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Polzin, Kurt

Publications and source records attributed to Polzin, Kurt.

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Electric Sail Propulsion for Deep Space Missions

Electrostatic Sail (E-sail) propulsion extracts momentum from the solar wind, which has minimal speeds of 400 km/sec, through electrostatic repulsion of the positively charged solar wind ions. This momentum exchange is accomplished by an array of multi-kilometer-length charged tethers biased to a high positive voltage, producing an electric field around the tethers that deflect the positively charged solar wind particles. This electric field grows in diameter as the spacecraft moves away from the sun, increasing the E-sail effective area. The growth of the E-sail effective area allows the created propulsive force to decrease at a rate closer to 1/r out to distances from the sun of up to 20 Astronomical Units (AU). This is unlike solar sail propulsion, where the thrust decreases at a rate of 1/r2 and is only effective out to distances of ~5 AU. The propulsive force is applied without expending any propellant. Although the thrust generated by an E-Sail is low, it can be applied continuously over a period of years (depending on the mission type), and can push a 500 kg spacecraft to tremendous velocities—as high as 12 AU per year—making the voyage to 600 AU in as few as 50 years after launch. The NASA Marshall Space Flight Center recently completed a NASA Innovative Advanced Concept (NIAC) study in which all of the major system- and subsystem-level aspects of an E-Sail-propelled vehicle were assessed and a maturation plan developed. If implemented, E-Sail propulsion could be realized and deployed as a viable option for space missions within a decade.

Interstellar Probe↗
NASA NTRS2017

Coilgun Acceleration Model Containing Interactions Between Multiple Coils

Electromagnetic (EM) accelerators have the potential to fill a performance range not currently being met by conventional chemical and electric propulsion systems by providing a specific impulse of 600-1000 seconds and a thrust-to-power ratio greater than 200 mN/kW. A propulsion system based on EM acceleration of small projectiles has the traditional advantages of using a pulsed system, including precise control over a range of thrust and power levels as well as rapid response and repetition rates. Furthermore, EM accelerators have lower power requirements than conventional electric propulsion systems since no plasma creation is necessary. A coilgun is a specific type of EM device where a high-current pulse through a coil of wire interacts with a conductive projectile via an induced magnetic field to accelerate the projectile. There are no physical or electrical connections to the projectile, which leads to less system degradation and a longer life expectancy. Multi-staging a coilgun by adding multiple turns on a single coil or on the projectile increases the inductance, thus permitting acceleration of the projectile to higher velocities. Previously, a simplified problem of modeling an inductively-coupled, single-coil coilgun using a circuit-based analysis coupled to the one-dimensional momentum equation through Lenz's law was solved; however, the analysis was only conducted on uncoupled coils. The problem is significantly more complicated when multiple, independently-powered coils simultaneously operate and interact with each other and the projectile through induced magnetic fields. This paper presents a multi-coil model developed with the magnetostatic finite element solver QuickField. In the model, mutual inductance values between pairs of conductors were found by first computing the magnetic field energy for different cases where individual coils or multiple coils carry current, then integrating over the entire finite element domain for each case, and finally using the definition of inductive energy storage to solve for the self and mutual inductance. The electric circuit model is coupled to the projectile through Lenz's law, with the coils coupled through mutual inductance but able to be independently triggered at different times to optimize the acceleration profile. This initial model to predict the behavior of a projectile's acceleration through a coupled, multi-coil coilgun increases the potential of building a highly efficient coilgun thruster with key advantages over other EM thruster systems, thus making it a promising candidate for satellite main propulsion or attitude control thrusters.

Liu, Connie↗
NASA NTRS2016

Overview of Iodine Propellant Hall Thruster Development Activities at NASA Glenn Research Center

NASA is continuing to invest in advancing Hall thruster technologies for implementation in commercial and government missions. There have been several recent iodine Hall propulsion system development activities performed by the team of the NASA Glenn Research Center, the NASA Marshall Space Flight Center, and Busek Co. Inc. In particular, the work focused on qualification of the 200 W Busek BHT-200-I and the continued development of the 600 W BHT-600-I Hall thruster propulsion systems. This paper presents an overview of these development activities and also reports on the results of short duration tests that were performed on the engineering model BHT-200-I and the development model BHT-600-I Hall thrusters.

