Showing papers on "Arcjet rocket published in 2001"

Diode-laser tomography for arcjet plume reconstruction

Abstract: Diode-laser absorption tomography is described with which the spatial temperature and the atomic number density distribution of a 3-kW class arcjet can be derived simultaneously by reconstruction of the absorption coefficient field of the arcjet's argon exhaust plume. One can perform various parameter measurements by changing the arcjet's mass-flow rates and discharge currents. The maximum temperature and atomic number density increase with the mass-flow rate and the discharge current. The trend for increase is not always found for a specific input power, although at a fixed mass-flow rate the power increases at that rate.

Entry Conditions in Planetary Atmospheres: Emission Spectroscopy of Molecular Plasma Arcjets

Abstract: A stationary arcjet, operating at low pressure (0.13 mbar in the vacuum chamber), is used for the simulation of the gas flow surrounding a vehicle during its entry into the atmosphere of Titan and Mars. For the Titan atmosphere the gas mixture 99% N 2 - 1% CH 4 is used and 97% CO 2 - 3% N 2 for the Martian atmosphere. The respective plasma arcjets are analyzed by optical emission spectroscopy for identifying the emitting molecules and atoms and also deducing the temperatures associated with their different internal modes. For the N 2 -CH 4 plasma the vibrational temperatures T v deduced from CN and NH spectra are found in concordance with the measured electron temperature, that is, 8000 K, whereas, from CH spectra, T v is obtained close to 3700 K. The rotational temperatures are found to be between 2500 and 2800 K for CH and NH and nearly 5000 K for both CN and N 2 + spectra. For the CO 2 -N 2 plasma no emission from the arcjet is detected in the UV-visible range; the feasibility of the infrared analysis of the Δν = 2 band of CO is demonstrated in a stationary plasma discharge experiment

Multi-mode Hall Thruster Development

Abstract: The BPT-4500 and method of tuning the acceleration by focusing a plasma lens is described. Test results show that overall thruster efficiency and plume shape is improved by adjusting the curvature of magnetic field lines in the acceleration zone. Thrust efficiency is improved to 50% from 21% at 150 V and from 58% to 62.5% at 350 V using this technique. At 125 V, the BPT-4500 operates at 45% thrust efficiency and 1030 s specific impulse yielding a thrust/power ratio only 10% less than a 600 s hydrazine arcjet. Introduction Multi-mode Hall thruster capability enables one thruster to serve constant power applications needing the highest possible thrust for orbit transfer and high specific impulse for station keeping. Towards this objective, we have evaluated a GENERAL DYNAMICS (formerly PRIMEX Aerospace Company) approach for tuning the BPT-4500 for the best possible performance over a range of accelerating potentials. This paper describes our initial exploratory efforts to improve thruster efficiency by actively changing magnetic field shape. Thruster Description The BPT-4500 is a 4.5 kW Hall field plasma accelerator based on an earlier 4000 watt device''. These thrusters employ a short discharge to minimize contact with the insulators and pole pieces to shape the curvature of the magnetic field in the ion creation zone. As is well known in Hall type accelerators, the magnetic field lines become nearly equipotential lines defining the acceleration vector of the ions. Also, the technique of adjusting electric potential shape by magnetized electrons in a magnetic field is commonly called a plasma lens. By virtue of the discharge properties in the BPT and shaping these fields, we have attempted to Copyright (c) 2001 by GENERAL DYNAMICS. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. create a plasma lens with adjustable focus. Plasma Lens Focusing Morozov in the Soviet Union has shown that using two sets of electromagnets to change the gradient of the radial magnetic field can have a significant effect on the behavior of ion acceleration in Hall type discharges. Gavryushin and Kim have shown that the axial gradient of the radial magnetic field "was regulated by varying the degree of screening of the accelerator channel." These accelerators are characterized by relatively long acceleration lengths in channels that typically have a length to width of 2:1. Insofar as these accelerators had a long discharge, the curvature of the field was not emphasized in the description of this early work. Instead, the researchers discuss primarily the axial gradient of Curvature of ion path changed by Anode focal length Anode potential contour Cathode potential contour Figure 1 Hall Acceleration with Plasma Lens Concept (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. the magnetic field. The BPT discharge is configured to take place in a very short channel, typically with a length to width of 1:2. As such, physical evidence from a variety of geometry studies suggests that the BPT discharge creates ions in a thin, -1-2 mm layer, occurring approximately 3-4 mm downstream of the anode, schematically illustrated in Fig 1. As such, we hypothesize the BPT discharge embodies Hall acceleration within a plasma lens. Our innovation is to exploit the features of this short discharge by adjusting the focal length of a plasma lens using a trim or focusing coil. Variation of the curvature of the fields in this layer has the effect of changing the initial trajectory of ions. Changes in the shape downstream of the ionization zone are minimized by the fact that magnetic fields are manipulated inside the primary magnetic structure. The BPT-4500 thruster shown in Fig 2 is composed of a primary magnetic structure with four outer coils and an inner coil. The BPT-4500 has a 13 cm mid-diameter and a 2 cm annulus providing slightly more discharge area than the BPT-4000. The schematic cross section of the thruster is shown in Fig. 3. This shows the anode is surrounded by a secondary magnetic structure to add or subtract from the magnetic field radius of curvature in the annulus between insulator rings. Figure 2 BPT-4500 Laboratory Style Thruster Flight Version A flight capable version of the thruster is under development in the Lockheed Martin funded Hall Thruster Propulsion System (HTPS) program. This flight thruster is referred to as the BPT-4000. This thruster has a fixed magnetic field and plasma lens taken from the optimal configuration found here. GENERAL DYNAMICS is on track to deliver a flight qualified thruster, power processor and xenon flow control system in 2003.

