"The pulsed plasma thruster (PPT) is an onboard electromagnetic propulsion device currently being considered for use in various small satellite missions. The work presented in this thesis is directed toward improving PPT performance using a control engineering approach along with externally applied magnetic fields. An improved one dimensional electromechanical model for PPT operation is developed. This slug model represents the PPT as an LRC circuit with a dynamics equation for the ablated plasma. The improved model includes detailed derivation for the induced magnetic field and a model for the plasma resistance. A modified electromechanical model for the case of externally applied magnetic fields is also derived for the parallel plate geometry. A software package with a graphical user interface (GUI) is developed for the simulation of various PPT types, geometric configurations, and parameters The simulations show excellent agreement with data from the Lincoln Experimental Satellite (LES)-6, the LES-8/9 PPT and the Univ. of Tokyo PPT. The control objective employed in this thesis involves the maximization of the specific impulse and thrust efficiency of the PPT, which are each directly related with the exhaust velocity of the thruster. This objective is achieved through the use of an externally applied magnetic field as a system actuator. To simulate an open-loop constant-input controller the modified electromechanical PPT model is applied to the various PPT configurations. In this controller the external magnetic field was applied as constant throughout or portions of the PPT channel. For the Univ. of Tokyo PPT a magnetic field applied over the entire 6-cm long channel increases the specific impulse and thrust efficiency by 10% over the case that the filed is applied in the first 1.75 cm of the PPT channel. The magnitude of these increases compare well with the results of the UOT applied B-field experiments. For the LES-6 and LES-9 PPTs, the simulations predicts significant performance enhancements with approximately linear increases for the specific impulse, thrust efficiency and impulse bit. "

"As new, smaller satellites are built, the need for improved on-board propulsion systems has grown. The pulsed plasma thruster has received attention due to its low power requirements, its simple propellant management, and the success of initial flight tests. Successful integration of PPTs on spacecraft requires the comprehensive evaluation of possible plume-spacecraft interactions. The PPT plume consists of neutrals and ions from the decomposition of the Teflon propellant, material from electrode erosion, as well as electromagnetic fields and optical emissions. To investigate the PPT plume, an on-going program is underway at WPI that combines experimental and computational investigations. Experimental investigation of the PPT plume is challenging due to the unsteady, pulsed as well as the partially ionized character of the plume. In this thesis, a triple Langmuir probe apparatus was designed and used to obtain electron temperature and density measurements in the plume of a PPT. This experimental investigation provides further characterization of the plume, much needed validation data for computational models, and is useful in thruster optimization studies. The pulsed plasma thruster used in this study is a rectangular geometry laboratory model built at NASA Lewis Research Center for component lifetime tests and plume studies. It is almost identical in size and performance to the LES 8/9 thruster, ablating 26.6 ug of Teflon, producing an impulse bit of 256 uN-s and a specific impulse of 986 s at 20 J. All experiments were carried out at NASA LeRC Electric Propulsion Laboratory. The experimental setup included triple Langmuir probes mounted on a moveable probe stand, to collect data over a wide range of locations and operating conditions. Triple probes have the ability to instantaneously measure electron temperature and density, and have the benefit of being relatively simple to use, compared to other methods used to measure these same properties. The implementation of this measuring technique is discussed in detail, to aid future work that utilizes these devices. Electron temperature and density was measured from up to 45 degrees from the centerline on planes parallel and perpendicular to the thruster electrodes, for thruster energy levels of 5, 20 and 40 J. Radial distances extend from 6 to 20 cm downstream from the Teflon surface. These locations cover the core of the PPT plume, over a range of energy levels that corresponds to proposed mission operating conditions. Data analysis shows the spatial and temporal variation of the plume. Maximum electron density near the exit of the thruster is 1.6 x 1020, 1.6 x 1021, and 1.8 x 1021 m-3 for the 5, 20 and 40 J discharges, respectively. At 20 cm downstream from the Teflon surface, densities are 1 x 1019, 1.5 x 1020 and 4.2 x 1020 for the 5, 20 and 40 J discharges, respectively. The average electron temperature at maximum density was found to vary between 3.75 and 4.0 eV for the above density measurements at the thruster exit, and 20 cm from the Teflon surface the temperatures are 0.5, 2.5, and 3 eV for the 5, 20 and 40 J discharges. Plume properties show a great degree of angular variation in the perpendicular plane and very little in the parallel plane, most likely due to the rectangular geometry of the PPT electrodes. Simultaneous electron temperature and density traces for a single thruster discharge show that the hottest electrons populate the leading edge of the plume. Analysis between pulses shows a 50% variation in density and a 25% variation in electron temperature. Error analysis estimates that maximum uncertainty in the temperature measurements to be approximately +/- 0.75 eV due to noise smoothing, and the maximum uncertainty in electron density to be +/- 60%, due to assumptions related to the triple probe theory. In addition, analysis of previously observed slow and fast ion components in the PPT plume was performed. The analysis shows that there is approximately a 3 us difference in creation time between the fast and slow ions, and that this correlates almost exactly with the half period of the oscillations in the thruster discharge current."