Hall Thrusters↗
NASA NTRS2016

Overview of Iodine Propellant Hall Thruster Development Activities at NASA Glenn Research Center

NASA is continuing to invest in advancing Hall thruster technologies for implementation in commercial and government missions. There have been several recent iodine Hall propulsion system development activities performed by the team of the NASA Glenn Research Center, the NASA Marshall Space Flight Center, and Busek Co. Inc. In particular, the work focused on qualification of the Busek BHT-200-I, 200 W and the continued development of the BHT-600-I Hall thruster propulsion systems. This presentation presents an overview of these development activities and also reports on the results of short duration tests that were performed on the engineering model BHT-200-I and the development model BHT-600-I Hall thrusters.

Electric Propulsion↗
NASA NTRS2016

Overview of NASA Iodine Hall Thruster Propulsion System Development

NASA is continuing to invest in advancing Hall thruster technologies for implementation in commercial and government missions. The most recent focus has been on increasing the power level for large-scale exploration applications. However, there has also been a similar push to examine applications of electric propulsion for small spacecraft in the range of 300 kg or less. There have been several recent iodine Hall propulsion system development activities performed by the team of the NASA Glenn Research Center, the NASA Marshall Space Flight Center, and Busek Co. Inc. In particular, the work focused on qualification of the Busek 200-W BHT-200-I and development of the 600-W BHT-600-I systems. This paper discusses the current status of iodine Hall propulsion system developments along with supporting technology development efforts.

Iodine↗
NASA NTRS2015

Electronegative Gas Thruster - Direct Thrust Measurement Project

This effort is an international collaboration and academic partnership to mature an innovative electric propulsion (EP) thruster concept to TRL 3 through direct thrust measurement. The initial target application is for Small Satellites, but can be extended to higher power. The Plasma propulsion with Electronegative GASES (PEGASES) concept simplifies ion thruster operation, eliminates a neutralizer requirement and should yield longer life capabilities and lower cost implementation over conventional gridded ion engines. The basic proof-of concept has been demonstrated and matured to TRL 2 over the past several years by researchers at the Laboratoire de Physique des Plasma in France. Due to the low maturity of the innovation, there are currently no domestic investments in electronegative gas thrusters anywhere within NASA, industry or academia. The end product of this Center Innovation Fund (CIF) project will be a validation of the proof-of-concept, maturation to TRL 3 and technology assessment report to summarize the potential for the PEGASES concept to supplant the incumbent technology. Information exchange with the foreign national will be one-way with the exception of the test results. Those test results will first go through a standard public release ITAR/export control review, and the results will be presented in a public technical forum, and the results will be presented in a public technical forum.

Dankanich, John↗
NASA NTRS2015

Direct Drive Solar-Powered Arcjet Thruster

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft-provided power to the voltage and current that a thruster needs for operation. NASA Marshall Space Flight Center has initiated fundamental studies on whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The proposed work will aim to refine the proof-of-concept presently being assembled and begin to identify and address technical questions related to power conditioning and noise suppression in the system, and heating of the thruster in long-duration operation. The apparatus proposed for investigation has a target power level of 400 to 1,000 W. The proposed direct-drive arcjet is potentially a highly scalable concept, applicable to spacecraft with up to hundreds of kilowatts and beyond. The design of the arcjet built for this effort was based on previous low power (1 kW class) arcjets.1-3 It has a precision machined 99.95% pure tungsten anode that also serves as the nozzle with a 0.040-in- (1-mm-) diameter, 0.040-in-long constrictor region. An additional anode with a 0.020-in- (0.5-mm-) diameter, 0.020-inlong constrictor region was purchased, but has not yet been used. The cathode is a 0.125-in-diameter tungsten welding electrode doped with lanthum-oxygen; its tip was precision ground to a 308deg angle and terminates in a blunt end. The two electrodes are separated by a boron-nitride insulator that also serves as the propellant manifold; it ends in six small holes which introduce the propellant gas in the diverging section of the nozzle, directly adjacent to the cathode. The electrodes and insulator are housed in a stainless-steel outer body, with a Macor insulator at the mid-plane to provide thermal isolation between the front and back halves of the device. The gas seals were made using Grafoil gaskets. Figure 1(a) shows the assembled thruster; figure 1(b) shows the thruster in the vacuum chamber with electrical and propellant connections.

Polzin, Kurt↗
NASA NTRS2015

Electronegative Gas Thruster

The project is an international collaboration and academic partnership to mature an innovative electric propulsion thruster concept to Technology Research Level-3 (TRL-3) through direct thrust measurement. The project includes application assessment of the technology ranging from small spacecraft to high power. The Plasma propulsion with Electronegative GASES(PEGASES) basic proof of concept has been matured to TRL-2 by Ane Aanesland of Laboratoire de Physique des Plasma at Ecole Polytechnique. The concept has advantages through eliminating the neutralizer requirement and should yield longer life and lower cost over conventional gridded ion engines. The objective of this research is to validate the proof of concept through the first direct thrust measurements and mature the concept to TRL-3.

