Design and Operation of a Multiple-Cathode, High-Power, Rectangular Discharge Chamber

Peer Reviewedhttp://deepblue.lib.umich.edu/bitstream/2027.42/77200/1/AIAA-2005-4407-602.pd

Electrical/Chemical Thruster using the Same Monopropellant and Method

A thruster operable in a chemical mode or in an electrospray mode using the same liquid monopropellant for operation in both modes is described having a multiplicity of a microthrusters made of a catalytic material having a bore therethrough, where, when operated in the chemical mode, the microthrusters are heated to decompose the monopropellant the monopropellant flows therethrough to generate relatively high thrust. An extractor is positioned downstream of the outlet ends of the microthrusters, such that when the system is operated in its electrospray mode the flowrate of the monopropellant through the microthrusters is substantially lower than in the chemical mode and the extractor is energized with an electric field so that ions and droplets are discharged from the microthrusters and accelerated so as to yield a relatively high specific impulse

Dormant Cathode Plasma Properties and Erosion Analysis in a Multiple-Cathode, High-Power, Rectangular Discharge Chamber

Peer Reviewedhttp://deepblue.lib.umich.edu/bitstream/2027.42/76546/1/AIAA-2005-4241-437.pd

Design and Development of a Multi-Mode Monopropellant Electrospray Mircropropulsion System

Multi-mode spacecraft propulsion is the use of two or more types of propulsive devices on a spacecraft that share some commonality in terms of either hardware or propellant. An example is the Mars Global Surveyor, which made use of hydrazine as both a monopropellant for attitude control and a bipropellant for primary maneuvering. Specific to this study is a multi-mode propulsion system making use of a high-thrust chemical mode and a high-specific impulse electric mode. Using these two modes can be beneficial in two primary ways. One way is by designing a mission such that the high-thrust and high-specific impulse maneuvers are conducted in such a way that it provides a more optimum trajectory over a single chemical or single electric maneuver. The second is to increase the mission flexibility of a single spacecraft architecture in that both high-thrust and high-specific impulse maneuvers are available to mission designers at will, perhaps even allowing for drastic changes in the mission plan while on-orbit or with a relatively short turnaround from concept to launch. For the second method, it is extremely beneficial to utilize a shared propellant for both modes as this provides the highest flexibility in terms of mission design choices. 9 Previous research has investigated a multi-mode system utilizing a single ionic liquid propellant for chemical monopropellant and electrospray modes. Two propellants were developed that may not only function, but theoretically perform well in both modes. These propellants, based on binary mixtures of ionic liquid fuels [Emim][EtSO4] and [Bmim][NO3] with ionic liquid oxidizer hydroxylammonium nitrate (HAN), have been previously synthesized and tested for thermal and catalytic decomposition in a microreactor and electrosprayed in a capillary emitter. This paper will present the design of a multi-mode micropropulsion system for nano- and picosatellites. The system uses a shared propellant, as described above, and shared hardware including tanks, feed lines, valves, and thruster to provide both monopropellant and electrospray propulsive capabilities. The thruster is a catalytic microtube integrated with a capillary electrospray emitter. Using experimental results investigating operation of each mode separately, power, mass, and flow control requirements are developed and scale favorably for small satellite systems. For a 6U cubesat, the system has a three times larger mission design space compared to a system using completely separate, state-of-the-art monopropellant and electrospray thrusters. The effect of using a fully integrated thruster is shown to increase delta-V capability at a given mission duration time by 40% compared to using separate, state-of-the-art thrusters. The final paper will include potential mission applications of the final propulsion system design for a cubesat system

