Type Of Thruster Research Articles

Measurement of ion acceleration characteristics of a laser-electrostatic hybrid microthruster for space propulsion applications

For space propulsion applications, a fundamental study of a laser-electrostatic hybrid thruster was conducted. A new type of thruster, in which a laser-produced plasma was accelerated by an additional electrostatic field, was tested to optimize the ion acceleration process. A time-of-flight measurement by a Faraday cup showed that the average speed of ions was about 15 km/s when only 0.04 mJ of laser impulse was introduced to a copper target. When an accelerator electrode with a 6-mm-diameter hole was placed in front of the laser target, it was observed that the average speed of ions increased. The maximum velocity was 23 km/s, which corresponded to the case where the accelerator grid was biased to +100 V for the target-to-electrode gap of 2 mm. It was found that the positively biased electrode was more effective than the negatively biased electrode for ion acceleration in the thruster.

Vortex Ring Thrusters for Maneuvering and Control of Underwater Robots

Inspired by the natural locomotion of squid and other cephalopod, a new type of thruster was designed which utilized pulsatile jet propulsion to generate controlling forces. The thrust production from this device was characterized in a static environment and seen to be well approximated by a simple fluid slug model. A virtual vehicle simulation was developed to test the thruster with unsteady driving signals. Within this environment using a simple proportional derivative feedback control algorithm, the thruster/controller was observed to provide sufficient control over the vehicle performing large scale maneuvers. However, compensation is required for accurate small scale maneuvers due to a thruster time delay which is unmodeled in the simple slug model. +

Recent advances in nuclear powered electric propulsion for space exploration

Nuclear and radioisotope powered electric thrusters are being developed as primary in space propulsion systems for potential future robotic and piloted space missions. Possible applications for high-power nuclear electric propulsion include orbit raising and maneuvering of large space platforms, lunar and Mars cargo transport, asteroid rendezvous and sample return, and robotic and piloted planetary missions, while lower power radioisotope electric propulsion could significantly enhance or enable some future robotic deep space science missions. This paper provides an overview of recent US high-power electric thruster research programs, describing the operating principles, challenges, and status of each technology. Mission analysis is presented that compares the benefits and performance of each thruster type for high priority NASA missions. The status of space nuclear power systems for high-power electric propulsion is presented. The paper concludes with a discussion of power and thruster development strategies for future radioisotope electric propulsion systems.

Examination of plasma-wall interactions in Hall effect thrusters by means of calibrated thermal imaging

A semiempirical time-dependent thermal model of Hall effect thrusters was developed to compute evolution of the thruster channel wall temperature during a transient regime. The model accounts for heat conduction through dielectric walls and radiative transfer as the channel is considered as a thermal enclosure. The graybody view factors are experimentally determined during the cooling period that follows a power shutdown. The model permits the assessment of the amount of power delivered to walls by fitting calculated temperature profiles to experimental profiles obtained by means of calibrated infrared imaging. Plasma-surface interactions are examined in this contribution for three types of thrusters, namely, SPT100-ML, PPS®1350-G, and the 5-kW-class PPSX000-ML. The electrical input power range that is probed stretches out from 500Wto5kW. The energy flux passed to the discharge chamber dielectric walls varies between 0.1 and 2W∕cm2. The overall amount of power delivered to the external and internal walls is found to vary linearly with the applied power. It is in favor of an energy loss mechanism driven by ion bombardment. A departure from the linear trend is observed at high discharge voltage. Production of multiply charged ion species could be an explanation. The impact of thruster design and wall material on power losses inside the thruster channel is also investigated from both a technological and a physics standpoint.

