Performance and Wear Test Results for a 20 kW-Class Ion Engine with Carbon-Carbon Grids

Polk, Jay; Goebel, Dan M.; Synder, John; Schneider, Analyn; Johnson, L. K.; Sengupta, Anita

doi:10.2514/6.2005-4393

Cited by 13 publications

(7 citation statements)

References 18 publications

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“…IV.A (with C cf 0:74), we find that neutral temperatures from 2100 to 5600 K are required. This finding conflicts with the common assumption that the neutral gas accommodates to the thruster wall temperature [30] that, for NEXT, does not exceed approximately 530 K in the steady state [31].…”

Section: Discharge-only With Plasmacontrasting

confidence: 72%

Thrust Stand Characterization of the NASA Evolutionary Xenon Thruster

Diamant¹,

Pollard²,

Crofton³

et al. 2011

Journal of Propulsion and Power

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Direct thrust measurements have been made on the NASA Evolutionary Xenon Thruster using a standard pendulum-style thrust stand constructed specifically for this application. Values have been obtained for the full 40-level throttle table, as well as for a few offnominal operating conditions. Measurements differ from the nominal NASA throttle table 10 values by 3.1% at most, while at 30 throttle levels, the difference is less than 2.0%. When measurements are compared with throttle table 10 values that have been corrected using recently obtained ion beam current density and charge-state data, they differ by 1.2% at most, and at 37 throttle levels, they differ by 1.0% or less. Thrust correction factors calculated from direct thrust measurements and from recently acquired plume data agree to within measurement error for all but one throttle level. Measurements of thrust due to cold flow, neutralizer, and discharge-only operation are also presented.Nomenclature A e = neutralizer exit aperture area c = neutral mean thermal speed C cf = cold flow thrust correction for offaxis flow= load cell reading F T , F = thrust g = acceleration due to gravity at sea level H = dimension J b = ion beam current k = Boltzmann constant, also ratio of specific heats l = distance from pivot to counterweight center of mass L C = dimension L CG = distance from pivot to thruster center of mass L F = distance from pivot to thruster centerline L S = dimension M e = moments due to gas and electrical connections m CW = counterweight mass _ m d = discharge chamber mass flow rate m n , m = neutral atomic mass _ m n = neutralizer mass flow rate m T = thruster mass p e = pressure at neutralizer exit q = elementary charge R = accelerator grid aperture radius, also specific gas constant t = accelerator grid aperture thickness T n = neutral gas temperature T o = neutralizer gas stagnation temperature u e = neutralizer gas effective exhaust velocity V b = ion beam voltage V bps = beam power supply voltage X C = dimension X S = dimension Y = dimension = angle = thrust correction factor for beam divergence = total thrust correction factor = dimension " = angle relative to aperture centerline = pendulum angle = angle = angle from grid centerline

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Section: Discharge-only With Plasmacontrasting

confidence: 72%

Thrust Stand Characterization of the NASA Evolutionary Xenon Thruster

Diamant¹,

Pollard²,

Crofton³

et al. 2011

Journal of Propulsion and Power

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“…An example of the probability of a primary making a collision before finding a cusp is shown in Figure 6 for the case of the NEXIS thruster 10,11 designed with either 4 or 6 cusps. For the design with 6 cusps, it is required to have cusp field strengths approaching 2000 G at the surface of the anode in order to ensure that the primaries are not lost prematurely.…”

Section: B Electron Confinement At the Anodementioning

confidence: 99%

“…While analytical calculations of the field strength in the multipole structure are possible, it is much more common to utilize a commercial available 2-D or 3-D magnetic field solver, like the commercial Maxwell 3-D code 12 , to find the fields. Figure 7 shows the measured and calculated contours of constant magnetic field strength for an example 6-ring cusp thruster design 10 . The 60-gauss contour is closed all along the boundary, which will be shown to provide good ion confinement between the cusps.…”

Section: Ion Confinement At the Anodementioning

confidence: 99%

“…Brophy's model required some measurements of parameters in the thruster or data from plasma and gas flow simulations to accurately predict the thruster behavior. An improved version of this model was developed 8 with plasma potential and discharge stability effects 9 , and was used to predict the performance of the NEXIS thruster 10,11 . This model still also required some plasma parameter inputs (such as electron temperature), and so was not a completely predictive tool.…”

Section: Introductionmentioning

confidence: 99%

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Analytical Ion Thruster Discharge Performance Model

Goebel

Wirz

Katz

2007

Journal of Propulsion and Power

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A particle and energy balance model of the plasma discharge in magnetic ring-cusp ion thrusters has been developed. The model follows the original work of Brophy in the development of global 0-D discharge models that utilize conservation of particles into and out of the thruster and conservation of energy into the discharge and out of the plasma in the form of charged particles to the walls and beam and plasma radiation. The present model is significantly expanded over Brophy's original work by including self-consistent calculations of the internal neutral pressure, electron temperature, primary electron density, electrostatic ion confinement (due to the ring-cusp fields), plasma potential, discharge stability, and time dependent behavior during recycling. The model only requires information on the thruster geometry, ion optics performance and electrical inputs such as discharge voltage and currents, etc. to produce accurate performance curves of discharge loss versus mass utilization efficiency. The model has been benchmarked against the NEXIS Laboratory Model (LM) and Development Model (DM) thrusters, and successfully predicts the thruster discharge loss as a function of mass utilization efficiency for a variety of thrusters. The discharge performance model will be presented and results showing ion thruster performance and stability given.

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“…The most probable thruster candidates for high-power EPS realization are the same as for abovementioned projects (see section 2): HETs [6,9,13,3942], ITs [7,8,43], and MPD thrusters [10,14,44]. Summarizing reference data mentioned above, typical requirements to EPS components can be obtained: 2050-kilowatt thruster with thrust and speci¦c impulse regulation (multimode thruster).…”

Section: Typical Requirements To High-power Electric Propulsion Systementioning

confidence: 99%

Concept of electric propulsion realization for high power space tug

Zakharenkov

Semenkin

Solodukhin

2016

Progress in Propulsion Physics

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Popular at the beginning of the Space Age, ambitious projects aimed at Moon, Mars, and other space objects exploration, have returned with new technology and design level. High power space tug with electric propulsion system (EPS) is mainly considered as a transport vehicle for such missions. Modern high power space tugs projects as well as their spacecraft (SC) power and propulsion systems are reviewed in the paper. The main technologies and design solutions needed for high-power EPS realization are considered.

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Performance and Wear Test Results for a 20 kW-Class Ion Engine with Carbon-Carbon Grids

Cited by 13 publications

References 18 publications

Thrust Stand Characterization of the NASA Evolutionary Xenon Thruster

Thrust Stand Characterization of the NASA Evolutionary Xenon Thruster

Analytical Ion Thruster Discharge Performance Model

Concept of electric propulsion realization for high power space tug

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