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306 records · Page 17
All sourcesNASA NTRSDOE OSTI
DOE OSTI2022

Silver Nanowire-Indium Zinc Oxide Composite Flexible Transparent Conducting Electrodes Made by Spin- coating and Photonic Curing

Realizing high-throughput, low-cost perovskite solar cell (PSC) manufacturing is highly sought-after in photovoltaic (PV) research in recent years. To fully achieve roll-to-roll (R2R) manufacturing of PSCs, it is important to consider the flexible transparent electrode (TE). PET/ITO is a commonly used substrate for making flexible PSCs. When optimizing transparent conducting materials, there is a tradeoff between sheet resistance (Rsh) and optical transparency. Because commercial PET/ITO substrates are made with slow (~1 m/min) vacuum deposition processes, they tend to be expensive. Therefore, it would be advantageous to develop a high-throughput, R2R compatible, solution-deposition approach for fabricating the TE on PET substrates. While various solution-deposition processes, such as blade coating or slot-die coating, can achieve the desired web speed of > 10 m/min, there is still a need to improve the post-deposition annealing step. One promising post-deposition processing technique is intense-pulsed-light processing, also known as photonic curing. Photonic curing delivers short (0.01 – 100 ms) pulses of broadband (200 – 1500 nm) light from a xenon flash lamp to the samples. Any materials in the sample stack that absorb light will convert the impinging light pulse into heat within the sample, which drives changes in the sample (calcination, phase change, crystallization, etc.). Photonic curing has three main advantages over thermal annealing: 1. Faster processing speed (milliseconds or seconds). 2. Compatibility with plastic substrates. 3. Smaller physical footprint and less wasted energy. Since the light pulses are on for a short time, the intensity can be high while the total energy delivered to the sample is low, minimizing damages to the plastic substrates. In this work, a hybrid TE material is fabricated on PET substrates using photonic curing. The hybrid TE material contains a layer of silver nanowires (AgNWs) and a layer of metal-oxide (InOx, ITO, IZO, etc.). The AgNWs increase the light absorbed by the film during the photonic curing process, which leads to higher processing temperatures, possibly improving the conversion of the metal-oxide layer. The AgNWs also enhance the electrical conductivity of the final TE layer after photonic curing. A AgNW and metal-oxide bilayer is formed by spin coating each solution onto the PET substrate sequentially followed by a single photonic curing process. We use average optical transmittance (Tavg) from 400 to 700 nm and average Rsh to evaluate the TE performance. The following photonic curing parameters are varied to optimize Tavg (maximize) and Rsh (minimize): Pulse voltage, pulse envelope, number of micro-pulses, duty cycle, number of pulses, and pulse repetition rate. Preliminarily, we also observe a significant impact on the TE properties by the volume of AgNW deposited during the spin coating deposition step. Using dispense volumes of 80 µL and 20 µL, we achieve samples with Tavg = 73%, Rsh = 19 Ω/sq, and roughness = 9 nm, and Tavg = 83%, Rsh = 58 Ω/sq, and roughness = 5.6 nm, respectively, after photonic curing.

14 SOLAR ENERGY↗
NASA NTRS1981

Nucleocosmochronology

Nucleocosmochronology is analyzed on the basis of elemental isotopic abundances of radionuclides. The production and depletion mechanisms for the cosmological chronometers are investigated and model-independent theories are reviewed. Emphasis is given to the significance of the mean age of the elements and the time resolution of the last nucleosynthetic events contributing to the solar system, and it is shown how the intermediate-lived chronometers may give information on the time dependence shape of the production function. Anomalous nucleocosmochronology is investigated and galactic and cosmological constraints are described in terms of time scales for the mean age of elements and galaxies and the age of the universe. Data resulting from both model-independent and dependent analyses are compared. Results based on the mean age for the r-process elements suggest a lower limit to the age of the galaxy and the universe. It is concluded that nucleocosmochronology, when coupled with other independent determinations for the age of the universe, defines a concordant universe age of 13.5-15.5 billion years. The shape of the production function is not well determined because of the uncertainties in the xenon data.

Symbalisty, E. M. D.↗
NASA NTRS1994

The NASA-JPL advanced propulsion program

The NASA Advanced Propulsion Concepts (APC) program at the Jet Propulsion Laboratory (JPL) consists of two main areas: The first involves cooperative modeling and research activities between JPL and various universities and industry; the second involves research at universities and industry that is directly supported by JPL. The cooperative research program consists of mission studies, research and development of ion engine technology using C-60 (Buckminsterfullerene) propellant, and research and development of lithium-propellant Lorentz-force accelerator (LFA) engine technology. The university/industry- supported research includes research (modeling and proof-of-concept experiments) in advanced, long-life electric propulsion, and in fusion propulsion. These propulsion concepts were selected primarily to cover a range of applications from near-term to far-term missions. For example, the long-lived pulsed-xenon thruster research that JPL is supporting at Princeton University addresses the near-term need for efficient, long-life attitude control and station-keeping propulsion for Earth-orbiting spacecraft. The C-60-propellant ion engine has the potential for good efficiency in a relatively low specific impulse (Isp) range (10,000 - 30,000 m/s) that is optimum for relatively fast (less than 100 day) cis-lunar (LEO/GEO/Lunar) missions employing near-term, high-specific mass electric propulsion vehicles. Research and modeling on the C-60-ion engine are currently being performed by JPL (engine demonstration), Caltech (C-60 properties), MIT (plume modeling), and USC (diagnostics). The Li-propellant LFA engine also has good efficiency in the modest Isp range (40,000 - 50,000 m/s) that is optimum for near-to-mid-term megawatt-class solar- and nuclear-electric propulsion vehicles used for Mars missions transporting cargo (in support of a piloted mission). Research and modeling on the Li-LFA engine are currently being performed by JPL (cathode development), Moscow Aviation Institute (engine testing), Thermacore (electrode development), as well as at MIT (plume modeling), and USC (diagnostics). Also, the mission performance of a nuclear-electric propulsion (NEP) Li-LFA Mars cargo vehicle is being modeled by JPL (mission analysis; thruster and power processor modeling) and the Rocketdyne Energy Technology and Engineering Center (ETEC) (power system modeling). Finally, the fusion propulsion research activities that JPL is supporting at Pennsylvania State University (PSU) and at Lawrenceville Plasma Physics (LPP) are aimed at far-term fast (less than 100 day round trip) piloted Mars missions and, in the very far term, interstellar missions.