Dankanich, John↗
NASA NTRS2013

High-Speed Imaging of the First Kink Mode Instability in a Magnetoplasmadynamic Thruster

One of the biggest challenges to efficient magnetoplasmadynamic thruster (MPDT) operation is the onset of high-frequency voltage oscillations as the discharge current is increased above a threshold value. The onset regime is closely related to magnetohydrodynamic instabilities known as kink modes. This work documents direct observation of the formation and quasi-steady state behavior of an argon discharge plasma in a MPDT operating at discharge currents of 8 to 10 kA for a pulse length of approximately 4 ms. A high-speed camera images the quasi-steady-state operation of the thruster at 26,143 fps with a frame exposure time of 10 micro s. A 0.9 neutral density filter and 488-nm argon line filter with a 10-nm bandwidth are used on separate trials to capture the time evolution of the discharge plasma. Frame-by-frame analysis of the power flux incident on the CCD sensor shows both the initial discharge plasma formation process and the steady-state behavior of the discharge plasma. Light intensity levels on the order of 4-6 W/m2 indicate radial and azimuthal asymmetries in the concentration of argon plasma in the discharge channel. The plasma concentration exhibits characteristics that suggest the presence of a helical plasma column. This helical behavior has been observed in previous experiments that characterize plasma kink mode instabilities indirectly. Therefore, the direct imaging of these plasma kink modes further supports the link between MPDT onset behavior and the excitation of the magnetohydrodynamic instabilities.

Walker, Jonathan A.↗
NASA NTRS2013

Single and Multi-Pulse Low-Energy Conical Theta Pinch Inductive Pulsed Plasma Thruster Performance

Fabricated and tested CTP IPPTs at cone angles of 20deg, 38deg, and 60deg, and performed direct single-pulse impulse bit measurements with continuous gas flow. Single pulse performance highest for 38deg angle with impulse bit of approx.1 mN-s for both argon and xenon. Estimated efficiencies low, but not unexpectedly so based on historical data trends and the direction of the force vector in the CTP. Capacitor charging system assembled to provide rapid recharging of capacitor bank, permitting repetition-rate operation. IPPT operated at repetition-rate of 5 Hz, at maximum average power of 2.5 kW, representing to our knowledge the highest average power for a repetitively-pulsed thruster. Average thrust in repetition-rate mode (at 5 kV, 75 sccm argon) was greater than simply multiplying the single-pulse impulse bit and the repetition rate.

Hallock, Ashley K.↗
NASA NTRS2012

Effect of Inductive Coil Geometry on the Thrust Efficiency of a Microwave Assisted Discharge Inductive Plasma Accelerator

Pulsed inductive plasma thrusters [1-3] are spacecraft propulsion devices in which electrical energy is capacitively stored and then discharged through an inductive coil. The thruster is electrodeless, with a time-varying current in the coil interacting with a plasma covering the face of the coil to induce a plasma current. Propellant is accelerated and expelled at a high exhaust velocity (O(10-100 km/s)) by the Lorentz body force arising from the interaction of the magnetic field and the induced plasma current. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, pulsed inductive plasma thrusters require high pulse energies to inductively ionize propellant. The Microwave Assisted Discharge Inductive Plasma Accelerator (MAD-IPA) [4, 5] is a pulsed inductive plasma thruster that addressees this issue by partially ionizing propellant inside a conical inductive coil via an electron cyclotron resonance (ECR) discharge. The ECR plasma is produced using microwaves and permanent magnets that are arranged to create a thin resonance region along the inner surface of the coil, restricting plasma formation, and in turn current sheet formation, to a region where the magnetic coupling between the plasma and the inductive coil is high. The use of a conical theta-pinch coil is under investigation. The conical geometry serves to provide neutral propellant containment and plasma plume focusing that is improved relative to the more common planar geometry of the Pulsed Inductive Thruster (PIT) [2, 3], however a conical coil imparts a direct radial acceleration of the current sheet that serves to rapidly decouple the propellant from the coil, limiting the direct axial electromagnetic acceleration in favor of an indirect acceleration mechanism that requires significant heating of the propellant within the volume bounded by the current sheet. In this paper, we describe thrust stand measurements performed to characterize the performance (specific impulse, thrust efficiency) of the MAD-IPA thruster. Impulse data are obtained at various pulse energies, mass flow rates and inductive coil. geometries. Dependencies on these experimental parameters are discussed in the context of the current sheet formation and electromagnetic plasma acceleration processes.

