Showing papers on "Arcjet rocket published in 1994"

Near and far-field plume studies of a 1 kW arcjet

Abstract: In order to begin the process of characterizing transport in arcjet plumes, profiles of electron temperature, number density, pressure, and flow field patterns were obtained over an extensive volume of the plnme of a 1 kW arcjet operating on hydrogen. Axial and radial measnrements made over a region of the plume that extends from the arcjet exit plane to over one meter downstream of it are reported. Emission spectroscopy measurements of electron temperature were made near the exit plane of the arcjet. All experiments were performed in a 6 m by 9 m vacuum chamber with a pumping speed of over 100,000 l/s on hydrogen. Tank pressure was typically maintained to less than 3.8 x lor4 Torr during arcjet operation. In addition to the plume study, arcjet performance measurements are reportcd. 2 re rP kT, = electron temperature, eV V = probe bias, V Z, = ion charge vi,, = thrust efficiency y = ratio.of specific heats AD = Debye length, em A, = electron mean free path, cm A, = ion mean free path, cm aP = plasma potential, V x p = dimensionless probe potential 0 = flow angle = nozzle exit radius, m = radius from plume centerline, m = Langmuir probe collector electrode diameter, em r

Multiplexed Laser Induced Fluorescence and Non-Equilibrium Processes in Arcjets.

Abstract: : For the past five years there has been an ongoing experimental and analytical program at the University of Tennessee Space Institute (UTSI) to improve our understanding of arcjet physics. A computational model that assumed local thermodynamic equilibrium was first used to simulate arcjet thrusters operating on ammonia, hydrogen, and argon. The UTSI arcjet code was later extended to include a two temperature, finite rate kinetic model for hydrogen plasma. Recently, this code has been used to simulate a radiation-cooled arcjet (MARC thruster) experiment and a water-cooled arcjet (TT1 thruster) experiment performed at The Universitat Stuttgart Institut fur Raumfahrtsysteme. The results of these simulations are presented along with a review of UTSI arcjet computation code development. A two-beam multiplexed laser induced fluorescence (LIF) technique was developed at UTSI to provide detailed measurements of arcjet flows near the nozzle exit plane. Comparison of detailed flowfield measurements with predictions of the computation model were used to provide insight into the physical models used in the arcjet code. The method was first demonstrated using a small, 300 W, water-cooled arcjet operated with argon propellant. The method was then applied to a 1 kW arcjet operated with hydrogen and nitrogen propellant mixtures using the Balmer alpha line of hydrogen. Recently, the method has been extended to use an excited state line in nitrogen. The results of this most recent research are presented. (MM)

Nonintrusive electron number density measurements in the plume of a 1 kW arcjet using a modern microwave interferometer

Abstract: Reported is the use of a microwave interferometric technique for making nonintrusive measurements of plasma electron number density in the plume of a space electric propulsion thruster. The technique is capable of providing good sensitivity and accuracy as well as resolution using modern microwave network analyzer technology. Density profiles were obtained throughout the plume of a 1 kW hydrogen arcjet based on accurate microwave differential phase measurements. Spatially resolved integrated phase shifts for a 17.5 GHz signal radiated through the plume at various radial positions were Abel inverted to calculate radial electron density profiles. All measurements were taken in the University of Michigan's Large Chamber Plasma Facility, a 6 m by 9 m vacuum chamber, at pressures of 2/spl middot/10/sup -4/ Torr or less. The interferometer measured electron densities as low as 1/spl middot/10/sup 15/ m/sup -3/ with a predicted capability to measure peak densities as high as 3/spl middot/10/sup 18/ m/sup -3/. The accuracy of this technique is estimated to be on the order of /spl plusmn/10%. Comparison with Langmuir probe electron number density measurements demonstrate relative agreement between the two methods. Further, a previously reported tendency for the Langmuir probe to underpredict electron number density was consistent with our measurements. It is postulated that this underprediction may be due to small Langmuir probe perturbations in the local plasma of the far-field plume and errors in estimation of the probe collection area. >

