Scholarly

A Xenon Mass Gauging through Heat Transfer Modeling for Electric Propulsion Thrusters

Year: 2017 Volume: 11 Issue: 1 94 - 105 Pages

Abstract: The current state-of-the-art methods of mass gauging of Electric Propulsion (EP) propellants in microgravity conditions rely on external measurements that are taken at the surface of the tank. The tanks are operated under a constant thermal duty cycle to store the propellant within a pre-defined temperature and pressure range. We demonstrate using computational fluid dynamics (CFD) simulations that the heat-transfer within the pressurized propellant generates temperature and density anisotropies. This challenges the standard mass gauging methods that rely on the use of time changing skin-temperatures and pressures. We observe that the domes of the tanks are prone to be overheated, and that a long time after the heaters of the thermal cycle are switched off, the system reaches a quasi-equilibrium state with a more uniform density. We propose a new gauging method, which we call the Improved PVT method, based on universal physics and thermodynamics principles, existing TRL-9 technology and telemetry data. This method only uses as inputs the temperature and pressure readings of sensors externally attached to the tank. These sensors can operate during the nominal thermal duty cycle. The improved PVT method shows little sensitivity to the pressure sensor drifts which are critical towards the end-of-life of the missions, as well as little sensitivity to systematic temperature errors. The retrieval method has been validated experimentally with CO2 in gas and fluid state in a chamber that operates up to 82 bar within a nominal thermal cycle of 38 °C to 42 °C. The mass gauging error is shown to be lower than 1% the mass at the beginning of life, assuming an initial tank load at 100 bar. In particular, for a pressure of about 70 bar, just below the critical pressure of CO2, the error of the mass gauging in gas phase goes down to 0.1% and for 77 bar, just above the critical point, the error of the mass gauging of the liquid phase is 0.6% of initial tank load. This gauging method improves by a factor of 8 the accuracy of the standard PVT retrievals using look-up tables with tabulated data from the National Institute of Standards and Technology.

Ion Thruster Grid Lifetime Assessment Based on Its Structural Failure

Year: 2015 Volume: 9 Issue: 11 1989 - 1992 Pages

Abstract: This article developed an ion thruster optic system sputter erosion depth numerical 3D model by IFE-PIC (Immersed Finite Element-Particle-in-Cell) and Mont Carlo method, and calculated the downstream surface sputter erosion rate of accelerator grid; compared with LIPS-200 life test data. The results of the numerical model are in reasonable agreement with the measured data. Finally, we predicted the lifetime of the 20cm diameter ion thruster via the erosion data obtained with the model. The ultimate result demonstrated that under normal operating condition, the erosion rate of the grooves wears on the downstream surface of the accelerator grid is 34.6μm⁄1000h, which means the conservative lifetime until structural failure occurring on the accelerator grid is 11500 hours.

Current Status of 5A Lab6 Hollow Cathode Life Tests in Lanzhou Institute of Physics, China

Year: 2015 Volume: 9 Issue: 11 675 - 678 Pages

Abstract: The current statuses of lifetime test of LaB6 hollow cathode at the Lanzhou Institute of Physics (LIP), China, was described. 5A LaB6 hollow cathode was design for LIPS-200 40mN Xenon ion thruster, and it could be used for LHT-100 80 mN Hall thruster, too. Life test of the discharge and neutralizer modes of LHC-5 hollow cathode were stared in October 2011, and cumulative operation time reached 17,300 and 16,100 hours in April 2015, respectively. The life of cathode was designed more than 11,000 hours. Parameters of discharge and key structure dimensions were monitored in different stage of life test indicated that cathodes were health enough. The test will continue until the cathode cannot work or operation parameter is not in normally. The result of the endurance test of cathode demonstrated that the LaB6 hollow cathode is satisfied for the required of thruster in life and performance.

Retarding Potential Analyzer Design and Result Analysis for Ion Energy Distribution Measurement of the Thruster Plume in the Laboratory

Year: 2012 Volume: 6 Issue: 11 2546 - 2549 Pages

Abstract: Plasma plume will be produced and arrive at spacecraft when the electric thruster operates on orbit. It-s important to characterize the thruster plasma parameters because the plume has significant effects or hazards on spacecraft sub-systems and parts. Through the ground test data of the desired parameters, the major characteristics of the thruster plume will be achieved. Also it is very important for optimizing design of Ion thruster. Retarding Potential Analyzer (RPA) is an effective instrument for plasma ion energy per unit charge distribution measurement. Special RPA should be designed according to certain plume plasma parameters range and feature. In this paper, major principles usable for good RPA design are discussed carefully. Conform to these principles, a four-grid planar electrostatic energy analyzer RPA was designed to avoid false data, and details were discussed including construction, materials, aperture diameter and so on. At the same time, it was designed more suitable for credible and long-duration measurements in the laboratory. In the end, RPA measurement results in the laboratory were given and discussed.

Evaluation of a Dual-Fluid Cold-Gas Thruster Concept

Year: 2012 Volume: 6 Issue: 8 1635 - 1640 Pages

Abstract: A new dual-fluid concept was studied that could eventually find application for cold-gas propulsion for small space satellites or other constant flow applications. In basic form, the concept uses two different refrigerant working fluids, each having a different saturation vapor pressure. The higher vapor pressure refrigerant remains in the saturation phase and is used to pressurize the lower saturation vapor pressure fluid (the propellant) which remains in the compressed liquid phase. A demonstration thruster concept based on this principle was designed and built to study its operating characteristics. An automotive-type electronic fuel injector was used to meter and deliver the propellant. Ejected propellant mass and momentum were measured for several combinations of refrigerants and hydrocarbon fluids. The thruster has the advantage of delivering relatively large total impulse at low tank pressure within a small volume.

Experiment Study on the Plasma Parameters Measurement in Backflow Region of Ion Thruster

Year: 2012 Volume: 6 Issue: 11 1527 - 1530 Pages

Abstract: The charge-exchange xenon (CEX) ion generated by ion thruster can backflow to the surface of spacecraft and threaten to the safety of spacecraft operation. In order to evaluate the effects of the induced plasma environment in backflow regions on the spacecraft, we designed a spherical single Langmuir probe of 5.8cm in diameter for measuring low-density plasma parameters in backflow region of ion thruster. In practice, the tests are performed in a two-dimensional array (40cm×60cm) composed of 20 sites. The experiment results illustrate that the electron temperature ranges from 3.71eV to 3.96eV, with the mean value of 3.82eV and the standard deviation of 0.064eV. The electron density ranges from 8.30×1012/m3 to 1.66×1013/m3, with the mean value of 1.30×1013/m3 and the standard deviation of 2.15×1012/m3. All data is analyzed according to the “ideal" plasma conditions of Maxwellian distributions.