Showing papers on "Arcjet rocket published in 2007"

Computational Simulation of High Enthalpy Arc Heater Flows

Abstract: The flowfields in segmented constrictor-type arc heaters are simulated using a new Navier-Stokes code named ARCFLO3. The validity of the radiation and turbulence modeling employed in ARCFLO3 is assessed by comparing the calculated results with the existing experimental data obtained in the 20 and 60 MW arcjet facilities at NASA Ames Research Center. Comparison is made between the calculated and the measured data for arc voltage, heater efficiency, mass-averaged enthalpy, chamber pressure, heat flux at wall, and total enthalpy in the centerline region of an arcjet flow

Calculation of Thermal Response of Ablator Under Arcjet Flow Condition

Abstract: An integrated computational method is developed to calculate thermal response of ablator under an arcjet flow condition. In the method, the arcjet freestream condition in the test section is evaluated by calculating the flows in the arcjet wind tunnel fully theoretically. The thermal response of the ablator is calculated by loosely coupling the shock layer computational fluid dynamics code and the 2-D version of ablation code using the arcjet freestream condition so evaluated. The method is applied to the heating tests conducted in the 1 MW arcjet wind tunnel for one operating condition. The influence of catalytic conditions of ablating surface and the effect of nitridation reaction and surface roughness on the thermal response of the ablator are investigated. Comparison of the temperature profile at the ablating surface between calculation and measurement suggests that the measured temperature profile can be reproduced with a low catalytic efficiency of the surface. It is found that the nitridation reaction increase the surface temperature moderately, and that the effect of the roughness on the surface were small for the present operating condition.

Laboratory Experiment of Plasma Flow Around Magnetic Sail

Abstract: To propel a spacecraft in the direction leaving the Sun, a magnetic sail (MagSail) blocks the hypersonic solar wind plasma flow by an artificial magnetic field. In order to simulate the interaction between the solar wind and the artificially deployed magnetic field produced around a magnetic sail spacecraft, a laboratory simulator was designed and constructed inside a space chamber. As a solar wind simulator, a high-power magnetoplasmadynamic arcjet is operated in a quasisteady mode of 0.8 ms duration. It can generate a simulated solar wind that is a high-speed (above 20 km/s), high-density (1018 m−3) hydrogen plasma plume of ∼0.7 m in diameter. A small coil (2 cm in diameter), which is to simulate a magnetic sail spacecraft and can obtain 1.9-T magnetic field strength at its center, was immersed inside the simulated solar wind. Using these devices, the formation of a magnetic cavity (∼8 cm in radius) was observed around the coil, which indicates successful simulation of the plasma flow of a MagSail in the laboratory.

Calculation of High-Enthalpy Aerothermal Environment in an Arcjet Facility

Abstract: S EGMENTED-CONSTRICTOR type arc-heated wind tunnels are used to test the heat shield materials for spacecraft thermal protection systems. The arc-heated wind tunnel consists of an upstream electrode (anode) chamber, constrictor section, downstream electrode (cathode) chamber, and a diverging–converging nozzle connecting to a test chamber. In the test section, the heat shield materials are exposed to a high-enthalpy flow environment produced by the facility. The high-enthalpy environment is often such that the flowdoes not reach equilibrium condition at the edge of the boundary layer over the tested material. In such a case, we need to calculate the flow properties at the surface of the material using a computational fluid dynamics approach to understand the thermal response of the material [1]. For this purpose, the arcjet freestream conditions must be known accurately. To calculate an arcjet freestream condition, two important physical processes occurring in the arcjet wind tunnel should be accounted for: the heating process in the arc heater region upstream of the nozzle throat and the relaxing process in the expanding nozzle region downstream of the nozzle throat. The ARCFLO3 code has been developed recently to calculate the flowfield in the segmentedconstrictor type of arc heaters [2]. Unlike the arc heater flowfield code named ARCFLO developed in the 1970s [3], which is able to calculate the flow in the constrictor section, this new code calculates the flow from the anode chamber to the nozzle throat [2]. Arcjet freestream conditions can be calculated fully theoretically if a nonequilibrium expanding nozzle calculation is made with the calculated flow properties at the nozzle throat obtained by using the ARCFLO3 code. In addition, because the radial distribution of the flow properties at the nozzle throat is calculated with the ARCFLO3 code, the unified computational method can give the radial flow properties in the arcjet freestream at the test section. We tried to make such a unified calculation very recently for one operating condition in an arcjet facility [1]. However, the question remains as to how well such a computational approach predicts the flow properties in an arcjet freestream. It is the purpose of the present work to test the validity of the unified method. The method is applied to calculate the flowfield in a 0.75-MW arcjet wind-tunnel facility at the Institute of Aerospace Technology of the Japan Aerospace Exploration Agency (IAT/JAXA) in Japan. This facility was chosen for the following reasons: 1) In the recent measurement [4] in the IAT/JAXA arcjet facility, the operational characteristic parameters for a wide range of conditions were obtained. The experimental data offer an opportunity to test the current state of the computational modeling in the proposed method. 2) In our previous work, the ARCFLO3 code was applied for the arc heater flowfield calculation only in a high-power-level arc heater, such as the 20or 60-MW arcjet facility at NASA Ames Research Center [2]. The applicability of the ARCFLO3 code to submegawatt class facilities is unknown.

