Ak Knoll

Thrust Balance Characterization of a 200 W Quad Confinement Thruster for High Thrust Regimes

A thrust balance characterization of a low powered Quad Confinement Thruster is presented for high levels of propellant flow. The nominal flow rate for this device is between 1 and 2 sccm of xenon propellant. This paper extends the operating range, and investigates the performance at two high flow conditions of 10 and 20 sccm. Power is varied incrementally between 20 and 200 W in order to characterize the performance versus power trends of the device. It was found that for these high flow regimes the propellant is underutilized, and a proportion of the increased thrust can likely be attributed to a hot gas expansion of the neutral xenon rather than the generation of additional accelerated ions. The thrust was increased from 1 (nominal) to 3.3 mN at 200 W of input power for the 20 sccm condition. However, the performance penalty in terms of the specific impulse was considerable. The specific impulse under these conditions dropped below 200 s, where the nominal condition is 1000 s. A compromise between increased thrust and decreased performance was found at 10 sccm of flow: 3 mN of thrust at 300 s specific impulse.

Impact of plasma noise on a direct thrust measurement system.

In order to evaluate the accuracy and sensitivity of a pendulum-type thrust measurement system, a linear variable differential transformer (LVDT) and a laser optical displacement sensor have been used simultaneously to determine the displacement resulting from an applied thrust. The LVDT sensor uses an analog interface, whereas the laser sensor uses a digital interface to communicate the displacement readings to the data acquisition equipment. The data collected by both sensors show good agreement for static mass calibrations and validation with a cold gas thruster. However, the data obtained using the LVDT deviate significantly from that of the laser sensor when operating two varieties of plasma thrusters: a radio frequency (RF) driven plasma thruster, and a DC powered plasma thruster. Results establish that even with appropriate shielding and signal filtering the LVDT sensor is subject to plasma noise and radio frequency interactions which result in anomalous thrust readings. Experimental data show that the thrust determined using the LVDT system in a direct current plasma environment and a RF discharge is approximately a factor of three higher than the thrust values obtained using a laser sensor system for the operating conditions investigated. These findings are of significance to the electric propulsion community as LVDT sensors are often utilized in thrust measurement systems and accurate thrust measurement and the reproducibility of thrust data is key to analyzing thruster performance. Methods are proposed to evaluate system susceptibility to plasma noise and an effective filtering scheme presented for DC discharges.

A simple isentropic model of electron transport in Hall thrusters

A simple model is presented for the time-averaged electron mobility within a Hall thruster. The model is predicated on the experimental evidence for isentropic electron flow and, when used in a one-dimensional simulation, captures plasma properties that are in reasonable agreement with experiment.

An Axial-Azimuthal Hybrid Simulation of Coaxial Hall Thrusters

We report on progress towards the development of a Hall thruster simulation in the axial-azimuthal (z - θ) computational space. Unlike most computational studies of closeddrift Hall accelerators which have been in one dimension (1D) along the axial direction or in two dimensions (2D) in the axial and radial dimensions, and which require some specification of the axial transport mechanism, this z - θ numerical simulation developed here self-consistently evolves the azimuthal electron drift velocity. The simulation is, in principal, capable of capturing correlated azimuthal disturbances in plasma properties which may give rise to cross-field transport, and makes no use of ad-hoc transport models. Preliminary analysis of the results indicates that azimuthal plasma instabilities may contribute to the axial electron transport process.

Investigating Formation Flying and COTS in an Integrated Simulation Environment

The Canadian Space Agency’s (CSA) Software and Ground Segment (SGS) section has the mandate to develop innovative software and ground segment technologies. The implementation of formation flying concepts for Canadian missions is also currently under investigation at CSAs Spacecraft Engineering section. To that end, there is an ongoing development of a simulation environment to test COTS (Commercial-Off-The-Shelf) and formation flying technologies. Some of todays spacecraft are laboriously custom designed for a specific mission and a limited set of tasks. Development time can be lengthy (several years), which means that designs do not take advantage of the most recent technology. Designs also tend to be extremely inflexible, creating a spacecraft that cannot be easily adapted to future missions. A design ethic that promotes reusability is much more cost-effective and increases the time available to advance new technologies. COTS offer advantages such as a reduced development time, an increased product selection, faster and cheaper parts replacement, and extensively tested advanced designs. The main drawbacks to COTS use in space are susceptibility to radiation and in some cases decreased reliability. Since one of the main advantages of formation flying is the reduced mission sensitivity to a spacecraft failure, the risk associated with COTS, which has hindered its use in conventional space missions, is less concerning in the context of a multiple spacecraft mission. Achieving some level of standardization is a problem currently confronting the space industry, which must be addressed to realize the cost savings that can come from mass production and spacecraft interoperability. The use of standard components with standard interfaces also reduces development time. As well, since part of the goal is to have spacecraft already in orbit regroup and possibly join with new spacecraft to accomplish other missions, some forwards and backwards compatibility between generations of spacecraft will be necessary. This paper describes an integrated simulation environment that uses COTS spacecraft and simulation components to investigate formation flying scenarios and their benefits and challenges. A few of the well-known industry software and hardware tools incorporated into the environment include Analytical Graphics STK, Mathworks Matlab/Simulink, CAE Electronics Real-time Object-Oriented Simulation Environment (ROSETM), Intels StrongARM processor, and the PC-104 architecture for rugged embedded systems. There are short term plans to incorporate goal decomposition hierarchies to implement autonomous operations and robust fault tolerance. This will be implemented on high-speed logic based controller cards, known as Q4 cards, developed by Xiphos Technologies Inc. The simulation environment was developed to allow spacecraft designers or mission operators to test their respective technologies or ideas in a modular structure rapidly, accurately, and cheaply. This will enhance the successful use of formation flying spacecraft and improve the technologies needed to make formation flying a feasible and a cost effective reality.

