Showing papers on "Arcjet rocket published in 2008"

Two-Dimensional Ablation, Thermal Response, and Sizing Program for Pyrolyzing Ablators

Abstract: The purpose of this paper is to describe and demonstrate new capabilities in the two-dimensional implicit thermal response and ablation program. These expanded capabilities include grid options for flight and arcjet geometries, a sizing algorithm for the flight-type geometry, and an orthotropic thermal conductivity model. Applications to analysis of an orthotropic low-density carbon-phenolic material in arcjet and flight environments relevant to the Orion crew module are presented. For the arcjet environment, multidimensional conduction effects strongly influence the in-depth thermal response. For a lunar return flight environment, in the shoulder region of the crew module (where the radius of curvature is smallest), the thermal response is influenced by multidimensional conduction and by the orientation of the orthotropic material.

An analytical optimization method for electric propulsion orbit transfer vehicles

Abstract: Due to electric propulsion’s inherent propellant mass savings over chemical propulsion, electric propulsion orbit transfer vehicles (EPOTVs) are a highly efficient mode of orbit transfer. When selecting an electric propulsion device (ion, MPD, or arcjet) and propellant for a particular mission, it is preferable to use quick, analytical system optimization methods instead of time intensive numerical integration methods. It is also of interest to determine each thruster’s optimal operating characteristics for a specific mission. Analytical expressions are derived which determine the optimal specific impulse (Isp) for each type of electric thruster to maximize payload fraction for a desired thrusting time. These expressions take into account the variation of thruster efficiency with specific impulse. Verification of the method is made with representative electric propulsion values on a LEO‐to‐GEO mission. Application of the method to specific missions is discussed.

Numerical Simulation of Electromagnetic Flow Control in a One-Kilowatt Class Argon Arcjet Windtunnel

Abstract: The electromagnetic flow control was numerically investigated to clarify the electromagnetic effects measured in a one-kilowatt argon arcjet windtunnel. As a result, an uniform pre-ionized flow assumption was found to be unacceptable to reproduce the measured electromagnetic drag increase because the strong Hall effect in the flow dissipates the circumferential current. Contrastingly, in the simulation using an artificial pre-ionized flow with a distinct insulative plume boundary, a strong Hall electric field is induced and the flow is electromagnetically controllable through E×B drift. To verify this result, a one-kilowatt class arcjet windtunnel is solved instead of such an artificial plume environment using the Naiver-Stokes equations with thermochemical nonequilibrium and steady magnetohydrodynamic equations. As a result, the simulated arcjet plume has a strong electroconductive non-uniformity in the radial direction, and this non-uniformity was found to activate a strong Hall electric field and the resulting electromagnetic flow control even if the Hall parameter is over 40.

Numerical Simulation of Electromagnetic Flow Control in an Arcjet Plume

Abstract: Electromagnetic flow control was simulated using magnetohydrodynamic equations to clarify the electromagnetic effects measured in an arcjet flow with a strong Hall parameter which is of order several tens. 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. Moreover, an arcjet plume with a strong Hall effect and ion slip was solved instead of using the artificial insulative plume boundary. The result showed that the shock standoff increases due to the activated Hall electric field and circumferential current whose values are identical to that in the flow with the artificial insulative boundary. Consequently, these results verified the supposition that the non-uniformity in the radial direction is required to activate the electromagnetic control in a flow with a strong Hall effect and ion slip.

