollow cylindrical permanent magnets. Thrust, low frequency discharge current oscillation, ion beam current, and plasma property measurements were conducted to characterize the manufactured prototype thruster for the proof of concept. The thrust performance, propellant ionization, and thruster erosion were discussed. Thrust generated by the XPT was on par with conventional Hall thrusters [stationary plasma thruster (SPT) or thruster with anode layer] at the same power level (˜11 mN at 250 W with 25% anode efficiency without any optimization), and discharge current had SPT-level stability (Δ design and provide a successful proof of concept experiment of the XPT.

Continuous Wheel Momentum Dumping Using Magnetic Torquers and Thrusters

Science.gov (United States)

Oh, Hwa-Suk; Choi, Wan-Sik; Eun, Jong-Won

1996-12-01

Two momentum management schemes using magnetic torquers and thrusters are sug-gested. The stability of the momentum dumping logic is proved at a general attitude equilibrium. Both momentum dumping control laws are implemented with Pulse-Width- Pulse-Frequency Modulated on-off control, and shown working equally well with the original continuous and variable strength control law. Thrusters are assummed to be asymmetrically configured as a contingency case. Each thruster is fired following separated control laws rather than paired thrusting. Null torque thrusting control is added on the thrust control calculated from the momentum control law for the gener-ation of positive thrusting force. Both magnetic and thrusting control laws guarantee the momentum dumping, however, the wheel inner loop control is needed for the "wheel speed" dumping, The control laws are simulated on the KOrea Multi-Purpose SATellite (KOMPSAT) model.

Iodine Hall Thruster Propellant Feed System for a CubeSat

Science.gov (United States)

Polzin, Kurt A.

2014-01-01

There has been significant work recently in the development of iodine-fed Hall thrusters for in-space propulsion applications.1 The use of iodine as a propellant provides many advantages over present xenon-gas-fed Hall thruster systems. Iodine is a solid at ambient temperature (no pressurization required) and has no special handling requirements, making it safe for secondary flight opportunities. It has exceptionally high ?I sp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing system level advantages over mid-term high power electric propulsion options. Iodine provides thrust and efficiency that are comparable to xenonfed Hall thrusters while operating in the same discharge current and voltage regime, making it possible to leverage the development of flight-qualified xenon Hall thruster power processing units for the iodine application. Work at MSFC is presently aimed at designing, integrating, and demonstrating a flight-like iodine feed system suitable for the Hall thruster application. This effort represents a significant advancement in state-of-the-art. Though Iodine thrusters have demonstrated high performance with mission enabling potential, a flight-like feed system has never been demonstrated and iodine compatible components do not yet exist. Presented in this paper is the end-to-end integrated feed system demonstration. The system includes a propellant tank with active feedback-control heating, fill and drain interfaces, latching and proportional flow control valves (PFCV), flow resistors, and flight-like CubeSat power and control electronics. Hardware is integrated into a CubeSat-sized structure, calibrated and tested under vacuum conditions, and operated under under hot-fire conditions using a Busek BHT-200 thruster designed for iodine. Performance of the system is evaluated thorugh accurate measurement of thrust and a calibrated of mass flow rate measurement, which is a function of

Experimental investigation of the catalytic decomposition and combustion characteristics of a non-toxic ammonium dinitramide (ADN)-based monopropellant thruster

Science.gov (United States)

Chen, Jun; Li, Guoxiu; Zhang, Tao; Wang, Meng; Yu, Yusong

2016-12-01

Low toxicity ammonium dinitramide (ADN)-based aerospace propulsion systems currently show promise with regard to applications such as controlling satellite attitude. In the present work, the decomposition and combustion processes of an ADN-based monopropellant thruster were systematically studied, using a thermally stable catalyst to promote the decomposition reaction. The performance of the ADN propulsion system was investigated using a ground test system under vacuum, and the physical properties of the ADN-based propellant were also examined. Using this system, the effects of the preheating temperature and feed pressure on the combustion characteristics and thruster performance during steady state operation were observed. The results indicate that the propellant and catalyst employed during this work, as well as the design and manufacture of the thruster, met performance requirements. Moreover, the 1 N ADN thruster generated a specific impulse of 223 s, demonstrating the efficacy of the new catalyst. The thruster operational parameters (specifically, the preheating temperature and feed pressure) were found to have a significant effect on the decomposition and combustion processes within the thruster, and the performance of the thruster was demonstrated to improve at higher feed pressures and elevated preheating temperatures. A lower temperature of 140 °C was determined to activate the catalytic decomposition and combustion processes more effectively compared with the results obtained using other conditions. The data obtained in this study should be beneficial to future systematic and in-depth investigations of the combustion mechanism and characteristics within an ADN thruster.

Ion thruster design and analysis

Science.gov (United States)

Kami, S.; Schnelker, D. E.

1976-01-01

Questions concerning the mechanical design of a thruster are considered, taking into account differences in the design of an 8-cm and a 30-cm model. The components of a thruster include the thruster shell assembly, the ion extraction electrode assembly, the cathode isolator vaporizer assembly, the neutralizer isolator vaporizer assembly, ground screen and mask, and the main isolator vaporizer assembly. Attention is given to the materials used in thruster fabrication, the advanced manufacturing methods used, details of thruster performance, an evaluation of thruster life, structural and thermal design considerations, and questions of reliability and quality assurance.

Concept Study of Radio Frequency (RF Plasma Thruster for Space Propulsion

Directory of Open Access Journals (Sweden)

Anna-Maria Theodora ANDREESCU

2016-12-01

Full Text Available Electric thrusters are capable of accelerating ions to speeds that are impossible to reach using chemical reaction. Recent advances in plasma-based concepts have led to the identification of electromagnetic (RF generation and acceleration systems as able to provide not only continuous thrust, but also highly controllable and wide-range exhaust velocities. For Future Space Propulsion there is a pressing need for low pressure, high mass flow rate and controlled ion energies. This paper explores the potential of using RF heated plasmas for space propulsion in order to mitigate the electric propulsion problems caused by erosion and gain flexibility in plasma manipulation. The main key components of RF thruster architecture are: a feeding system able to provide the required neutral gas flow, plasma source chamber, antenna/electrodes wrapped around the discharge tube and optimized electromagnetic field coils for plasma confinement. A preliminary analysis of system performance (thrust, specific impulse, efficiency is performed along with future plans of Space Propulsion based on this new concept of plasma mechanism.

Thermo-mechanical design aspects of mercury bombardment ion thrusters.

Science.gov (United States)

Schnelker, D. E.; Kami, S.

1972-01-01

The mechanical design criteria are presented as background considerations for solving problems associated with the thermomechanical design of mercury ion bombardment thrusters. Various analytical procedures are used to aid in the development of thruster subassemblies and components in the fields of heat transfer, vibration, and stress analysis. Examples of these techniques which provide computer solutions to predict and control stress levels encountered during launch and operation of thruster systems are discussed. Computer models of specific examples are presented.

Iodine Hall Thruster

Science.gov (United States)

Szabo, James

2015-01-01

Iodine enables dramatic mass and cost savings for lunar and Mars cargo missions, including Earth escape and near-Earth space maneuvers. The demonstrated throttling ability of iodine is important for a singular thruster that might be called upon to propel a spacecraft from Earth to Mars or Venus. The ability to throttle efficiently is even more important for missions beyond Mars. In the Phase I project, Busek Company, Inc., tested an existing Hall thruster, the BHT-8000, on iodine propellant. The thruster was fed by a high-flow iodine feed system and supported by an existing Busek hollow cathode flowing xenon gas. The Phase I propellant feed system was evolved from a previously demonstrated laboratory feed system. Throttling of the thruster between 2 and 11 kW at 200 to 600 V was demonstrated. Testing showed that the efficiency of iodine fueled BHT-8000 is the same as with xenon, with iodine delivering a slightly higher thrust-to-power (T/P) ratio. In Phase II, a complete iodine-fueled system was developed, including the thruster, hollow cathode, and iodine propellant feed system. The nominal power of the Phase II system is 8 kW; however, it can be deeply throttled as well as clustered to much higher power levels. The technology also can be scaled to greater than 100 kW per thruster to support megawatt-class missions. The target thruster efficiency for the full-scale system is 65 percent at high specific impulse (Isp) (approximately 3,000 s) and 60 percent at high thrust (Isp approximately 2,000 s).

A mechanical, thermal and electrical packaging design for a prototype power management and control system for the 30 cm mercury ion thruster

Science.gov (United States)

Sharp, G. R.; Gedeon, L.; Oglebay, J. C.; Shaker, F. S.; Siegert, C. E.

1978-01-01

A prototype electric power management and thruster control system for a 30 cm ion thruster is described. The system meets all of the requirements necessary to operate a thruster in a fully automatic mode. Power input to the system can vary over a full two to one dynamic range (200 to 400 V) for the solar array or other power source. The power management and control system is designed to protect the thruster, the flight system and itself from arcs and is fully compatible with standard spacecraft electronics. The system is easily integrated into flight systems which can operate over a thermal environment ranging from 0.3 to 5 AU. The complete power management and control system measures 45.7 cm (18 in.) x 15.2 cm (6 in.) x 114.8 cm (45.2 in.) and weighs 36.2 kg (79.7 lb). At full power the overall efficiency of the system is estimated to be 87.4 percent. Three systems are currently being built and a full schedule of environmental and electrical testing is planned.

A mechanical, thermal and electrical packaging design for a prototype power management and control system for the 30 cm mercury ion thruster

Science.gov (United States)

Sharp, G. R.; Gedeon, L.; Oglebay, J. C.; Shaker, F. S.; Siegert, C. E.

1978-01-01

A prototype Electric Power Management and Thruster Control System for a 30 cm ion thruster has been built and is ready to support a first mission application. The system meets all of the requirements necessary to operate a thruster in a fully automatic mode. Power input to the system can vary over a full two to one dynamic range (200 to 400 V) for the solar array or other power source. The Power Management and Control system is designed to protect the thruster, the flight system and itself from arcs and is fully compatible with standard spacecraft electronics. The system is designed to be easily integrated into flight systems which can operate over a thermal environment ranging from 0.3 to 5 AU. The complete Power Management and Control system measures 45.7 cm x 15.2 cm x 114.8 cm and weighs 36.2 kg. At full power the overall efficiency of the system is estimated to be 87.4 percent. Three systems are currently being built and a full schedule of environmental and electrical testing is planned.

Testing of a Liquid Oxygen/Liquid Methane Reaction Control Thruster in a New Altitude Rocket Engine Test Facility

Science.gov (United States)

Meyer, Michael L.; Arrington, Lynn A.; Kleinhenz, Julie E.; Marshall, William M.

2012-01-01

A relocated rocket engine test facility, the Altitude Combustion Stand (ACS), was activated in 2009 at the NASA Glenn Research Center. This facility has the capability to test with a variety of propellants and up to a thrust level of 2000 lbf (8.9 kN) with precise measurement of propellant conditions, propellant flow rates, thrust and altitude conditions. These measurements enable accurate determination of a thruster and/or nozzle s altitude performance for both technology development and flight qualification purposes. In addition the facility was designed to enable efficient test operations to control costs for technology and advanced development projects. A liquid oxygen-liquid methane technology development test program was conducted in the ACS from the fall of 2009 to the fall of 2010. Three test phases were conducted investigating different operational modes and in addition, the project required the complexity of controlling propellant inlet temperatures over an extremely wide range. Despite the challenges of a unique propellant (liquid methane) and wide operating conditions, the facility performed well and delivered up to 24 hot fire tests in a single test day. The resulting data validated the feasibility of utilizing this propellant combination for future deep space applications.

Prediction of plasma properties in mercury ion thrusters

Science.gov (United States)

Longhurst, G. R.

1978-01-01

A simplified theoretical model was developed which obtains to first order the plasma properties in the discharge chamber of a mercury ion thruster from basic thruster design and controllable operating parameters. The basic operation and design of ion thrusters is discussed, and the important processes which influence the plasma properties are described in terms of the design and control parameters. The conservation for mass, charge and energy were applied to the ion production region, which was defined as the region of the discharge chamber having as its outer boundary the surface of revolution of the innermost field line to intersect the anode. Mass conservation and the equations describing the various processes involved with mass addition and removal from the ion production region are satisfied by a Maxwellian electron density spatial distribution in that region.

Firing Control Optimization of Impulse Thrusters for Trajectory Correction Projectiles

Directory of Open Access Journals (Sweden)

Min Gao

2015-01-01

Full Text Available This paper presents an optimum control scheme of firing time and firing phase angle by taking impact point deviation as optimum objective function which takes account of the difference of longitudinal and horizontal correction efficiency, firing delay, roll rate, flight stability, and so forth. Simulations indicate that this control scheme can assure lateral impulse thrusters are activated at time and phase angle when the correction efficiency is higher. Further simulations show that the impact point dispersion is mainly influenced by the total impulse deployed, and the impulse, number, and firing interval need to be optimized to reduce the impact point dispersion of rockets. Live firing experiments with two trajectory correction rockets indicate that the firing control scheme works effectively.

Electric Propellant Solid Rocket Motor Thruster Results Enabling Small Satellites

OpenAIRE

Koehler, Frederick; Langhenry, Mark; Summers, Matt; Villarreal, James; Villarreal, Thomas

2017-01-01

Raytheon Missile Systems has developed and tested true on/off/restart solid propellant thrusters which are controlled only by electrical current. This new patented class of energetic rocket propellant is safe, controllable and simple. The range of applications for this game changing technology includes attitude control systems and a safe alternative to higher impulse space satellite thrusters. Described herein are descriptions and performance data for several small electric propellant solid r...

Diagnosis and Fault-Tolerant Control for Thruster-Assisted Position Mooring System

DEFF Research Database (Denmark)

Nguyen, Trong Dong; Blanke, Mogens; Sørensen, Asgeir

2007-01-01

Development of fault-tolerant control systems is crucial to maintain safe operation of o®shore installations. The objective of this paper is to develop a fault- tolerant control for thruster-assisted position mooring (PM) system with faults occurring in the mooring lines. Faults in line......'s pretension or line breaks will degrade the performance of the positioning of the vessel. Faults will be detected and isolated through a fault diagnosis procedure. When faults are detected, they can be accommodated through the control action in which only parameter of the controlled plant has to be updated...... to cope with the faulty condition. Simulations will be carried out to verify the advantages of the fault-tolerant control strategy for the PM system....

Advanced laboratory for testing plasma thrusters and Hall thruster measurement campaign

Directory of Open Access Journals (Sweden)

Szelecka Agnieszka

2016-06-01

Full Text Available Plasma engines are used for space propulsion as an alternative to chemical thrusters. Due to the high exhaust velocity of the propellant, they are more efficient for long-distance interplanetary space missions than their conventional counterparts. An advanced laboratory of plasma space propulsion (PlaNS at the Institute of Plasma Physics and Laser Microfusion (IPPLM specializes in designing and testing various electric propulsion devices. Inside of a special vacuum chamber with three performance pumps, an environment similar to the one that prevails in space is created. An innovative Micro Pulsed Plasma Thruster (LμPPT with liquid propellant was built at the laboratory. Now it is used to test the second prototype of Hall effect thruster (HET operating on krypton propellant. Meantime, an improved prototype of krypton Hall thruster is constructed.

Control Valve for Miniature Xenon Ion Thruster, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — NASA is continuing its development of electric propulsion engines for various applications. Efforts have been directed toward both large and small thrusters,...

Thermal Modeling for Pulsed Inductive FRC Plasmoid Thrusters

Science.gov (United States)

Pfaff, Michael

Due to the rising importance of space based infrastructure, long-range robotic space missions, and the need for active attitude control for spacecraft, research into Electric Propulsion is becoming increasingly important. Electric Propulsion (EP) systems utilize electric power to accelerate ions in order to produce thrust. Unlike traditional chemical propulsion, this means that thrust levels are relatively low. The trade-off is that EP thrusters have very high specific impulses (Isp), and can therefore make do with far less onboard propellant than cold gas, monopropellant, or bipropellant engines. As a consequence of the high power levels used to accelerate the ionized propellant, there is a mass and cost penalty in terms of solar panels and a power processing unit. Due to the large power consumption (and waste heat) from electric propulsion thrusters, accurate measurements and predictions of thermal losses are needed. Excessive heating in sensitive locations within a thruster may lead to premature failure of vital components. Between the fixed cost required to purchase these components, as well as the man-hours needed to assemble (or replace) them, attempting to build a high-power thruster without reliable thermal modeling can be expensive. This paper will explain the usage of FEM modeling and experimental tests in characterizing the ElectroMagnetic Plasmoid Thruster (EMPT) and the Electrodeless Lorentz Force (ELF) thruster at the MSNW LLC facility in Redmond, Washington. The EMPT thruster model is validated using an experimental setup, and steady state temperatures are predicted for vacuum conditions. Preliminary analysis of the ELF thruster indicates possible material failure in absence of an active cooling system for driving electronics and for certain power levels.

Optimization of Cylindrical Hall Thrusters

International Nuclear Information System (INIS)

Raitses, Yevgeny; Smirnov, Artem; Granstedt, Erik; Fisch, Nathaniel J.

2007-01-01

The cylindrical Hall thruster features high ionization efficiency, quiet operation, and ion acceleration in a large volume-to-surface ratio channel with performance comparable with the state-of-the-art annular Hall thrusters. These characteristics were demonstrated in low and medium power ranges. Optimization of miniaturized cylindrical thrusters led to performance improvements in the 50-200W input power range, including plume narrowing, increased thruster efficiency, reliable discharge initiation, and stable operation.

Optimization of Cylindrical Hall Thrusters

International Nuclear Information System (INIS)

Raitses, Yevgeny; Smirnov, Artem; Granstedt, Erik; Fi, Nathaniel J.

2007-01-01

The cylindrical Hall thruster features high ionization efficiency, quiet operation, and ion acceleration in a large volume-to-surface ratio channel with performance comparable with the state-of-the-art annular Hall thrusters. These characteristics were demonstrated in low and medium power ranges. Optimization of miniaturized cylindrical thrusters led to performance improvements in the 50-200W input power range, including plume narrowing, increased thruster efficiency, reliable discharge initiation, and stable operation

Diagnostics Systems for Permanent Hall Thrusters Development

Science.gov (United States)

Ferreira, Jose Leonardo; Soares Ferreira, Ivan; Santos, Jean; Miranda, Rodrigo; Possa, M. Gabriela

-Effect Thruster (PMHET), developed at the Plasma Physics Laboratory of UnB. The idea of using an array of permanent magnets, instead of an electromagnet, to produce a radial magnetic field inside the cylindrical plasma drift channel of the thruster is very attractive, especially because of the possibility of developing a HET with power consumption low enough to be used in small satellites or medium-size satellites with low on board power. Hall-Effect Thrusters are now a very good option for spacecraft primary propulsion and also for station-keeping of medium and large satellites. This is because of their high specific impulse, efficient use of propellant mass and combined low and precise thrust capabilities, which are related to an economy in terms of propellant mass utilization , longer satellite lifetime and easier spacecraft maneuvering in microgravity environment. The first HETs were developed in the mid 1950’s, and they were first called Closed Drift Thrusters. Today, the successful use of electric thrusters for attitude control and orbit modification on hundreds of satellites shows the advanced stage of development of this technology. In addition to this, after the success of space missions such as Deep Space One and Dawn (NASA), Hayabusa (JAXA) and Smart-1 (ESA), the employment of electric thrusters is also consolidated for the primary propulsion of spacecraft. This success is mainly due to three factors: reliability of this technology; efficiency of propellant utilization, and therefore reduction of the initial mass of the ship; possibility of operation over long time intervals, with practically unlimited cycling and restarts. This thrusting system is designed to be used in satellite attitude control and long term space missions. One of the greatest advantage of this kind of thruster is the production of a steady state magnetic field by permanent magnets providing electron trapping and Hall current generation within a significant decrease on the electric energy supply

Cylindrical Hall Thrusters with Permanent Magnets

International Nuclear Information System (INIS)

Raitses, Yevgeny; Merino, Enrique; Fisch, Nathaniel J.

2010-01-01

The use of permanent magnets instead of electromagnet coils for low power Hall thrusters can offer a significant reduction of both the total electric power consumption and the thruster mass. Two permanent magnet versions of the miniaturized cylindrical Hall thruster (CHT) of different overall dimensions were operated in the power range of 50W-300 W. The discharge and plasma plume measurements revealed that the CHT thrusters with permanent magnets and electromagnet coils operate rather differently. In particular, the angular ion current density distribution from the permanent magnet thrusters has an unusual halo shape, with a majority of high energy ions flowing at large angles with respect to the thruster centerline. Differences in the magnetic field topology outside the thruster channel and in the vicinity of the channel exit are likely responsible for the differences in the plume characteristics measured for the CHTs with electromagnets and permanent magnets. It is shown that the presence of the reversing-direction or cusp-type magnetic field configuration inside the thruster channel without a strong axial magnetic field outside the thruster channel does not lead to the halo plasma plume from the CHT.

HG ion thruster component testing

Science.gov (United States)

Mantenieks, M. A.

1979-01-01

Cathodes, isolators, and vaporizers are critical components in determining the performance and lifetime of mercury ion thrusters. The results of life tests of several of these components are reported. A 30-cm thruster CIV test in a bell jar has successfully accumulated over 26,000 hours. The cathode has undergone 65 restarts during the life test without requiring any appreciable increases in starting power. Recently, all restarts have been achieved with only the 44 volt keeper supply with no change required in the starting power. Another ongoing 30-cm Hg thruster cathode test has successfully passed the 10,000 hour mark. A solid-insert, 8-cm thruster cathode has accumulated over 4,000 hours of thruster operation. All starts have been achieved without the use of a high voltage ignitor. The results of this test indicate that the solid impregnated insert is a viable neutralizer cathode for the 8-cm thruster.

Integrated Stirling Convertor and Hall Thruster Test Conducted

Science.gov (United States)

Mason, Lee S.

2002-01-01

An important aspect of implementing Stirling Radioisotope Generators on future NASA missions is the integration of the generator and controller with potential spacecraft loads. Some recent studies have indicated that the combination of Stirling Radioisotope Generators and electric propulsion devices offer significant trip time and payload fraction benefits for deep space missions. A test was devised to begin to understand the interactions between Stirling generators and electric thrusters. An electrically heated RG- 350 (350-W output) Stirling convertor, designed and built by Stirling Technology Company of Kennewick, Washington, under a NASA Small Business Innovation Research agreement, was coupled to a 300-W SPT-50 Hall-effect thruster built for NASA by the Moscow Aviation Institute (RIAME). The RG-350 and the SPT-50 shown, were installed in adjacent vacuum chamber ports at NASA Glenn Research Center's Electric Propulsion Laboratory, Vacuum Facility 8. The Stirling electrical controller interfaced directly with the Hall thruster power-processing unit, both of which were located outside of the vacuum chamber. The power-processing unit accepted the 48 Vdc output from the Stirling controller and distributed the power to all the loads of the SPT-50, including the magnets, keeper, heater, and discharge. On February 28, 2001, the Glenn test team successfully operated the Hall-effect thruster with the Stirling convertor. This is the world's first known test of a dynamic power source with electric propulsion. The RG-350 successfully managed the transition from the purely resistive load bank within the Stirling controller to the highly capacitive power-processing unit load. At the time of the demonstration, the Stirling convertor was operating at a hot temperature of 530 C and a cold temperature of -6 C. The linear alternator was producing approximately 250 W at 109 Vac, while the power-processing unit was drawing 175 W at 48 Vdc. The majority of power was delivered to the

Enhanced Performance of Cylindrical Hall Thrusters

International Nuclear Information System (INIS)

Raitses, Y.; Smirnov, A.; Fisch, N.J.

2007-01-01

The cylindrical thruster differs significantly in its underlying physical mechanisms from the conventional annular Hall thruster. It features high ionization efficiency, quiet operation, ion acceleration in a large volume-to-surface ratio channel, and performance comparable with the state-of-the-art conventional Hall thrusters. Very significant plume narrowing, accompanied by the increase of the energetic ion fraction and improvement of ion focusing, led to 50-60% increase of the thruster anode efficiency. These improvements were achieved by overrunning the discharge current in the magnetized thruster plasma

Project of an ion thruster

International Nuclear Information System (INIS)

Perche, G.E.

1983-07-01

The mercury bombardment electrostatic ion thruster is the most successful electric thruster available today. This work describes a 5 cm diameter ion thruster with 3.000 s specific impulse and 5 mN thrust. The advantages of electric propulsion and the tests that will be performed are also presented. (Author) [pt

Single Cathode Ion Thruster

Data.gov (United States)

National Aeronautics and Space Administration — Objective is to design an electrostatic ion thruster that is more efficient, simpler, and lower cost than the current gridded ion thruster. Initial objective is to...

Theoretical and experimental study of a thruster discharging a weight

Science.gov (United States)

Michaels, Dan; Gany, Alon

2014-06-01

An innovative concept for a rocket type thruster that can be beneficial for spacecraft trajectory corrections and station keeping was investigated both experimentally and theoretically. It may also be useful for divert and attitude control systems (DACS). The thruster is based on a combustion chamber discharging a weight through an exhaust tube. Calculations with granular double-base propellant and a solid ejected weight reveal that a specific impulse based on the propellant mass of well above 400 s can be obtained. An experimental thruster was built in order to demonstrate the new idea and validate the model. The thruster impulse was measured both directly with a load cell and indirectly by using a pressure transducer and high speed photography of the weight as it exits the tube, with both ways producing very similar total impulse measurement. The good correspondence between the computations and the measured data validates the model as a useful tool for studying and designing such a thruster.

Fuzzy based attitude controller for flexible spacecraft with on/off thrusters

Science.gov (United States)

Knapp, Roger Glenn

1993-05-01

A fuzzy-based attitude controller is designed for attitude control of a generic spacecraft with on/off thrusters. The controller is comprised of packages of rules dedicated to addressing different objectives (e.g., disturbance rejection, low fuel consumption, avoiding the excitation of flexible appendages, etc.). These rule packages can be inserted or removed depending on the requirements of the particular spacecraft and are parameterized based on vehicle parameters such as inertia or operational parameters such as the maneuvering rate. Individual rule packages can be 'weighted' relative to each other to emphasize the importance of one objective relative to another. Finally, the fuzzy controller and rule packages are demonstrated using the high-fidelity Space Shuttle Interactive On-Orbit Simulator (IOS) while performing typical on-orbit operations and are subsequently compared with the existing shuttle flight control system performance.

Analysis of state-of-the-art single-thruster attitude control techniques for spinning penetrator

Science.gov (United States)

Raus, Robin; Gao, Yang; Wu, Yunhua; Watt, Mark

2012-07-01

The attitude dynamics and manoeuvre survey in this paper is performed for a mission scenario involving a penetrator-type spacecraft: an axisymmetric prolate spacecraft spinning around its minor axis of inertia performing a 90° spin axis reorientation manoeuvre. In contrast to most existing spacecraft only one attitude control thruster is available, providing a control torque perpendicular to the spin axis. Having only one attitude thruster on a spinning spacecraft could be preferred for spacecraft simplicity (lower mass, lower power consumption etc.), or it could be imposed in the context of redundancy/contingency operations. This constraint does yield restrictions on the thruster timings, depending on the ratio of minor to major moments of inertia among other parameters. The Japanese Lunar-A penetrator spacecraft proposal is a good example of such a single-thruster spin-stabilised prolate spacecraft. The attitude dynamics of a spinning rigid body are first investigated analytically, then expanded for the specific case of a prolate and axisymmetric rigid body and finally a cursory exploration of non-rigid body dynamics is made. Next two well-known techniques for manoeuvring a spin-stabilised spacecraft, the Half-cone/Multiple Half-cone and the Rhumb line slew, are compared with two new techniques, the Sector-Arc Slew developed by Astrium Satellites and the Dual-cone developed at Surrey Space Centre. Each technique is introduced and characterised by means of simulation results and illustrations based on the penetrator mission scenario and a brief robustness analysis is performed against errors in moments of inertia and spin rate. Also, the relative benefits of each slew algorithm are discussed in terms of slew accuracy, energy (propellant) efficiency and time efficiency. For example, a sequence of half-cone manoeuvres (a Multi-half-cone manoeuvre) tends to be more energy-efficient than one half-cone for the same final slew angle, but more time-consuming. As another

Advanced Development of a Compact 5-15 lbf Lox/Methane Thruster for an Integrated Reaction Control and Main Engine Propulsion System

Science.gov (United States)

Hurlbert, Eric A.; McManamen, John Patrick; Sooknanen, Josh; Studak, Joseph W.

2011-01-01

This paper describes the advanced development and testing of a compact 5 to 15 lbf LOX/LCH4 thruster for a pressure-fed integrated main engine and RCS propulsion system to be used on a spacecraft "vertical" test bed (VTB). The ability of the RCS thruster and the main engine to operate off the same propellant supply in zero-g reduces mass and improves mission flexibility. This compact RCS engine incorporates several features to dramatically reduce mass and parts count, to ease manufacturing, and to maintain acceptable performance given that specific impulse (Isp) is not the driver. For example, radial injection holes placed on the chamber body for easier drilling, and high temperature Haynes 230 were selected for the chamber over other more expensive options. The valve inlets are rotatable before welding allowing different orientations for vehicle integration. In addition, the engine design effort selected a coil-on-plug ignition system which integrates a relay and coil with the plug electrode, and moves some exciter electronics to avionics driver board. The engine injector design has small dribble volumes to target minimum pulse widths of 20 msec. and an efficient minimum impulse bit of less than 0.05 lbf-sec. The propellants, oxygen and methane, were chosen because together they are a non-toxic, Mars-forward, high density, space storable, and high performance propellant combination that is capable of pressure-fed and pump-fed configurations and integration with life support and power subsystems. This paper will present the results of the advanced development testing to date of the RCS thruster and the integration with a vehicle propulsion system.

Contamination Study of Micro Pulsed Plasma Thruster

National Research Council Canada - National Science Library

Kesenek, Ceylan

2008-01-01

.... Micro-Pulsed Plasma Thrusters (PPTs) are highly reliable and simple micro propulsion systems that will offer attitude control, station keeping, constellation flying, and drag compensation for such satellites...

NASA's Evolutionary Xenon Thruster (NEXT) Project Qualification Propellant Throughput Milestone: Performance, Erosion, and Thruster Service Life Prediction After 450 kg

Science.gov (United States)

Herman, Daniel A.

2010-01-01

The NASA s Evolutionary Xenon Thruster (NEXT) program is tasked with significantly improving and extending the capabilities of current state-of-the-art NSTAR thruster. The service life capability of the NEXT ion thruster is being assessed by thruster wear test and life-modeling of critical thruster components, such as the ion optics and cathodes. The NEXT Long-Duration Test (LDT) was initiated to validate and qualify the NEXT thruster propellant throughput capability. The NEXT thruster completed the primary goal of the LDT; namely to demonstrate the project qualification throughput of 450 kg by the end of calendar year 2009. The NEXT LDT has demonstrated 28,500 hr of operation and processed 466 kg of xenon throughput--more than double the throughput demonstrated by the NSTAR flight-spare. Thruster performance changes have been consistent with a priori predictions. Thruster erosion has been minimal and consistent with the thruster service life assessment, which predicts the first failure mode at greater than 750 kg throughput. The life-limiting failure mode for NEXT is predicted to be loss of structural integrity of the accelerator grid due to erosion by charge-exchange ions.

Performance and Qualification of the Power Supply and Control Unit for the HEMP Thruster

Science.gov (United States)

Brag, R.; Herty, F.

2014-08-01

In 2013, Astrium GmbH delivered several flight model electronics for Electric Propulsion (EP) systems or corresponding components. One of the elements is a Power Supply and Control Unit (PSCU) for the Thales development "High Efficiency Multistage Plasma Thruster" (HEMP-T) (see Figure 1). This paper presents the PSCU specification and results of the qualification and acceptance phase of the EQM and the PFM.

Spacecraft attitude and velocity control system

Science.gov (United States)

Paluszek, Michael A. (Inventor); Piper, Jr., George E. (Inventor)

1992-01-01

A spacecraft attitude and/or velocity control system includes a controller which responds to at least attitude errors to produce command signals representing a force vector F and a torque vector T, each having three orthogonal components, which represent the forces and torques which are to be generated by the thrusters. The thrusters may include magnetic torquer or reaction wheels. Six difference equations are generated, three having the form ##EQU1## where a.sub.j is the maximum torque which the j.sup.th thruster can produce, b.sub.j is the maximum force which the j.sup.th thruster can produce, and .alpha..sub.j is a variable representing the throttling factor of the j.sup.th thruster, which may range from zero to unity. The six equations are summed to produce a single scalar equation relating variables .alpha..sub.j to a performance index Z: ##EQU2## Those values of .alpha. which maximize the value of Z are determined by a method for solving linear equations, such as a linear programming method. The Simplex method may be used. The values of .alpha..sub.j are applied to control the corresponding thrusters.

A Small Modular Laboratory Hall Effect Thruster

Science.gov (United States)

Lee, Ty Davis

Electric propulsion technologies promise to revolutionize access to space, opening the door for mission concepts unfeasible by traditional propulsion methods alone. The Hall effect thruster is a relatively high thrust, moderate specific impulse electric propulsion device that belongs to the class of electrostatic thrusters. Hall effect thrusters benefit from an extensive flight history, and offer significant performance and cost advantages when compared to other forms of electric propulsion. Ongoing research on these devices includes the investigation of mechanisms that tend to decrease overall thruster efficiency, as well as the development of new techniques to extend operational lifetimes. This thesis is primarily concerned with the design and construction of a Small Modular Laboratory Hall Effect Thruster (SMLHET), and its operation on argon propellant gas. Particular attention was addressed at low-cost, modular design principles, that would facilitate simple replacement and modification of key thruster parts such as the magnetic circuit and discharge channel. This capability is intended to facilitate future studies of device physics such as anomalous electron transport and magnetic shielding of the channel walls, that have an impact on thruster performance and life. Preliminary results demonstrate SMLHET running on argon in a manner characteristic of Hall effect thrusters, additionally a power balance method was utilized to estimate thruster performance. It is expected that future thruster studies utilizing heavier though more expensive gases like xenon or krypton, will observe increased efficiency and stability.

Colloid Thruster for Attitude Control Systems (ACS) and Tip-off Control Applications, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — Busek proposes to develop and deliver a complete engineering model colloid thruster system, capable of thrust levels and lifetimes required for spacecraft...

Artificial Neural Network Test Support Development for the Space Shuttle PRCS Thrusters

Science.gov (United States)

Lehr, Mark E.

2005-01-01

A significant anomaly, Fuel Valve Pilot Seal Extrusion, is affecting the Shuttle Primary Reaction Control System (PRCS) Thrusters, and has caused 79 to fail. To help address this problem, a Shuttle PRCS Thruster Process Evaluation Team (TPET) was formed. The White Sands Test Facility (WSTF) and Boeing members of the TPET have identified many discrete valve current trace characteristics that are predictive of the problem. However, these are difficult and time consuming to identify and trend by manual analysis. Based on this exhaustive analysis over months, 22 thrusters previously delivered by the Depot were identified as high risk for flight failures. Although these had only recently been installed, they had to be removed from Shuttles OV103 and OV104 for reprocessing, by directive of the Shuttle Project Office. The resulting impact of the thruster removal, replacement, and valve replacement was significant (months of work and hundreds of thousands of dollars). Much of this could have been saved had the proposed Neural Network (NN) tool described in this paper been in place. In addition to the significant benefits to the Shuttle indicated above, the development and implementation of this type of testing will be the genesis for potential Quality improvements across many areas of WSTF test data analysis and will be shared with other NASA centers. Future tests can be designed to incorporate engineering experience via Artificial Neural Nets (ANN) into depot level acceptance of hardware. Additionally, results were shared with a NASA Engineering and Safety Center (NESC) Super Problem Response Team (SPRT). There was extensive interest voiced among many different personnel from several centers. There are potential spin-offs of this effort that can be directly applied to other data acquisition systems as well as vehicle health management for current and future flight vehicles.

Plasma property and performance prediction for mercury ion thrusters

Science.gov (United States)

Longhurst, G. R.; Wilbur, P. J.

1979-01-01

The discharge chambers of mercury ion thrusters are modelled so the principal effects and processes which govern discharge plasma properties and thruster performance are described. The conservation relations for mass, charge and energy when applied to the Maxwellian electron population in the ion production region yield equations which may be made one-dimensional by the proper choice of coordinates. Solutions to these equations with the appropriate boundary conditions give electron density and temperature profiles which agree reasonably well with measurements. It is then possible to estimate plasma properties from thruster design data and those operating parameters which are directly controllable. By varying the operating parameter inputs to the computer code written to solve these equations, perfromance curves are obtained which agree quite well with measurements.

Anode sheath in Hall thrusters

International Nuclear Information System (INIS)

Dorf, L.; Semenov, V.; Raitses, Y.

2003-01-01

A set of hydrodynamic equations is used to describe quasineutral plasma in ionization and acceleration regions of a Hall thruster. The electron distribution function and Poisson equation are invoked for description of a near-anode region. Numerical solutions suggest that steady-state operation of a Hall thruster can be achieved at different anode sheath regimes. It is shown that the anode sheath depends on the thruster operating conditions, namely the discharge voltage and the mass flow rate

Low power arcjet thruster pulse ignition

Science.gov (United States)

Sarmiento, Charles J.; Gruber, Robert P.

1987-01-01

An investigation of the pulse ignition characteristics of a 1 kW class arcjet using an inductive energy storage pulse generator with a pulse width modulated power converter identified several thruster and pulse generator parameters that influence breakdown voltage including pulse generator rate of voltage rise. This work was conducted with an arcjet tested on hydrogen-nitrogen gas mixtures to simulate fully decomposed hydrazine. Over all ranges of thruster and pulser parameters investigated, the mean breakdown voltages varied from 1.4 to 2.7 kV. Ignition tests at elevated thruster temperatures under certain conditions revealed occasional breakdowns to thruster voltages higher than the power converter output voltage. These post breakdown discharges sometimes failed to transition to the lower voltage arc discharge mode and the thruster would not ignite. Under the same conditions, a transition to the arc mode would occur for a subsequent pulse and the thruster would ignite. An automated 11 600 cycle starting and transition to steady state test demonstrated ignition on the first pulse and required application of a second pulse only two times to initiate breakdown.

Krypton Ion Thruster Performance

Science.gov (United States)

Patterson, Michael J.; Williams, George J.

1992-01-01

Preliminary data were obtained from a 30 cm ion thruster operating on krypton propellant over the input power range of 0.4 to 5.5 kW. The data presented are compared and contrasted to the data obtained with xenon propellant over the same input power envelope. Typical krypton thruster efficiency was 70 percent at a specific impulse of approximately 5000 s, with a maximum demonstrated thrust to power ratio of approximately 42 mN/kW at 2090 s specific impulse and 1580 watts input power. Critical thruster performance and component lifetime issues were evaluated. Order of magnitude power throttling was demonstrated using a simplified power-throttling strategy.

Oxygen-Methane Thruster, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Orion Propulsion, Inc. proposes to develop an Oxygen and Methane RCS Thruster to advance the technology of alternate fuels. A successful Oxygen/CH4 RCS Thruster will...

Inert gas thrusters

Science.gov (United States)

Kaufman, H. R.; Robinson, R. S.

1980-01-01

Some advances in component technology for inert gas thrusters are described. The maximum electron emission of a hollow cathode with Ar was increased 60-70% by the use of an enclosed keeper configuration. Operation with Ar, but without emissive oxide, was also obtained. A 30 cm thruster operated with Ar at moderate discharge voltages give double-ion measurements consistent with a double ion correlation developed previously using 15 cm thruster data. An attempt was made to reduce discharge losses by biasing anodes positive of the discharge plasma. The reason this attempt was unsuccessful is not yet clear. The performance of a single-grid ion-optics configuration was evaluated. The ion impingement on the single grid accelerator was found to approach the value expected from the projected blockage when the sheath thickness next to the accelerator was 2-3 times the aperture diameter.

NSTAR Ion Thruster and Breadboard Power Processor Functional Integration Test Results

Science.gov (United States)

Hamley, John A.; Pinero, Luis R.; Rawlin, Vincent K.; Miller, John R.; Myers, Roger M.; Bowers, Glen E.

1996-01-01

A 2.3 kW Breadboard Power Processing Unit (BBPPU) was developed as part of the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) Program. The NSTAR program will deliver an electric propulsion system based on a 30 cm xenon ion thruster to the New Millennium (NM) program for use as the primary propulsion system for the initial NM flight. The final development test for the BBPPU, the Functional Integration Test, was carried out to demonstrate all aspects of BBPPU operation with an Engineering Model Thruster. Test objectives included: (1) demonstration and validation of automated thruster start procedures, (2) demonstration of stable closed loop control of the thruster beam current, (3) successful response and recovery to thruster faults, and (4) successful safing of the system during simulated spacecraft faults. These objectives were met over the specified 80-120 VDC input voltage range and 0.5-2.3 output power capability of the BBPPU. Two minor anomalies were noted in discharge and neutralizer keeper current. These anomalies did not affect the stability of the system and were successfully corrected.

Parametric Investigation of Miniaturized Cylindrical and Annular Hall Thrusters

International Nuclear Information System (INIS)

Smirnov, A.; Raitses, Y.; Fisch, N.J.

2002-01-01

Conventional annular Hall thrusters become inefficient when scaled to low power. An alternative approach, a 2.6-cm miniaturized cylindrical Hall thruster with a cusp-type magnetic field distribution, was developed and studied. Its performance was compared to that of a conventional annular thruster of the same dimensions. The cylindrical thruster exhibits discharge characteristics similar to those of the annular thruster, but it has a much higher propellant ionization efficiency. Significantly, a large fraction of multi-charged xenon ions might be present in the outgoing ion flux generated by the cylindrical thruster. The operation of the cylindrical thruster is quieter than that of the annular thruster. The characteristic peak in the discharge current fluctuation spectrum at 50-60 kHz appears to be due to ionization instabilities. In the power range 50-300 W, the cylindrical and annular thrusters have comparable efficiencies (15-32%) and thrusts (2.5-12 mN). For the annular configuration, a voltage less than 200 V was not sufficient to sustain the discharge at low propellant flow rates. The cylindrical thruster can operate at voltages lower than 200 V, which suggests that a cylindrical thruster can be designed to operate at even smaller power

Scale Model Thruster Acoustic Measurement Results

Science.gov (United States)

Vargas, Magda; Kenny, R. Jeremy

2013-01-01

The Space Launch System (SLS) Scale Model Acoustic Test (SMAT) is a 5% scale representation of the SLS vehicle, mobile launcher, tower, and launch pad trench. The SLS launch propulsion system will be comprised of the Rocket Assisted Take-Off (RATO) motors representing the solid boosters and 4 Gas Hydrogen (GH2) thrusters representing the core engines. The GH2 thrusters were tested in a horizontal configuration in order to characterize their performance. In Phase 1, a single thruster was fired to determine the engine performance parameters necessary for scaling a single engine. A cluster configuration, consisting of the 4 thrusters, was tested in Phase 2 to integrate the system and determine their combined performance. Acoustic and overpressure data was collected during both test phases in order to characterize the system's acoustic performance. The results from the single thruster and 4- thuster system are discussed and compared.

Q-Thruster Breadboard Campaign Project

Science.gov (United States)

White, Harold

2014-01-01

Dr. Harold "Sonny" White has developed the physics theory basis for utilizing the quantum vacuum to produce thrust. The engineering implementation of the theory is known as Q-thrusters. During FY13, three test campaigns were conducted that conclusively demonstrated tangible evidence of Q-thruster physics with measurable thrust bringing the TRL up from TRL 2 to early TRL 3. This project will continue with the development of the technology to a breadboard level by leveraging the most recent NASA/industry test hardware. This project will replace the manual tuning process used in the 2013 test campaign with an automated Radio Frequency (RF) Phase Lock Loop system (precursor to flight-like implementation), and will redesign the signal ports to minimize RF leakage (improves efficiency). This project will build on the 2013 test campaign using the above improvements on the test implementation to get ready for subsequent Independent Verification and Validation testing at Glenn Research Center (GRC) and Jet Propulsion Laboratory (JPL) in FY 2015. Q-thruster technology has a much higher thrust to power than current forms of electric propulsion (7x Hall thrusters), and can significantly reduce the total power required for either Solar Electric Propulsion (SEP) or Nuclear Electric Propulsion (NEP). Also, due to the high thrust and high specific impulse, Q-thruster technology will greatly relax the specific mass requirements for in-space nuclear reactor systems. Q-thrusters can reduce transit times for a power-constrained architecture.

One-millipound mercury ion thruster

Science.gov (United States)

Hyman, J., Jr.; Dulgeroff, C. R.; Kami, S.; Williamson, W. S.

1975-01-01

A mercury ion thruster has been developed for efficient operation at the nominal 1-mlb thrust level with a specific impulse of about 3,000 sec and a total power consumption of about 120 W. At a beam voltage of 1,200 V and beam current of 72 mA, the discharge chamber operates with a propellant efficiency of 93.8% at an ion-generation energy of 276 eV/ion. The 8-cm diameter thruster advances proven component technology to assure the capability for thruster operation over an accumulated beam-on time in excess of 20,000 hours with a capability for 10,000 on-off duty cycles. Discharge chamber optimization has combined stable current-voltage characteristics with high performance efficiency by careful placement of the discharge cathode near the location of a magnetic-field zero just upstream of the thruster endplate.

Electrostatic ion thrusters - towards predictive modeling

Energy Technology Data Exchange (ETDEWEB)

Kalentev, O.; Matyash, K.; Duras, J.; Lueskow, K.F.; Schneider, R. [Ernst-Moritz-Arndt Universitaet Greifswald, D-17489 (Germany); Koch, N. [Technische Hochschule Nuernberg Georg Simon Ohm, Kesslerplatz 12, D-90489 Nuernberg (Germany); Schirra, M. [Thales Electronic Systems GmbH, Soeflinger Strasse 100, D-89077 Ulm (Germany)

2014-02-15

The development of electrostatic ion thrusters so far has mainly been based on empirical and qualitative know-how, and on evolutionary iteration steps. This resulted in considerable effort regarding prototype design, construction and testing and therefore in significant development and qualification costs and high time demands. For future developments it is anticipated to implement simulation tools which allow for quantitative prediction of ion thruster performance, long-term behavior and space craft interaction prior to hardware design and construction. Based on integrated numerical models combining self-consistent kinetic plasma models with plasma-wall interaction modules a new quality in the description of electrostatic thrusters can be reached. These open the perspective for predictive modeling in this field. This paper reviews the application of a set of predictive numerical modeling tools on an ion thruster model of the HEMP-T (High Efficiency Multi-stage Plasma Thruster) type patented by Thales Electron Devices GmbH. (copyright 2014 WILEY-VCH Verlag GmbH and Co. KGaA, Weinheim) (orig.)

NASA HERMeS Hall Thruster Electrical Configuration Characterization

Science.gov (United States)

Peterson, Peter; Kamhawi, Hani; Huang, Wensheng; Yim, John; Herman, Daniel; Williams, George; Gilland, James; Hofer, Richard

2016-01-01

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight ready propulsion system. Part of the technology maturation was to test the TDU-1 thruster in several ground based electrical configurations to assess the thruster robustness and suitability to successful in-space operation. The ground based electrical configuration testing has recently been demonstrated as an important step in understanding and assessing how a Hall thruster may operate differently in space compared to ground based testing, and to determine the best configuration to conduct development and qualification testing. This presentation will cover the electrical configuration testing of the TDU-1 HERMeS Hall thruster in NASA Glenn Research Centers Vacuum Facility 5. The three electrical configurations examined are the thruster body tied to facility ground, thruster floating, and finally the thruster body electrically tied to cathode common. The TDU-1 HERMeS was configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

Iodine Hall Thruster Propellant Feed System for a CubeSat

Science.gov (United States)

Polzin, Kurt A.; Peeples, Steven

2014-01-01

The components required for an in-space iodine vapor-fed Hall effect thruster propellant management system are described. A laboratory apparatus was assembled and used to produce iodine vapor and control the flow through the application of heating to the propellant reservoir and through the adjustment of the opening in a proportional flow control valve. Changing of the reservoir temperature altered the flowrate on the timescale of minutes while adjustment of the proportional flow control valve changed the flowrate immediately without an overshoot or undershoot in flowrate with the requisite recovery time associated with thermal control systems. The flowrates tested spanned a range from 0-1.5 mg/s of iodine, which is sufficient to feed a 200-W Hall effect thruster.

Fabrication of LTCC based Micro Thruster for Precision Controlled Spaceflight

DEFF Research Database (Denmark)

Larsen, Jack; Jørgensen, John Leif

2011-01-01

The paper at hand presents the initial investigations on the development and fabrication of a micro thruster based on LTCC technology, delivering a thrust in the micro Newton regime. Using smaller segments of an observation system distributed on two or more spacecrafts, one can realize an observa...

Validation Test Results for Orthogonal Probe Eddy Current Thruster Inspection System

Science.gov (United States)

Wincheski, Russell A.

2007-01-01

Recent nondestructive evaluation efforts within NASA have focused on an inspection system for the detection of intergranular cracking originating in the relief radius of Primary Reaction Control System (PCRS) Thrusters. Of particular concern is deep cracking in this area which could lead to combustion leakage in the event of through wall cracking from the relief radius into an acoustic cavity of the combustion chamber. In order to reliably detect such defects while ensuring minimal false positives during inspection, the Orthogonal Probe Eddy Current (OPEC) system has been developed and an extensive validation study performed. This report describes the validation procedure, sample set, and inspection results as well as comparing validation flaws with the response from naturally occuring damage.

The development of the micro-solid propellant thruster array with improved repeatability

International Nuclear Information System (INIS)

Seo, Daeban; Kwon, Sejin; Lee, Jongkwang

2012-01-01

This paper presents the development of a micro-solid propellant thruster array with improved repeatability. The repeatability and low performance variation of each thruster unit with a high ignition success rate is essential in micro-solid propellant thruster array. To date, the study on the improvement of the repeatability has not yet been reported. As the first step for this study, we propose a new type of micro igniter, using a glass wafer called the heater-contact micro igniter. This igniter is also designed to improve the ignition characteristics of a glass-based micro igniter. The prototype of the igniter array is designed and fabricated to establish its fabrication process and to conduct its performance evaluation. Through the firing test, the performance of the heater-contact micro igniter is verified. The 5 × 5 sized micro-solid propellant thruster array is designed and fabricated applying the developed heater-contact igniter. The measured average thrust of each thruster unit is 2.542 N, and calculated standard deviation is 0.369 N. The calculated average total impulse and its standard deviation are 0.182 and 0.04 mNs, respectively. Based on these results, the improvement of repeatability is verified. Finally, the ignition control system of the micro-thruster array is developed. (paper)

Hall Thruster Thermal Modeling and Test Data Correlation

Science.gov (United States)

Myers, James; Kamhawi, Hani; Yim, John; Clayman, Lauren

2016-01-01

The life of Hall Effect thrusters are primarily limited by plasma erosion and thermal related failures. NASA Glenn Research Center (GRC) in cooperation with the Jet Propulsion Laboratory (JPL) have recently completed development of a Hall thruster with specific emphasis to mitigate these limitations. Extending the operational life of Hall thursters makes them more suitable for some of NASA's longer duration interplanetary missions. This paper documents the thermal model development, refinement and correlation of results with thruster test data. Correlation was achieved by minimizing uncertainties in model input and recognizing the relevant parameters for effective model tuning. Throughout the thruster design phase the model was used to evaluate design options and systematically reduce component temperatures. Hall thrusters are inherently complex assemblies of high temperature components relying on internal conduction and external radiation for heat dispersion and rejection. System solutions are necessary in most cases to fully assess the benefits and/or consequences of any potential design change. Thermal model correlation is critical since thruster operational parameters can push some components/materials beyond their temperature limits. This thruster incorporates a state-of-the-art magnetic shielding system to reduce plasma erosion and to a lesser extend power/heat deposition. Additionally a comprehensive thermal design strategy was employed to reduce temperatures of critical thruster components (primarily the magnet coils and the discharge channel). Long term wear testing is currently underway to assess the effectiveness of these systems and consequently thruster longevity.

Temperature Gradient in Hall Thrusters

International Nuclear Information System (INIS)

Staack, D.; Raitses, Y.; Fisch, N.J.