Numerical Study of the Electromagnetic Control of a Weakly Ionized Flow around a Blunt Body: Role of an Insulative Boundary in the Flow

Abstract: The electromagnetic flow control was simulated using magnetohydrodynamic equations including the Hall effect to clarify the electromagnetic effects measured in an arcjet flow. The result showed that the combined effect of the Hall effect and an insulative boundary in the flow activates the electromagnetic flow control: an ideal uniform ionized flow without any insulative boundary in the flow was found to be unsuitable to investigate the electromagnetic flow control experimentally. In addition, the present study computes the flow with an artificial insulative plume boundary to imitate the non-uniformity of the experimental flow. As a result, the measured drag increase was reproduced if an artificial insulative boundary is set to the location identical to the plume radius estimated on the basis of the measurements. Consequently, the measured drag increase in the arcjet flow results from the plume boundary serving as an insulative boundary.

Local Enthalpy Measurements in a Supersonic Arcjet Facility

Abstract: Results of the application of a mass-injection probe to estimate the local mass-specific enthalpy in a supersonic air plasma flow are presented. Up to now this probe has only been applied to subsonic flows. Numerical calculations to interpret the probe behavior are shown and compared to experimental results. Furthermore, to interpret the measured local mass-specific enthalpies, classical enthalpy estimation methods are used together with analytical calculations of the turbulent free plasma jet and fairly good agreement is found.

Generation of supersonic plasma flows using an applied-field MPD arcjet and ICRF heating

Abstract: A quasi-steady, supersonic plasma flow was produced by using a magneto-plasma-dynamic arcjet (MPDA) in a divergent magnetic nozzle. The ion acoustic Mach number, Mi, evaluated by spectroscopy and Mach probe measurement, was unity in the uniform field near the MPDA and reached almost 3 in the divergent magnetic nozzle. In the case of a subsonic flow near the exit of the MPDA, a magnetic Laval nozzle was effective in converting it to a supersonic flow with Mi = 1 at the throat. We heated ions of the fast-flowing plasma by ICRF (ion-cyclotron-range of frequency) in a magnetic beach field. The increased thermal energy was converted to flow energy as it passed through the divergent magnetic nozzle in accordance with the constant magnetic moment. We were able to successfully control the plasma flow with a wide range of densities for basic magnetohydrodynamic (MHD) studies and space thruster applications.

Computational analysis of semi-elliptical nozzle arcjet experiments : Calibration plate, wing leading edge

Abstract: This paper reports computational analysis in support of experiments in a high enthalpy arcjet wind tunnel at NASA Ames Research Center. These experiments were conducted in the NASA Ames 60-MW Interaction Heating Facility and include surface temperature measurements of swept-wing leading edge shaped pylon models. Surface temperatures of the arcjet pylon models were measured with thermocouples, an infrared camera, and a pyrometer. During the facility characterization runs, surface pressure and heat flux measurements on a water-cooled calibration plate were obtained. The present analysis comprises computational simulations of the nonequilibrium flowfield in the facility (the nozzle and the test box) and comparisons with the experimental measurements. The value of computational fluid dynamics simulations in planning and analysis of a complex arcjet test configuration is demonstrated.

TIHTUS Thrust Measurement with a Baffle Plate

Abstract: A novel two-stage, electric thruster TIHTUS is under development at IRS. The first stage is an arcjet, while the second stage is inductive heating of the arcjet plume. Due to the complex set-up of the thruster, it cannot be placed in a thrust balance. Thrust is therefore measured by a baffle plate. Over a variation of power- and mass-flow-ratios between the two stages, thrust is measured. In order to interpret the results, the present paper is foreseen to disclose radially resolved measurement of total pressure from which thrust can be determined. It is shown that the thrust obtained from the two measurement methods are in good agreement with each other and range between 1.7 and 2.5 N for the respective operating conditions of the thruster. Thrust decreases, the more power (of a sum of 50 kW) is coupled into the second, inductive stage. However, when heating is active in the second stage, at increasing mass flow rate (of a total of 300 mg/s) thrust increases.