Hallock, Ashley↗
NASA NTRS2011

Demonstration of a Fiber Optic Regression Probe in a High-Temperature Flow

The capability to provide localized, real-time monitoring of material regression rates in various applications has the potential to provide a new stream of data for development testing of various components and systems, as well as serving as a monitoring tool in flight applications. These applications include, but are not limited to, the regression of a combusting solid fuel surface, the ablation of the throat in a chemical rocket or the heat shield of an aeroshell, and the monitoring of erosion in long-life plasma thrusters. The rate of regression in the first application is very fast, while the second and third are increasingly slower. A recent fundamental sensor development effort has led to a novel regression, erosion, and ablation sensor technology (REAST). The REAST sensor allows for measurement of real-time surface erosion rates at a discrete surface location. The sensor is optical, using two different, co-located fiber-optics to perform the regression measurement. The disparate optical transmission properties of the two fiber-optics makes it possible to measure the regression rate by monitoring the relative light attenuation through the fibers. As the fibers regress along with the parent material in which they are embedded, the relative light intensities through the two fibers changes, providing a measure of the regression rate. The optical nature of the system makes it relatively easy to use in a variety of harsh, high temperature environments, and it is also unaffected by the presence of electric and magnetic fields. In addition, the sensor could be used to perform optical spectroscopy on the light emitted by a process and collected by fibers, giving localized measurements of various properties. The capability to perform an in-situ measurement of material regression rates is useful in addressing a variety of physical issues in various applications. An in-situ measurement allows for real-time data regarding the erosion rates, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over an operating envelope could also be useful in the modeling detailed physical processes. The sensor has been embedded in many regressing media to demonstrate the capabilities in a number of regressing environments. In the present work, sensors were installed in the eroding/regressing throat region of a converging-diverging flow, with the working gas heated to high temperatures by means of a high-pressure arc discharge at steady-state discharge power levels up to 500 kW. The amount of regression observed in each material sample was quantified using a later profilometer, which was compared to the in-situ erosion measurements to demonstrate the efficacy of the measurement technique in very harsh, high-temperature environments.

Korman, Valentin↗
NASA NTRS2011

Performance Testing of a Liquid Metal Pump for In-Space Power Systems

Fission surface power (FSP) systems could be used to provide power on the surface of the moon, Mars, or other planets and moons of our solar system. Fission power systems could provide excellent performance at any location, including those near the poles or other permanently shaded regions, and offer the capability to provide on demand power at any time, even at large distances from the sun. Fission-based systems also offer the potential for outposts, crew and science instruments to operate in a power-rich environment. NASA has been exploring technologies with the goal of reducing the cost and technical risk of employing FSP systems. A reference 40 kWe option has been devised that is cost-competitive with alternatives while providing more power for less mass anywhere on the lunar surface. The reference FSP system is also readily extensible for use on Mars, where it would be capable of operating through global dust storms and providing year-round power at any Martian latitude. Detailed development of the FSP concept and the reference mission are documented in various other reports. The development discussed in this paper prepares the way for testing of the Technology Demonstration Unit (TDU), which is a 10 kWe end-to-end test of FSP technologies intended to raise the entire FSP system to technology readiness level (TRL) 6. The Early Flight Fission Test Facility (EFF-TF) was established by NASA s Marshall Space Flight Center (MSFC) to provide a capability for performing hardware-directed activities to support multiple in-space nuclear reactor concepts by using a nonnuclear test methodology. This includes fabrication and testing at both the module/component level and at near prototypic reactor components and configurations allowing for realistic thermal-hydraulic evaluations of systems. The liquid-metal pump associated with the FSP system must be compatible with the liquid NaK coolant and have adequate performance to enable a viable flight system. Idaho National Laboratory (INL) was tasked with the modeling, design, and fabrication of an ALIP suitable for the FSP reference mission. A prototypic ALIP was fabricated under the direction of INL and shipped to MSFC for inclusion in the Technology Demonstration Unit (TDU), a quarter-scale end-to-end reactor simulator system that is scheduled for testing at NASA-GRC. Before inclusion in the TDU, the ALIP was tested in the ALIP test circuit (ATC), which is a rig developed and operated at MSFC for the specific purpose of providing accurate quantification of liquid metal pump performance. Data showing the pump performance curves (pressure, flowrate, and pump efficiency) are presented for various operating power levels, demonstrating the full performance envelope of the pump.