2003-01-01

Plasma potentials and electron temperatures were deduced from emissive and cold floating probe measurements in a 2 kW Hall thruster, operated in the discharge voltage range of 200-400 V. An almost linear dependence of the electron temperature on the plasma potential was observed in the acceleration region of the thruster both inside and outside the thruster. This result calls into question whether secondary electron emission from the ceramic channel walls plays a significant role in electron energy balance. The proportionality factor between the axial electron temperature gradient and the electric field is significantly smaller than might be expected by models employing Ohmic heating of electrons

Micro Cathode Arc Thruster for PhoneSat: Development and Potential Applications

Science.gov (United States)

Gazulla, Oriol Tintore; Perez, Andres Dono; Agasid, Elwood; Uribe, Eddie; Trinh, Greenfield; Keidar, Michael; Teel, George; Haque, Samudra; Lukas, Joseph; Salas, Alberto Guillen;

Coaxial plasma thrusters for high specific impulse propulsion

Science.gov (United States)

Schoenberg, Kurt F.; Gerwin, Richard A.; Barnes, Cris W.; Henins, Ivars; Mayo, Robert; Moses, Ronald, Jr.; Scarberry, Richard; Wurden, Glen

1991-01-01

A fundamental basis for coaxial plasma thruster performance is presented and the steady-state, ideal MHD properties of a coaxial thruster using an annular magnetic nozzle are discussed. Formulas for power usage, thrust, mass flow rate, and specific impulse are acquired and employed to assess thruster performance. The performance estimates are compared with the observed properties of an unoptimized coaxial plasma gun. These comparisons support the hypothesis that ideal MHD has an important role in coaxial plasma thruster dynamics.

Performance and flow characteristics of MHD seawater thruster

Energy Technology Data Exchange (ETDEWEB)

Doss, E.D.

1990-01-01

The main goal of the research is to investigate the effects of strong magnetic fields on the electrical and flow fields inside MHD thrusters. The results of this study is important in the assessment of the feasibility of MHD seawater propulsion for the Navy. To accomplish this goal a three-dimensional fluid flow computer model has been developed and applied to study the concept of MHD seawater propulsion. The effects of strong magnetic fields on the current and electric fields inside the MHD thruster and their interaction with the flow fields, particularly those in the boundary layers, have been investigated. The results of the three-dimensional computations indicate that the velocity profiles are flatter over the sidewalls of the thruster walls in comparison to the velocity profiles over the electrode walls. These nonuniformities in the flow fields give rise to nonuniform distribution of the skin friction along the walls of the thrusters, where higher values are predicted over the sidewalls relative to those over the electrode walls. Also, a parametric study has been performed using the three-dimensional MHD flow model to analyze the performance of continuous electrode seawater thrusters under different operating parameters. The effects of these parameters on the fluid flow characteristics, and on the thruster efficiency have been investigated. Those parameters include the magnetic field (10--20 T), thruster diameter, surface roughness, flow velocity, and the electric load factor. The results show also that the thruster performance improves with the strength of the magnetic field and thruster diameter, and the efficiency decreases with the flow velocity and surface roughness.

Chemical Characterization and Reactivity of Fuel-Oxidizer Reaction Product

Science.gov (United States)

David, Dennis D.; Dee, Louis A.; Beeson, Harold D.

1997-01-01

Fuel-oxidizer reaction product (FORP), the product of incomplete reaction of monomethylhydrazine and nitrogen tetroxide propellants prepared under laboratory conditions and from firings of Shuttle Reaction Control System thrusters, has been characterized by chemical and thermal analysis. The composition of FORP is variable but falls within a limited range of compositions that depend on three factors: the fuel-oxidizer ratio at the time of formation; whether the composition of the post-formation atmosphere is reducing or oxidizing; and the reaction or post-reaction temperature. A typical composition contains methylhydrazinium nitrate, ammonium nitrate, methylammonium nitrate, and trace amounts of hydrazinium nitrate and 1,1-dimethylhydrazinium nitrate. Thermal decomposition reactions of the FORP compositions used in this study were unremarkable. Neither the various compositions of FORP, the pure major components of FORP, nor mixtures of FORP with propellant system corrosion products showed any unusual thermal activity when decomposed under laboratory conditions. Off-limit thruster operations were simulated by rapid mixing of liquid monomethylhydrazine and liquid nitrogen tetroxide in a confined space. These tests demonstrated that monomethylhydrazine, methylhydrazinium nitrate, ammonium nitrate, or Inconel corrosion products can induce a mixture of monomethylhydrazine and nitrogen tetroxide to produce component-damaging energies. Damaging events required FORP or metal salts to be present at the initial mixing of monomethylhydrazine and nitrogen tetroxide.

Status of the J-series 30-cm mercury ion thruster

Science.gov (United States)

Kami, S.; Dulgeroff, C. R.; Bechtel, R. T.

1982-01-01

This paper describes the status of the 30-cm J-series mercury ion thruster. This thruster was baselined for the Solar Electric Propulsion System (SEPS) vehicle. This thruster is described and several modifications plus suggested modifications are presented. Some of the modifications resulted from tests performed with the thruster. The operational characteristics of eight J-series thrusters are presented. Isolator contamination and flake formation are also discussed.

Los Alamos NEP research in advanced plasma thrusters

Science.gov (United States)

Schoenberg, Kurt; Gerwin, Richard

1991-01-01

Research was initiated in advanced plasma thrusters that capitalizes on lab capabilities in plasma science and technology. The goal of the program was to examine the scaling issues of magnetoplasmadynamic (MPD) thruster performance in support of NASA's MPD thruster development program. The objective was to address multi-megawatt, large scale, quasi-steady state MPD thruster performance. Results to date include a new quasi-steady state operating regime which was obtained at space exploration initiative relevant power levels, that enables direct coaxial gun-MPD comparisons of thruster physics and performance. The radiative losses are neglible. Operation with an applied axial magnetic field shows the same operational stability and exhaust plume uniformity benefits seen in MPD thrusters. Observed gun impedance is in close agreement with the magnetic Bernoulli model predictions. Spatial and temporal measurements of magnetic field, electric field, plasma density, electron temperature, and ion/neutral energy distribution are underway. Model applications to advanced mission logistics are also underway.

Performance of a Cylindrical Hall-Effect Thruster Using Permanent Magnets

Science.gov (United States)

Polzin, Kurt A.; Raitses, Y.; Merino, E.; Fisch, N. J.

2009-01-01

electromagnets. Data are presented to expose the effect different controllable parameters have on the discharge and to summarize performance measurements (thrust, Isp, efficiency) obtained using a thrust stand. In addition, beam current data are presented to show the effect of the magnetic field topology on the plume profile and current utilization and to gain insight into the thruster s operation. These data extend and improve upon the results previously presented by the authors in Ref. [1].

Particle simulation of grid system for krypton ion thrusters

Directory of Open Access Journals (Sweden)

Maolin CHEN

2018-04-01

Full Text Available The transport processes of plasmas in grid systems of krypton (Kr ion thrusters at different acceleration voltages were simulated with a 3D-PIC model, and the result was compared with xenon (Xe ion thrusters. The variation of the screen grid transparency, the accelerator grid current ratio and the divergence loss were explored. It is found that the screen grid transparency increases with the acceleration voltage and decreases with the beam current, while the accelerator grid current ratio and divergence loss decrease first and then increase with the beam current. This result is the same with Xe ion thrusters. Simulation results also show that Kr ion thrusters have more advantages than Xe ion thrusters, such as higher screen grid transparency, smaller accelerator grid current ratio, larger cut-off current threshold, and better divergence loss characteristic. These advantages mean that Kr ion thrusters have the ability of operating in a wide range of current. Through comprehensive analyses, it can be concluded that using Kr as propellant is very suitable for a multi-mode ion thruster design. Keywords: Grid system, Ion thrusters, Krypton, Particle in cell method, Plasma

The Plasmoid Thruster Experiment (PTX)

Science.gov (United States)

Eskridge, Richard; Martin, Adam; Koelfgen, Syri; Lee, Mike; Smith, James W.

2003-01-01

A plasmoid is a compact plasma structure with an integral magnetic field. They have been studied extensively in controlled fusion research and are categorized according to the relative strength of the poloidal and toroidal magnetic field (B(phi), and B(tau), respectively). An object with B(phi)/B(tau) >> 1 is classified as a Field Reverse Configuration (FRC); if B(phi) = B(tau), it is called a Spheromak. There are a number of possible advantages to using accelerated plasmoids for in-space propulsion. A thruster based on this concept would operate by repetitively producing plasmoids and ejecting them from the device at high velocity. The plasmoid is formed inside of a single turn conical theta-pinch coil; as this process is inductive, there are no life-limiting electrodes. Similar experiments have yielded plasmoid velocities of at least 50 km/s (l), and calculations indicate that velocities in excess of 100 km/s are possible. A thruster based on this concept would be capable of producing an I(sp) in the range of 5,000 - 10,OOO s, with thrust densities of order 10(exp 5) N/m(exp 2). The current experiment is designed to produce jet powers in the range of 5-10 kW, although the concept should be scalable to higher power. The purpose of this experiment is to determine the feasibility of this plasma propulsion concept. To accomplish this, it will be necessary to determine: a.) specific impulse and thrust, b.) efficiency and mass utilization, c.) which type of plasmoid (FRC-like or Spheromak-like) gives the best performance, and d.) the characteristics required of actual thruster components (i.e., switch and capacitor technology). The plasmoid mass and velocity will be measured with a variety of diagnostics, including internal and external B-dot probes, flux loops, Langmuir probes, high-speed cameras, and an interferometer. Simulations of the plasmoid thruster using MOQUI, a time dependent MHD code, will be carried out concurrently with experimental testing. The PTX

Ion thruster performance model

International Nuclear Information System (INIS)

Brophy, J.R.

1984-01-01

A model of ion thruster performance is developed for high flux density cusped magnetic field thruster designs. This model is formulated in terms of the average energy required to produce an ion in the discharge chamber plasma and the fraction of these ions that are extracted to form the beam. The direct loss of high energy (primary) electrons from the plasma to the anode is shown to have a major effect on thruster performance. The model provides simple algebraic equations enabling one to calculate the beam ion energy cost, the average discharge chamber plasma ion energy cost, the primary electron density, the primary-to-Maxwellian electron density ratio and the Maxwellian electron temperature. Experiments indicate that the model correctly predicts the variation in plasma ion energy cost for changes in propellant gas (Ar, Kr, and Xe), grid transparency to neutral atoms, beam extraction area, discharge voltage, and discharge chamber wall temperature

Pressure History Measurement in a Microwave Beaming Thruster

International Nuclear Information System (INIS)

Oda, Yasuhisa; Ushio, Masato; Komurasaki, Kimiya; Takahashi, Koji; Kasugai, Atsushi; Sakamoto, Keishi

2006-01-01

In a microwave beaming thruster with a 1-dimensional nozzle, plasma and shock wave propagates in the nozzle absorbing microwave power. In this study, pressure histories in the thruster are measured using pressure gauges. Measured pressure history at the thruster wall shows constant pressure during plasma propagation in the nozzle. The result of measurement of the propagating velocities of shock wave and plasma shows that both propagate in the same velocity. These result shows that thrust producing model of analogy of pulse detonation engine is successful for the 1D thruster

Investigations of Probe Induced Perturbations in a Hall Thruster

International Nuclear Information System (INIS)

D. Staack; Y. Raitses; N.J. Fisch

2002-01-01

An electrostatic probe used to measure spatial plasma parameters in a Hall thruster generates perturbations of the plasma. These perturbations are examined by varying the probe material, penetration distance, residence time, and the nominal thruster conditions. The study leads us to recommendations for probe design and thruster operating conditions to reduce discharge perturbations, including metal shielding of the probe insulator and operation of the thruster at lower densities

A Numerical Study on Hydrodynamic Interactions between Dynamic Positioning Thrusters

Energy Technology Data Exchange (ETDEWEB)

Jin, Doo Hwa; Lee, Sang Wook [University of Ulsan, Ulsan (Korea, Republic of)

2017-06-15

In this study, we conducted computational fluid dynamics (CFD) simulations for the unsteady hydrodynamic interaction of multiple thrusters by solving Reynolds averaged Navier-Stokes equations. A commercial CFD software, STAR-CCM+ was used for all simulations by employing a ducted thruster model with combination of a propeller and No. 19a duct. A sliding mesh technique was used to treat dynamic motion of propeller rotation and non-conformal hexahedral grid system was considered. Four different combinations in tilting and azimuth angles of the thrusters were considered to investigate the effects on the propulsion performance. We could find that thruster-hull and thruster-thruster interactions has significant effect on propulsion performance and further study will be required for the optimal configurations with the best tilting and relative azimuth angle between thrusters.

Magnetically enhanced vacuum arc thruster

International Nuclear Information System (INIS)

Keidar, Michael; Schein, Jochen; Wilson, Kristi; Gerhan, Andrew; Au, Michael; Tang, Benjamin; Idzkowski, Luke; Krishnan, Mahadevan; Beilis, Isak I

2005-01-01

A hydrodynamic model of the vacuum arc thruster and its plume is described. Primarily an effect of the magnetic field on the plume expansion and plasma generation is considered. Two particular examples are investigated, namely the magnetically enhanced co-axial vacuum arc thruster (MVAT) and the vacuum arc thruster with ring electrodes (RVAT). It is found that the magnetic field significantly decreases the plasma plume radial expansion under typical conditions. Predicted plasma density profiles in the plume of the MVAT are compared with experimental profiles, and generally a good agreement is found. In the case of the RVAT the influence of the magnetic field leads to plasma jet deceleration, which explains the non-monotonic dependence of the ion current density, on an axial magnetic field observed experimentally

Magnetically enhanced vacuum arc thruster

Energy Technology Data Exchange (ETDEWEB)

Keidar, Michael [University of Michigan, Ann Arbor 48109 MI (United States); Schein, Jochen [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Wilson, Kristi [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Gerhan, Andrew [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Au, Michael [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Tang, Benjamin [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Idzkowski, Luke [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Krishnan, Mahadevan [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Beilis, Isak I [Tel Aviv University, Tel Aviv (Israel)

2005-11-01

A hydrodynamic model of the vacuum arc thruster and its plume is described. Primarily an effect of the magnetic field on the plume expansion and plasma generation is considered. Two particular examples are investigated, namely the magnetically enhanced co-axial vacuum arc thruster (MVAT) and the vacuum arc thruster with ring electrodes (RVAT). It is found that the magnetic field significantly decreases the plasma plume radial expansion under typical conditions. Predicted plasma density profiles in the plume of the MVAT are compared with experimental profiles, and generally a good agreement is found. In the case of the RVAT the influence of the magnetic field leads to plasma jet deceleration, which explains the non-monotonic dependence of the ion current density, on an axial magnetic field observed experimentally.

Vacuum arc plasma thrusters with inductive energy storage driver

Science.gov (United States)

Krishnan, Mahadevan (Inventor)

2009-01-01

A plasma thruster with a cylindrical inner and cylindrical outer electrode generates plasma particles from the application of energy stored in an inductor to a surface suitable for the formation of a plasma and expansion of plasma particles. The plasma production results in the generation of charged particles suitable for generating a reaction force, and the charged particles are guided by a magnetic field produced by the same inductor used to store the energy used to form the plasma.

Arcjet space thrusters

Science.gov (United States)

Keefer, Dennis; Rhodes, Robert

1993-05-01

Electrically powered arc jets which produce thrust at high specific impulse could provide a substantial cost reduction for orbital transfer and station keeping missions. There is currently a limited understanding of the complex, nonlinear interactions in the plasma propellant which has hindered the development of high efficiency arc jet thrusters by making it difficult to predict the effect of design changes and to interpret experimental results. A computational model developed at the University of Tennessee Space Institute (UTSI) to study laser powered thrusters and radio frequency gas heaters has been adapted to provide a tool to help understand the physical processes in arc jet thrusters. The approach is to include in the model those physical and chemical processes which appear to be important, and then to evaluate our judgement by the comparison of numerical simulations with experimental data. The results of this study have been presented at four technical conferences. The details of the work accomplished in this project are covered in the individual papers included in the appendix of this report. We present a brief description of the model covering its most important features followed by a summary of the effort.

Thermodynamic Vent System for an On-Orbit Cryogenic Reaction Control Engine

Science.gov (United States)

Hurlbert, Eric A.; Romig, Kris A.; Jimenez, Rafael; Flores, Sam

2012-01-01

A report discusses a cryogenic reaction control system (RCS) that integrates a Joule-Thompson (JT) device (expansion valve) and thermodynamic vent system (TVS) with a cryogenic distribution system to allow fine control of the propellant quality (subcooled liquid) during operation of the device. It enables zero-venting when coupled with an RCS engine. The proper attachment locations and sizing of the orifice are required with the propellant distribution line to facilitate line conditioning. During operations, system instrumentation was strategically installed along the distribution/TVS line assembly, and temperature control bands were identified. A sub-scale run tank, full-scale distribution line, open-loop TVS, and a combination of procured and custom-fabricated cryogenic components were used in the cryogenic RCS build-up. Simulated on-orbit activation and thruster firing profiles were performed to quantify system heat gain and evaluate the TVS s capability to maintain the required propellant conditions at the inlet to the engine valves. Test data determined that a small control valve, such as a piezoelectric, is optimal to provide continuously the required thermal control. The data obtained from testing has also assisted with the development of fluid and thermal models of an RCS to refine integrated cryogenic propulsion system designs. This system allows a liquid oxygenbased main propulsion and reaction control system for a spacecraft, which improves performance, safety, and cost over conventional hypergolic systems due to higher performance, use of nontoxic propellants, potential for integration with life support and power subsystems, and compatibility with in-situ produced propellants.

Non-Toxic Dual Thrust Reaction Control Engine Development for On-Orbit APS Applications

Science.gov (United States)

Robinson, Philip J.; Veith, Eric M.

2003-01-01

A non-toxic dual thrust proof-of-concept demonstration engine was successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the Next Generation Launch Technology (NGLT) program. The demonstration engine utilized the existing Kistler K-1 870 lbf LOX/Ethanol orbital maneuvering engine ( O m ) coupled with some special test equipment (STE) that enabled engine operation at 870 lbf in the primary mode and 25 lbf in the vernier mode. Ambient testing in primary mode varied mixture ratio (MR) from 1.28 to 1.71 and chamber pressure (P(c) from 110 to 181 psia, and evaluated electrical pulse widths (EPW) of 0.080, 0.100 and 0.250 seconds. Altitude testing in vernier mode explored igniter and thruster pulsing characteristics, long duration steady state operation (greater than 420 sec) and the impact of varying the percent fuel film cooling on vernier performance and chamber thermal response at low PC (4 psia). Data produced from the testing provided calibration of the performance and thermal models used in the design of the next version of the dual thrust Reaction Control Engine (RCE).

Technology for Transient Simulation of Vibration during Combustion Process in Rocket Thruster

Science.gov (United States)

Zubanov, V. M.; Stepanov, D. V.; Shabliy, L. S.

2018-01-01

The article describes the technology for simulation of transient combustion processes in the rocket thruster for determination of vibration frequency occurs during combustion. The engine operates on gaseous propellant: oxygen and hydrogen. Combustion simulation was performed using the ANSYS CFX software. Three reaction mechanisms for the stationary mode were considered and described in detail. The way for obtaining quick CFD-results with intermediate combustion components using an EDM model was found. The way to generate the Flamelet library with CFX-RIF was described. A technique for modeling transient combustion processes in the rocket thruster was proposed based on the Flamelet library. A cyclic irregularity of the temperature field like vortex core precession was detected in the chamber. Frequency of flame precession was obtained with the proposed simulation technique.

Real time hardware-in-loop simulation of ESMO satellite attitude control system

Directory of Open Access Journals (Sweden)

Rune Finnset

2006-04-01

Full Text Available This paper studies attitude control of the ESMO satellite using six reaction thrusters. Bang-bang control with dead-zone and Pulse-Width Modulation (PWM for the modulation of the on-time of the thrusters are treated. Closed loop hardware-in-loop simulations, using themicrocontroller unit (MCU Microchip PIC18F452 for implementation of attitude control and MatLab in a standard PC for simulating satellite dynamics, are carried out. Results for real time simulation are compared with autonomous simulations. The controller gives a satisfactory performance in the real time environment.

Performance of a Permanent-Magnet Cylindrical Hall-Effect Thruster

Science.gov (United States)

Polzin, K. A.; Sooby, E. S.; Kimberlin, A. C.; Raites, Y.; Merino, E.; Fisch, N. J.

2009-01-01

The performance of a low-power cylindrical Hall thruster, which more readily lends itself to miniaturization and low-power operation than a conventional (annular) Hall thruster, was measured using a planar plasma probe and a thrust stand. The field in the cylindrical thruster was produced using permanent magnets, promising a power reduction over previous cylindrical thruster iterations that employed electromagnets to generate the required magnetic field topology. Two sets of ring-shaped permanent magnets are used, and two different field configurations can be produced by reorienting the poles of one magnet relative to the other. A plasma probe measuring ion flux in the plume is used to estimate the current utilization for the two magnetic topologies. The measurements indicate that electron transport is impeded much more effectively in one configuration, implying higher thrust efficiency. Thruster performance measurements on this configuration were obtained over a power range of 70-350 W and with the cathode orifice located at three different axial positions relative to the thruster exit plane. The thrust levels over this power range were 1.25-6.5 mN, with anode efficiencies and specific impulses spanning 4-21% and 400-1950 s, respectively. The anode efficiency of the permanent-magnet thruster compares favorable with the efficiency of the electromagnet thruster when the power consumed by the electromagnets is taken into account.

Comparisons in Performance of Electromagnet and Permanent-Magnet Cylindrical Hall-Effect Thrusters

Science.gov (United States)

Polzin, K. A.; Raitses, Y.; Gayoso, J. C.; Fisch, N. J.

2010-01-01

Three different low-power cylindrical Hall thrusters, which more readily lend themselves to miniaturization and low-power operation than a conventional (annular) Hall thruster, are compared to evaluate the propulsive performance of each. One thruster uses electromagnet coils to produce the magnetic field within the discharge channel while the others use permanent magnets, promising power reduction relative to the electromagnet thruster. A magnetic screen is added to the permanent magnet thruster to improve performance by keeping the magnetic field from expanding into space beyond the exit of the thruster. The combined dataset spans a power range from 50-350 W. The thrust levels over this range were 1.3-7.3 mN, with thruster efficiencies and specific impulses spanning 3.5-28.7% and 400-1940 s, respectively. The efficiency is generally higher for the permanent magnet thruster with the magnetic screen, while That thruster s specific impulse as a function of discharge voltage is comparable to the electromagnet thruster.

Mechanical design of SERT 2 thruster system

Science.gov (United States)

Zavesky, R. J.; Hurst, E. B.

1972-01-01

The mechanical design of the mercury bombardment thruster that was tested on SERT is described. The report shows how the structural, thermal, electrical, material compatibility, and neutral mercury coating considerations affected the design and integration of the subsystems and components. The SERT 2 spacecraft with two thrusters was launched on February 3, 1970. One thruster operated for 3782 hours and the other for 2011 hours. A high voltage short resulting from buildup of loose eroded material was believed to be the cause of failure.

Studies of Non-Conventional Configuration Closed Electron Drift Thrusters

International Nuclear Information System (INIS)

Y. Raitses; D. Staack; A. Smirnov; A.A. Litvak; L.A. Dorf; T. Graves; N.J. Fisch

2001-01-01

In this paper, we review recent results obtained for segmented electrode and cylindrical Hall thrusters. A low sputtering graphite segmented electrode, placed at the exit of the annular thruster, is shown to affect the plasma potential distribution in the ceramic channel. This effect appears to be correlated with an observed plume reduction compared to a conventional, nonsegmented thruster. In preliminary experiments a 3-cm thruster was operated in the 50-200 W power range. Two operating regimes, stable and oscillating, were observed and investigated

High-Power Ion Thruster Technology

Science.gov (United States)

Beattie, J. R.; Matossian, J. N.

1996-01-01

Performance data are presented for the NASA/Hughes 30-cm-diam 'common' thruster operated over the power range from 600 W to 4.6 kW. At the 4.6-kW power level, the thruster produces 172 mN of thrust at a specific impulse of just under 4000 s. Xenon pressure and temperature measurements are presented for a 6.4-mm-diam hollow cathode operated at emission currents ranging from 5 to 30 A and flow rates of 4 sccm and 8 sccm. Highly reproducible results show that the cathode temperature is a linear function of emission current, ranging from approx. 1000 C to 1150 C over this same current range. Laser-induced fluorescence (LIF) measurements obtained from a 30-cm-diam thruster are presented, suggesting that LIF could be a valuable diagnostic for real-time assessment of accelerator-arid erosion. Calibration results of laminar-thin-film (LTF) erosion badges with bulk molybdenum are presented for 300-eV xenon, krypton, and argon sputtering ions. Facility-pressure effects on the charge-exchange ion current collected by 8-cm-diam and 30-cm-diam thrusters operated on xenon propellant are presented to show that accel current is nearly independent of facility pressure at low pressures, but increases rapidly under high-background-pressure conditions.

Effect of Anode Dielectric Coating on Hall Thruster Operation

International Nuclear Information System (INIS)

Dorf, L.; Raitses, Y.; Fisch, N.J.; Semenov, V.

2003-01-01

An interesting phenomenon observed in the near-anode region of a Hall thruster is that the anode fall changes from positive to negative upon removal of the dielectric coating, which is produced on the anode surface during the normal course of Hall thruster operation. The anode fall might affect the thruster lifetime and acceleration efficiency. The effect of the anode coating on the anode fall is studied experimentally using both biased and emissive probes. Measurements of discharge current oscillations indicate that thruster operation is more stable with the coated anode

Numerical simulation of ammonium dinitramide (ADN)-based non-toxic aerospace propellant decomposition and combustion in a monopropellant thruster

International Nuclear Information System (INIS)

Zhang, Tao; Li, Guoxiu; Yu, Yusong; Sun, Zuoyu; Wang, Meng; Chen, Jun

2014-01-01

Highlights: • Decomposition and combustion process of ADN-based thruster are studied. • Distribution of droplets is obtained during the process of spray hit on wire mesh. • Two temperature models are adopted to describe the heat transfer in porous media. • The influences brought by different mass flux and porosity are studied. - Abstract: Ammonium dinitramide (ADN) monopropellant is currently the most promising among all ‘green propellants’. In this paper, the decomposition and combustion process of liquid ADN-based ternary mixtures for propulsion are numerically studied. The R–R distribution model is used to study the initial boundary conditions of droplet distribution resulting from spray hit on a wire mesh based on PDA experiment. To simulate the heat-transfer characteristics between the gas–solid phases, a two-temperature porous medium model in a catalytic bed is used. An 11-species and 7-reactions chemistry model is used to study the catalytic and combustion processes. The final distribution of temperature, pressure, and other kinds of material component concentrations are obtained using the ADN thruster. The results of simulation conducted in the present study are well agree with previous experimental data, and the demonstration of the ADN thruster confirms that a good steady-state operation is achieved. The effects of spray inlet mass flux and porosity on monopropellant thruster performance are analyzed. The numerical results further show that a larger inlet mass flux results in better thruster performance and a catalytic bed porosity value of 0.5 can exhibit the best thruster performance. These findings can serve as a key reference for designing and testing non-toxic aerospace monopropellant thrusters

Pattern Recognition Control Design

Science.gov (United States)

Gambone, Elisabeth A.

2018-01-01

Spacecraft control algorithms must know the expected vehicle 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 was used to investigate the relationship between the control effector commands and spacecraft responses. Instead of supplying the approximated vehicle properties and the thruster performance characteristics, a database of information relating the thruster ring commands and the desired vehicle response was used for closed-loop control. A Monte Carlo simulation data set of the spacecraft dynamic response to effector commands was analyzed to establish the influence a command has on the behavior of the spacecraft. A tool developed at NASA Johnson Space Center to analyze flight dynamics Monte Carlo data sets through pattern recognition methods was used to perform this analysis. Once a comprehensive data set relating spacecraft responses with commands was established, it was used in place of traditional control methods and gains set. This pattern recognition approach was compared with traditional control algorithms to determine the potential benefits and uses.

Retrofit and verification test of a 30-cm ion thruster

Science.gov (United States)

Dulgeroff, C. R.; Poeschel, R. L.

1980-01-01

Twenty modifications were found to be necessary and were approved by design review. These design modifications were incorporated in the thruster documents (drawings and procedures) to define the J series thruster. Sixteen of the design revisions were implemented in a 900 series thruster by retrofit modification. A standardized set of test procedures was formulated, and the retrofit J series thruster design was verified by test. Some difficulty was observed with the modification to the ion optics assembly, but the overall effect of the design modification satisfies the design objectives. The thruster was tested over a wide range of operating parameters to demonstrate its capabilities.

Electronegative Gas Thruster - Direct Thrust Measurement Project

Science.gov (United States)

Dankanich, John (Principal Investigator); Aanesland, Ane; Polzin, Kurt; Walker, Mitchell

2015-01-01

This effort is an international collaboration and academic partnership to mature an innovative electric propulsion (EP) thruster concept to TRL 3 through direct thrust measurement. The initial target application is for Small Satellites, but can be extended to higher power. The Plasma propulsion with Electronegative GASES (PEGASES) concept simplifies ion thruster operation, eliminates a neutralizer requirement and should yield longer life capabilities and lower cost implementation over conventional gridded ion engines. The basic proof-of concept has been demonstrated and matured to TRL 2 over the past several years by researchers at the Laboratoire de Physique des Plasma in France. Due to the low maturity of the innovation, there are currently no domestic investments in electronegative gas thrusters anywhere within NASA, industry or academia. The end product of this Center Innovation Fund (CIF) project will be a validation of the proof-of-concept, maturation to TRL 3 and technology assessment report to summarize the potential for the PEGASES concept to supplant the incumbent technology. Information exchange with the foreign national will be one-way with the exception of the test results. Those test results will first go through a standard public release ITAR/export control review, and the results will be presented in a public technical forum, and the results will be presented in a public technical forum.

Advanced electrostatic ion thruster for space propulsion

Science.gov (United States)

Masek, T. D.; Macpherson, D.; Gelon, W.; Kami, S.; Poeschel, R. L.; Ward, J. W.

1978-01-01

The suitability of the baseline 30 cm thruster for future space missions was examined. Preliminary design concepts for several advanced thrusters were developed to assess the potential practical difficulties of a new design. Useful methodologies were produced for assessing both planetary and earth orbit missions. Payload performance as a function of propulsion system technology level and cost sensitivity to propulsion system technology level are among the topics assessed. A 50 cm diameter thruster designed to operate with a beam voltage of about 2400 V is suggested to satisfy most of the requirements of future space missions.

Experimental test of 200 W Hall thruster with titanium wall

Science.gov (United States)

Ding, Yongjie; Sun, Hezhi; Peng, Wuji; Xu, Yu; Wei, Liqiu; Li, Hong; Li, Peng; Su, Hongbo; Yu, Daren

2017-05-01

We designed a 200 W Hall thruster based on the technology of pushing down a magnetic field with two permanent magnetic rings. Boron nitride (BN) is an important insulating wall material for Hall thrusters. The discharge characteristics of the designed Hall thruster were studied by replacing BN with titanium (Ti). Experimental results show that the designed Hall thruster can discharge stably for a long time under a Ti channel. Experiments were performed to determine whether the channel and cathode are electrically connected. When the channel wall and cathode are insulated, the divergence angle of the plume increases, but the performance of the Hall thruster is improved in terms of thrust, specific impulse, anode efficiency, and thrust-to-power ratio. Ti exhibits a powerful antisputtering capability, a low emanation rate of gas, and a large structural strength, making it a potential candidate wall material in the design of low-power Hall thrusters.

Numerical investigation of two interacting parallel thruster-plumes and comparison to experiment

Science.gov (United States)

Grabe, Martin; Holz, André; Ziegenhagen, Stefan; Hannemann, Klaus

2014-12-01

Clusters of orbital thrusters are an attractive option to achieve graduated thrust levels and increased redundancy with available hardware, but the heavily under-expanded plumes of chemical attitude control thrusters placed in close proximity will interact, leading to a local amplification of downstream fluxes and of back-flow onto the spacecraft. The interaction of two similar, parallel, axi-symmetric cold-gas model thrusters has recently been studied in the DLR High-Vacuum Plume Test Facility STG under space-like vacuum conditions, employing a Patterson-type impact pressure probe with slot orifice. We reproduce a selection of these experiments numerically, and emphasise that a comparison of numerical results to the measured data is not straight-forward. The signal of the probe used in the experiments must be interpreted according to the degree of rarefaction and local flow Mach number, and both vary dramatically thoughout the flow-field. We present a procedure to reconstruct the probe signal by post-processing the numerically obtained flow-field data and show that agreement to the experimental results is then improved. Features of the investigated cold-gas thruster plume interaction are discussed on the basis of the numerical results.

50 KW Class Krypton Hall Thruster Performance

Science.gov (United States)

Jacobson, David T.; Manzella, David H.

2003-01-01

The performance of a 50-kilowatt-class Hall thruster designed for operation on xenon propellant was measured using kryton propellant. The thruster was operated at discharge power levels ranging from 6.4 to 72.5 kilowatts. The device produced thrust ranging from 0.3 to 2.5 newtons. The thruster was operated at discharge voltages between 250 and 1000 volts. At the highest anode mass flow rate and discharge voltage and assuming a 100 percent singly charged condition, the discharge specific impulse approached the theoretical value. Discharge specific impulse of 4500 seconds was demonstrated at a discharge voltage of 1000 volts. The peak discharge efficiency was 64 percent at 650 volts.

MEMS-Based Solid Propellant Rocket Array Thruster

Science.gov (United States)

Tanaka, Shuji; Hosokawa, Ryuichiro; Tokudome, Shin-Ichiro; Hori, Keiichi; Saito, Hirobumi; Watanabe, Masashi; Esashi, Masayoshi

The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3 4 W with RK and 4 6 W without RK. Measured impulse thrusts were from 2×10-5 Ns to 3×10-4 Ns after the calculation of compensation for air dumping.

Influence of Triply-Charged Ions and Ionization Cross-Sections in a Hybrid-PIC Model of a Hall Thruster Discharge

Science.gov (United States)

Smith, Brandon D.; Boyd, Iain D.; Kamhawi, Hani

2014-01-01

The sensitivity of xenon ionization rates to collision cross-sections is studied within the framework of a hybrid-PIC model of a Hall thruster discharge. A revised curve fit based on the Drawin form is proposed and is shown to better reproduce the measured crosssections at high electron energies, with differences in the integrated rate coefficients being on the order of 10% for electron temperatures between 20 eV and 30 eV. The revised fit is implemented into HPHall and the updated model is used to simulate NASA's HiVHAc EDU2 Hall thruster at discharge voltages of 300, 400, and 500 V. For all three operating points, the revised cross-sections result in an increase in the predicted thrust and anode efficiency, reducing the error relative to experimental performance measurements. Electron temperature and ionization reaction rates are shown to follow the trends expected based on the integrated rate coefficients. The effects of triply-charged xenon are also assessed. The predicted thruster performance is found to have little or no dependence on the presence of triply-charged ions. The fraction of ion current carried by triply-charged ions is found to be on the order of 1% and increases slightly with increasing discharge voltage. The reaction rates for the 0?III, I?III, and II?III ionization reactions are found to be of similar order of magnitude and are about one order of magnitude smaller than the rate of 0?II ionization in the discharge channel.

Magnetoelectrostatic thruster physical geometry tests

Science.gov (United States)

Ramsey, W. D.

1981-01-01

Inert gas tests are conducted with several magnetoelectrostatic containment discharge chamber geometries. The configurations tested include three discharge chamber lengths; three boundary magnet patterns; two different flux density magnet materials; hemispherical and conical shaped thrusters having different surface-to-volume ratios; and two and three grid ion optics. Argon mass utilizations of 60 to 79% are attained at 210 to 280 eV/ion in different test configurations. Short hemi thruster configurations are found to produce 70 to 92% xenon mass utilization at 185 to 220 eV/ion.

Anode Fall Formation in a Hall Thruster

International Nuclear Information System (INIS)

Dorf, Leonid A.; Raitses, Yevgeny F.; Smirnov, Artem N.; Fisch, Nathaniel J.

2004-01-01

As was reported in our previous work, accurate, nondisturbing near-anode measurements of the plasma density, electron temperature, and plasma potential performed with biased and emissive probes allowed the first experimental identification of both electron-repelling (negative anode fall) and electron-attracting (positive anode fall) anode sheaths in Hall thrusters. An interesting new phenomenon revealed by the probe measurements is that the anode fall changes from positive to negative upon removal of the dielectric coating, which appears on the anode surface during the course of Hall thruster operation. As reported in the present work, energy dispersion spectroscopy analysis of the chemical composition of the anode dielectric coating indicates that the coating layer consists essentially of an oxide of the anode material (stainless steel). However, it is still unclear how oxygen gets into the thruster channel. Most importantly, possible mechanisms of anode fall formation in a Hall thruster with a clean and a coated anodes are analyzed in this work; practical implication of understanding the general structure of the electron-attracting anode sheath in the case of a coated anode is also discussed

MPD thruster research issues, activities, strategies

Science.gov (United States)

1991-01-01

The following activities and plans in the MPD thruster development are summarized: (1) experimental and theoretical research (magnetic nozzles at present and high power levels, MPD thrusters with applied fields extending into the thrust chamber, and improved electrode performance); and (2) tools (MACH2 code for MPD and nozzle flow calculation, laser diagnostics and spectroscopy for non-intrusive measurements of flow conditions, and extension to higher power). National strategies are also outlined.

Low Frequency Plasma Oscillations in a 6-kW Magnetically Shielded Hall Thruster

Science.gov (United States)

Jorns, Benjamin A.; Hofery, Richard R.

2013-01-01

The oscillations from 0-100 kHz in a 6-kW magnetically shielded thruster are experimen- tally characterized. Changes in plasma parameters that result from the magnetic shielding of Hall thrusters have the potential to significantly alter thruster transients. A detailed investigation of the resulting oscillations is necessary both for the purpose of determin- ing the underlying physical processes governing time-dependent behavior in magnetically shielded thrusters as well as for improving thruster models. In this investigation, a high speed camera and a translating ion saturation probe are employed to examine the spatial extent and nature of oscillations from 0-100 kHz in the H6MS thruster. Two modes are identified at 8 kHz and 75-90 kHz. The low frequency mode is azimuthally uniform across the thruster face while the high frequency oscillation is concentrated close to the thruster centerline with an m = 1 azimuthal dependence. These experimental results are discussed in the context of wave theory as well as published observations from an unshielded variant of the H6MS thruster.

Effects of Enhanced Eathode Electron Emission on Hall Thruster Operation

International Nuclear Information System (INIS)

Raitses, Y.; Smirnov, A.; Fisch, N.J.

2009-01-01

Interesting discharge phenomena are observed that have to do with the interaction between the magnetized Hall thruster plasma and the neutralizing cathode. The steadystate parameters of a highly ionized thruster discharge are strongly influenced by the electron supply from the cathode. The enhancement of the cathode electron emission above its self-sustained level affects the discharge current and leads to a dramatic reduction of the plasma divergence and a suppression of large amplitude, low frequency discharge current oscillations usually related to an ionization instability. These effects correlate strongly with the reduction of the voltage drop in the region with the fringing magnetic field between the thruster channel and the cathode. The measured changes of the plasma properties suggest that the electron emission affects the electron cross-field transport in the thruster discharge. These trends are generalized for Hall thrusters of various configurations.

Integration Tests of the 4 kW-class High Voltage Hall Accelerator Power Processing Unit with the HiVHAc and the SPT-140 Hall Effect Thrusters

Science.gov (United States)

Kamhawi, Hani; Pinero, Luis; Haag, Thomas; Huang, Wensheng; Ahern, Drew; Liang, Ray; Shilo, Vlad

2016-01-01

NASAs Science Mission Directorate is sponsoring the development of a 4 kW-class Hall propulsion system for implementation in NASA science and exploration missions. The main components of the system include the High Voltage Hall Accelerator (HiVHAc), an engineering model power processing unit (PPU) developed by Colorado Power Electronics, and a xenon flow control module (XFCM) developed by VACCO Industries. NASA Glenn Research Center is performing integrated tests of the Hall thruster propulsion system. This presentation presents results from integrated tests of the PPU and XFCM with the HiVHAc engineering development thruster and a SPT-140 thruster provided by Space System Loral. The results presented in this paper demonstrate thruster discharge initiation, open-loop and closed-loop control of the discharge current with anode flow for both the HiVHAc and the SPT-140 thrusters. Integrated tests with the SPT-140 thruster indicated that the PPU was able to repeatedly initiate the thrusters discharge, achieve steady state operation, and successfully throttle the thruster between 1.5 and 4.5 kW. The measured SPT-140 performance was identical to levels reported by Space Systems Loral.

A data acquisition and storage system for the ion auxiliary propulsion system cyclic thruster test

Science.gov (United States)

Hamley, John A.

1989-01-01

A nine-track tape drive interfaced to a standard personal computer was used to transport data from a remote test site to the NASA Lewis mainframe computer for analysis. The Cyclic Ground Test of the Ion Auxiliary Propulsion System (IAPS), which successfully achieved its goal of 2557 cycles and 7057 hr of thrusting beam on time generated several megabytes of test data over many months of continuous testing. A flight-like controller and power supply were used to control the thruster and acquire data. Thruster data was converted to RS232 format and transmitted to a personal computer, which stored the raw digital data on the nine-track tape. The tape format was such that with minor modifications, mainframe flight data analysis software could be used to analyze the Cyclic Ground Test data. The personal computer also converted the digital data to engineering units and displayed real time thruster parameters. Hardcopy data was printed at a rate dependent on thruster operating conditions. The tape drive provided a convenient means to transport the data to the mainframe for analysis, and avoided a development effort for new data analysis software for the Cyclic test. This paper describes the data system, interfacing and software requirements.

Design and Testing of a Hall Effect Thruster with Additively Manufactured Components

Science.gov (United States)

Hopping, Ethan

The UAH-78AM is a low-power Hall effect thruster developed at the University of Alabama in Huntsville to study the application of low-cost additive manufacturing in the design and fabrication of Hall thrusters. The goal of this project is to assess the feasibility of using unconventional materials to produce a low-cost functioning Hall effect thruster and consider how additive manufacturing can expand the design space and provide other benefits. The thruster features channel walls and a propellant distributor that were manufactured using 3D printing with a variety of materials including ABS, ULTEM, and glazed ceramic. A version of the thruster was tested at NASA Glenn Research Center to obtain performance metrics and to validate the ability of the thruster to produce thrust and sustain a discharge. The design of the thruster and the transient performance measurements are presented here. Measured thrust ranged from 17.2 mN to 30.4 mN over a discharge power of 280 W to 520 W with an anode Isp range of 870 s to 1450 s. Temperature limitations of materials used for the channel walls and propellant distributor limit the ability to run the thruster at thermal steady-state. While the current thruster design is not yet ready for continuous operation, revisions to the device that could enable longer duration tests are discussed.

High Power MPD Thruster Development at the NASA Glenn Research Center

Science.gov (United States)

LaPointe, Michael R.; Mikellides, Pavlos G.; Reddy, Dhanireddy (Technical Monitor)

2001-01-01

Propulsion requirements for large platform orbit raising, cargo and piloted planetary missions, and robotic deep space exploration have rekindled interest in the development and deployment of high power electromagnetic thrusters. Magnetoplasmadynamic (MPD) thrusters can effectively process megawatts of power over a broad range of specific impulse values to meet these diverse in-space propulsion requirements. As NASA's lead center for electric propulsion, the Glenn Research Center has established an MW-class pulsed thruster test facility and is refurbishing a high-power steady-state facility to design, build, and test efficient gas-fed MPD thrusters. A complimentary numerical modeling effort based on the robust MACH2 code provides a well-balanced program of numerical analysis and experimental validation leading to improved high power MPD thruster performance. This paper reviews the current and planned experimental facilities and numerical modeling capabilities at the Glenn Research Center and outlines program plans for the development of new, efficient high power MPD thrusters.

Development of an Ion Thruster and Power Processor for New Millennium's Deep Space 1 Mission

Science.gov (United States)

Sovey, James S.; Hamley, John A.; Haag, Thomas W.; Patterson, Michael J.; Pencil, Eric J.; Peterson, Todd T.; Pinero, Luis R.; Power, John L.; Rawlin, Vincent K.; Sarmiento, Charles J.;

Electron Cross-field Transport in a Miniaturized Cylindrical Hall Thruster

International Nuclear Information System (INIS)

Smirnov Artem; Raitses Yevgeny; Fisch Nathaniel J

2005-01-01

Conventional annular Hall thrusters become inefficient when scaled to low power. Cylindrical Hall thrusters, which have lower surface-to-volume ratio, are more promising for scaling down. They presently exhibit performance comparable with conventional annular Hall thrusters. The present paper gives a review of the experimental and numerical investigations of electron crossfield transport in the 2.6 cm miniaturized cylindrical Hall thruster (100 W power level). We show that, in order to explain the discharge current observed for the typical operating conditions, the electron anomalous collision frequency ν b has to be on the order of the Bohm value, ν B ∼ ω c /16. The contribution of electron-wall collisions to cross-field transport is found to be insignificant. The optimal regimes of thruster operation at low background pressure (below 10 -5 Torr) in the vacuum tank appear to be different from those at higher pressure (∼ 10 -4 Torr)

Plasma Characterization of Hall Thruster with Active and Passive Segmented Electrodes

International Nuclear Information System (INIS)

Raitses, Y.; Staack, D.; Fisch, N.J.

2002-01-01

Non-emissive electrodes and ceramic spacers placed along the Hall thruster channel are shown to affect the plasma potential distribution and the thruster operation. These effects are associated with physical properties of the electrode material and depend on the electrode configuration, geometry and the magnetic field distribution. An emissive segmented electrode was able to maintain thruster operation by supplying an additional electron flux to sustain the plasma discharge between the anode and cathode neutralizer. These results indicate the possibility of new configurations for segmented electrode Hall thruster

The FAST (FRC Acceleration Space Thruster) Experiment

Science.gov (United States)

Martin, Adam; Eskridge, R.; Lee, M.; Richeson, J.; Smith, J.; Thio, Y. C. F.; Slough, J.; Rodgers, Stephen L. (Technical Monitor)

2001-01-01

The Field Reverse Configuration (FRC) is a magnetized plasmoid that has been developed for use in magnetic confinement fusion. Several of its properties suggest that it may also be useful as a thruster for in-space propulsion. The FRC is a compact toroid that has only poloidal field, and is characterized by a high plasma beta = (P)/(B (sup 2) /2Mu0), the ratio of plasma pressure to magnetic field pressure, so that it makes efficient use of magnetic field to confine a plasma. In an FRC thruster, plasmoids would be repetitively formed and accelerated to high velocity; velocities of = 250 km/s (Isp = 25,000s) have already been achieved in fusion experiments. The FRC is inductively formed and accelerated, and so is not subject to the problem of electrode erosion. As the plasmoid may be accelerated over an extended length, it can in principle be made very efficient. And the achievable jet powers should be scalable to the MW range. A 10 kW thruster experiment - FAST (FRC Acceleration Space Thruster) has just started at the Marshall Space Flight Center. The design of FAST and the status of construction and operation will be presented.

Experimental and theoretical studies of cylindrical Hall thrusters

International Nuclear Information System (INIS)

Smirnov, Artem; Raitses, Yegeny; Fisch, Nathaniel J.

2007-01-01

The Hall thruster is a mature electric propulsion device that holds considerable promise in terms of the propellant saving potential. The annular design of the conventional Hall thruster, however, does not naturally scale to low power. The efficiency tends to be lower and the lifetime issues are more aggravated. Cylindrical geometry Hall thrusters have lower surface-to-volume ratio than conventional thrusters and, thus, seem to be more promising for scaling down. The cylindrical Hall thruster (CHT) is fundamentally different from the conventional design in the way the electrons are confined and the ion space charge is neutralized. The performances of both the large (9-cm channel diameter, 600-1000 W) and miniaturized (2.6-cm channel diameter, 50-300 W) CHTs are comparable with those of the state-of-the-art conventional (annular) design Hall thrusters of similar sizes. A comprehensive experimental and theoretical study of the CHT physics has been conducted, addressing the questions of electron cross-field transport, propellant ionization, plasma-wall interaction, and formation of the electron distribution function. Probe measurements in the harsh plasma environment of the microthruster were performed. Several interesting effects, such as the unusually high ionization efficiency and enhanced electron transport, were observed. Kinetic simulations suggest the existence of the strong fluctuation-enhanced electron diffusion and predict the non-Maxwellian shape of the electron distribution function. Through the acquired understanding of the new physics, ways for further optimization of this means for low-power space propulsion are suggested. Substantial flexibility in the magnetic field configuration of the CHT is the key tool in achieving the high-efficiency operation

Particle-in-cell simulations of Hall plasma thrusters

Science.gov (United States)

Miranda, Rodrigo; Ferreira, Jose Leonardo; Martins, Alexandre

2016-07-01

Hall plasma thrusters can be modelled using particle-in-cell (PIC) simulations. In these simulations, the plasma is described by a set of equations which represent a coupled system of charged particles and electromagnetic fields. The fields are computed using a spatial grid (i.e., a discretization in space), whereas the particles can move continuously in space. Briefly, the particle and fields dynamics are computed as follows. First, forces due to electric and magnetic fields are employed to calculate the velocities and positions of particles. Next, the velocities and positions of particles are used to compute the charge and current densities at discrete positions in space. Finally, these densities are used to solve the electromagnetic field equations in the grid, which are interpolated at the position of the particles to obtain the acting forces, and restart this cycle. We will present numerical simulations using software for PIC simulations to study turbulence, wave and instabilities that arise in Hall plasma thrusters. We have sucessfully reproduced a numerical simulation of a SPT-100 Hall thruster using a two-dimensional (2D) model. In addition, we are developing a 2D model of a cylindrical Hall thruster. The results of these simulations will contribute to improve the performance of plasma thrusters to be used in Cubesats satellites currenty in development at the Plasma Laboratory at University of Brasília.

Mission and System Advantages of Iodine Hall Thrusters

Science.gov (United States)

Dankanich, John W.; Szabo, James; Pote, Bruce; Oleson, Steve; Kamhawi, Hani

2014-01-01

The exploration of alternative propellants for Hall thrusters continues to be of interest to the community. Investments have been made and continue for the maturation of iodine based Hall thrusters. Iodine testing has shown comparable performance to xenon. However, iodine has a higher storage density and resulting higher ?V capability for volume constrained systems. Iodine's vapor pressure is low enough to permit low-pressure storage, but high enough to minimize potential adverse spacecraft-thruster interactions. The low vapor pressure also means that iodine does not condense inside the thruster at ordinary operating temperatures. Iodine is safe, it stores at sub-atmospheric pressure, and can be stored unregulated for years on end; whether on the ground or on orbit. Iodine fills a niche for both low power (10kW) electric propulsion regimes. A range of missions have been evaluated for direct comparison of Iodine and Xenon options. The results show advantages of iodine Hall systems for both small and microsatellite application and for very large exploration class missions.

Optimization of a coaxial electron cyclotron resonance plasma thruster with an analytical model

Energy Technology Data Exchange (ETDEWEB)

Cannat, F., E-mail: felix.cannat@onera.fr, E-mail: felix.cannat@gmail.com; Lafleur, T. [Physics and Instrumentation Department, Onera -The French Aerospace Lab, Palaiseau, Cedex 91123 (France); Laboratoire de Physique des Plasmas, CNRS, Sorbonne Universites, UPMC Univ Paris 06, Univ Paris-Sud, Ecole Polytechnique, 91128 Palaiseau (France); Jarrige, J.; Elias, P.-Q.; Packan, D. [Physics and Instrumentation Department, Onera -The French Aerospace Lab, Palaiseau, Cedex 91123 (France); Chabert, P. [Laboratoire de Physique des Plasmas, CNRS, Sorbonne Universites, UPMC Univ Paris 06, Univ Paris-Sud, Ecole Polytechnique, 91128 Palaiseau (France)

2015-05-15

A new cathodeless plasma thruster currently under development at Onera is presented and characterized experimentally and analytically. The coaxial thruster consists of a microwave antenna immersed in a magnetic field, which allows electron heating via cyclotron resonance. The magnetic field diverges at the thruster exit and forms a nozzle that accelerates the quasi-neutral plasma to generate a thrust. Different thruster configurations are tested, and in particular, the influence of the source diameter on the thruster performance is investigated. At microwave powers of about 30 W and a xenon flow rate of 0.1 mg/s (1 SCCM), a mass utilization of 60% and a thrust of 1 mN are estimated based on angular electrostatic probe measurements performed downstream of the thruster in the exhaust plume. Results are found to be in fair agreement with a recent analytical helicon thruster model that has been adapted for the coaxial geometry used here.

Hardware in the Loop Testing of an Iodine-Fed Hall Thruster

Science.gov (United States)

Polzin, Kurt A.; Peeples, Steven R.; Cecil, Jim; Lewis, Brandon L.; Molina Fraticelli, Jose C.; Clark, James P.

