Arcjet rocket

About: Arcjet rocket is a research topic. Over the lifetime, 1121 publications have been published within this topic receiving 9687 citations. The topic is also known as: Arcjet.

Multidimensional Testing of Thermal Protection Materials in the Arcjet Test Facility

Abstract: Many thermal protection system materials used for spacecraft heatshields have anisotropic thermal properties, causing them to display significantly different thermal characteristics in different directions, when subjected to a heating environment during flight or arcjet tests. The anisotropic effects are enhanced in the presence of sidewall heating. This paper investigates the effects of anisotropic thermal properties of thermal protection materials coupled with sidewall heating in the arcjet environment. Phenolic Impregnated Carbon Ablator (PICA) and LI-2200 materials (the insulation material of Shuttle tiles) were used for this study. First, conduction-based thermal response simulations were carried out, using the Marc.Mentat finite element solver, to study the effects of sidewall heating on PICA arcjet coupons. The simulation showed that sidewall heating plays a significant role in thermal response of these models. Arcjet tests at the Aerodynamic Heating Facility (AHF) at NASA Ames Research Center were performed later on instrumented coupons to obtain temperature history at sidewall and various radial locations. The details of instrumentation and experimental technique are the prime focus of this paper. The results obtained from testing confirmed that sidewall heating plays a significant role in thermal response of these models. The test results were later used to validate the two-dimensional ablation, thermal response, and sizing program, TITAN. The test data and model predictions were found to be in excellent agreement.

Atomic hydrogen concentration in a diamond depositing dc arcjet determined by calorimetry

Abstract: The fraction of hydrogen dissociated in the plume of a dc arcjet used for diamond deposition is determined by calorimetry to be 0.33±0.12. A dc arc is struck in a mixture of argon and hydrogen at 90 psi and the effluent is expanded through a converging/diverging nozzle into a reactor maintained at 25 Torr. Methane (<1%) is added to the luminous gas plume in the diverging nozzle. This supersonic jet impinges on a water cooled molybdenum substrate, and diamond thin film grows from the reactive mixture. The electrical power input of the arcjet (1.6 kW) is balanced by the power losses due to cooling of the nozzle, enthalpy change in the gas, ionization of the gas, dissociation of H2, and the directed velocity of the gas phase. The gas temperature is determined by linear laser-induced fluorescence (LIF) measurements of several rotational lines of NO seeded to the gas plume. The velocity of the gas plume is obtained via the Doppler shift between LIF signals measured simultaneously in a stationary reference cell...

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.

Hydrogen arcjet technology

Abstract: During the 1960's, a substantial research effort was centered on the development of arcjets for space propulsion applications. The majority of the work was at the 30 kW power level with some work at 1-2 kW. At the end of the research effort, the hydrogen arcjet had demonstrated over 700 hours of life in a continuous endurance test at 30 kW, at a specific impulse over 1000 s, and at an efficiency of 0.41. Another high power design demonstrated 500 h life with an efficiency of over 0.50 at the same specific impulse and power levels. At lower power levels, a life of 150 hours was demonstrated at 2 kW with an efficiency of 0.31 and a specific impulse of 935 s. Lack of a space power source hindered arcjet acceptance and research ceased. Over three decades after the first research began, renewed interest exists for hydrogen arcjets. The new approach includes concurrent development of the power processing technology with the arcjet thruster. Performance data were recently obtained over a power range of 0.3-30 kW. The 2 kW performance has been repeated; however, the present high power performance is lower than that obtained in the 1960's at 30 kW, and lifetimes of present thrusters have not yet been demonstrated. Laboratory power processing units have been developed and operated with hydrogen arcjets for the 0.1 kW to 5 kW power range. A 10 kW power processing unit is under development and has been operated at design power into a resistive load.