Efficiency Analysis of a Two-Stage Hybrid Electric Thruster

Abstract: A RCJET thrusters have been developed throughout the past few decades. They provide relatively high thrust at moderate exhaust velocities, and their plasma flows are characterized by high specific enthalpy and high flow velocity. These characteristics are, however, combined with the presence of steep radial property gradients as in a hot, energy-rich core with a relatively cold gas layer at its edge. When the limits to the nozzle exit velocity are imposed bywall and electrode materials, another possibility to transfer more power into an arc heated plasma flow is to reheat the relatively cold plume edges by means of another heating mechanism, as done in an afterburner. TIHTUS has been developed over the past four years and is a novel two-stage plasma thruster where reheating of the arc heated plume is realized bymeans of inductive heating, as sketched in Fig. 1 [1–3]. It is therefore considered the predecessor of a future propulsion system for the transport of heavy payload on interplanetary trajectories [1]. The principal question within the development of this thruster is whether it is possible to specifically heat the outer edges of an arcjet plume so that higher exhaust velocity can be attained. A dependency of the power staging between the two thruster stages and of the gas mass flow rate staging is expected. For the presented research, the thruster was investigated with a cavity calorimeter which provides data of plasma power so that a total efficiency can be derived. Moreover, a method for deducing the two-stage system’s efficiency from operational data is presented and its data are compared with the measured calorimetric data.

Radiated Emissions and Magnetic Field Characterization of a 2-kW Electrothermal Propulsion System

Abstract: Earth-orbiting spacecraft that include increasingly sensitive telecommunications, Earth-observing/sensing, and/or scientific instrument payloads often utilize electric propulsion systems. Such systems can generate high levels of radiated and conducted electromagnetic interference (EMI) making the accurate characterization of propulsion system EMI critical to spacecraft design. Radiated emissions and magnetic field measurements were performed on a 2-kW hydrazine arcjet electrothermal propulsion engine at the Large Dielectric Pulsed Propulsion facility at the Princeton University Electric Propulsion and Plasma Dynamics Laboratory. Plume/thruster emissions were characterized during steady state and transient startup operating conditions per MIL-STD-461E method RE102 over the 200-MHz to 18-GHz frequency range while magnetic field measurements were performed per MIL-STD-461E method RE101 from dc to 100 kHz. Results demonstrated a high level of compatibility between the arcjet electrothermal propulsion system with even the most sensitive radio- frequency payloads or science instruments to levels of -8 dBldrmuV/m and below.

Generation of Supersonic and Super-Alfvénic Flow Using ICRF Heating and a Magnetic Nozzle

Abstract: Fast-flowing plasmas in supersonic and super-Alfvenic regime are generated in combined experiments of ion cyclotron resonance heating (ICRH) and acceleration in a magnetic nozzle. During radio-frequency (RF) wave excitation in a fast-flowing plasma produced by a magnet-plasma-dynamic arcjet (MPDA), strong ion cyclotron heating is clearly observed. Thermal energy in the heated plasma is converted into flow energy in a diverging magnetic nozzle, where the magnetic moment μ is nearly kept constant. Plasma flow energy can be controlled by changing the input RF power and/or modifying the magnetic nozzle configuration. In a strongly diverging magnetic nozzle, an Alfven Mach number as well as ion acoustic Mach number are more than unity, that is, supersonic and super-Alfvenic plasma flow is realized.

Operational testing of the power conditioning unit for a 30 kWe arcjet

Abstract: Two arcjet Power Conditioning Units (PCU’s) were recently tested wtih off‐site arcjet thrusters. The arcjet PCU’s used in these in these tests were designed and built by SPI. The design of PCU I.1 was based on our previous work in arcjet power conditioning funded by AFAL. SDIO‐SBIR funded the development of PCU II, modification of PCU I.1 and PCU I.2, as well as the testing of these PCUs. The first of these recent tests was conducted at NASA Lewis Research Center with a hydrogen arcjet. The second one was conducted in Jet Propulsion Laboratory with an ammonia arcjet. The major thrust of the tests was on thruster performance, so the technical objectives and the interests of these tests were not limited to the PCU performance. SPI’s major objectives of these tests were to demonstrate the stable operation of the PCU’s with arcjet thrusters, to prove the capability of initiating the arc breakdown with its built‐in starter, and to demonstrate the endurance of the PCU. The personnel at NASA LeRC and JPL successfully operated the SPI PCUs with their arcjet and obtained valuable test data.