2015-01-01

chamber (it is under 10(exp -6) torr at -75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. In the present paper, we describe testing performed using an iodine-fed 200 W Hall thruster mounted to a thrust stand and operated in conjunction with MSFCs Small Projects Rapid Integration and Test Environment (SPRITE) Portable Hardware In the Loop (PHIL) hardware. This work is performed in support of the iodine satellite (iSAT) project, which aims to fly a 200-W iodine-fed thruster on a 12-U CubeSat. The SPRITE PHIL hardware allows a given vehicle to do a checkout of its avionics algorithm by allowing it to monitor and feed data to simulated sensors and effectors in a digital environment. These data are then used to determine the attitude of the vehicle and a separate computer is used to interpret the data set and visualize it using a 3D graphical interface. The PHIL hardware allows the testing of the vehicles bus by providing 'real' hardware interfaces (in the case of this test a real RS422 bus) and specific components can be modeled to show their interactions with the avionics algorithm (e.g. a thruster model). For the iSAT project the PHIL is used to visualize the operating cycle of the thruster and the subsequent effect this thrusting has on the attitude of the satellite over a given period of time. The test is controlled using software running on an Andrews Space Cortex 160 flight computer. This computer is the current baseline for a full iSAT mission. While the test could be conducted with a lab computer and software, the team chose to exercise the propulsion system with a representative CubeSat-class computer. For purposes of this test, the "flight" software monitored the propulsion and PPU systems, controlled operation of the thruster, and provided thruster state data to the PHIL simulation. Commands to operate the thruster were

Numerical simulation of SMART-1 Hall-thruster plasma interactions

NARCIS (Netherlands)

Tajmar, Martin; Sedmik, René; Scharlemann, Carsten

2009-01-01

SMART-1 has been the first European mission using a Hall thruster to reach the moon. An onboard plasma diagnostic package allowed a detailed characterization of the thruster exhaust plasma and its interactions with the spacecraft. Analysis of in-flight data revealed, amongst others, an unpredicted

Performance prediction of electrohydrodynamic thrusters by the perturbation method

International Nuclear Information System (INIS)

Shibata, H.; Watanabe, Y.; Suzuki, K.

2016-01-01

In this paper, we present a novel method for analyzing electrohydrodynamic (EHD) thrusters. The method is based on a perturbation technique applied to a set of drift-diffusion equations, similar to the one introduced in our previous study on estimating breakdown voltage. The thrust-to-current ratio is generalized to represent the performance of EHD thrusters. We have compared the thrust-to-current ratio obtained theoretically with that obtained from the proposed method under atmospheric air conditions, and we have obtained good quantitative agreement. Also, we have conducted a numerical simulation in more complex thruster geometries, such as the dual-stage thruster developed by Masuyama and Barrett [Proc. R. Soc. A 469, 20120623 (2013)]. We quantitatively clarify the fact that if the magnitude of a third electrode voltage is low, the effective gap distance shortens, whereas if the magnitude of the third electrode voltage is sufficiently high, the effective gap distance lengthens.

Global Controllability of Chemical Reactions

OpenAIRE

Drexler, Dániel András; Tóth, János

2015-01-01

Controllability of chemical reactions is an important problem in chemical engineering science. In control theory, analysis of the controllability of linear systems is well-founded, however the dynamics of chemical reactions is usually nonlinear. Global controllability properties of chemical reactions are analyzed here based on the Lie-algebra of the vector fields associated to elementary reactions. A chemical reaction is controllable almost everywhere if all the reaction rate coefficients can...

Experimental Studies of Anode Sheath Phenomena in a Hall Thruster Discharge

International Nuclear Information System (INIS)

Dorf, L.; Raitses, Y.; Fisch, N.J.

2004-01-01

Both electron-repelling and electron-attracting anode sheaths in a Hall thruster were characterized by measuring the plasma potential with biased and emissive probes [L. Dorf, Y. Raitses, V. Semenov, and N.J. Fisch, Appl. Phys. Let. 84 (2004) 1070]. In the present work, two-dimensional structures of the plasma potential, electron temperature, and plasma density in the near-anode region of a Hall thruster with clean and dielectrically coated anodes are identified. Possible mechanisms of anode sheath formation in a Hall thruster are analyzed. The path for current closure to the anode appears to be the determining factor in the anode sheath formation process. The main conclusion of this work is that the anode sheath formation in Hall thrusters differs essentially from that in the other gas discharge devices, like a glow discharge or a hollow anode, because the Hall thruster utilizes long electron residence times to ionize rather than high neutral pressures

Operation of a Segmented Hall Thruster with Low-sputtering Carbon-velvet Electrodes

International Nuclear Information System (INIS)

Raitses, Y.; Staack, D.; Dunaevsky, A.; Fisch, N.J.

2005-01-01

Carbon fiber velvet material provides exceptional sputtering resistance properties exceeding those for graphite and carbon composite materials. A 2 kW Hall thruster with segmented electrodes made of this material was operated in the discharge voltage range of 200-700 V. The arcing between the floating velvet electrodes and the plasma was visually observed, especially, during the initial conditioning time, which lasted for about 1 h. The comparison of voltage versus current and plume characteristics of the Hall thruster with and without segmented electrodes indicates that the magnetic insulation of the segmented thruster improves with the discharge voltage at a fixed magnetic field. The observations reported here also extend the regimes wherein the segmented Hall thruster can have a narrower plume than that of the conventional nonsegmented thruster

Thermal stability of the krypton Hall effect thruster

Directory of Open Access Journals (Sweden)

Szelecka Agnieszka

2017-03-01

Full Text Available The Krypton Large IMpulse Thruster (KLIMT ESA/PECS project, which has been implemented in the Institute of Plasma Physics and Laser Microfusion (IPPLM and now is approaching its final phase, was aimed at incremental development of a ~500 W class Hall effect thruster (HET. Xenon, predominantly used as a propellant in the state-of-the-art HETs, is extremely expensive. Krypton has been considered as a cheaper alternative since more than fifteen years; however, to the best knowledge of the authors, there has not been a HET model especially designed for this noble gas. To address this issue, KLIMT has been geared towards operation primarily with krypton. During the project, three subsequent prototype versions of the thruster were designed, manufactured and tested, aimed at gradual improvement of each next exemplar. In the current paper, the heat loads in new engine have been discussed. It has been shown that thermal equilibrium of the thruster is gained within the safety limits of the materials used. Extensive testing with both gases was performed to compare KLIMT’s thermal behaviour when supplied with krypton and xenon propellants.

Parametric studies of the Hall Thruster at Soreq

International Nuclear Information System (INIS)

Ashkenazy, J.; Rattses, Y.; Appelbaum, G.

1997-01-01

An electric propulsion program was initiated at Soreq a few years ago, aiming at the research and development of advanced Hall thrusters for various space applications. The Hall thruster accelerates a plasma jet by an axial electric field and an applied radial magnetic field in an annular ceramic channel. A relatively large current density (> 0.1 A/cm 2 ) can be obtained, since the acceleration mechanism is not limited by space charge effects. Such a device can be used as a small rocket engine onboard spacecraft with the advantage of a large jet velocity compared with conventional rocket engines (10,000-30,000 m/s vs. 2,000-4,800 m/s). An experimental Hall thruster was constructed at Soreq and operated under a broad range of operating conditions and under various configurational variations. Electrical, magnetic and plasma diagnostics, as well as accurate thrust and gas flow rate measurements, have been used to investigate the dependence of thruster behavior on the applied voltage, gas flow rate, magnetic field, channel geometry and wall material. Representative results highlighting the major findings of the studies conducted so far are presented

Experimental Investigation of a Direct-drive Hall Thruster and Solar Array System at Power Levels up to 10 kW

Science.gov (United States)

Snyder, John S.; Brophy, John R.; Hofer, Richard R.; Goebel, Dan M.; Katz, Ira

2012-01-01

As NASA considers future exploration missions, high-power solar-electric propulsion (SEP) plays a prominent role in achieving many mission goals. Studies of high-power SEP systems (i.e. tens to hundreds of kilowatts) suggest that significant mass savings may be realized by implementing a direct-drive power system, so NASA recently established the National Direct-Drive Testbed to examine technical issues identified by previous investigations. The testbed includes a 12-kW solar array and power control station designed to power single and multiple Hall thrusters over a wide range of voltages and currents. In this paper, single Hall thruster operation directly from solar array output at discharge voltages of 200 to 450 V and discharge powers of 1 to 10 kW is reported. Hall thruster control and operation is shown to be simple and no different than for operation on conventional power supplies. Thruster and power system electrical oscillations were investigated over a large range of operating conditions and with different filter capacitances. Thruster oscillations were the same as for conventional power supplies, did not adversely affect solar array operation, and were independent of filter capacitance from 8 to 80 ?F. Solar array current and voltage oscillations were very small compared to their mean values and showed a modest dependence on capacitor size. No instabilities or anomalous behavior were observed in the thruster or power system at any operating condition investigated, including near and at the array peak power point. Thruster startup using the anode propellant flow as the power 'switch' was shown to be simple and reliable with system transients mitigated by the proper selection of filter capacitance size. Shutdown via cutoff of propellant flow was also demonstrated. A simple electrical circuit model was developed and is shown to have good agreement with the experimental data.

Chemical Characterization and Reactivity Testing of Fuel-Oxidizer Reaction Product (Test Report)

Science.gov (United States)

1996-01-01

The product of incomplete reaction of monomethylhydrazine (MMH) and nitrogen tetroxide (NTO) propellants, or fuel-oxidizer reaction product (FORP), has been hypothesized as a contributory cause of an anomaly which occurred in the chamber pressure (PC) transducer tube on the Reaction Control Subsystem (RCS) aft thruster 467 on flight STS-51. A small hole was found in the titanium-alloy PC tube at the first bend below the pressure transducer. It was surmised that the hole may have been caused by heat and pressure resulting from ignition of FORP. The NASA Johnson Space Center (JSC) White Sands Test Facility (WSTF) was requested to define the chemical characteristics of FORP, characterize its reactivity, and simulate the events in a controlled environment which may have lead to the Pc-tube failure. Samples of FORP were obtained from the gas-phase reaction of MMH with NTO under laboratory conditions, the pulsed firings of RCS thrusters with modified PC tubes using varied oxidizer or fuel lead times, and the nominal RCS thruster firings at WSTF and Kaiser-Marquardt. Fourier transform infrared spectroscopy (FTIR), differential scanning calorimetry (DSC), accelerating rate calorimetry (ARC), ion chromatography (IC), inductively coupled plasma (ICP) spectrometry, thermogravimetric analysis (TGA) coupled to FTIR (TGA/FTIR), and mechanical impact testing were used to qualitatively and quantitatively characterize the chemical, thermal, and ignition properties of FORP. These studies showed that the composition of FORP is variable but falls within a limited range of compositions that depends on the fuel loxidizer ratio at the time of formation, composition of the post-formation atmosphere (reducing or oxidizing), and reaction or postreaction temperature. A typical composition contains methylhydrazinium nitrate (MMHN), ammonium nitrate (AN), methylammonium nitrate (MAN), and trace amounts of hydrazinium nitrate and 1,1-dimethylhydrazinium nitrate. The thermal decomposition

Orbital Dynamics of a Simple Solar Photon Thruster

OpenAIRE

Guerman, Anna D.; Smirnov, Georgi V.; Pereira, Maria Cecilia

2009-01-01

We study orbital dynamics of a compound solar sail, namely, a Simple Solar Photon Thruster and compare its behavior to that of a common version of sailcraft. To perform this analysis, development of a mathematical model for force created by light reflection on all sailcraft elements is essential. We deduce the equations of sailcraft's motion and compare performance of two schemes of solar propulsion for two test time-optimal control problems of trajectory transfer.

Bi-Modal Micro-Cathode Arc Thruster for Cube Satellites

Science.gov (United States)

Chiu, Dereck

A new concept design, named the Bi-Modal Micro-Cathode Arc Thruster (BM-muCAT), has been introduced utilizing features from previous generations of muCATs and incorporating a multi-propellant functionality. This arc thruster is a micro-Newton level thruster based off of vacuum arc technology utilizing an enhanced magnetic field. Adjusting the magnetic field allows the thrusters performance to be varied. The goal of this thesis is to present a new generation of micro-cathode arc thrusters utilizing a bi-propellant, nickel and titanium, system. Three experimental procedures were run to test the new designs capabilities. Arc rotation experiment was used as a base experiment to ensure erosion was occurring uniformly along each electrode. Ion utilization efficiency was found, using an ion collector, to be up to 2% with the nickel material and 2.5% with the titanium material. Ion velocities were also studied using a time-of-flight method with an enhanced ion detection system. This system utilizes double electrostatic probes to measure plasma propagation. Ion velocities were measured to be 10km/s and 20km/s for nickel and titanium without a magnetic field. With an applied magnetic field of 0.2T, nickel ion velocities almost doubled to about 17km/s, while titanium ion velocities also increased to about 30km/s.

Geometrical characterization and performance optimization of monopropellant thruster injector

Directory of Open Access Journals (Sweden)

T.R. Nada

2012-12-01

Full Text Available The function of the injector in a monopropellant thruster is to atomize the liquid hydrazine and to distribute it over the catalyst bed as uniformly as possible. A second objective is to place the maximum amount of catalyst in contact with the propellant in as short time as possible to minimize the starting transient time. Coverage by the spray is controlled mainly by cone angle and diameter of the catalyst bed, while atomization quality is measured by the Sauter Mean Diameter, SMD. These parameters are evaluated using empirical formulae. In this paper, two main types of injectors are investigated; plain orifice and full cone pressure swirl injectors. The performance of these two types is examined for use with blow down monopropellant propulsion system. A comprehensive characterization is given and design charts are introduced to facilitate optimizing the performance of the injector. Full-cone injector is a more suitable choice for monopropellant thruster and it might be available commercially.

Oxygen-Methane Thruster, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — Two main innovations will be developed in the Phase II effort that are fundamentally associated with our gaseous oxygen/gaseous methane RCS thruster. The first...

Performance, Facility Pressure Effects, and Stability Characterization Tests of NASA's Hall Effect Rocket with Magnetic Shielding Thruster

Science.gov (United States)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Yim, John; Herman, Daniel; Williams, George; Gilland, James; Peterson, Peter; Hofer, Richard; Mikellides, Ioannis

2016-01-01

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for flight system development. Part of the technology maturation effort included experimental evaluation of the TDU-1 thruster with conducting and dielectric front pole cover materials in two different electrical configurations. A graphite front pole cover thruster configuration with the thruster body electrically tied to cathode and an alumina front pole cover thruster configuration with the thruster body floating were evaluated. Both configurations were also evaluated at different facility background pressure conditions to evaluate background pressure effects on thruster operation. Performance characterization tests found that higher thruster performance was attained with the graphite front pole cover configuration with the thruster electrically tied to cathode. A total thrust efficiency of 68 and a total specific impulse of 2,820 s was demonstrated at a discharge voltage of 600 V and a discharge power of 12.5 kW. Thruster stability regimes were characterized with respect to the thruster discharge current oscillations and with maps of the current-voltage-magnetic field (IVB). Analysis of TDU-1 discharge current waveforms found that lower normalized discharge current peak-to-peak and root mean square magnitudes were attained when the thruster was electrically floated with alumina front pole covers. Background pressure effects characterization tests indicated that the thruster performance and stability was mostly invariant to changes in the facility background pressure for vacuum chamber pressure below 110-5 Torr-Xe (for thruster flow rate above 8 mgs). Power spectral density analysis of the discharge current waveform showed that increasing the vacuum chamber background pressure resulted in a higher discharge current dominant frequency. Finally the IVB maps of the TDU-1

Low-Cost, High-Performance Hall Thruster Support System

Science.gov (United States)

Hesterman, Bryce

2015-01-01

Colorado Power Electronics (CPE) has built an innovative modular PPU for Hall thrusters, including discharge, magnet, heater and keeper supplies, and an interface module. This high-performance PPU offers resonant circuit topologies, magnetics design, modularity, and a stable and sustained operation during severe Hall effect thruster current oscillations. Laboratory testing has demonstrated discharge module efficiency of 96 percent, which is considerably higher than current state of the art.

Thermal Environmental Testing of NSTAR Engineering Model Ion Thrusters

Science.gov (United States)

Rawlin, Vincent K.; Patterson, Michael J.; Becker, Raymond A.

1999-01-01

NASA's New Millenium program will fly a xenon ion propulsion system on the Deep Space 1 Mission. Tests were conducted under NASA's Solar Electric Propulsion Technology Applications Readiness (NSTAR) Program with 3 different engineering model ion thrusters to determine thruster thermal characteristics over the NSTAR operating range in a variety of thermal environments. A liquid nitrogen-cooled shroud was used to cold-soak the thruster to -120 C. Initial tests were performed prior to a mature spacecraft design. Those results and the final, severe, requirements mandated by the spacecraft led to several changes to the basic thermal design. These changes were incorporated into a final design and tested over a wide range of environmental conditions.

Retrofit and acceptance test of 30-cm ion thrusters

Science.gov (United States)

Poeschel, R. L.

1981-01-01

Six 30 cm mercury thrusters were modified to the J-series design and evaluated using standardized test procedures. The thruster performance meets the design objectives (lifetime objective requires verification), and documentation (drawings, etc.) for the design is completed and upgraded. The retrofit modifications are described and the test data for the modifications are presented and discussed.

Density and velocity measurements of a sheath plasma from MPD thruster

Energy Technology Data Exchange (ETDEWEB)

Ko, J.J.; Cho, T.S.; Choi, M.C.; Choi, E.H.; Cho, G.S.; Uhm, H.S.

1999-07-01

Magnetoplasma is the plasma that the electron and ion orbits are strongly confined by intense magnetic field. Recently, magnetoplasma dynamics (MPD) has been investigated in connection with applications to the rocket thruster in USA, Germany, etc. It can be widely applicable, including modification of satellite position and propulsion of the interplanetary space shuttle. A travel for a long distance journey is possible because a little amount of neutral gases is needed for the plasma source. Besides, this will provide a pollution free engine for future generations. MPD thruster is not a chemical engine. The authors have built a Mather type MPD thruster, which has 1 kV max charging, 10 kA max current flows, and has about 1 ms characteristic operation time. The Paschen curve of this thruster is measured and its minimum breakdown voltage occurs in the pressure range of 0.1 to 1 Torr. Langmuir and double probes are fabricated to diagnose the sheath plasma from the thruster. The temperature and density are calculated to be 2.5 eV and 10{sup 15} cm {sup {minus}3}, respectively, from the probe data. Making use of photo diode, an optical probe is fabricated to measure propagation velocity of the sheath plasma. The sheath plasma from the MPD thruster in the experiment propagates with velocity of 1 cm/{micro}s.

Optical Diagnostic Characterization of High-Power Hall Thruster Wear and Operation

Science.gov (United States)

Williams, George J., Jr.; Soulas, George C.; Kamhawi, Hani

2012-01-01

Optical emission spectroscopy is employed to correlate BN insulator erosion with high-power Hall thruster operation. Specifically, actinometry leveraging excited xenon states is used to normalize the emission spectra of ground state boron as a function of thruster operating condition. Trends in the strength of the boron signal are correlated with thruster power, discharge voltage, and discharge current. In addition, the technique is demonstrated on metallic coupons embedded in the walls of the HiVHAc EM thruster. The OES technique captured the overall trend in the erosion of the coupons which boosts credibility in the method since there are no data to which to calibrate the erosion rates of high-power Hall thrusters. The boron signals are shown to trend linearly with discharge voltage for a fixed discharge current as expected. However, the boron signals of the higher-power NASA 300M and NASA 457Mv2 trend with discharge current and show an unexpectedly weak to inverse dependence on discharge voltage. Electron temperatures measured optically in the near-field plume of the thruster agree well with Langmuir probe data. However, the optical technique used to determine Te showed unacceptable sensitivity to the emission intensities. Near-field, single-frequency imaging of the xenon neutrals is also presented as a function of operating condition for the NASA 457 Mv2.

Fuel Oxidizer Reaction Products (FORP) Contamination of Service Module (SM) and Release of N-nitrosodimethylamine(NDMA)in a Humid Environment from Crew EVA Suits Contaminated with FORP

Science.gov (United States)

Schmidl, William; Mikatarian, Ron; Lam, Chiu-Wing; West, Bil; Buchanan, Vanessa; Dee, Louis; Baker, David; Koontz, Steve

2004-01-01

The Service Module (SM) is an element of the Russian Segment of the International Space Station (ISS). One of the functions of the SM is to provide attitude control for the ISS using thrusters when the U.S. Control Moment Gyros (CMG's) must be desaturated. Prior to an Extravehicular Activity (EVA) on the Russian Segment, the Docking Compartment (DC1) is depressurized, as it is used as an airlock. When the DC1 is depressurized, the CMG's margin of momentum is insufficient and the SM attitude control thrusters need to fire to desaturate the CMG's. SM roll thruster firings induce contamination onto adjacent surfaces with Fuel Oxidizer Reaction Products (FORP). FORP is composed of both volatile and non-volatile components. One of the components of FORP is the potent carcinogen N-nitrosdimethylamine (NDMA). Since the EVA crewmembers often enter the area surrounding the thrusters for tasks on the aft end of the SM and when translating to other areas of the Russian Segment, the presence of FORP is a concern. This paper will discuss FORP contamination of the SM surfaces, the release of NDMA in a humid environment from crew EVA suits, if they happen to be contaminated with FORP, and the toxicological risk associated with the NDMA release.

Electric arc discharge damage to ion thruster grids

Science.gov (United States)

Beebe, D. D.; Nakanishi, S.; Finke, R. C.

1974-01-01

Arcs representative of those occurring between the grids of a mercury ion thruster were simulated. Parameters affecting an arc and the resulting damage were studied. The parameters investigated were arc energy, arc duration, and grid geometry. Arc attenuation techniques were also investigated. Potentially serious damage occurred at all energy levels representative of actual thruster operating conditions. Of the grids tested, the lowest open-area configuration sustained the least damage for given conditions. At a fixed energy level a long duration discharge caused greater damage than a short discharge. Attenuation of arc current using various impedances proved to be effective in reducing arc damage. Faults were also deliberately caused using chips of sputtered materials formed during the operation of an actual thruster. These faults were cleared with no serious grid damage resulting using the principles and methods developed in this study.

Orbital Dynamics of a Simple Solar Photon Thruster

Directory of Open Access Journals (Sweden)

Anna D. Guerman

2009-01-01

Full Text Available We study orbital dynamics of a compound solar sail, namely, a Simple Solar Photon Thruster and compare its behavior to that of a common version of sailcraft. To perform this analysis, development of a mathematical model for force created by light reflection on all sailcraft elements is essential. We deduce the equations of sailcraft's motion and compare performance of two schemes of solar propulsion for two test time-optimal control problems of trajectory transfer.

Performance Test Results of the NASA-457M v2 Hall Thruster

Science.gov (United States)

Soulas, George C.; Haag, Thomas W.; Herman, Daniel A.; Huang, Wensheng; Kamhawi, Hani; Shastry, Rohit

2012-01-01

Performance testing of a second generation, 50 kW-class Hall thruster labeled NASA-457M v2 was conducted at the NASA Glenn Research Center. This NASA-designed thruster is an excellent candidate for a solar electric propulsion system that supports human exploration missions. Thruster discharge power was varied from 5 to 50 kW over discharge voltage and current ranges of 200 to 500 V and 15 to 100 A, respectively. Anode efficiencies varied from 0.56 to 0.71. The peak efficiency was similar to that of other state-of-the-art high power Hall thrusters, but outperformed these thrusters at lower discharge voltages. The 0.05 to 0.18 higher anode efficiencies of this thruster compared to its predecessor were primarily due to which of two stable discharge modes the thruster was operated. One stable mode was at low magnetic field strengths, which produced high anode efficiencies, and the other at high magnetic fields where its predecessor was operated. Cathode keeper voltages were always within 2.1 to 6.2 V and cathode voltages were within 13 V of tank ground during high anode efficiency operation. However, during operation at high magnetic fields, cathode-to-ground voltage magnitudes increased dramatically, exceeding 30 V, due to the high axial magnetic field strengths in the immediate vicinity of the centrally-mounted cathode. The peak thrust was 2.3 N and this occurred at a total thruster input power of 50.0 kW at a 500 V discharge voltage. The thruster demonstrated a thrust-to-power range of 76.4 mN/kW at low power to 46.1 mN/kW at full power, and a specific impulse range of 1420 to 2740 s. For a discharge voltage of 300 V, where specific impulses would be about 2000 s, thrust efficiencies varied from 0.57 to 0.63.

Pocket rocket: An electrothermal plasma micro-thruster

Science.gov (United States)

Greig, Amelia Diane

Recently, an increase in use of micro-satellites constructed from commercial off the shelf (COTS) components has developed, to address the large costs associated with designing, testing and launching satellites. One particular type of micro-satellite of interest are CubeSats, which are modular 10 cm cubic satellites with total weight less than 1.33 kg. To assist with orbit boosting and attitude control of CubeSats, micro-propulsion systems are required, but are currently limited. A potential electrothermal plasma micro-thruster for use with CubeSats or other micro-satellites is under development at The Australian National University and forms the basis for this work. The thruster, known as ‘Pocket Rocket’, utilises neutral gas heating from ion-neutral collisions within a weakly ionised asymmetric plasma discharge, increasing the exhaust thermal velocity of the propellant gas, thereby producing higher thrust than if the propellant was emitted cold. In this work, neutral gas temperature of the Pocket Rocket discharge is studied in depth using rovibrational spectroscopy of the nitrogen (N2) second positive system (C3Πu → B3Πg), using both pure N2 and argon/N2 mixtures as the operating gas. Volume averaged steady state gas temperatures are measured for a range of operating conditions, with an analytical collisional model developed to verify experimental results. Results show that neutral gas heating is occurring with volume averaged steady state temperatures reaching 430 K in N2 and 1060 K for argon with 1% N2 at standard operating conditions of 1.5 Torr pressure and 10 W power input, demonstrating proof of concept for the Pocket Rocket thruster. Spatiotemporal profiles of gas temperature identify that the dominant heating mechanisms are ion-neutral collisions within the discharge and wall heating from ion bombardment of the thruster walls. To complement the experimental results, computational fluid dynamics (CFD) simulations using the commercial CFD

Laser injection of ultra-short electron bursts for the diagnosis of Hall thruster plasma

International Nuclear Information System (INIS)

Albarede, L; Gibert, T; Lazurenko, A; Bouchoule, A

2006-01-01

The present developments of Hall thrusters for satellite control and space mission technologies represent a new step towards their routine use in place of conventional thermal thrusters. In spite of their long R and D history, the complex physics of the E x B discharge at work in these structures has prevented, up to now, the availability of predictive simulations. The electron transport in the accelerating layers of these thrusters is one of the remaining challenges in this direction. From the experimental point of view, any diagnostics of electron transport and electric field in this critical layer would be welcome for comparison with code predictions. Appropriate diagnostics are difficult, due to the very aggressive local plasma conditions. This paper presents the first step in the development of a new tool for characterization of the plasma electric field in the very near exhaust thruster plume and comparison with simulation code predictions. The main idea is to use very short bursts of electrons, probing local electron dynamics in this critical plume area. Such bursts can be obtained through photoelectric emission induced by a UV pulsed laser beam on a convenient target. A specific study, devoted to the characterization of the electron burst emission, is presented in the first section of the paper; the implementation and testing of the injection of electrons in the critical layer of Hall thruster plasma is described in the second section. The design and testing of a fast and sensitive system for characterizing the transport of injected bursts will be the next step of this program. It requires a preliminary evaluation of electron trajectories which was achieved by using simulation code. Simulation data are presented in the last section of the paper, with the full diagnostic design to be tested in the near future, when runs will be available in the renewed PIVOINE facility. The same electron burst injection could also be a valuable input in the present

Ion ejection from a permanent-magnet mini-helicon thruster

Energy Technology Data Exchange (ETDEWEB)

Chen, Francis F. [Electrical Engineering Department, University of California, Los Angeles 90095-1594 (United States)

2014-09-15

A small helicon source, 5 cm in diameter and 5 cm long, using a permanent magnet (PM) to create the DC magnetic field B, is investigated for its possible use as an ion spacecraft thruster. Such ambipolar thrusters do not require a separate electron source for neutralization. The discharge is placed in the far-field of the annular PM, where B is fairly uniform. The plasma is ejected into a large chamber, where the ion energy distribution is measured with a retarding-field energy analyzer. The resulting specific impulse is lower than that of Hall thrusters but can easily be increased to relevant values by applying to the endplate of the discharge a small voltage relative to spacecraft ground.

An evaluation of krypton propellant in Hall thrusters

Science.gov (United States)

Linnell, Jesse Allen

Due to its high specific impulse and low price, krypton has long sparked interest as an alternate Hall thruster propellant. Unfortunately at the moment, krypton's relatively poor performance precludes it as a legitimate option. This thesis presents a detailed investigation into krypton operation in Hall thrusters. These findings suggest that the performance gap can be decreased to 4% and krypton can finally become a realistic propellant option. Although krypton has demonstrated superior specific impulse, the xenon-krypton absolute efficiency gap ranges between 2 and 15%. A phenomenological performance model indicates that the main contributors to the efficiency gap are propellant utilization and beam divergence. Propellant utilization and beam divergence have relative efficiency deficits of 5 and 8%, respectively. A detailed characterization of internal phenomena is conducted to better understand the xenon-krypton efficiency gap. Krypton's large beam divergence is found to be related to a defocusing equipotential structure and a weaker magnetic field topology. Ionization processes are shown to be linked to the Hall current, the magnetic mirror topology, and the perpendicular gradient of the magnetic field. Several thruster design and operational suggestions are made to optimize krypton efficiency. Krypton performance is optimized for discharge voltages above 500 V and flow rates corresponding to an a greater than 0.015 mg/(mm-s), where alpha is a function of flow rate and discharge channel dimensions (alpha = m˙alphab/Ach). Performance can be further improved by increasing channel length or decreasing channel width for a given flow rate. Also, several magnetic field design suggestions are made to enhance ionization and beam focusing. Several findings are presented that improve the understanding of general Hall thruster physics. Excellent agreement is shown between equipotential lines and magnetic field lines. The trim coil is shown to enhance beam focusing

Combined tunable diode laser absorption spectroscopy and monochromatic radiation thermometry in ammonium dinitramide-based thruster

Science.gov (United States)

Zeng, Hui; Ou, Dongbin; Chen, Lianzhong; Li, Fei; Yu, Xilong

2018-02-01

Nonintrusive temperature measurements for a real ammonium dinitramide (ADN)-based thruster by using tunable diode laser absorption spectroscopy and monochromatic radiation thermometry are proposed. The ADN-based thruster represents a promising future space propulsion employing green, nontoxic propellant. Temperature measurements in the chamber enable quantitative thermal analysis for the thruster, providing access to evaluate thermal properties of the thruster and optimize thruster design. A laser-based sensor measures temperature of combustion gas in the chamber, while a monochromatic thermometry system based on thermal radiation is utilized to monitor inner wall temperature in the chamber. Additional temperature measurements of the outer wall temperature are conducted on the injector, catalyst bed, and combustion chamber of the thruster by using thermocouple, respectively. An experimental ADN thruster is redesigned with optimizing catalyst bed length of 14 mm and steady-state firing tests are conducted under various feed pressures over the range from 5 to 12 bar at a typical ignition temperature of 200°C. A threshold of feed pressure higher than 8 bar is required for the thruster's normal operation and upstream movement of the heat release zone is revealed in the combustion chamber out of temperature evolution in the chamber.

Improvement of Flow Characteristics for an Advanced Plasma Thruster

International Nuclear Information System (INIS)

Inutake, M.; Hosokawa, Y.; Sato, R.; Ando, A.; Tobari, H.; Hattori, K.

2005-01-01

A higher specific impulse and a larger thrust are required for a manned interplanetary space thruster. Until the realization of a fusion-plasma thruster, a magneto-plasma-dynamic arcjet (MPDA) powered by a fission reactor is one of the promising candidates for a manned Mars space thruster. The MPDA plasma is accelerated axially by a self-induced j x B force. Thrust performance of the MPDA is expected to increase by applying a magnetic nozzle instead of a solid nozzle. In order to get a much higher thruster performance, two methods have been investigated in the HITOP device, Tohoku University. One is to use a magnetic Laval nozzle in the vicinity of the MPDA muzzle for converting the high ion thermal energy to the axial flow energy. The other is to heat ions by use of an ICRF antenna in the divergent magnetic nozzle. It is found that by use of a small-sized Laval-type magnetic nozzle, the subsonic flow near the muzzle is converted to be supersonic through the magnetic Laval nozzle. A fast-flowing plasma is successfully heated by use of an ICRF antenna in the magnetic beach configuration

Brayton-Cycle Power-Conversion Unit Tested With Ion Thruster

Science.gov (United States)

Hervol, David S.

2005-01-01

Nuclear electric propulsion has been identified as an enabling technology for future NASA space science missions, such as the Jupiter Icy Moons Orbiter (JIMO) now under study. An important element of the nuclear electric propulsion spacecraft is the power conversion system, which converts the reactor heat to electrical power for use by the ion propulsion system and other spacecraft loads. The electrical integration of the power converter and ion thruster represents a key technical challenge in making nuclear electric propulsion technology possible. This technical hurdle was addressed extensively on December 1, 2003, when a closed- Brayton-cycle power-conversion unit was tested with a gridded ion thruster at the NASA Glenn Research Center. The test demonstrated end-to-end power throughput and marked the first-ever coupling of a Brayton turbo alternator and a gridded ion thruster, both of which are candidates for use on JIMO-type missions. The testing was conducted at Glenn's Vacuum Facility 6, where the Brayton unit was installed in the 3-m-diameter vacuum test port and the ion thruster was installed in the 7.6-m-diameter main chamber.

Empirical electron cross-field mobility in a Hall effect thruster

International Nuclear Information System (INIS)

Garrigues, L.; Perez-Luna, J.; Lo, J.; Hagelaar, G. J. M.; Boeuf, J. P.; Mazouffre, S.

2009-01-01

Electron transport across the magnetic field in Hall effect thrusters is still an open question. Models have so far assumed 1/B 2 or 1/B scaling laws for the 'anomalous' electron mobility, adjusted to reproduce the integrated performance parameters of the thruster. We show that models based on such mobility laws predict very different ion velocity distribution functions (IVDF) than measured by laser induced fluorescence (LIF). A fixed spatial mobility profile, obtained by analysis of improved LIF measurements, leads to much better model predictions of thruster performance and IVDF than 1/B 2 or 1/B mobility laws for discharge voltages in the 500-700 V range.

Carbon Nanotube Based Electric Propulsion Thruster with Low Power Consumption, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — This SBIR project is to develop field emission electric propulsion (FEEP) thruster using carbon nanotubes (CNT) integrated anode. FEEP thrusters have gained...

Pulsed inductive thruster performance data base for megawatt-class engine applications

International Nuclear Information System (INIS)

Dailey, C.L.; Lovberg, R.H.

1993-01-01

The pulsed inductive thruster (PIT) is an electrodeless plasma accelerator employing a large (1m diameter) spiral coil energized by a capacitor bank discharge. The bank can be repetitively recharged by a nuclear electric generator for continuous MW level operation. The coil can be designed as a transformer that permits thruster operation at the generator voltage, which results in a low thruster specific mass. Specific impulse (I sp ) can be readily altered by changing the propellant valve plenum pressure. Performance curves generated from mesausred impulse, injected mass and capacitor bank energy are presented for argon, ammonia, hydrazine, carbon dioxide and helium. The highest performance measured to date is 48% efficiency at 4000 seconds I sp with ammonia. The development of a theoretical model of the thruster, which assumes a fully ionized plasma, is presented in an appendix

Testing of an Arcjet Thruster with Capability of Direct-Drive Operation

Science.gov (United States)

Martin, Adam K.; Polzin, Kurt A.; Eskridge, Richard H.; Smith, James W.; Schoenfeld, Michael P.; Riley, Daniel P.

2015-01-01

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft provided power to the voltage-current that a thruster needs for operation. Testing has been initiated to study whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The work aims to identify and address technical questions related to power conditioning and noise suppression in the system and heating of the thruster in long-duration operation. The apparatus under investigation has a target power level from 400-1,000 W. However, the proposed direct-drive arcjet is potentially a highly scalable concept, applicable to solar-electric spacecraft with up to 100's of kW and beyond. A direct-drive electric propulsion system would be comprised of a thruster that operates with the power supplied directly from the power source (typically solar arrays) with no further power conditioning needed between those two components. Arcjet thrusters are electric propulsion devices, with the power supplied as a high current at low voltage; of all the different types of electric thruster, they are best suited for direct drive from solar arrays. One advantage of an arcjet over Hall or gridded ion thrusters is that for comparable power the arcjet is a much smaller device and can provide more thrust and orders of magnitude higher thrust density (approximately 1-10 N/sq m), albeit at lower I(sub sp) (approximately 800-1000 s). In addition, arcjets are capable of operating on a wide range of propellant options, having been demonstrated on H2, ammonia, N2, Ar, Kr, Xe, while present SOA Hall and ion thrusters are primarily limited to Xe propellant. Direct-drive is often discussed in terms of Hall thrusters, but they require 250-300 V for operation, which is difficult even with high-voltage solar

Mathematical Modeling of Liquid-fed Pulsed Plasma Thruster

Directory of Open Access Journals (Sweden)

Kaartikey Misra

2018-01-01

Full Text Available Liquid propellants are fast becoming attractive for pulsed plasma thrusters due to their high efficiency and low contamination issues. However, the complete plasma interaction and acceleration processes are still not very clear. Present paper develops a multi-layer numerical model for liquid propellant PPTs (pulsed plasma thrusters. The model is based on a quasi-steady flow assumption. The model proposes a possible acceleration mechanism for liquid-fed pulsed plasma thrusters and accurately predicts the propellant utilization capabilities and estimations for the fraction of propellant gas that is completely ionized and accelerated to high exit velocities. Validation of the numerical model and the assumptions on which the model is based on is achieved by comparing the experimental results and the simulation results for two different liquid-fed thrusters developed at the University of Tokyo. Simulation results shows that up-to 50 % of liquid propellant injected is completely ionized and accelerated to high exit velocities (>50 Km/s, whereas, neutral gas contribute to only 7 % of the total specific impulse and accelerated to low exit velocity (<4 Km/s. The model shows an accuracy up-to 92 % . Optimization methods are briefly discussed to ensure efficient propellant utilization and performance. The model acts as a tool to understand the background physics and to optimize the performance for liquid-fed PPTs.

Electromagnetic Spacecraft Propulsion Motor and a Permanent Magnet (PM-Drive) Thruster

Science.gov (United States)

Ahmadov, B. A.

2018-04-01

Ion thrusters are designed to be used for realization of a Mars Sample Return mission. The competing technologies with ion thrusters are electromagnetic spacecraft propulsion motors. I'm an engineer and engage in the creation of the new electromagnetic propulsion motors.

ExB Measurements of a 200 W Xenon Hall Thruster (Preprint)

National Research Council Canada - National Science Library

Ekholm, Jared M; Hargus, Jr, William A

2007-01-01

Angularly resolved ion species fractions of Xe+1, Xe+2, and Xe+3 in a low power xenon Hall thruster Busek BHT-200 plume were measured using an ExB probe under a variety of thruster operating conditions and background pressures...

Analysis and design of ion thruster for large space systems

Science.gov (United States)

Poeschel, R. L.; Kami, S.

1980-01-01

Design analyses showed that an ion thruster of approximately 50 cm in diameter will be required to produce a thrust of 0.5 N using xenon or argon as propellants, and operating the thruster at a specific impulse of 3530 sec or 6076 sec respectively. A multipole magnetic confinement discharge chamber was specified.

Effects of facility backpressure on the performance and plume of a Hall thruster

Science.gov (United States)

Walker, Mitchell Louis Ronald

2005-07-01

This dissertation presents research aimed at understanding the relationship between facility background pressure, Hall thruster performance, and plume characteristics. Due to the wide range of facilities used in Hall thruster testing, it is difficult for researchers to make adequate comparisons between data sets because of both dissimilar instrumentation and backpressures. The differences in the data sets are due to the ingestion of background gas into the Hall thruster discharge channel and charge-exchange collisions in the plume. Thus, this research aims to understand facility effects and to develop the tools needed to allow researchers to obtain relevant plume and performance data for a variety of chambers and backpressures. The first portion of this work develops a technique for calibrating a vacuum chamber in terms of pressure to account for elevated backpressures while testing Hall thrusters. Neutral gas background pressure maps of the Large Vacuum Test Facility are created at a series of cold anode flow rates and one hot flow rate at two UM/AFRL P5 5 kW Hall thruster operating conditions. These data show that a cold flow pressure map can be used to approximate the neutral background pressure in the chamber with the thruster in operation. In addition, the data are used to calibrate a numerical model that accurately predicts facility backpressure within a vacuum chamber of specified geometry and pumping speed. The second portion of this work investigates how facility backpressure influences the plume, plume diagnostics, and performance of the P5 Hall thruster. Measurements of the plume and performance characteristics over a wide range of pressures show that ingestion, a decrease in the downstream plasma potential, and broadening of the ion energy distribution function cause the increase in thrust with backpressure. Furthermore, a magnetically-filtered Faraday probe accurately measures ion current density at elevated operating pressures. The third portion of

Electronegative Gas Thruster - Direct Thrust Measurement

Data.gov (United States)

National Aeronautics and Space Administration — This effort is an international collaboration and academic partnership to mature an innovative electric propulsion (EP) thruster concept to TRL 3 through direct...

A cavity ring-down spectroscopy sensor for real-time Hall thruster erosion measurements

International Nuclear Information System (INIS)

Lee, B. C.; Huang, W.; Tao, L.; Yamamoto, N.; Yalin, A. P.; Gallimore, A. D.

2014-01-01

A continuous-wave cavity ring-down spectroscopy sensor for real-time measurements of sputtered boron from Hall thrusters has been developed. The sensor uses a continuous-wave frequency-quadrupled diode laser at 250 nm to probe ground state atomic boron sputtered from the boron nitride insulating channel. Validation results from a controlled setup using an ion beam and target showed good agreement with a simple finite-element model. Application of the sensor for measurements of two Hall thrusters, the H6 and SPT-70, is described. The H6 was tested at power levels ranging from 1.5 to 10 kW. Peak boron densities of 10 ± 2 × 10 14 m −3 were measured in the thruster plume, and the estimated eroded channel volume agreed within a factor of 2 of profilometry. The SPT-70 was tested at 600 and 660 W, yielding peak boron densities of 7.2 ± 1.1 × 10 14 m −3 , and the estimated erosion rate agreed within ∼20% of profilometry. Technical challenges associated with operating a high-finesse cavity in the presence of energetic plasma are also discussed

Facility Effect Characterization Test of NASA's HERMeS Hall Thruster

Science.gov (United States)

Huang, Wensheng; Kamhawi, Hani; Haag, Thomas W.; Ortega, Alejandro Lopez; Mikellides, Ioannis G.

2016-01-01

A test to characterize the effect of varying background pressure on NASA's 12.5-kW Hall Effect Rocket with Magnetic Shielding had being completed. This thruster is the baseline propulsion system for the Solar Electric Propulsion Technology Demonstration Mission (SEP TDM). Potential differences in thruster performance and oscillation characteristics when in ground facilities versus on-orbit are considered a primary risk for the propulsion system of the Asteroid Redirect Robotic Mission, which is a candidate for SEP TDM. The first primary objective of this test was to demonstrate that the tools being developed to predict the zero-background-pressure behavior of the thruster can provide self-consistent results. The second primary objective of this test was to provide data for refining a physics-based model of the thruster plume that will be used in spacecraft interaction studies. Diagnostics deployed included a thrust stand, Faraday probe, Langmuir probe, retarding potential analyzer, Wien filter spectrometer, and high-speed camera. From the data, a physics-based plume model was refined. Comparisons of empirical data to modeling results are shown.

Electric field measurement in microwave discharge ion thruster with electro-optic probe.

Science.gov (United States)

Ise, Toshiyuki; Tsukizaki, Ryudo; Togo, Hiroyoshi; Koizumi, Hiroyuki; Kuninaka, Hitoshi

2012-12-01

In order to understand the internal phenomena in a microwave discharge ion thruster, it is important to measure the distribution of the microwave electric field inside the discharge chamber, which is directly related to the plasma production. In this study, we proposed a novel method of measuring a microwave electric field with an electro-optic (EO) probe based on the Pockels effect. The probe, including a cooling system, contains no metal and can be accessed in the discharge chamber with less disruption to the microwave distribution. This method enables measurement of the electric field profile under ion beam acceleration. We first verified the measurement with the EO probe by a comparison with a finite-difference time domain numerical simulation of the microwave electric field in atmosphere. Second, we showed that the deviations of the reflected microwave power and the beam current were less than 8% due to inserting the EO probe into the ion thruster under ion beam acceleration. Finally, we successfully demonstrated the measurement of the electric-field profile in the ion thruster under ion beam acceleration. These measurements show that the electric field distribution in the thruster dramatically changes in the ion thruster under ion beam acceleration as the propellant mass flow rate increases. These results indicate that this new method using an EO probe can provide a useful guide for improving the propulsion of microwave discharge ion thrusters.

Magnetic Field Effects on the Plume of a Diverging Cusped-Field Thruster

KAUST Repository

Matlock, Taylor

2010-07-25

The Diverging Cusped-Field Thruster (DCFT) uses three permanent ring magnets of alternating polarity to create a unique magnetic topology intended to reduce plasma losses to the discharge chamber surfaces. The magnetic field strength within the DCFT discharge chamber (up to 4 kG on axis) is much higher than in thrusters of similar geometry, which is believed to be a driving factor in the high measured anode efficiencies. The field strength in the near plume region is large as well, which may bear on the high beam divergences measured, with peaks in ion current found at angles of around 30-35 from the thruster axis. Characterization of the DCFT has heretofore involved only one magnetic topology. It is then the purpose of this study to investigate changes to the near-field plume caused by altering the shape and strength of the magnetic field. A thick magnetic collar, encircling the thruster body, is used to lower the field strength outside of the discharge chamber and thus lessen any effects caused by the external field. Changes in the thruster plume with field topology are monitored by the use of normal Langmuir and emissive probes interrogating the near-field plasma. Results are related to other observations that suggest a unified conceptual framework for the important near-exit region of the thruster.

Propellantless Attitude Control of Solar Sail Technology Utilizing Reflective Control Devices

Science.gov (United States)

Munday, Jeremy

2016-01-01

Solar sails offer an opportunity for a CubeSatscale, propellant-free spacecraft technology that enables long-term and long-distance missions not possible with traditional methods. Solar sails operate using the transfer of linear momentum from photons of sunlight reflected from the surface of the sail. To propel the spacecraft, no mechanically moving parts, thrusters, or propellant are needed. However, attitude control, or orientation, is still performed using traditional methods involving reaction wheels and propellant ejection, which severely limit mission lifetime. For example, the current state of the art solutions employed by upcoming missions couple solar sails with a state of the art propellant ejection gas system. Here, the use of the gas thruster has limited the lifetime of the mission. To solve the limited mission lifetime problem, the Propellantless Attitude Control of Solar Sail Technology Utilizing Reflective Control Devices project team is working on propellantless attitude control using thin layers of material, an optical film, electrically switchable from transparent to reflective. The technology is based on a polymer-dispersed liquid crystal (PDLC), which allows this switch upon application of a voltage. This technology removes the need for propellant, which reduces weight and cost while improving performance and lifetime.

Development of a 30-cm ion thruster thermal-vacuum power processor

Science.gov (United States)

Herron, B. G.

1976-01-01

The 30-cm Hg electron-bombardment ion thruster presently under development has reached engineering model status and is generally accepted as the prime propulsion thruster module to be used on the earliest solar electric propulsion missions. This paper presents the results of a related program to develop a transistorized 3-kW Thermal-Vacuum Breadboard (TVBB) Power Processor for this thruster. Emphasized in the paper are the implemented electrical and mechanical designs as well as the resultant system performance achieved over a range of test conditions. In addition, design modifications affording improved performance are identified and discussed.

Overview of NASA GRCs Green Propellant Infusion Mission Thruster Testing and Plume Diagnostics

Science.gov (United States)

Deans, Matthew C.; Reed, Brian D.; Yim, John T.; Arrington, Lynn A.; Williams, George J.; Kojima, Jun J.; McLean, Christopher H.

2014-01-01

The Green Propellant Infusion Mission (GPIM) is sponsored by NASA's Space Technology Mission Directorate (STMD) Technology Demonstration Mission (TDM) office. The goal of GPIM is to advance the technology readiness level of a green propulsion system, specifically, one using the monopropellant, AF-M315E, by demonstrating ground handling, spacecraft processing, and on-orbit operations. One of the risks identified for GPIM is potential contamination of sensitive spacecraft surfaces from the effluents in the plumes of AF-M315E thrusters. NASA Glenn Research Center (GRC) is conducting activities to characterize the effects of AF-M315E plume impingement and deposition. GRC has established individual plume models of the 22-N and 1-N thrusters that will be used on the GPIM spacecraft. The models describe the pressure, temperature, density, Mach number, and species concentration of the AF-M315E thruster exhaust plumes. The models are being used to assess the impingement effects of the AF-M315E thrusters on the GPIM spacecraft. The model simulations will be correlated with plume measurement data from Laboratory and Engineering Model 22-N, AF-M315E thrusters. The thrusters will be tested in a small rocket, altitude facility at NASA GRC. The GRC thruster testing will be conducted at duty cycles representatives of the planned GPIM maneuvers. A suite of laser-based diagnostics, including Raman spectroscopy, Rayleigh spectroscopy, Schlieren imaging, and physical probes will be used to acquire plume measurements of AFM315E thrusters. Plume data will include temperature, velocity, relative density, and species concentration. The plume measurement data will be compared to the corresponding simulations of the plume model. The GRC effort will establish a data set of AF-M315E plume measurements and a plume model that can be used for future AF-M315E applications.

Power Dependence of the Electron Mobility Profile in a Hall Thruster

Science.gov (United States)

Jorns, Benjamin A.; Hofery, Richard H.; Mikellides, Ioannis G.

2014-01-01

The electron mobility profile is estimated in a 4.5 kW commercial Hall thruster as a function of discharge power. Internal measurements of plasma potential and electron temperature are made in the thruster channel with a high-speed translating probe. These measurements are presented for a range of throttling conditions from 150 - 400 V and 0.6 - 4.5 kW. The fluid-based solver, Hall2De, is used in conjunction with these internal plasma parameters to estimate the anomalous collision frequency profile at fixed voltage, 300 V, and three power levels. It is found that the anomalous collision frequency profile does not change significantly upstream of the location of the magnetic field peak but that the extent and magnitude of the anomalous collision frequency downstream of the magnetic peak does change with thruster power. These results are discussed in the context of developing phenomenological models for how the collision frequency profile depends on thruster operating conditions.

Low-Power Operation and Plasma Characterization of a Qualification Model SPT-140 Hall Thruster for NASA Science Missions

Science.gov (United States)

Garner, Charles E.; Jorns, Benjamin A.; van Derventer, Steven; Hofer, Richard R.; Rickard, Ryan; Liang, Raymond; Delgado, Jorge

2015-01-01

Hall thruster systems based on commercial product lines can potentially lead to lower cost electric propulsion (EP) systems for deep space science missions. A 4.5-kW SPT-140 Hall thruster presently under qualification testing by SSL leverages the substantial heritage of the SPT-100 being flown on Russian and US commercial satellites. The Jet Propulsion Laboratory is exploring the use of commercial EP systems, including the SPT-140, for deep space science missions, and initiated a program to evaluate the SPT-140 in the areas of low power operation and thruster operating life. A qualification model SPT-140 designated QM002 was evaluated for operation and plasma properties along channel centerline, from 4.5 kW to 0.8 kW. Additional testing was performed on a development model SPT-140 designated DM4 to evaluate operation with a Moog proportional flow control valve (PFCV). The PFCV was commanded by an SSL engineering model PPU-140 Power Processing Unit (PPU). Performance measurements on QM002 at 0.8 kW discharge power were 50 mN of thrust at a total specific impulse of 1250 s, a total thruster efficiency of 0.38, and discharge current oscillations of under 3% of the mean current. Steady-state operation at 0.8 kW was demonstrated during a 27 h firing. The SPT-140 DM4 was operated in closed-loop control of the discharge current with the PFCV and PPU over discharge power levels of 0.8-4.5 kW. QM002 and DM4 test data indicate that the SPT-140 design is a viable candidate for NASA missions requiring power throttling down to low thruster input power.

Modeling of physical processes in radio-frequency plasma thrusters

OpenAIRE

Tian, Bin

2017-01-01

This Thesis presents an investigation of the plasma-wave interaction in Helicon Plasma Thrusters (HPT). The HPT is a new concept of electric space propulsion, which generates plasmas with RF heating and provides thrust by the electrodeless acceleration of plasmas in a magnetic nozzle. An in-depth and extensive literature review of the state of the art of the models and experiments of plasma-wave interaction in helicon plasma sources and thrusters is carried out. Then, a theoret...

Characteristics of the LeRC/Hughes J-series 30-cm engineering model thruster

Science.gov (United States)

Collett, C. R.; Poeschel, R. L.; Kami, S.