Sub-Milli-Newton Class Miniature Microwave Ion Thruster

A miniaturized microwave ion source with a 1.6-cm beam diameter grid system was designed and then evaluated experimentally. Based on the HAYABUSA μ10 neutralizer, we fabricated a small 18-nun-diam discharge chamber, into which 4.2 GHz microwaves were launched through an L-shaped antenna that was located in a magnetic field created by permanent magnets and iron yokes. Ion beams were emitted from the small discharge chamber when operated with a grid system whose respective hole diameters of the screen grid and acceleration grid were 0.72 and 0.43 mm, and the total number of grid holes was 211. For a beam voltage of 1500 V and a microwave input power of 10 W, the typical thruster performance was thrust of 0.34 mN, a thrust/power ratio of 16 mN/kW, propellant utilization efficiency of 68%, and a specific impulse of 3200 s. If we were able to further reduce the ion production cost (circa 3000 W/A in the current experiment), this thruster would be a candidate for main propulsion of a small satellite or precise attitude control of various sizes of satellites.

Comparison of the rocket engines efficiency in the case of low thrust orbit-to-orbit transfers

The main task of this paper is to compare two types of low thrust rocket engines: constant thrust vs. variable-thrust engines. We will be concerned with efficiency, where efficiency is evaluated in the case of the orbit-to-orbit transfer with maximum payload mass in the central Newtonian gravity field. The launch mass of the space vehicle is supposed to be fixed. The traditional solution is the decomposition of the problem into parametric and dynamical parts. The corresponding variational problems differ for two rocket thruster types under consideration. We propose change of variables, which makes it possible to reduce averaged equations of optimal motion of a spacecraft with the mentioned engines to the unified form. Using this unified form comparison of the performance of constant- and variable-thrust engines is conducted.

A calibrated infrared imaging study on the steady state thermal behaviour of Hall effect thrusters

The thermal behaviour of Hall effect thrusters was investigated by means of calibrated infrared thermal imaging performed in the 8–9 µm spectral domain. Study on the variation of the steady state temperature of Hall thruster elements like discharge chamber (channel) walls and anodes along with discharge voltage and propellant (xenon) mass flow rate confirms that energy loss mechanisms, which are responsible for the heating of the thrusters, are a direct consequence of interactions between charged particles and surfaces. In order to obtain new insights into plasma surface interactions inside a thruster, the channel wall temperature was monitored over a broad range of electrical power stretching from 400 W to 5.5 kW for three types of thrusters with different designs, dimensions and operation domains, namely SPT100-ML, PPS ® 1350-G and PPSX000-ML. Note that over the range of thruster operating conditions the facility backpressure varies from 10−5 to 6 × 10−5 mbar. In addition, the effect of discharge chamber wall material on temperature field was also investigated using dielectric BN–SiO2 and AlN walls as well as conducting graphite walls. For a given thruster geometry and material, a simple relationship between the mean wall temperature and the input electrical power can be established, in contradiction to the complex dynamics of such a magnetized plasma medium. Besides, thruster thermal history and degree of wear do not have a strong impact on power losses inside the channel.

Thermal Analysis for Propellant Stream in Thruster'sInjection Tube During Start Process

According to the leading design features of the monopropellant attitude-control thruster and the cryogenic condition in space, the coupling heat-transfer physical model and a set of mathematical equations are put forward to calculate the temperature variation of the injection tube and three other components of the thruster, which are the aggregate organ, the bracket, and the injection plate. The wall temperature of the injection tube is taken as the thermal boundary conditions for the propellant stream temperature drop calculation. The temperature calculation of the propellant stream flowing inside the injection tube during the start process in a typical 10 N thruster has been performed, and two simulation experiments are carried out on the same type thruster in the vacuum cryogenic chamber. The results of the calculation agree well with the experiments. The validated calculation model can be used to judge whether the propellant stream freezes or not in the injection tube during the start process after the thruster has been exposed to the cryogenic space for a long time. In addition, four factors that affect the front temperature of the propellant stream at the injection tube exit during the propellant filled process are studied, and according to the calculation results, three correlation relationships are deduced, respectively.