Thrust termination dynamics of solid propellant rocket motors

Abstract: A CTIVE thrust termination or reversing of a separatingmotor are ways to facilitate the staging process and improve the performance of a multistage rocket [1–4]. This is done by opening a secondary nozzle located on the motor head and causing rapid depressurization of the motor chamber. During a rapid change in the chamber pressure the burning rate becomes a nonlinear function of pressure and its time rate of change. The strong dependence of the solid propellant burning rate on the chamber pressure results in a transient nonlinear behavior that may cause premature dynamic extinguishment of the motor. Solid propellants’ transient burning and extinguishment during rapid pressure change have been the subject of a number of studies and several transient burning models have been developed [5–8]. Zeldovich has proposed a comprehensive model based on the unsteady rate of heat transfer to the propellant surface and the time rate of change of the temperature distribution in the solid propellant [6]. This model requires specification of many experimentally determined parameters. The transient burning rate model proposed by Von Elbe and McHale is easy to implement and requires instantaneous pressure and its time rate of change for the regression rate prediction [5]. References [1,2] studying motor transient behavior have used a control volume approach for the gas dynamic simulation of the motor’s internal ballistics. However during the active thrust termination (or reversing) process there is a considerable pressure variation along the motor port and there is a rapid propagation of expansionwaves, through the thrust termination nozzles. These expansion waves play an important role in the transient thrust behavior and a more comprehensive flow model should be used for the internal ballistics’ simulation.

Photogrammetric recession measurements of an ablating surface

Abstract: An instrument and method for measuring the time history of recession of an ablating surface of a test article during testing in a high enthalpy thermal test facility, such as an arcjet. The method advances prior art by providing time-history data over the full ablating surface without targets and without any modifications to the test article. The method is non-intrusive, simple to implement, requires no external light source, and does not interfere with normal operations of the arcjet facility.

An Advanced Ammonia Propellant Feed System for the Thermal Arcjet TALOS

Abstract: Previous investigation of the propellant feed system for the thermal arcjet thruster of the lunar mission Bw1 concluded in the necessity that the pressure regulator of the propellant feed system needs to be modified to reach a higher pressure level inside the propellant feed system. The mass flow rate showed fluctuations of 15%, which are not acceptable for stable arcjet operation. To sustain a high temperature behind the vaporizer the propellant feed system is insulated by use of space blankets wrapped around the pipings. A new vaporizer concept is developed increasing the heat transfer into the ammonia by use of stainless steel srub inside the vaporizer. The necessary heating power for complete vaporization is calculated to be 50W. The results of the tests with the modificated propellant feed system show that first the fluctuations of the mass flow rate disappear by use of the modified pressure regulator. Secondly, the higher pressure level is reached but the mass flow is to high. Consequently, the diameter of the flow aperture has to be decreased.

Microarcjet Microthruster for Nanosat Applications

Abstract: This paper summarizes experiments and analysis of the micro-discharge microarcjet thruster. A small, very low power (2-5 Watt) micro-discharge was maintained between two electrodes in geometries compatible with application to a cold gas thruster. To evaluate the efficiency of the discharge in providing an increase in Isp and thrust, a special torsional micro-balance thrust stand, capable of micro-Newton resolution, was designed and constructed. The micro-balance thrust stand was installed in a large dielectric chamber with high pumping speed to eliminate stray coupling of the discharge with the vacuum chamber or background gas. A battery operated driver provided the 240-280 volts required to sustain the very low current (5-40 mA) arc discharge. The discharge was also ballasted inductively to avoid capacitive effects as well as resistive losses. Thrust measurements using a variety of electrode geometries and propellants were carried out. It was found that the optimum scale for the discharge bore for these low power levels was ~ 300 micron. Smaller bores resulted in too large a power loss in thermal transfer to the thruster/nozzle body. A larger bore led to large mass flows at the pressures required to produce a stable discharge. All ring or cavity electrode structures showed no appreciable gain and were rapidly eroded by the discharge. The only electrode configuration to show an increase in thrust from the application of the discharge with no measurable electrode erosion was in a configuration similar to the standard arcjet. Most of the studies were conducted in Argon, but other gases from H to Xe were also used. In Argon the optimal mass flow was ~ 2 mg/s with between 10 to 20 kPa upstream of the nozzle. With 5 W of electrical power into the discharge the thrust was increased 1.1 to 1.8 mN with an effective Isp of 90 s. The highest efficiencies (~ 40%) were obtained at lower power indicating enhanced thermal losses at higher Isp (gas temperature).

Photogrammetric Recession Measurements of Ablative Materials in the NASA Ames Sixty-Megawatt Arcjet

Abstract: The photogrammetric recession measurement (PRM) technique has been applied in several tests of ablative thermal-protection system materials in the 60-MW arcjet at NASA Ames Research Center. This paper presents selected results from those tests to illustrate the capabilities and limitations of the technique.