1981-01-01

As a consequence of endurance and structural tests performed on 900-series engineering model thrusters (EMT), several modifications in design were found to be necessary for achieving performance goals. The modified thruster is known as the J-series EMT. The most important of the design modifications affect the accelerator grid, gimbal mount, cathode polepiece, and wiring harness. The paper discusses the design modifications incorporated, the condition(s) they corrected, and the characteristics of the modified thruster.

Design and numerical evaluation of full-authority flight control systems for conventional and thruster-augmented helicopters employed in NOE operations

Science.gov (United States)

Perri, Todd A.; Mckillip, R. M., Jr.; Curtiss, H. C., Jr.

1987-01-01

The development and methodology is presented for development of full-authority implicit model-following and explicit model-following optimal controllers for use on helicopters operating in the Nap-of-the Earth (NOE) environment. Pole placement, input-output frequency response, and step input response were used to evaluate handling qualities performance. The pilot was equipped with velocity-command inputs. A mathematical/computational trajectory optimization method was employed to evaluate the ability of each controller to fly NOE maneuvers. The method determines the optimal swashplate and thruster input histories from the helicopter's dynamics and the prescribed geometry and desired flying qualities of the maneuver. Three maneuvers were investigated for both the implicit and explicit controllers with and without auxiliary propulsion installed: pop-up/dash/descent, bob-up at 40 knots, and glideslope. The explicit controller proved to be superior to the implicit controller in performance and ease of design.

Thruster allocation for dynamical positioning

NARCIS (Netherlands)

Poppe, K.; van den Berg, J.B.; Blank, E.; Archer, C.; Redeker, M.; Kutter, M.; Hemker, P.

2010-01-01

Positioning a vessel at a fixed position in deep water is of great importance when working offshore. In recent years a Dynamical Positioning (DP) system was developed at Marin [2]. After the measurement of the current position and external forces (like waves, wind etc.), each thruster of the vessel

In-Situ Measurement of Hall Thruster Erosion Using a Fiber Optic Regression Probe

Science.gov (United States)

Polzin, Kurt; Korman, Valentin

2009-01-01

One potential life-limiting mechanism in a Hall thruster is the erosion of the ceramic material comprising the discharge channel. This is especially true for missions that require long thrusting periods and can be problematic for lifetime qualification, especially when attempting to qualify a thruster by analysis rather than a test lasting the full duration of the mission. In addition to lifetime, several analytical and numerical models include electrode erosion as a mechanism contributing to enhanced transport properties. However, there is still a great deal of dispute over the importance of erosion to transport in Hall thrusters. The capability to perform an in-situ measurement of discharge channel erosion is useful in addressing both the lifetime and transport concerns. An in-situ measurement would allow for real-time data regarding the erosion rates at different operating points, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over a thruster s operating envelope would also be useful in the modeling of the detailed physics inside the discharge chamber. There are many different sensors and techniques that have been employed to quantify discharge channel erosion in Hall thrusters. Snapshots of the wear pattern can be obtained at regular shutdown intervals using laser profilometry. Many non-intrusive techniques of varying complexity and sensitivity have been employed to detect the time-varying presence of erosion products in the thruster plume. These include the use quartz crystal microbalances, emission spectroscopy, laser induced flourescence, and cavity ring-down spectroscopy. While these techniques can provide a very accurate picture of the level of eroded material in the thruster plume, it is more difficult to use them to determine the location from which the material was eroded. Furthermore, none of the methods cited provide a true in-situ measure of erosion at the channel surface while

Spatiotemporal study of gas heating mechanisms in a radio-frequency electrothermal plasma micro-thruster

Directory of Open Access Journals (Sweden)

Amelia eGreig

2015-10-01

Full Text Available A spatiotemporal study of neutral gas temperature during the first 100 s of operation for a radio-frequency electrothermal plasma micro-thruster operating on nitrogen at 60 W and 1.5 Torr is performed to identify the heating mechanisms involved. Neutral gas temperature is estimated from rovibrational band fitting of the nitrogen second positive system. A set of baffles are used to restrict the optical image and separate the heating mechanisms occurring in the central bulk discharge region and near the thruster walls.For each spatial region there are three distinct gas heating mechanisms being fast heating from ion-neutral collisions with timescales of tens of milliseconds, intermediate heating with timescales of 10 s from ion bombardment on the inner thruster tube surface creating wall heating, and slow heating with timescales of 100 s from gradual warming of the entire thruster housing. The results are discussed in relation to optimising the thermal properties of future thruster designs.

Integration Testing of a Modular Discharge Supply for NASA's High Voltage Hall Accelerator Thruster

Science.gov (United States)

Pinero, Luis R.; Kamhawi, hani; Drummond, Geoff

2010-01-01

NASA s In-Space Propulsion Technology Program is developing a high performance Hall thruster that can fulfill the needs of future Discovery-class missions. The result of this effort is the High Voltage Hall Accelerator thruster that can operate over a power range from 0.3 to 3.5 kW and a specific impulse from 1,000 to 2,800 sec, and process 300 kg of xenon propellant. Simultaneously, a 4.0 kW discharge power supply comprised of two parallel modules was developed. These power modules use an innovative three-phase resonant topology that can efficiently supply full power to the thruster at an output voltage range of 200 to 700 V at an input voltage range of 80 to 160 V. Efficiencies as high as 95.9 percent were measured during an integration test with the NASA103M.XL thruster. The accuracy of the master/slave current sharing circuit and various thruster ignition techniques were evaluated.

Global Linear Stability Analysis of the Spoke Oscillation in Hall Effect Thrusters

Science.gov (United States)

2014-07-15

meνeχ 2 nTe qex (4.1f) ddc dx = 2cpl vix ≡ γ (4.1g) where x is the axial coordinate along the thruster channel; e, me and mi are the electron charge...mi ) P − ( 5 2 Te mi nvex + qex mi ) 1 dc ddc dξ (4.25i) ddc dξ = Pγ (4.25j) Distribution A: Approved for public release; distribution is unlimited...Thruster. PhD thesis, Standford University , 2011. [128] D. Liu, R.E. Huffman, R.D. Branam, and W.A. Hargus. Ultrahigh-speed imaging of hall-thruster

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

Science.gov (United States)

Soulas, George C.; Shastry, Rohit

2016-01-01

A Long Duration Test (LDT) was initiated in June 2005 as a part of NASAs Evolutionary Xenon Thruster (NEXT) service life validation approach. Testing was voluntarily terminated in February 2014, with the thruster accumulating 51,184 hours of operation, processing 918 kg of xenon propellant, and delivering 35.5 MN-s of total impulse. This presentation will present the post-test inspection results to date for the thrusters ion optics.

Chaotic waves in Hall thruster plasma

International Nuclear Information System (INIS)

Peradzynski, Zbigniew; Barral, S.; Kurzyna, J.; Makowski, K.; Dudeck, M.

2006-01-01

The set of hyperbolic equations of the fluid model describing the acceleration of plasma in a Hall thruster is analyzed. The characteristic feature of the flow is the existence of a trapped characteristic; i.e. there exists a characteristic line, which never intersects the boundary of the flow region in the thruster. To study the propagation of short wave perturbations, the approach of geometrical optics (like WKB) can be applied. This can be done in a linear as well as in a nonlinear version. The nonlinear version describes the waves of small but finite amplitude. As a result of such an approach one obtains so called transport equation, which are governing the wave amplitude. Due to the existence of trapped characteristics this transport equation appears to have chaotic (turbulent) solutions in both, linear and nonlinear versions

The Green Propellant Infusion Mission Thruster Performance Testing for Plume Diagnostics

Science.gov (United States)

Deans, Matthew C.; Reed, Brian D.; Arrington, Lynn A.; Williams, George J.; Kojima, Jun J.; Kinzbach, McKenzie I.; McLean, Christopher H.

2014-01-01

The Green Propellant Infusion Mission (GPIM) is sponsored by NASA's Space Technology Mission Directorate (STMD) Technology Demonstration Mission (TDM) office. The goal of GPIM is to advance the technology readiness level of a green propulsion system, specifically, one using the monopropellant, AF-M315E, by demonstrating ground handling, spacecraft processing, and on-orbit operations. One of the risks identified for GPIM is potential contamination of sensitive spacecraft surfaces from the effluents in the plumes of AF-M315E thrusters. NASA Glenn Research Center (GRC) is conducting activities to characterize the effects of AF-M315E plume impingement and deposition. GRC has established individual plume models of the 22-N and 1-N thrusters that will be used on the GPIM spacecraft. The model simulations will be correlated with plume measurement data from Laboratory and Engineering Model 22-N, AF-M315E thrusters. The thrusters are currently being tested in a small rocket, altitude facility at NASA GRC. A suite of diagnostics, including Raman spectroscopy, Rayleigh spectroscopy, and Schlieren imaging are being used to acquire plume measurements of AF-M315E thrusters. Plume data will include temperature, velocity, relative density, and species concentration. The plume measurement data will be compared to the corresponding simulations of the plume model. The GRC effort will establish a data set of AF-M315E plume measurements and a plume model that can be used for future AF-M315E applications.

Micro-cathode Arc Thruster PhoneSat Experiment

Data.gov (United States)

National Aeronautics and Space Administration — The Micro-cathode Arc Thruster Phonesat Experiment  was a joint project between George Washington University and NASA Ames Research Center that successfully...

Hot-Fire Testing of a 1N AF-M315E Thruster

Science.gov (United States)

Burnside, Christopher G.; Pedersen, Kevin; Pierce, Charles W.

2015-01-01

This hot-fire test continues NASA investigation of green propellant technologies for future missions. To show the potential for green propellants to replace some hydrazine systems in future spacecraft, NASA Marshall Space Flight Center (MSFC) is continuing to embark on hot-fire test campaigns with various green propellant blends. NASA completed a hot-fire test of a 1N AF-M315E monopropellant thruster at the Marshall Space Flight Center in the small altitude test stand located in building 4205. The thruster is a ground test article used for basic performance determination and catalyst studies. The purpose of the hot-fire testing was for performance determination of a 1N size thruster and form a baseline from which to study catalyst performance and life with follow-on testing to be conducted at a later date. The thruster performed as expected. The result of the hot-fire testing are presented in this paper and presentation.

Plasma simulation in space propulsion : the helicon plasma thruster

OpenAIRE

Navarro Cavallé, Jaume

2017-01-01

The Helicon Plasma Thruster (HPT) is an electrodynamic rocket proposed in the early 2000s. It matches an Helicon Plasma Source (HPS), which ionizes the neutral gas and heats up the plasma, with aMagneticNozzle (MN),where the plasma is supersonically accelerated resulting in thrust. Although the core of this thruster inherits the knowledge on Helicon Plasma sources, dated from the seventies, the HPT technology is still not developed and remains below TRL 4. A deep review of the HPT State-of-ar...

Performance optimization of 20 cm xenon ion thruster discharge chamber

International Nuclear Information System (INIS)

Chen Juanjuan; Zhang Tianping; Jia Yanhui; Li Xiaoping

2012-01-01

This paper describes the performance of the LIPS-200 ion thruster discharge chamber which was developed by Lanzhou Institute of Physics. Based on the discharge chamber geometric configuration and magnetic field, the completely self-consistent analytical model is utilized to discuss performance optimization of the discharge chamber of the LIPS-200. The thrust is enhanced from 40 mN up to 60 mN at rated impulse and efficiency. The results show that the 188.515 W/A beam ion production cost at a propellant flow rate of 2.167 × 10 17 m -3 requires that the thruster runs at a discharge current of 6.9 A to produce 1.2 A ion beam current. Also, during the process of LIPS-200 ion thruster discharge chamber performance optimization, the sheath potential is always within 3.80 ∼ 6.65 eV. (authors)

Mode transition of a Hall thruster discharge plasma

International Nuclear Information System (INIS)

Hara, Kentaro; Sekerak, Michael J.; Boyd, Iain D.; Gallimore, Alec D.

2014-01-01

A Hall thruster is a cross-field plasma device used for spacecraft propulsion. An important unresolved issue in the development of Hall thrusters concerns the effect of discharge oscillations in the range of 10–30 kHz on their performance. The use of a high speed Langmuir probe system and ultra-fast imaging of the discharge plasma of a Hall thruster suggests that the discharge oscillation mode, often called the breathing mode, is strongly correlated to an axial global ionization mode. Stabilization of the global oscillation mode is achieved as the magnetic field is increased and azimuthally rotating spokes are observed. A hybrid-direct kinetic simulation that takes into account the transport of electronically excited atoms is used to model the discharge plasma of a Hall thruster. The predicted mode transition agrees with experiments in terms of the mean discharge current, the amplitude of discharge current oscillation, and the breathing mode frequency. It is observed that the stabilization of the global oscillation mode is associated with reduced electron transport that suppresses the ionization process inside the channel. As the Joule heating balances the other loss terms including the effects of wall loss and inelastic collisions, the ionization oscillation is damped, and the discharge oscillation stabilizes. A wide range of the stable operation is supported by the formation of a space charge saturated sheath that stabilizes the electron axial drift and balances the Joule heating as the magnetic field increases. Finally, it is indicated from the numerical results that there is a strong correlation between the emitted light intensity and the discharge current.

Development and Control of the Naval Postgraduate School Planar Autonomous Docking Simulator (NPADS)

Science.gov (United States)

Porter, Robert D.

2002-09-01

The objective of this thesis was to design, construct and develop the initial autonomous control algorithm for the NPS Planar Autonomous Docking Simulator (NPADS) The effort included hardware design, fabrication, installation and integration; mass property determination; and the development and testing of control laws utilizing MATLAB and Simulink for modeling and LabView for NPADS control, The NPADS vehicle uses air pads and a granite table to simulate a 2-D, drag-free, zero-g space environment, It is a completely self-contained vehicle equipped with eight cold-gas, bang-bang type thrusters and a reaction wheel for motion control, A 'star sensor' CCD camera locates the vehicle on the table while a color CCD docking camera and two robotic arms will locate and dock with a target vehicle, The on-board computer system leverages PXI technology and a single source, simplifying systems integration, The vehicle is powered by two lead-acid batteries for completely autonomous operation, A graphical user interface and wireless Ethernet enable the user to command and monitor the vehicle from a remote command and data acquisition computer. Two control algorithms were developed and allow the user to either control the thrusters and reaction wheel manually or simply specify a desired location and rotation angle,

Design and Testing of a Hall Effect Thruster with 3D Printed Channel and Propellant Distributor

Science.gov (United States)

Hopping, Ethan P.; Xu, Kunning G.

2017-01-01

The UAH-78AM is a low-power Hall effect thruster developed at the University of Alabama in Huntsville with channel walls and a propellant distributor manufactured using 3D printing. The goal of this project is to assess the feasibility of using unconventional materials to produce a low-cost functioning Hall effect thruster and consider how additive manufacturing can expand the design space and provide other benefits. A version of the thruster was tested at NASA Glenn Research Center to obtain performance metrics and to validate the ability of the thruster to produce thrust and sustain a discharge. An overview of the thruster design and transient performance measurements are presented here. Measured thrust ranged from 17.2 millinewtons to 30.4 millinewtons over a discharge power of 280 watts to 520 watts with an anode I (sub SP)(Specific Impulse) range of 870 seconds to 1450 seconds. Temperature limitations of materials used for the channel walls and propellant distributor limit the ability to run the thruster at thermal steady-state.

Rarefied gas electro jet (RGEJ) micro-thruster for space propulsion

Science.gov (United States)

Blanco, Ariel; Roy, Subrata

2017-11-01

This article numerically investigates a micro-thruster for small satellites which utilizes plasma actuators to heat and accelerate the flow in a micro-channel with rarefied gas in the slip flow regime. The inlet plenum condition is considered at 1 Torr with flow discharging to near vacuum conditions (consumption and the thrust effectiveness of the thruster are predicted based on these results. The ionized gas is modelled using local mean energy approximation. An electrically induced body force and a thermal heating source are calculated based on the space separated charge distribution and the ion Joule heating, respectively. The rarefied gas flow with these electric force and heating source is modelled using density-based compressible flow equations with slip flow boundary conditions. The results show that a significant improvement of specific impulse can be achieved over highly optimized cold gas thrusters using the same propellant.

Plasma Reactors and Plasma Thrusters Modeling by Ar Complete Global Models

Directory of Open Access Journals (Sweden)

Chloe Berenguer

2012-01-01

Full Text Available A complete global model for argon was developed and adapted to plasma reactor and plasma thruster modeling. It takes into consideration ground level and excited Ar and Ar+ species and the reactor and thruster form factors. The electronic temperature, the species densities, and the ionization percentage, depending mainly on the pressure and the absorbed power, have been obtained and commented for various physical conditions.

The microwave thermal thruster and its application to the launch problem

Science.gov (United States)

Parkin, Kevin L. G.

Nuclear thermal thrusters long ago bypassed the 50-year-old specific impulse (Isp) limitation of conventional thrusters, using nuclear powered heat exchangers in place of conventional combustion to heat a hydrogen propellant. These heat exchanger thrusters experimentally achieved an Isp of 825 seconds, but with a thrust-to-weight ratio (T/W) of less than ten they have thus far been too heavy to propel rockets into orbit. This thesis proposes a new idea to achieve both high Isp and high T/W The Microwave Thermal Thruster. This thruster covers the underside of a rocket aeroshell with a lightweight microwave absorbent heat exchange layer that may double as a re-entry heat shield. By illuminating the layer with microwaves directed from a ground-based phased array, an Isp of 700--900 seconds and T/W of 50--150 is possible using a hydrogen propellant. The single propellant simplifies vehicle design, and the high Isp increases payload fraction and structural margins. These factors combined could have a profound effect on the economics of building and reusing rockets. A laboratory-scale microwave thermal heat exchanger is constructed using a single channel in a cylindrical microwave resonant cavity, and new type of coupled electromagnetic-conduction-convection model is developed to simulate it. The resonant cavity approach to small-scale testing reveals several drawbacks, including an unexpected oscillatory behavior. Stable operation of the laboratory-scale thruster is nevertheless successful, and the simulations are consistent with the experimental results. In addition to proposing a new type of propulsion and demonstrating it, this thesis provides three other principal contributions: The first is a new perspective on the launch problem, placing it in a wider economic context. The second is a new type of ascent trajectory that significantly reduces the diameter, and hence cost, of the ground-based phased array. The third is an eclectic collection of data, techniques, and

Hall-Effect Thruster Simulations with 2-D Electron Transport and Hydrodynamic Ions

Science.gov (United States)

Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard H.; Goebel, Dan M.

2009-01-01

A computational approach that has been used extensively in the last two decades for Hall thruster simulations is to solve a diffusion equation and energy conservation law for the electrons in a direction that is perpendicular to the magnetic field, and use discrete-particle methods for the heavy species. This "hybrid" approach has allowed for the capture of bulk plasma phenomena inside these thrusters within reasonable computational times. Regions of the thruster with complex magnetic field arrangements (such as those near eroded walls and magnets) and/or reduced Hall parameter (such as those near the anode and the cathode plume) challenge the validity of the quasi-one-dimensional assumption for the electrons. This paper reports on the development of a computer code that solves numerically the 2-D axisymmetric vector form of Ohm's law, with no assumptions regarding the rate of electron transport in the parallel and perpendicular directions. The numerical challenges related to the large disparity of the transport coefficients in the two directions are met by solving the equations in a computational mesh that is aligned with the magnetic field. The fully-2D approach allows for a large physical domain that extends more than five times the thruster channel length in the axial direction, and encompasses the cathode boundary. Ions are treated as an isothermal, cold (relative to the electrons) fluid, accounting for charge-exchange and multiple-ionization collisions in the momentum equations. A first series of simulations of two Hall thrusters, namely the BPT-4000 and a 6-kW laboratory thruster, quantifies the significance of ion diffusion in the anode region and the importance of the extended physical domain on studies related to the impact of the transport coefficients on the electron flow field.

HiVHAc Thruster Wear and Structural Tests

Data.gov (United States)

National Aeronautics and Space Administration — NASA GRC is developing a 4.5 kW-class Hall propulsion system. This system includes a long life high performance Hall Effect Thruster (HET), a highly efficient...

1000 Hours of Testing Completed on 10-kW Hall Thruster

Science.gov (United States)

Mason, Lee S.

2001-01-01

Between the months of April and August 2000, a 10-kW Hall effect thruster, designated T- 220, was subjected to a 1000-hr life test evaluation. Hall effect thrusters are propulsion devices that electrostatically accelerate xenon ions to produce thrust. Hall effect propulsion has been in development for many years, and low-power devices (1.35 kW) have been used in space for satellite orbit maintenance. The T-220, shown in the photo, produces sufficient thrust to enable efficient orbital transfers, saving hundreds of kilograms in propellant over conventional chemical propulsion systems. This test is the longest operation ever achieved on a high-power Hall thruster (greater than 4.5 kW) and is a key milestone leading to the use of this technology for future NASA, commercial, and military missions.

Reduced power processor requirements for the 30-cm diameter HG ion thruster

Science.gov (United States)

Rawlin, V. K.

1979-01-01

The characteristics of power processors strongly impact the overall performance and cost of electric propulsion systems. A program was initiated to evaluate simplifications of the thruster-power processor interface requirements. The power processor requirements are mission dependent with major differences arising for those missions which require a nearly constant thruster operating point (typical of geocentric and some inbound planetary missions) and those requiring operation over a large range of input power (such as outbound planetary missions). This paper describes the results of tests which have indicated that as many as seven of the twelve power supplies may be eliminated from the present Functional Model Power Processor used with 30-cm diameter Hg ion thrusters.

High Fidelity Modeling of Field-Reversed Configuration (FRC) Thrusters (Briefing Charts)

Science.gov (United States)

2017-05-24

THRUSTERS (Briefing Charts) Robert Martin , Eder Sousa, Jonathan Tran Air Force Research Laboratory (AFMC) AFRL/RQRS 1 Ara Drive Edwards AFB, CA 93524... Martin N/A HIGH FIDELITY MODELING OF FIELD-REVERSED CONFIGURATION (FRC) THRUSTERS Robert Martin1, Eder Sousa2, Jonathan Tran2 1AIR FORCE RESEARCH...Distribution is unlimited. PA Clearance No. 17314 MARTIN , SOUSA, TRAN (AFRL/RQRS) DISTRIBUTION A - APPROVED FOR PUBLIC RELEASE; DISTRIBUTION UNLIMITED. PA

Recycle Requirements for NASA's 30 cm Xenon Ion Thruster

Science.gov (United States)

Pinero, Luis R.; Rawlin, Vincent K.

1994-01-01

Electrical breakdowns have been observed during ion thruster operation. These breakdowns, or arcs, can be caused by several conditions. In flight systems, the power processing unit must be designed to handle these faults autonomously. This has a strong impact on power processor requirements and must be understood fully for the power processing unit being designed for the NASA Solar Electric Propulsion Technology Application Readiness program. In this study, fault conditions were investigated using a NASA 30 cm ion thruster and a power console. Power processing unit output specifications were defined based on the breakdown phenomena identified and characterized.

Optimized Magnetic Nozzles for MPD Thrusters, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Magnetoplasmadynamic (MPD) thrusters can provide the high-specific impulse, high-power propulsion required to enable ambitious human and robotic exploration missions...

Demonstration of a Non-Toxic Reaction Control Engine

Science.gov (United States)

Robinson, Philip J.; Turpin, Alicia A.; Veith, Eric M.

2007-01-01

T:hree non-toxic demonstration reaction control engines (RCE) were successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the auspices of the Exploration Systems Mission Directorate. The demonstration engine utilized Liquid Oxygen (LOX) and Ethanol as propellants to produce 870 lbf thrust. The Aerojet RCE's were successfully acceptance tested over a broad range of operating conditions. Steady state tests evaluated engine response to varying chamber pressures and mixture ratios. In addition to the steady state tests, a variety of pulsing tests were conducted over a wide range of electrical pulse widths (EPW). Each EPW condition was also tested over a range of percent duty cycles (DC), and bit impulse and pulsing specific impulse were determined for each of these conditions. Subsequent to acceptance testing at Aerojet, these three engines were delivered to the NASA White Sands Test Facility (WSTF) in April 2005 for incorporation into a cryogenic Auxiliary Propulsion System Test Bed (APSTB). The APSTB is a test article that will be utilized in an altitude test cell to simulate anticipated mission applications. The objectives of this APSTB testing included evaluation of engine performance over an extended duty cycle map of propellant pressure and temperature, as well as engine and system performance at typical mission duty cycles over extended periods of time. This paper provides acceptance test results and a status of the engine performance as part of the system level testing.

Determination of the Hall Thruster Operating Regimes

International Nuclear Information System (INIS)

L. Dorf; V. Semenov; Y. Raitses; N.J. Fisch

2002-04-01

A quasi one-dimensional (1-D) steady-state model of the Hall thruster is presented. For the same discharge voltage two operating regimes are possible -- with and without the anode sheath. For given mass flow rate, magnetic field profile and discharge voltage a unique solution can be constructed, assuming that the thruster operates in one of the regimes. However, we show that for a given temperature profile the applied discharge voltage uniquely determines the operating regime: for discharge voltages greater than a certain value, the sheath disappears. That result is obtained over a wide range of incoming neutral velocities, channel lengths and widths, and cathode plane locations. It is also shown that a good correlation between the quasi 1-D model and experimental results can be achieved by selecting an appropriate electron mobility and temperature profile

Rarefied gas electro jet (RGEJ) micro-thruster for space propulsion

International Nuclear Information System (INIS)

Blanco, Ariel; Roy, Subrata

2017-01-01

This article numerically investigates a micro-thruster for small satellites which utilizes plasma actuators to heat and accelerate the flow in a micro-channel with rarefied gas in the slip flow regime. The inlet plenum condition is considered at 1 Torr with flow discharging to near vacuum conditions (<0.05 Torr). The Knudsen numbers at the inlet and exit planes are ∼0.01 and ∼0.1, respectively. Although several studies have been performed in micro-hallow cathode discharges at constant pressure, to our knowledge, an integrated study of the glow discharge physics and resulting fluid flow of a plasma thruster under these low pressure and low Knudsen number conditions is yet to be reported. Numerical simulations of the charge distribution due to gas ionization processes and the resulting rarefied gas flow are performed using an in-house code. The mass flow rate, thrust, specific impulse, power consumption and the thrust effectiveness of the thruster are predicted based on these results. The ionized gas is modelled using local mean energy approximation. An electrically induced body force and a thermal heating source are calculated based on the space separated charge distribution and the ion Joule heating, respectively. The rarefied gas flow with these electric force and heating source is modelled using density-based compressible flow equations with slip flow boundary conditions. The results show that a significant improvement of specific impulse can be achieved over highly optimized cold gas thrusters using the same propellant. (paper)

High Fidelity Multi-Objective Design Optimization of a Downscaled Cusped Field Thruster

Directory of Open Access Journals (Sweden)

Thomas Fahey

2017-11-01

Full Text Available The Cusped Field Thruster (CFT concept has demonstrated significantly improved performance over the Hall Effect Thruster and the Gridded Ion Thruster; however, little is understood about the complexities of the interactions and interdependencies of the geometrical, magnetic and ion beam properties of the thruster. This study applies an advanced design methodology combining a modified power distribution calculation and evolutionary algorithms assisted by surrogate modeling to a multi-objective design optimization for the performance optimization and characterization of the CFT. Optimization is performed for maximization of performance defined by five design parameters (i.e., anode voltage, anode current, mass flow rate, and magnet radii, simultaneously aiming to maximize three objectives; that is, thrust, efficiency and specific impulse. Statistical methods based on global sensitivity analysis are employed to assess the optimization results in conjunction with surrogate models to identify key design factors with respect to the three design objectives and additional performance measures. The research indicates that the anode current and the Outer Magnet Radius have the greatest effect on the performance parameters. An optimal value for the anode current is determined, and a trend towards maximizing anode potential and mass flow rate is observed.

Performance of an iodine-fueled radio-frequency ion-thruster

Science.gov (United States)

Holste, Kristof; Gärtner, Waldemar; Zschätzsch, Daniel; Scharmann, Steffen; Köhler, Peter; Dietz, Patrick; Klar, Peter J.

2018-01-01

Two sets of performance data of the same radio-frequency ion-thruster (RIT) have been recorded using iodine and xenon, respectively, as propellant. To characterize the thruster's performance, we have recorded the radio-frequency DC-power, required for yielding preset values of the extracted ion-beam currents, as a function of mass flow. For that purpose, an iodine mass flow system had to be developed, calibrated, and integrated into a newly-built test facility for studying corrosive propellants. The performance mappings for iodine and xenon differ significantly despite comparable operation conditions. At low mass flows, iodine exhibits the better performance. The situation changes at higher mass flows where the performance of iodine is significantly poorer than that of xenon. The reason is very likely related to the molecular nature of iodine. Our results show that iodine as propellant is compatible with RIT technology. Furthermore, it is a viable alternative as propellant for dedicated space missions. In particular, when taking into account additional benefits such as possible storage as a solid and its low price the use of iodine as propellant in ion thrusters is competitive.

Overview of Iodine Propellant Hall Thruster Development Activities at NASA Glenn Research Center

Science.gov (United States)

Kamhawi, Hani; Benavides, Gabriel; Haag, Thomas; Hickman, Tyler; Smith, Timothy; Williams, George; Myers, James; Polzin, Kurt; Dankanich, John; Byrne, Larry;

Performance and Facility Background Pressure Characterization Tests of NASAs 12.5-kW Hall Effect Rocket with Magnetic Shielding Thruster

Science.gov (United States)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Shastry, Rohit; Thomas, Robert; Yim, John; Herman, Daniel; Williams, George; Myers, James; Hofer, Richard;

Femtosecond laser control of chemical reactions

CSIR Research Space (South Africa)

Du Plessis, A

2010-08-31

Full Text Available Femtosecond laser control of chemical reactions is made possible through the use of pulse-shaping techniques coupled to a learning algorithm feedback loop – teaching the laser pulse to control the chemical reaction. This can result in controllable...

Pickup ion processes associated with spacecraft thrusters: Implications for solar probe plus

Energy Technology Data Exchange (ETDEWEB)

Clemens, Adam, E-mail: a.j.clemens@qmul.ac.uk; Burgess, David [School of Physics and Astronomy, Queen Mary University of London, London (United Kingdom)

2016-03-15

Chemical thrusters are widely used in spacecraft for attitude control and orbital manoeuvres. They create an exhaust plume of neutral gas which produces ions via photoionization and charge exchange. Measurements of local plasma properties will be affected by perturbations caused by the coupling between the newborn ions and the plasma. A model of neutral expansion has been used in conjunction with a fully three-dimensional hybrid code to study the evolution and ionization over time of the neutral cloud produced by the firing of a mono-propellant hydrazine thruster as well as the interactions of the resulting ion cloud with the ambient solar wind. Results are presented which show that the plasma in the region near to the spacecraft will be perturbed for an extended period of time with the formation of an interaction region around the spacecraft, a moderate amplitude density bow wave bounding the interaction region and evidence of an instability at the forefront of the interaction region which causes clumps of ions to be ejected from the main ion cloud quasi-periodically.

A Plasmoid Thruster for Space Propulsion

Science.gov (United States)

Koelfgen, Syri J.; Hawk, Clark W.; Eskridge, Richard; Smith, James W.; Martin, Adam K.

2003-01-01

There are a number of possible advantages to using accelerated plasmoids for in-space propulsion. A plasmoid is a compact plasma structure with an integral magnetic field. They have been studied extensively in controlled fusion research and are classified according to the relative strength of the poloidal and toroidal magnetic field (B(sub p), and B(sub t), respectively). An object with B(sub p), / B(sub t) much greater than 1 is classified as a Field Reversed Configuration (FRC); if B(sub p) approximately equal to B(sub t), it is called a Spheromak. The plasmoid thruster operates by producing FRC-like plasmoids and subsequently ejecting them from the device at a high velocity. The plasmoid is formed inside of a single-turn conical theta-pinch coil. As this process is inductive, there are no electrodes. Similar experiments have yielded plasmoid velocities of at least 50 km/s, and calculations indicate that velocities in excess of 100 km/s should be possible. This concept should be capable of producing Isp's in the range of 5,000 - 15,000 s with thrust densities on the order of 10(exp 5) N per square meters. The current experiment is designed to produce jet powers in the range of 5 - 10 kW, although the concept should be scalable to several MW's. The plasmoid mass and velocity will be measured with a variety of diagnostics, including internal and external B-dot probes, flux loops, Langmuir probes, high-speed cameras and a laser interferometer. Also of key importance will be measurements of the efficiency and mass utilization. Simulations of the plasmoid thruster using MOQUI, a time-dependent MHD code, will be carried out concurrently with experimental testing.

Hollow Cathode Assembly Development for the HERMeS Hall Thruster

Science.gov (United States)

Sarver-Verhey, Timothy R.; Kamhawi, Hani; Goebel, Dan M.; Polk, James E.; Peterson, Peter Y.; Robinson, Dale A.

2016-01-01

To support the operation of the HERMeS 12.5 kW Hall Thruster for NASA's Asteroid Redirect Robotic Mission, hollow cathodes using emitters based on barium oxide impregnate and lanthanum hexaboride are being evaluated through wear-testing, performance characterization, plasma modeling, and review of integration requirements. This presentation will present the development approach used to assess the cathode emitter options. A 2,000-hour wear-test of development model Barium Oxide (BaO) hollow cathode is being performed as part of the development plan. Specifically this test is to identify potential impacts cathode emitter life during operation in the HERMeS thruster. The cathode was operated with a magnetic field-equipped anode that simulates the HERMeS hall thruster operating environment. Cathode discharge performance has been stable with the device accumulating 743 hours at the time of this report. Observed voltage changes are attributed to keeper surface condition changes during testing. Cathode behavior during characterization sweeps exhibited stable behavior, including cathode temperature. The details of the cathode assembly operation of the wear-test will be presented.

Nonlinear dynamics of mini-satellite respinup by weak internal controllable torques

Science.gov (United States)

Somov, Yevgeny

2014-12-01

Contemporary space engineering advanced new problem before theoretical mechanics and motion control theory: a spacecraft directed respinup by the weak restricted control internal forces. The paper presents some results on this problem, which is very actual for energy supply of information mini-satellites (for communication, geodesy, radio- and opto-electronic observation of the Earth et al.) with electro-reaction plasma thrusters and gyro moment cluster based on the reaction wheels or the control moment gyros. The solution achieved is based on the methods for synthesis of nonlinear robust control and on rigorous analytical proof for the required spacecraft rotation stability by Lyapunov function method. These results were verified by a computer simulation of strongly nonlinear oscillatory processes at respinuping of a flexible spacecraft.

Nonlinear dynamics of mini-satellite respinup by weak internal controllable torques

Energy Technology Data Exchange (ETDEWEB)

Somov, Yevgeny, E-mail: e-somov@mail.ru [Samara State Technical University, Department for Guidance, Navigation and Control, 244 Molodogvardeyskaya Str., Samara 443100 (Russian Federation)

2014-12-10

Contemporary space engineering advanced new problem before theoretical mechanics and motion control theory: a spacecraft directed respinup by the weak restricted control internal forces. The paper presents some results on this problem, which is very actual for energy supply of information mini-satellites (for communication, geodesy, radio- and opto-electronic observation of the Earth et al.) with electro-reaction plasma thrusters and gyro moment cluster based on the reaction wheels or the control moment gyros. The solution achieved is based on the methods for synthesis of nonlinear robust control and on rigorous analytical proof for the required spacecraft rotation stability by Lyapunov function method. These results were verified by a computer simulation of strongly nonlinear oscillatory processes at respinuping of a flexible spacecraft.

Dual Mode Low Power Hall Thruster, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Sample and return missions desire and missions like Saturn Observer require a low power Hall thruster that can operate at high thrust to power as well as high...

Advanced-technology 30-cm-diameter mercury ion thruster

Science.gov (United States)

Beattie, J. R.; Kami, S.

1982-01-01

An advanced-technology mercury ion thruster designed for operation at high thrust and high thrust-to-power ratio is described. The laboratory-model thruster employs a highly efficient discharge-chamber design that uses high-field-strength samarium-cobalt magnets arranged in a ring-cusp configuration. Ion extraction is achieved using an advanced three-grid ion-optics assembly which utilizes flexible mounts for supporting the screen, accel, and decel electrodes. Performance results are presented for operation at beam currents in the range from 1 to 5 A. The baseline specific discharge power is shown to be about 125 eV/ion, and the acceptable range of net-to-total accelerating-voltage ratio is shown to be in the range of 0.2-0.8 for beam currents in the range of 1-5 A.

Hall Thruster Modeling with a Given Temperature Profile

International Nuclear Information System (INIS)

Dorf, L.; Semenov, V.; Raitses, Y.; Fisch, N.J.

2002-01-01

A quasi one-dimensional steady-state model of the Hall thruster is presented. For given mass flow rate, magnetic field profile, and discharge voltage the unique solution can be constructed, assuming that the thruster operates in one of the two regimes: with or without the anode sheath. It is shown that for a given temperature profile, the applied discharge voltage uniquely determines the operating regime; for discharge voltages greater than a certain value, the sheath disappears. That result is obtained over a wide range of incoming neutral velocities, channel lengths and widths, and cathode plane locations. A good correlation between the quasi one-dimensional model and experimental results can be achieved by selecting an appropriate temperature profile. We also show how the presented model can be used to obtain a two-dimensional potential distribution

Particle-in-cell numerical simulations of a cylindrical Hall thruster with permanent magnets

Science.gov (United States)

Miranda, Rodrigo A.; Martins, Alexandre A.; Ferreira, José L.

2017-10-01

The cylindrical Hall thruster (CHT) is a propulsion device that offers high propellant utilization and performance at smaller dimensions and lower power levels than traditional Hall thrusters. In this paper we present first results of a numerical model of a CHT. This model solves particle and field dynamics self-consistently using a particle-in-cell approach. We describe a number of techniques applied to reduce the execution time of the numerical simulations. The specific impulse and thrust computed from our simulations are in agreement with laboratory experiments. This simplified model will allow for a detailed analysis of different thruster operational parameters and obtain an optimal configuration to be implemented at the Plasma Physics Laboratory at the University of Brasília.

Use of an ions thruster to dispose of type II long-lived fission products into outer space

International Nuclear Information System (INIS)

Takahashi, H.; Yu, A.

1997-01-01

To dispose of long-lived fission products (LLFPs) into outer space, an ions thruster can be used instead of a static accelerator. The specifications of the ions thrusters which are presently studies for space propulsion are presented, and their usability discussed. Using of a rocket with an ions thruster for disposing of the LLFPs directly into the sun required a larger amount of energy than does the use of an accelerator

Orbital and angular motion construction for low thrust interplanetary flight

Science.gov (United States)

Yelnikov, R. V.; Mashtakov, Y. V.; Ovchinnikov, M. Yu.; Tkachev, S. S.

2016-11-01

Low thrust interplanetary flight is considered. Firstly, the fuel-optimal control is found. Then the angular motion is synthesized. This motion provides the thruster tracking of the required by optimal control direction. And, finally, reaction wheel control law for tracking this angular motion is proposed and implemented. The numerical example is given and total operation time for thrusters is found. Disturbances from solar pressure, thrust eccentricity, inaccuracy of reaction wheels installation and errors of inertia tensor are taken into account.

STS-39: OMS Pod Thruster Removal/Replace

Science.gov (United States)

1991-01-01

Shown is the removal and replacement of the Discovery's orbital maneuvering systems (OMS) pod thruster. The OMS engine will be used to propel Discovery north, off of its previous orbital groundtrack, without changing the spacecraft's altitude. A burn with this lateral effect is known as "out-of-plane."

Attitude Control System Design for the Solar Dynamics Observatory

Science.gov (United States)

Starin, Scott R.; Bourkland, Kristin L.; Kuo-Chia, Liu; Mason, Paul A. C.; Vess, Melissa F.; Andrews, Stephen F.; Morgenstern, Wendy M.

2005-01-01

The Solar Dynamics Observatory mission, part of the Living With a Star program, will place a geosynchronous satellite in orbit to observe the Sun and relay data to a dedicated ground station at all times. SDO remains Sun- pointing throughout most of its mission for the instruments to take measurements of the Sun. The SDO attitude control system is a single-fault tolerant design. Its fully redundant attitude sensor complement includes 16 coarse Sun sensors, a digital Sun sensor, 3 two-axis inertial reference units, 2 star trackers, and 4 guide telescopes. Attitude actuation is performed using 4 reaction wheels and 8 thrusters, and a single main engine nominally provides velocity-change thrust. The attitude control software has five nominal control modes-3 wheel-based modes and 2 thruster-based modes. A wheel-based Safehold running in the attitude control electronics box improves the robustness of the system as a whole. All six modes are designed on the same basic proportional-integral-derivative attitude error structure, with more robust modes setting their integral gains to zero. The paper details the mode designs and their uses.

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

Science.gov (United States)

Beattie, J. R.

1983-01-01

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.

Space Charge Saturated Sheath Regime and Electron Temperature Saturation in Hall Thrusters

International Nuclear Information System (INIS)

Raitses, Y.; Staack, D.; Smirnov, A.; Fisch, N.J.

2005-01-01

Secondary electron emission in Hall thrusters is predicted to lead to space charge saturated wall sheaths resulting in enhanced power losses in the thruster channel. Analysis of experimentally obtained electron-wall collision frequency suggests that the electron temperature saturation, which occurs at high discharge voltages, appears to be caused by a decrease of the Joule heating rather than by the enhancement of the electron energy loss at the walls due to a strong secondary electron emission

Determination of the Hall Thruster Operating Regimes; TOPICAL

International Nuclear Information System (INIS)

L. Dorf; V. Semenov; Y. Raitses; N.J. Fisch

2002-01-01

A quasi one-dimensional (1-D) steady-state model of the Hall thruster is presented. For the same discharge voltage two operating regimes are possible - with and without the anode sheath. For given mass flow rate, magnetic field profile and discharge voltage a unique solution can be constructed, assuming that the thruster operates in one of the regimes. However, we show that for a given temperature profile the applied discharge voltage uniquely determines the operating regime: for discharge voltages greater than a certain value, the sheath disappears. That result is obtained over a wide range of incoming neutral velocities, channel lengths and widths, and cathode plane locations. It is also shown that a good correlation between the quasi 1-D model and experimental results can be achieved by selecting an appropriate electron mobility and temperature profile

The physics, performance and predictions of the PEGASES ion-ion thruster

Science.gov (United States)

Aanesland, Ane

2014-10-01

Electric propulsion (EP) is now used systematically in space applications (due to the fuel and lifetime economy) to the extent that EP is now recognized as the next generation space technology. The uses of EP systems have though been limited to attitude control of GEO-stationary satellites and scientific missions. Now, the community envisages the use of EP for a variety of other applications as well; such as orbit transfer maneuvers, satellites in low altitudes, space debris removal, cube-sat control, challenging scientific missions close to and far from earth etc. For this we need a platform of EP systems providing much more variety in performance than what classical Hall and Gridded thrusters can provide alone. PEGASES is a gridded thruster that can be an alternative for some new applications in space, in particular for space debris removal. Unlike classical ion thrusters, here positive and negative ions are alternately accelerated to produce thrust. In this presentation we will look at the fundamental aspects of PEGASES. The emphasis will be put on our current understanding, obtained via analytical models, PIC simulations and experimental measurements, of the alternate extraction and acceleration process. We show that at low grid bias frequencies (10 s of kHz), the system can be described as a sequence of negative and positive ions accelerated as packets within a classical DC mode. Here secondary electrons created in the downstream chamber play an important role in the beam space charge compensation. At higher frequencies (100 s of kHz) the transit time of the ions in the grid gap becomes comparable to the bias period, leading to an ``AC acceleration mode.'' Here the beam is fully space charge compensated and the ion energy and current are functions of the applied frequency and waveform. A generalization of the Child-Langmuir space charge limited law is developed for pulsed voltages and allows evaluating the optimal parameter space and performance of PEGASES

Experimental Investigation of the Near-Wall Region in the NASA HiVHAc EDU2 Hall Thruster

Science.gov (United States)

Shastry, Rohit; Kamhawi, Hani; Huang, Wensheng; Haag, Thomas W.

2015-01-01

The HiVHAc propulsion system is currently being developed to support Discovery-class NASA science missions. Presently, the thruster meets the required operational lifetime by utilizing a novel discharge channel replacement mechanism. As a risk reduction activity, an alternative approach is being investigated that modifies the existing magnetic circuit to shift the ion acceleration zone further downstream such that the magnetic components are not exposed to direct ion impingement during the thruster's lifetime while maintaining adequate thruster performance and stability. To measure the change in plasma properties between the original magnetic circuit configuration and the modified, "advanced" configuration, six Langmuir probes were flush-mounted within each channel wall near the thruster exit plane. Plasma potential and electron temperature were measured for both configurations across a wide range of discharge voltages and powers. Measurements indicate that the upstream edge of the acceleration zone shifted downstream by as much as 0.104 channel lengths, depending on operating condition. The upstream edge of the acceleration zone also appears to be more insensitive to operating condition in the advanced configuration, remaining between 0.136 and 0.178 channel lengths upstream of the thruster exit plane. Facility effects studies performed on the original configuration indicate that the plasma and acceleration zone recede further upstream into the channel with increasing facility pressure. These results will be used to inform further modifications to the magnetic circuit that will provide maximum protection of the magnetic components without significant changes to thruster performance and stability.

3D ion velocity distribution function measurement in an electric thruster using laser induced fluorescence tomography

Science.gov (United States)

Elias, P. Q.; Jarrige, J.; Cucchetti, E.; Cannat, F.; Packan, D.

2017-09-01

Measuring the full ion velocity distribution function (IVDF) by non-intrusive techniques can improve our understanding of the ionization processes and beam dynamics at work in electric thrusters. In this paper, a Laser-Induced Fluorescence (LIF) tomographic reconstruction technique is applied to the measurement of the IVDF in the plume of a miniature Hall effect thruster. A setup is developed to move the laser axis along two rotation axes around the measurement volume. The fluorescence spectra taken from different viewing angles are combined using a tomographic reconstruction algorithm to build the complete 3D (in phase space) time-averaged distribution function. For the first time, this technique is used in the plume of a miniature Hall effect thruster to measure the full distribution function of the xenon ions. Two examples of reconstructions are provided, in front of the thruster nose-cone and in front of the anode channel. The reconstruction reveals the features of the ion beam, in particular on the thruster axis where a toroidal distribution function is observed. These findings are consistent with the thruster shape and operation. This technique, which can be used with other LIF schemes, could be helpful in revealing the details of the ion production regions and the beam dynamics. Using a more powerful laser source, the current implementation of the technique could be improved to reduce the measurement time and also to reconstruct the temporal evolution of the distribution function.

2D Electrostatic Potential Solver for Hall Thruster Simulation

National Research Council Canada - National Science Library

Koo, Justin W

2006-01-01

...) for Hall thruster simulation. It is based on a finite volume discretization of a current conservation equation where the electron current density is described by a Generalized Ohm's law description...

Electromagnetic properties of a modular MHD thruster

Science.gov (United States)

Kom, C. H.; Brunet, Y.

1999-04-01

The magnetic field of an annular MHD thruster made of independent superconducting modules has been studied with analytical and numerical methods. This configuration allows to obtain large magnetized volumes and high induction levels with rapidly decreasing stray fields. When some inductors are out of order, the thruster remains still operational, but the stray fields increase in the vicinity of the failure. For given structural materials and superconductors, it is possible to determine the size of the conductor in order to reduce the electromagnetic forces and the peak field supported by the conductors. For an active field of 10 T in a 6 m ray annular active channel of a thruster with 24 modules, the peak field is exactly 15.6 T in the Nb3Sn conductors and the structure has to sustain 10^8 N/m forces. The necessity to place some magnetic or superconducting shield is discussed, particularly when the thruster is in a degraded regime. Nous présentons une étude analytique et numérique du champ magnétique d'un propulseur MHD naval annulaire, constitué de secteurs inducteurs supraconducteurs. Cette configuration nécessite des champs magnétiques élevés dans des volumes importants, et permet une décroissance rapide des champs de fuite. Lorsque quelques inducteurs sont en panne, le propulseur reste toujours opérationnel, mais les champs de fuite sont importants aux environs des modules hors service. Étant donné un matériau supraconducteur, il est possible de déterminer la forme des inducteurs dans le but de réduire à la fois les forces électromagnétiques et le surchamp supporté par le bobinage. Pour un propulseur annulaire constitué de 24 modules inducteurs, et un champ actif de 10 T au centre de la partie active du canal (r = 6 m) on obtient avec du Nb3Sn un champ maximun sur le conducteur de 15,5 T et la structure supporte une force de 10^8 N/m. De plus, la nécessité de placer des écrans magnétique ou supraconducteur en régime dégradé (mise

Electronegative Gas Thruster

Science.gov (United States)

Dankanich, John; Polzin, Kurt; Walker, Mitchell

2015-01-01

The project is an international collaboration and academic partnership to mature an innovative electric propulsion thruster concept to Technology Research Level-3 (TRL-3) through direct thrust measurement. The project includes application assessment of the technology ranging from small spacecraft to high power. The Plasma propulsion with Electronegative GASES(PEGASES) basic proof of concept has been matured to TRL-2 by Ane Aanesland of Laboratoire de Physique des Plasma at Ecole Polytechnique. The concept has advantages through eliminating the neutralizer requirement and should yield longer life and lower cost over conventional gridded ion engines. The objective of this research is to validate the proof of concept through the first direct thrust measurements and mature the concept to TRL-3.

Non-Maxwellian electron energy probability functions in the plume of a SPT-100 Hall thruster

Science.gov (United States)

Giono, G.; Gudmundsson, J. T.; Ivchenko, N.; Mazouffre, S.; Dannenmayer, K.; Loubère, D.; Popelier, L.; Merino, M.; Olentšenko, G.

2018-01-01

We present measurements of the electron density, the effective electron temperature, the plasma potential, and the electron energy probability function (EEPF) in the plume of a 1.5 kW-class SPT-100 Hall thruster, derived from cylindrical Langmuir probe measurements. The measurements were taken on the plume axis at distances between 550 and 1550 mm from the thruster exit plane, and at different angles from the plume axis at 550 mm for three operating points of the thruster, characterized by different discharge voltages and mass flow rates. The bulk of the electron population can be approximated as a Maxwellian distribution, but the measured distributions were seen to decline faster at higher energy. The measured EEPFs were best modelled with a general EEPF with an exponent α between 1.2 and 1.5, and their axial and angular characteristics were studied for the different operating points of the thruster. As a result, the exponent α from the fitted distribution was seen to be almost constant as a function of the axial distance along the plume, as well as across the angles. However, the exponent α was seen to be affected by the mass flow rate, suggesting a possible relationship with the collision rate, especially close to the thruster exit. The ratio of the specific heats, the γ factor, between the measured plasma parameters was found to be lower than the adiabatic value of 5/3 for each of the thruster settings, indicating the existence of non-trivial kinetic heat fluxes in the near collisionless plume. These results are intended to be used as input and/or testing properties for plume expansion models in further work.

Preventing Corrosion by Controlling Cathodic Reaction Kinetics

Science.gov (United States)

2016-03-25

3. DATES COVERED (From - To) 09/23/15 - 04/22/16 4. TITLE AND SUBTITLE Sa. CONTRACT NUMBER Preventing Corrosion by Controlling Cathodic Reaction...Preventing corrosion by controlling cathodic reaction kinetics Progress Report for Period: 1 SEP 2015-31 MAR 2016 John Keith Department of...25 March 2016 Preventing corrosion by controlling cathodic reaction kinetics Annual Summary Report: FY16 PI: John Keith, 412-624-7016,jakeith

A concept of ferroelectric microparticle propulsion thruster

International Nuclear Information System (INIS)

Yarmolich, D.; Vekselman, V.; Krasik, Ya. E.

2008-01-01

A space propulsion concept using charged ferroelectric microparticles as a propellant is suggested. The measured ferroelectric plasma source thrust, produced mainly by microparticles emission, reaches ∼9x10 -4 N. The obtained trajectories of microparticles demonstrate that the majority of the microparticles are positively charged, which permits further improvement of the thruster

Deposition of fluorocarbon films by Pulsed Plasma Thruster on the anode side

International Nuclear Information System (INIS)

Zhang, Rui; Zhang, Daixian; Zhang, Fan; He, Zhen; Wu, Jianjun

2013-01-01

Fluorocarbon thin films were deposited by Pulsed Plasma Thruster at different angles on the anode side of the thruster. Density and velocity of the plasma in the plume of the Pulsed Plasma Thruster were determined using double and triple Langmuir probe apparatus respectively. The deposited films were characterized by X-ray photoelectron spectroscopy (XPS), scanning probe microscope (SPM) and UV–vis spectrometer. Low F/C ratio (0.64–0.86) fluorocarbon films are deposited. The F/C ratio decreases with angle increasing from 0 degree to 30 degree; however it turns to increase with angle increasing from 45 degree to 90 degree. The films deposited at center angles appear rougher compared with that prepared at angles beyond 45 degree. These films basically show having strong absorption properties for wavelength below 600 nm and having enhanced reflective characteristics. Due to the influence of the chemical composition and the surface morphology of the films, the optical properties of these films also show significant angular dependence.