小型マイクロ波イオン推進機の実験的評価

For a main propulsion system of small satellite or precise attitude control of various size satellites, a miniaturized microwave ion thruster based on HAYABUSA μ10 neutralizer was fabricated. This thruster has a 1.6cm beam diameter grid system with high hole number density for the improvement of beam extraction. Through the experiments, it was evaluated that the typical thruster performance was the thrust of 0.34mN, the thrust/power ratio of 16mN/kW, the propellant utilization efficiency of 68% and the specific impulse of 3,200s. Judging from the comparison with the other small thruster performance, it was estimated that this thruster can be a candidate for the small thruster.

Mitigation of Current-Driven, Vortex-Induced Vibrations of a Spar Platform via “SMART” Thrusters

The possibility of using active-control technology to mitigate long-period (100–300 s) motions of spar platforms was investigated. In particular, the technical feasibility and cost-effectiveness of using a thruster-based active-control system was examined. Only vortex-induced vibrations (VIV) of a moored spar were considered for this study, although it is believed that similar active-control systems can contend equally well with low-frequency wave-drift and wind-induced motions. VIV has been selected for this demonstration due to the considerable attention it has received for spars, because of the high cost for passive VIV-mitigation means. The investigation employed numerical simulations of the hydrodynamics around the spar in two and three dimensions. In particular, for this study, Navier-Stokes methods were used to compute the strongly nonlinear interactions between the current-flow field and the spar hull. The fluid forces (hull-integrated pressures) were then used to compute the unsteady motions of the moored spar. The effect of the motion-controlling thrusters was included as an additional external force—that also varies with time. Of the several different control strategies that were examined, it was found that the most effective one is “D-Control” wherein thruster forces are proportional to spar velocities. Using such control, spar motions can be kept below 35% of the spar diameter, whereas uncontrolled responses were computed to be as high as 80% of the spar diameter. Similar uncontrolled motions have actually been observed in current-tank model tests of spars. As part of this study, discussions were held with engineers from three different thruster companies to ascertain the number and type of thrusters that would be required to achieve desirable VIV mitigation—given the computed thruster-forces, and to provide prudent reliability. Associated costs were also very much of interest. For the spar system examined here, which is representative of those introduced into or being considered for the Gulf of Mexico, it was found that the cost of the thrusters and their electric motors is around $8 million—a cost that is significantly less than that estimated for passive VIV-control devices such as helical strakes, or discontinuous hull geometries. Furthermore, the proposed thruster system could also be employed to contend with (to mitigate) other undesirable, large-amplitude, near-resonant spar motions, e.g., low-frequency drift, for which there are no alternative remedies at present.

On a similarity criterion for plasma accelerators of the stationary plasma thruster type

Operation of a stationary plasma accelerator of the ATON type employing various working gases is considered. Using a proposed similarity criterion, an optimum operation regime of the accelerator for krypton is selected. Comparative integral parameters of an accelerator operating on krypton and xenon are presented.

Diode Laser-Driven Microthrusters: A New Departure for Micropropulsion

We developed an entirely new type of orientation thruster for micro- and nanosatellites. The laser plasma thruster is based on the recent commercial availability of diode lasers with sufficient brightness and 100% duty cycle to produce a repetitively pulsed or continuous vapor or plasma jet on a surface in vacuum. A low-voltage semiconductor switch can drive the laser. A lens focuses the laser diode output on the ablation target, producing a miniature jet that provides the thrust. Single-impulse dynamic range is nearly five orders of magnitude, and the minimum impulse bit is 1 nN/s in a 100-„s pulse. Even with diffraction-limited focusing optics, at least 0.5-W optical power is needed to produce thrust from selected ablator materials. Thrust-to-power ratio Cm is 50 to 100„N/W and specific impulse Isp is 200‐500 s with a 1-W laser, depending partially on the illumination mode. Transmission and reflection (R) illumination modes are discussed. R mode gives about 50% better Isp and two times better Cm. Improved results are anticipated from higher laser power in the reflection mode. The prototype engine we are developing is intended to provide lifetime on-orbit steering for a 5-kg satellite, as well as reentering it from low Earth orbit.