Polzin, Kurt↗
NASA NTRS2011

Local Leak Detection and Health Monitoring of Pressurized Tanks

An optical gas-detection sensor safely monitors pressurized systems (such as cryogenic tanks) and distribution systems for leaks. This sensor system is a fiber-coupled, solid optical body interferometer that allows for the miniaturized sensing element of the device to be placed in the smallest of recesses, and measures a wide range of gas species and densities (leaks). The deflection of the fringe pattern is detected and recorded to yield the time-varying gas density in the gap. This technology can be used by manufacturers or storage facilities with toxic, hazardous, or explosive gases. The approach is to monitor the change in the index of refraction associated with low-level gas leaks into a vacuum environment. The completion of this work will provide NASA with an enabling capability to detect gas system leaks in space, and to verify that pressurized systems are in a safe (i.e. non-leaking) condition during manned docking and transit operations. By recording the output of the sensor, a time-history of the leak can be constructed to indicate its severity. Project risk is mitigated by having several interferometric geometries and detection techniques available, each potentially leveraging hardware and lessons learned to enhance detectability.

Polzin, Kurt↗
NASA NTRS2010

Non-Nuclear Testing of Space Nuclear Systems at NASA MSFC

Highly realistic non-nuclear testing can be used to investigate and resolve potential issues with space nuclear power and propulsion systems. Non-nuclear testing is particularly useful for systems designed with fuels and materials operating within their demonstrated nuclear performance envelope. Non-nuclear testing allows thermal hydraulic, heat transfer, structural, integration, safety, operational, performance, and other potential issues to be investigated and resolved with a greater degree of flexibility and at reduced cost and schedule compared to nuclear testing. The primary limit of non-nuclear testing is that nuclear characteristics and potential nuclear issues cannot be directly investigated. However, non-nuclear testing can be used to augment the potential benefit from any nuclear testing that may be required for space nuclear system design and development. This paper describes previous and ongoing non-nuclear testing related to space nuclear systems at NASA's Marshall Space Flight Center (MSFC).

Houts, Michael G.↗
NASA NTRS2009

In-Situ Measurement of Hall Thruster Erosion Using a Fiber Optic Regression Probe

One potential life-limiting mechanism in a Hall thruster is the erosion of the ceramic material comprising the discharge channel. This is especially true for missions that require long thrusting periods and can be problematic for lifetime qualification, especially when attempting to qualify a thruster by analysis rather than a test lasting the full duration of the mission. In addition to lifetime, several analytical and numerical models include electrode erosion as a mechanism contributing to enhanced transport properties. However, there is still a great deal of dispute over the importance of erosion to transport in Hall thrusters. The capability to perform an in-situ measurement of discharge channel erosion is useful in addressing both the lifetime and transport concerns. An in-situ measurement would allow for real-time data regarding the erosion rates at different operating points, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over a thruster s operating envelope would also be useful in the modeling of the detailed physics inside the discharge chamber. There are many different sensors and techniques that have been employed to quantify discharge channel erosion in Hall thrusters. Snapshots of the wear pattern can be obtained at regular shutdown intervals using laser profilometry. Many non-intrusive techniques of varying complexity and sensitivity have been employed to detect the time-varying presence of erosion products in the thruster plume. These include the use quartz crystal microbalances, emission spectroscopy, laser induced flourescence, and cavity ring-down spectroscopy. While these techniques can provide a very accurate picture of the level of eroded material in the thruster plume, it is more difficult to use them to determine the location from which the material was eroded. Furthermore, none of the methods cited provide a true in-situ measure of erosion at the channel surface while the thruster is in operation (i.e. none yield a continuous channel erosion measurement). A recent fundamental sensor development effort has led to a novel regression, erosion, and ablation sensor technology (REAST). The REAST sensor allows for measurement of real-time surface erosion rates at a discrete surface location. The sensor was tested using a linear Hall thruster geometry (see Fig. 1), which served as a means of producing plasma erosion of a ceramic discharge chamber. The mass flow rate, discharge voltage, and applied magnetic field strength could be varied, allowing for erosion measurements over a broad thruster operating envelope. Results are presented demonstrating the ability of the REAST sensor to capture not only the insulator erosion rates but also changes in these rates as a function of the discharge parameters.

Polzin, Kurt↗