Near-Field Measurement and Modeling Results for Flight-Type Arcjet: Hydrogen Atom

Abstract: Hydrogen atoms were probed in the near-e eld plume of a 1.8-kW e ight-type arcjet thruster, using two-photon laser-induced e uorescence. The arcjet was operated on an N 2/H2 gas mix simulating hydrazine decomposition products. Thevelocity distributionsof Hatoms were used to obtain theirtranslational temperatures. Comparisons of the laser-induced e uorescence signal intensity, radial linewidth, and derived density distribution with direct simulation Monte Carlo results were in good agreement except near the nozzle exit plane. A relatively weak signal was observed near the exit plane, while the maximum signal was unexpectedly obtained 0.8 cm downstream. Fluorescence quenching, linewidth variation, and self-absorption effects are proposed to account for most of the discrepancy in signal intensity. The results improve the knowledge of near-e eld e ow parameters for the arcjet and ultimately enhance the accuracy and understanding of predicted plume impingement torques and other arcjet plume phenomenology on spacecraft.

Thrust and Electrical Power from Solid Propellant Microrocka

Abstract: Progress has been made in the development of flying silicon. Fully self-contained h4EMS rockets have been designed, fabricated and tested. A peak thrust of 4" was obtained from a rocket with a mass of 0.98grams. This rocket was comprised of a 3.2" inside diameter ceramic tube bonded to a silicon chip containing the rocket nozzle, igniter and an array of twelve thermopiles. One thermopile was measured during combustion and yielded 20microWatts of electrical power during a combustion event of approximately 10 seconds. Third-generation devices have demonstrated liftoff.

