John J. Blandino

Electric Propulsion and Controller Design for Drag-Free Spacecraft Operation

A study is presented detailing the simulation of a drag-free follow-up mission to NASA’s Gravity Recovery and Climate Experiment. This work evaluates controller performance, as well as thrust, power, and propellant mass requirements for drag-free spacecraft operation at orbital altitudes of 160–225 km. In addition, sensitivities to thermospheric wind, Global Positioning System signal accuracy, and availability of ephemeris data are studied. Thruster (control actuator) models are based on two different Hall thrusters for providing the orbital along-track acceleration, colloid thrusters for the normal acceleration, and a miniature xenon ion thruster for the cross-track acceleration. At an altitude of 160 km, the maximum along-track thrust component is calculated to be 98 mNwith a required dynamic (throttling) response of 42 mN=s. Themaximum position error at this altitude was shown to be in the along-track direction with a magnitude of 3314.9 nm. At 225 km, the maximum along-track thrust component reduces to 10.3 mN. The maximum dynamic response at this altitude is 4:0 mN=s. For the spacecraft point design considered with a propellant mass fraction of 0.18, the mission lifetime for the 160 km case was calculated to be 0.76 years. This increases 2.27 years at an altitude of 225 km.

Propulsion Requirements for Drag-Free Operation of Spacecraft in Low Earth Orbit

The use of drag-free spacecraft in low Earth orbit has the potential for enabling gravitational surveys with substantial improvement in sensitivity over the Gravity Recovery and Climate Experiment (GRACE) mission. We evaluate the required thrust envelope, maximum thruster dynamic response, and ΔV for a 5-year drag-free mission over a range of low-Earth-orbit altitudes (150-450 km). The analysis uses the Jacchia-Roberts atmospheric density model and an adaptive, proportional-derivative control algorithm in which full-state error feedback is used to command compensation of disturbing forces by the thrusters. The actuator (thruster) control resolution and control loop frequency are limited to 0.1 pN and approximately 0.8-31 Hz, respectively. Detailed results at 450 km show that the thrust ranges from 10 to 145 μN, the maximum thruster dynamic response is 5.9 μN/s, and the ΔV for a 5-year mission is approximately 20 m/s. Although drag is the dominant disturbance force, the disturbance due to solar radiation pressure is on the same order at 450 km. At 250 km, the thrust ranges from 1.3 to 5.4 mN, the maximum thruster dynamic response is 6.9 pN/s, and the A V for a 5-year mission increases to 865 m/s. We consider seven different electric propulsion options as possible candidates for this mission. The selection will be strongly dependent on the altitude, which for the altitudes considered, increases the required thrust (and power) through several orders of magnitude. We conclude that existing thruster technology should be adequate for altitudes above 250 km, but below this the required power and AV for a 5-year mission may be prohibitive.

Overview of Electric Propulsion Research in U.S. Academia

An overview is presented of electric propulsion research carried out in U.S. academic institutions. Universities in the U.S. are engaged in a wide range of research, varying from fundamental studies in micro thruster concepts, to future flight involving magneto plasma sails.

Application of Micronewton Thrusters for Control of Multispacecraft Formations in Earth Orbit

Previous studies have explored options for separated spacecraft interferometers for planet detection as well as optical imaging of astrophysical objects. In this study, the feasibility of micronewton thrusters (100μN maximum) to provide a “drag -free” platform for earth orbital missions is investigated. The representative mission selected for this study consists of two identical spacecraft in a near polar (i=88.5o), elliptic (e = 0.001), orbit with a nominal 475 km altitude and 220 km in-track separation. Using a full-state feedback control law, it is found that the assumed 100μN maximum thrust level (in each axis) is adequate to maintain the desired orbital altitude and in-track separation with a total ΔV of 24.9 m/s for a five year mission. In addition, the required maximum dynamic response from the thrusters for the assumed control law was evaluated and found to be approximately 0.89 μN/s.

Cubesat Design and Attitude Control with Micro Pulsed Plasma Thrusters

A feasibility study is presented for using micro Pulsed Plasma Thrusters (μPPT) on a 3U CubeSat to perform attitude control to meet pointing requirements at altitudes between 400 and 1000 km. The orbits and attitude control are consistent with science missions for ionospheric and solar observations. The positioning of the thrusters is such that they provide full, three-axis rotational control while maximizing the efficiency of each pulse. The aerodynamic, magnetic, solar radiation, and gravity gradient disturbance torques are included in the formulation. Sensor modeling is also included. The study involves realistic power constraints anticipated on the 3U CubeSat. The paper presents the overall design of the 3U CubeSat and focuses on the attitude determination and control (ADC) approach and implementation. The paper also introduces the control strategy and numerical simulation results for stabilization, pointing and spinning applications.

