Showing papers on "Arcjet rocket published in 1999"

Ablation and Thermal Response Program for Spacecraft Heatshield Analysis

Abstract: An implicit ablation and thermal response program is presented for simulation of one-dimensional transient thermal energy transport in a multilayer stack of isotropic materials and structure which can ablate from a front surface and decompose in-depth. The governing equations and numerical procedures for solution are summarized. Solutions are compared with those of an existing code, the Aerotherm Charring Material Thermal Response and Ablation Program, and also with arcjet data Numerical experiments show that the new code is numerically more stable and solves a much wider range of problems compared with the older code. To demonstrate its capability, applications for thermal analysis and sizing of aeroshell heatshields for planetary missions, such as Stardust, Mars Microprobe (Deep Space n), Saturn Entry Probe, and Mars 2001, using advanced light-weight ceramic ablators developed at NASA Ames Research Center, are presented and discussed.

Arcjet flow properties determined from laser-induced fluorescence of atomic nitrogen

Abstract: Flow property measurements that were recently acquired in the Ames Research Center Aerodynamic Heating Facility arcjet using two-photon laser-induced fluorescence (LIF) of atomic nitrogen (N) are reported. The flow properties, which include velocity, translational temperature, and N concentration, were measured simultaneously over a range of facility operating conditions for N2–argon test gas flows in the 30-cm-diameter nozzle. A recent measurement of the two-photon excitation cross section for the 3p4D° ← 2p4S° transition of atomic nitrogen is used to convert the relative nitrogen concentration measurements to absolute values, and a nitrogen flow reactor is used to provide a room-temperature, reference-wavelength calibration of the translational temperature and velocity measurements. When combined with information from facility measurements, an analysis of the flow properties obtained using two-photon LIF of N yields the total free-stream flow enthalpy.

Preliminary Evaluation of a 10 kW Hall Thruster

Abstract: A 10 kW Hall thruster was characterized over a range of discharge voltages from 300-500 V and a range of discharge currents from 15-23 A. This corresponds to power levels from a low of 4.6 kW to a high of 10.7 kW. Over this range of discharge powers, thrust varied from 278 mN to 524 mN, specific impulse ranged from 1644 to 2392 seconds, and efficiency peaked at approximately 59%. A continuous 40 hour test was also undertaken in an attempt to gain insight with regard to long term operation of the engine. For this portion of the testing the thruster was operated at a discharge voltage of 500 V and a discharge current of 20 A. Steady-state temperatures were achieved after 3-5 hrs and very little variation in performance was detected.

Diagnostics of an argon arcjet plume with a diode laser

Abstract: The diode-laser absorption technique was applied for simultaneous velocity and temperature measurements of an argon plume exhausted by an arcjet. The Ar i absorption line at 811.531 nm was taken as the center absorption line. The velocity and the temperature were derived from the Doppler shift in the absorption profiles and the full width at half-maximum of the plume absorption profile, respectively. From the measured plume velocity and temperature, the total enthalpy of the exhausted plume, the thrust efficiency, and the thermal efficiency of the arcjet were derived, and the performance of the arcjet was examined. The results are demonstrated to agree with results derived by other methods, and the technique can be applied to the measurement of other arcjet systems without much modification.

Spectroscopic Measurements of Shock-Layer Flows in an Arcjet Facility

Abstract: Under two different test conditions, radiation emanating from the freestream and shock-layer flow over a 15.24-cm-diam blunt-body test article as measured in NASA Ames Research Center's 20 MW Aerodynamic Heating Facility Arcjet. The test gas was a mixture of air and argon. Spatially resolved emission spectra were obtained over a 2000-8900 A wavelength range using a charge-coupled device camera (1024 × 256 array) attached to a spectrograph. The optical system was calibrated using tungsten and deuterium radiation sources. Previously developed analytical tools were used to determine the following line-of-sight-averaged thermodynamic properties from the calibrated spectra: 1) rotational temperature of the freestream and 2) rotational and vibrational temperatures within the shock layer. Based on the variation in intensity of emission spectra along the stagnation streamline, shock standoff distance was determined. Two sets of data for each test condition were compared to evaluate the repeatability of measurements. Considering likely sources of errors, an uncertainty analysis was performed to estimate the error bounds of the determined properties

