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Active Oldest Votes. Improve this answer. That comes as a bit of a surprise. Roughly speaking, how much pressure are we talking about here? Assuming you're not using reactive or radioactive Stronger containers generally weigh more, and weight decreases efficiency. By using a gas with a high molar mass, you increase the mass required for the vessel , and that's a bad thing.
Show 2 more comments. Russell Borogove Russell Borogove k 9 9 gold badges silver badges bronze badges. The drawback with fuels that are not gases at very low temperatures is that: They have to be vaporized, which requires extra energy and engineering.
If they are not gaseous at low temperatures then there is unvalidated concern that the ions could condense back onto the spacecraft. Lysander Lysander 6 6 bronze badges.
Sign up or log in Sign up using Google. Sign up using Facebook. Sign up using Email and Password. Presently the test article for the NEXT LDT has been exposed to atmosphere and placed within a clean room environment, with post-test disassembly and inspection underway. Experimental investigations of argon and xenon ion sources. The multipole thruster was used to investigate the use of argon and xenon propellants as possible alternatives to the electric thruster propellants of mercury and cesium.
The multipole approach was used because of its general high performance level. The design employed, using flat and cylindrical rolled sections of sheet metal, was selected for ease of fabrication, design, assembly, and modification. All testing was conducted in a vacuum facility and the pumping was accomplished by a 0. Minimum discharge losses were in the ev. Flatness parameters were typically in the 0. Plasma particle simulation of electrostatic ion thrusters. Charge exchange collisons between beam ions and neutral propellant gas can result in erosion of the accelerator grid surfaces of an ion engine.
A particle in cell PIC is developed along with a Monte Carlo method to simulate the ion dynamics and charge exchange processes in the grid region of an ion thruster. The simulation is two-dimensional axisymmetric and uses three velocity components 2d3v to investigate the influence of charge exchange collisions on the ion sputtering of the accelerator grid surfaces. An example calculation has been performed for an ion thruster operated on xenon propellant. The simulation shows that the greatest sputtering occurs on the downstream surface of the grid, but some sputtering can also occur on the upstream surface as well as on the interior of the grid aperture.
Sputtering Erosion in the Ion Thruster. During the first phase of this research, the sputtering yields of molybdenum by low energy eV and higher xenon ions were measured by using the methods of secondary neutral mass spectrometry SNMS and Rutherford backscattering spectrometry RBS.
However, the measured sputtering yields were found to be far too low to explain the sputtering erosions observed in the long-duration tests of ion thrusters. The only difference between the sputtering yield measurement experiments and the ion thruster tests was that the later are conducted at high ion fluences.
Hence, a study was initiated to investigate if any linkage exists between high ion fluence and an enhanced sputtering yield. The objective of this research is to gain an understanding of the causes of the discrepancies between the sputtering rates of molybdenum grids in an ion thruster and those measured from our experiments. We are developing a molecular dynamics simulation technique for studying low-energy xenon ion interactions with molybdenum.
It is difficult to determine collision sequences analytically for primary ions below the eV energy range where the ion energy is too low to be able to employ a random cascade model with confidence and it is too high to have to consider only single collision at or near the surface. At these low energies, the range of primary ions is about 1 to 2 nm from the surface and it takes less than 4 collisions on the average to get an ion to degrade to such an energy that it can no longer migrate.
The fine details of atomic motion during the sputtering process are revealed through computer simulation schemes. By using an appropriate interatomic potential, the positions and velocities of the incident ion together with a sufficient number of target atoms are determined in small time steps. Hence, it allows one to study the evolution of damages in the target and its effect on the sputtering yield.
We are at the preliminary stages of setting up the simulation program. A liquid nitrogen-cooled shroud was used to cold-soak the thruster to C. Initial tests were performed prior to a mature spacecraft design.