Small satellite electric propulsion options

Abstract: Low power electric propulsion offers attractive options for small satellite propulsion. Applications include orbit raising, orbit maintenance, attitude control, repositioning, and deorbit of both Earth-space and planetary spacecraft. Potential electric propu+lsion technologies include very low power arcjets, Hall thrusters, and pulsed plasma thrusters, all of which have been shown to operate at power levels consistent with currently planned small satellites. Mission analyses show that insertion of electric propulsion technology enables and/or greatly enhances many currently planned small satellite missions. Examples of commercial, DoD, and NASA missions are shown to illustrate the potential benefits of using electric propulsion. Introduction The current emphasis on cost reduction and spacecraft downsizing has forced a reevaluation of technologies with critical impact on spacecraft mass. For a large number of commercial, scientific, and DoD near-Earth missions, on-board propulsion is the largest single spacecraft mass driver. Therefore, high performance electric propulsion systems offer high leverage for reducing injected mass requirements. Additional issues resulting from the use of new launch vehicles, spacecraft architectures, and the costs associated with ground testing and handling toxic fuels have also led to the consideration of electric propulsion technologies. Small spacecraft require propulsion for a wide range of on-orbit functions, including orbit raising and adjustment, drag make-up and stationkeeping, sunsynchronous orbit maintenance, and satellite orientation control. In addition, new communications and remote sensing markets and requirements for constellation maintenance and de-orbit are emerging. This diverse set of propulsion functions results in a wide range of propulsion requirements. Figure 1 shows the total impulse required by a number of planned NASA, DoD, and commercial small spacecraft. The values range from a low of 1.4 x lo4 N-s for the HETE spacecraft1 to a high of 2.5 x lo6 N-s for the Vesta asteroid rendezvous m i ~ s i o n . ~ Commercial spacecraft, not identified by name in the figure because of their proprietary nature, also require a wide range of total impulses. For all these spacecraft the propulsion system wet mass is the largest mass bus subsystem and This paper is declared a work of the U.S. Government and is not subject ot copyright protection in the United States. thus improvements in this subsystem will have the largest payoff for satellite mass reduction. Near-term electric propulsion options for small, power limited, spacecraft include very low power arcjets, Hall thrusters, and pulsed plasma thrusters (PPTs). While the planned spacecraft power range, shown in Fig. 2, is quite large, there is a clear need for electric propulsion systems which consume less than 500 W of power. 1.8 kW arcjets are currently flying on AT&T's Telstar 4 satellite, and arcjets have been successfully operated at power levels below 100 W. However, arcjet performance was found to degrade substantially at power levels below 500 w . ~ Hall thrusters were flown on over 60 Soviet and Russian spacecraft4 PPTs, which use solid cloroflourocarbon propellant, have been operational on several spacecraft for over 20 years.5 PPTs have several unique features which make them attractive for small satellite missions, including their simplicity, use of inert, non-toxic propellants, and their ability to operate over a wide input power range at constant performance via simple changes in pulse frequency.5 This paper reviews the status of these small satellite electric propulsion options, and provides examples of commercial, DoD, and NASA missions for which electric propulsion offers significant benefits. Electric Propulsion Options Verv Low Power Arciets A highly simplified schematic of an arcjet thruster system is shown in Fig. 3. In operation, an arc initiates from the tip of the cathode and is forced by the propellant flow through the throat to seat diffusely in the diverging section of the nozzle which also functions as the anode of the device. Current generation arcjets use hydrazine propellant so as to be compatible with flight qualified propellant feed systems. The arcjet power processing unit (PPU) must ignite the discharge and reliably operate the thruster in both the period of transition immediately following startup and in the steady state mode. Arcjet loads exhibit negative slope impedance characteristics and a high impedance, or constant current, output is required for stable steady state erati ti on.^^^ Operating voltages are on the order of 100 V. For arcjet applications to date, pulse-widthmodulated (PWM) current mode has been used for the PPU. The selection of power stages has largely been determined by the power level required and the spacecraft bus voltage. In general, for applications below approximately 2 kW, push-pull or parallel converters have been used6 To ignite the discharge, a pulsed high voltage starting technique is typically ~ s e d . ~ . ~ In practice, this pulse is generated by an integral pulse winding on the output inductor. Current spikes at startup have been found to be a major cause of electrode damage and the current mode control of the PWM system provides excellent control of arc current at startup as it limits the inrush current following ignition. Typical thruster performance during steady state operation at power levels between 500 and 800 W ranges from 26 to 41 percent efficiency at between 320 and 530 s specific impulse (Isp). PPU efficiencies are over 90 percent, and arcjet system masses (excluding the propellant storage and feed system) are near 7 kg.