Performance Computation of a Low-Power Hydrogen Arcjet

Abstract: A numerical analysis of a low-power hydrogen arcjet has been conducted by taking account of the chemical and thermal nonequilibria by separating the electron temperature from the heavy species temperature. A sheath model is introduced to the electrode boundaries to evaluate the electrode potential drops in a coupled manner with the flow calculation. Comparison of calculated performance for the IRS ARTUS-4 thruster with experimental results shows good agreement regarding the discharge voltage and specific impulses, although the heat loss is somewhat underestimated. To account for these discrepancies, the radiative loss is computed in detail, taking account of the bound-bound, free-bound, and free-free electronic transitions in the hydrogen plasma. The estimated radiation is found to become significant as the electron temperature exceeds approximately 12,000 K. Contribution of the radiative heat transfer to the total heat loss is discussed briefly. The results suggest that the radiation analysis may be necessary for further accurate prediction of thruster performance, even under low-power operation.

Low-power microwave arcjet testing: Plasma and plume diagnostics and performance evaluation

Abstract: The results of an experimental study of a microwave thruster are presented. Performance evaluation of a 7.5 GHz engine was done using different propellants under both atmospheric and vacuum conditions. Helium, nitrogen and ammonia were tested providing mean chamber stagnation temperature values. Other experiments include plasma and plume diagnostics. Emission spectroscopy of the plasma was made in order to measure the plasma electron temperature at different specific power levels, and the commonly-made assumption of Local Thermodynamic Equilibrium (LTE) was examined. In order to obtain thrust and specific impulse data under vacuum conditions, the Doppler shift between light emitted by the exhaust plume parallel to the gas velocity and perpendicular to the flow was measured using a high spectral resolution Fabry-Perot interferometer, yielding centerline specific impulse values for helium propellant at various specific powers. Thrust and mean specific impulses for all three propellants are being measured using a vertical mechanical thrust stand mounted inside a vacuum tank. Nomenclature

Low-Power Helium Pulsed Arcjet

Abstract: An electrothermal thruster that operates in a rapid-pulse mode at low power ( 85%of the initial stored energy is transferred to the arc in a unipolar pulse. The arc discharges occur in a cylindrical capillary upstream of a converging–diverging nozzle, and all of the energy additionoccurs in the subsonic region. Tests with heliumpropellant are conducted for two 20-deghalf-angle conicalnozzleswith area ratiosof 20 and230.Thrust levels from14 to 31mNaremeasured,andamaximumspeciŽ c impulse of 313 s is achieved with 36% efŽ ciency at 119 W. A time-dependent, quasi-one-dimensional numerical model is developed to evaluate energy losses using a time-marching procedure, and comparisonwith experimental results is good. Heat transfer losses to the wall in the subsonic region are found to be the primary energy-loss mechanisms. SpeciŽ c impulse is strongly affected by wall temperature. Viscous effects become important as the speciŽ c energy increases above 12 MJ/kg and the throat Reynolds number falls below 1 £ 103.