Those results and the final, severe, requirements mandated by the spacecraft led to several changes to the basic thermal design. These changes were incorporated into a final design and tested over a wide range of environmental conditions. A high power ion thruster for deep space missions. Polk, James E. This engine was designed to operate at power levels ranging from 13 to 28 kW at specific impulses of s and for burn times of up to 10 years.
State-of-the-art performance and life assessment tools were used to design the thruster , which featured cm-diameter carbon-carbon composite grids operating at voltages of 3. Preliminary validation of the thruster performance was accomplished with a laboratory model thruster , while in parallel, a flight-like development model DM thruster was completed and two DM thrusters fabricated.
The first thruster completed full performance testing and a h wear test. The second successfully completed vibration tests at the full protoflight levels defined for this NEP program and then passed performance validation testing. The thruster design, performance, and the experimental validation of the design tools are discussed in this paper.
Hall thruster. Three operating conditions are considered with variations to the magnetic field strength and chamber background pressure in an effort to capture their effects on ion acceleration and centerline ion energy distributions.
Under nominal conditions, xenon ions are accelerated to an energy of 25 eV within the thruster with an additional eV gain in the thruster plume. At a position 40 mm into the plume,. The NASA s Evolutionary Xenon Thruster NEXT program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost.
As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test LDT was initiated to validate and qualify the NEXT propellant throughput capability to a qualification-level of kg, 1.
This wear test is being conducted with a modified, flight-representative NEXT engineering model ion thruster , designated EM3. As of June 21, , the thruster has accumulated 10, hr of operation at the thruster full-input-power of 6. The thruster has processed kg of xenon and demonstrated a total impulse of 8. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.
Overall ion thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. Lifetime-limiting component erosion rates have been consistent with the NEXT service life assessment, which predicts the earliest failure sometime after kg of xenon propellant throughput; well beyond the mission-derived lifetime requirement. The NEXT wear test data confirm that the erosion of the discharge keeper orifice, enlarging of nominal-current-density accelerator grid aperture cusps, and the decrease in cold grid-gap observed during the NSTAR Extended Life Test have been mitigated.
The test was voluntarily terminated in April after accumulating 51, hours of high voltage operation, processing kg of xenon , and delivering This presentation covers the post-test inspection of the thruster hardware, in particular of the discharge chamber and other miscellaneous components such as propellant isolators and electrical cabling. The PM1 thruster exhibits operational behavior consistent with its predecessors, the engineering model thrusters , with substantial mass savings, enhanced thermal margins, and design improvements for environmental testing compliance.
The dry mass of PM1 is Modifications made in the thruster design have resulted in improved performance and operating margins, as anticipated. PM1 beginning-of-life performance satisfies all of the electric propulsion thruster mission-derived technical requirements. It demonstrates a wide range of throttleability by processing input power levels from 0.
The flat beam profile, flatness parameters vary from 0. The thruster throughput capability is predicted to exceed kg of xenon , an equivalent of 36, hr of continuous operation at the full-power operating condition. Erosion rate diagnostics in ion thrusters using laser-induced fluorescence.
Gaeta, C. We have used laser-induced fluorescence LIF to monitor the charge-exchange ion erosion of the molybdenum accelerator electrode in ion thrusters. This real-time, nonintrusive method was implemented by operating a 30cm-diam ring-cusp thruster using xenon propellant.
With the thruster operating at a total power of 5 kW, laser radiation at a wavelength of nm corresponding to a ground state atomic transition of molybdenum was directed through the extracted ion beam adjacent to the downstream surface of the molybdenum accelerator electrode. Molybdenum atoms, sputtered from this surface as a result of charge-exchange ion erosion, were excited by the laser radiation. The intensity of the laser-induced fluorescence radiation, which is proportional to the sputter rate of the molybdenum atoms, was measured and correlated with variations in thruster operating conditions such as accelerator electrode voltage, accelerator electrode current, and test facility background pressure.