<p>The commercial CubeSat industry has experienced explosive growth recently, and with falling  costs  and  growing  numbers  of  launch  providers,  the  trend  is  likely  to  continue.  The scientific missions CubeSats could complete are expanding, and this has resulted in a demand for reliable  high  specific  impulse  nanosatellite  propulsion  systems.  Interest  in  liquid-fed  pulsed plasma thrusters (LF-PPTs) to fulfill this role has grown lately. Prior work on a nanosatellite LF-PPT was done in the Purdue Electric Propulsion and Plasma Laboratory, but its high operational voltage and electrode size would be disadvantageous for integration on a CubeSat, which have strict volume limitations and provide only tens of Watts in power at low voltages. This work aims to address those disadvantages and further advance the development of a nanosatellite LF-PPT by reducing the operating voltage and removing long plate electrodes to prevent energy losses on components other than the expelled plasma sheet. Two major objectives are pursued: to construct a  coaxial  pulsed  plasma  thruster  operating  with  10s  to  100s  of  volts  and  to  characterize  the temporal evolution of the discharge parameters in this low-voltage operation scenario. </p> <p>It  took  three  experimental  design  iterations,  all  of  which  used  a  260  <em>uF</em> ,  400 <em>V</em> film capacitor, to arrive at a functional coaxial pulsed plasma thruster. First, a button gun was tested. It produced  a  peak  current  of ~16<em> kA</em>,  which  serves  as  the  expected  maximum  for  the  later experiments. Due to the presence of parasitic arcing, it revealed that electrical lines needed to be removed from vacuum chamber to enable testing at a wide range of pressures. Second, a coaxial PPT was designed, built, and tested. This design confirmed operation at discharge voltages <100 <em>V</em> across the plasma, achieving one of the project’s aims, and produced a peak current of 7.4 <em>kA</em>. However,  necessity  to  better  align  the  cathode and  provide  an  unobstructed  camera  view  for observation of the discharge column attachment to the cathode surface forced additional system redesign. Third, a revised coaxial PPT was built and tested. Using air as a propellant, the discharge generated a peak current of 10.4 <em>kA</em> at a mass flow rate of 2 mgs. The PPT cathode was imaged with an ICCD camera over a wide range of pressures, and the photos indicated “spotless” diffuse arc attachment to the cathode, which serves as evidence to expect low erosion rates. The direct measurements of the cathode erosion rate are planned for future. </p>