An overview of a solar thermal propulsion and power system demonstration applicable to heds

Abstract: One of the key requirements for future HEDS missions will be the simple generation of reliable propulsion and power. A possible solution to this problem will be described. An overview will be given of a solar propulsion and power system experiment that could be demonstrated as an ISS payload. The experiment will use a focusing solar energy collection system to heat multiple dynamic engine and solar thermal receivers. In addition to the clean burning solar thermal engine, the dynamically generated power will in turn power a variety of electric propulsion systems being investigated by the NASA for International Space Station (ISS) reboost, potentially including Arcjet, Hall, ion, and Variable Specific Impulse Magneto Plasma (VaSIMR) engines. The dynamic power system demonstrations will validate one or more very high efficiency methods of power production that should be of interest to a variety of high power space applications envisioned for the next decade including Space Based Radar, broadband communication satellites and second generation power production on the ISS itself.

Near-Field Measurement and Modeling Results for Flight-Type Arcjet: NH Molecule

Abstract: Density, velocity, and temperature data were obtained in the near-e eld plume of a 1.8-kW hydrazine arcjet thruster, using the NH molecule as a probe. This was the e rst laser spectroscopic study on a e ight-type arcjet. The rotational temperature of NH was obtained in v = 0;1 of the electronic ground state and the vibrational population in v = 1;2. Rotational and vibrational temperatures were similar, and their variation along the thrust axis was minor. Line shapes as well as radial and axial proe les of the peak line intensity were obtained. Maximum laser-induced e uorescence signal was obtained on-axis 1.4 cm downstream from the exit plane, suggesting rapid NH production in the plume. Axial velocity components were determined at various plume locations. The power distribution of the available channels was found, on summation, to be in good agreement with arcjet input power. Direct simulation Monte Carlo predictions have been generated for comparison with much of the experimental data and to provide additional e owe eld information. Good agreement was achieved in most cases.

Plume Mass Spectrometry with a Hydrazine Arcjet Thruster

Abstract: Molecular beam mass spectrometry was used to measure the chemical composition, angular distribution, and speed distribution in the plume of an arcjet thruster at 127 mm from the nozzle The arcjet operating conditions were1660Wofinputpower,464mg /sofsimulatedhydrazinepropellant,anda background pressureof085 mtorr Time-of-e ight spectra yielded speed distributions for N 2, N, H2, and H as a function of angle µ relative to plume centerline Absolute e uxes were obtained from the experimental relative e ux distributions by normalizing to the totalmasse owrate Dissociation fractions forN 2 andH2 were 11 and 14%Molecules nearthe centerlinehad most probable speeds of 74 km /s with the high end of the distribution extending to 10 ‐12 km/s The N2 speed decreased to 27 km/s at µ=80 deg The N 2 and N accounted for 89% of the arcjet thrust, and the mass e ux for both species decreased by a factor of 10 4 between µ=0 and 90 deg The angular distributions of H 2 and H were much broader, with mass e uxes that decreased by less than a factor of 10 between µ=0 and 90 deg Other quantities derived from time-of-e ight spectra were the momentum e ux, kinetic power e ux, average speed, and translational temperature

High Density Plasma Neutralization of a Hall Thruster

Abstract: In the attempt to fill the performance niche between the arcjet and Hall thruster, we propose the use of a helium arcjet to neutralize Hall thrusters. Since the arcjet is a high plasma density device, one arcjet can potentially neutralize a cluster of Hall thrusters. In this preliminary study, we used a surrogate planar anode in the place of a Hall thruster anode to determine the effects of drawing electron current from a low power arcjet plume on the operation and performance of the arcjet. In all tests we are able to draw currents to the surrogate anode that were greater than the arc current. It is found that biasing the surrogate anode does lead to a perturbation in the arcjet discharge voltage, in some cases resulting in a voltage decrease of up to 40%. The helium arcjet exhibits arc voltage instabilities in the same spectral range of the instabilities that are intrinsic in Hall thrusters. We find however, that these arcjet instabilities decreased in strength and in bandwidth when a bias is applied to the anode. The effect of biasing on the overall arcjet performance (e.g., thrust, specific impulse) is still under investigation. However, preliminary studies made using an impact pressure probe confirms that there is little compromise in the arcjet thrust during the current draw. The use of laser-induced fluorescence to measure the velocity of the arcjet plume flowfield during an applied bias was hindered by the finding that the lower energy state of the helium transition used for the LIF measurements was effectively depopulated by the biasing a result supported by optical emission measurements of the plume.