Water rocket - Electrolysis propulsion and fuel cell power

Abstract: Water Rocket is the collective name for an integrated set of technologies that offer new options for spacecraft propulsion, power, energy storage, and structure. Low pressure water stored on the spacecraft is electrolyzed to generate, separate, and pressurize gaseous hydrogen and oxygen. These gases, stored in lightweight pressure tanks, can be burned to generate thrust or recombined to produce electric power. As a rocket propulsion system, Water Rocket provides the highest feasible chemical specific impulse (-400 seconds). Even higher specific impulse propulsion can be achieved by combining Water Rocket with other advanced propulsion technologies, such as arcjet or electric thrusters. With innovative pressure tank technology, Water Rocket's specific energy [Wh/kg] can exceed that of the best foreseeable batteries by an order of magnitude, and the tanks can often serve as vehicle structural elements. For pulsed power applications, Water Rocket propellants can be used to drive very high power density generators, such as MHD devices or detonation-driven pulse generators. A space vehicle using Water Rocket propulsion can be totally inert and non-hazardous during assembly and launch. These features are particularly important for the timely development and flight qualification of new classes of spacecraft, such as microsats, nanosats, and refuelable spacecraft.

Comparisons between numerical model and experiments for a direct current plasma flow

Abstract: The governing equations describing a flowing stream of a hydrogen plasma encountered in applications, such as diamond deposition, and in devices, such as arcjet thrusters, are solved numerically using the linearized implicit (LBI) Method of Briley and McDonald. The results of simulations under the assumption that the plasma can be described by a single temperature are compared with detailed experimental measurements of flow characteristics and species concentrations in a 1 kW arcjet. These comparisons show that by formulating the problem in terms of known experimental operating conditions, such as mass flow rate, power, and current levels, it is possible to predict many of the characteristics of the flowing plasma. As expected, predictions from this one-temperature model show that some deviations from the experimental results occur near the exit plane of the channel, where unequal electron and heavy particle temperatures are encountered because of lower pressures.

Preliminary Analysis of Contamination Measurements from the ESEX 26 kW Ammonia Arcjet Flight Experiment

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 (and 33 min. 26 sec operating time), the following preliminary results are reported. The microbalances show no measurable deposition from the arcjet, in agreement with predictions. The radiometer near the thruster, viewing the arcjet plume and body, experiences a change in the thermal properties of its coating from the firings. Radiometers with no view of the arcjet, or a view of only the plume, show no change. During firings, the solar cell segments, near the thruster, show decreasing open-circuit voltage; probably attributable to an additional electrical load provided by the plume plasma. The solar cells also exhibit a 3% decrease in non-firing, solar-illuminated over the eight arcjet firings, attributable to decreased solar transmission of the cover glass. However, no effects associated with the arcjet are observed on the spacecraft solar arrays. These data are in good agreement with model predictions, where available. 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.

Preliminary Orbital Performance Analysis 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 from on-board systems, GPS, and radar ranging taken during these firings are used in this paper to determine thruster performance. The on-board Servo Accelerometer Assembly (SAA) measured spacecraft acceleration continually at 10 Hz. Although the design prohibited precise acceleration measurement in the range nominally expected during the firings, estimates of acceleration were obtained. The uncertainties of the acceleration estimates are on the order of 5% of nominal due primarily to the discretization error of the A/D converter. Mean performance figures are calculated based on acceleration and other on-board measurements. The final estimates of specific impulse and thrust efficiency are 787.0 +/- 49.8 seconds and 0.284 +/- 0.029, respectively.

Contamination Measurements from the ESEX 26 kW Ammonia Arcjet Flight Experiment

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 (and 33 min. 26 sec operating time), the following preliminary results are reported. The microbalances show no measurable deposition from the arcjet, in agreement with predictions. The radiometer near the thruster, viewing the arcjet plume and body, experiences a change in the thermal properties of its coating from the firings. Radiometers with no view of the arcjet, or a view of only the plume, show no change. During firings, the solar cell segments, near the thruster, show decreasing open-circuit voltage; probably attributable to an additional electrical load provided by the plume plasma. The solar cells also exhibit a 3% decrease in non-firing, solar-illuminated short-circuit current over the eight arcjet firings, attributable to decreased solar transmission of the cover glass. However, no effects associated with the arcjet are observed on the spacecraft solar arrays. These data are in good agreement with model predictions, where available. 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.