Numerical Study of Low-Power MPD Arcjet

Abstract: In spite of many experimental studies of low-power applied-field magnetoplasmadynamic (AFMPD) thrusters, thrust efficiencies of the past thrusters are very low. Hence, drastic improvement in thrust performance is required for AF-MPD thrusters to compete against other types of electric propulsion in a moderate power regime around 10 kW. For the optimization of AF-MPD thrusters, a numerical code for the flowfield simulation is now under development. A preliminary result shows that the code can deal with a complicated mixture of the induced and applied magnetic fields, which will lead to a combination of the self-field, swirl, Hall, as well as electrothermal accelerations.

Advanced materials for thermal protection system

Abstract: Reticulated open‐cell ceramic foams (both vitreous carbon and silicon carbide) and ceramic composites (SiC‐based, both monolithic and fiber‐reinforced) were evaluated as candidate materials for use in a heat shield sandwich panel design as an advanced thermal protection system (TPS) for unmanned single‐use hypersonic reentry vehicles. These materials were fabricated by chemical vapor deposition/infiltration (CVD/CVI) and evaluated extensively for their mechanical, thermal, and erosion/ablation performance. In the TPS, the ceramic foams were used as a structural core providing thermal insulation and mechanical load distribution, while the ceramic composites were used as facesheets providing resistance to aerodynamic, shear, and erosive forces. Tensile, compressive, and shear strength, elastic and shear modulus, fracture toughness, Poisson’s ratio, and thermal conductivity were measured for the ceramic foams, while arcjet testing was conducted on the ceramic composites at heat flux levels up to 5.90 MW/m2 (520 Btu/ft2⋅sec). Two prototype test articles were fabricated and subjected to arcjet testing at heat flux levels of 1.70–3.40 MW/m2 (150–300 Btu/ft2⋅sec) under simulated reentry trajectories.

Numerical Modeling of Hydrazine-Fueled Arcjet Thruster

Abstract: The computational fluid dynamic analysis has been conducted for the thermo-chemical flow field in an arcjet thruster with mono-propellant Hydrazine (N2H4) as a working fluid. The Reynolds Averaged Navier-Stokes (RANS) equations are modified to analyze compressible flows with the thermal radiation and electric field. The Maxwell equation, which is loosely coupled with the fluid dynamic equations through the Ohm heating and Lorentz forces, is adopted to analyze the electric field induced by the electric arc. The chemical reactions of Hydrazine were assumed to be infinitely fast due to the high temperature field inside the arcjet thruster. The chemical and the thermal radiation models for the nitrogen-hydrogen mixture and optically thick media respectively, were incorporated with the fluid dynamic equations. The results show that performance indices of the arcjet thruster with 1kW arc heating are improved by amount of 180% in thrust and 200% in specific impulse more than frozen flow. In addition to thermo-physical process inside the arcjet thruster is understood from the flow field results.

Effect of anode structure on the performance of arcjet

Abstract: The instability of arcjet under low power and flow rate leads to the decrease of performance.Using nitrogen as propellant,experiments under different conditions are compared between the new and old anode structure.For the old arcjet,thrust and impulse are low at low power,thermal efficiency is reduced at low mass flux.While for the new arcjet,it is clear that arc voltage,thrust,specific impulse,efficiency and stability are all improved remarkably.So the improvement of anode structure can control the place of the arc anode attachment,which has the effect on the low power arcjet.