An Overview of the On-Orbit Results from the Electric Propulsion Space Experiment (ESEX)

Abstract: : The United States Air Force (USAF) Research Laboratory's Electric Propulsion Space Experiment (ESEX) was launched and successfully operated - demonstrating the compatibility and readiness of a 26 kW ammonia arcjet subsystem for satellite applications. ESEX is one of nine experiments on the USAF's Advanced Research and Global Observation Satellite (ARGOS). Data were acquired to characterize the thruster in four different areas: electromagnetic interactions, contamination effects, optical properties of the plume, and thruster system performance. The results demonstrated that the critical system components (including the arcjet, power processor, and propellant system) operated well, and verified the interoperability of high power electric propulsion with generic satellite operations.

A Study of Low-Power MPD Arcjets for Future High-Power Evolution

Abstract: Preliminary concepts of low-power MPD (Magnetoplasmadynamic) arcjets for highpower evolution have been studied in ISAS (Institute of Space and Astronautical Science). There are two categories of these concepts, pulsed-mode and steady-state MPD arcjets. Both are the candidates, because ISAS has several flight experiences of pulsed MPD arcjet, and also R & D experiences of quasi-steady MPD arcjets and low power DC arcjets. This study shows that the cluster of single MPD arcjet will be the reasonable way to realize large-scale high-power MPD arcjet from the viewpoint of ground tests and facility availability. The focal points for such design will be the thermal analysis and electrode erosion for single MPD arcjet that is just a portion clipped out from the cluster configuration. The simple and robust MPD arcjets at high-power evolution will propose a lot of large-scale transportation in the near future.

Numerical Evaluation of Inelastic Mechanisms in a Nitrogen Arcjet

Abstract: Introduction T HE electrons in a low-power arcjet thruster are thermally excited through strong interactions with the electric Ž eld via ohmic heating andmay be out of equilibriumdue to low density and small thrusterdimensions.Numericalevaluationof suchnonequilibrium states is not only important for accurate prediction of thruster performance but also of interest from the viewpoint of  uid dynamics. For a propellant composed of molecular species, the rate of electron energy loss due to inelastic collisions with molecules can be of many orders of magnitude larger than that due to elastic collisions.One common approach taking into account these effects is to multiply the collision frequencies for electron–molecule encounters by the inelastic loss factor ±, but there exist considerable uncertaintiesin determiningthe valueof ±, whichwas assumed to be constant in Refs. 1 and 2. In fact, inelasticmechanismsare expected to be strongly nonlinearwith respect to electron energy distribution as pointed out in Ref. 3. This Note attempts to evaluate the degree of the electron– molecule inelastic effects for a nitrogen arcjet by consistently accounting for each energy transfer mechanism, not assuming a constant ±. This is essentially the same approach as the work of Megli et al. in the sense that the ± value can be evaluated from the solution obtained, but emphasis is placed on the detailed description of the different mechanisms involved and on the determination of the relative importance among them.

Development of Electromagnetic Acceleration Plasma Arcjet Generators for Titanium Nitride Reactive Spray Coatings.