Measured exit plane properties of a low-power simulated-hydrazine arcjet

Abstract: Introduction I N two previouspaperson the 1-kWhydrazinearcjet, two types of electrostatic probe diagnostics were used at the nozzle exit plane to measure electron density ne , electron temperature Te, ion Mach number, plasma velocity u i , gas temperature Tg , and related quantities. All measurements were performed at 10 A and 50 mg/s propellantmass  ow on a NASA Lewis Research Center design,3 at a speciŽ c power of P= P mD 22:4 MJ/kg. These measurements were then compared to the two-temperaturechemicalnonequilibriumnumerical arcjet model of Megli et al. The quantitiesne , Te, and ion Mach numberwere determinedwith a novel quadrupleelectrostatic probe swept across the exit plane,with the probe axis parallel to the  ow velocity.1;6 The plasma velocity was measured with a double electrostatic probe using a technique similar to current modulation velocimetry (CMV).2;6 Both radial and axial variations of plasma propertieswere measured and compared to the numerical results. This Note presents further results for P= P mD 19:8 MJ/kg and P= P mD 26:0 MJ/kg (Ref. 6). The arcjet operating conditions are given in Table 1. This range of P= P m is representative of the stable operating range of this arcjet on simulated hydrazine (1 mole N2, 2 moles H2). The speciŽ c impulse is interpolatedfromdata acquired elsewhere on a comparable thruster. The quadruple probe combines the triple probe method of Chen and Sekiguchi, Ž rst applied to electric thruster plumes by Tilley et al.,8;9 with the crossed probe theory of Kanal,10 Ž rst applied to a single-speciesplasma by Johnson andMurphree and to a two-species plasma by Burton and Bufton.1 The quadrupleprobe was Ž rst appled to an electric thruster plume by DelMedico and Burton et al.13 The plasma velocity, assumed to equal the ion velocity u i , is determined by a modiŽ cation of a technique Ž rst developed by Pobst et al., in which the arc current is brie y interruptedand the resulting deŽ cit in electron density is convected at the plasma velocity and either detected optically or by a double time-of- ight (TOF) electrostaticprobe. Radial velocity proŽ les generatedby this technique are accurate to §500 m/s and provide excellent agreement with those generated by numericalmodeling.

On-Orbit Optical Observations of the ESEX 26kW Ammonia Arcjet

Abstract: : During the course of eight flight firings of the ESEX 26 kW arcjet in March and April, 1999, optical observations from on-board and ground-based sensors were obtained. Images from the on-board still camera indicate that the nozzle temperature distribution is consistent with arcjet heating models and ground observations. Images of the thruster plume at 656 nm confirm predictions that the luminescent plume in the space environment is more diffuse and compact than the plume from a thruster operated in the laboratory at a higher background pressure. Finally, observations using a ground-based telescope reveal a mixed greybody/line emission spectrum over the range 325-675 nm. The spectral features and line ratios are similar to those observed in ground-based measurements.

Rf plasma thrusters for smallsat applications

Abstract: The following is the first of a series of tests, dedicated to determining the feasibility of using the RF glow discharge phenomenon for a low power propulsion system. The main goals of this preliminary study are to explore electrode wear and characteristic operating parameters such as frequency, RF power, mass flow rate and gap distance. The RF thruster concept is characterized by a low-power RF glow discharge established between co-axial electrodes, analogous in design to an arcjet thruster. RF power sustains the plasma and ‘suspends’ charged particles within the electrode gap. The time duration of a half cycle does not allow the statistical majority of ions and electrons in the electrode gap to transverse the gap distance before the polarity reverses and causes the charged particles to be attracted in the opposite direction. The glow discharge is powered by an RF power supply encompassing a frequency range between 10 and 175 MHz. The power delivered to the glow discharge is in the range of 20 to 110 W and depends on the particular settings of the fundamental operating parameters. This paper explored the issue of electrode damage that may be caused by a glow discharge and compared it to DC glow discharge with equivalent operating conditions. In addition, it discusses the approach to a sensitivity study exploring the relations between the fundamental operating parameters (frequency, power, mass flow rate, and gap distance) and temperature.

