Laser induced fluorescence measurements on a laboratory Hall thruster
14
Citation
17
Reference
10
Related Paper
Citation Trend
Abstract:
In this paper, we describe the results of a study of laser induced fluorescence velocimetry of neutral xenon in the plume of a Hall type thruster operating at powers ranging from 250 to 725 W. Neutral velocities are seen to increase with thruster discharge voltage. There is no evidence for neutrals being accelerated in the near field plume. Velocities appear to remain constant past the cathode plane. In preparation for future ion velocimetry studies, the plume plasma potential profile is measured for a number of conditions. For a low power condition, the plasma potential profile is mapped through the ionization region into the interior of the thruster. For this condition, the electric field profile is calculated. We also find evidence of neutral xenon streaming toward the Hall thruster. These backstreaming neutrals make determination of neutral xenon velocities difficult. We believe the neutrals originate from the thruster plume wall impingement approximately 2 m from the thruster.Keywords:
Langmuir Probe
Ion thruster
Energetic neutral atom
Experiments show electrostatic thrusters with components such as the discharge chamber or acceleration channel, solenoid or permanent magnets, hollow cathode, and keeper can be replaced by a simple, propellant‐selective, solid‐state, ion‐conducting membrane (Wilbur et al., 2007; Wilbur, Wilson, and Williams, 2005). In addition, analyzes show these membranes can be shaped, structured, and assembled into integrated thruster systems that will operate at much greater thrust densities and thruster efficiencies than those for state‐of‐the‐art, Hall and ion thrusters (Wilbur, Farnell, and Williams, 2005). The implications of these findings are revolutionary and promise an electrostatic propulsion system much less massive, more reliable, and less costly than ion and Hall thruster systems as they can be fabricated readily using traditional ceramic manufacturing techniques. The status of the Emissive Membrane Ion Thruster (EMIT) concept is described and recent measurements are used to estimate the performance of a propulsion system based on this concept. Estimates are also provided for the specific masses of various components required for it to perform typical satellite missions and comparisons are made to conventional electric propulsion systems currently in use. The emissive membrane thruster is shown to enable operation at 20% to 50% greater thrust‐to‐power ratios at specific impulses from 1000 s to 5000 s. Related performance advantages will also be discussed and analyses will be presented that show why an EMIT system is less expensive, more reliable, easily scalable, and simpler compared to existing electric thruster systems.
Ion thruster
Solenoid
Cite
Citations (1)
The development of electric propulsion systems is discussed and the benefits of these systems to various space mission requirements are outlined. The characteristics and development status of 8 and 30 cm mercury ion thrusters and solar electric propulsion systems are reported. In addition the advantages of an inert gas thruster for Earth orbital missions are examined and include its capability for operation at higher values of specific impulse, the ease at which it can be integrated with space systems, and it's low pollution potential.
Ion thruster
Specific impulse
In-space propulsion technologies
Laser Propulsion
Mercury
Cite
Citations (0)
Abstract A new approach for the development of an innovative, low-cost, long lifespan, small size and versatile metal ion thruster (MIT), able to independently control thrust and specific impulse generation, is proposed. The concept of pulsed thermionic vacuum arc (PTVA) is used to generate and accelerate metal ions, without using acceleration grids. Operating under high or ultra-high vacuum conditions makes PTVA discharge suitable to be used in the vacuum of space. The proposed electric propulsion system can provide significant thrust and specific impulse levels due to an efficient metal ion acceleration process in the electric field of a double layer structure developed in PTVA plasma. For certain experimental conditions, the performance parameters values of this MIT–PTVA approach, and even exceed, those of the classical ion thrusters like xenon ion propulsion system and stationary plasma Hall thruster.
Specific impulse
Ion thruster
Vacuum arc
Thermionic emission
Electric arc
Cite
Citations (3)
There is common agreement within the scientific community that in order to understand our local galactic environment it will be necessary to send a spacecraft into the region beyond the solar wind termination shock. Considering distances of 200 AU for a new mission, one needs a spacecraft travelling at a speed of close to 10 AU/yr in order to keep the mission duration in the range of less than 25 yrs, a transfer time postulated by ESA. Two propulsion options for the mission have been proposed and discussed so far: the solar sail propulsion and the ballistic/radioisotope electric propulsion. As a further alternative, we here investigate a combination of solar-electric propulsion and radioisotope-electric propulsion. The solar-electric propulsion stage consists of six 22 cm diameter “RIT-22”ion thrusters working with a high specific impulse of 7377 s corresponding to a positive grid voltage of 5 kV. Solar power of 53 kW BOM is provided by a light-weight solar array. The REP-stage consists of four space-proven 10 cm diameter “RIT-10” ion thrusters that will be operating one after the other for 9 yrs in total. Four advanced radioisotope generators provide 648 W at BOM. The scientific instrument package is oriented at earlier studies. For its mass and electric power requirement 35 kg and 35 W are assessed, respectively. Optimized trajectory calculations, treated in a separate contribution, are based on our “InTrance” method. The program yields a burn out of the REP stage in a distance of 79.6 AU for a usage of 154 kg of Xe propellant. With a C3 = 45,1 (km/s) 2 a heliocentric probe velocity of 10 AU/yr is reached at this distance, provided a close Jupiter gravity assist adds a velocity increment of 2.7 AU/yr. A transfer time of 23.8 yrs results for this scenario requiring about 450 kg Xe for the SEP stage, jettisoned at 3 AU. We interpret the SEP/REP propulsion as a competing alternative to solar sail and ballistic/REP propulsion. Omiting a Jupiter fly-by even allows more launch flexibility, leaving the mission duration in the range of the ESA specification.