Abstract: An electromagnetic acceleration plasma arcjet generator, which is called Magneto-Plasma-Dynamics (MPD) arcjet generator, has a coaxial electrode structure similar to those of conventional thermal arcjet generators. However, their acceleration mechanisms are different ; that is, in MPD arcjet generator, plasma is accelerated by the electromagnetic interaction between the discharge current and the magnetic field induced by it in MW-class input power operations during the discharge, although in thermal arcjet generators the working gas is accelerated aerodynamically through a straight or convergent-divergent nozzle. As a result, the MPD arcjet generator can produce higher-velocity, higher-temperature, higher-energy-density, and larger-area plasma than those of other conventional plasma torches can. These MPD plasma properties are effective for various material processes.This paper deals with a development of MPD arcjet generator for titanium nitride (TiN) reactive coatings. The MPD arcjet generator is equipped with a titanium cathode, and its working gas is nitrogen. The coatings were deposited onto steel substrate. The phase structure and the composition of the coatings were analyzed by means of scanning electron microscopy (SEM) and X-ray diffraction (XRD), and their Vickers hardnesses were measured. In addition, titanium particles that were generated by MPD arcjet generator ablation process were observed by means of SEM and electron probe micro analysis (EPMA). These analyses showed that the MPD spray process could successfully form dense and uniform titanium nitride coatings. The properties of the titanium nitride coatings were highly sensitive to the titanium cathode diameter and discharge current.

A small thrust measurement system of electric thruster

Abstract: A small thrust measurement method was presented in theory for precision measurement of small thrust of low power electric thruster, namely when thrust measure balance is in the state of indifferent equilibrium (or near to the state of indifferent equilibrium), electromagnetism force feedback compensation can eliminate the influence of connection of gas line and electric wire and feedback electromagnetism force can compensation thrust of thruster. Thrust values of different operational mode of arcjet thruster were measured. Experiment validated that this measure method can measure small thrust of electric thruster accurately.

Radiometric Analysis from the 26-kW Electric Propulsion Space Experiment (ESEX) Flight

Abstract: : The United States Air Force Research Laboratory's Electric Propulsion Space Experiment (ESEX) was launched and operated in early 1999 in order to demonstrate the compatibility and readiness of a 30 kW class ammonia arcjet for satellite propulsion applications. As part of this flight, an array of on-board contamination sensors was used to assess the effect of the arcjet and other environments on the spacecraft. The sensors consisted of microbalances to measure material deposition, radiometers to assess material degradation due to thermal radiation, and solar cell segments to investigate solar array degradation. Over eight firings of the ESEX arcjet, (3 minutes, 26 seconds operating time) the radiometer near the thruster, viewing the arcjet plume and body, experienced a change in the thermal properties of its coating (repetitive) Radiometers with no view of the arcjet, or a view of only the plume, show no change. In general, contamination effects are observed only on sensors near the thruster exhaust nozzle, a location unlikely to be used in an operational high-power electric propulsion system. No contamination effects are observed in the backplane of the thruster. For future programs, while engineering measures may be needed for spacecraft equipment in the immediate vicinity of the thruster body, the arcjet environment is generally benign.

Solar Cell Degradation During the 26kW Electric Propulsion Space Experiment (ESEX) Flight

Abstract: : The United States Air Force Research Laboratory's Electric Propulsion Space Experiment (ESEX) was launched and operated in early 1999 in order to demonstrate the compatibility and readiness of a 30-kW class ammonia arcjet for satellite propulsion applications. As part of this flight, an array of on-board contamination sensors was used to assess the effect of the arcjet and other environments on the spacecraft. The sensors consisted of microbalances to measure material deposition, radiometers to assess material degradation due to thermal radiation, and solar cell segments to investigate solar array degradation. During firings, the solar cell segments show decreasing open-circuit voltage probably attributable to an additional electrical load provided by a short through the plume plasma. Over eight firings of the ESEX arcjet, and 33 minutes, 26 seconds operating time, the solar cells also exhibit a 3% decrease in short-circuit current, attributable to decreased solar transmission of the cover glass. However, no effects associated with the arcjet are observed on the spacecraft solar arrays. In general, the contamination effects are observed only on the solar cells sensor segments placed very near the thruster exhaust nozzle. In the backplane of the thruster, where the main arrays are located, no deleterious effects are observed, indicating that while engineering measures may be required for equipment in the immediate vicinity of the thruster, the arcjet environment is generally benign.