We also demonstrated that the LIF technique has sufficient sensitivity and spatial resolution to evaluate accelerator electrode lifetime in ground-based test facilities. Measurements of neutral and ion velocity distribution functions in a Hall thruster. Hall thruster is a plasma device for space propulsion. It utilizes a cross-field discharge to generate a partially ionized weakly collisional plasma with magnetized electrons and non-magnetized ions.
The ions are accelerated by the electric field to produce the thrust. There is a relatively large number of studies devoted to characterization of accelerated ions , including measurements of ion velocity distribution function using laser-induced fluorescence diagnostic. Interactions of these accelerated ions with neutral atoms in the thruster and the thruster plume is a subject of on-going studies, which require combined monitoring of ion and neutral velocity distributions.
Herein, laser-induced fluorescence technique has been employed to study neutral and single-charged ion velocity distribution functions in a W cylindrical Hall thruster operating with xenon propellant. An optical system is installed in the vacuum chamber enabling spatially resolved axial velocity measurements.
The fluorescence signals are well separated from the plasma background emission by modulating the laser beam and using lock-in detectors. Measured velocity distribution functions of neutral atoms and ions at different operating parameters of the thruster are reported and analyzed. Component testing to-date has consisted of long-duration high voltage propellant isolator and high-cycle heater life validation testing.
The high voltage propellant isolator, a heritage design, will be operated under different environmental condition in the NEXT ion thruster requiring verification testing.
The life test of two NEXT isolators was initiated with comparable voltage and pressure conditions with a higher temperature than measured for the NEXT prototype-model thruster. To date the NEXT isolators have accumulated 18, h of operation. Measurements indicate a negligible increase in leakage current over the testing duration to date. The heater fabrication processes, developed for the International Space Station ISS plasma contactor hollow cathode assembly, were utilized with modification of heater dimensions to accommodate a larger cathode.
To date two of the heaters have been cycled to 10, cycles and suspended to preserve hardware. Three of the heaters have been cycled to failure giving a B10 life of 12, cycles, approximately 6, more cycles than the established qualification B10 life of the ISS plasma contactor heaters.
Miniature ion thruster ring-cusp discharge performance and behavior. Miniature ion thrusters are an attractive option for a wide range of space missions due to their low power levels and high specific impulse. Thrusters using ring-cusp plasma discharges promise the highest performance, but are still limited by the challenges of efficiently maintaining a plasma discharge at such small scales typically cm diameter. This effort significantly advances the understanding of miniature-scale plasma discharges by comparing the performance and xenon plasma confinement behavior for 3-ring, 4-ring, and 5-ring cusp by using the 3 cm Miniature Xenon Ion thruster as a modifiable platform.
By measuring and comparing the plasma and electron energy distribution maps throughout the discharge, we find that miniature ring-cusp plasma behavior is dominated by the high magnetic fields from the cusps; this can lead to high loss rates of high-energy primary electrons to the anode walls.
Even though these design modifications still present some challenges, they show promise to bypassing what were previously seen as inherent limitations to ring-cusp discharge efficiency at miniature scales. The mass of the xenon stored onboard DS1 was about 81 kg, and the spacecraft wet mass was approximately kg. The DS1 spacecraft was launched on October 24, , and on July 29, , it flew within 16 miles of the small asteroid Braille formerly KD at a relative speed of 35, mph. As of November , the ion propulsion system had performed flawlessly for nearly days of thrusting.
NASA has approved an extension to the mission, which will allow DS1 to continue thrusting to encounters with two comets in The DS1 optical and plasma diagnostic instruments will be used to investigate the comet and space environments.
The spacecraft is scheduled to fly past the dormant comet Wilson- Harrington in January and the very active comet Borrelly in September , at which time. Hall thruster is investigated. The MFFP is designed to eliminate the collection of low-energy, charge-exchange CEX ions by using a variable magnetic field as an ion filter.