Magnesium Hall Thruster for Solar System Exploration, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — The innovation being developed in this program is a Mg Hall Effect Thruster system that would open the door for In-Situ Resource Utilization based solar system...

Development and characterization of high-efficiency, high-specific impulse xenon Hall thrusters

Science.gov (United States)

Hofer, Richard Robert

This dissertation presents research aimed at extending the efficient operation of 1600 s specific impulse Hall thruster technology to the 2000--3000 s range. While recent studies of commercially developed Hall thrusters demonstrated greater than 4000 s specific impulse, maximum efficiency occurred at less than 3000 s. It was hypothesized that the efficiency maximum resulted as a consequence of modern magnetic field designs, optimized for 1600 s, which were unsuitable at high-specific impulse. Motivated by the industry efforts and mission studies, the aim of this research was to develop and characterize xenon Hall thrusters capable of both high-specific impulse and high-efficiency operation. The research divided into development and characterization phases. During the development phase, the laboratory-model NASA-173M Hall thrusters were designed with plasma lens magnetic field topographies and their performance and plasma characteristics were evaluated. Experiments with the NASA-173M version 1 (v1) validated the plasma lens design by showing how changing the magnetic field topography at high-specific impulse improved efficiency. Experiments with the NASA-173M version 2 (v2) showed there was a minimum current density and optimum magnetic field topography at which efficiency monotonically increased with voltage. Between 300--1000 V, total specific impulse and total efficiency of the NASA-173Mv2 operating at 10 mg/s ranged from 1600--3400 s and 51--61%, respectively. Comparison of the thrusters showed that efficiency can be optimized for specific impulse by varying the plasma lens design. During the characterization phase, additional plasma properties of the NASA-173Mv2 were measured and a performance model was derived accounting for a multiply-charged, partially-ionized plasma. Results from the model based on experimental data showed how efficient operation at high-specific impulse was enabled through regulation of the electron current with the magnetic field. The

Discharge Oscillations in a Permanent Magnet Cylindrical Hall-Effect Thruster

Science.gov (United States)

Polzin, K. A.; Sooby, E. S.; Raitses, Y.; Merino, E.; Fisch, N. J.

2009-01-01

Measurements of the discharge current in a cylindrical Hall thruster are presented to quantify plasma oscillations and instabilities without introducing an intrusive probe into the plasma. The time-varying component of the discharge current is measured using a current monitor that possesses a wide frequency bandwidth and the signal is Fourier transformed to yield the frequency spectra present, allowing for the identification of plasma oscillations. The data show that the discharge current oscillations become generally greater in amplitude and complexity as the voltage is increased, and are reduced in severity with increasing flow rate. The breathing mode ionization instability is identified, with frequency as a function of discharge voltage not increasing with discharge voltage as has been observed in some traditional Hall thruster geometries, but instead following a scaling similar to a large-amplitude, nonlinear oscillation mode recently predicted in for annular Hall thrusters. A transition from lower amplitude oscillations to large relative fluctuations in the oscillating discharge current is observed at low flow rates and is suppressed as the mass flow rate is increased. A second set of peaks in the frequency spectra are observed at the highest propellant flow rate tested. Possible mechanisms that might give rise to these peaks include ionization instabilities and interactions between various oscillatory modes.

Enabling Ring-Cusp Ion Thruster Technology for NASA Missions

Data.gov (United States)

National Aeronautics and Space Administration — ESA is flying T6 Kaufman ion thrusters on the BepiColombo Mission to Mercury in 2018. They are planning to develop a longer life, higher performing, 30-cm ring-cusp...

High Input Voltage Hall Thruster Discharge Converter, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — The overall scope of this Phase I/II effort is the development of a high efficiency 15kW (nominal) Hall thruster discharge converter. In Phase I, Busek Co. Inc. will...

Experimental Investigations of a Krypton Stationary Plasma Thruster

Directory of Open Access Journals (Sweden)

A. I. Bugrova

2013-01-01

Full Text Available Stationary plasma thrusters are attractive electric propulsion systems for spacecrafts. The usual propellant is xenon. Among the other suggested propellants, krypton could be one of the best candidates. Most studies have been carried out with a Hall effect thruster previously designed for xenon. The ATON A-3 developed by MSTU MIREA (Moscow initially defined for xenon has been optimized for krypton. The stable high-performance ATON A-3 operation in Kr has been achieved after optimization of its magnetic field configuration and its optimization in different parameters: length and width of the channel, buffer volume dimensions, mode of the cathode operation, and input parameters. For a voltage of 400 V and the anode mass flow rate of 2.5 mg/s the anode efficiency reaches 60% and the specific impulse reaches 2900 s under A-3 operating with Kr. The achieved performances under operation A-3 with Kr are presented and compared with performances obtained with Xe.

(abstract) ARGOS: a System to Monitor Ulysses Nutation and Thruster Firings from Variations of the Spacecraft Radio Signal

Science.gov (United States)

McElrath, T. P.; Cangahuala, L. A.; Miller, K. J.; Stravert, L. R.; Garcia-Perez, Raul

1995-01-01

Ulysses is a spin-stabilized spacecraft that experienced significant nutation after its launch in October 1990. This was due to the Sun-spacecraft-Earth geometry, and a study of the phenomenon predicted that the nutation would again be a problem during 1994-95. The difficulty of obtaining nutation estimates in real time from the spacecraft telemetry forced the ESA/NASA Ulysses Team to explore alternative information sources. The work performed by the ESA Operations Team provided a model for a system that uses the radio signal strength measurements to monitor the spacecraft dynamics. These measurements (referred to as AGC) are provided once per second by the tracking stations of the DSN. The system was named ARGOS (Attitude Reckoning from Ground Observable Signals) after the ever-vigilant, hundred-eyed giant of Greek Mythology. The ARGOS design also included Doppler processing, because Doppler shifts indicate thruster firings commanded by the active nutation control carried out onboard the spacecraft. While there is some visibility into thruster activity from telemetry, careful processing of the high-sample-rate Doppler data provides an accurate means of detecting the presence and time of thruster firings. DSN Doppler measurements are available at a ten-per-second rate in the same tracking data block as the AGC data.

Effect of plasma distribution on propulsion performance in electrodeless plasma thrusters

Science.gov (United States)

Takao, Yoshinori; Takase, Kazuki; Takahashi, Kazunori

2016-09-01

A helicon plasma thruster consisting of a helicon plasma source and a magnetic nozzle is one of the candidates for long-lifetime thrusters because no electrodes are employed to generate or accelerate plasma. A recent experiment, however, detected the non-negligible axial momentum lost to the lateral wall boundary, which degrades thruster performance, when the source was operated with highly ionized gases. To investigate this mechanism, we have conducted two-dimensional axisymmetric particle-in-cell (PIC) simulations with the neutral distribution obtained by Direct Simulation Monte Carlo (DSMC) method. The numerical results have indicated that the axially asymmetric profiles of the plasma density and potential are obtained when the strong decay of neutrals occurs at the source downstream. This asymmetric potential profile leads to the accelerated ion towards the lateral wall, leading to the non-negligible net axial force in the opposite direction of the thrust. Hence, to reduce this asymmetric profile by increasing the neutral density at downstream and/or by confining plasma with external magnetic field would result in improvement of the propulsion performance. These effects are also analyzed by PIC/DSMC simulations.

Experimental Investigations from the Operation of a 2 Kw Brayton Power Conversion Unit and a Xenon Ion Thruster

Science.gov (United States)

Mason, Lee; Birchenough, Arthur; Pinero, Luis

2004-01-01

A 2 kW Brayton Power Conversion Unit (PCU) and a xenon ion thruster were integrated with a Power Management and Distribution (PMAD) system as part of a Nuclear Electric Propulsion (NEP) Testbed at NASA's Glenn Research Center. Brayton converters and ion thrusters are potential candidates for use on future high power NEP missions such as the proposed Jupiter Icy Moons Orbiter (JIMO). The use of existing lower power test hardware provided a cost-effective means to investigate the critical electrical interface between the power conversion system and ion propulsion system. The testing successfully demonstrated compatible electrical operations between the converter and the thruster, including end-to-end electric power throughput, high efficiency AC to DC conversion, and thruster recycle fault protection. The details of this demonstration are reported herein.

Study and Developement of Compact Permanent Magnet Hall Thrusters for Future Brazillian Space Missions

Science.gov (United States)

Ferreira, Jose Leonardo; Martins, Alexandre; Cerda, Rodrigo

2016-07-01

The Plasma Physics Laboratory of UnB has been developing a Permanent Magnet Hall Thruster (PHALL) for the UNIESPAÇO program, part of the Space Activities Program conducted by AEB- The Brazillian Space Agency since 2004. Electric propulsion is now a very successful method for primary and secondary propulsion systems. It is essential for several existing geostationary satellite station keeping systems and for deep space long duration solar system missions, where the thrusting system can be designed to be used on orbit transfer maneuvering and/or for satellite attitude control in long term space missions. Applications of compact versions of Permanent Magnet Hall Thrusters on future brazillian space missions are needed and foreseen for the coming years beginning with the use of small divergent cusp field (DCFH) Hall Thrusters type on CUBESATS ( 5-10 kg , 1W-5 W power consumption) and on Micro satellites ( 50- 100 kg, 10W-100W). Brazillian (AEB) and German (DLR) space agencies and research institutions are developing a new rocket dedicated to small satellite launching. The VLM- Microsatellite Launch Vehicle. The development of PHALL compact versions can also be important for the recently proposed SBG system, a future brazillian geostationary satellite system that is already been developed by an international consortium of brazillian and foreign space industries. The exploration of small bodies in the Solar System with spacecraft has been done by several countries with increasing frequency in these past twenty five years. Since their historical beginning on the sixties, most of the Solar System missions were based on gravity assisted trajectories very much depended on planet orbit positioning relative to the Sun and the Earth. The consequence was always the narrowing of the mission launch window. Today, the need for Solar System icy bodies in situ exploration requires less dependence on gravity assisted maneuvering and new high precision low thrust navigation methods

Long Life Cold Cathodes for Hall effect Thrusters, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — An electron source incorporating long life, high current density cold cathodes inside a microchannel plate for use with ion thrusters is proposed. Cathode lifetime...

Microfluidic Array of Externally Fed Electrospray Thrusters for Micro-Propulsion

Data.gov (United States)

National Aeronautics and Space Administration — The goal of this proposal is to design an electrospray micropropulsion thruster that utilizes a novel propellant transport mechanism. This project is a collaboration...

Modeling of the near field plume of a Hall thruster

International Nuclear Information System (INIS)

Boyd, Iain D.; Yim, John T.

2004-01-01

In this study, a detailed numerical model is developed to simulate the xenon plasma near-field plume from a Hall thruster. The model uses a detailed fluid model to describe the electrons and a particle-based kinetic approach is used to model the heavy xenon ions and atoms. The detailed model is applied to compute the near field plume of a small, 200 W Hall thruster. Results from the detailed model are compared with the standard modeling approach that employs the Boltzmann model. The usefulness of the model detailed is assessed through direct comparisons with a number of different measured data sets. The comparisons illustrate that the detailed model accurately predicts a number of features of the measured data not captured by the simpler Boltzmann approach

Hybrid-PIC Computer Simulation of the Plasma and Erosion Processes in Hall Thrusters

Science.gov (United States)

Hofer, Richard R.; Katz, Ira; Mikellides, Ioannis G.; Gamero-Castano, Manuel

2010-01-01

HPHall software simulates and tracks the time-dependent evolution of the plasma and erosion processes in the discharge chamber and near-field plume of Hall thrusters. HPHall is an axisymmetric solver that employs a hybrid fluid/particle-in-cell (Hybrid-PIC) numerical approach. HPHall, originally developed by MIT in 1998, was upgraded to HPHall-2 by the Polytechnic University of Madrid in 2006. The Jet Propulsion Laboratory has continued the development of HPHall-2 through upgrades to the physical models employed in the code, and the addition of entirely new ones. Primary among these are the inclusion of a three-region electron mobility model that more accurately depicts the cross-field electron transport, and the development of an erosion sub-model that allows for the tracking of the erosion of the discharge chamber wall. The code is being developed to provide NASA science missions with a predictive tool of Hall thruster performance and lifetime that can be used to validate Hall thrusters for missions.

Real time control of the flexible dynamics of orbital launch vehicles

NARCIS (Netherlands)

Bos, van den J.; Steinbuch, M.; Gutierrez, H.M.

2011-01-01

During this traineeship the flexible dynamics of orbital launch vehicles are estimated and controlled in real time, using distributed fiber-Bragg sensor arrays for motion estimation and cold gas thrusters for control. The use of these cold-gas thrusters to actively control flexible modes is the main

High-Power Krypton Hall Thruster Technology Being Developed for Nuclear-Powered Applications

Science.gov (United States)

Jacobson, David T.; Manzella, David H.

2004-01-01

The NASA Glenn Research Center has been performing research and development of moderate specific impulse, xenon-fueled, high-power Hall thrusters for potential solar electric propulsion applications. These applications include Mars missions, reusable tugs for low-Earth-orbit to geosynchronous-Earth-orbit transportation, and missions that require transportation to libration points. This research and development effort resulted in the design and fabrication of the NASA-457M Hall thruster that has been tested at input powers up to 95 kW. During project year 2003, NASA established Project Prometheus to develop technology in the areas of nuclear power and propulsion, which are enabling for deep-space science missions. One of the Project-Prometheus-sponsored Nuclear Propulsion Research tasks is to investigate alternate propellants for high-power Hall thruster electric propulsion. The motivation for alternate propellants includes the disadvantageous cost and availability of xenon propellant for extremely large scale, xenon-fueled propulsion systems and the potential system performance benefits of using alternate propellants. The alternate propellant krypton was investigated because of its low cost relative to xenon. Krypton propellant also has potential performance benefits for deep-space missions because the theoretical specific impulse for a given voltage is 20 percent higher than for xenon because of krypton's lower molecular weight. During project year 2003, the performance of the high-power NASA-457M Hall thruster was measured using krypton as the propellant at power levels ranging from 6.4 to 72.5 kW. The thrust produced ranged from 0.3 to 2.5 N at a discharge specific impulse up to 4500 sec.

Laser-Induced Fluorescence Measurements within a Laboratory Hall Thruster (Postprint)

National Research Council Canada - National Science Library

Hargus, Jr., W. A; Cappelli, M. A

1999-01-01

In this paper, we describe the results of a study of laser induced fluorescence velocimetry of ionic xenon in the plume and interior acceleration channel of a laboratory Hall type thruster operating...

Investigation of excited states populations density of Hall thruster plasma in three dimensions by laser-induced fluorescence spectroscopy

Science.gov (United States)

Krivoruchko, D. D.; Skrylev, A. V.

2018-01-01

The article deals with investigation of the excited states populations distribution of a low-temperature xenon plasma in the thruster with closed electron drift at 300 W operating conditions were investigated by laser-induced fluorescence (LIF) over the 350-1100 nm range. Seven xenon ions (Xe II) transitions were analyzed, while for neutral atoms (Xe I) just three transitions were explored, since the majority of Xe I emission falls into the ultraviolet or infrared part of the spectrum and are difficult to measure. The necessary spontaneous emission probabilities (Einstein coefficients) were calculated. Measurements of the excited state distribution were made for points (volume of about 12 mm3) all over the plane perpendicular to thruster axis in four positions on it (5, 10, 50 and 100 mm). Measured LIF signal intensity have differences for each location of researched point (due to anisotropy of thruster plume), however the structure of states populations distribution persisted at plume and is violated at the thruster exit plane and cathode area. Measured distributions show that for describing plasma of Hall thruster one needs to use a multilevel kinetic model, classic model can be used just for far plume region or for specific electron transitions.

Non-Toxic Orbiter Maneuvering System (OMS) and Reaction Control System

Science.gov (United States)

Hurlbert, Eric A.; Nicholson, Leonard S. (Technical Monitor)

1999-01-01

NASA is pursuing the technology and advanced development of a non-toxic (NT) orbital maneuvering system (OMS) and reaction control system (RCS) for shuttle upgrades, RLV, and reusable first stages. The primary objectives of the shuttle upgrades program are improved safety, improved reliability, reduced operations time and cost, improved performance or capabilities, and commonality with future space exploration needs. Non-Toxic OMS/RCS offers advantages in each of these categories. A non-toxic OMS/RCS eliminates the ground hazards and the flight safety hazards of the toxic and corrosive propellants. The cost savings for ground operations are over $24M per year for 7 flights, and the savings increase with increasing flight rate up to $44M per year. The OMS/RCS serial processing time is reduced from 65 days to 13 days. The payload capability can be increased up to 5100 Ibms. The non-toxic OMS/RCS also provides improved space station reboost capability up to 20 nautical miles over the current toxic system of 14 nautical miles. A NT OMS/RCS represents a clear advancement in the SOA over MMH/NTO. Liquid oxygen and ethanol are clean burning, high-density propellants that provide a high degree of commonality with other spacecraft subsystems including life support, power, and thermal control, and with future human exploration and development of space missions. The simple and reliable pressure-fed design uses sub-cooled liquid oxygen at 250 to 350 psia, which allows a propellant to remain cryogenic for longer periods of time. The key technologies are thermal insulation and conditioning techniques are used to maintain the sub-cooling. Phase I successfully defined the system architecture, designed an integrated OMS/RCS propellant tank, analyzed the feed system, built and tested the 870 lbf RCS thrusters, and tested the 6000 lbf OMS engine. Phase 11 is currently being planned for the development and test of full-scale prototype of the system in 1999 and 2000

Near-Surface Plasma Characterization of the 12.5-kW NASA TDU1 Hall Thruster

Science.gov (United States)

Shastry, Rohit; Huang, Wensheng; Kamhawi, Hani

2015-01-01

To advance the state-of-the-art in Hall thruster technology, NASA is developing a 12.5-kW, high-specific-impulse, high-throughput thruster for the Solar Electric Propulsion Technology Demonstration Mission. In order to meet the demanding lifetime requirements of potential missions such as the Asteroid Redirect Robotic Mission, magnetic shielding was incorporated into the thruster design. Two units of the resulting thruster, called the Hall Effect Rocket with Magnetic Shielding (HERMeS), were fabricated and are presently being characterized. The first of these units, designated the Technology Development Unit 1 (TDU1), has undergone extensive performance and thermal characterization at NASA Glenn Research Center. A preliminary lifetime assessment was conducted by characterizing the degree of magnetic shielding within the thruster. This characterization was accomplished by placing eight flush-mounted Langmuir probes within each discharge channel wall and measuring the local plasma potential and electron temperature at various axial locations. Measured properties indicate a high degree of magnetic shielding across the throttle table, with plasma potential variations along each channel wall being less than or equal to 5 eV and electron temperatures being maintained at less than or equal to 5 eV, even at 800 V discharge voltage near the thruster exit plane. These properties indicate that ion impact energies within the HERMeS will not exceed 26 eV, which is below the expected sputtering threshold energy for boron nitride. Parametric studies that varied the facility backpressure and magnetic field strength at 300 V, 9.4 kW, illustrate that the plasma potential and electron temperature are insensitive to these parameters, with shielding being maintained at facility pressures 3X higher and magnetic field strengths 2.5X higher than nominal conditions. Overall, the preliminary lifetime assessment indicates a high degree of shielding within the HERMeS TDU1, effectively

Long Life Miniature Hall Thruster Enabling Low Cost Human Precursor Missions

Data.gov (United States)

National Aeronautics and Space Administration — Key and Central Objectives: This investigation aims to demonstrate that the application of magnetic shielding technology on miniature Hall thrusters will...

Experimental Investigation from the Operation of a 2 kW Brayton Power Conversion Unit and a Xenon Ion Thruster

Science.gov (United States)

Hervol, David; Mason, Lee; Birchenough, Art; Pinero, Luis

2004-01-01

A 2kW Brayton Power Conversion Unit (PCU) and a xenon ion thruster were integrated with a Power Management and Distribution (PMAD) system as part of a Nuclear Electric Propulsion (NEP) Testbed at NASA's Glenn Research Center. Brayton Converters and ion thrusters are potential candidates for use on future high power NEP mission such as the proposed Jupiter Icy Moons Orbiter (JIMO). The use of a existing lower power test hardware provided a cost effective means to investigate the critical electrical interface between the power conversion system and the propulsion system. The testing successfully demonstrated compatible electrical operations between the converter and the thruster, including end-to-end electric power throughput, high efficiency AC to DC conversion, and thruster recycle fault protection. The details of this demonstration are reported herein.

Hot-Fire Testing of 5N and 22N HPGP Thrusters

Science.gov (United States)

Burnside, Christopher G.; Pedersen, Kevin W.; Pierce, Charles W.

2015-01-01

This hot-fire test continues NASA investigation of green propellant technologies for future missions. To show the potential for green propellants to replace some hydrazine systems in future spacecraft, NASA Marshall Space Flight Center (MSFC) is continuing to embark on hot-fire test campaigns with various green propellant blends.NASA completed hot-fire testing of 5N and 22N HPGP thrusters at the Marshall Space Flight Center’s Component Development Area altitude test stand in April 2015. Both thrusters are ground test articles and not flight ready units, but are representative of potential flight hardware with a known path towards flight application. The purpose of the 5N testing was to perform facility check-outs and generate a small set of data for comparison to ECAPS and Orbital ATK data sets. The 5N thruster performed as expected with thrust and propellant flow-rate data generated that are similar to previous testing at Orbital ATK. Immediately following the 5N testing, and using the same facility, the 22N testing was conducted on the same test stand with the purpose of demonstrating the 22N performance. The results of 22N testing indicate it performed as expected.The results of the hot-fire testing are presented in this paper and presentation.

Transit-time instability in Hall thrusters

International Nuclear Information System (INIS)

Barral, Serge; Makowski, Karol; Peradzynski, Zbigniew; Dudeck, Michel

2005-01-01

Longitudinal waves characterized by a phase velocity of the order of the velocity of ions have been recurrently observed in Hall thruster experiments and simulations. The origin of this so-called ion transit-time instability is investigated with a simple one-dimensional fluid model of a Hall thruster discharge in which cold ions are accelerated between two electrodes within a quasineutral plasma. A short-wave asymptotics applied to linearized equations shows that plasma perturbations in such a device consist of quasineutral ion acoustic waves superimposed on a background standing wave generated by discharge current oscillations. Under adequate circumstances and, in particular, at high ionization levels, acoustic waves are amplified as they propagate, inducing strong perturbation of the ion density and velocity. Responding to the subsequent perturbation of the column resistivity, the discharge current generates a standing wave, the reflection of which sustains the generation of acoustic waves at the inlet boundary. A calculation of the frequency and growth rate of this resonance mechanism for a supersonic ion flow is proposed, which illustrates the influence of the ionization degree on their onset and the approximate scaling of the frequency with the ion transit time. Consistent with experimental reports, the traveling wave can be observed on plasma density and velocity perturbations, while the plasma potential ostensibly oscillates in phase along the discharge

Engineering Model Propellant Feed System Development for an Iodine Hall Thruster Demonstration Mission

Science.gov (United States)

Polzin, Kurt A.

2016-01-01

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cu cm and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high (Delta)v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ?Isp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (engineering model propellant feed system for iSAT (see Fig. 1). The feed system is based around an iodine propellant reservoir and two proportional control valves (PFCVs) that meter the iodine flow to the cathode and anode. The flow is split upstream of the PFCVs to both components can be fed from a common reservoir. Testing of the reservoir is reported to demonstrate that the design is capable of delivering the required propellant flow rates to operate the thruster. The tubing and reservoir are fabricated from hastelloy to resist corrosion by the heated gaseous iodine propellant. The reservoir, tubing, and PFCVs are heated to ensure the sublimed propellant will not re

NASA Brief: Q-Thruster Physics

Science.gov (United States)

White, Harold

2013-01-01

Q-thrusters are a low-TRL form of electric propulsion that operates on the principle of pushing off of the quantum vacuum. A terrestrial analog to this is to consider how a submarine uses its propeller to push a column of water in one direction, while the sub recoils in the other to conserve momentum -the submarine does not carry a "tank" of sea water to be used as propellant. In our case, we use the tools of Magnetohydrodynamics (MHD) to show how the thruster pushes off of the quantum vacuum which can be thought of as a sea of virtual particles -principally electrons and positrons that pop into and out of existence, and where fields are stronger, there are more virtual particles. The idea of pushing off the quantum vacuum has been in the technical literature for a few decades, but to date, the obstacle has been the magnitude of the predicted thrust which has been derived analytically to be very small, and therefore not likely to be useful for human spaceflight. Our recent theoretical model development and test data suggests that we can greatly increase the magnitude of the negative pressure of the quantum vacuum and generate a specific force such that technology based on this approach can be competitive for in-space propulsion approx. 0.1N/kW), and possibly for terrestrial applications (approx. 10N/kW). As an additional validation of the approach, the theory allows calculation of physics constants from first principles: Gravitational constant, Planck constant, Bohr radius, dark energy fraction, electron mass.

Real-Tme Boron Nitride Erosion Measurements of the HiVHAc Thruster via Cavity Ring-Down Spectroscopy

Science.gov (United States)

Lee, Brian C.; Yalin, Azer P.; Gallimore, Alec; Huang, Wensheng; Kamhawi, Hani

2013-01-01

Cavity ring-down spectroscopy was used to make real-time erosion measurements from the NASA High Voltage Hall Accelerator thruster. The optical sensor uses 250 nm light to measure absorption of atomic boron in the plume of an operating Hall thruster. Theerosion rate of the High Voltage Hall Accelerator thruster was measured for discharge voltages ranging from 330 to 600 V and discharge powers ranging from 1 to 3 kW. Boron densities as high as 6.5 x 10(exp 15) per cubic meter were found within the channel. Using a very simple boronvelocity model, approximate volumetric erosion rates between 5.0 x 10(exp -12) and 8.2 x 10(exp -12) cubic meter per second were found.

Development and control of a three-axis satellite simulator for the bifocal relay mirror initiative

OpenAIRE

Chernesky, Vincent S.

2001-01-01

The Three Axis Satellite Simulator (TASS) is a 4-foot diameter octagonal platform supported on a spherical air bearing. The platform hosts several satellite subsystems, including rate gyros, reaction wheels, thrusters, sun sensors, and an onboard control computer. This free-floating design allows for realistic emulation of satellite attitude dynamics in a laboratory environment. The bifocal relay mirror spacecraft system is composed of two optically coupled telescopes used to redirect the las...

Simulations of a Plasma Thruster Utilizing the FRC Configuration

Energy Technology Data Exchange (ETDEWEB)

Rognlien, T. D. [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States); Cohen, B. I. [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States)

2016-10-10

This report describes work performed by LLNL to model the behavior and performance of a reverse-field configuration (FRC) type of plasma device as a plasma thruster as summarized by Razin et al. [1], which also describes the MNX device at PPPL used to study this concept.

A cellular automata approach to chemical reactions : 1 reaction controlled systems

NARCIS (Netherlands)

Korte, de A.C.J.; Brouwers, H.J.H.

2013-01-01

A direct link between the chemical reaction controlled (shrinking core) model and cellular automata, to study the dissolution of particles, is derived in this paper. Previous research on first and second order reactions is based on the concentration of the reactant. The present paper describes the

Development of the Multiple Use Plug Hybrid for Nanosats (MUPHyN) miniature thruster

Science.gov (United States)

Eilers, Shannon

The Multiple Use Plug Hybrid for Nanosats (MUPHyN) prototype thruster incorporates solutions to several major challenges that have traditionally limited the deployment of chemical propulsion systems on small spacecraft. The MUPHyN thruster offers several features that are uniquely suited for small satellite applications. These features include 1) a non-explosive ignition system, 2) non-mechanical thrust vectoring using secondary fluid injection on an aerospike nozzle cooled with the oxidizer flow, 3) a non-toxic, chemically-stable combination of liquid and inert solid propellants, 4) a compact form factor enabled by the direct digital manufacture of the inert solid fuel grain. Hybrid rocket motors provide significant safety and reliability advantages over both solid composite and liquid propulsion systems; however, hybrid motors have found only limited use on operational vehicles due to 1) difficulty in modeling the fuel flow rate 2) poor volumetric efficiency and/or form factor 3) significantly lower fuel flow rates than solid rocket motors 4) difficulty in obtaining high combustion efficiencies. The features of the MUPHyN thruster are designed to offset and/or overcome these shortcomings. The MUPHyN motor design represents a convergence of technologies, including hybrid rocket regression rate modeling, aerospike secondary injection thrust vectoring, multiphase injector modeling, non-pyrotechnic ignition, and nitrous oxide regenerative cooling that address the traditional challenges that limit the use of hybrid rocket motors and aerospike nozzles. This synthesis of technologies is unique to the MUPHyN thruster design and no comparable work has been published in the open literature.

Carbon Nanotube Based Electric Propulsion Thruster with Low Power Consumption, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — Field emission electric propulsion (FEEP) thrusters have gained considerable attention for spacecrafts disturbance compensation because of excellent characteristics....

Optimisation of a quantum pair space thruster

Directory of Open Access Journals (Sweden)

Valeriu DRAGAN

2012-06-01

Full Text Available The paper addresses the problem of propulsion for long term space missions. Traditionally a space propulsion unit has a propellant mass which is ejected trough a nozzle to generate thrust; this is also the case with inert gases energized by an on-board power unit. Unconventional methods for propulsion include high energy LASERs that rely on the momentum of photons to generate thrust. Anti-matter has also been proposed for energy storage. Although the momentum of ejected gas is significantly higher, the LASER propulsion offers the perspective of unlimited operational time – provided there is a power source. The paper will propose the use of the quantum pair formation for generating a working mass, this is different than conventional anti-matter thrusters since the material particles generated are used as propellant not as energy storage.Two methods will be compared: LASER and positron-electron, quantum pair formation. The latter will be shown to offer better momentum above certain energy levels.For the demonstrations an analytical solution is obtained and provided in the form of various coefficients. The implications are, for now, theoretical however the practicality of an optimized thruster using such particles is not to be neglected for long term space missions.

Laser-Driven Mini-Thrusters

International Nuclear Information System (INIS)

Sterling, Enrique; Lin Jun; Sinko, John; Kodgis, Lisa; Porter, Simon; Pakhomov, Andrew V.; Larson, C. William; Mead, Franklin B. Jr.

2006-01-01

Laser-driven mini-thrusters were studied using Delrin registered and PVC (Delrin registered is a registered trademark of DuPont) as propellants. TEA CO2 laser (λ = 10.6 μm) was used as a driving laser. Coupling coefficients were deduced from two independent techniques: force-time curves measured with a piezoelectric sensor and ballistic pendulum. Time-resolved ICCD images of the expanding plasma and combustion products were analyzed in order to determine the main process that generates the thrust. The measurements were also performed in a nitrogen atmosphere in order to test the combustion effects on thrust. A pinhole transmission experiment was performed for the study of the cut-off time when the ablation/air breakdown plasma becomes opaque to the incoming laser pulse

Laser-Driven Mini-Thrusters

Science.gov (United States)

Sterling, Enrique; Lin, Jun; Sinko, John; Kodgis, Lisa; Porter, Simon; Pakhomov, Andrew V.; Larson, C. William; Mead, Franklin B.

2006-05-01

Laser-driven mini-thrusters were studied using Delrin® and PVC (Delrin® is a registered trademark of DuPont) as propellants. TEA CO2 laser (λ = 10.6 μm) was used as a driving laser. Coupling coefficients were deduced from two independent techniques: force-time curves measured with a piezoelectric sensor and ballistic pendulum. Time-resolved ICCD images of the expanding plasma and combustion products were analyzed in order to determine the main process that generates the thrust. The measurements were also performed in a nitrogen atmosphere in order to test the combustion effects on thrust. A pinhole transmission experiment was performed for the study of the cut-off time when the ablation/air breakdown plasma becomes opaque to the incoming laser pulse.

Effects of magnetic field strength in the discharge channel on the performance of a multi-cusped field thruster

Directory of Open Access Journals (Sweden)

Peng Hu

2016-09-01

Full Text Available The performance characteristics of a Multi-cusped Field Thruster depending on the magnetic field strength in the discharge channel were investigated. Four thrusters with different outer diameters of the magnet rings were designed to change the magnetic field strength in the discharge channel. It is found that increasing the magnetic field strength could restrain the radial cross-field electron current and decrease the radial width of main ionization region, which gives rise to the reduction of propellant utilization and thruster performance. The test results in different anode voltage conditions indicate that both the thrust and anode efficiency are higher for the weaker magnetic field in the discharge channel.

Silicon Carbide (SiC) Power Processing Unit (PPU) for Hall Effect Thrusters

Science.gov (United States)

Reese, Bradley

2015-01-01

Arkansas Power Electronics International (APEI), Inc., is developing a high-efficiency, radiation-hardened 3.8-kW SiC power supply for the PPU of Hall effect thrusters. This project specifically targets the design of a PPU for the high-voltage Hall accelerator (HiVHAC) thruster, with target specifications of 80- to 160-V input, 200- to 700-V/5A output, efficiency greater than 96 percent, and peak power density in excess of 2.5 kW/kg. The PPU under development uses SiC junction field-effect transistor power switches, components that APEI, Inc., has irradiated under total ionizing dose conditions to greater than 3 MRad with little to no change in device performance.

Plume Characterization of a Laboratory Model 22 N GPIM Thruster via High-Frequency Raman Spectroscopy

Science.gov (United States)

Williams, George J.; Kojima, Jun J.; Arrington, Lynn A.; Deans, Matthew C.; Reed, Brian D.; Kinzbach, McKenzie I.; McLean, Christopher H.

2015-01-01

The Green Propellant Infusion Mission (GPIM) will demonstrate the capability of a green propulsion system, specifically, one using the monopropellant, AF-M315E. One of the risks identified for GPIM is potential contamination of sensitive areas of the spacecraft from the effluents in the plumes of AF-M315E thrusters. Plume characterization of a laboratory-model 22 N thruster via optical diagnostics was conducted at NASA GRC in a space-simulated environment. A high-frequency pulsed laser was coupled with an electron-multiplied ICCD camera to perform Raman spectroscopy in the near-field, low-pressure plume. The Raman data yielded plume constituents and temperatures over a range of thruster chamber pressures and as a function of thruster (catalyst) operating time. Schlieren images of the near-field plume enabled calculation of plume velocities and revealed general plume structure of the otherwise invisible plume. The measured velocities are compared to those predicted by a two-dimensional, kinetic model. Trends in data and numerical results are presented from catalyst mid-life to end-of-life. The results of this investigation were coupled with the Raman and Schlieren data to provide an anchor for plume impingement analysis presented in a companion paper. The results of both analyses will be used to improve understanding of the nature of AF-M315E plumes and their impacts to GPIM and other future missions.

Effects of thruster firings on the shuttle's plasma and electric field environment

International Nuclear Information System (INIS)

Machuzak, J.S.; Burke, W.J.; Retterer, J.M.; Hunton, D.E.; Jasperse, J.R.; Smiddy, M.

1993-01-01

Simultaneous plasma and AC/DC electric field measurements taken during the space shuttle mission STS-4 at times of prolonged thruster firings are analyzed and cross correlated. Depending on the orientation of the shuttle's velocity vector to the magnetic field, ion densities and electric field wave spectra were enhanced or decreased. The systematic picture of interactions within the shuttle's plasma/neutral gas environment of Cairns and Gurnett (1991b) is confirmed and extended. Waves are excited by outgassed and thruster-ejected molecules that ionize in close proximity to the shuttle. On time scales significantly less than an ion gyroperiod, the newly created ions act as beams in the background plasma. These beams are sources of VLF waves that propagate near the shuttle and intensify during thruster firings. Plasma density depletions and/or the shuttle's geometry may hinder wave detection in the payload bay. A modified two-stream analysis indicates that beam components propagating at large angles to the magnetic field are unstable to the growth of lower hybrid waves. The beam-excited, lower hybrid waves heat some electrons to sufficient energies to produce impact ionization. Empirical evidence for other wave-growth mechanisms outside the lower-hybrid band is presented. 42 refs., 15 figs., 3 tabs

Investigation of a subsonic-arc-attachment thruster using segmented anodes

Science.gov (United States)

Berns, Darren H.; Sankovic, John M.; Sarmiento, Charles J.

1993-01-01

To investigate high frequency arc instabilities observed in subsonic-arc-attachment thrusters, a 3 kW, segmented-anode arcjet was designed and tested using hydrogen as the propellant. The thruster nozzle geometry was scaled from a 30 kW design previously tested in the 1960's. By observing the current to each segment and the arc voltage, it was determined that the 75-200 kHz instabilities were results of axial movements of the arc anode attachment point. The arc attachment point was fully contained in the subsonic portion of the nozzle for nearly all flow rates. The effects of isolating selected segments were investigated. In some cases, forcing the arc downstream caused the restrike to cease. Finally, decreasing the background pressure from 18 Pa to 0.05 Pa affected the pressure distribution in the nozzle, including the pressure in the subsonic arc chamber.

Investigation of a subsonic-arc-attachment thruster using segmented anodes

Science.gov (United States)

Berns, Darren H.; Sankovic, John M.; Sarmiento, Charles J.

1993-01-01

To investigate high frequency arc instabilities observed in subsonic-arc-attachment thrusters, a 3 kW, segmented-anode arc jet was designed and tested using hydrogen as the propellant. The thruster nozzle geometry was scaled from a 30 kW design previously tested in the 1960's. By observing the current to each segment and the arc voltage, it was determined that the 75-200 kHz instabilities were results of axial movements of the arc anode attachment point. The arc attachment point was fully contained in the subsonic portion of the nozzle for nearly all flow rates. The effects of isolating selected segments were investigated. In some cases, forcing the arc downstream caused the restrike to cease. Finally, decreasing the background pressure from 18 to 0.05 Pa affected the pressure distribution in the nozzle including the pressure in the subsonic arc chamber.

HIGH ENERGY REPLACEMENT FOR TEFLON PROPELLANT IN PULSED PLASMA THRUSTERS, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — This program will utilize a well-characterized Pulsed Plasma Thruster (PPT) to test experimental high-energy extinguishable solid propellants (HE), instead of...

Clearance of short circuited ion optics electrodes by capacitive discharge. [in ion thrusters

Science.gov (United States)

Poeschel, R. L.

1976-01-01

The ion optics electrodes of low specific impulse (3000 sec) mercury electron bombardment ion thrusters are vulnerable to short circuits by virtue of their relatively small interelectrode spacing (0.5 mm). Metallic flakes from backsputtered deposits are the most probable cause of such 'shorts' and 'typical' flakes have been simulated here using refractory wire that has a representative, but controllable, cross section. Shorting wires can be removed by capacitive discharge without significant damage to the electrodes. This paper describes an evaluation of 'short' removal versus electrode damage for several combinations of capacitor voltage, stored energy, and short circuit conditions.

Vacuum Chamber Construction and Contamination Study of A Micro Pulsed Plasma Thruster

National Research Council Canada - National Science Library

Debevec, Jacob H

2006-01-01

.... This study examines the deposition profile and rate of particle emission from the thruster so that satellite designers understand any potential contamination issues with sensitive instruments and solar panels...

Measurement of erosion rate by absorption spectroscopy in a Hall thruster

International Nuclear Information System (INIS)

Yamamoto, Naoji; Yokota, Shigeru; Matsui, Makoto; Komurasaki, Kimiya; Arakawa, Yoshihiro

2005-01-01

The erosion rate of a Hall thruster was estimated with the objective of building a real-time erosion rate monitoring system using a 1 kW class anode layer type Hall thruster. This system aids the understanding of the tradeoff between lifetime and performance. To estimate the flux of the sputtered wall material, the number density of the sputtered iron was measured by laser absorption spectroscopy using an absorption line from ground atomic iron at 371.9935 nm. An ultravioletAl x In y Ga (1-x-y) N diode laser was used as the probe. The estimated number density of iron was 1.1x10 16 m -3 , which is reasonable when compared with that measured by duration erosion tests. The relation between estimated erosion rate and magnetic flux density also agreed with that measured by duration erosion tests

Feasibility of a 5mN Laser-Driven Mini-Thruster, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — We have developed a next-generation thruster under a Phase II SBIR which we believe can meet NASA requirements after some modifications and improvements. It is the...

On the Application of Hall Thruster Working with Ambient Atmospheric Gas for Orbital Station-Keeping

Directory of Open Access Journals (Sweden)

D. V. Duhopel'nikov

2016-01-01

Full Text Available The paper considers the application of the Hall thruster using the ambient atmospheric air for orbital station keeping. This is a relevant direction at the up-to-date development stage of propulsion systems. Many teams of designers of electric rocket thrusters evaluate the application of different schemes of particle acceleration at the low-earth orbit. Such technical solution allows us to abandon the storage systems of the working agent on the spacecraft board. Thus, lifetime of such a system at the orbit wouldn`t be limited by fuel range. The paper suggests a scheme of the propulsion device with a parabolic confuser that provides a required compression ratio of the ambient air for correct operation. Formulates physical and structural restrictions on ambient air to be used as a working agent of the thruster. Pointes out that the altitudes from 200 to 300 km are the most promising for such propulsion devices. Shows that for operation at lower altitudes are required the higher capacities that are not provided by modern onboard power supply systems. For the orbit heightening the air intakes with significant compression rate are of necessity. The size of such air intakes would exceed nose fairing of exploited space launch systems. To perform further design calculations are shown dependencies that allow us to calculate an effective diameter of the thruster channel and a critical voltage to be desirable for thrust force excess over air resistance. The dependencies to calculate minimal and maximal fluxes of neutral particles of oxygen and nitrogen, that are necessary for normal thruster operation, are also shown. Calculation results of the propulsion system parameters for the spacecrafts with cross-sectional area within 1 - 3 m2 and inlet diameter of air intake within 1 - 3 m are demonstrated. The research results have practical significance in design of advanced propulsion devices for lowaltitude spacecrafts. The work has been supported by the RFFR

A centre-triggered magnesium fuelled cathodic arc thruster uses sublimation to deliver a record high specific impulse

Science.gov (United States)

Neumann, Patrick R. C.; Bilek, Marcela; McKenzie, David R.

2016-08-01

The cathodic arc is a high current, low voltage discharge that operates in vacuum and provides a stream of highly ionised plasma from a solid conducting cathode. The high ion velocities, together with the high ionisation fraction and the quasineutrality of the exhaust stream, make the cathodic arc an attractive plasma source for spacecraft propulsion applications. The specific impulse of the cathodic arc thruster is substantially increased when the emission of neutral species is reduced. Here, we demonstrate a reduction of neutral emission by exploiting sublimation in cathode spots and enhanced ionisation of the plasma in short, high-current pulses. This, combined with the enhanced directionality due to the efficient erosion profiles created by centre-triggering, substantially increases the specific impulse. We present experimentally measured specific impulses and jet power efficiencies for titanium and magnesium fuels. Our Mg fuelled source provides the highest reported specific impulse for a gridless ion thruster and is competitive with all flight rated ion thrusters. We present a model based on cathode sublimation and melting at the cathodic arc spot explaining the outstanding performance of the Mg fuelled source. A further significant advantage of an Mg-fuelled thruster is the abundance of Mg in asteroidal material and in space junk, providing an opportunity for utilising these resources in space.

System analysis and test-bed for an atmosphere-breathing electric propulsion system using an inductive plasma thruster

Science.gov (United States)

Romano, F.; Massuti-Ballester, B.; Binder, T.; Herdrich, G.; Fasoulas, S.; Schönherr, T.

2018-06-01

Challenging space mission scenarios include those in low altitude orbits, where the atmosphere creates significant drag to the S/C and forces their orbit to an early decay. For drag compensation, propulsion systems are needed, requiring propellant to be carried on-board. An atmosphere-breathing electric propulsion system (ABEP) ingests the residual atmosphere particles through an intake and uses them as propellant for an electric thruster. Theoretically applicable to any planet with atmosphere, the system might allow to orbit for unlimited time without carrying propellant. A new range of altitudes for continuous operation would become accessible, enabling new scientific missions while reducing costs. Preliminary studies have shown that the collectible propellant flow for an ion thruster (in LEO) might not be enough, and that electrode erosion due to aggressive gases, such as atomic oxygen, will limit the thruster lifetime. In this paper an inductive plasma thruster (IPT) is considered for the ABEP system. The starting point is a small scale inductively heated plasma generator IPG6-S. These devices are electrodeless and have already shown high electric-to-thermal coupling efficiencies using O2 and CO2 . The system analysis is integrated with IPG6-S tests to assess mean mass-specific energies of the plasma plume and estimate exhaust velocities.

Overview of the Development of the Solar Electric Propulsion Technology Demonstration Mission 12.5-kW Hall Thruster

Science.gov (United States)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Yim, John; Chang, Li; Clayman, Lauren; Herman, Daniel; Shastry, Rohit; Thomas, Robert; Verhey, Timothy;

A New Method for Analyzing Near-Field Faraday Probe Data in Hall Thrusters

Science.gov (United States)

Huang, Wensheng; Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Kamhawi, Hani

2013-01-01

This paper presents a new method for analyzing near-field Faraday probe data obtained from Hall thrusters. Traditional methods spawned from far-field Faraday probe analysis rely on assumptions that are not applicable to near-field Faraday probe data. In particular, arbitrary choices for the point of origin and limits of integration have made interpretation of the results difficult. The new method, called iterative pathfinding, uses the evolution of the near-field plume with distance to provide feedback for determining the location of the point of origin. Although still susceptible to the choice of integration limits, this method presents a systematic approach to determining the origin point for calculating the divergence angle. The iterative pathfinding method is applied to near-field Faraday probe data taken in a previous study from the NASA-300M and NASA-457Mv2 Hall thrusters. Since these two thrusters use centrally mounted cathodes the current density associated with the cathode plume is removed before applying iterative pathfinding. A procedure is presented for removing the cathode plume. The results of the analysis are compared to far-field probe analysis results. This paper ends with checks on the validity of the new method and discussions on the implications of the results.

Engineering Risk Assessment of Space Thruster Challenge Problem

Science.gov (United States)

Mathias, Donovan L.; Mattenberger, Christopher J.; Go, Susie

2014-01-01

The Engineering Risk Assessment (ERA) team at NASA Ames Research Center utilizes dynamic models with linked physics-of-failure analyses to produce quantitative risk assessments of space exploration missions. This paper applies the ERA approach to the baseline and extended versions of the PSAM Space Thruster Challenge Problem, which investigates mission risk for a deep space ion propulsion system with time-varying thruster requirements and operations schedules. The dynamic mission is modeled using a combination of discrete and continuous-time reliability elements within the commercially available GoldSim software. Loss-of-mission (LOM) probability results are generated via Monte Carlo sampling performed by the integrated model. Model convergence studies are presented to illustrate the sensitivity of integrated LOM results to the number of Monte Carlo trials. A deterministic risk model was also built for the three baseline and extended missions using the Ames Reliability Tool (ART), and results are compared to the simulation results to evaluate the relative importance of mission dynamics. The ART model did a reasonable job of matching the simulation models for the baseline case, while a hybrid approach using offline dynamic models was required for the extended missions. This study highlighted that state-of-the-art techniques can adequately adapt to a range of dynamic problems.

Ultra-Compact Center-Mounted Hollow Cathodes for Hall Effect Thrusters, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — The proposed innovation is a long lifetime, compact hollow cathode that can be mounted along the axis of a 600 W-class Hall effect thruster. Testing at kilowatt...

Three Dimensional Simulation of Ion Thruster Plume-Spacecraft Interaction Based on a Graphic Processor Unit

International Nuclear Information System (INIS)

Ren Junxue; Xie Kan; Qiu Qian; Tang Haibin; Li Juan; Tian Huabing

2013-01-01

Based on the three-dimensional particle-in-cell (PIC) method and Compute Unified Device Architecture (CUDA), a parallel particle simulation code combined with a graphic processor unit (GPU) has been developed for the simulation of charge-exchange (CEX) xenon ions in the plume of an ion thruster. Using the proposed technique, the potential and CEX plasma distribution are calculated for the ion thruster plume surrounding the DS1 spacecraft at different thrust levels. The simulation results are in good agreement with measured CEX ion parameters reported in literature, and the GPU's results are equal to a CPU's. Compared with a single CPU Intel Core 2 E6300, 16-processor GPU NVIDIA GeForce 9400 GT indicates a speedup factor of 3.6 when the total macro particle number is 1.1×10 6 . The simulation results also reveal how the back flow CEX plasma affects the spacecraft floating potential, which indicates that the plume of the ion thruster is indeed able to alleviate the extreme negative floating potentials of spacecraft in geosynchronous orbit

EADS-ST's Latest Bipropellant 10N Thruster and 400 N Engine: The Fully European Solution

Science.gov (United States)

Fick, M.; Dreer, T.; Gotzig, U.; Schulte, G.; Bachmann, J.; Lagier, F.; Benoit, E.

2004-10-01

Increasing restrictions, complications and bureaucratic hurdles for obtaining export licenses from the US government for American components to be used on certain projects and for certain launch sites or end customers, required the development of new or upgraded European flow control valves to guarantee an independent and unrestricted marketing of EADS-ST's orbital propulsion products to commercial customers worldwide. The development and qualification of the European flow control valves for EADS-ST's 10 N bipropellant thruster and 400 N bipropellant engine is highlighted, together with verification tests and the qualification programs at engine level. The 400 N engine under qualification with the new valves is an enhanced version with an increased area ratio of the nozzle.

An Effective Approach Control Scheme for the Tethered Space Robot System

Directory of Open Access Journals (Sweden)

Zhongjie Meng

2014-09-01

Full Text Available The tethered space robot system (TSR, which is composed of a platform, a gripper and a space tether, has great potential in future space missions. Given the relative motion among the platform, tether, gripper and the target, an integrated approach model is derived. Then, a novel coordinated approach control scheme is presented, in which the tether tension, thrusters and the reaction wheel are all utilized. It contains the open-loop trajectory optimization, the feedback trajectory control and attitude control. The numerical simulation results show that the rendezvous between TSR and the target can be realized by the proposed coordinated control scheme, and the propellant consumption is efficiently reduced. Moreover, the control scheme performs well in the presence of the initial state's perturbations, actuator characteristics and sensor errors.

Characterization of Aerodynamic Interactions with the Mars Science Laboratory Reaction Control System Using Computation and Experiment

Science.gov (United States)

Schoenenberger, Mark; VanNorman, John; Rhode, Matthew; Paulson, John

2013-01-01

On August 5 , 2012, the Mars Science Laboratory (MSL) entry capsule successfully entered Mars' atmosphere and landed the Curiosity rover in Gale Crater. The capsule used a reaction control system (RCS) consisting of four pairs of hydrazine thrusters to fly a guided entry. The RCS provided bank control to fly along a flight path commanded by an onboard computer and also damped unwanted rates due to atmospheric disturbances and any dynamic instabilities of the capsule. A preliminary assessment of the MSL's flight data from entry showed that the capsule flew much as predicted. This paper will describe how the MSL aerodynamics team used engineering analyses, computational codes and wind tunnel testing in concert to develop the RCS system and certify it for flight. Over the course of MSL's development, the RCS configuration underwent a number of design iterations to accommodate mechanical constraints, aeroheating concerns and excessive aero/RCS interactions. A brief overview of the MSL RCS configuration design evolution is provided. Then, a brief description is presented of how the computational predictions of RCS jet interactions were validated. The primary work to certify that the RCS interactions were acceptable for flight was centered on validating computational predictions at hypersonic speeds. A comparison of computational fluid dynamics (CFD) predictions to wind tunnel force and moment data gathered in the NASA Langley 31-Inch Mach 10 Tunnel was the lynch pin to validating the CFD codes used to predict aero/RCS interactions. Using the CFD predictions and experimental data, an interaction model was developed for Monte Carlo analyses using 6-degree-of-freedom trajectory simulation. The interaction model used in the flight simulation is presented.

Effects of the Phoenix Lander descent thruster plume on the Martian surface

Science.gov (United States)

Plemmons, D. H.; Mehta, M.; Clark, B. C.; Kounaves, S. P.; Peach, L. L.; Renno, N. O.; Tamppari, L.; Young, S. M. M.