Conceptual Design of a Synergistic Propulsion System for the International Space Station (ISS)

Abstract: In this study, three different concepts of synergistic propulsion systems for the orbit control system of the International Space Station (ISS) are proposed. Their application was examined by means of numerical simulation. As propellants, selected waste fluids from the Environmental Control and Life Support System (ECLSS) are used. Possible future evolutions towards a highly regenerative life support system are taken into account. Resistojets and/or arcjets were chosen as thrusters. A detailed concept of an arcjet fueled by waste fluids was developed. It is based on the development of direct-current-arc-plasma generators suitable for the utilization of oxygen- containing gases like water steam, oxygen and carbon dioxide. The operation of the propulsion concepts onboard ISS was examined numerically with the software tool IRIS++, which allows the attitude and orbit dynamics of spacecrafts to be simulated. Results demonstrate that the drift phase during a reboost cycle can be increased significantly. A comparison of the system mass between the synergistic propulsion systems and the conventional system shows a high potential of reducing mass supplies to ISS.

Preliminary Study of Arcjet Neutralization of Hall Thruster Clusters

Abstract: Clustered Hall thrusters have emerged as a favored choice for extending Hall thruster options to very high powers (50 kW – 150 kW). This paper examines the possible use of an arcjet to neutralize clustered Hall thrusters, as the hybrid arcjet-Hall thruster concept can fill a performance niche amongst available propulsion options. We examine missions on which this hybrid concept would be a competitive or favored thruster option, report on fundamental experiments to understand how much electron current can be drawn to a surrogate anode from the plume of low power arcjets operating on hydrogen and helium, and then demonstrate the first successful operation of a low power Hall thruster-arcjet neutralizer package. In the surrogate anode studies, we find that the drawing of current from the arcjet plume has only a weak effect on overall arcjet performance (thrust), with a slight decrease in arc voltage with increased extracted current. A single Hall thruster – arcjet neutralizer package was constructed for the hybrid concept demonstration. The arcjet operated at very low powers (~ 70-120W) on helium, at a mass flow rate of 4.5 mg/s, and was able to effectively neutralize the ~ 200 – 900W xenon Hall thruster causing little measurable departure from the hollow-cathode neutralized Hall thruster VI characteristics up to 250V. At higher helium mass flow rates, the Hall discharge current is slightly perturbed from its expected values, due most likely to the ingestion of helium. Further developments of the hybrid concept to clustered configurations and higher powers will require a vacuum facility that can pump tens of milligrams of helium while maintaining the low pressures needed for normal xenon Hall thruster operation.

Numerical Analysis of the Rarefied Plume from an Arcjet Thruster: Effect of the Nozzle Exit Shape

Abstract: The plume flow field of an arcjet thruster with hydrazine decomposed gas as the propellant has been numerically analyzed to reveal a backflow field, using Direct Simulation Monte Carlo (DSMC). Hydrazine decomposed gas is approximated by a mixture of gas composed of N2, H2, and H. Four different nozzle lip shapes have been prepared to investigate their effects on the backflow, and species existing in the backflow have also been examined. The present results of mass flux at various angles measured from the plume axis have been compared with experiments, and a qualitatively fairly good agreement has been obtained.

Simulation and design of a hydrogen arcjet thruster seeded with cesium

Abstract: Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.

Experimental and Numerical Studies of a Low-Power N2 Arcjet Thruster

Abstract: This paper describes the experimental and numerical studies of a laboratory model of the low-power nitrogen arcjet thruster that was developed to provide thruster performance data to validate numerical results. The arcjet thruster was operated by using a nozzle 1.0mm in constrictor diameter. Thrust and input power were measured for various arc currents and nitrogen mass flow rates. The operation was done at power levels ranging from 156W to 540W and nitrogen mass flow rates from 5mg/s to 30mg/s. Typical specific impulse obtained in the experiment was 188s at 542W. Numerical simulation was conducted by using the physical model of a thermochemical nonequilibrium gaseous flow, a two-temperature model consisting of heavy particle and electron temperatures. The flowfield equations were numerically solved by combining with the Maxwell’s equation and the generalized Ohm’s law. It is shown that the predicted thruster performance is higher than the experimental data for the specific impulse, and the possible causes for this trend are discussed.