Comparison of Magnetic Probe Calibration at Nano and Millitesla Magnitudes

Magnetic field probes are invaluable diagnostics for pulsed inductive plasma devices where field magnitudes on the order of tenths of tesla or larger are common. Typical methods of providing a broadband calibration of Ḃ probes involve either a Helmholtz coil driven by a function generator or a network analyzer. Both calibration methods typically produce field magnitudes of tens of microtesla or less, at least three and as many as six orders of magnitude lower than their intended use. This calibration factor is then assumed constant regardless of magnetic field magnitude and the effects of experimental setup are ignored. This work quantifies the variation in calibration factor observed when calibrating magnetic field probes in low field magnitudes. Calibration of two B? probe designs as functions of frequency and field magnitude are presented. The first Ḃ probe design is the most commonly used design and is constructed from two hand-wound inductors in a differential configuration. The second probe uses surface mounted inductors in a differential configuration with balanced shielding to further reduce common mode noise. Calibration factors are determined experimentally using an 80.4 mm radius Helmholtz coil in two separate configurations over a frequency range of 100-1000 kHz. A conventional low magnitude calibration using a vector network analyzer produced a field magnitude of 158 nT and yielded calibration factors of 15 663 ± 1.7% and 4920 ± 0.6% T [over]Vs at 457 kHz for the surface mounted and hand-wound probes, respectively. A relevant magnitude calibration using a pulsed-power setup with field magnitudes of 8.7-354 mT yielded calibration factors of 14 615 ± 0.3% and 4507 ± 0.4% T [over]Vs at 457 kHz for the surface mounted inductor and hand-wound probe, respectively. Low-magnitude calibration resulted in a larger calibration factor, with an average difference of 9.7% for the surface mounted probe and 12.0% for the hand-wound probe. The maximum difference between relevant and low magnitude tests was 21.5%

Simple Penning Ion Source for Laboratory Research and Development Applications

A simple Penning ion generator (PIG) that can be easily fabricated with simple machining skills and standard laboratory accessories is described. The PIG source uses an iron cathode body, samarium cobalt permanent magnet, stainless steel anode, and iron cathode faceplate to generate a plasma discharge that yields a continuous 1 mA beam of positively charged hydrogen ions at 1 mTorr of pressure. This operating condition requires 5.4 kV and 32.4 W of power. Operation with helium is similar to hydrogen. The ion source is being designed and investigated for use in a sealed-tube neutron generator; however, this ion source is thoroughly described so that it can be easily implemented by other researchers for other laboratory research and development applications

Comparison of Magnetic Probe Calibration at Nano and Millitesla Magnitudes

Magnetic field probes are invaluable diagnostics for pulsed inductive plasma devices where field magnitudes on the order of tenths of tesla or larger are common. Typical methods of providing a broadband calibration of Ḃ probes involve either a Helmholtz coil driven by a function generator or a network analyzer. Both calibration methods typically produce field magnitudes of tens of microtesla or less, at least three and as many as six orders of magnitude lower than their intended use. This calibration factor is then assumed constant regardless of magnetic field magnitude and the effects of experimental setup are ignored. This work quantifies the variation in calibration factor observed when calibrating magnetic field probes in low field magnitudes. Calibration of two B? probe designs as functions of frequency and field magnitude are presented. The first Ḃ probe design is the most commonly used design and is constructed from two hand-wound inductors in a differential configuration. The second probe uses surface mounted inductors in a differential configuration with balanced shielding to further reduce common mode noise. Calibration factors are determined experimentally using an 80.4 mm radius Helmholtz coil in two separate configurations over a frequency range of 100-1000 kHz. A conventional low magnitude calibration using a vector network analyzer produced a field magnitude of 158 nT and yielded calibration factors of 15 663 ± 1.7% and 4920 ± 0.6% T [over]Vs at 457 kHz for the surface mounted and hand-wound probes, respectively. A relevant magnitude calibration using a pulsed-power setup with field magnitudes of 8.7-354 mT yielded calibration factors of 14 615 ± 0.3% and 4507 ± 0.4% T [over]Vs at 457 kHz for the surface mounted inductor and hand-wound probe, respectively. Low-magnitude calibration resulted in a larger calibration factor, with an average difference of 9.7% for the surface mounted probe and 12.0% for the hand-wound probe. The maximum difference between relevant and low magnitude tests was 21.5%

Performance Measurements of Electric Solid Propellant in an Ablative Pulsed Electric Thruster

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. These propellants may also be used for electric in-space propulsion, specifically in the ablative pulsed plasma thruster. In this paper, we will investigate the performance of an electric solid propellant operating in an ablation-fed pulsed plasma device by use of an inverted pendulum micro-Newton thrust stand. Namely, the impulse-per-pulse and the specific impulse of the device using the electric solid propellant will be reported for test runs of 100 pulses and energy levels of 5, 10, 15 and 20 J. Further, the device will also be tested using the current state-of-the-art pulsed plasma thruster propellant, polytetrafluoroethylene. The performance of each propellant will be compared for each energy level using an identical setup and apparatus. This comparison of performance between propellants in a controlled setting will allow for better understanding of previous experimental observations