2008-08-01

The exhaust plume of Phoenix's hydrazine monopropellant pulsed descent thrusters will impact the surface of Mars during its descent and landing phase in the northern polar region. Experimental and computational studies have been performed to characterize the chemical compounds in the thruster exhausts. No undecomposed hydrazine is observed above the instrument detection limit of 0.2%. Forty-five percent ammonia is measured in the exhaust at steady state. Water vapor is observed at a level of 0.25%, consistent with fuel purity analysis results. Moreover, the dynamic interactions of the thruster plumes with the ground have been studied. Large pressure overshoots are produced at the ground during the ramp-up and ramp-down phases of the duty cycle of Phoenix's pulsed engines. These pressure overshoots are superimposed on the 10 Hz quasi-steady ground pressure perturbations with amplitude of about 5 kPa (at touchdown altitude) and have a maximum amplitude of about 20-40 kPa. A theoretical explanation for the physics that causes these pressure perturbations is briefly described in this article. The potential for soil erosion and uplifting at the landing site is also discussed. The objectives of the research described in this article are to provide empirical and theoretical data for the Phoenix Science Team to mitigate any potential problem. The data will also be used to ensure proper interpretation of the results from on-board scientific instrumentation when Martian soil samples are analyzed.

Chemical potential and reaction electronic flux in symmetry controlled reactions.

Science.gov (United States)

Vogt-Geisse, Stefan; Toro-Labbé, Alejandro

2016-07-15

In symmetry controlled reactions, orbital degeneracies among orbitals of different symmetries can occur along a reaction coordinate. In such case Koopmans' theorem and the finite difference approximation provide a chemical potential profile with nondifferentiable points. This results in an ill-defined reaction electronic flux (REF) profile, since it is defined as the derivative of the chemical potential with respect to the reaction coordinate. To overcome this deficiency, we propose a new way for the calculation of the chemical potential based on a many orbital approach, suitable for reactions in which symmetry is preserved. This new approach gives rise to a new descriptor: symmetry adapted chemical potential (SA-CP), which is the chemical potential corresponding to a given irreducible representation of a symmetry group. A corresponding symmetry adapted reaction electronic flux (SA-REF) is also obtained. Using this approach smooth chemical potential profiles and well defined REFs are achieved. An application of SA-CP and SA-REF is presented by studying the Cs enol-keto tautomerization of thioformic acid. Two SA-REFs are obtained, JA'(ξ) and JA'' (ξ). It is found that the tautomerization proceeds via an in-plane delocalized 3-center 4-electron O-H-S hypervalent bond which is predicted to exist only in the transition state (TS) region. © 2016 Wiley Periodicals, Inc. © 2016 Wiley Periodicals, Inc.

East–West GEO Satellite Station-Keeping with Degraded Thruster Response

Directory of Open Access Journals (Sweden)

Stoian Borissov

2015-09-01

Full Text Available The higher harmonic terms of Earth’s gravitational potential slowly modify the nominal longitude of geostationary Earth orbit (GEO satellites, while the third-body presence (Moon and Sun mainly affects their latitude. For this reason, GEO satellites periodically need to perform station-keeping maneuvers, namely, east–west and north–south maneuvers to compensate for longitudinal and latitudinal variations, respectively. During the operational lifetime of GEO satellites, the thrusters’ response when commanded to perform these maneuvers slowly departs from the original nominal impulsive behavior. This paper addresses the practical problem of how to perform reliable east–west station-keeping maneuvers when thruster response is degraded. The need for contingency intervention from ground-based satellite operators is reduced by breaking apart the scheduled automatic station-keeping maneuvers into smaller maneuvers. Orbital alignment and attitude are tracked on-board during and in between sub-maneuvers, and any off nominal variations are corrected for with subsequent maneuvers. These corrections are particularly important near the end of the lifetime of GEO satellites, where thruster response is farthest from nominal performance.

Comparison study of exhaust plume impingement effects of small mono- and bipropellant thrusters using parallelized DSMC method.

Directory of Open Access Journals (Sweden)

Kyun Ho Lee

Full Text Available A space propulsion system is important for the normal mission operations of a spacecraft by adjusting its attitude and maneuver. Generally, a mono- and a bipropellant thruster have been mainly used for low thrust liquid rocket engines. But as the plume gas expelled from these small thrusters diffuses freely in a vacuum space along all directions, unwanted effects due to the plume collision onto the spacecraft surfaces can dramatically cause a deterioration of the function and performance of a spacecraft. Thus, aim of the present study is to investigate and compare the major differences of the plume gas impingement effects quantitatively between the small mono- and bipropellant thrusters using the computational fluid dynamics (CFD. For an efficiency of the numerical calculations, the whole calculation domain is divided into two different flow regimes depending on the flow characteristics, and then Navier-Stokes equations and parallelized Direct Simulation Monte Carlo (DSMC method are adopted for each flow regime. From the present analysis, thermal and mass influences of the plume gas impingements on the spacecraft were analyzed for the mono- and the bipropellant thrusters. As a result, it is concluded that a careful understanding on the plume impingement effects depending on the chemical characteristics of different propellants are necessary for the efficient design of the spacecraft.

Human Outer Solar System Exploration via Q-Thruster Technology

Science.gov (United States)

Joosten, B. Kent; White, Harold G.

2014-01-01

Propulsion technology development efforts at the NASA Johnson Space Center continue to advance the understanding of the quantum vacuum plasma thruster (QThruster), a form of electric propulsion. Through the use of electric and magnetic fields, a Q-thruster pushes quantum particles (electrons/positrons) in one direction, while the Qthruster recoils to conserve momentum. This principle is similar to how a submarine uses its propeller to push water in one direction, while the submarine recoils to conserve momentum. Based on laboratory results, it appears that continuous specific thrust levels of 0.4 - 4.0 N/kWe are achievable with essentially no onboard propellant consumption. To evaluate the potential of this technology, a mission analysis tool was developed utilizing the Generalized Reduced Gradient non-linear parameter optimization engine contained in the Microsoft Excel® platform. This tool allowed very rapid assessments of "Q-Ship" minimum time transfers from earth to the outer planets and back utilizing parametric variations in thrust acceleration while enforcing constraints on planetary phase angles and minimum heliocentric distances. A conservative Q-Thruster specific thrust assumption (0.4 N/kWe) combined with "moderate" levels of space nuclear power (1 - 2 MWe) and vehicle specific mass (45 - 55 kg/kWe) results in continuous milli-g thrust acceleration, opening up realms of human spaceflight performance completely unattainable by any current systems or near-term proposed technologies. Minimum flight times to Mars are predicted to be as low as 75 days, but perhaps more importantly new "retro-phase" and "gravity-augmented" trajectory shaping techniques were revealed which overcome adverse planetary phasing and allow virtually unrestricted departure and return opportunities. Even more impressively, the Jovian and Saturnian systems would be opened up to human exploration with round-trip times of 21 and 32 months respectively including 6 to 12 months of

Design, fabrication and testing of porous tungsten vaporizers for mercury ion thrusters

Science.gov (United States)

Zavesky, R.; Kroeger, E.; Kami, S.

1983-01-01

The dispersions in the characteristics, performance and reliability of vaporizers for early model 30-cm thrusters were investigated. The purpose of the paper is to explore the findings and to discuss the approaches that were taken to reduce the observed dispersion and present the results of a program which validated those approaches. The information that is presented includes porous tungsten materials specifications, a discussion of assembly procedures, and a description of a test program which screens both material and fabrication processes. There are five appendices providing additional detail in the areas of vaporizer contamination, nitrogen flow testing, bubble testing, porosimeter testing, and mercury purity. Four neutralizers, seven cathodes and five main vaporizers were successfully fabricated, tested, and operated on thrusters. Performance data from those devices is presented and indicates extremely repeatable results from using the design and fabrication procedures.

Control of Maillard Reactions in Foods: Strategies and Chemical Mechanisms.

Science.gov (United States)

Lund, Marianne N; Ray, Colin A

2017-06-14

Maillard reactions lead to changes in food color, organoleptic properties, protein functionality, and protein digestibility. Numerous different strategies for controlling Maillard reactions in foods have been attempted during the past decades. In this paper, recent advances in strategies for controlling the Maillard reaction and subsequent downstream reaction products in food systems are critically reviewed. The underlying mechanisms at play are presented, strengths and weaknesses of each strategy are discussed, and reasonable reaction mechanisms are proposed to reinforce the evaluations. The review includes strategies involving addition of functional ingredients, such as plant polyphenols and vitamins, as well as enzymes. The resulting trapping or modification of Maillard targets, reactive intermediates, and advanced glycation endproducts (AGEs) are presented with their potential unwanted side effects. Finally, recent advances in processing for control of Maillard reactions are discussed.

Experimental investigation of the effects of variable expanding channel on the performance of a low-power cusped field thruster

Directory of Open Access Journals (Sweden)

Hui Liu

2018-04-01

Full Text Available Due to a special magnetic field structure, the multi-cusped field thruster shows advantages of low wall erosion, low noise and high thrust density over a wide range of thrust. In this paper, expanding discharge channels are employed to make up for deficiencies on the range of thrust and plume divergence, which often emerges in conventional straight cylindrical channels. Three thruster geometries are fabricated with different expanding-angle channels, and a group of experiments are carried out to find out their influence on the performance and discharge characteristics of the thruster. A retarding potential analyzer and a Faraday probe are employed to analyze the structures of the plume in these three models. The results show that when the thrusters operate at low mass flow rate, the gradually-expanding channels exhibit lower propellant utilization and lower overall performance by amounts not exceeding 44.8% in ionization rate and 19.5% in anode efficiency, respectively. But the weakening of magnetic field intensity near the exit of expanding channels leads to an extended thrust throttling ability, a smaller plume divergence angle, and a relatively larger stable operating space without mode converting and the consequent performance degradation.

Plasma-Sheath Instability in Hall Thrusters Due to Periodic Modulation of the Energy of Secondary Electrons in Cyclotron Motion

International Nuclear Information System (INIS)

Sydorenko, D.; Smolyakov, A.; Kaganovich, I.; Raitses, Y.

2008-01-01

Particle-in-cell simulation of Hall thruster plasmas reveals a plasma-sheath instability manifesting itself as a rearrangement of the plasma sheath near the thruster channel walls accompanied by a sudden change of many discharge parameters. The instability develops when the sheath current as a function of the sheath voltage is in the negative conductivity regime. The major part of the sheath current is produced by beams of secondary electrons counter-streaming between the walls. The negative conductivity is the result of nonlinear dependence of beam-induced secondary electron emission on the plasma potential. The intensity of such emission is defined by the beam energy. The energy of the beam in crossed axial electric and radial magnetic fields is a quasi-periodical function of the phase of cyclotron rotation, which depends on the radial profile of the potential and the thruster channel width. There is a discrete set of stability intervals determined by the final phase of the cyclotron rotation of secondary electrons. As a result, a small variation of the thruster channel width may result in abrupt changes of plasma parameters if the plasma state jumps from one stability interval to another

Communication: Control of chemical reactions using electric field gradients

Energy Technology Data Exchange (ETDEWEB)

Deshmukh, Shivaraj D.; Tsori, Yoav, E-mail: tsori@bgu.ac.il [Department of Chemical Engineering, Ben-Gurion University of the Negev, Beer-Sheva 84105 (Israel)

2016-05-21

We examine theoretically a new idea for spatial and temporal control of chemical reactions. When chemical reactions take place in a mixture of solvents, an external electric field can alter the local mixture composition, thereby accelerating or decelerating the rate of reaction. The spatial distribution of electric field strength can be non-trivial and depends on the arrangement of the electrodes producing it. In the absence of electric field, the mixture is homogeneous and the reaction takes place uniformly in the reactor volume. When an electric field is applied, the solvents separate and the reactants are concentrated in the same phase or separate to different phases, depending on their relative miscibility in the solvents, and this can have a large effect on the kinetics of the reaction. This method could provide an alternative way to control runaway reactions and to increase the reaction rate without using catalysts.

Communication: Control of chemical reactions using electric field gradients.

Science.gov (United States)

Deshmukh, Shivaraj D; Tsori, Yoav

2016-05-21

We examine theoretically a new idea for spatial and temporal control of chemical reactions. When chemical reactions take place in a mixture of solvents, an external electric field can alter the local mixture composition, thereby accelerating or decelerating the rate of reaction. The spatial distribution of electric field strength can be non-trivial and depends on the arrangement of the electrodes producing it. In the absence of electric field, the mixture is homogeneous and the reaction takes place uniformly in the reactor volume. When an electric field is applied, the solvents separate and the reactants are concentrated in the same phase or separate to different phases, depending on their relative miscibility in the solvents, and this can have a large effect on the kinetics of the reaction. This method could provide an alternative way to control runaway reactions and to increase the reaction rate without using catalysts.

Internal plasma potential measurements of a Hall thruster using xenon and krypton propellant

International Nuclear Information System (INIS)

Linnell, Jesse A.; Gallimore, Alec D.

2006-01-01

For krypton to become a realistic option for Hall thruster operation, it is necessary to understand the performance gap between xenon and krypton and what can be done to reduce it. A floating emissive probe is used with the Plasmadynamics and Electric Propulsion Laboratory's High-speed Axial Reciprocating Probe system to map the internal plasma potential structure of the NASA-173Mv1 Hall thruster [R. R. Hofer, R. S. Jankovsky, and A. D. Gallimore, J. Propulsion Power 22, 721 (2006); and ibid.22, 732 (2006)] using xenon and krypton propellant. Measurements are taken for both propellants at discharge voltages of 500 and 600 V. Electron temperatures and electric fields are also reported. The acceleration zone and equipotential lines are found to be strongly linked to the magnetic-field lines. The electrostatic plasma lens of the NASA-173Mv1 Hall thruster strongly focuses the xenon ions toward the center of the discharge channel, whereas the krypton ions are defocused. Krypton is also found to have a longer acceleration zone than the xenon cases. These results explain the large beam divergence observed with krypton operation. Krypton and xenon have similar maximum electron temperatures and similar lengths of the high electron temperature zone, although the high electron temperature zone is located farther downstream in the krypton case

An axially propagating two-stream instability in the Hall thruster plasma

Czech Academy of Sciences Publication Activity Database

Tsikata, S.; Cavalier, Jordan; Héron, A.; Honore, C.; Lemoine, N.; Gresillon, D.; Coulette, D.

2014-01-01

Roč. 21, č. 7 (2014), 072116-072116 ISSN 1070-664X Institutional support: RVO:61389021 Keywords : Collective Thomson scattering * Hall thruster * kinetic theory * electrostatic modes Subject RIV: BL - Plasma and Gas Discharge Physics Impact factor: 2.142, year: 2014 http://dx.doi.org/10.1063/1.4890025

Hall Thruster Thermal Modeling and Test Data Correlation

Science.gov (United States)

Myers, James

2016-01-01

HERMeS - Hall Effect Rocket with Magnetic Shielding. Developed through a joint effort by NASA/GRC and the Jet Propulsion Laboratory (JPL). Design goals: High power (12.5 kW) high Isp (3000 sec), high efficiency (> 60%), high throughput (10,000 kg), reduced plasma erosion and increased life (5 yrs) to support Asteroid Redirect Robotic Mission (ARRM). Further details see "Performance, Facility Pressure Effects and Stability Characterization Tests of NASAs HERMeS Thruster" by H. Kamhawi and team. Hall Thrusters (HT) inherently operate at elevated temperatures approx. 600 C (or more). Due to electric magnetic (E x B) fields used to ionize and accelerate propellant gas particles (i.e., plasma). Cooling is largely limited to radiation in vacuum environment.Thus the hardware components must withstand large start-up delta-T's. HT's are constructed of multiple materials; assorted metals, non-metals and ceramics for their required electrical and magnetic properties. To mitigate thermal stresses HT design must accommodate the differential thermal growth from a wide range of material Coef. of Thermal Expansion (CTEs). Prohibiting the use of some bolted/torqued interfaces.Commonly use spring loaded interfaces, particularly at the metal-to-ceramic interfaces to allow for slippage.However most component interfaces must also effectively conduct heat to the external surfaces for dissipation by radiation.Thus contact pressure and area are important.

Plasma Perturbations in High-Speed Probing of Hall Thruster Discharge Chambers: Quantification and Mitigation

Science.gov (United States)

Jorns, Benjamin A.; Goebel, Dan M.; Hofer, Richard R.

2015-01-01

An experimental investigation is presented to quantify the effect of high-speed probing on the plasma parameters inside the discharge chamber of a 6-kW Hall thruster. Understanding the nature of these perturbations is of significant interest given the importance of accurate plasma measurements for characterizing thruster operation. An array of diagnostics including a high-speed camera and embedded wall probes is employed to examine in real time the changes in electron temperature and plasma potential induced by inserting a high-speed reciprocating Langmuir probe into the discharge chamber. It is found that the perturbations onset when the scanning probe is downstream of the electron temperature peak, and that along channel centerline, the perturbations are best characterized as a downstream shift of plasma parameters by 15-20% the length of the discharge chamber. A parametric study is performed to investigate techniques to mitigate the observed probe perturbations including varying probe speed, probe location, and operating conditions. It is found that the perturbations largely disappear when the thruster is operated at low power and low discharge voltage. The results of this mitigation study are discussed in the context of recommended methods for generating unperturbed measurements of the discharge chamber plasma.

The direct wave-drive thruster

Science.gov (United States)

Feldman, Matthew Solomon

A propulsion concept relying on the direct, steady-state acceleration of a plasma by an inductive wave-launching antenna is presented. By operating inductively in steady state, a Direct Wave-Drive Thruster avoids drawbacks associated with electrode erosion and pulsed acceleration. The generalized relations for the scaling of thrust and efficiency with the antenna current are derived analytically; thrust is shown to scale with current squared, and efficiency is shown to increase with increasing current or power. Two specific configurations are modeled to determine nondimensional parameters governing the antenna-plasma coupling: an annular antenna pushing against a finite-conductivity plasma, and a linear antenna targeting the magnetosonic wave. Calculations from the model show that total thrust improves for increasing excitation frequencies, wavenumbers, plasma densities, and device sizes. To demonstrate the magnetosonic wave as an ideal candidate to drive a DWDT, it is shown to be capable of carrying substantial momentum and able to drive a variable specific impulse. The magnetosonic wave-driven mass flow is compared to mass transport due to thermal effects and cross-field diffusion in order to derive critical power requirements that ensure the thruster channel is dominated by wave dynamics. A proof-of-concept experiment is constructed that consists of a separate plasma source, a confining magnetic field, and a wave-launching antenna. The scaling of the increase of exhaust velocity is analytically modeled and is dependent on a nondimensional characteristic wavenumber that is proportional to the excitation frequency and plasma density and inversely proportional to the magnetic field strength. Experimental validation of the derived scaling behavior is carried out using a Mach probe to measure the flow velocity in the plume. Increases in exhaust velocity are measured as the antenna current increases for varying excitation frequencies and applied magnetic field

Design of a Laboratory Hall Thruster with Magnetically Shielded Channel Walls, Phase III: Comparison of Theory with Experiment

Science.gov (United States)

Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard R.; Goebel, Dan M.

2012-01-01

A proof-of-principle effort to demonstrate a technique by which erosion of the acceleration channel in Hall thrusters of the magnetic-layer type can be eliminated has been completed. The first principles of the technique, now known as "magnetic shielding," were derived based on the findings of numerical simulations in 2-D axisymmetric geometry. The simulations, in turn, guided the modification of an existing 6-kW laboratory Hall thruster. This magnetically shielded (MS) thruster was then built and tested. Because neither theory nor experiment alone can validate fully the first principles of the technique, the objective of the 2-yr effort was twofold: (1) to demonstrate in the laboratory that the erosion rates can be reduced by >order of magnitude, and (2) to demonstrate that the near-wall plasma properties can be altered according to the theoretical predictions. This paper concludes the demonstration of magnetic shielding by reporting on a wide range of comparisons between results from numerical simulations and laboratory diagnostics. Collectively, we find that the comparisons validate the theory. Near the walls of the MS thruster, theory and experiment agree: (1) the plasma potential has been sustained at values near the discharge voltage, and (2) the electron temperature has been lowered by at least 2.5-3 times compared to the unshielded (US) thruster. Also, based on carbon deposition measurements, the erosion rates at the inner and outer walls of the MS thruster are found to be lower by at least 2300 and 1875 times, respectively. Erosion was so low along these walls that the rates were below the resolution of the profilometer. Using a sputtering yield model with an energy threshold of 25 V, the simulations predict a reduction of 600 at the MS inner wall. At the outer wall ion energies are computed to be below 25 V, for which case we set the erosion to zero in the simulations. When a 50-V threshold is used the computed ion energies are below the threshold at both

A structural and thermal packaging approach for power processing units for 30-cm ion thrusters

Science.gov (United States)

Maloy, J. E.; Sharp, G. R.

1975-01-01

Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near earth and planetary missions. The thruster subsystem for these missions would consist of 30 centimeter ion thrusters with Power Processor Units (PPU) clustered in assemblies of from two to ten units. A preliminary design study of the electronic packaging of the PPU has been completed at Lewis Research Center of NASA. This study evaluates designs meeting the competing requirements of low system weight and overall mission flexibility. These requirements are evaluated regarding structural and thermal design, electrical efficiency, and integration of the electrical circuits into a functional PPU layout.

Molecular controls of the oxygenation and redox reactions of hemoglobin.

Science.gov (United States)

Bonaventura, Celia; Henkens, Robert; Alayash, Abdu I; Banerjee, Sambuddha; Crumbliss, Alvin L

2013-06-10

The broad classes of O(2)-binding proteins known as hemoglobins (Hbs) carry out oxygenation and redox functions that allow organisms with significantly different physiological demands to exist in a wide range of environments. This is aided by allosteric controls that modulate the protein's redox reactions as well as its O(2)-binding functions. The controls of Hb's redox reactions can differ appreciably from the molecular controls for Hb oxygenation and come into play in elegant mechanisms for dealing with nitrosative stress, in the malarial resistance conferred by sickle cell Hb, and in the as-yet unsuccessful designs for safe and effective blood substitutes. An important basic principle in consideration of Hb's redox reactions is the distinction between kinetic and thermodynamic reaction control. Clarification of these modes of control is critical to gaining an increased understanding of Hb-mediated oxidative processes and oxidative toxicity in vivo. This review addresses emerging concepts and some unresolved questions regarding the interplay between the oxygenation and oxidation reactions of structurally diverse Hbs, both within red blood cells and under acellular conditions. Developing methods that control Hb-mediated oxidative toxicity will be critical to the future development of Hb-based blood substitutes.

Nonlinear control of the Salnikov model reaction

DEFF Research Database (Denmark)

Recke, Bodil; Jørgensen, Sten Bay

1999-01-01

This paper explores different nonlinear control schemes, applied to a simple model reaction. The model is the Salnikov model, consisting of two ordinary differential equations. The control strategies investigated are I/O-linearisation, Exact linearisation, exact linearisation combined with LQR...

Development of a Methodology for Conducting Hall Thruster EMI Tests in Metal Vacuum Chambers of Arbitrary Shape and Size

Science.gov (United States)

Gallimore, Alec D.

2000-01-01

While the closed-drift Hall thruster (CDT) offers significant improvement in performance over conventional chemical rockets and other advanced propulsion systems such as the arcjet, its potential impact on spacecraft communication signals must be carefully assessed before widespread use of this device can take place. To this end, many of the potentially unique issues that are associated with these thrusters center on its plume plasma characteristics and the its interaction with electromagnetic waves. Although a great deal of experiments have been made in characterizing the electromagnetic interference (EMI) potential of these thrusters, the interpretation of the resulting data is difficult because most of these measurements have been made in vacuum chambers with metal walls which reflect radio waves emanating from the thruster. This project developed a means of assessing the impact of metal vacuum chambers of arbitrary size or shape on EMI experiments, thereby allowing for test results to be interpreted properly. Chamber calibration techniques were developed and initially tested at RIAME using their vacuum chamber. Calibration experiments were to have been made at Tank 5 of NASA GRC and the 6 m by 9 m vacuum chamber at the University of Michigan to test the new procedure, however the subcontract to RIAME was cancelled by NASA memorandum on Feb. 26. 1999.

Polyol synthesis of silver nanocubes via moderate control of the reaction atmosphere.

Science.gov (United States)

Jeon, Seog-Jin; Lee, Jae-Hwang; Thomas, Edwin L

2014-12-01

Silver nanocubes were successfully synthesized at high yield in variously controlled reaction atmospheres by balancing etching of O2/Cl(-) and reduction of glycolaldehyde. There have been efforts to control the O2 content in reaction atmospheres by purging of O2 or Ar gas for the balancing, but we found that moderate control of reaction atmosphere, just by careful timing of the opening and the capping of the reaction vial, greatly enhanced reproducibility. Enhanced reproducibility is attributed to alleviation of evaporation and condensation of glycolaldehyde (b.p.=131°C) by using capping at reaction temperatures higher than the b.p. of glycolaldehyde rather than purging with gas. The most important finding is that seeding is initiated by HNO3 induced deoxygenation reaction in the gas phase. O2 is consumed by oxidation of NO generated from the silver etching reaction by HNO3, which effectively controls the reaction atmosphere without introduction of gas. Our simple method to control reaction atmosphere reduces the overall reaction time to one fifth of the previous result and provides excellent size and distribution selectivity of the Ag nanocube product. Copyright © 2014 Elsevier Inc. All rights reserved.

Molecular Controls of the Oxygenation and Redox Reactions of Hemoglobin

Science.gov (United States)

Henkens, Robert; Alayash, Abdu I.; Banerjee, Sambuddha; Crumbliss, Alvin L.

2013-01-01

Abstract Significance: The broad classes of O2-binding proteins known as hemoglobins (Hbs) carry out oxygenation and redox functions that allow organisms with significantly different physiological demands to exist in a wide range of environments. This is aided by allosteric controls that modulate the protein's redox reactions as well as its O2-binding functions. Recent Advances: The controls of Hb's redox reactions can differ appreciably from the molecular controls for Hb oxygenation and come into play in elegant mechanisms for dealing with nitrosative stress, in the malarial resistance conferred by sickle cell Hb, and in the as-yet unsuccessful designs for safe and effective blood substitutes. Critical Issues: An important basic principle in consideration of Hb's redox reactions is the distinction between kinetic and thermodynamic reaction control. Clarification of these modes of control is critical to gaining an increased understanding of Hb-mediated oxidative processes and oxidative toxicity in vivo. Future Directions: This review addresses emerging concepts and some unresolved questions regarding the interplay between the oxygenation and oxidation reactions of structurally diverse Hbs, both within red blood cells and under acellular conditions. Developing methods that control Hb-mediated oxidative toxicity will be critical to the future development of Hb-based blood substitutes. Antioxid. Redox Signal. 18, 2298–2313. PMID:23198874

Direct measurement of axial momentum imparted by an electrothermal radiofrequency plasma micro-thruster

Science.gov (United States)

Charles, Christine; Boswell, Roderick; Bish, Andrew; Khayms, Vadim; Scholz, Edwin

2016-05-01

Gas flow heating using radio frequency plasmas offers the possibility of depositing power in the centre of the flow rather than on the outside, as is the case with electro-thermal systems where thermal wall losses lower efficiency. Improved systems for space propulsion are one possible application and we have tested a prototype micro-thruster on a thrust balance in vacuum. For these initial tests, a fixed component radio frequency matching network weighing 90 grams was closely attached to the thruster in vacuum with the frequency agile radio frequency generator power being delivered via a 50 Ohm cable. Without accounting for system losses (estimated at around 50%), for a few 10s of Watts from the radio frequency generator the specific impulse was tripled to ˜48 seconds and the thrust tripled from 0.8 to 2.4 milli-Newtons.

Direct measurement of axial momentum imparted by an electrothermal radiofrequency plasma micro-thruster

Directory of Open Access Journals (Sweden)

Christine eCharles

2016-05-01

Full Text Available Gas flow heating using radio frequency plasmas offers the possibility of depositing power in the centre of the flow rather than on the outside, as is the case with electro-thermal systems where thermal wall losses lower efficiency. Improved systems for space propulsion are one possible application and we have tested a prototype micro-thruster on a thrust balance in vacuum. For these initial tests, a fixed component radio frequency matching network weighing 90 grams was closely attached to the thruster in vacuum with the frequency agile radio frequency generator power being delivered via a 50 Ohm cable. Without accounting for system losses (estimated at around 50~$%$, for a few 10s of Watts from the radio frequency generator the specific impulse was tripled to $sim$48 seconds and the thrust tripled from 0.8 to 2.4 milli-Newtons.

Study on the Attitude Control of Spacecraft Using Reaction Wheels

Directory of Open Access Journals (Sweden)

Ju-Young Du

1998-06-01

Full Text Available Attitude determination and control of satellite is important component which determines the accomplish satellite missions. In this study, attitude control using reaction wheels and momentum dumping of wheels are considered. Attitude control law is designed by Sliding control and LQR. Attitude maneuver control law is obtained by Shooting method. Wheels momentum dumping control law is designed by Bang-Bang control. Four reaction wheels are configurated for minimized the electric power consumption. Wheels control torque and magnetic moment of magnetic torquer are limited.

A direct-measurement technique for estimating discharge-chamber lifetime. [for ion thrusters

Science.gov (United States)

Beattie, J. R.; Garvin, H. L.

1982-01-01

The use of short-term measurement techniques for predicting the wearout of ion thrusters resulting from sputter-erosion damage is investigated. The laminar-thin-film technique is found to provide high precision erosion-rate data, although the erosion rates are generally substantially higher than those found during long-term erosion tests, so that the results must be interpreted in a relative sense. A technique for obtaining absolute measurements is developed using a masked-substrate arrangement. 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 (approximately 1 hr) tests. Results obtained using the direct-measurement technique are shown to agree with sputter-erosion depths calculated for the plasma conditions of the test. The direct-measurement approach is found to be applicable to both mercury and argon discharge-plasma environments and will be useful for estimating the lifetimes of inert gas and extended performance mercury ion thrusters currently under development.

Two-Dimensional Modelling of the Hall Thruster Discharge: Final Report

Science.gov (United States)

2007-09-10

ion energy flux to wall, qWi, and electron energy flux to wall, qWe for Vd= 300 V, 600 V and 750 V. All variables are evaluated at the outer wall (r... qWe for Vd= 300 V, 600 V and 750 V. All variables are evaluated at the outer wall (r=0.05m). The vertical dashed line represents the thruster exit

Electron energy distribution function in a low-power Hall thruster discharge and near-field plume

Science.gov (United States)

Tichý, M.; Pétin, A.; Kudrna, P.; Horký, M.; Mazouffre, S.

2018-06-01

Electron temperature and plasma density, as well as the electron energy distribution function (EEDF), have been obtained inside and outside the dielectric channel of a 200 W permanent magnet Hall thruster. Measurements were carried out by means of a cylindrical Langmuir probe mounted onto a compact fast moving translation stage. The 3D particle-in cell numerical simulations complement experiments. The model accounts for the crossed electric and magnetic field configuration in a weakly collisional regime where only electrons are magnetized. Since only the electron dynamics is of interest in this study, an artificial mass of ions corresponding to mi = 30 000me was used to ensure ions could be assumed at rest. The simulation domain is located at the thruster exit plane and does not include the cathode. The measured EEDF evidences a high-energy electron population that is superimposed onto the low energy bulk population outside the channel. Inside the channel, the EEDF is close to Maxwellian. Both the experimental and numerical EEDF depart from an equilibrium distribution at the channel exit plane, a region of high magnetic field. We therefore conclude that the fast electron group found in the experiment corresponds to the electrons emitted by the external cathode that reach the thruster discharge without experiencing collision events.

Diagnostic Setup for Characterization of Near-Anode Processes in Hall Thrusters

International Nuclear Information System (INIS)

Dorf, L.; Raitses, Y.; Fisch, N.J.

2003-01-01

A diagnostic setup for characterization of near-anode processes in Hall-current plasma thrusters consisting of biased and emissive electrostatic probes, high-precision positioning system and low-noise electronic circuitry was developed and tested. Experimental results show that radial probe insertion does not cause perturbations to the discharge and therefore can be used for accurate near-anode measurements

Satellite Integration of a PhoneSat-EDSN Bus with a Micro Cathode Arc Thruster

Data.gov (United States)

National Aeronautics and Space Administration —  NASA Ames Research Center and GWU are investigating applications of Micro-Cathode Arc Thrusters (μCAT) sub-systems for attitude and orbit correction of a PhoneSat...

Femtosecond laser control of chemical reaction of carbon monoxide and hydrogen

CSIR Research Space (South Africa)

Du Plessis, A

2010-09-01

Full Text Available Femtosecond laser control of chemical reactions is made possible through the use of pulse-shaping techniques coupled to a learning algorithm feedback loop – teaching the laser pulse to control the chemical reaction. This can result in controllable...

Effect of Ambipolar Potential on the Propulsive Performance of the GDM Plasma Thruster, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — The Gasdynamic Mirror (GDM) thruster is an electric propulsion device, without electrodes, that will magnetically confine a plasma with such density and temperature...

Versatile Dual Photoresponsive System for Precise Control of Chemical Reactions.

Science.gov (United States)

Xu, Can; Bing, Wei; Wang, Faming; Ren, Jinsong; Qu, Xiaogang

2017-08-22

A versatile method for photoregulation of chemical reactions was developed through a combination of near-infrared (NIR) and ultraviolet (UV) light sensitive materials. This regulatory effect was achieved through photoresponsive modulation of reaction temperature and pH values, two prominent factors influencing reaction kinetics. Photothermal nanomaterial graphene oxide (GO) and photobase reagent malachite green carbinol base (MGCB) were selected for temperature and pH regulation, respectively. Using nanocatalyst- and enzyme-mediated chemical reactions as model systems, we demonstrated the feasibility and high efficiency of this method. In addition, a photoresponsive, multifunctional "Band-aid"-like hydrogel platform was presented for programmable wound healing. Overall, this simple, efficient, and reversible system was found to be effective for controlling a wide variety of chemical reactions. Our work may provide a method for remote and sustainable control over chemical reactions for industrial and biomedical applications.

High-Pressure Lightweight Thrusters

Science.gov (United States)

Holmes, Richard; McKechnie, Timothy; Shchetkovskiy, Anatoliy; Smirnov, Alexander

2013-01-01

Returning samples of Martian soil and rock to Earth is of great interest to scientists. There were numerous studies to evaluate Mars Sample Return (MSR) mission architectures, technology needs, development plans, and requirements. The largest propulsion risk element of the MSR mission is the Mars Ascent Vehicle (MAV). Along with the baseline solid-propellant vehicle, liquid propellants have been considered. Similar requirements apply to other lander ascent engines and reaction control systems. The performance of current state-ofthe- art liquid propellant engines can be significantly improved by increasing both combustion temperature and pressure. Pump-fed propulsion is suggested for a single-stage bipropellant MAV. Achieving a 90-percent stage propellant fraction is thought to be possible on a 100-kg scale, including sufficient thrust for lifting off Mars. To increase the performance of storable bipropellant rocket engines, a high-pressure, lightweight combustion chamber was designed. Iridium liner electrodeposition was investigated on complex-shaped thrust chamber mandrels. Dense, uniform iridium liners were produced on chamber and cylindrical mandrels. Carbon/carbon composite (C/C) structures were braided over iridium-lined mandrels and densified by chemical vapor infiltration. Niobium deposition was evaluated for forming a metallic attachment flange on the carbon/ carbon structure. The new thrust chamber was designed to exceed state-of-the-art performance, and was manufactured with an 83-percent weight savings. High-performance C/Cs possess a unique set of properties that make them desirable materials for high-temperature structures used in rocket propulsion components, hypersonic vehicles, and aircraft brakes. In particular, more attention is focused on 3D braided C/Cs due to their mesh-work structure. Research on the properties of C/Cs has shown that the strength of composites is strongly affected by the fiber-matrix interfacial bonding, and that weakening

Bond-selective control of a gas-surface reaction

Science.gov (United States)

Killelea, Daniel R.

The prospect of using light to selectively control chemical reactions has tantalized chemists since the development of the laser. Unfortunately, the realization of laser-directed chemistry is frequently thwarted by the randomization of energy within the molecule through intramolecular vibrational energy distribution (IVR). However, recent results showing vibrational mode-specific reactivity on metal surfaces suggest that IVR may not always be complete for gas-surface reactions. Here, we combine molecular beam techniques and direct laser excitation to characterize the bond-specific reactivity of trideuteromethane on a Ni(111) surface. Our results reveal important details about how vibrational energy is distributed in the reactive molecule. We use a molecular beam to direct state-selected trideuteromethane (CHD 3) molecules onto a nickel single crystal sample and use the results we obtain to describe the flow of vibrational energy in the methane-surface reaction complex. We show that CHD3 molecules initially excited to v=1, J=2, K=0 of the v 1 symmetric C-H stretching mode will dissociate exclusively via C-H cleavage on Ni(111). This result highlights the localization of vibrational energy in the reaction complex, despite the presence of many energy exchange channels with the high state-density surface. We demonstrate, for the first time, highly parallel bond-selective control of a heterogeneously catalyzed reaction. We place our results in the context of recent experiments investigating IVR for molecules in both the gas phase and liquid solutions. If IVR is fast on the reaction timescale, vibrational energy would be randomly distributed throughout the nascent methane-surface reaction complex and vibrational mode-specific behavior would not occur. The short timescale of a direct gas-surface collision may explain how the exchange of energy via IVR is limited to only a small subset of the energetic configurations available to the reaction complex. This framework

Effect of Ambipolar Potential on the Propulsive Performance of the GDM Plasma Thruster, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — The gasdynamic mirror (GDM) plasma thruster has the ability to confine high-density plasma for the length of time required to heat it to the temperatures...

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

Science.gov (United States)

Pinero, Luis R.; Scheidegger, Robert J.; Aulsio, Michael V.; Birchenough, Arthur G.

2014-01-01

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

Experimental study of effect of magnetic field on anode temperature distribution in an ATON-type Hall thruster

Science.gov (United States)

Liu, Jinwen; Li, Hong; Mao, Wei; Ding, Yongjie; Wei, Liqiu; Li, Jianzhi; Yu, Daren; Wang, Xiaogang

2018-05-01

The energy deposition caused by the absorption of electrons by the anode is an important cause of power loss in a Hall thruster. The resulting anode heating is dangerous, as it can potentially reduce the thruster lifetime. In this study, by considering the ring shape of the anode of an ATON-type Hall thruster, the effects of the magnetic field strength and gradient on the anode ring temperature distribution are studied via experimental measurement. The results show that the temperature distribution is not affected by changes in the magnetic field strength and that the position of the peak temperature is essentially unchanged; however, the overall temperature does not change monotonically with the increase of the magnetic field strength and is positively correlated with the change in the discharge current. Moreover, as the magnetic field gradient increases, the position of the peak temperature gradually moves toward the channel exit and the temperature tends to decrease as a whole, regardless of the discharge current magnitude; in any case, the position of the peak temperature corresponds exactly to the intersection of the magnetic field cusp with the anode ring. Further theoretical analysis shows that the electrons, coming from the ionization region, travel along two characteristic paths to reach the anode under the guidance of the cusped magnetic field configuration. The change of the magnetic field strength or gradient changes the transfer of momentum and energy of the electrons in these two paths, which is the main reason for the changes in the temperature and distribution. This study is instructive for matching the design of the ring-shaped anode and the cusp magnetic field of an ATON-type Hall thruster.

Long life reaction control system design

Science.gov (United States)

Fanciullo, Thomas J.; Judd, Craig

1993-02-01

Future single stage to orbit systems will utilize oxygen/hydrogen propellants in their main propulsion means due to the propellant's high energy content and environmental acceptability. Operational effectiveness studies and life cycle cost studies have indicated that minimizing the number of different commodities on a given vehicle not only reduces cost, but reduces the ground span times in both the pre- and postflight operations. Therefore, oxygen and hydrogen should be used for the reaction controls systems, eliminating the need to deal with toxic or corrosive fluids. When the hydrogen scramjet powered NASP design development began in 1985, new system design studies considered overall integration of subsystems; in the context of that approach, O2/H2 reaction controls system were more than competitive with storable propellant systems and had the additional benefits of lower life cycle cost, rapid turnaround times, and O2 and H2 commodities for use throughout the vehicle. Similar benefits were derived in rocket-powered SSTO vehicles.

Development of Long-Lifetime Pulsed Gas Valves for Pulsed Electric Thrusters

Science.gov (United States)

Burkhardt, Wendel M.; Crapuchettes, John M.; Addona, Brad M.; Polzin, Kurt A.

2015-01-01

It is advantageous for gas-fed pulsed electric thrusters to employ pulsed valves so propellant is only flowing to the device during operation. The propellant utilization of the thruster will be maximized when all the gas injected into the thruster is acted upon by the fields produced by the electrical pulse. Gas that is injected too early will diffuse away from the thruster before the electrical pulse can act to accelerate the propellant. Gas that is injected too late will miss being accelerated by the already-completed electrical pulse. As a consequence, the valve must open quickly and close equally quickly, only remaining open for a short duration. In addition, the valve must have only a small amount of volume between the sealing body and the thruster so the front and back ends of the pulse are as coincident as possible with the valve cycling, with very little latent propellant remaining in the feed lines after the valve is closed. For a real mission of interest, a pulsed thruster can be expected to pulse at least 10(exp 10) - 10(exp 11) times, setting the range for the number of times a valve must open and close. The valves described in this paper have been fabricated and tested for operation in an inductive pulsed plasma thruster (IPPT) for in-space propulsion. In general, an IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged, producing a high-current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed, it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. The valve characteristics needed for the IPPT application require a fast-acting valve capable of a minimum of 10(exp 10) valve actuation cycles. Since

Development and Testing of High Current Hollow Cathodes for High Power Hall Thrusters

Science.gov (United States)

Kamhawi, Hani; Van Noord, Jonathan

2012-01-01

NASA's Office of the Chief Technologist In-Space Propulsion project is sponsoring the testing and development of high power Hall thrusters for implementation in NASA missions. As part of the project, NASA Glenn Research Center is developing and testing new high current hollow cathode assemblies that can meet and exceed the required discharge current and life-time requirements of high power Hall thrusters. This paper presents test results of three high current hollow cathode configurations. Test results indicated that two novel emitter configurations were able to attain lower peak emitter temperatures compared to state-of-the-art emitter configurations. One hollow cathode configuration attained a cathode orifice plate tip temperature of 1132 degC at a discharge current of 100 A. More specifically, test and analysis results indicated that a novel emitter configuration had minimal temperature gradient along its length. Future work will include cathode wear tests, and internal emitter temperature and plasma properties measurements along with detailed physics based modeling.

The effect of magnetic mirror on near wall conductivity in Hall thrusters

International Nuclear Information System (INIS)

Yu, D.; Liu, H.; Fu, H.; Cao, Y.

2008-01-01

The effect of magnetic mirror on near wall conductivity is studied in the acceleration region of Hall thrusters. The electron dynamics process in the plasma is described by test particle method, in which electrons are randomly emitted from the centerline towards the inner wall of the channel. It is found that the effective collision coefficient, i.e. the rate of electrons colliding with the wall, changes dramatically with the magnetic mirror effect being considered; and that it decreases further with the increase of magnetic mirror ratio to enhance the electron mobility accordingly. In particular, under anistropic electron velocity distribution conditions, the magnetic mirror effect becomes even more prominent. Furthermore, due to decrease in magnetic mirror ratio from the exhaust plane to the anode in Hall thrusters, the axial gradient of electron mobility with magnetic mirror effect is greater than without it. The magnetic mirror effects on electron mobility are derived analytically and the results are found in agreement with the simulation. (copyright 2008 WILEY-VCH Verlag GmbH and Co. KGaA, Weinheim) (orig.)

Test Results of a 200 W Class Hall Thruster

Science.gov (United States)

Jacobson, David; Jankovsky, Robert S.

1999-01-01

The performance of a 200 W class Hall thruster was evaluated. Performance measurements were taken at power levels between 90 W and 250 W. At the nominal 200 W design point, the measured thrust was 11.3 mN. and the specific impulse was 1170 s excluding cathode flow in the calculation. A laboratory model 3 mm diameter hollow cathode was used for all testing. The engine was operated on laboratory power supplies in addition to a breadboard power processing unit fabricated from commercially available DC to DC converters.

CASTOR: Cathode/Anode Satellite Thruster for Orbital Repositioning

Science.gov (United States)

Mruphy, Gloria A.

2010-01-01

The purpose of CASTOR (Cathode/Anode Satellite Thruster for Orbital Repositioning) satellite is to demonstrate in Low Earth Orbit (LEO) a nanosatellite that uses a Divergent Cusped Field Thruster (DCFT) to perform orbital maneuvers representative of an orbital transfer vehicle. Powered by semi-deployable solar arrays generating 165W of power, CASTOR will achieve nearly 1 km/s of velocity increment over one year. As a technology demonstration mission, success of CASTOR in LEO will pave the way for a low cost, high delta-V orbital transfer capability for small military and civilian payloads in support of Air Force and NASA missions. The educational objective is to engage graduate and undergraduate students in critical roles in the design, development, test, carrier integration and on-orbit operations of CASTOR as a supplement to their curricular activities. This program is laying the foundation for a long-term satellite construction program at MIT. The satellite is being designed as a part of AFRL's University Nanosatellite Program, which provides the funding and a framework in which student satellite teams compete for a launch to orbit. To this end, the satellite must fit within an envelope of 50cmx50cmx60cm, have a mass of less than 50kg, and meet stringent structural and other requirements. In this framework, the CASTOR team successfully completed PDR in August 2009 and CDR in April 2010 and will compete at FCR (Flight Competition Review) in January 2011. The complexity of the project requires implementation of many systems engineering techniques which allow for development of CASTOR from conception through FCR and encompass the full design, fabrication, and testing process.

Magnetically filtered Faraday probe for measuring the ion current density profile of a Hall thruster

International Nuclear Information System (INIS)

Rovey, Joshua L.; Walker, Mitchell L.R.; Gallimore, Alec D.; Peterson, Peter Y.

2006-01-01

The ability of a magnetically filtered Faraday probe (MFFP) to obtain the ion current density profile of a Hall thruster is investigated. The MFFP is designed to eliminate the collection of low-energy, charge-exchange (CEX) ions by using a variable magnetic field as an ion filter. In this study, a MFFP, Faraday probe with a reduced acceptance angle (BFP), and nude Faraday probe are used to measure the ion current density profile of a 5 kW Hall thruster operating over the range of 300-500 V and 5-10 mg/s. The probes are evaluated on a xenon propellant Hall thruster in the University of Michigan Large Vacuum Test Facility at operating pressures within the range of 4.4x10 -4 Pa Xe (3.3x10 -6 Torr Xe) to 1.1x10 -3 Pa Xe (8.4x10 -6 Torr Xe) in order to study the ability of the Faraday probe designs to filter out CEX ions. Detailed examination of the results shows that the nude probe measures a greater ion current density profile than both the MFFP and BFP over the range of angular positions investigated for each operating condition. The differences between the current density profiles obtained by each probe are attributed to the ion filtering systems employed. Analysis of the results shows that the MFFP, operating at a +5 A solenoid current, provides the best agreement with flight-test data and across operating pressures

Effect of the Thruster Configurations on a Laser Ignition Microthruster

Science.gov (United States)

Koizumi, Hiroyuki; Hamasaki, Kyoichi; Kondo, Ryo; Okada, Keisuke; Nakano, Masakatsu; Arakawa, Yoshihiro

Research and development of small spacecraft have advanced extensively throughout the world and propulsion devices suitable for the small spacecraft, microthruster, is eagerly anticipated. The authors proposed a microthruster using 1—10-mm-size solid propellant. Small pellets of solid propellant are installed in small combustion chambers and ignited by the irradiation of diode laser beam. This thruster is referred as to a laser ignition microthruster. Solid propellant enables large thrust capability and compact propulsion system. To date theories of a solid-propellant rocket have been well established. However, those theories are for a large-size solid propellant and there are a few theories and experiments for a micro-solid rocket of 1—10mm class. This causes the difficulty of the optimum design of a micro-solid rocket. In this study, we have experimentally investigated the effect of thruster configurations on a laser ignition microthruster. The examined parameters are aperture ratio of the nozzle, length of the combustion chamber, area of the nozzle throat, and divergence angle of the nozzle. Specific impulse dependences on those parameters were evaluated. It was found that large fraction of the uncombusted propellant was the main cause of the degrading performance. Decreasing the orifice diameter in the nozzle with a constant open aperture ratio was an effective method to improve this degradation.

Determining two-step control in heterogeneous catalytic reactions

Energy Technology Data Exchange (ETDEWEB)

Chakrabarty, T; Silveston, P L; Hudgins, R R

1979-10-01

The data by Thaller and Thodos on the sec.-butanol dehydrogenation to methyl ethyl ketone on brass catalyst indicated that a dual site surface reaction was rate-controlling below 575/sup 0/K and hydrogen desorption was rate-controlling above 616/sup 0/K (Vertical BarAIChE J.

Diffusion-controlled reaction. V. Effect of concentration-dependent diffusion coefficient on reaction rate in graft polymerization

International Nuclear Information System (INIS)

Imre, K.; Odian, G.

1979-01-01

The effect of diffusion on radiation-initiated graft polymerization has been studied with emphasis on the single- and two-penetrant cases. When the physical properties of the penetrants are similar, the two-penetrant problems can be reduced to the single-penetrant problem by redefining the characteristic parameters of the system. The diffusion-free graft polymerization rate is assumed to be proportional to the upsilon power of the monomer concentration respectively, and, in which the proportionality constant a = k/sub p/R/sub i//sup w//k/sub t//sup z/, where k/sub p/ and k/sub t/ are the propagation and termination rate constants, respectively, and R/sub i/ is the initiation rate. The values of upsilon, w, and z depend on the particular reaction system. The results of earlier work were generalized by allowing a non-Fickian diffusion rate which predicts an essentially exponential dependence on the monomer concentration of the diffusion coefficient, D = D 0 [exp(deltaC/M)], where M is the saturation concentration. A reaction system is characterized by the three dimensionless parameters, upsilon, delta, and A = (L/2)[aM/sup (upsilon--1)//D 0 ]/sup 1/2/, where L is the polymer film thickness. Graft polymerization tends to become diffusion controlled as A increases. Larger values of delta and ν cause a reaction system to behave closer to the diffusion-free regime. Transition from diffusion-free to diffusion-controlled reaction involves changes in the dependence of the reaction rate on film thickness, initiation rate, and monomer concentration. Although the diffusion-free rate is w order in initiation rate, upsilon order in monomer, and independent of film thickness, the diffusion-controlled rate is w/2 order in initiator rate and inverse first-order in film thickness. Dependence of the diffusion-controlled rate on monomer is dependent in a complex manner on the diffusional characteristics of the reaction system. 11 figures, 4 tables

An experimental investigation of the internal magnetic field topography of an operating Hall thruster

International Nuclear Information System (INIS)

Peterson, Peter Y.; Gallimore, Alec D.; Haas, James M.

2002-01-01

Magnetic field measurements were made in the discharge channel of the 5 kW-class P5 laboratory-model Hall thruster to investigate what effect the Hall current has on the static, applied magnetic field topography. The P5 was operated at 1.6 and 3.0 kW with a discharge voltage of 300 V. A miniature inductive loop probe (B-Dot probe) was employed to measure the radial magnetic field profile inside the discharge channel of the P5 with and without the plasma discharge. These measurements are accomplished with minimal disturbance to thruster operation with the High-speed Axial Reciprocating Probe system. The results of the B-Dot probe measurements indicate a change in the magnetic field topography from that of the vacuum field measurements. The measured magnetic field profiles are then examined to determine the possible nature and source of the difference between the vacuum and plasma magnetic field profiles

Confidence Testing of Shell 405 and S-405 Catalysts in a Monopropellant Hydrazine Thruster

Science.gov (United States)

McRight, Patrick; Popp, Chris; Pierce, Charles; Turpin, Alicia; Urbanchock, Walter; Wilson, Mike

2005-01-01

As part of the transfer of catalyst manufacturing technology from Shell Chemical Company (Shell 405 catalyst manufactured in Houston, Texas) to Aerojet (S-405 manufactured in Redmond, Washington), Aerojet demonstrated the equivalence of S-405 and Shell 405 at beginning of life. Some US aerospace users expressed a desire to conduct a preliminary confidence test to assess end-of-life characteristics for S-405. NASA Marshall Space Flight Center (MSFC) and Aerojet entered a contractual agreement in 2004 to conduct a confidence test using a pair of 0.2-lbf MR-103G monopropellant hydrazine thrusters, comparing S-405 and Shell 405 side by side. This paper summarizes the formulation of this test program, explains the test matrix, describes the progress of the test, and analyzes the test results. This paper also includes a discussion of the limitations of this test and the ramifications of the test results for assessing the need for future qualification testing in particular hydrazine thruster applications.

Asymmetrical Capacitors for Propulsion and the ISR Asymmetrical Capacitator Thruster, Experimental Results and Improved Designs

Science.gov (United States)

Canning, Francis; Winet, Ed; Ice, Bob; Melcher, Cory; Pesavento, Phil; Holmes, Alan; Butler, Carey; Cole, John; Campbell, Jonathan

2004-01-01

The outline of this viewgraph presentation on asymmetrical capacitor thruster development includes: 1) Test apparatus; 2) Devices tested; 3) Circuits used; 4) Data collected (Time averaged, Time resolved); 5) Patterns observed; 6) Force calculation; 7) Electrostatic modeling; 8) Understand it all.