Fundamental study on laser plasma accelerator for propulsion applications

Recent advances of laser plasma accelerators that promise very high gradients of the accelerating fields and enable small sized accelerators were reviewed, and a feasibility study of the utilization of these accelerators for space propulsion was conducted. Forward plasma acceleration through a foil target with a low-power laser pulse was observed, and resultant impulse was measured for various targets. The significant merit of this thruster was discussed in the case of operation of this type of thruster on earth and/or solar orbit, in which solar power is available. Related issues on this point were also discussed.

Gas-Surface Interaction Model Influence on Predicted Performance of Microelectromechanical System Resistojet

The free molecule micro-resistojet was designed as a micropropulsion system capable of performing attitude controlandprimarymaneuversfornanospacecraftwithmasslessthan10kg.Thedetailsofgas ‐surfaceinteractions between propellant molecules and surfaces held at elevated temperature are critical in predicting the propulsion system’ s performance and efe ciency. The aim is to assess parametrically the performance of a typical thruster geometry using a general Maxwell scattering model and two versions of the Cercignani ‐Lampis‐Lord model (Lord, R. G., “ Some Extensions of the Cercignani ‐Lampis Gas‐Surface Scattering Kernal,” Physics of Fluids, A , Vol.3,No.4,1991,pp.706 ‐710 andLord, R.G., “ SomeFurtherExtensionsoftheCercignani ‐LampisGas‐Surface Interaction Model,” Physics of Fluids, A , Vol. 7, No. 5, 1995, pp. 1159 ‐1161). The models are incorporated into a direct simulation Monte Carlo numerical code and are used to bound the predicted performance characteristics of the thruster. The total specie c impulse varies by approximately 20% over range of accommodation coefe cients from specular to diffuse surface scattering. However, there was only a maximum difference of about 5% between the models for a given accommodation coefe cient. Other more microscopic parameters, such as axial velocity distribution functions, appear to depend more on the scattering model assumed.

Three-Dimensional Numerical Simulation of Field-Emission-Electric-Propulsion Backflow Contamination

A three-dimensional full particle particle-in-cell with Monte Carlo collision simulation model is developed to study charge-exchange ion backe ow from a e eld-emission-electric-propulsion thruster using liquid cesium as a propellant. Contamination from backe ow of low-energy charge-exchange ions generated in the plume is an important spacecraft integration issue. Simulations show that the backe ow ion current is on the order of 0.01% of the total emitting current under typical thruster operating conditions. Nearly all charge-exchange ions are collected by the accelerator electrode close to the emitter. Operating a e eld emission thruster at lower neutral e ux, lower emitter current, or higher emitter potentials helps to reduce the backe ow current and possible spacecraft contamination. The operation of a neutralizer does not have a signie cant effect on the backe ow current under typical space plasma environments. A semi-analytical expression is also derived for quick estimations of the e eld emission thruster backe ow current. Nomenclature Aacc = accelerator area, m 2 Ae; Ai; An = exit area of the electron, ion, and neutral beam, m 2 a = emitter‐accelerator distance, m b = half of accelerator aperature, m C

Numerical Model of Cesium Ion Beams in Field Emission Electric Propulsion Thrusters

A numerical model of the ion beam in field emission electric propulsion thrusters is presented. The objective of the study is to develop a computational tool capable of predicting the effects of thruster configuration parameters (geometry, voltage distribution between the electrodes) on thruster performance (beam divergence, ion exhaust velocity). A two-dimensional, steady model has been adopted. A simple emission model has been assumed, neglecting the detailed, microscopic description of the physical emission mechanism. The electric field is evaluated by a finite element scheme; boundary conditions are provided by an auxiliary technique based on the placement of fictitious charges in proximity to the electrode surfaces. The beam is simulated through a discrete number of charged particles; the effects of the associated space charge density distribution on the electric field is considered. The results of the simulations of typical thruster configurations are presented and discussed. The need to take space charge effects into account to attain a sufficiently accurate solution is indicated: the beam charge density modifies the electric potential distribution generated by the electrodes voltage difference, causing a decrease in the local electric field at emitter tip and an increase in beam divergence. Neutralization at a finite distance must be taken into account to fully appreciate the effects of different voltage distributions between the electrodes: a higher accelerator electrode voltage determines a lower ion exhaust velocity and a larger beam divergence. The capability of an additional, decelerating neutral electrode to focus the ion beam is also investigated, showing that only little improvement in beam containment is attained.