Micronewton Thruster Requirements for Earth-Orbiting Imaging Formations

Previous studies have explored options for separated spacecraft interferometers for planet detection as well as optical imaging of astrophysical objects. As others have noted, one of the benefits of Earth-orbiting formations is the potential to reduce the propellant required to complete image (u-v) plane filling maneuvers by taking advantage of gravity-gradient forces. In this work, the equations of motion for each of the three spacecraft in a linear formation are numerically integrated using an acceleration profile previously proposed by DeCou to produce a gradual filling of the image plane. The analysis also includes the addition of the J 2 perturbing acceleration and provides the detailed spacecraft position, velocity, and thrust histories for a maneuver as well as the maximum baseline rate of change and ΔV. Results suggest that micronewton thrusters under development for other missions requiring precision spacecraft control might be enabling for the Earth-orbiting imaging formation considered here. For one of the reference cases in geostationary orbit with a maximum baseline of 1 km and an allocated maneuver time of five days, the maximum thrust and throttling rate are approximately 87 pN and 0.24 μN/min, respectively, with a maximum baseline rate of change of 2.4 cm/s and ΔV of 6 cm/s for a single maneuver (one observation). The same reference case with thruster cancellation of the J 2 acceleration results in a maximum thrust of 2.1 mN and ΔV of 3.6 The The maximum thrust, baseline rate of change, and ΔV are presented for maneuver times of one to five days, maximum baselines of 250 m and 1 km, and a formation altitude of 35,786 and 20,000 km. Three propulsion technology options-pulsed plasma, field emission electric propulsion, and colloid-are discussed as possible candidates for this role. Two solution strategies are proposed to the problem raised by the J 2 perturbation.

Analysis of a Lithium Vaporizer for the Advanced Lithium-Fed Applied-Field Lorentz Force Accelerator

The lithium vaporizer for a high power (magnetoplasmadynamic) MPD thruster is modeled using a one-dimensional, thermal-resistive network. We use this model to calculate the required vaporizer length and power as a function of mass flowrate, channel geometry, and material properties. After comparing results from this model with preheat power data for a 200 kW thruster developed at the Moscow Aviation Institute, we investigate performance over a parameter space of interest for the Advanced Lithium-Fed, Applied-field, Lorentz Force Accelerator (ALFA 2 ) thruster. The cold-start heater power for 80 mg/s is found to range from 2.9 to 3.4 kW, corresponding to a vaporizer (axial) length of 11.5 to 24 cm. Heater power sensitivity to cathode tube emissivity, mass flowrate, and vapor superheat are also presented. Finally, the network model results are used to provide boundary conditions for a finite element thermal model of the cathode assembly. This model is used to calculate a higher resolution temperature distribution throughout the cathode assembly.

Feasibility for Orbital Life Extension of a CubeSat Flying in the Lower Thermosphere

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Analysis of triple Langmuir probe measurements in the near-exit region of a gas-fed pulsed plasma thruster

Triple Langmuir probes were used to measure electron number density, and electron temperature in the near-exit region of a laboratory model gas-fed pulsed plasma thruster. Triple Langmuir probe data was obtained on a plane parallel to the thruster electrodes at radial distances of 3 and 7 cm downstream of the propellant inlet and angular positions of 0, 10,20, and 30 degrees. The thruster was operated with Xe propellant, 2 J per pulse, and a mass flow rate of 3 mg/s. Analysis shows that average density at the thruster exit plane is in the range of 5x10” to 2.5~10’~ mv3 and temperature is in the range of 0.5 to 4 eV. At a radial distance of 4 cm downstream from the exit, the density is in the range of 2x10” to 1~10’~ mm3 and temperature in the range of 0.2 to 1.4 eV. Temperature averaged over the duration of a pulse is in the range of 0.4 to 1.3 eV and shows angular and radial variation. A APPT 4 Nomenclature probe area ablative pulsed plasma thruster sheath thickness GFPPT I ? J J; gas-fed pulse plasma thruster current sample-average current random electron thermal current ion flux at the surface of an electrode Kn Knudsen number L probe length M mass of ion i n number density % sample-average lectron number * Graduate Fellow, CGPL, Mechanical Engineering Dept., WPI, 100 Institute Rd., MA 01609. Member AlAA. + Assistant Professor, CGPL Mechanical Engineering Dept., WPI, 100 Institute Rd., MA 01609. Senior Member AIAA. t Member Technical Staff, NASA JPL, Pasadena, CA 9 1109. Senior Member AIAA. B Graduate Student:Research Assistant, EPPDyL, MEA Dept., Princeton University, NJ 08544. Member AIAA. **Assistant Professor, EPPDyL, MAE Dept., Princeton University, NJ 08544. Senior Member AIAA. Copyright