Space Vehicle Dynamics and Control, Second Edition

Propulsion is required for satellite motion in outer space. The displacement of a satellite in space, orbit transfer and its attitude control are the task of space propulsion, which is carried out by rocket engines. Electric propulsion uses electric energy to energize or accelerate the propellant. The electric propulsion, which uses electrical energy to accelerate propellant in the form of plasma, is known as plasma propulsion. Plasma propulsion utilizes the electric energy to first, ionize the propellant and then, deliver energy to the resulting plasma leading to plasma acceleration. Many types of plasma thrusters have been developed over last 50 years. The variety of these devices can be divided into three main categories dependent on the mechanism of acceleration: (i) electrothermal, (ii) electrostatic and (iii) electromagnetic. Recent trends in space exploration associate with the paradigm shift towards small and efficient satellites, or micro- and nano-satellites. A particular example of microthruster considered in this paper is the micro-cathode arc thruster (µCAT). The µCAT is based on vacuum arc discharge. Thrust is produced when the arc discharge erodes some of the cathode at high velocity and is accelerated out the nozzle by a Lorentz force. The thrust amount is controlled by varying the frequency of pulses with demonstrated range to date of 1‐50Hz producing thrust ranging from 1 µN to 0.05mN.

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Electric propulsion for small satellites

Presented in this work is a guide to standardize the methods of performing correct and accurate Langmuir probe measurements within the plasma environment produced by an electric propulsion device. ...

Recommended Practice for Use of Langmuir Probes in Electric Propulsion Testing

At present, no Cubesat has flown in space featuring propulsion. This was acceptable as long as CubeSats were flown mostly as university experiments. As CubeSats become of interest to other users in the government and industry communities as well, a larger range of capabilities may be required than exhibited so far, while maintaining the uniqueness of the Cubesat platform. Propulsion capability is crucial in increasing mission capabilities of future CubeSats, such as orbit change and raising, formation flying, proximity operations, fine attitude control, or drag-make-up and de-orbit. While some of these tasks may be accomplished with propellantless devices, their applications are limited, applicable mostly to a single task, and bear their own risks. In this study, a survey was conducted of propulsion technologies applicable to CubeSats. Only few off-the-shelf design solutions exist today. The survey was thus expanded to such devices as well that are under significant development, and are approaching the required design envelope for CubeSats with respect to mass, volume, and power. In some cases, such as electric propulsion devices, CubeSat architectures themselves may need to be adapted, required to feature deployable solar arrays to increase power capabilities. Given the vast scope of this survey, only thruster technologies could be surveyed. However, valves and other feed system components, as well as their integration, are equally important, but have to be left to a future survey. Three major propulsion technology areas applicable to CubeSats emerged when conducting this review: (1) Existing technologies, such as butane systems, pulsed plasma thrusters, and vacuum arc thrusters are applicable to CubeSats today with no or only minor changes, (2) New thruster technologies under significant development, such as hydrazine monopropellant systems, ion engines, or colloid thrusters could be adapted to CubeSats with some further development, especially also in other subsystem areas such as feed systems and power processing units. They will also require increased power capabilities, and (3) emerging technologies, such as micro electrospray arrays and micro cavity discharge arrays that offer even higher flexibility due to scalability for CubeSats, and enable compact integration.

Survey of Propulsion Technologies Applicable to Cubesats

We present in this paper a case study of the probabilistic approach to model-based diagnosis. Here, the diagnosed system is a real-world electrical power system (EPS), i.e., the Advanced Diagnostic and Prognostic Testbed (ADAPT) located at the NASA Ames Research Center. Our probabilistic approach is formally well founded and based on Bayesian networks (BNs) and arithmetic circuits (ACs). We pay special attention to meeting two of the main challenges often associated with real-world application of model-based diagnosis technologies: model development and real-time reasoning. To address the challenge of model development, we develop a systematic approach to representing EPSs as BNs, supported by an easy-to-use specification language. To address the real-time reasoning challenge, we compile BNs into ACs. AC evaluation (ACE) supports real-time diagnosis by being predictable, fast, and exact. In experiments with the ADAPT BN, which contains 503 discrete nodes and 579 edges and produces accurate results, the time taken to compute the most probable explanation using ACs has a mean of 0.2625 ms and a standard deviation of 0.2028 ms. In comparative experiments, we found that, while the variable elimination and join tree propagation algorithms also perform very well in the ADAPT setting, ACE was an order of magnitude or more faster.