Arcjet for a space flying body

Abstract: A thermal arcjet propulsion plant is particularly constructed for position and orbit corrections in a space flying body or spacecraft. A housing forms an anode and an expansion nozzle. An elongated rod mounted in an electrically insulating manner centrally in the housing forms the cathode which reaches with its tip into a plenum chamber just upstream of the nozzle. Propellant components formed by dissociating ammonia are injected into the plenum chamber. A small gap is provided between the tip of the cathode and the restrictor of the nozzle. An arc is ignited in this gap and the thermal arc energy is taken up by the by propellant. A dissociation chamber for the by propellant is arranged around the housing. The arrangement is such that the chamber in which the gas dissociation takes place is arranged in series with and upstream of the plenum chamber. The cathode is also arranged in series with the restrictor and the nozzle. The cathode is connected through a flexible electrical conductor to an electric power supply. Similarly, the housing forming the anode is connected to the same power supply. The dissociation chamber includes a catalyst for a catalytic and/or thermal dissociation of the ammonia into its components of hydrogen and nitrogen.

Multifluid nonequilibrium simulation of arcjet thrusters

Abstract: Thesis (Sc. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1994.

Discharge features in a steady-state nitrogen arcjet engine with an expansion nozzle

Abstract: Nitrogen discharges in a 10-kW-class water-cooled arcjet engine were studied by means of spectroscopic diagnostics. The atomic excitation temperature and electron density in the constrictor increased from 15,000 to 18,000 K and from 1/spl times/10/sup 15/ to 2/spl times/10/sup 16/ cm/sup /spl minus/3/, respectively, with input power levels of 7-11 kW. An increase in the mass flow rate raised the electron density because of enhanced thermal pinch effect of the arc column. In the expansion nozzle, the pressure and electron density decreased drastically downstream, and therefore the plasma was expected to be in high non-equilibrium. The N/sub 2//sup +/ vibrational temperature reached about 6,000 K at the nozzle exit, and the N/sub 2//sup +/ rotational temperature decreased gradually resulting in a range of 1,000-2,500 K. >

Interior Spectroscopic Investigation of the Propellant Energy Modes in an Arcjet Nozzle

Abstract: : Internal emission spectroscopy measurements were performed in the nozzle expansion region of a 26 kW class ammonia arcjet. A series of three optical access ports (0.020-inch diameter) were equally spaced along the center line of the expansion nozzle. Atomic and ionic excitation temperatures of H and NII were determined through the use of Boltzmann plots. Electron density measurements were also taken based upon Stark broadening of the hydrogen beta Balmer transition. It is believed that a change in the mechanism governing the electron density occurred just downstream of the constrictor between the first two access ports lending evidence that arc attachment is occurring in this region. Although the trends for excitation temperatures and electron density were similar to those previously observed in 1 kW arcjets, the temperature values were found to be significantly less. In addition, vibrational and rotational temperatures of NH were measured and appeared to be frozen throughout the expansion nozzle.

Radiation transfer computation in cylindrical arc columns using a Monte Carlo method

Abstract: A one-dimensional, axisymmetric, Monte Carlo radiative transfer method has been developed to calculate the energy transfer and temperature profiles of moderate to high pressure cylindrical electric arc columns. The model includes the energy transfer due to the emission and absorption of a non-gray gas and thermal conduction but neglects convection. The thermal and electrical conductivities and the wavelength dependent spectral absorptivities are calculated based on assumption of local thermodynamical equilibrium (LTE). To verify the method, calculations were performed for the case of an electric arc in air at 30 atm pressure. The results were compared to those obtained from an earlier calculation which used the diffusion approximation for the radiative transfer. This case is representative of conditions that are typical of constrictors in arc heaters. The method was then applied to an arc in NH 3 at 1 atm for conditions that are typical of the constrictor in arcjet thrusters. The geometrical part of the method was then extended to enable a fully two-dimensional, axisymmetric simulation of radiative transfer for generic geometrical shapes.

Discharge ignition behavior of the Space Station plasma contactor

Abstract: Ignition testing of hollow cathode assemblies being developed for the Space Station plasma contactor system has been initiated to validate reliable multiple restart capability. An ignition approach was implemented that was derived from an earlier arcjet program that successfully demonstrated over 11,600 ignitions. For this, a test profile was developed to allow accelerated cyclic testing at expected operating conditions. To date, one hollow cathode assembly has been used to demonstrate multiple ignitions. A prototype hollow cathode assembly has achieved 3,615 successful ignitions at a nominal anode voltage of 18.0 V. During the ignition testing several parameters were investigated, of which the heater power and pre-heat time were the only parameters found to significantly impact ignition rate.