Characteristics of a non-volatile liquid propellant in liquid-fed ablative pulsed plasma thrusters

Science.gov (United States)

Ling, William Yeong Liang; Schönherr, Tony; Koizumi, Hiroyuki

2017-02-01

In the past several decades, the use of electric propulsion in spacecraft has experienced tremendous growth. With the increasing adoption of small satellites in the kilogram range, suitable propulsion systems will be necessary in the near future. Pulsed plasma thrusters (PPTs) were the first form of electric propulsion to be deployed in orbit, and are highly suitable for small satellites due to their inherent simplicity. However, their lifetime is limited by disadvantages such as carbon deposition leading to thruster failure, and complicated feeding systems required due to the conventional use of solid propellants (usually polytetrafluoroethylene (PTFE)). A promising alternative to solid propellants has recently emerged in the form of non-volatile liquids that are stable in vacuum. This study presents a broad comparison of the non-volatile liquid perfluoropolyether (PFPE) and solid PTFE as propellants on a PPT with a common design base. We show that liquid PFPE can be successfully used as a propellant, and exhibits similar plasma discharge properties to conventional solid PTFE, but with a mass bit that is an order of magnitude higher for an identical ablation area. We also demonstrate that the liquid PFPE propellant has exceptional resistance to carbon deposition, completely negating one of the major causes of thruster failure, while solid PTFE exhibited considerable carbon build-up. Energy dispersive X-ray spectroscopy was used to examine the elemental compositions of the surface deposition on the electrodes and the ablation area of the propellant (or PFPE encapsulator). The results show that based on its physical characteristics and behavior, non-volatile liquid PFPE is an extremely promising propellant for use in PPTs, with an extensive scope available for future research and development.

Mars Ascent Vehicle Reaction Control System, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — During this Phase I NASA program, Valley Tech Systems (VTS) will develop an innovative solid Reaction Control System (RCS) architecture concept design that can...

Hybrid-Particle-In-Cell Simulation of Backsputtered Carbon Transport in the Near-Field Plume of a Hall Thruster

Science.gov (United States)

Choi, Maria; Yim, John T.; Williams, George J.; Herman, Daniel A.; Gilland, James H.

2018-01-01

Magnetic shielding has eliminated boron nitride erosion as the life limiting mechanism in a Hall thruster but has resulted in erosion of the front magnetic field pole pieces. Recent experiments show that the erosion of graphite pole covers, which are added to protect the magnetic field pole pieces, causes carbon to redeposit on other surfaces, such as boron nitride discharge channel and cathode keeper surfaces. As a part of the risk-reduction activities for Advanced Electric Propulsion System thruster development, this study models transport of backsputtered carbon from the graphite front pole covers and vacuum facility walls. Fluxes, energy distributions, and redeposition rates of backsputtered carbon on the anode, discharge channel, and graphite cathode keeper surfaces are predicted.

Design of an embedded inverse-feedforward biomolecular tracking controller for enzymatic reaction processes.

Science.gov (United States)

Foo, Mathias; Kim, Jongrae; Sawlekar, Rucha; Bates, Declan G

2017-04-06

Feedback control is widely used in chemical engineering to improve the performance and robustness of chemical processes. Feedback controllers require a 'subtractor' that is able to compute the error between the process output and the reference signal. In the case of embedded biomolecular control circuits, subtractors designed using standard chemical reaction network theory can only realise one-sided subtraction, rendering standard controller design approaches inadequate. Here, we show how a biomolecular controller that allows tracking of required changes in the outputs of enzymatic reaction processes can be designed and implemented within the framework of chemical reaction network theory. The controller architecture employs an inversion-based feedforward controller that compensates for the limitations of the one-sided subtractor that generates the error signals for a feedback controller. The proposed approach requires significantly fewer chemical reactions to implement than alternative designs, and should have wide applicability throughout the fields of synthetic biology and biological engineering.

Modeling Reaction Control System Effects on Mars Odyssey

National Research Council Canada - National Science Library

Hanna, Jill

2002-01-01

...) simulations to determine rotational motion of the spacecraft. The main objective of this study was to assess the reaction control system models and their effects on the atmospheric flight of Odyssey...

ION ACOUSTIC TURBULENCE, ANOMALOUS TRANSPORT, AND SYSTEM DYNAMICS IN HALL EFFECT THRUSTERS

Science.gov (United States)

2017-06-30

NUMBER (Include area code) 30 June 2017 Briefing Charts 26 May 2017 - 30 June 2017 ION ACOUSTIC TURBULENCE, ANOMALOUS TRANSPORT, AND SYSTEM DYNAMICS ...Robert Martin N/A ION ACOUSTIC TURBULENCE, ANOMALOUS TRANSPORT, AND SYSTEM DYNAMICS IN HALL EFFECT THRUSTERS Robert Martin1, Jonathan Tran2 1AIR FORCE...Approved for Public Release; Distribution is Unlimited. PA# 17394 1 / 13 OUTLINE 1 INTRODUCTION 2 TRANSPORT 3 DYNAMIC SYSTEM 4 SUMMARY AND CONCLUSION

Investigation of the Hall Effect Thruster Breathing Mode and Spoke Mode Instabilities in the Very Near Field

Data.gov (United States)

National Aeronautics and Space Administration — One of the most practical forms of electric propulsion is the Hall Effect Thruster (HET), which makes use of electric and magnetic fields to create and eject a...

Effect of Anode Magnetic Shield on Magnetic Field and Ion Beam in Cylindrical Hall Thruster

International Nuclear Information System (INIS)

Zhao Jie; Wang Shiqing; Liu Jian; Xu Li; Tang Deli; Geng Shaofei

2010-01-01

Numerical simulation of the effect of the anode magnetic shielding on the magnetic field and ion beam in a cylindrical Hall thruster is presented. The results show that after the anode is shielded by the magnetic shield, the magnetic field lines near the anode surface are obviously convex curved, the ratio of the magnetic mirror is enhanced, the width of the positive magnetic field gradient becomes larger than that without the anode magnetic shielding, the radial magnetic field component is enhanced, and the discharge plasma turbulence is reduced as a result of keeping the original saddle field profile and the important role the other two saddle field profiles play in restricting electrons. The results of the particle in cell (PIC) numerical simulation show that both the ion number and the energy of the ion beam increase after the anode is shielded by the magnetic shield. In other words, the specific impulse of the cylindrical Hall thruster is enhanced.

Synthesis of porous gold nanoshells by controlled transmetallation reaction

Energy Technology Data Exchange (ETDEWEB)

Pattabi, Manjunatha, E-mail: manjupattabi@yahoo.com; M, Krishnaprabha [Department of Materials Science, Mangalore University, Mangalagangothri-574199 (India)

2015-06-24

Aqueous synthesis of porous gold nanoshells in one step is carried out through controlled transmetallation (TM) reaction using a naturally available egg shell membrane (ESM) as a barrier between the sacrificial silver particles (AgNPs) and the gold precursor solution (HAuCl{sub 4}). The formation of porous gold nanoshells via TM reaction is inferred from UV-Vis spectroscopy and the scanning electron microscopic (SEM) studies.

Reaction parameters for controlled sonosynthesis of gold nanoparticles

Energy Technology Data Exchange (ETDEWEB)

Gonzalez M, A. L. [Universidad Autonoma del Estado de Mexico, Facultad de Quimica, Paseo Colon esq. Paseo Tollocan s/n, 50120 Toluca, Estado de Mexico (Mexico); Cabrera L, L. I. [UNAM-UAEM, Centro Conjunto de Investigacion en Quimica Sustentable, Km 14.5 Carretera Toluca-Atlacomulco, 50200 San Cayetano-Toluca, Estado de Mexico (Mexico)

2015-07-01

The synthesis of gold nanoparticles by sonochemical technique has been previously performed with excellent results. The synthesis has been carried out in the presence of citric acid, a strong reducing agent, which allows the nucleation and growth of gold nanoparticles, at the same time that controls particle size. In this work we report the use of sodium tartrate as a mild reducing agent that allows a better understanding of the effect of the reaction parameters during gold nanoparticle synthesis. A conventional sonication bath (37 k Hz) was used for the sonochemical synthesis. This work focuses on the reaction temperature effect and the effect of sodium tartrate concentration. It was confirmed that particle size, and particle morphology is dependent of these two reaction parameters. Equally, colloidal stabilization was related to reaction temperature and sodium tartrate concentration. It was also determined that Ostwald ripening takes place during sonochemical reaction under our conditions, allowing to understand the mechanism that takes place during synthesis. Gold nanoparticles with main particle size of 17 nm were achieved by this method. Characterization techniques used: Fourier transform infrared spectra (Ftir), X-ray diffraction and Atomic Force Microscope was used in order to determine particle size of the synthetic product of reaction M10c by tapping mode. (Author)

Reaction parameters for controlled sonosynthesis of gold nanoparticles

International Nuclear Information System (INIS)

Gonzalez M, A. L.; Cabrera L, L. I.

2015-01-01

The synthesis of gold nanoparticles by sonochemical technique has been previously performed with excellent results. The synthesis has been carried out in the presence of citric acid, a strong reducing agent, which allows the nucleation and growth of gold nanoparticles, at the same time that controls particle size. In this work we report the use of sodium tartrate as a mild reducing agent that allows a better understanding of the effect of the reaction parameters during gold nanoparticle synthesis. A conventional sonication bath (37 k Hz) was used for the sonochemical synthesis. This work focuses on the reaction temperature effect and the effect of sodium tartrate concentration. It was confirmed that particle size, and particle morphology is dependent of these two reaction parameters. Equally, colloidal stabilization was related to reaction temperature and sodium tartrate concentration. It was also determined that Ostwald ripening takes place during sonochemical reaction under our conditions, allowing to understand the mechanism that takes place during synthesis. Gold nanoparticles with main particle size of 17 nm were achieved by this method. Characterization techniques used: Fourier transform infrared spectra (Ftir), X-ray diffraction and Atomic Force Microscope was used in order to determine particle size of the synthetic product of reaction M10c by tapping mode. (Author)

Understanding newly discovered oscillation modes in magnetically shielded Hall thrusters utilizing state of the art high speed diagnostics.

Data.gov (United States)

National Aeronautics and Space Administration — I propose to investigate the newly discovered oscillation modes specific to Magnetically Shied (MS) Hall Effect Thrusters (HET). Although HETs are classified as a...

A high sensitivity momentum flux measuring instrument for plasma thruster exhausts and diffusive plasmas.

Science.gov (United States)

West, Michael D; Charles, Christine; Boswell, Rod W

2009-05-01

A high sensitivity momentum flux measuring instrument based on a compound pendulum has been developed for use with electric propulsion devices and radio frequency driven plasmas. A laser displacement system, which builds upon techniques used by the materials science community for surface stress measurements, is used to measure with high sensitivity the displacement of a target plate placed in a plasma thruster exhaust. The instrument has been installed inside a vacuum chamber and calibrated via two different methods and is able to measure forces in the range of 0.02-0.5 mN with a resolution of 15 microN. Measurements have been made of the force produced from the cold gas flow and with a discharge ignited using argon propellant. The plasma is generated using a Helicon Double Layer Thruster prototype. The instrument target is placed about 1 mean free path for ion-neutral charge exchange collisions downstream of the thruster exit. At this position, the plasma consists of a low density ion beam (10%) and a much larger downstream component (90%). The results are in good agreement with those determined from the plasma parameters measured with diagnostic probes. Measurements at various flow rates show that variations in ion beam velocity and plasma density and the resulting momentum flux can be measured with this instrument. The instrument target is a simple, low cost device, and since the laser displacement system used is located outside the vacuum chamber, the measurement technique is free from radio frequency interference and thermal effects. It could be used to measure the thrust in the exhaust of other electric propulsion devices and the momentum flux of ion beams formed by expanding plasmas or fusion experiments.

Study on the plasma generation characteristics of an induction-triggered coaxial pulsed plasma thruster

Science.gov (United States)

Weisheng, CUI; Wenzheng, LIU; Jia, TIAN; Xiuyang, CHEN

2018-02-01

At present, spark plugs are used to trigger discharge in pulsed plasma thrusters (PPT), which are known to be life-limiting components due to plasma corrosion and carbon deposition. A strong electric field could be formed in a cathode triple junction (CTJ) to achieve a trigger function under vacuum conditions. We propose an induction-triggered electrode structure on the basis of the CTJ trigger principle. The induction-triggered electrode structure could increase the electric field strength of the CTJ without changing the voltage between electrodes, contributing to a reduction in the electrode breakdown voltage. Additionally, it can maintain the plasma generation effect when the breakdown voltage is reduced in the discharge experiments. The induction-triggered electrode structure could ensure an effective trigger when the ablation distance of Teflon increases, and the magnetic field produced by the discharge current could further improve the plasma density and propagation velocity. The induction-triggered coaxial PPT we propose has a simplified trigger structure, and it is an effective attempt to optimize the micro-satellite thruster.

Addressing EO-1 Spacecraft Pulsed Plasma Thruster EMI Concerns

Science.gov (United States)

Zakrzwski, C. M.; Davis, Mitch; Sarmiento, Charles; Bauer, Frank H. (Technical Monitor)

2001-01-01

The Pulsed Plasma Thruster (PPT) Experiment on the Earth Observing One (EO-1) spacecraft has been designed to demonstrate the capability of a new generation PPT to perform spacecraft attitude control. Results from PPT unit level radiated electromagnetic interference (EMI) tests led to concerns about potential interference problems with other spacecraft subsystems. Initial plans to address these concerns included firing the PPT at the spacecraft level both in atmosphere, with special ground support equipment. and in vacuum. During the spacecraft level tests, additional concerns where raised about potential harm to the Advanced Land Imager (ALI). The inadequacy of standard radiated emission test protocol to address pulsed electromagnetic discharges and the lack of resources required to perform compatibility tests between the PPT and an ALI test unit led to changes in the spacecraft level validation plan. An EMI shield box for the PPT was constructed and validated for spacecraft level ambient testing. Spacecraft level vacuum tests of the PPT were deleted. Implementation of the shield box allowed for successful spacecraft level testing of the PPT while eliminating any risk to the ALI. The ALI demonstration will precede the PPT demonstration to eliminate any possible risk of damage of ALI from PPT operation.

Predicting Hall Thruster Operational Lifetime Using a Kinetic Plasma Model and a Molecular Dynamics Simulation Method, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Hall thrusters are being considered for many space missions because their high specific impulse delivers a larger payload mass fraction than chemical rockets. With a...

Control of transversal instabilities in reaction-diffusion systems

Science.gov (United States)

Totz, Sonja; Löber, Jakob; Totz, Jan Frederik; Engel, Harald

2018-05-01

In two-dimensional reaction-diffusion systems, local curvature perturbations on traveling waves are typically damped out and vanish. However, if the inhibitor diffuses much faster than the activator, transversal instabilities can arise, leading from flat to folded, spatio-temporally modulated waves and to spreading spiral turbulence. Here, we propose a scheme to induce or inhibit these instabilities via a spatio-temporal feedback loop. In a piecewise-linear version of the FitzHugh–Nagumo model, transversal instabilities and spiral turbulence in the uncontrolled system are shown to be suppressed in the presence of control, thereby stabilizing plane wave propagation. Conversely, in numerical simulations with the modified Oregonator model for the photosensitive Belousov–Zhabotinsky reaction, which does not exhibit transversal instabilities on its own, we demonstrate the feasibility of inducing transversal instabilities and study the emerging wave patterns in a well-controlled manner.

Controlled trifluoromethylation reactions of alkynes through visible-light photoredox catalysis.

Science.gov (United States)

Iqbal, Naeem; Jung, Jaehun; Park, Sehyun; Cho, Eun Jin

2014-01-07

The control of a reaction that can form multiple products is a highly attractive and challenging concept in synthetic chemistry. A set of valuable CF3 -containing molecules, namely trifluoromethylated alkenyl iodides, alkenes, and alkynes, were selectively generated from alkynes and CF3 I by environmentally benign and efficient visible-light photoredox catalysis. Subtle differences in the combination of catalyst, base, and solvent enabled the control of reactivity and selectivity for the reaction between an alkyne and CF3 I. Copyright © 2014 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

A Comprehensive Investigation of Facility Effects on the Testing of High-Power Monolithic and Clustered Hall Thruster Systems

National Research Council Canada - National Science Library

Gallimore, Alec D; Walker, Mitchell M; Beal, Brian E; Smith, Timothy B

2006-01-01

.... It is difficult for researchers to make adequate comparisons between data sets because of both differences in instrumentation and back pressures due to the wide range of facilities used in Hall thruster testing...

Synchronization criteria for generalized reaction-diffusion neural networks via periodically intermittent control.

Science.gov (United States)

Gan, Qintao; Lv, Tianshi; Fu, Zhenhua

2016-04-01

In this paper, the synchronization problem for a class of generalized neural networks with time-varying delays and reaction-diffusion terms is investigated concerning Neumann boundary conditions in terms of p-norm. The proposed generalized neural networks model includes reaction-diffusion local field neural networks and reaction-diffusion static neural networks as its special cases. By establishing a new inequality, some simple and useful conditions are obtained analytically to guarantee the global exponential synchronization of the addressed neural networks under the periodically intermittent control. According to the theoretical results, the influences of diffusion coefficients, diffusion space, and control rate on synchronization are analyzed. Finally, the feasibility and effectiveness of the proposed methods are shown by simulation examples, and by choosing different diffusion coefficients, diffusion spaces, and control rates, different controlled synchronization states can be obtained.

Autonomous Reconfigurable Control Allocation (ARCA) for Reusable Launch Vehicles

Science.gov (United States)

Hodel, A. S.; Callahan, Ronnie; Jackson, Scott (Technical Monitor)

2002-01-01

The role of control allocation (CA) in modern aerospace vehicles is to compute a command vector delta(sub c) is a member of IR(sup n(sub a)) that corresponding to commanded or desired body-frame torques (moments) tou(sub c) = [L M N](sup T) to the vehicle, compensating for and/or responding to inaccuracies in off-line nominal control allocation calculations, actuator failures and/or degradations (reduced effectiveness), or actuator limitations (rate/position saturation). The command vector delta(sub c) may govern the behavior of, e.g., acrosurfaces, reaction thrusters, engine gimbals and/or thrust vectoring. Typically, the individual moments generated in response to each of the n(sub a) commands does not lie strictly in the roll, pitch, or yaw axes, and so a common practice is to group or gang actuators so that a one-to-one mapping from torque commands tau(sub c) actuator commands delta(sub c) may be achieved in an off-line computed CA function.

Improvement of the low frequency oscillation model for Hall thrusters

Energy Technology Data Exchange (ETDEWEB)

Wang, Chunsheng, E-mail: wangcs@hit.edu.cn; Wang, Huashan [Yanshan University, College of Vehicles and Energy, Qinhuangdao 066004, Hebei (China)

2016-08-15

The low frequency oscillation of the discharge current in Hall thrusters is a major aspect of these devices that requires further study. While the existing model captures the ionization mechanism of the low frequency oscillation, it unfortunately fails to express the dynamic characteristics of the ion acceleration. The analysis in this paper shows this is because of the simplification of the electron equation, which affects both the electric field distribution and the ion acceleration process. Additionally, the electron density equation is revised and a new model that is based on the physical properties of ion movement is proposed.

Simulation of Main Plasma Parameters of a Cylindrical Asymmetric Capacitively Coupled Plasma Micro-Thruster using Computational Fluid Dynamics

Directory of Open Access Journals (Sweden)

Amelia eGreig

2015-01-01

Full Text Available Computational fluid dynamics (CFD simulations of a radio-frequency (13.56 MHz electro-thermal capacitively coupled plasma (CCP micro-thruster have been performed using the commercial CFD-ACE+ package. Standard operating conditions of a 10 W, 1.5 Torr argon discharge were used to compare with previously obtained experimental results for validation. Results show that the driving force behind plasma production within the thruster is ion-induced secondary electrons ejected from the surface of the discharge tube, accelerated through the sheath to electron temperatures up to 33.5 eV. The secondary electron coefficient was varied to determine the effect on the discharge, with results showing that full breakdown of the discharge did not occur for coefficients coefficients less than or equal to 0.01.

Controlling Chemical Reactions in Confined Environments: Water Dissociation in MOF-74

Directory of Open Access Journals (Sweden)

Erika M. A. Fuentes-Fernandez

2018-02-01

Full Text Available The confined porous environment of metal organic frameworks (MOFs is an attractive system for studying reaction mechanisms. Compared to flat oxide surfaces, MOFs have the key advantage that they exhibit a well-defined structure and present significantly fewer challenges in experimental characterization. As an example of an important reaction, we study here the dissociation of water—which plays a critical role in biology, chemistry, and materials science—in MOFs and show how the knowledge of the structure in this confined environment allows for an unprecedented level of understanding and control. In particular, combining in-situ infrared spectroscopy and first-principles calculations, we show that the water dissociation reaction can be selectively controlled inside Zn-MOF-74 by alcohol, through both chemical and physical interactions. Methanol is observed to speed up water dissociation by 25% to 100%, depending on the alcohol partial pressure. On the other hand, co-adsorption of isopropanol reduces the speed of the water reaction, due mostly to steric interactions. In addition, we also investigate the stability of the product state after the water dissociation has occurred and find that the presence of additional water significantly stabilizes the dissociated state. Our results show that precise control of reactions within nano-porous materials is possible, opening the way for advances in fields ranging from catalysis to electrochemistry and sensors.

Field emission electric propulsion thruster modeling and simulation

Science.gov (United States)

Vanderwyst, Anton Sivaram

Electric propulsion allows space rockets a much greater range of capabilities with mass efficiencies that are 1.3 to 30 times greater than chemical propulsion. Field emission electric propulsion (FEEP) thrusters provide a specific design that possesses extremely high efficiency and small impulse bits. Depending on mass flow rate, these thrusters can emit both ions and droplets. To date, fundamental experimental work has been limited in FEEP. In particular, detailed individual droplet mechanics have yet to be understood. In this thesis, theoretical and computational investigations are conducted to examine the physical characteristics associated with droplet dynamics relevant to FEEP applications. Both asymptotic analysis and numerical simulations, based on a new approach combining level set and boundary element methods, were used to simulate 2D-planar and 2D-axisymmetric probability density functions of the droplets produced for a given geometry and electrode potential. The combined algorithm allows the simulation of electrostatically-driven liquids up to and after detachment. Second order accuracy in space is achieved using a volume of fluid correction. The simulations indicate that in general, (i) lowering surface tension, viscosity, and potential, or (ii) enlarging electrode rings, and needle tips reduce operational mass efficiency. Among these factors, surface tension and electrostatic potential have the largest impact. A probability density function for the mass to charge ratio (MTCR) of detached droplets is computed, with a peak around 4,000 atoms per electron. High impedance surfaces, strong electric fields, and large liquid surface tension result in a lower MTCR ratio, which governs FEEP droplet evolution via the charge on detached droplets and their corresponding acceleration. Due to the slow mass flow along a FEEP needle, viscosity is of less importance in altering the droplet velocities. The width of the needle, the composition of the propellant, the

Magnetically Filtered Faraday Probe for Measuring the Ion Current Density Profile of a Hall Thruster

National Research Council Canada - National Science Library

Rovey, Joshua L; Walker, Mitchell L. R; Gallimore, Alec D; Peterson, Peter Y

2006-01-01

.../s. The probes are evaluated on a xenon propellant Hall thruster in the University of Michigan Large Vacuum Test Facility at operating pressures within the range of 4.4 x 10(-4) Pa Xe (3.3 x 10(-6) Torr Xe) to 1.1 10(-3) Pa Xe (8.4 x 10(-6) Torr Xe...

Metabolic control analysis of biochemical pathways based on a thermokinetic description of reaction rates

DEFF Research Database (Denmark)

Nielsen, Jens Bredal

1997-01-01

Metabolic control analysis is a powerful technique for the evaluation of flux control within biochemical pathways. Its foundation is the elasticity coefficients and the flux control coefficients (FCCs). On the basis of a thermokinetic description of reaction rates it is here shown...... that the elasticity coefficients can be calculated directly from the pool levels of metabolites at steady state. The only requirement is that one thermodynamic parameter be known, namely the reaction affinity at the intercept of the tangent in the inflection point of the curve of reaction rate against reaction...... of the thermokinetic description of reaction rates to include the influence of effecters. Here the reaction rate is written as a linear function of the logarithm of the metabolite concentrations. With this type of rate function it is shown that the approach of Delgado and Liao [Biochem. J. (1992) 282, 919-927] can...

Combined Contamination and Space Environmental Effects on Solar Cells and Thermal Control Surfaces

Science.gov (United States)

Dever, Joyce A.; Bruckner, Eric J.; Scheiman, David A.; Stidham, Curtis R.

1994-01-01

For spacecraft in low Earth orbit (LEO), contamination can occur from thruster fuel, sputter contamination products and from products of silicone degradation. This paper describes laboratory testing in which solar cell materials and thermal control surfaces were exposed to simulated spacecraft environmental effects including contamination, atomic oxygen, ultraviolet radiation and thermal cycling. The objective of these experiments was to determine how the interaction of the natural LEO environmental effects with contaminated spacecraft surfaces impacts the performance of these materials. Optical properties of samples were measured and solar cell performance data was obtained. In general, exposure to contamination by thruster fuel resulted in degradation of solar absorptance for fused silica and various thermal control surfaces and degradation of solar cell performance. Fused silica samples which were subsequently exposed to an atomic oxygen/vacuum ultraviolet radiation environment showed reversal of this degradation. These results imply that solar cells and thermal control surfaces which are susceptible to thruster fuel contamination and which also receive atomic oxygen exposure may not undergo significant performance degradation. Materials which were exposed to only vacuum ultraviolet radiation subsequent to contamination showed slight additional degradation in solar absorptance.

Two-dimensional magnetic field evolution measurements and plasma flow speed estimates from the coaxial thruster experiment

International Nuclear Information System (INIS)

Black, D.C.; Mayo, R.M.; Gerwin, R.A.; Schoenberg, K.F.; Scheuer, J.T.; Hoyt, R.P.; Henins, I.

1994-01-01

Local, time-dependent magnetic field measurements have been made in the Los Alamos coaxial thruster experiment (CTX) [C. W. Barnes et al., Phys. Fluids B 2, 1871 (1990); J. C. Fernandez et al., Nucl. Fusion 28, 1555 (1988)] using a 24 coil magnetic probe array (eight spatial positions, three axis probes). The CTX is a magnetized, coaxial plasma gun presently being used to investigate the viability of high pulsed power plasma thrusters for advanced electric propulsion. Previous efforts on this device have indicated that high pulsed power plasma guns are attractive candidates for advanced propulsion that employ ideal magnetohydrodynamic (MHD) plasma stream flow through self-formed magnetic nozzles. Indirect evidence of magnetic nozzle formation was obtained from plasma gun performance and measurements of directed axial velocities up to v z ∼10 7 cm/s. The purpose of this work is to make direct measurement of the time evolving magnetic field topology. The intent is to both identify that applied magnetic field distortion by the highly conductive plasma is occurring, and to provide insight into the details of discharge evolution. Data from a magnetic fluctuation probe array have been used to investigate the details of applied magnetic field deformation through the reconstruction of time-dependent flux profiles. Experimentally observed magnetic field line distortion has been compared to that predicted by a simple one-dimensional (1-D) model of the discharge channel. Such a comparison is utilized to estimate the axial plasma velocity in the thruster. Velocities determined in this manner are in approximate agreement with the predicted self-field magnetosonic speed and those measured by a time-of-flight spectrometer

Effects of catalyst-bed’s structure parameters on decomposition and combustion characteristics of an ammonium dinitramide (ADN)-based thruster

International Nuclear Information System (INIS)

Yu, Yu-Song; Li, Guo-Xiu; Zhang, Tao; Chen, Jun; Wang, Meng

2015-01-01

Highlights: • The decomposition and combustion process is investigated by numerical method. • Heat transfer in catalyst bed is modeled using non-isothermal and radiation model. • The wall heat transfer can impact on the distribution of temperature and species. • The value of catalyst bed length, diameter and wall thickness are optimized. - Abstract: The present investigation numerically studies the evolutions of decomposition and combustion within an ADN-based thruster, and the effects of the catalyst-bed’s three structure parameters (length, diameter, and wall thickness) on the general performance of ADN-based thruster have been systematically investigated. Based upon the calculated results, it can be known that the distribution of temperature gives a Gaussian manner at the exits of the catalyst-bed and the combustion chamber, and the temperature can be obviously effected by each the three structure parameters of the catalyst-bed. With the rise of each the three structure parameter, the temperature will first increases and decreases, and there exists an optimal design value making the temperature be the highest. Via the comparison on the maximal temperature at combustion chamber’s exit and the specific impulse, it can be obtained that the wall thickness plays an important role in the influences on the general performance of ADN-based thruster while the catalyst-bed’s length has the weak effects on the general performance among the three structure parameters.

Reaction Control Engine for Space Launch Initiative

Science.gov (United States)

2002-01-01

Engineers at the Marshall Space Flight Center (MSFC) have begun a series of engine tests on a new breed of space propulsion: a Reaction Control Engine developed for the Space Launch Initiative (SLI). The engine, developed by TRW Space and Electronics of Redondo Beach, California, is an auxiliary propulsion engine designed to maneuver vehicles in orbit. It is used for docking, reentry, attitude control, and fine-pointing while the vehicle is in orbit. The engine uses nontoxic chemicals as propellants, a feature that creates a safer environment for ground operators, lowers cost, and increases efficiency with less maintenance and quicker turnaround time between missions. Testing includes 30 hot-firings. This photograph shows the first engine test performed at MSFC that includes SLI technology. Another unique feature of the Reaction Control Engine is that it operates at dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The low-level thrust of 25 pounds of force allows the vehicle to fine-point maneuver and dock while the high-level thrust of 1,000 pounds of force is used for reentry, orbit transfer, and coarse positioning. SLI is a NASA-wide research and development program, managed by the MSFC, designed to improve safety, reliability, and cost effectiveness of space travel for second generation reusable launch vehicles.

Numerical investigation of a Hall thruster plasma

International Nuclear Information System (INIS)

Roy, Subrata; Pandey, B.P.

2002-01-01

The dynamics of the Hall thruster is investigated numerically in the framework of a one-dimensional, multifluid macroscopic description of a partially ionized xenon plasma using finite element formulation. The model includes neutral dynamics, inelastic processes, and plasma-wall interaction. Owing to disparate temporal scales, ions and neutrals have been described by set of time-dependent equations, while electrons are considered in steady state. Based on the experimental observations, a third order polynomial in electron temperature is used to calculate ionization rate. The results show that in the acceleration channel the increase in the ion number density is related to the decrease in the neutral number density. The electron and ion velocity profiles are consistent with the imposed electric field. The electron temperature remains uniform for nearly two-thirds of the channel; then sharply increases to a peak before dropping slightly at the exit. This is consistent with the predicted electron gyration velocity distribution

Method and apparatus for controlling gas evolution from chemical reactions

Science.gov (United States)

Skorpik, James R.; Dodson, Michael G.

1999-01-01

The present invention is directed toward monitoring a thermally driven gas evolving chemical reaction with an acoustic apparatus. Signals from the acoustic apparatus are used to control a heater to prevent a run-away condition. A digestion module in combination with a robotic arm further automate physical handling of sample material reaction vessels. The invention is especially useful for carrying out sample procedures defined in EPA Methods SW-846.

Carbon Back Sputter Modeling for Hall Thruster Testing

Science.gov (United States)

Gilland, James H.; Williams, George J.; Burt, Jonathan M.; Yim, John T.

2016-01-01

In support of wear testing for the Hall Effect Rocket with Magnetic Shielding (HERMeS) program, the back sputter from a Hall effect thruster plume has been modeled for the NASA Glenn Research Centers Vacuum Facility 5. The predicted wear at a near-worst case condition of 600 V, 12.5 kW was found to be on the order of 3 4 mkhour in a fully carbon-lined chamber. A more detailed numerical monte carlo code was also modified to estimate back sputter for a detailed facility and pumping configuration. This code demonstrated similar back sputter rate distributions, but is not yet accurately modeling the magnitudes. The modeling has been benchmarked to recent HERMeS wear testing, using multiple microbalance measurements. These recent measurements have yielded values, on the order of 1.5- 2 microns/khour.

Spectrum Diagnosis for Fuchsia Plume of Hall Effect Thruster with Xenon as Propellant

International Nuclear Information System (INIS)

Yu Daren; Ding Jiapeng; Dai Jingmin

2006-01-01

The colour of the Hall effect thruster's plume is often light-green, and sometimes a fuchsia plume appears during experiments. Based on a spectrum and colour analysis, and a comparison with normal plumes, a conclusion is made that the density of the Xe ions and the temperature of electrons are low when the plume appears fuchsia. In this condition, most of the components of the plume are Xe atoms, and the ionization rate of the propellant is low

Neural control of vascular reactions: impact of emotion and attention.

Science.gov (United States)

Okon-Singer, Hadas; Mehnert, Jan; Hoyer, Jana; Hellrung, Lydia; Schaare, Herma Lina; Dukart, Juergen; Villringer, Arno

2014-03-19

This study investigated the neural regions involved in blood pressure reactions to negative stimuli and their possible modulation by attention. Twenty-four healthy human subjects (11 females; age = 24.75 ± 2.49 years) participated in an affective perceptual load task that manipulated attention to negative/neutral distractor pictures. fMRI data were collected simultaneously with continuous recording of peripheral arterial blood pressure. A parametric modulation analysis examined the impact of attention and emotion on the relation between neural activation and blood pressure reactivity during the task. When attention was available for processing the distractor pictures, negative pictures resulted in behavioral interference, neural activation in brain regions previously related to emotion, a transient decrease of blood pressure, and a positive correlation between blood pressure response and activation in a network including prefrontal and parietal regions, the amygdala, caudate, and mid-brain. These effects were modulated by attention; behavioral and neural responses to highly negative distractor pictures (compared with neutral pictures) were smaller or diminished, as was the negative blood pressure response when the central task involved high perceptual load. Furthermore, comparing high and low load revealed enhanced activation in frontoparietal regions implicated in attention control. Our results fit theories emphasizing the role of attention in the control of behavioral and neural reactions to irrelevant emotional distracting information. Our findings furthermore extend the function of attention to the control of autonomous reactions associated with negative emotions by showing altered blood pressure reactions to emotional stimuli, the latter being of potential clinical relevance.

Femtosecond laser induced and controlled chemical reaction of carbon monoxide and hydrogen

CSIR Research Space (South Africa)

Du Plessis, A

2011-07-01

Full Text Available Results from experiments aimed at bimolecular chemical reaction control of CO and H2 at room temperature and pressure, without any catalyst, using shaped femtosecond laser pulses are presented. A stable reaction product (CO2) was measured after...

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

Science.gov (United States)

Shastry, Rohit; Soulas, George C.

2016-01-01

The NEXT Long-Duration Test is part of a comprehensive thruster service life assessment intended to demonstrate overall throughput capability, validate service life models, quantify wear rates as a function of time and operating condition, and identify any unknown life-limiting mechanisms. The test was voluntarily terminated in February 2014 after demonstrating 51,184 hours of high-voltage operation, 918 kg of propellant throughput, and 35.5 MN-s of total impulse. The post-test inspection of the thruster hardware began shortly afterwards with a combination of non-destructive and destructive analysis techniques, and is presently nearing completion. This paper presents relevant results of the post-test inspection for both discharge and neutralizer cathodes. Discharge keeper erosion was found to be significantly reduced from what was observed in the NEXT 2 kh wear test and NSTAR Extended Life Test, providing adequate protection of vital cathode components throughout the test with ample lifetime remaining. The area of the discharge cathode orifice plate that was exposed by the keeper orifice exhibited net erosion, leading to cathode plate material building up in the cathode-keeper gap and causing a thermally-induced electrical short observed during the test. Significant erosion of the neutralizer cathode orifice was also found and is believed to be the root cause of an observed loss in flow margin. Deposition within the neutralizer keeper orifice as well as on the downstream surface was thicker than expected, potentially resulting in a facility-induced impact on the measured flow margin from plume mode. Neutralizer keeper wall erosion on the beam side was found to be significantly lower compared to the NEXT 2 kh wear test, likely due to the reduction in beam extraction diameter of the ion optics that resulted in decreased ion impingement. Results from the post-test inspection have led to some minor thruster design improvements.

Effect of Magnetic Mirror on the Asymmetry of the Radial Profile of Near-Wall Conductivity in Hall Thrusters

International Nuclear Information System (INIS)

Yu Daren; Liu Hui; Fu Haiyang

2009-01-01

Considering the actual magnetic field configuration in a Hall thruster, the effect of magnetic mirror on the radial profile of near-wall conductivity (NWC) is studied in this paper. The plasma electron dynamic process is described by the test particle method. The Monte Carlo scheme is used to solve this model. The radial profile of electron mobility is obtained and the role of magnetic mirror in NWC is analysed both theoretically and numerically. The numerical results show that the electron mobility peak due to NWC is inversely proportional to the magnetic mirror ratio and the asymmetry of electron mobility along the radial direction gets greater when the magnetic mirror is considered. This effect indicates that apart from the disparity in the magnetic field strength, the difference in the magnetic mirror ratio near the inner and outer walls would actually augment the asymmetry of the radial profile of NWC in Hall thrusters.

Low frequency azimuthal stability of the ionization region of the Hall thruster discharge. II. Global analysis

International Nuclear Information System (INIS)

Escobar, D.; Ahedo, E.

2015-01-01

The linear stability of the Hall thruster discharge is analysed against axial-azimuthal perturbations in the low frequency range using a time-dependent 2D code of the discharge. This azimuthal stability analysis is spatially global, as opposed to the more common local stability analyses, already afforded previously (D. Escobar and E. Ahedo, Phys. Plasmas 21(4), 043505 (2014)). The study covers both axial and axial-azimuthal oscillations, known as breathing mode and spoke, respectively. The influence on the spoke instability of different operation parameters such as discharge voltage, mass flow, and thruster size is assessed by means of different parametric variations and compared against experimental results. Additionally, simplified models are used to unveil and characterize the mechanisms driving the spoke. The results indicate that the spoke is linked to azimuthal oscillations of the ionization process and to the Bohm condition in the transition to the anode sheath. Finally, results obtained from local and global stability analyses are compared in order to explain the discrepancies between both methods

Slew Maneuver Control for Spacecraft Equipped with Star Camera and Reaction Wheels

DEFF Research Database (Denmark)

Wisniewski, Rafal; Kulczycki, P.

2005-01-01

A configuration consisting of a star camera, four reaction wheels and magnetorquers for momentum unloading has become standard for many spacecraft missions. This popularity has motivated numerous agencies and private companies to initiate work on the design of an imbedded attitude control system...... realized on an integrated circuit. This paper provides an easily implementable control algorithm for this type of configuration. The paper considers two issues: slew maneuver with a feature of avoiding direct exposure of the camera's CCD chip to the Sun %, three-axis attitude control and optimal control...... torque distribution in a reaction wheel assembly. The attitude controller is synthesized applying the energy shaping technique, where the desired potential function is carefully designed using a physical insight into the nature of the problem. The system stability is thoroughly analyzed and the control...

General Attitude Control Algorithm for Spacecraft Equipped with Star Camera and Reaction Wheels

DEFF Research Database (Denmark)

Wisniewski, Rafal; Kulczycki, P.

A configuration consisting of a star camera, four reaction wheels and magnetorquers for momentum unloading has become standard for many spacecraft missions. This popularity has motivated numerous agencies and private companies to initiate work on the design of an imbedded attitude control system...... realized on an integrated circuit. This paper considers two issues: slew maneuver with a feature of avoiding direct exposure of the camera's CCD chip to the Sun %, three-axis attitude control and optimal control torque distribution in a reaction wheel assembly. The attitude controller is synthesized...

Plasma and controlled thermonuclear reaction

International Nuclear Information System (INIS)

Kapitsa, P.

1980-01-01

The principle and prospects are given of three methods of achieving controlled thermonuclear reaction. The original and so far most promising TOKAMAK method is presented invented in the USSR. Another method is the heating of a sphere about 1 mm in diameter from a mixture of deuterium and tritium by focused laser light from all sides. The third method consists in continuous plasma heating. A rope-like plasma discharge at a temperature of more than a million K results in the gas from microwave oscillations. The discharge is placed in a magnetic field and the ion temperature is increased by magneto-acoustic waves. A reactor is proposed operating on this principle and problems are pointed out which will have to be resolved. (M.S.)

Plasma and controlled thermonuclear reaction

Energy Technology Data Exchange (ETDEWEB)

Kapitsa, P

1980-06-01

The principle and prospects are given of three methods of achieving controlled thermonuclear reaction. The original and so far most promising TOKAMAK method is presented invented in the USSR. Another method is the heating of a sphere about 1 mm in diameter from a mixture of deuterium and tritium by focused laser light from all sides. The third method consists in continuous plasma heating. A rope-like plasma discharge at a temperature of more than a million K results in the gas from microwave oscillations. The discharge is placed in a magnetic field and the ion temperature is increased by magneto-acoustic waves. A reactor is proposed operating on this principle and problems are pointed out which will have to be resolved.

One-dimensional hybrid-direct kinetic simulation of the discharge plasma in a Hall thruster

International Nuclear Information System (INIS)

Hara, Kentaro; Boyd, Iain D.; Kolobov, Vladimir I.

2012-01-01

In order to model the non-equilibrium plasma within the discharge region of a Hall thruster, the velocity distribution functions (VDFs) must be obtained accurately. A direct kinetic (DK) simulation method that directly solves the plasma Boltzmann equation can achieve better resolution of VDFs in comparison to particle simulations, such as the particle-in-cell (PIC) method that inherently include statistical noise. In this paper, a one-dimensional hybrid-DK simulation, which uses a DK simulation for heavy species and a fluid model for electrons, is developed and compared to a hybrid-PIC simulation. Time-averaged results obtained from the hybrid-DK simulation are in good agreement with hybrid-PIC results and experimental data. It is shown from a comparison of using a kinetic simulation and solving the continuity equation that modeling of the neutral atoms plays an important role for simulations of the Hall thruster discharge plasma. In addition, low and high frequency plasma oscillations are observed. Although the kinetic nature of electrons is not resolved due to the use of a fluid model, the hybrid-DK model provides spatially and temporally well-resolved plasma properties and an improved resolution of VDFs for heavy species with less statistical noise in comparison to the hybrid-PIC method.

RHETT2/EPDM Hall Thruster Propulsion System Electromagnetic Compatibility Evaluation

Science.gov (United States)

Sarmiento, Charles J.; Sankovic, John M.; Freitas, Joseph; Lynn, Peter R.

1997-01-01

Electromagnetic compatibility measurements were obtained as part of the Electric Propulsion Demonstration Module (EPDM) flight qualification program. Tests were conducted on a Hall thruster system operating at a nominal 66O W discharge power. Measurements of conducted and radiated susceptibility and emissions were obtained and referenced to MEL-STD-461 C. The power processor showed some conducted susceptibility below 4 kHz for the magnet current and discharge voltage. Radiated susceptibility testing yielded a null result. Conducted emissions showed slight violations of the specified limit for MIL-461C CE03. Radiated emissions exceeded the RE02 standard at low frequencies, below 300 MHz, by up to 40 dB RV/m/MHz.

Overview of NASA Iodine Hall Thruster Propulsion System Development

Science.gov (United States)

Smith, Timothy D.; Kamhawi, Hani; Hickman, Tyler; Haag, Thomas; Dankanich, John; Polzin, Kurt; Byrne, Lawrence; Szabo, James

2016-01-01

NASA is continuing to invest in advancing Hall thruster technologies for implementation in commercial and government missions. The most recent focus has been on increasing the power level for large-scale exploration applications. However, there has also been a similar push to examine applications of electric propulsion for small spacecraft in the range of 300 kg or less. There have been several recent iodine Hall propulsion system development activities performed by the team of the NASA Glenn Research Center, the NASA Marshall Space Flight Center, and Busek Co. Inc. In particular, the work focused on qualification of the Busek 200-W BHT-200-I and development of the 600-W BHT-600-I systems. This paper discusses the current status of iodine Hall propulsion system developments along with supporting technology development efforts.

From automatic control of vessels designed to operate in offshore conditions to automatic control of a large barge convoy in a river/channel: lessons learned; Do controle automatico de embarcacoes projetadas para condicoes operacionais de offshore para o controle automatico de um grande comboio de barcacas em um rio/canal: licoes aprendidas

Energy Technology Data Exchange (ETDEWEB)

Pontes, Anselmo Carvalho [Symmetry Ltda., Rio de Janeiro, RJ (Brazil); Santos, Edson Mesquita do [Marinha do Brasil, Rio de Janeiro, RJ (Brazil)

2008-07-01

The waterway transportation sector, despite its problems and bottlenecks, is continually seeking alternatives to improve its operations. With goals of increased efficiency, safety and cost reduction, the sector is modernizing and acquiring new technology. The Dynamic Position control system for a large barge convoy is one of the new technology highlights. This innovation may make development in the Brazilian waterway transportation sector feasible by permitting navigational safety and greater flexibility in transport and distribution not only for mined products but for petroleum and its derivatives. This paper presents the development process of The Dynamic Position control system for a large barge convoy in the Parana-Paraguay River, for Rio Tinto Mining Company. The hydrodynamics effects in a river are quite different than those of the traditional offshore environment. The degree of difficulty involved in solving this problem is highly dependent on a ship's characteristics, which can be analyzed in function of their hydrodynamic coefficients. Model tests in towing tanks were made to adjust the hydrodynamic coefficients and characteristics of the thrusters. The choice of the 'backstepping' control law and observer (state estimator) presented in this paper is a result of these particular conditions. The authors point out the advantage of using this technique, since all the properties of the coefficients of the hull, propeller(s) and thruster(s) have already been tested and tuned and are able to make a reliable set-up that optimizes navigation in restricted waterways. (author)

Reaction-diffusion systems in intracellular molecular transport and control.

Science.gov (United States)

Soh, Siowling; Byrska, Marta; Kandere-Grzybowska, Kristiana; Grzybowski, Bartosz A

2010-06-07

Chemical reactions make cells work only if the participating chemicals are delivered to desired locations in a timely and precise fashion. Most research to date has focused on active-transport mechanisms, although passive diffusion is often equally rapid and energetically less costly. Capitalizing on these advantages, cells have developed sophisticated reaction-diffusion (RD) systems that control a wide range of cellular functions-from chemotaxis and cell division, through signaling cascades and oscillations, to cell motility. These apparently diverse systems share many common features and are "wired" according to "generic" motifs such as nonlinear kinetics, autocatalysis, and feedback loops. Understanding the operation of these complex (bio)chemical systems requires the analysis of pertinent transport-kinetic equations or, at least on a qualitative level, of the characteristic times of the constituent subprocesses. Therefore, in reviewing the manifestations of cellular RD, we also describe basic theory of reaction-diffusion phenomena.

Genetic control of the angular leaf spot reaction in common bean leaves and pods

Directory of Open Access Journals (Sweden)

Jerônimo Constantino Borel

2011-12-01

Full Text Available Information about genetic control of plant reaction to pathogens is essential in plant breeding programs focusing resistance. This study aimed to obtain information about genetic control of the angular leaf spot reaction in leaves and pods from common bean (Phaseolus vulgaris L. line ESAL 686. This line was crossed with cultivars Jalo EEP 558 (resistant, Cornell 49-242 (resistant and Carioca MG (susceptible. Generations F1, F2 and backcrosses (BC11 and BC21 were obtained. In the dry season (2009, parents and respective populations were evaluated for angular leaf spot reaction under field conditions. Disease severity was evaluated on leaves and pods using diagrammatic scales. Severity scores were obtained and mean and variance genetic components were estimated for both. Segregation of F2 generation was analyzed for some crosses. Different genes control angular leaf spot reaction in leaves and pods. Mean and variance components showed predominance of additive effects. Heritability was high, however, was greater on pods than on leaves which indicated that leaf reaction is more influenced by the environment.

Fault-tolerant Control of Unmanned Underwater Vehicles with Continuous Faults: Simulations and Experiments

Directory of Open Access Journals (Sweden)

Qian Liu

2010-02-01

Full Text Available A novel thruster fault diagnosis and accommodation method for open-frame underwater vehicles is presented in the paper. The proposed system consists of two units: a fault diagnosis unit and a fault accommodation unit. In the fault diagnosis unit an ICMAC (Improved Credit Assignment Cerebellar Model Articulation Controllers neural network information fusion model is used to realize the fault identification of the thruster. The fault accommodation unit is based on direct calculations of moment and the result of fault identification is used to find the solution of the control allocation problem. The approach resolves the continuous faulty identification of the UV. Results from the experiment are provided to illustrate the performance of the proposed method in uncertain continuous faulty situation.

Fault-tolerant Control of Unmanned Underwater Vehicles with Continuous Faults: Simulations and Experiments

Directory of Open Access Journals (Sweden)

Qian Liu

2009-12-01

Full Text Available A novel thruster fault diagnosis and accommodation method for open-frame underwater vehicles is presented in the paper. The proposed system consists of two units: a fault diagnosis unit and a fault accommodation unit. In the fault diagnosis unit an ICMAC (Improved Credit Assignment Cerebellar Model Articulation Controllers neural network information fusion model is used to realize the fault identification of the thruster. The fault accommodation unit is based on direct calculations of moment and the result of fault identification is used to find the solution of the control allocation problem. The approach resolves the continuous faulty identification of the UV. Results from the experiment are provided to illustrate the performance of the proposed method in uncertain continuous faulty situation.

A General Strategy for Nanohybrids Synthesis via Coupled Competitive Reactions Controlled in a Hybrid Process

OpenAIRE

Wang, Rongming; Yang, Wantai; Song, Yuanjun; Shen, Xiaomiao; Wang, Junmei; Zhong, Xiaodi; Li, Shuai; Song, Yujun

2015-01-01

A new methodology based on core alloying and shell gradient-doping are developed for the synthesis of nanohybrids, realized by coupled competitive reactions, or sequenced reducing-nucleation and co-precipitation reaction of mixed metal salts in a microfluidic and batch-cooling process. The latent time of nucleation and the growth of nanohybrids can be well controlled due to the formation of controllable intermediates in the coupled competitive reactions. Thus, spatiotemporal-resolved synthesi...

Hall effect thruster with an AlN chamber

International Nuclear Information System (INIS)

Barral, S.; Jayet, Y.; Mazouffre, S.; Veron, E.; Echegut, P.; Dudeck, M.

2005-01-01

The plasma discharge of a Hall-effect thruster (SPT) is strongly depending of the plasma-insulated wall interactions. These interactions are mainly related to the energy deposition, potential sheath effect and electron secondary emission rate (e.s.e.). In usual SPT, the annular channel is made of BN-SiO 2 . The SPT100-ML (laboratory model will be tested with an AlN chamber in the French test facility Pivoine in the laboratoire d'Aerothermique (Orleans-France). The different parameters such as discharge current, thrust, plasma oscillations and wall temperature will studied for several operating conditions. The results will be compared with a fluid model developed in IPPT (Warsaw-Poland) taking into account electron emission from the internal and external walls and using previous experimental measurements of e.s.e. for AlN from ONERA (Toulouse-France). The surface state of AlN will be analysed before and after experiments by an Environmental Scanning Electron Microscope and by a Strength Electron Microscope. (author)

Chemical-Reaction-Controlled Phase Separated Drops: Formation, Size Selection, and Coarsening

Science.gov (United States)

Wurtz, Jean David; Lee, Chiu Fan

2018-02-01

Phase separation under nonequilibrium conditions is exploited by biological cells to organize their cytoplasm but remains poorly understood as a physical phenomenon. Here, we study a ternary fluid model in which phase-separating molecules can be converted into soluble molecules, and vice versa, via chemical reactions. We elucidate using analytical and simulation methods how drop size, formation, and coarsening can be controlled by the chemical reaction rates, and categorize the qualitative behavior of the system into distinct regimes. Ostwald ripening arrest occurs above critical reaction rates, demonstrating that this transition belongs entirely to the nonequilibrium regime. Our model is a minimal representation of the cell cytoplasm.

Global characteristics of an ATON stationary plasma thruster operating with krypton and xenon

International Nuclear Information System (INIS)

Bugrova, A.I.; Lipatov, A.S.; Solomatina, L.V.; Morozov, A.I.

2002-01-01

Paper contains the experimental results of operation of the ATON plasma thruster operating with krypton and xenon. It is shown that consumption of a working gas for consumption of a working gas the krypton base thrust is higher in contrast to xenon base one at lower efficiency. In case of krypton use one obtained the efficiency constituting ∼ 60% at specific pulse reaching 3000 s. Jet divergence in case of krypton use is ∼ ± 22 deg in contrast to ∼ ± 11 deg in case of xenon use [ru

Comparison of Medium Power Hall Effect Thruster Ion Acceleration for Krypton and Xenon Propellants

Science.gov (United States)

2016-09-14

Pumping is provided by four single-stage cryogenic panels (single-stage cold heads at 25 K) and one 50 cm two stage cryogenic pump (12 K). This vacuum...test chamber has a mea- sured pumping speed of 36 kL/s on xenon. The Hall thruster used in this study is a medium power laboratory Hall effect...The first compo- nent passes through a krypton opto-galvanic cell and is terminated by a beam dump . The opto-galvanic cell current is capacitively

Nanoscale control of reversible chemical reaction between fullerene C60 molecules using scanning tunneling microscope.