The Nuclear Electric Xenon Ion System NEXIS research and development activity within NASA's Project Prometheus, was one of three proposals selected by NASA to develop thruster technologies for long life, high power, high specific impulse nuclear electric propulsion systems that would enable more robust and ambitious science exploration missions to the outer solar system.
The NEXIS technology uses erosion resistant carbon-carbon grids, a graphite keeper, a new reservoir hollow cathode, a cm diameter chamber masked to produce a cm diameter ion beam, and a shared neutralizer architecture to achieve these goals. The accomplishments of the NEXIS activity so far include performance testing of a laboratory model thruster , successful completion of a proof of concept reservoir cathode hour wear test, structural and thermal analysis of a completed development model thruster design, fabrication of most of the development model piece parts, and the nearly complete vacuum facility modifications to allow long duration wear testing of high power ion thrusters.
A nominal Watt Hall Effect Thruster was developed to propel unmanned space vehicles. Both xenon and iodine compatible versions were demonstrated. Evolution of the thruster channel due to ion erosion was predicted through numerical models and calibrated with experimental measurements.
Estimated xenon throughput is greater than kg. The thruster is well sized for satellite station keeping and orbit maneuvering, either by itself or within a cluster. Analysis and design of ion thrusters for large space systems. This study undertakes the analysis and conceptual design of a 0. Either argon or xenon gas shall be used as propellant. A 50 cm diameter discharge chamber was selected to meet stipulated performance goals. The discharge plasma is contained at the boundary by a periodic structure of alternating permanent magnets generating a series of line cusps.
Anode strips between the magnets collect Maxwellian electrons generated by a central cathode. Ion extraction utilizes either two or three grid optics at the user's choice. An extensive analysis was undertaken to investigate optics behavior in the high power environment of this large thruster. A plasma bridge neutralizer operating on inert gas provides charge neutralizing electrons to complete the design. The resulting conceptual thruster and the necessary power management and control requirements are described.
Mercury ion thruster technology. The Mercury Ion Thruster Technology program was an investigation for improving the understanding of state-of-the-art mercury ion thrusters. Emphasis was placed on optimizing the performance and simplifying the design of the 30 cm diameter ring-cusp discharge chamber.
Comprehensive Langmuir-probe surveys were obtained and compared with similar measurements for a J-series thruster. A successful volume-averaging scheme was developed to correlate thruster performance with the dominant plasma processes that prevail in the two thruster designs.
The average Maxwellian electron temperature in the optimized ring-cusp design is as much as 1 eV higher than it is in the J-series thruster. Advances in ion -extraction electrode fabrication technology were made by improving materials selection criteria, hydroforming and stress-relieving tooling, and fabrications procedures.
An ion -extraction performance study was conducted to assess the effect of screen aperture size on ion -optics performance and to verify the effectiveness of a beam-vectoring model for three-grid ion optics. An assessment of the technology readiness of the J-series thruster was completed, and operation of an 8 cm IAPS thruster using a simplified power processor was demonstrated. Improved accelerator and screen grids for an ion accelerator have been designed and tested in a continuing effort to increase the sustainable power and thrust at the high end of the accelerator throttling range.
The improved accelerator and screen grids could also be incorporated into ion accelerators used in such industrial processes as ion implantation and ion milling. NEXT is a successor to the NASA Solar Electric Propulsion Technology Application Readiness NSTAR thruster - a state-of-the-art ion thruster characterized by, among other things, a beam-extraction diameter of 28 cm, a span-to-gap ratio defined as this diameter divided by the distance between the grids of about , and a rated peak input power of 2.
To enable the NEXT thruster to operate at the required higher peak power, the beam-extraction diameter was increased to 40 cm almost doubling the beam-extraction area over that of NSTAR see figure. The span-to-gap ratio was increased to to enable throttling to the low end of the required input-power range. The geometry of the apertures in the grids was selected on the basis of experience in the use of grids of similar geometry in the NSTAR thruster.