Axial emission measurements on a medium power hydrogen arcjet thruster

Abstract: A numerical and experimental investigation of the axial emission of a 5 kW class radiatively-cooled hydrogen arcjet thruster is presented. The complete visible spectrum is modeled, taking into account graybody thermal emission from the cathode and plasma radiation from hydrogen atoms and free electrons. The cathode temperature and the arc electron number density are obtained from the measured emission spectrum. The cathode temperature is found to be in the neighbourhood of the melting point of tungsten and increases with arcjet power at a constant massflow rate, suggesting an increase in the current density at the arc attachment point. The measured near-cathode electron number density is also found to increase with power, suppofling the idea of an increasing arc current density with arcjet power. The measured and modeled arc electron densities are, however, in discrepancy by a factor of approximately three.

Interior plasma diagnostics of arcjet thrusters

Abstract: We review past experimental measurements of internal flow properties of arcjet thrusters. These measurements are generally classified as either intrusive, requiring design changes to prototype thrusters, or nonintrusive, and include measurements of cathode temperature, as well as static pressure, flow temperatures (vibrational, rotational, electronic, translational), electron density, and velocity throughout the interior region extending to the exit plane. Comparisons are made to available model predictions. These measurements, performed on a wide range of thrusters, and operating on variety of propellants, indicate that the nozzle plasma flow may be removed from local thermodynamic equilibrium.

Increased efficiency arcjet thruster

Abstract: There is disclosed an anode for an electrothermal arcjet thruster. The anode includes a divergent nozzle having, in tandem, a recombination portion and an expansion portion. The expansion portion has a greater rate of divergence than the recombination portion. A regeneration channel containing flowing propellent gas extends internally within the anode adjacent the recombination portion. The combination of the biangle nozzle and the regeneration channel produces an arcjet thruster having a markedly increased thrust efficiency.

Plasma diagnostics of a direct current arcjet diamond reactor. I. Electrostatic probe analysis

Abstract: Plasma parameters were determined for a direct current arcjet diamond chemical vapor deposition reactor. Electron temperatures and densities were measured using a floating double probe technique, yielding values ranging from 2.3 eV and 7×1011 cm−3 near the plasma gun nozzle to 0.4 eV and 1×109 cm−3 in the downstream and peripheral regions of the plasma. These results show that the plasma is weakly ionized, indicating that reactions involving electrons play a minor role in the macroscopic gas–phase chemistry though high energy products from these reactions may significantly affect film quality.

System-level requirements for an operational solar electric orbital transfer vehicle

Abstract: The concept of using a solar electric orbital transfer vehicle (SEOTV) propulsion systems as a replacement for chemical propulsion systems has been examined on the merits of performance enhancement, economic benefit, and operability. This paper summarizes the system-level requirements for an operational SEOTV that were generated over the course of the study. The requirements provided are the result of detailed system-level trades covering complete vehicle performance, costs, and operational characteristics. Top-level requirements are defined for the SEOTV system and each of the major subsystems. Major operational requirements are also given. Results indicate that advanced, low-cost solar arrays will be required to provide the necessary economic pay-offs as will highly efficient hydrogen arcjet thruster systems. An advanced technology cryogenic propellant storage system will also be required for maximum payload capability. 20 refs.

New Developments and Research Findings: NASA Hydrazine Arcjets

Abstract: In 1984, the market for commercial geosynchronous communications satellites (comsats) was expanding and there was strong competition between spacecraft builders for market share. The propellant required for the north-south stationkeeping (NSSK) function was a major mission limiter, and the small chemical and resistojet systems then in use were at or near their physical limits. Thus, conditions were right for the development of a high performance NSSK system, and after an extensive survey of both propulsion technologies and the aerospace community, the NASA program chose hydrazine arcjets for development. A joint government/industry development program ensued which culminated in the acceptance of arcjet technology. NASA efforts included fundamental feasibility assessments, hardware development and verification, and multiple efforts aimed at the demonstration of critical operational characteristics of arcjet systems. Throughout the program, constant contact with the user community was maintained to determine system requirements. Both contracted and cooperative programs with industry were supported. First generation, kW-class arcjets are now operational for NSSK on the Telstar 401 satellite launched in December of 1993 and are baselined for use on multiple future satellite series (Intelsat 8, AsiaSat, Echostar). Arcjet development efforts are now focusing on the development of both high performance (600 s), 2 kW thrusters for application on next generation comsats and low power (Pe approximately 0.5 kW) for a variety of applications on power limited satellites. This paper presents a review of the NASA's role in the development of hydrazine arcjets with a focus on approaches, lessons learned, and the future.