The effect of entrance conditions on an expanding arcjet flow

Abstract: Flow development with various entrance conditions in a conical arcjet nozzle is simulated for a test gas consisting of nitrogen and argon. Simulations with uniform nozzle entrance conditions are performed and nozzle exit centerline conditions are compared with available experimental data. Simulations with entrance conditions consisting of a hot core and cooler annulus are then performed, and it is found that the non-uniformity persists through the nozzle, affecting the exit centerline temperature and velocity. Pitot pressure and heat transfer profiles determined horn simulated exit conditions are used to evaluate the effectiveness of pitot pressure and calorimeter measurements in determining the extent of flow uniformity. It is found that heat transfer measurements are far superior to pitot pressure measurements in assessing flow uniformity.

In-Flight Contamination Measurements of the ESEX 26 kW Ammonia Arcjet

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 on the spacecraft surfaces and environment. The contamination sensors have also been used to assess the spacecraft outgassing and to provide information during in-flight anomalies. The sensors consisted of microbalances to measure material deposition, radiometers to assess material degradation due to thermal irradiation, and solar cell segments to quantify the potential for solar array degradation. No material deposition is observed during six arcjet firings, with the exception of the first firing. Deposition during the first firing is observed only on the microbalance near the arcjet exit, and is attributed to the expulsion of foreign material within the arcjet that presumably accumulated during the fabrication, integration, and storage of the device prior to launch. Solar cell degradation is observed during the firings and is attributed to the exhaust plasma forming slightly conductive paths away from the cell load. Once the firing ends, the solar cell returns to full operation with no observable degradation. Initial evaluation of the radiometer data suggests some material degradation of the 513-GLO coating, however a more detailed transient analysis is required for a conclusive result. Flight data is continuing to be collected on all contamination sensors to further quantify changes to the spacecraft atmosphere due to operation of the other 8 experiments aboard the ARGOS spacecraft.

Laser-induced fluorescence characteristics of a helium arcjet flow

Abstract: Laser-induced fluorescence is used to characterize the axial velocity and temperature field at the exit plane of a low-power helium arcjet. Two cases were examined, one in which the mass-flow rate was changed at a constant current, and the other where the current was changed at a constant mass-flow rate. At constant mass-flow rate, the velocity scales with the increase in power. At constant current, a higher specific energy results in an unexpected lower mean exit velocity. The temperature profiles show that the nozzle-wall temperature is greater than the mean exit temperature. Along the axis of the arcjet, these measurements indicate the presence of a shock less than one nozzle diameter downstream of the exit.

A Numerical and Experimental Study of Arcjet Fluctuation

Abstract: : The work described here is a combined numerical/experimental study of arcjets. The focus of the experimental work has been a parametric study of the onset of instability in arc thruster devices. The focus of the numerical work has been the development of numerical methods for solving the magnetohydrodynamic (MHD) equations in complex geometries.

Electromagnetic Effects of the ESEX 26 kW Ammonia Arcjet on Normal Spacecraft Communications and Operations

Abstract: : Arcjet thrusters employ an arc discharge to heat propellant which expands through a nozzle to produce thrust. Spacecraft designers who desire to exploit arcjet technology have expressed concern about interference from the electromagnetic environment produced by the thrusters. One of the four major interest areas of the ESEX program was to determine the electromagnetic effect of operating a 30 kW class arcjet upon normal spacecraft communications and operations. To accomplish this task, noise levels were recorded in four frequency bands (2, 4, 8, and 12 GHz) by two onboard antennas. The electromagnetic noise levels observed during all arcjet firings were indistinguishable from those observed during the baseline non-firing periods. Communication bit error rates were also measured during arcjet firings and comparison with non-firing BER data revealed that spacecraft command and control communications were not significantly affected by arcjet firings.