Spacecraft Electric Propulsion-An Overview

A short review of the status of electric propulsion (EP) is presented to serve as an introduction to the more specialized technical papers also appearing in this Special Issue (Journal of Propulsion and Power, Vol. 14, No. 5, Sept. –Oct. 1998). The principles of operation and the several types of thrusters that are either operational or in advanced development are discussed Ž rst, followed by some considerations on the necessary power sources. A few prototypical missions are then described to highlight the operational peculiarities of EP, including spacecraft interactions. We conclude with a historical summary of the accumulated  ight experience using this technology.

Ion sources for space thrusters (invited)

One of the main tasks of the creation of spacecraft power plants is raising the thrust producing jet velocity. Conventional chemical engines create jet velocities in the range of 3000–4500 m/s. This situation can be drastically changed if beams of charged particles accelerated by electric and magnetic fields are used to produce thrust. In such cases practically any jet velocity might be created, which considerably enlarges the number of tasks being fulfilled by spacecraft having such types of thruster. Several types of electric propulsion thrusters exist nowadays. They differ in the principles of acceleration of charged particles, for example, arc jets, magnetic plasma dynamic thrusters, stationary plasma thrusters, pulse thrusters, and ion thrusters. Electric propulsion thrusters are practically the accelerators of charged particles which operate under rather strict requirements concerning energy consumption and lifetime. Since the mid-fifties in Russia there have been intensive studies of practically all types of electric propulsion thrusters, including their tests in space, and beginning with the mid-seventies they have been practically used aboard spacecraft with a long, active lifetime. The study of the physical process involved together with the research design allowed Russian scientists to develop electric propulsion thrusters in the power range from hundreds of watts to tens of kilowatts, with jet velocities between 20000 and 50000 m/s and lifetime more than several thousand hours.

A magnetically-nozzled, quasi-steady, multimegawatt, coaxial plasma thruster

The Los Alamos National Laboratory Coaxial Thruster Experiment (CTX) has been upgraded to enable 10 ms quasi-steady-state (QSS) operation of magnetohydrodynamic type thrusters at power levels from 1 to 40 MW. Here we report on experimental observations and associated analysis for QSS discharges using deuterium, helium, and argon propellants. Measurements of thruster current and voltage, magnetic field fluctuations, electron density and temperature, and plasma potential indicate true quasi-steady operation over 10 ms with propellant exhaust velocities of 1-2/spl times/10/sup 5/ m/s in helium and deuterium. The application of unique applied magnetic nozzle configurations result in a substantial reduction of the measured anode fall potential. Data analysis and analytical MHD modeling imply an impulsive start in the propellant flow that is limited to the Alfven critical ionization velocity and the presence of an effective constriction or nozzling in the how channel. This research indicates that efficiency optimization of MHD class thrusters will most likely occur at megawatt power levels. This implies repetitively pulsed, high-power operation for near term, low-mean-power solar electric propulsion missions. >

Cathode phenomena in plasma thrusters

For the optimization of cathodes of steady-state plasma accelerators, an extensive experimental program has been carried out at IRS to investigate the erosion mechanisms. On different thruster types and with an additional fundamental experiment, the influences of geometry, chamber pressure, amount of ThO2, cathode temperature, power, and gas species were explored. Propellants Ar, H2, N2 were used which had been purified with a special filter to remove residual water and oxygen. The erosion rates during the ignition phase as well as during continuous operation were measured. The temporal evolution of the cathode erosion rate was measured by surface layer activation. At high current densities within the cathode, severe problems were encountered with all thruster geometries. The cathodes developed cracks and started to melt. To clarify the reason for this behavior, some cathodes were investigated metallurgically.