Scale-model Experiment of Magnetoplasma Sail for Future Deep Space Missions

Abstract: When Magnetic sail (MagSail) spacecraft is operated in space, the supersonic solar wind plasma flow is blocked by an artificially produced magnetic cavity to accelerate the spacecraft in the direction leaving the Sun. To evaluate the momentum transferring process from the solar wind to the coil onboard the MagSail spacecraft, we arranged a laboratory experiment of MagSail spacecraft. Based on scaling considerations, a solenoidal coil was immersed into the plasma flow from a magnetoplasmadynamic arcjet in a quasi‐steady mode of about 1 ms duration. In this setup, it is confirmed that a magnetic cavity, which is similar to that of the geomagnetic field, was formed around the coil to produce thrust in the ion Larmor scale interaction. Also, the controllability of magnetic cavity size by a plasma jet from inside the coil of MagSail is demonstrated, although the thrust characteristic of the MagSail with plasma jet, which is so called plasma sail, is to be clarified in our next step.

Low power H_2-N_2 arcjet thruster operation and performance

Abstract: Gas mixtures of hydrogen and nitrogen with the mole ratio(r) between 1.0~2.0 were used as the propellant,and operation parameters,such as arc current,arc voltage,inlet pressure,chamber pressure,gas flow rates and nozzle temperature,were measured in-situ.The dependence of the thrust,specific impulse and efficiency on the specific power,r and total mass flow rate of the propellant were analyzed.Results showed that the thrust increases appreciably at high mass flow rate together with high r.However,for the improvement of specific impulse and efficiency,high r plays a more important role.Especially,the specific impulse increases significantly with the increasing r.

The Electric Propulsion Space Experiment (ESEX)—A demonstration of high power arcjets for orbit transfer applications

Abstract: The Electric Propulsion Space Experiment (ESEX) is a high power (30 kW) ammonia arcjet space demonstration sponsored by the Propulsion Directorate of the Phillips Laboratory with TRW as the prime contractor. ESEX is one of nine experiments being launched in early 1998 on board the Advanced Research and Global Observation Satellite (ARGOS). ESEX will demonstrate the feasibility of using a high power arcjet for orbit transfer. ESEX is instrumented with various sensors to address all of the expected interactions with ARGOS including electromagnetic interference, contamination, and radiated thermal loading. The performance of the arcjet will also be measured using ground tracking, an on-board GPS receiver, and on-board accelerometer. In addition to the performance and spacecraft interaction studies, ground-based spectroscopic and radiometric measurements will be performed to observe plume species as well as determine the effect of the arcjet firing on the space environment. ESEX is currently undergoing integrated testing with the spacecraft bus and the eight other experiments to verify the full operability of ARGOS while on-orbit. These tests include basic functionality of the system in addition to the normal suite of environmental tests including electromagnetic interference and compatibility, acoustic and pyroshock testing, and thermal vacuum tests.

High-Power and High-Thrust-Density Electric Propulsion for In-Space Transportation

Abstract: M scenarios such as building a scientifi c outpost on the moon or human exploration of Mars require new propulsion systems for in-space transportation of heavy payloads [1]. A thrust level of at least 100 N and a specifi c impulse level of 30 km/s are of central importance to increase the payload capacity and shorten the trip time as much as possible. The propulsion system mass should be as lightweight as possible. Therefore, both thrust density and effi ciency of the propulsion device should be high. Nuclear and solar thermal propulsion offer very high thrust densities but are still weak in their specifi c enthalpy level and thus specifi c impulse (see Chapter 1), which is clearly below 10 km/s (see Table 1). Ion and Hall-ion thrusters offer both the required exit velocity level but their thrust density is still relatively low. Promising in-space propulsion candidates for heavy payloads are currently thermal arcjet thrusters with which an exit velocity of 20 km/s at 100 kW and a thrust density of more than 2100 N/m2 can already be achieved. This technology is in an advanced developmental stage; low power devices have been implemented in commercial applications with hydrazine as propellant and are exceptionally reliable. The highest thrusts and thrust densities reached to date have been achieved with MPD self-fi eld thrusters. However, the effective exit velocity is still limited to 15 km/s. Although the achievable thrust density is an order of magnitude lower, applied-fi eld MPD thrusters should still be considered for this