Spraying Using Electromagnetically Accelerated Plasma

Abstract: In magneto-plasma-dynamic (MPD) arcjet generators, plasma is accelerated by electromagnetic body forces. Silicon nitride reactive spraying was carried out using an MPD arcjet generator with crystal silicon rods and nitrogen gas. Because higher-velocity, higher-temperature and higher-density and larger-area plasmas are produced with the MPD arcjet generator than those with conventional thermal plasma torches, nitriding of silicon can be enhanced. A dense and uniform β-Si3N4 coating with 30 μm in thickness was formed after 200 shots at a repetitive frequency of 0.03 Hz with a discharge current of 9 kA and a substrate temperature of 700 °C. The Vickers hardness reached about 1300. Furthermore, silicon carbide and aluminum nitride sprayings were conducted with some spraying systems. All results showed that the MPD arcjet generator had high potentials for spraying.

Preliminary Study of Arcjet Neutralization of Hall Thruster Clusters (Postprint)

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.

Effect of parameter variation on low power arcjet performance and stability

Abstract: A kind of arcjet structure for low power below 500 W and a related experiment system are presented.Performance parameters are obtained under different working conditions with nitrogen as propellant.The effect of the change of key dimensions and main working parameters on the performance and stability of the thruster is analyzed.Stable working parameters and laws are studied.

Calculation of Radiation from a Shock Layer Flow in an Arc-Jet Facility

Abstract: The radiation from the shock layer flow over the blunt body placed in a 20MW arcjet facility is analyzed. The nozzle throat condition is calculated using an arc heater flowfield code (ARCFLO3). Two-temperature thermochemical nonequilibrium codes are used to calculate the expanding flow in the nozzle region downstream of the throat and to calculate the shock layer flowfield over the blunt body. The radiation spectrum incident on the stagnation point of the blunt body is computed using a line-by-line radiation method. The calculated result is compared with the experimental spectrum over the wavelength region from 1000 to 9000˚ The result shows that the calculation underestimates the measurement by up to a factor of about 1000. Possible reasons for the discrepancy between calculation and measurement are discussed. The present result shows that it is likely that the measured spectrum include the radiation from the hot flow in the settling chamber region.

Scale-Model Experiment of Magnetoplasma Sail : Preliminary Results

Abstract: ** Magnetic sail (MagSail) is a deep space propulsion system, which uses the energy of the solar wind. MagSail produces an artificial magnetic field and captures the energy of the solar wind plasma to propel a spacecraft in the direction of the solar wind. In order to conduct a scale-model experiment of the plasma flow of a MagSail, we developed a solar wind simulator based on a magnetoplasmadynamic arcjet in a quasi-steady mode of about 1 ms duration. Based on scaling considerations, a solenoidal coil was designed and it was immersed into the plasma flow. In this setup, a magnetic cavity, which is similar to that of the geomagnetic field, was observed, although the magnetic cavity of MagSail is usually much smaller than the geomagnetic cavity of the Earth. It was experimentally confirmed that MagSail could produce thrust even in such ion Larmor scale. Also, an extension of this MagSail experiment to MagSail with plasma jet (M2P2 or MPS) is in progress, and some preliminary results are reported.

Plasma Flow Measurement by Mach Probes in GAMMA 10

Abstract: Calibration of an up-down type Mach probe is performed using a fast-flowing plasma produced by a magneto-plasma-dynamic arcjet. Mach probe data are compared with ion acoustic Mach numbers M i , which are calculated using a plasma flow velocity U p and an ion temperature T i measured by spectroscopy and electron temperature T e by Langmuir probe. The obtained data are also compared with Hutchinson's PIC simulation results in an unmagnetized plasma and are in good agreement with each other. First attempts to measure plasma flow field using a Mach probe are performed at the open-end section in GAMMA10. It is found that M i at r=0 is more than 2, namely a supersonic plasma flow is formed in the end-cell region.

Measurement of flow velocity of MPD arcjet in gamma 10

Abstract: At both end cells of GAMMA 10, two of Magneto-Plasma-Dynamic Arcjet (MPDA) are installed. MPDAs consists of two coaxial electrodes (tungsten stick and molybdic cylinder), and can produce quasi-stationary hydrogen plasmas. In GAMMA 10, the plasma generated by the MPDAs are used as the seed plasma. The main plasma is produced by ion cyclotron range of frequency (ICRF) waves in combination with the hydrogen gas injection. In the center at r=0 near the mirror throat of plug/barrier cell, the flow velocity and the mach number are measured with a mach probe in cases of several discharge voltages. The flow velocity and the mach number increase with the discharge voltage. The mach number of the seed plasma and the main plasma (from the confined region) is also measured in the radial direction. The flow velocity and the mach number of the seed plasma become small in the core region than in the peripheral region.