Refined Orbital Performance Measurements of the Air Force Electric Propulsion Space Experiment (ESEX) Ammonia Arcjet

Abstract: : During the Electric Propulsion Space Experiment (ESEX) mission, eight firings of the 26 kW ammonia arcjet were performed. Data taken from on-board systems, GPS and ground tracking during these firings are used to determine thruster performance. The on-board Servo Accelerometer Assembly (SAA) measured spacecraft acceleration. The mean values of thrust, specific impulse and thrust efficiency are 1.93 +/- 0.06 Newtons, 786.2 +/- 43.0 seconds and 0.267 +/- 0.021, respectively. This measured performance is lower than expected based on ground test. The most likely cause of this discrepancy is onboard measurement error in discharge power due to a 6% drift in the power processing unit current shunt. At the corrected power, performance falls within the expected envelope.

Plural voltage minima in an arc-heated channel flow

Abstract: In flows through a channel with varying cross-sectional area, the impulse and total enthalpy can be increased by superimposing an electrical discharge. The flow field is determined from the inlet flow condition, channel geometry, and discharge specifications. In this study, steady-state, quasi-one-dimensional flows interacting with an arc discharge are computed numerically. Once the arc column configuration is given, the discharge voltage is computed from the solution of flow field variables. For a constant discharge current, there exist plural column configurations which yield a minimum discharge voltage. This result explains the fluid-dynamic mechanisms of the existence of plural voltage modes in an arcjet operation.

High-energy plasma acceleration by means of an ideal arcjet thruster

Abstract: A theoretical study is presented on the entropy acceleration of a plasma by means of an ideal arcjet thruster made up of two major parts; an arc column of a uniform temperature between the cathode and the anode, and the heating zone of the injected cold gas. The plasma flow through the arc column follows the Navier–Stokes, since its dynamic viscosity is quite high. Proper boundary values on the cathode surface give a solution that yields sharp plasma acceleration to a high energy by the viscous force. In the heating zone, the flow is driven up mainly by a large temperature difference between the arc and the injected cold gas. The viscous force plasma driving model is applied to the solar wind; good agreement is confirmed between the theory and the observed data.

Experiments and numerical simulations in low power arc jets

Abstract: This paper describes experimental setup and performance evaluation of laboratory model of low power arcjet thrusters. Radiation-cooled arcjet thrusters were operated in electric power below IkW with nitrogen and argon as propellant respectively, using a nozzle with 0.5~1.Omm in constrictor diameter and propellant mass flow rates varying from 20mg/sec to 40 mg/sec. Corresponding test parameters were measured including thrust, arc current, arc voltage, propellant mass flow rate, inlet pressure. Actual specific impulse, thrust efficiency and ratio of thrust and power were calculated. Parallel to the experiments a numerical code system was developed to compare with experimental data and further optimize the arcjet thruster. The numerical method is referred as a Non-Oscillatory and Non-free-Parameters Dissipative finite difference scheme (NND). Twodimension, axisymmetric, compressible channel flow N-S equation coupled with electromagnetism equation was solved. Experiments and calculations showed the agreement that operating performances using N2 as propellant got much better than that of Ar, and presented that thruster with different structure and size had an optimal propellant mass flow rate with the best operating state. Nomenclature B Magnetic field vector E, F Convection vector in general coordinate E Electric field vector EV,FV Diffusion vector F Thrust * PH.D. student, School of Astronautics, BUAA f Professor, School of Astronautics, BUAA Copyright © 2001 by the American Institute of Aeronautics and Astronautics Inc. All right reserved. Fc Thrust of cold gas h Specific enthalpy H, Hv Source vector / Current