/pdf/probabilistic-model-based-diagnosis-an-electrical-power-z6cpirzx5m.pdf

Probabilistic Model-Based Diagnosis: An Electrical Power System Case Study

Abstract : A class of miniaturized pulsed plasma thrusters (PPT), known as MicroPPTs, is currently in development at the Air Force Research Laboratory. Operating by means of a surface discharge across a Teflon (TM) propellant fuel bar only a few millimeters in diameter, the MicroPPT delivers a thrust-to-power ratio of 5-10 Micro-N-s/J. In the near term, these thrusters can provide propulsive attitude control on 150-kilogram-class spacecraft using one-fifth the dry mass of conventional torque rods and reaction wheels. Eventually these thrusters are designed for primary and attitude control propulsion on future 25-kilogram-class spacecraft. Efforts to characterize MicroPPT performance and the thruster plume are underway. To this end, a modified torsional thrust stand has been developed for the purpose of accurately measuring the low-level thrust generated by the MicroPPT. A Herriott Cell interferometer is introduced to establish the plume electron and neutral densities. Comparison of the measured electron density with modeling predictions shows close agreement. Additionally, a Pockels cell has been developed to provide a zero-impedance MicroPPT breakdown voltage measurement, and an intensified CCD array has been used to characterize the divergence of both the thruster plume and the late-time particulate emission. A synopsis is presented of the status of the thruster's development, including lifetime, thermal, and environmental testing. (2 refs.)

AFRL MicroPPT development for small spacecraft propulsion

: The Electric Propulsion Space Experiment (ESEX) is one of eight experiments being launched in early 1997 on the Advanced Research and Global Observation Satellite (ARGOS). ESEX has successfully completed flight qualification and is awaiting integration. ESEX will be the first flight of a high power arcjet propulsion subsystem. The arcjet experiment will nominally fire 10 times for a total of 150 minutes. The flight data will include measurements of both performance and spacecraft integration issues, as well as ground and space observations of the arcjet plume.

Electric Propulsion Space Experiment (ESEX) flight qualification and operations

The U.S. Air Force Research Laboratory’ s Electric Propulsion Space Experiment was launched and successfully operated on the Advanced Research and Global Observation Satellite, demonstrating the compatibility and readiness of high-power electric propulsion. This e ight was the culmination of an Air Force technology program to demonstrate the applicability of high-power electric propulsion for satellite applications. A brief history of the program is presented, followed by a description of the e ight operations, which successfully demonstrated the critical system components, the arcjet, power processor, and propellant feed system, verifying the interoperability of high-power electric propulsion with satellite operations. The two anomalies experienced during on-orbit operations, and their proposed causes, are also described.

Review of the Electric Propulsion Space Experiment (ESEX) Program

A versatile double Langmuir probe technique has been developed by incorporating analytical fits to Laframboise's numerical results for ion current collection by biased electrodes of various sizes relative to the local electron Debye length. Application of these fits to the double probe circuit has produced a set of coupled equations that express the potential of each electrode relative to the plasma potential as well as the resulting probe current as a function of applied probe voltage. These equations can be readily solved via standard numerical techniques in order to determine electron temperature and plasma density from probe current and voltage measurements. Because this method self-consistently accounts for the effects of sheath expansion, it can be readily applied to plasmas with a wide range of densities and low ion temperature (Ti/Te ≪ 1) without requiring probe dimensions to be asymptotically large or small with respect to the electron Debye length. The presented approach has been successfully ap...

Improved analysis techniques for cylindrical and spherical double probes.

During the Electric Propulsion Space Experiment mission, eight e rings of the 26-kW ammonia arcjet were performed. Data from onboard systems,including an accelerometerandglobal positioning system unit, are used to determine thruster performance. In addition, ground-based tracking is used to determine velocity change during these e rings. The mean values of thrust, specie c impulse, and thrust efe ciency are estimated to be 1.93 § 0.06 N, 786.2§ 43.0s,and 0.267 § 0.021,respectively.Thismeasuredperformanceislowerthanexpectedbasedon ground test.Themostlikelycauseofthisdiscrepancyiserrorinonboardmeasurementofdischargepowerduetoa6%drift inthepowerprocessing unitcurrent shunt.Atthecorrected power, performancefallswithintheexpectedenvelope.

D. R. Bromaghim

Papers

AFRL MicroPPT development for small spacecraft propulsion

Electric Propulsion Space Experiment (ESEX) flight qualification and operations

Review of the Electric Propulsion Space Experiment (ESEX) Program

Improved analysis techniques for cylindrical and spherical double probes.

Orbital Performance Measurements of Air Force Electric Propulsion Space Experiment Ammonia Arcjet