Ion thruster
In-space propulsion technologies
Specific impulse
Laser Propulsion
Cite
Citations (2)
Plasma potential measurements using the conventional Langmuir probe may cause an error due to the space charge effect. To solve the problem, a tube probe is proposed in this study which can minimize the space charge effect by collecting electrons with an orifice instead of the solid surface of the Langmuir probe. The I-V characteristic of the tube probe exhibits a clear turning point, accurately indicating the plasma potential. Comparing with the results of the conventional Langmuir probe, it suggests that the plasma potential measured by the Langmuir probe may be underestimated by about 0.1-0.2 Te/e, which may cause underestimation of the electron density by about 10%-20%. Combination use of the tube probe and the Langmuir probe is suggested for accurate measurement of the electron density.
Langmuir Probe
Electron temperature
Cite
Citations (4)
Abstract : The Report commences with a discussion of the merits of electric propulsion technology, explaining that the very high exhaust velocities attainable allow the propellant masses required for most missions to be drastically reduced. The various types of electric thruster are then described briefly. The most highly developed and potentially useful thruster, the Kaufman electron bombardment ion thruster, is covered in greater detail, with particular reference to the T5 device developed in the UK. Candidate missions are discussed, ranging from attitude and orbit control functions to the application of ion propulsion to the deployment of solar power satellites. Important terrestrial applications of electric propulsion technology are also mentioned. (Author)
Ion thruster
In-space propulsion technologies
Orbit (dynamics)
Cite
Citations (13)
Ion thruster
Faraday cup
Characterization
Faraday cage
Cite
Citations (14)
Complex space missions involving formation flying or drag compensation are driving the need for spacecraft propulsion systems capable of providing low but also highly accurate thrust levels. Currently, no single propulsion device exists that is able to provide both precision and coarse thrust capability over the micro-Newton to milli-Newton thrust range required by these missions. A need for a precision, low thrust, miniature electric propulsion device with a wide throttling range therefore exists. The concept of a differential ion thruster was initially proposed by the Ion Propulsion Group of QinetiQ to address this requirement. It was proposed that an unprecedented throttling range and thrust resolution could be achieved through differential control of opposing ion beams, by which very small net offsets in thrust could be achieved. Single ion beam operation, as for conventional gridded ion thrusters, would permit higher thrust levels to be achieved with high specific impulse. The extraction and independent control of two ion beams from a single gridded ion thruster has never previously been reported. Prototype and breadboard models of the proposed Miniaturised Differential Gridded Ion Thruster (MiDGIT) were designed and manufactured in collaboration with QinetiQ to provide a proof-of-concept and to demonstrate preliminary performance. Test campaigns were conducted at the QinetiQ Large European Electric Propulsion Facilities and within the EP1 vacuum chamber at the University of Southampton. The work reported in this thesis contributes to the first detailed characterisation of a twin-ended radio frequency gridded ion thruster utilising a common plasma discharge. Two control methods were identified which permitted independent control of the ion beams extracted from either end of the thruster. These were: variation of the accelerator grid potential in order to induce changes in the plasma sheath geometry upstream of each screen grid leading to variations in the extracted ion currents, and variation of the RF power delivered to each end of the thruster to generate a higher plasma density on one end of the discharge and ultimately a net thrust out of that end of the thruster. The performance of the MiDGIT thruster has been evaluated with regards to both coarse thrust and fine thrust control requirements. Though the MiDGIT thruster has demonstrated a wide thrust range surpassing competing single-ended miniature ion thrusters, the extraction of two ion beams to achieve very low thrust levels leads to low specific impulse and high specific power for the MiDGIT thruster compared to any other single-ended ion thruster that can achieve the same thrust levels. Recommendations to improve efficiency are made and suggestions for future work and further development of the MiDGIT thruster are given.
Ion thruster
Specific impulse
Cite
Citations (0)
The mercury bombardment electrostatic ion thruster is the most successful electric thruster available today. A 5 cm diameter ion thruster with 3,000 specific impulse and 5mN thrust is described. The advantages of electric propulsion and the tests that will be performed are also presented.
Ion thruster
Specific impulse
Mercury
Cite
Citations (0)
An electric propulsion thrust system has the capability of providing a high specific impulse for long duration scientific missions in space. The EMI from the elements of an ion engine was characterized. The compatibility of ion drive electric propulsion systems with typical interplanetary spacecraft engineering was predicted.
Ion thruster
Specific impulse
In-space propulsion technologies
Laser Propulsion
Electromagnetic Compatibility
Cite
Citations (2)