Experimental Demonstration of an Aluminum-Fueled Propulsion System for CubeSat Applications

He inherent mass, volume, and power constraints of CubeSats puts a stringent limitation on the resources that can be allocated to the propulsion system without overly detracting from the other subsystems on board [1]. Both Chemical Propulsion (CP) and Electric Propulsion (EP) have been proposed for CubeSat applications, with different key constraints in each case. The limited on-board power restricts the use of highly fuel-efficient EP systems, while the CP systems have scaling related issues and require a larger mass and volume of the 1 CubeSat for equivalent levels of ∆ [2]. Nevertheless, future CubeSat missions involving constellation or formation flight will require that the satellites be equipped with a propulsion system that can operate within the stringent resource limitations of the satellite and deliver enough ∆ for their orbital maintenance [3]. There are ongoing research activities in this regard aimed at miniaturizing the space proven and conventional propulsion systems, which can be found in [1, 4-5]. However, scaling of conventional propulsion systems to the size and power limitations of 1 CubeSats, while retaining their operational performances is difficult and complex, and therefore requires the investigation of alternative approaches [6]. This study considers a CP alternative that utilizes aluminium wool as fuel and a mixture of sodium hydroxide and water as an oxidiser [7]. The novelty over conventional propellant combinations is that the reaction takes place at moderate temperatures and at a slow reaction rate. Additionally, the reactants are low cost, easy to handle and can be stored over a long duration without decomposing. The proposed configuration of the propulsion system shown in Fig. 1, is designed to take about one-third of the volume of a 1 CubeSat. It contains a reaction chamber with a nozzle, a plenum volume, two oxidiser tanks with bladders, two cool gas generators and three Lee extended performance valves (IEP Series with FFKM seal option). Table 1 shows the mass budget of the propulsion system. The total system mass is 126 grams, which is about 10% of a 1 CubeSat. The volume is 10×10×3.25 (see Error! Reference source not found.), which represents about 30% of the total volume of the 1 CubeSat.

Experimental Performance Characterization of a Novel Direct Current Cold Cathode Neutralizer for Electric Thruster Applications

The performance of a novel neutralizer for space applications based on a ExB discharge is presented. Preliminary tests were carried out with argon gas and flow rates in the range of 5-10 SCCM. Electrons were extracted through an orifice of diameter 1.8 mm. The maximum extracted current versus input power reported was 2.4 mA/W. The total power input, given by the sum of discharge power plus the extraction power, was in the range of 40-90 W. During extraction tests, the discharge current was limited at 0.2 A due to limit in the cooling system. Future work will be focused on tests at various extraction orifice diameters and cathode materials. Ultimately, xenon and non-conventional gases would be tested as working gases.

Numerical Simulation of High Frequency Wave Coupling within a Hall Thruster

A 2-dimensional Hall thruster simulation has been developed in the axial-azimuthal coordinate plane. The goal of this simulation is to numerically model high frequency plasma waves within the discharge channel of the Hall thruster, and study the contribution of these waves to the time-averaged axial electron drift. This model uses a continuum (fluid) representation for both the electrons and ions. In order to simulate oscillations in the electron field it was necessary to model the electrons dynamically, as opposed to assuming a steady state solution at each time step. The electron momentum equations also include electron inertia terms that are normally neglected in typical Hall thruster models. These inertia terms provide a wave coupling mechanism between axially and azimuthally propagating waves. This numerical model was able to reproduce two dominant high frequency plasma oscillations in the Hall thruster: a 74MHz Kelvin-Helmholtz type shearing instability, and a 7MHz oscillation in the plasma density that has also been observed experimentally. The simulation was successful at predicting the axial electron drift in good agreement with experiment. The results of this study suggest that the plasma oscillations play a dominant role in the electron transport process. In particular, contributions to the electron transport resulting from perturbations in the azimuthal electron velocity were found to be greater than 300% of classical collisional transport.

Experimental Investigation of High Frequency Plasma Oscillations Within Hall Thrusters

[Abstract] An experimental setup has been developed to measure high frequency plasma oscillations within the acceleration channel of a l aboratory Hall thruster. The plasma oscillations are measured with three Langmuir probes separated by small axial and azimuthal offsets. This configuration permits the oscillations to be correlated with direction and wave number. This work is motivated by the anomalous electron transport phenomena, as plasma instabilities may play a crucial role in this transport process. Preliminary data has been gathered downstream of the exit plane of the thruster and suggests high frequency oscillations in the 1 to 10MHz range predominately in the axial direction. Work is currently underway to measure the high frequency oscillations within the acceleration channel at various axial locations.