Science.gov (United States)

Nakaya, Masato; Kuwahara, Yuji; Aono, Masakazu; Nakayama, Tomonobu

2011-04-01

The nanoscale control of reversible chemical reactions, the polymerization and depolymerization between C60 molecules, has been investigated. Using a scanning tunneling microscope (STM), the polymerization and depolymerization can be controlled at designated positions in ultrathin films of C60 molecules. One of the two chemical reactions can be selectively induced by controlling the sample bias voltage (V(s)); the application of negative and positive values of V(s) results in polymerization and depolymerization, respectively. The selectivity between the two chemical reactions becomes extremely high when the thickness of the C60 film increases to more than three molecular layers. We conclude that STM-induced negative and positive electrostatic ionization are responsible for the control of the polymerization and depolymerization, respectively.

Momentum Management for the NASA Near Earth Asteroid Scout Solar Sail Mission

Science.gov (United States)

Heaton, Andrew; Diedrich, Benjamin L.; Orphee, Juan; Stiltner, Brandon; Becker, Christopher

2017-01-01

The Momentum Management (MM) system is described for the NASA Near Earth Asteroid Scout (NEA Scout) cubesat solar sail mission. Unlike many solar sail mission proposals that used solar torque as the primary or only attitude control system, NEA Scout uses small reaction wheels (RW) and a reaction control system (RCS) with cold gas thrusters, as described in the abstract "Solar Sail Attitude Control System for Near Earth Asteroid Scout Cubesat Mission." The reaction wheels allow fine pointing and higher rates with low mass actuators to meet the science, communication, and trajectory guidance requirements. The MM system keeps the speed of the wheels within their operating margins using a combination of solar torque and the RCS.

Pump-shaped dump optimal control reveals the nuclear reaction pathway of isomerization of a photoexcited cyanine dye.

Science.gov (United States)

Dietzek, Benjamin; Brüggemann, Ben; Pascher, Torbjörn; Yartsev, Arkady

2007-10-31

Using optimal control as a spectroscopic tool we decipher the details of the molecular dynamics of the essential multidimensional excited-state photoisomerization - a fundamental chemical reaction of key importance in biology. Two distinct nuclear motions are identified in addition to the overall bond-twisting motion: Initially, the reaction is dominated by motion perpendicular to the torsion coordinate. At later times, a second optically active vibration drives the system along the reaction path to the bottom of the excited-state potential. The time scales of the wavepacket motion on a different part of the excited-state potential are detailed by pump-shaped dump optimal control. This technique offers new means to control a chemical reaction far from the Franck-Condon point of absorption and to map details of excited-state reaction pathways revealing unique insights into the underlying reaction mechanism.

E × B electron drift instability in Hall thrusters: Particle-in-cell simulations vs. theory

Science.gov (United States)

Boeuf, J. P.; Garrigues, L.

2018-06-01

The E × B Electron Drift Instability (E × B EDI), also called Electron Cyclotron Drift Instability, has been observed in recent particle simulations of Hall thrusters and is a possible candidate to explain anomalous electron transport across the magnetic field in these devices. This instability is characterized by the development of an azimuthal wave with wavelength in the mm range and velocity on the order of the ion acoustic velocity, which enhances electron transport across the magnetic field. In this paper, we study the development and convection of the E × B EDI in the acceleration and near plume regions of a Hall thruster using a simplified 2D axial-azimuthal Particle-In-Cell simulation. The simulation is collisionless and the ionization profile is not-self-consistent but rather is given as an input parameter of the model. The aim is to study the development and properties of the instability for different values of the ionization rate (i.e., of the total ion production rate or current) and to compare the results with the theory. An important result is that the wavelength of the simulated azimuthal wave scales as the electron Debye length and that its frequency is on the order of the ion plasma frequency. This is consistent with the theory predicting destruction of electron cyclotron resonance of the E × B EDI in the non-linear regime resulting in the transition to an ion acoustic instability. The simulations also show that for plasma densities smaller than under nominal conditions of Hall thrusters the field fluctuations induced by the E × B EDI are no longer sufficient to significantly enhance electron transport across the magnetic field, and transit time instabilities develop in the axial direction. The conditions and results of the simulations are described in detail in this paper and they can serve as benchmarks for comparisons between different simulation codes. Such benchmarks would be very useful to study the role of numerical noise (numerical

Triana Safehold: A New Gyroless, Sun-Pointing Attitude Controller

Science.gov (United States)

Chen, J.; Morgenstern, Wendy; Garrick, Joseph

2001-01-01

Triana is a single-string spacecraft to be placed in a halo orbit about the sun-earth Ll Lagrangian point. The Attitude Control Subsystem (ACS) hardware includes four reaction wheels, ten thrusters, six coarse sun sensors, a star tracker, and a three-axis Inertial Measuring Unit (IMU). The ACS Safehold design features a gyroless sun-pointing control scheme using only sun sensors and wheels. With this minimum hardware approach, Safehold increases mission reliability in the event of a gyroscope anomaly. In place of the gyroscope rate measurements, Triana Safehold uses wheel tachometers to help provide a scaled estimation of the spacecraft body rate about the sun vector. Since Triana nominally performs momentum management every three months, its accumulated system momentum can reach a significant fraction of the wheel capacity. It is therefore a requirement for Safehold to maintain a sun-pointing attitude even when the spacecraft system momentum is reasonably large. The tachometer sun-line rate estimation enables the controller to bring the spacecraft close to its desired sun-pointing attitude even with reasonably high system momentum and wheel drags. This paper presents the design rationale behind this gyroless controller, stability analysis, and some time-domain simulation results showing performances with various initial conditions. Finally, suggestions for future improvements are briefly discussed.

Attitude Dynamics and Stability of a Simple Solar Photon Thruster

Directory of Open Access Journals (Sweden)

Anna D. Guerman

2013-01-01

Full Text Available This paper is dedicated to the development of a model of the attitude dynamics for a nonideal Simple Solar Photon Thruster (SSPT and to the analysis of sailcraft motions with respect to their centre of mass. Derivation of the expressions for force and torque due to solar radiation that is valid for the case, when there is a misalignment of the SSPT axis with the sun direction, is followed by study of sailcraft dynamics and stability properties. Analysis of stability shows that an ideally reflecting sail is unstable, while for a sailcraft with nonideal collector, the symmetry axis is stable with respect to the Sun direction for large variety of system parameters. The motion around symmetry axis is always unstable and requires an active stabilizer.

Design of an adaptive backstepping controller for auto-berthing a cruise ship under wind loads

Directory of Open Access Journals (Sweden)

Park Jong-Yong

2014-06-01

Full Text Available The auto-berthing of a ship requires excellent control for safe accomplishment. Crabbing, which is the pure sway motion of a ship without surge velocity, can be used for this purpose. Crabbing is induced by a peculiar operation procedure known as the push-pull mode. When a ship is in the push-pull mode, an interacting force is induced by complex turbulent flow around the ship generated by the propellers and side thrusters. In this paper, three degrees of freedom equations of the motions of crabbing are derived. The equations are used to apply the adaptive backstepping control method to the auto-berthing controller of a cruise ship. The controller is capable of handling the system nonlinearity and uncertainty of the berthing process. A control allocation algorithm for a ship equipped with two propellers and two side thrusters is also developed, the performance of which is validated by simulation of auto-berthing.

Inflight performance of the Ulysses reaction control system

Science.gov (United States)

McGarry, Andrew; Berry, William; Parker, David

1997-01-01

The Ulysses spacecraft has been exploring the heliosphere since October 1990 in a six-year polar orbit. Despite varying operational demands, the pressure-fed monopropellant hydrazine reaction control system (RCS) has experienced few problems. The observed anomalies, having minimal operational impact, include plume impingement effects, electrical power overload effects and hydrazine gas generation effects. These anomalies are presented and discussed, with emphasis on the first observation of gas in the hydrazine propellant. The relatively low gas generation rate is attributed to: the use of high purity hydrazine; the configuration of the spin-stabilized spacecraft; the extensive use of titanium alloys; and the efficiency of the thermal control of the propellant tank which maintains a temperature of 21 C.

Artificial control of biocatalytic reaction; Seitai shokubai hanno no jin`iteki seigyo

Energy Technology Data Exchange (ETDEWEB)

Nakamura, K. [Kyoto University, Kyoto (Japan). Institute for Chemical Research

1997-08-25

Selective composition of optically active compounds by use of biocatalysts is discussed. No search is made for any particular microbes or enzymes, but predetermined ones are used. For an increase in the selective yield of L-type carnitine by reducing 4-chloroacetoacetic acid ester using baker`s yeast, the ester length should be enlarged to that of octyl ester. Just as in this case, steric control by ground substance modification is often effective. Lipase helps on esterification which is contrary to hydrolysis in an organic solvent and, even in the optical division in this process, steric control by ground substance modification (for example by changing the structure of the acyl section) is effective. Immobilization of biocatalysts for use in reaction occasionally exerts some effect on stereoselectivity. Two types of enzymes may be participating in a reaction and inhibiting selectivity, and then a two-layer system of water and organic solvent may be effective in performing steric control over the situation. Another measure is to inhibit the activity of either of the two enzymes by use of a selective inhibitor utilizing enzyme reaction. The kind of solvent is also an influential factor. 11 refs., 7 figs.

Modeling the Influence of Diffusion-Controlled Reactions and Residual Termination and Deactivation on the Rate and Control of Bulk ATRP at High Conversions

Directory of Open Access Journals (Sweden)

Ali Mohammad Rabea

2015-04-01

Full Text Available In high-conversion atom transfer radical polymerization (ATRP, all the reactions, such as radical termination, radical deactivation, dormant chain activation, monomer propagation, etc. could become diffusion controlled sooner or later, depending on relative diffusivities of the involved reacting species. These diffusion-controlled reactions directly affect the rate of polymerization and the control of polymer molecular weight. A model is developed to investigate the influence of diffusion-controlled reactions on the high conversion ATRP kinetics. Model simulation reveals that diffusion-controlled termination slightly increases the rate, but it is the diffusion-controlled deactivation that causes auto-acceleration in the rate (“gel effect” and loss of control. At high conversions, radical chains are “trapped” because of high molecular weight. However, radical centers can still migrate through (1 radical deactivation–activation cycles and (2 monomer propagation, which introduce “residual termination” reactions. It is found that the “residual termination” does not have much influence on the polymerization kinetics. The migration of radical centers through propagation can however facilitate catalytic deactivation of radicals, which improves the control of polymer molecular weight to some extent. Dormant chain activation and monomer propagation also become diffusion controlled and finally stop the polymerization when the system approaches its glass state.

Effect of dust on tilted electrostatic resistive instability in a Hall thruster

Science.gov (United States)

Tyagi, Jasvendra; Singh, Sukhmander; Malik, Hitendra K.

2018-03-01

Effect of negatively charged dust on resistive instability corresponding to the electrostatic wave is investigated in a Hall thruster plasma when this purely azimuthal wave is tilted and strong axial component of wave vector is developed. Analytical calculations are done to obtain the relevant dispersion equation, which is solved numerically to investigate the growth rate of the instability. The magnitude of the growth rate in the plasma having dust particles is found to be much smaller than the case of pure plasma. However, the instability grows faster for the increasing dust density and the higher charge on the dust particles. The higher magnetic field is also found to support the instability.

Multiphoton control of the 1,3-cyclohexadiene ring-opening reaction in the presence of competing solvent reactions.

Science.gov (United States)

Carroll, Elizabeth C; White, James L; Florean, Andrei C; Bucksbaum, Philip H; Sension, Roseanne J

2008-07-31

Although physical chemistry has often concentrated on the observation and understanding of chemical systems, the defining characteristic of chemistry remains the direction and control of chemical reactivity. Optical control of molecular dynamics, and thus of chemical reactivity provides a path to use photon energy as a smart reagent in a chemical system. In this paper, we discuss recent research in this field in the context of our studies of the multiphoton optical control of the photo-initiated ring-opening reaction of 1,3-cyclohexadiene (CHD) to form 1,3,5- cis-hexatriene (Z-HT). Closed-loop feedback and learning algorithms are able to identify pulses that increase the desired target state by as much as a factor of two. Mechanisms for control are discussed through the influence of the intensity dependence, the nonlinear power spectrum, and the projection of the pulses onto low orders of polynomial phase. Control measurements in neat solvents demonstrate that competing solvent fragmentation reactions must also be considered. In particular, multiphoton excitation of cyclohexane alone is capable of producing hexatriene. Statistical analyses of data sets obtained in learning algorithm searches in neat cyclohexane and for CHD in hexane and cyclohexane highlight the importance of linear and quadratic chirp, while demonstrating that the control features are not so easily defined. Higher order phase components are also important. On the basis of these results the involvement of low-frequency ground-state vibrational modes is proposed. When the population is transferred to the excited state, momentum along the torsional coordinate may keep the wave packet localized as it moves toward the conical intersections controlling the yield of Z-HT.

Fault Tolerant Position-mooring Control for Offshore Vessels

DEFF Research Database (Denmark)

Blanke, Mogens; Nguyen, Trong Dong

2018-01-01

Fault-tolerance is crucial to maintain safety in offshore operations. The objective of this paper is to show how systematic analysis and design of fault-tolerance is conducted for a complex automation system, exemplified by thruster assisted Position-mooring. Using redundancy as required....... Functional faults that are only detectable, are rendered isolable through an active isolation approach. Once functional faults are isolated, they are handled by fault accommodation techniques to meet overall control objectives specified by class requirements. The paper illustrates the generic methodology...... by a system to handle faults in mooring lines, sensors or thrusters. Simulations and model basin experiments are carried out to validate the concept for scenarios with single or multiple faults. The results demonstrate that enhanced availability and safety are obtainable with this design approach. While...

On-Orbit Propulsion OMS/RCS

Science.gov (United States)

Hurlbert, Eric A.

2001-01-01

This slide presentation reviews the Space Shuttle's On-Orbit Propulsion systems: the Orbital Maneuvering System (OMS) and the Reaction Control System (RCS). The functions of each of the systems is described, and the diagrams of the systems are presented. The OMS/RCS thruster is detailed and a trade study comparison of non-toxic propellants is presented.

High-thrust and low-power operation of a 30-cm-diameter mercury ion thruster

Science.gov (United States)

Beattie, J. R.; Kami, S.

1981-01-01

An investigation of a 30-cm-diameter mercury ion thruster designed for high-thrust and low-power operation is described. Experimental results are presented which indicate that good performance and long lifetime are achieved by using a boundary magnetic field arrangement to confine the ionizing electrons. Details of advanced ion-optics designs are discussed, and performance measurements obtained with an advanced two-grid ion-optics assembly are presented. Scaling of the state-of-the-art hollow cathode for higher emission-current capability is described, and performance and lifetime measurements are presented for the scaled cathode.

The Maillard reaction and its control during food processing. The potential of emerging technologies.

Science.gov (United States)

Jaeger, H; Janositz, A; Knorr, D

2010-06-01

The Maillard reaction between reducing sugars and amino acids is a common reaction in foods which undergo thermal processing. Desired consequences like the formation of flavor and brown color of some cooked foods but also the destruction of essential amino acids and the production of anti-nutritive compounds require the consideration of the Maillard reaction and relevant mechanisms for its control. This paper aims to exemplify the recent advances in food processing with regard to the controllability of heat-induced changes in the food quality. Firstly, improved thermal technologies, such as ohmic heating, which allows direct heating of the product and overcoming the heat transfer limitations of conventional thermal processing are presented in terms of their applicability to reduce the thermal exposure during food preservation. Secondly, non-thermal technologies such as high hydrostatic pressure and pulsed electric fields and their ability to extend the shelf life of food products without the application of heat, thus also preserving the quality attributes of the food, will be discussed. Finally, an innovative method for the removal of Maillard reaction substrates in food raw materials by the application of pulsed electric field cell disintegration and extraction as well as enzymatic conversion is presented in order to demonstrate the potential of the combination of processes to control the occurrence of the Maillard reaction in food processing. (c) 2009 Elsevier Masson SAS. All rights reserved.

High thrust-to-power ratio micro-cathode arc thruster

Directory of Open Access Journals (Sweden)

Joseph Lukas

2016-02-01

Full Text Available The Micro-Cathode Arc Thruster (μCAT is an electric propulsion device that ablates solid cathode material, through an electrical vacuum arc discharge, to create plasma and ultimately produce thrust in the μN to mN range. About 90% of the arc discharge current is conducted by electrons, which go toward heating the anode and contribute very little to thrust, with only the remaining 10% going toward thrust in the form of ion current. A preliminary set of experiments were conducted to show that, at the same power level, thrust may increase by utilizing an ablative anode. It was shown that ablative anode particles were found on a collection plate, compared to no particles from a non-ablative anode, while another experiment showed an increase in ion-to-arc current by approximately 40% at low frequencies compared to the non-ablative anode. Utilizing anode ablation leads to an increase in thrust-to-power ratio in the case of the μCAT.

Ground reaction force comparison of controlled resistance methods to isoinertial loading of the squat exercise - biomed 2010.

Science.gov (United States)

Paulus, David C; Reynolds, Michael C; Schilling, Brian K

2010-01-01

The ground reaction force during the concentric (raising) portion of the squat exercise was compared to that of isoinertial loading (free weights) for three pneumatically controlled resistance methods: constant resistance, cam force profile, and proportional force control based on velocity. Constant force control showed lower ground reaction forces than isoinertial loading throughout the range of motion (ROM). The cam force profile exhibited slightly greater ground reaction forces than isoinertial loading at 10 and 40% ROM with fifty-percent greater loading at 70% ROM. The proportional force control consistently elicited greater ground reaction force than isoinertial loading, which progressively ranged from twenty to forty percent increase over isoinertial loading except for being approximately equal at 85% ROM. Based on these preliminary results, the proportional control shows the most promise for providing loading that is comparable in magnitude to isoinertial loading. This technology could optimize resistance exercise for sport-specific training or as a countermeasure to atrophy during spaceflight.

The Iodine Satellite (iSAT) Hall Thruster Demonstration Mission Concept and Development

Science.gov (United States)

Dankanich, John W.; Polzin, Kurt A.; Calvert, Derek; Kamhawi, Hani

2014-01-01

The use of iodine propellant for Hall thrusters has been studied and proposed by multiple organizations due to the potential mission benefits over xenon. In 2013, NASA Marshall Space Flight Center competitively selected a project for the maturation of an iodine flight operational feed system through the Technology Investment Program. Multiple partnerships and collaborations have allowed the team to expand the scope to include additional mission concept development and risk reduction to support a flight system demonstration, the iodine Satellite (iSAT). The iSAT project was initiated and is progressing towards a technology demonstration mission preliminary design review. The current status of the mission concept development and risk reduction efforts in support of this project is presented.

Analysis on oscillating actuator frequency influence of the fluid flow characterization for 2D contractile water jet thruster

Science.gov (United States)

Shaari, M. F.; Abu Bakar, H.; Nordin, N.; Saw, S. K.; Samad, Z.

2013-12-01

Contractile body is an alternative mechanism instead of rotating blade propeller to generate water jet for locomotion. The oscillating motion of the actuator at different frequencies varies the pressure and volume of the pressure chamber in time to draw in and jet out the water at a certain mass flow rate. The aim of this research was to analyze the influence of the actuating frequency of the fluid flow in the pressure chamber of the thruster during this inflation-deflation process. A 70mm × 70mm × 18mm (L × W × T) 2D water jet thruster was fabricated for this purpose. The contractile function was driven using two lateral pneumatic actuators where the fluid flow analysis was focused on the X-Y plane vector. Observation was carried out using a video camera and Matlab image measurement technique to determine the volume of the flowing mass. The result demonstrated that the greater actuating frequency decreases the fluid flow rate and the Reynolds number. This observation shows that the higher frequency would give a higher mass flow rate during water jet generation.

Directing Reaction Pathways through Controlled Reactant Binding at Pd-TiO2 Interfaces.

Science.gov (United States)

Zhang, Jing; Wang, Bingwen; Nikolla, Eranda; Medlin, J Will

2017-06-01

Recent efforts to design selective catalysts for multi-step reactions, such as hydrodeoxygenation (HDO), have emphasized the preparation of active sites at the interface between two materials having different properties. However, achieving precise control over interfacial properties, and thus reaction selectivity, has remained a challenge. Here, we encapsulated Pd nanoparticles (NPs) with TiO 2 films of regulated porosity to gain a new level of control over catalyst performance, resulting in essentially 100 % HDO selectivity for two biomass-derived alcohols. This catalyst also showed exceptional reaction specificity in HDO of furfural and m-cresol. In addition to improving HDO activity by maximizing the interfacial contact between the metal and metal oxide sites, encapsulation by the nanoporous oxide film provided a significant selectivity boost by restricting the accessible conformations of aromatics on the surface. © 2017 Wiley-VCH Verlag GmbH & Co. KGaA, Weinheim.

High Voltage Solar Array Arc Testing for a Direct Drive Hall Effect Thruster System

Science.gov (United States)

Schneider, Todd; Carruth, M. R., Jr.; Vaughn, J. A.; Jongeward, G. A.; Mikellides, I. G.; Ferguson, D.; Kerslake, T. W.; Peterson, T.; Snyder, D.; Hoskins, A.

2004-01-01

The deleterious effects of spacecraft charging are well known, particularly when the charging leads to arc events. The damage that results from arcing can severely reduce system lifetime and even cause critical system failures. On a primary spacecraft system such as a solar array, there is very little tolerance for arcing. Motivated by these concerns, an experimental investigation was undertaken to determine arc thresholds for a high voltage (200-500 V) solar array in a plasma environment. The investigation was in support of a NASA program to develop a Direct Drive Hall-Effect Thruster (D2HET) system. By directly coupling the solar array to a Hall-effect thruster, the D2HET program seeks to reduce mass, cost and complexity commonly associated with the power processing in conventional power systems. In the investigation, multiple solar array technologies and configurations were tested. The cell samples were biased to a negative voltage, with an applied potential difference between them, to imitate possible scenarios in solar array strings that could lead to damaging arcs. The samples were tested in an environment that emulated a low-energy, HET-induced plasma. Short duration trigger arcs as well as long duration sustained arcs were generated. Typical current and voltage waveforms associated with the arc events are presented. Arc thresholds are also defined in terms of voltage, current and power. The data will be used to propose a new, high-voltage (greater than 300 V) solar array design for which the likelihood of damage from arcing is minimal.

Recent activities in the development of the MOA thruster

Science.gov (United States)

Frischauf, Norbert; Hettmer, Manfred; Grassauer, Andreas; Bartusch, Tobias; Koudelka, Otto

2008-07-01

More than 60 years after the later Nobel laureate Hannes Alfvén had published a letter stating that oscillating magnetic fields can accelerate ionised matter via magneto-hydrodynamic interactions in a wave like fashion, the technical implementation of Alfvén waves for propulsive purposes has been proposed, patented and examined for the first time by a group of inventors. The name of the concept, utilising Alfvén waves to accelerate ionised matter for propulsive purposes, is MOA-magnetic field oscillating amplified thruster. Alfvén waves are generated by making use of two coils, one being permanently powered and serving also as magnetic nozzle, the other one being switched on and off in a cyclic way, deforming the field lines of the overall system. It is this deformation that generates Alfvén waves, which are in the next step used to transport and compress the propulsive medium, in theory leading to a propulsion system with a much higher performance than any other electric propulsion system. Based on computer simulations, which were conducted to get a first estimate on the performance of the system, MOA is a corrosion free and highly flexible propulsion system, whose performance parameters might easily be adapted in flight, by changing the mass flow and/or the power level. As such the system is capable to deliver a maximum specific impulse of 13 116 s (12.87 mN) at a power level of 11.16 kW, using Xe as propellant, but can also be attuned to provide a thrust of 236.5 mN (2411 s) at 6.15 kW of power. First tests-that are further described in this paper-have been conducted successfully and underline the feasibility of the concept. While space propulsion is expected to be the prime application for MOA and is supported by numerous applications such as Solar and/or Nuclear Electric Propulsion or even as an "afterburner system" for nuclear thermal propulsion, other terrestrial applications can be thought of as well, making the system highly suited for a common space

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

Science.gov (United States)

Spektor, R.; Tighe, W. G.; Kamhawi, H.

2016-01-01

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

Controlling Behaviors in Middle School Youth's Dating Relationships: Reactions and Help-Seeking Behaviors

Science.gov (United States)

Elias-Lambert, Nada; Black, Beverly M.; Chigbu, Kingsley U.

2014-01-01

This exploratory study examined middle school students' (N = 380) help-seeking behaviors and other reactions to controlling behaviors in their dating relationships. Over three-fourths of the participants perpetrated and were victimized by controlling behaviors in their dating relationships. Youth used emotional/verbal and dominance/isolation forms…

Localized temperature and chemical reaction control in nanoscale space by nanowire array.

Science.gov (United States)

Jin, C Yan; Li, Zhiyong; Williams, R Stanley; Lee, K-Cheol; Park, Inkyu

2011-11-09

We introduce a novel method for chemical reaction control with nanoscale spatial resolution based on localized heating by using a well-aligned nanowire array. Numerical and experimental analysis shows that each individual nanowire could be selectively and rapidly Joule heated for local and ultrafast temperature modulation in nanoscale space (e.g., maximum temperature gradient 2.2 K/nm at the nanowire edge; heating/cooling time chemical reactions such as polymer decomposition/cross-linking and direct and localized hydrothermal synthesis of metal oxide nanowires were demonstrated.

Measurements and calculations of 10B(n,He) reaction rates in a control rod in ZPPR

International Nuclear Information System (INIS)

Brumbach, S.B.; Collins, P.J.; Grasseschi, G.L.; Oliver, B.M.

1986-01-01

The helium accumulation fluence monitor (HAFM) technique has been used to measure the 10 B(n,He) reaction rate within B 4 C pellets in a control rod in ZPPR. Knowledge of this reaction rate is important to control rod design studies because helium production leads to control rod swelling, buildup of gas pressure and a reduction in thermal conductivity which can limit the lifetime of a control rod. We believe these to be the first measurements of boron capture within boron pins in a fast reactor spectrum. Previously reported measurements used 235 U foils to measure fission rates in a control rod, and to infer boron capture rates

Thermally activated reaction–diffusion-controlled chemical bulk reactions of gases and solids

Directory of Open Access Journals (Sweden)

S. Möller

2015-01-01

Full Text Available The chemical kinetics of the reaction of thin films with reactive gases is investigated. The removal of thin films using thermally activated solid–gas to gas reactions is a method to in-situ control deposition inventory in vacuum and plasma vessels. Significant scatter of experimental deposit removal rates at apparently similar conditions was observed in the past, highlighting the need for understanding the underlying processes. A model based on the presence of reactive gas in the films bulk and chemical kinetics is presented. The model describes the diffusion of reactive gas into the film and its chemical interaction with film constituents in the bulk using a stationary reaction–diffusion equation. This yields the reactive gas concentration and reaction rates. Diffusion and reaction rate limitations are depicted in parameter studies. Comparison with literature data on tokamak co-deposit removal results in good agreement of removal rates as a function of pressure, film thickness and temperature.

Ion velocities in a micro-cathode arc thruster

International Nuclear Information System (INIS)

Zhuang Taisen; Shashurin, Alexey; Keidar, Michael; Beilis, Isak

2012-01-01

Ion velocities in the plasma jet generated by the micro-cathode arc thruster are studied by means of time-of-flight method using enhanced ion detection system (EIDS). The EIDS triggers perturbations (spikes) on arc current waveform, and the larger current in the spike generates denser plasma bunches propagating along with the mainstream plasma. The EIDS utilizes double electrostatic probes rather than single probes. The average Ti ion velocity is measured to be around 2×10 4 m/s without a magnetic field. It was found that the application of a magnetic field does not change ion velocities in the interelectrode region while leads to ion acceleration in the free expanding plasma plume by a factor of about 2. Ion velocities of about 3.5×10 4 m/s were detected for the magnetic field of about 300 mT at distance of about 100–200 mm from the cathode. It is proposed that plasma is accelerated due to Lorentz force. The average thrust is calculated using the ion velocity measurements and the cathode mass consumption rate, and its increase with the magnetic field is demonstrated.

Diffusion-controlled reactions modeling in Geant4-DNA

Energy Technology Data Exchange (ETDEWEB)

Karamitros, M., E-mail: matkara@gmail.com [CNRS, IN2P3, CENBG, UMR 5797, F-33170 Gradignan (France); CNRS, INCIA, UMR 5287, F-33400 Talence (France); Luan, S. [University of New Mexico, Department of Computer Science, Albuquerque, NM (United States); Bernal, M.A. [Instituto de Física Gleb Wataghin, Universidade Estadual de Campinas, SP (Brazil); Allison, J. [Geant4 Associates International Ltd (United Kingdom); Baldacchino, G. [CEA Saclay, IRAMIS, LIDYL, Radiation Physical Chemistry Group, F-91191 Gif sur Yvette Cedex (France); CNRS, UMR3299, SIS2M, F-91191 Gif sur Yvette Cedex (France); Davidkova, M. [Nuclear Physics Institute of the ASCR, Prague (Czech Republic); Francis, Z. [Saint Joseph University, Faculty of Sciences, Department of Physics, Mkalles, Beirut (Lebanon); Friedland, W. [Helmholtz Zentrum München, German Research Center for Environmental Health, Institute of Radiation Protection, Ingolstädter Landstr. 1, 85764 Neuherberg (Germany); Ivantchenko, V. [Ecoanalytica, 119899 Moscow (Russian Federation); Geant4 Associates International Ltd (United Kingdom); Ivantchenko, A. [Geant4 Associates International Ltd (United Kingdom); Mantero, A. [SwHaRD s.r.l., via Buccari 9, 16153 Genova (Italy); Nieminem, P.; Santin, G. [ESA-ESTEC, 2200 AG Noordwijk (Netherlands); Tran, H.N. [Division of Nuclear Physics and Faculty of Applied Sciences, Ton Duc Thang University, Tan Phong Ward, District 7, Ho Chi Minh City (Viet Nam); Stepan, V. [CNRS, IN2P3, CENBG, UMR 5797, F-33170 Gradignan (France); Nuclear Physics Institute of the ASCR, Prague (Czech Republic); Incerti, S., E-mail: incerti@cenbg.in2p3.fr [CNRS, IN2P3, CENBG, UMR 5797, F-33170 Gradignan (France)

2014-10-01

Context Under irradiation, a biological system undergoes a cascade of chemical reactions that can lead to an alteration of its normal operation. There are different types of radiation and many competing reactions. As a result the kinetics of chemical species is extremely complex. The simulation becomes then a powerful tool which, by describing the basic principles of chemical reactions, can reveal the dynamics of the macroscopic system. To understand the dynamics of biological systems under radiation, since the 80s there have been on-going efforts carried out by several research groups to establish a mechanistic model that consists in describing all the physical, chemical and biological phenomena following the irradiation of single cells. This approach is generally divided into a succession of stages that follow each other in time: (1) the physical stage, where the ionizing particles interact directly with the biological material; (2) the physico-chemical stage, where the targeted molecules release their energy by dissociating, creating new chemical species; (3) the chemical stage, where the new chemical species interact with each other or with the biomolecules; (4) the biological stage, where the repairing mechanisms of the cell come into play. This article focuses on the modeling of the chemical stage. Method This article presents a general method of speeding-up chemical reaction simulations in fluids based on the Smoluchowski equation and Monte-Carlo methods, where all molecules are explicitly simulated and the solvent is treated as a continuum. The model describes diffusion-controlled reactions. This method has been implemented in Geant4-DNA. The keys to the new algorithm include: (1) the combination of a method to compute time steps dynamically with a Brownian bridge process to account for chemical reactions, which avoids costly fixed time step simulations; (2) a k–d tree data structure for quickly locating, for a given molecule, its closest reactants. The

Design of an embedded inverse-feedforward biomolecular tracking controller for enzymatic reaction processes

OpenAIRE

Foo, Mathias; Kim, Jongrae; Sawlekar, Rucha; Bates, Declan G.

2017-01-01

Feedback control is widely used in chemical engineering to improve the performance and robustness of chemical processes. Feedback controllers require a ‘subtractor’ that is able to compute the error between the process output and the reference signal. In the case of embedded biomolecular control circuits, subtractors designed using standard chemical reaction network theory can only realise one-sided subtraction, rendering standard controller design approaches inadequate. Here, we show how a b...

The MOA thruster. A high performance plasma accelerator for nuclear power and propulsion applications

International Nuclear Information System (INIS)

Frischauf, Norbert; Hettmer, Manfred; Grassauer, Andreas; Bartusch, Tobias; Koudelka, Otto

2009-01-01

More than 60 years after the late Nobel laureate Hannes Alfven had published a letter stating that oscillating magnetic fields can accelerate ionised matter via magneto-hydrodynamic interactions in a wave like fashion, the technical implementation of Alfven waves for propulsive purposes has been proposed, patented and examined for the first time by a group of inventors. The name of the concept, utilising Alfven waves to accelerate ionised matter for propulsive purposes, is MOA - Magnetic field Oscillating Amplified thruster. Alfven waves are generated by making use of two coils, one being permanently powered and serving also as magnetic nozzle, the other one being switched on and off in a cyclic way, deforming the field lines of the overall system. It is this deformation that generates Alfven waves, which are in the next step used to transport and compress the propulsive medium, in theory leading to a propulsion system with a much higher performance than any other electric propulsion system. While space propulsion is expected to be the prime application for MOA and is supported by numerous applications such as Solar and/or Nuclear Electric Propulsion or even as an 'afterburner system' for Nuclear Thermal Propulsion, other, terrestrial applications, like coating, semiconductor implantation and manufacturing as well as steel cutting can be thought of as well, making the system highly suited for a common space-terrestrial application research and utilisation strategy. This paper presents the recent developments of the MOA Thruster R and D activities at QASAR, the company in Vienna, Austria, which has been set up to further develop and test the Alfven wave technology and its applications. (author)

Resonant and Ground Experimental Study on the Microwave Plasma Thruster

Science.gov (United States)

Yang, Juan; He, Hongqing; Mao, Genwang; Qu, Kun; Tang, Jinlan; Han, Xianwei

2002-01-01

chemistry. Therefore, the application of EP for the attitude control and station keeping of satellite, the propulsion of deep space exploration craft allows to reduce substantially the mass of on-board propellant and the launching cost. The EP research is now receiving high interest everywhere. microwave generating subsystem, the propellant supplying subsystem and the resonator (the thruster). Its principle is that the magnetron of the microwave generating subsystem transfers electric energy into microwave energy at given frequency which is introduced into a resonant cavity. Microwave will resonate within the cavity when it is adjusted. When the propellant gas (N2, Ar, He, NH3 or H2) is put into the cavity and coupled with microwave energy at the maximal electric intensity place, it will be broken down to form free-floating plasma, which flows from nozzle with high speed to produce thrust. Its characteristic is high efficiency, simple power supply and without electrode ablation, its specific impulse is greater than arcjet. 2450MHz, have been developed. The microwave generating subsystem and resonator of lower power MPT, 70-200W, are coaxial. The resonator with TEM resonating mode is section of coaxial wave-guide, of which one end is shorted, another is semi-opened. The maximal electric intensity field is in the lumped capacity formed between the end surface of inner conductor, retracting in the cavity, and the semi-opened surface of outer conductor. It provides favorable condition for gas breakdown. The microwave generating system and resonator of middle power MPT, 500-1,000W, are wave-guide cavity. The resonator with TM011 resonating mode is cylinder wave-guide cavity, of which two end surface are shorted. The distribution of electromagnetic field is axial symmetry, its maximal electric intensity field locates on the axis and closes to the exit of nozzle, where the propellant gas is breakdown to form free floating plasma. The plasma is free from the wall of

Two-pulse atomic coherent control spectroscopy of Eley-Rideal reactions: An application of an atom laser

International Nuclear Information System (INIS)

Joergensen, Solvejg; Kosloff, Ronnie

2003-01-01

A spectroscopic application of the atom laser is suggested. The spectroscopy termed 2PACC (two-pulse atomic coherent control) employs the coherent properties of matter waves from a two-pulse atom laser. These waves are employed to control a gas-surface chemical recombination reaction. The method is demonstrated for an Eley-Rideal reaction of a hydrogen or alkali atom-laser pulse where the surface target is an adsorbed hydrogen atom. The reaction yields either a hydrogen or alkali hydride molecule. The desorbed gas-phase molecular yield and its internal state is shown to be controlled by the time and phase delay between two atom-laser pulses. The calculation is based on solving the time-dependent Schroedinger equation in a diabatic framework. The probability of desorption which is the predicted 2PACC signal has been calculated as a function of the pulse parameters

ATS-6 engineering performance report. Volume 2: Orbit and attitude controls

Science.gov (United States)

Wales, R. O. (Editor)

1981-01-01

Attitude control is reviewed, encompassing the attitude control subsystem, spacecraft attitude precision pointing and slewing adaptive control experiment, and RF interferometer experiment. The spacecraft propulsion system (SPS) is discussed, including subsystem, SPS design description and validation, orbital operations and performance, in-orbit anomalies and contingency operations, and the cesium bombardment ion engine experiment. Thruster failure due to plugging of the propellant feed passages, a major cause for mission termination, are considered among the critical generic failures on the satellite.

Chirality-controlled spontaneous twisting of crystals due to thermal topochemical reaction.

Science.gov (United States)

Rai, Rishika; Krishnan, Baiju P; Sureshan, Kana M

2018-03-20

Crystals that show mechanical response against various stimuli are of great interest. These stimuli induce polymorphic transitions, isomerizations, or chemical reactions in the crystal and the strain generated between the daughter and parent domains is transcribed into mechanical response. We observed that the crystals of modified dipeptide LL (N 3 -l-Ala-l-Val-NHCH 2 C≡CH) undergo spontaneous twisting to form right-handed twisted crystals not only at room temperature but also at 0 °C over time. Using various spectroscopic techniques, we have established that the twisting is due to the spontaneous topochemical azide-alkyne cycloaddition (TAAC) reaction at room temperature or lower temperatures. The rate of twisting can be increased by heating, exploiting the faster kinetics of the TAAC reaction at higher temperatures. To address the role of molecular chirality in the direction of twisting the enantiomer of dipeptide LL, N 3 -d-Ala-d-Val-NHCH 2 C≡CH (DD), was synthesized and topochemical reactivity and mechanoresponse of its crystals were studied. We have found that dipeptide DD not only underwent TAAC reaction, giving 1,4-triazole-linked pseudopolypeptides of d-amino acids, but also underwent twisting with opposite handedness (left-handed twisting), establishing the role of molecular chirality in controlling the direction of mechanoresponse. This paper reports ( i ) a mechanical response due to a thermal reaction and ( ii ) a spontaneous mechanical response in crystals and ( iii ) explains the role of molecular chirality in the handedness of the macroscopic mechanical response.

Control rod effects on reaction rate distributions in tight pitched PuO2-UO2 fuel assembly

International Nuclear Information System (INIS)

Gil, Choong-Sup; Okumura, Keisuke; Ishiguro, Yukio

1991-11-01

Investigations were made for the heterogeneity effects caused by insertion or withdrawal of a B 4 C control rod on fine structure of reaction rates distributions in a tight pitched PuO 2 -UO 2 fuel assembly. Analysis was carried out by using the VIM and SRAC codes with the libraries based on JENDL-2 for the hexagonal fuel assembly basically corresponding to the PROTEUS-LWHCR experimental core. The reaction rates are affected more remarkably by the withdrawal of the control rod rather than its insertion. The changes of the reaction rates were decomposed into three terms of spectrum shifts, the changes of effective cross sections with fine groups, and their higher order components. From the analysis, it is concluded that most changes of reaction rates are caused by spectral shifts. The SRAC code with fine group constants can predict the distribution of reaction rates and their ratios with the accuracy of about 5 % except for the values related to Pu-242 capture rate, as compared with the VIM results. To increase the accuracy, it is necessary to generate the effective cross sections of the fuel near control rods with consideration of the heterogeneities in the fuel assembly. (author)

Classical Example of Total Kinetic and Thermodynamic Control: The Diels-Alder Reaction between DMAD and Bis-furyl Dienes.

Science.gov (United States)

Borisova, Kseniya K; Kvyatkovskaya, Elizaveta A; Nikitina, Eugeniya V; Aysin, Rinat R; Novikov, Roman A; Zubkov, Fedor I

2018-04-20

A rare example of chemospecificity in the tandem Diels-Alder reaction of activated alkynes and bis-dienes has been revealed. The reaction between bis-furyl dienes and DMAD occurs at 25-80 °C and leads to kinetically controlled "pincer" adducts, 4a,8a-disubstituted 1,4:5,8-diepoxynaphthalenes. On the contrary, only thermodynamically controlled "domino" adducts (2,3-disubstituted 1,4:5,8-diepoxynaphthalenes) are formed in the same reaction at 140 °C. The "pincer" adducts can be transformed to the "domino" adducts at heating. The rate constants for reactions of both types were calculated using dynamic 1 H NMR spectroscopy.

A novel approach to sports concussion assessment: Computerized multilimb reaction times and balance control testing.

Science.gov (United States)

Vartiainen, Matti V; Holm, Anu; Lukander, Jani; Lukander, Kristian; Koskinen, Sanna; Bornstein, Robert; Hokkanen, Laura

2016-01-01

Mild traumatic brain injuries (MTBI) or concussions often result in problems with attention, executive functions, and motor control. For better identification of these diverse problems, novel approaches integrating tests of cognitive and motor functioning are needed. The aim was to characterize minor changes in motor and cognitive performance after sports-related concussions with a novel test battery, including balance tests and a computerized multilimb reaction time test. The cognitive demands of the battery gradually increase from a simple stimulus response to a complex task requiring executive attention. A total of 113 male ice hockey players (mean age = 24.6 years, SD = 5.7) were assessed before a season. During the season, nine concussed players were retested within 36 hours, four to six days after the concussion, and after the season. A control group of seven nonconcussed players from the same pool of players with comparable demographics were retested after the season. Performance was measured using a balance test and the Motor Cognitive Test battery (MotCoTe) with multilimb responses in simple reaction, choice reaction, inhibition, and conflict resolution conditions. The performance of the concussed group declined at the postconcussion assessment compared to both the baseline measurement and the nonconcussed controls. Significant changes were observed in the concussed group for the multilimb choice reaction and inhibition tests. Tapping and balance showed a similar trend, but no statistically significant difference in performance. In sports-related concussions, complex motor tests can be valuable additions in assessing the outcome and recovery. In the current study, using subtasks with varying cognitive demands, it was shown that while simple motor performance was largely unaffected, the more complex tasks induced impaired reaction times for the concussed subjects. The increased reaction times may reflect the disruption of complex and integrative cognitive

New technologies for ammonium dinitramide based monopropellant thrusters - The project RHEFORM

Science.gov (United States)

Negri, Michele; Wilhelm, Marius; Hendrich, Christian; Wingborg, Niklas; Gediminas, Linus; Adelöw, Leif; Maleix, Corentin; Chabernaud, Pierre; Brahmi, Rachid; Beauchet, Romain; Batonneau, Yann; Kappenstein, Charles; Koopmans, Robert-Jan; Schuh, Sebastian; Bartok, Tobias; Scharlemann, Carsten; Gotzig, Ulrich; Schwentenwein, Martin

2018-02-01

New technologies are developed in the project RHEFORM to enable the replacement of hydrazine with liquid propellants based on ammonium dinitramide (ADN). The replacement of hydrazine with green propellants will make space propulsion more sustainable and better suitable for the requirements of future missions. In the RHEFORM project investigation on the composition of the propellants are conducted to enable the use of materials for catalysts and combustion chambers which are not subject to the International Traffic in Arms Regulations (ITAR). New igniters are under development aiming at a reduction of required energy and a more prompt ignition. Two different types of igniters are considered: improved catalytic igniters and thermal igniters. The technologies developed in RHEFORM will be implemented in two thruster demonstrators, aiming at a technology readiness level (TRL) of 5. In the present work the results obtained in the first half of the project are presented.

Modeling an Iodine Hall Thruster Plume in the Iodine Satellite (ISAT)

Science.gov (United States)

Choi, Maria

2016-01-01

An iodine-operated 200-W Hall thruster plume has been simulated using a hybrid-PIC model to predict the spacecraft surface-plume interaction for spacecraft integration purposes. For validation of the model, the plasma potential, electron temperature, ion current flux, and ion number density of xenon propellant were compared with available measurement data at the nominal operating condition. To simulate iodine plasma, various collision cross sections were found and used in the model. While time-varying atomic iodine species (i.e., I, I+, I2+) information is provided by HPHall simulation at the discharge channel exit, the molecular iodine species (i.e., I2, I2+) are introduced as Maxwellian particles at the channel exit. Simulation results show that xenon and iodine plasma plumes appear to be very similar under the assumptions of the model. Assuming a sticking coefficient of unity, iodine deposition rate is estimated.

Colloid Microthruster Feed System Development for Fine Pointing and Drag-Free Control of Multi-Year Astronomical Observatories

Science.gov (United States)

Ziemer, John; Mueller, J.; Spence, D.; Hruby, V.

2014-01-01

A new Colloid Microthruster feed system, including a propellant tank and redundant Microvalves, is being developed for fine pointing and drag-free operations of multi-year astronomical observatories under the PCOS SAT program. Almost all Gravitational Wave Observatory (GWO) concepts require microthrusters to maintain a drag-free environment for the inertial sensor instrument to meet the mission science objectives. The current state-of-the-art microthruster in the US is the Busek Colloid Micro-Newton Thruster (CMNT) originally developed under the New Millennium Program for the Space Technology 7 (ST7) and ESA's LISA Pathfinder (LPF) technology demonstration mission. The ST7 CMNT design includes a bellows propellant storage tank that is sized to provide up to 90 days of maximum thrust (30 µN). The new propellant tank is based on a blow-down, metal-diaphragm spherical tank design with enough capacity for a 5-year GWO mission. The new feed system will also include the third generation of Busek’s Microvalve, currently being developed under a NASA Phase II SBIR. The Microvalve is responsible for the picoliter per second control of the propellant from the tank to the thruster head, demanding parts with micron-level tolerances, critical alignments, and challenging acceptance test protocols. This microthruster system could also be considered for replacement of reaction wheels for slewing and fine pointing of other astronomical observatories, including Exo-Planet Observatory concepts. The goal of the PCOS SAT effort is to raise the new system to TRL 5 with performance and environmental testing within the next two years.

Numerical study on the electron—wall interaction in a Hall thruster with segmented electrodes placed at the channel exit

International Nuclear Information System (INIS)

Qing Shao-Wei; E Peng; Xu Dian-Guo; Duan Ping

2013-01-01

Electron—wall interaction is always recognized as an important physical problem because of its remarkable influences on thruster discharge and performance. Based on existing theories, an electrode is predicted to weaken electron—wall interaction due to its low secondary electron emission characteristic. In this paper, the electron—wall interaction in an Aton-type Hall thruster with low-emissive electrodes placed near the exit of discharge channel is studied by a fully kinetic particle-in-cell method. The results show that the electron—wall interaction in the region of segmented electrode is indeed weakened, but it is significantly enhanced in the remaining region of discharge channel. It is mainly caused by electrode conductive property which makes equipotential lines convex toward channel exit and even parallel to wall surface in near-wall region; this convex equipotential configuration results in significant physical effects such as repelling electrons, which causes the electrons to move toward the channel center, and the electrons emitted from electrodes to be remarkably accelerated, thereby increasing electron temperature in the discharge channel, etc. Furthermore, the results also indicate that the discharge current in the segmented electrode case is larger than in the non-segmented electrode case, which is qualitatively in accordance with previous experimental results. (physics of gases, plasmas, and electric discharges)

The Reaction Wheel Pendulum

CERN Document Server

Block, Daniel J; Spong, Mark W

2007-01-01

This monograph describes the Reaction Wheel Pendulum, the newest inverted-pendulum-like device for control education and research. We discuss the history and background of the reaction wheel pendulum and other similar experimental devices. We develop mathematical models of the reaction wheel pendulum in depth, including linear and nonlinear models, and models of the sensors and actuators that are used for feedback control. We treat various aspects of the control problem, from linear control of themotor, to stabilization of the pendulum about an equilibrium configuration using linear control, t

Effect of the Gouy phase on the coherent phase control of chemical reactions.

Science.gov (United States)

Gordon, Robert J; Barge, Vishal J

2007-11-28

We show how the spatial phase of a focused laser beam may be used as a tool for controlling the branching ratio of a chemical reaction. Guoy discovered [Acad. Sci., Paris, C. R. 110, 1250 (1890)] that when an electromagnetic wave passes through a focus its phase increases by pi. In a coherent control scheme involving the absorption of n photons of frequency omega(m) and m photons of frequency omega(n), the overall phase shift produced by the Gouy phase is (n-m)pi. At any given point in space, this phase shift is identical for all reaction products. Nevertheless, if the yields for different reaction channels have different intensity dependencies, the Gouy phase produces a net phase lag between the products that varies with the axial coordinate of the laser focus. We obtain here analytical and numerical values of this phase as the laser focus is scanned across the diameter of the molecular beam, taking into account the Rayleigh range and astigmatism of the laser beam and saturation of the transition. We also show that the modulation depth of the interference pattern may be increased by optimizing the relative intensities of the two fields.

Neural Network Control of CSTR for Reversible Reaction Using Reverence Model Approach

Directory of Open Access Journals (Sweden)

Duncan ALOKO

2007-01-01

Full Text Available In this work, non-linear control of CSTR for reversible reaction is carried out using Neural Network as design tool. The Model Reverence approach in used to design ANN controller. The idea is to have a control system that will be able to achieve improvement in the level of conversion and to be able to track set point change and reject load disturbance. We use PID control scheme as benchmark to study the performance of the controller. The comparison shows that ANN controller out perform PID in the extreme range of non-linearity.This paper represents a preliminary effort to design a simplified neutral network control scheme for a class of non-linear process. Future works will involve further investigation of the effectiveness of thin approach for the real industrial chemical process

Assessment of High-Voltage Photovoltaic Technologies for the Design of a Direct Drive Hall Effect Thruster Solar Array

Science.gov (United States)

Mikellides, I. G.; Jongeward, G. A.; Schneider, T.; Carruth, M. R.; Peterson, T.; Kerslake, T. W.; Snyder, D.; Ferguson, D.; Hoskins, A.

2004-01-01

A three-year program to develop a Direct Drive Hall-Effect Thruster system (D2HET) begun in 2001 as part of the NASA Advanced Cross-Enterprise Technology Development initiative. The system, which is expected to reduce significantly the power processing, complexity, weight, and cost over conventional low-voltage systems, will employ solar arrays that operate at voltages higher than (or equal to) 300 V. The lessons learned from the development of the technology also promise to become a stepping-stone for the production of the next generation of power systems employing high voltage solar arrays. This paper summarizes the results from experiments conducted mainly at the NASA Marshal Space Flight Center with two main solar array technologies. The experiments focused on electron collection and arcing studies, when the solar cells operated at high voltages. The tests utilized small coupons representative of each solar array technology. A hollow cathode was used to emulate parts of the induced environment on the solar arrays, mostly the low-energy charge-exchange plasma (1012-1013 m-3 and 0.5-1 eV). Results and conclusions from modeling of electron collection are also summarized. The observations from the total effort are used to propose a preliminary, new solar array design for 2 kW and 30-40 kW class, deep space missions that may employ a single or a cluster of Hall- Effect thrusters.

Real-time nonlinear feedback control of pattern formation in (bio)chemical reaction-diffusion processes: a model study.

Science.gov (United States)

Brandt-Pollmann, U; Lebiedz, D; Diehl, M; Sager, S; Schlöder, J

2005-09-01

Theoretical and experimental studies related to manipulation of pattern formation in self-organizing reaction-diffusion processes by appropriate control stimuli become increasingly important both in chemical engineering and cellular biochemistry. In a model study, we demonstrate here exemplarily the application of an efficient nonlinear model predictive control (NMPC) algorithm to real-time optimal feedback control of pattern formation in a bacterial chemotaxis system modeled by nonlinear partial differential equations. The corresponding drift-diffusion model type is representative for many (bio)chemical systems involving nonlinear reaction dynamics and nonlinear diffusion. We show how the computed optimal feedback control strategy exploits the system inherent physical property of wave propagation to achieve desired control aims. We discuss various applications of our approach to optimal control of spatiotemporal dynamics.

Space robot simulator vehicle

Science.gov (United States)

Cannon, R. H., Jr.; Alexander, H.