Characteristics of the aperture geometry include a high open-area fraction in the screen grid to reduce discharge losses and a low open-area fraction in the accelerator grid to reduce losses of electrically neutral gas atoms or molecules.
The NEXT accelerator grid was made thicker than that of the NSTAR to make more material available for erosion, thereby increasing the service life and, hence, the total impulse. The NEXT grids are made of molybdenum, which was chosen because its combination of high strength and low thermal expansion helps to minimize thermally and inertially induced.
Ion thruster project. The mercury bombardment electrostatic ion thruster is the most successful electric thruster available today.
A 5 cm diameter ion thruster with 3, specific impulse and 5mN thrust is described. The advantages of electric propulsion and the tests that will be performed are also presented. Sputtering erosion in ion and plasma thrusters. An experimental set-up to measure low-energy below 1 keV sputtering of materials is described.
The materials to be bombarded represent ion thruster components as well as insulators used in the stationary plasma thruster. The sputtering takes place in a 9 inch diameter spherical vacuum chamber. Ions of argon, krypton and xenon are used to bombard the target materials. The sputtered neutral atoms are detected by a secondary neutral mass spectrometer SNMS.
Samples of copper, nickel, aluminum, silver and molybdenum are being sputtered initially to calibrate the spectrometer. The base pressure of the chamber is approximately 2 x 10 exp -9 Torr. The ion beam can be focused to a size approximately 1 mm in diameter. The mass spectrometer is positioned 10 mm from the target and at 90 deg angle to the primary ion beam direction. The ion beam impinges on the target at 45 deg.
For sputtering of insulators, charge neutralization is performed by flooding the sample with electrons generated from an electron gun. Preliminary sputtering results, methods of calculating the instrument response function of the spectrometer and the relative sensitivity factors of the sputtered elements will be discussed.
Ion properties in a Hall current thruster operating at high voltage. In this model, anomalous electron transport is fitted from ion mean velocity measurements, and energy losses due to electron—wall interactions are used as a tuned parameter to match expected electron temperature strength for same class of thruster. Doubly charged ions production has been taken into account and detailed collisions between heavy species included. Results show that the region of ion production of each species is located at the same place inside the thruster channel.
Because collision processes mean free path is larger than the acceleration region, each type of ions experiences same potential drop, and ion energy distributions of singly and doubly charged are very similar. Sub-component performance as well as overall thruster performance is presented and compared to results over the course of the test. Overall wear of critical thruster components is also described, and an update on the first failure mode of the thruster is provided.
Development of a miniature microwave electron cyclotron resonance plasma ion thruster for exospheric micro-propulsion.
The discharge source uses both radial and axial magnetostatic field confinement to facilitate electron cyclotron resonance and increase the electron dwell time in the volume, thereby enhancing plasma production efficiency.
Performance of the ion thruster is studied at 3 microwave frequencies 1. The discharge geometry is found to operate most efficiently at an input microwave frequency of 1.
Electric ion thrusters are the preferred engines for deep space missions, because of very high specific impulse. The ion engine consists of screen and accelerator grids containing thousands of concentric very small holes. The xenon gas accelerates between the two grids, thus developing the impulse force. The dominant life-limiting mechanism in the state-of-the-art molybdenum thrusters is the xenon ion sputter erosion of the accelerator grid.
Early effort to develop CCC composite thrusters had a limited success because of limitations of the drilling technology and the damage caused by drilling. The proposed is an in-situ manufacturing of holes while the CCC is made. First, a manufacture process for cm diameter thruster grids will be developed and verified. Quality of holes, density, CTE, tension, flexure, transverse fatigue and sputter yield properties will be measured.
After establishing the acceptable quality and properties, the process will be scaled to manufacture cm diameter grids. The properties of the two grid sizes are compared with each other. As part of a comprehensive thruster service life assessment, the NEXT Long-Duration Test LDT was initiated in June to demonstrate throughput capability and validate thruster service life modeling.