"The ablative Teflon pulsed plasma thruster (PPT) is an onboard electromagnetic propulsion enabling technology for small spacecraft missions. The integration of PPTs onboard spacecraft requires the understanding and evaluation of possible thruster/spacecraft interactions. To aid in this effort the work presented in this thesis is directed towards the development and application of Langmuir probe techniques for use in the plume of PPTs. Double and triple Langmuir probes were developed and used to measure electron temperature and density of the PPT plume. The PPT used in this thesis was a laboratory model parallel plate ablative Teflon® PPT similar in size to the Earth Observing (EO-1) PPT operating in discharge energies between 5 and 40 Joules. The triple Langmuir probe was operated in the current-mode technique that requires biasing all three electrodes and measuring the resulting probe currents. This new implementation differs from the traditional voltage-mode technique that keeps one probe floating and requires a voltage measurement that is often susceptible to noise in the fluctuating PPT plume environment. The triple Langmuir probe theory developed in this work incorporates Laframboise’s current collection model for Debye length to probe radius ratios less than 100 in order to account for sheath expansion effects on ion collection, and incorporates the thin-sheath current collection model for Debye length to probe radius ratios greater than 100. Error analysis of the non-linear system of current collection equations that describe the operation of the current-mode triple Langmuir probe is performed as well. Measurements were taken at three radial locations, 5, 10, and 15 cm from the Teflon® surface of the PPT and at angles of 20 and 40 degrees to either side of the thruster centerline as well as at the centerline. These measurements were taken on two orthogonal planes, parallel and perpendicular to the PPT electrodes. A data-processing software was developed and implements the current-mode triple Langmuir probe theory and associated error analysis. Results show the time evolution of the electron temperature and density. Characteristic to all the data is the presence of hot electrons of approximately 5 to 10 eV at the beginning of the pulse, occurring near the peak of the discharge current. The electron temperature quickly drops off from its peak values to 1-2 eV for the remainder of the pulse. Peak electron densities occur after the peak temperatures. The maximum electron density values on the centerline of the plume of a laboratory PPT 10 cm from the Teflon® surface are 6.6x10^19 +/- 1.3x10^19 m^-3 for the 5 J PPT, 7.2x10^20 +/- 1.4x10^20 m^-3 for the 20 J PPT, and 1.2x10^21 +/- 2.7x10^20 m^-3 for the 40 J PPT. Results from the double Langmuir probe taken at r=10 cm, theta perpendicular=70 degrees and 90 degrees of a laboratory PPT showed good agreement with the triple probe method."

No description available.

<p>Prior to the mid-20th century, the idea of electric propulsion had been all but a foreign one that manifested itself in the topic of science fiction. It was around this time when companies and agencies like NASA began to take interest in the topic of space propulsion, as most famously seen in the landing of the Apollo 11 mission on the moon. It was not until the early-1960s where the idea of a pulsed plasma thruster was first realized, with its first test being in 1964 aboard the Russian Zond-2 satellite which contained 6 ablative Polytetrafluoroethylene (PTFE, or “Teflon”) pulsed plasma thrusters.</p> <p>In this paper, a new low-voltage, liquid-fed pulsed plasma thruster was developed, tested, and characterized. This project took influence from the previous low-voltage, liquid-fed pulsed plasma thruster in Purdue’s EPPL and desired to transition it from a current gas-fed system to its intended liquid-fed system. The two main objectives for this project included conducting direct studies of the cathode’s erosion rate using a simple weighing method after simulating a lifetime of discharging the thruster, and completing the initial design of the liquid-fed pulsed plasma thruster using AF-M315E as its propellant while gathering data on its required breakdown voltage, exhaust velocity, and specific impulse.</p> <p>Both objectives were successfully completed, with the following parameters being measured or calculated. The required breakdown voltage was seen to be less than 26kV to keep the ignition spark inside the chamber. For the subsequent results measured however, the breakdown voltage was kept between 10-16kV for all successive tests. The peak current measured for all discharges was an average of 11kA, far exceeding similar geometries such as MPD thrusters. The operational voltage was less than 200V, although an operational voltage closer to 100V is expected after further optimization of the system is completed. The erosion rate of the tungsten cathode at this operational setting was found to be 15.4046 +/- 0.592 microgram/Coulomb which is much less than the cathode spot erosion rate reported for tungsten in literature of about 60 microgram/Coulomb and is beneficial for extending system lifetime. The exhaust velocity was calculated to be 30.6 +/- 4.8km/s which is typical of state-of-the-art PPT electric propulsion devices. The specific impulse was also extrapolated from the ion’s exhaust velocity, calculating to be 3,119 +/- 489 seconds. Future work would require optimization of the propellant injection mechanism to minimize propellant loss.</p>