To date, the NEXT LDT has set records for electric propulsion lifetime and has demonstrated 50, h of operation, processed kg of propellant, and delivered Various component erosion rates compare favorably to the pretest predictions based upon semi-empirical ion thruster models. The NEXT LDT either met or exceeded all of its original goals regarding lifetime demonstration, performance and wear characterization, and modeling validation.
As part of this termination procedure, a comprehensive post-test performance characterization was conducted across all operating conditions of the NEXT throttle table. These measurements were found to be consistent with prior data that show minimal degradation of performance over the thruster 's 50 kh lifetime. Repair of various diagnostics within the test facility is presently planned while keeping the thruster under high vacuum conditions. These diagnostics will provide additional critical.
To date, the NEXT LDT has set records for electric propulsion lifetime and has demonstrated 50, hours of operation, processed kg of propellant, and delivered Multipole gas thruster design. The development of a low field strength multipole thruster operating on both argon and xenon is described. Experimental results were obtained with a cm diameter multipole thruster and are presented for a wide range of discharge-chamber configurations.
Ion beam flatness parameters in the plane of the accelerator grid ranged from 0. Thruster performance is correlated for a range of ion chamber sizes and operating conditions as well as propellant type and accelerator system open area.
A cm diameter ion source designed and built using the procedure and theory presented here-in is shown capable of low discharge losses and flat ion -beam profiles without optimization. This indicates that by using the low field strength multipole design, as well as general performance correlation information provided herein, it should be possible to rapidly translate initial performance specifications into easily fabricated, high performance prototypes.
Values have been obtained for the full level throttle table, as well as for a few off-nominal operating conditions. When measurements are compared to TT10 values that have been corrected using ion beam current density and charge state data obtained at The Aerospace Corporation, they differ by 1. Thrust correction factors calculated from direct thrust measurements and from The Aerospace Corporation s plume data agree to within measurement error for all but one TL.
Thrust due to cold flow and "discharge only" operation has been measured, and analytical expressions are presented which accurately predict thrust based on thermal thrust generation mechanisms. Several researchers have measured ions leaving ion thruster discharge chambers with energies far greater than measured discharge chamber potentials.
Presented in this paper is a new mechanism for the generation of high energy ions and a comparison with measured ion spectra. The source of high energy ions has been a puzzle because they not only have energies in excess of measured steady state potentials, but as reported by Goebel et. The mechanism relies on the charge exchange neutralization of xenon ions accelerated radially into the potential trough in front of the discharge cathode.
Previous researchers [2] have identified the importance of charge exchange in this region as a mechanism for protecting discharge cathode surfaces from ion bombardment. This paper is the first to identify how charge exchange in this region can lead to ion energy enhancement. Polk, Jay E. NASA is investigating high power, high specific impulse propulsion technologies that could enable ambitious flights such as multi-body rendezvous missions, outer planet orbiters and interstellar precursor missions.
The requirements for these missions are much more demanding than those for state-of-the-art solar-powered ion propulsion applications. The purpose of the NEXIS program is to develop advanced ion thruster technologies that satisfy the requirements for high power, high specific impulse operation, high efficiency and long thruster life.
These performance and throughput goals will be achieved by applying a combination of advanced technologies including a large discharge chamber, erosion resistant carbon-carbon grids, an advanced reservoir hollow cathode and techniques for increasing propellant efficiency such as grid masking and accelerator grid aperture diameter tailoring.
This paper provides an overview of the challenges associated with these requirements and how they are being addressed in the NEXIS program. Hamley, John A. Test objectives included: 1 demonstration and validation of automated thruster start procedures, 2 demonstration of stable closed loop control of the thruster beam current, 3 successful response and recovery to thruster faults, and 4 successful safing of the system during simulated spacecraft faults.
These objectives were met over the specified VDC input voltage range and 0. Two minor anomalies were noted in discharge and neutralizer keeper current. These anomalies did not affect the stability of the system and were successfully corrected. Ion thruster performance model. A model of ion thruster performance is developed for high flux density, cusped magnetic field thruster designs. This model is formulated in terms of the average energy required to produce an ion in the discharge chamber plasma and the fraction of these ions that are extracted to form the beam.
The direct loss of high energy primary electrons from the plasma to the anode is shown to have a major effect on thruster performance. The model provides simple algebraic equations enabling one to calculate the beam ion energy cost, the average discharge chamber plasma ion energy cost, the primary electron density, the primary-to-Maxwellian electron density ratio and the Maxwellian electron temperature.
Experiments indicate that the model correctly predicts the variation in plasma ion energy cost for changes in propellant gas Ar, Kr and Xe , grid transparency to neutral atoms, beam extraction area, discharge voltage, and discharge chamber wall temperature.
The model and experiments indicate that thruster performance may be described in terms of only four thruster configuration dependent parameters and two operating parameters. The model also suggests that improved performance should be exhibited by thruster designs which extract a large fraction of the ions produced in the discharge chamber, which have good primary electron and neutral atom containment and which operate at high propellant flow rates.
Hall Thruster in a high vacuum environment. The ionized propellant velocities were measured using laser induced fluorescence of the excited state xenon ionic transition at Ion velocities were interrogated from the channel exit plane to a distance 30 mm from it. Both axial and cross-field along the electron Hall current direction velocities were measured. The results presented here, combined with those of previous work, highlight the high sensitivity of electron mobility inside and outside the channel, depending on the background gas density, type of wall.
The thruster was operated at a constant xenon flow rate of 10 milligrams per second and discharge voltages of to V. The ExB probe was placed two meters downstream of the thruster exit plane on the thruster centerline.
Hall thruster on spacecraft, a number of plume properties have been measured. These include current density using a Faraday probe, ion energy distribution using a retarding potential analyzer, and ion species fractions using an E x B probe. Air Force Research Laboratory. Plume characterization of Hall thrusters is required to fully understand the impacts of thruster operation on spacecraft.
Much of these plume data are. Xenon ion propulsion systems are being developed by NASA Lewis Research Center and the Jet Propulsion Laboratory to provide flight qualification and validation for planetary and earth-orbital missions.
In the ground-test element of this program, light-weight less than 7 kg , 30 cm diameter ion thrusters have been fabricated, and preliminary design verification tests have been conducted. An engineering model thruster is now undergoing a h wear-test. A breadboard power processor is being developed to operate from an 80 V to V power bus with inverter switching frequencies of 50 kHz. The power processor design is a pathfinder and uses only three power supplies. Preliminary integration tests of the neutralizer power supply and the ion thruster have been completed.
Banks, Bruce A. The anticipated approx. However, the mass of the atom being accelerated and shot out as propellant makes a big difference, because momentum is mass times velocity.
Xenon is times more massive than hydrogen, so it imparts times more momentum to the spacecraft then hydrogen, even though they are accelerated by the same electrostatic force. To possibly get a little closer to the question, and I'll happily accept any corrections: Higher specific impulse reduces the amount of propellant needed to achieve a sought mission velocity and so reduces the propellant ratio, so the effort to increase this with conventional engines.
With ion engines the Isp, and so propellant ratio, is in comparison to other concerns, vastly reduced. The larger issues with ion engines and their much higher Isp are thrust, power density, and energy density correct terms? The larger mass of xenon actually increases the energy efficiency of the engine because it actually lowers the exhaust velocity and a rocket engine's energy efficiency increases as the exhaust velocity and forward velocity of the vehicle close on each other.
I think so. In the early stages of development, they were. Cesium vapor, having a very high mass to ionization energy ratio was abandoned due to its high reactivity, mercury vapor also had bad handling characteristics, until xenon was settled on.
Or at least that's what Rocket Propulsion Elements says.
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