Frisbee, Robert H.↗
NASA NTRS2004

SEP Mission to Titan NEXT Aerocapture In-Space Propulsion (Quicktime Movie)

The ion thruster is one of the most promising solar electric propulsion (SEP) technologies to support future Outer Planet missions (place provided link below here) for NASA's Office of Space Science. Typically, ion thrusters are used in high Isp- low thrust applications that require long lifetimes, as well as, higher efficiency over state-of-the-art chemical propulsion systems.Today, the standard for ion thrusters is the SEP Technology Application Readiness (NSTAR) thruster. Jet Propulsion Laboratory's (JPL's) extended life test (ELT) of the DS 1 flight spare NSTAR thruster began in October 1998. This test successfully demonstrated lifetime of the NSTAR flight spare thruster, which will provide a solid basis for selection of ion thrusters for future Code S missions. The NSTAR ELT was concluded on June 30,2003 after 30,352 hours. The purpose of the Next Generation Ion (NGI) activities is to advance Ion propulsion system technologies through the development of NASA's Evolutionary Xenon Thruster (NEXT). The goal of NEXT is to more than double the power capability and lifetime throughput (the total amount of propellant which can be processed) while increasing the Isp by 30% and the thrust by 120%.

Baggett, Randy↗
NASA NTRS2002

Ion Engine and Hall Thruster Development at the NASA Glenn Research Center

NASA's Glenn Research Center has been selected to lead development of NASA's Evolutionary Xenon Thruster (NEXT) system. The central feature of the NEXT system is an electric propulsion thruster (EPT) that inherits the knowledge gained through the NSTAR thruster that successfully propelled Deep Space 1 to asteroid Braille and comet Borrelly, while significantly increasing the thruster power level and making improvements in performance parameters associated with NSTAR. The EPT concept under development has a 40 cm beam diameter, twice the effective area of the Deep-Space 1 thruster, while maintaining a relatively-small volume. It incorporates mechanical features and operating conditions to maximize the design heritage established by the flight NSTAR 30 cm engine, while incorporating new technology where warranted to extend the power and throughput capability. The NASA Hall thruster program currently supports a number of tasks related to high power thruster development for a number of customers including the Energetics Program (formerly called the Space-based Program), the Space Solar Power Program, and the In-space Propulsion Program. In program year 2002, two tasks were central to the NASA Hall thruster program: 1.) the development of a laboratory Hall thruster capable of providing high thrust at high power; 2.) investigations into operation of Hall thrusters at high specific impulse. In addition to these two primary thruster development activities, there are a number of other on-going activities supported by the NASA Hall thruster program, These additional activities are related to issues such as thruster lifetime and spacecraft integration.

Domonkos, Matthew T.↗
NASA NTRS2014

Conventional and Bimodal Nuclear Thermal Rocket (NTR) Artificial Gravity Mars Transfer Vehicle Concepts

A variety of countermeasures have been developed to address the debilitating physiological effects of "zero-gravity" (0-g) experienced by cosmonauts and astronauts during their approximately 0.5-1.2 year long stays in LEO (Low Earth Orbit). Longer interplanetary flights, combined with possible prolonged stays in Mars orbit, could subject crewmembers to up to approximately 2.5 years of weightlessness. In view of known and recently diagnosed problems associated with 0-g, an artificial gravity spacecraft offers many advantages and may indeed be an enabling technology for human flights to Mars. A number of important human factors must be taken into account in selecting the rotation radius, rotation rate, and orientation of the habitation module or modules. These factors include the gravity gradient effect, radial and tangential Coriolis forces, along with cross-coupled acceleration effects. Artificial gravity (AG) Mars transfer vehicle (MTV) concepts are presented that utilize both conventional NTR, as well as, enhanced "bimodal" nuclear thermal rocket (BNTR) propulsion. The NTR is a proven technology that generates high thrust and has a specific impulse (I (sub sp)) capability of approximately 900 s - twice that of today's best chemical rockets. The AG/MTV concepts using conventional NTP carry twin cylindrical "ISS-type" habitation modules with their long axes oriented either perpendicular or parallel to the longitudinal spin axis of the MTV and utilize photovoltaic arrays (PVAs) for spacecraft power. The twin habitat modules are connected to a central operations hub located at the front of the MTV via two pressurized tunnels that provide the rotation radius for the habitat modules. For the BNTR AG/MTV option, each engine has its own "closed" secondary helium-xenon gas loop and Brayton rotating unit that can generate tens of kilowatts (kW (sub e)) of spacecraft electrical power during the mission coast phase eliminating the need for large PVAs. A single inflatable "TransHab-type" habitation module is also used with multiple vertical floors oriented radial to the MTV spin axis. The BNTR MTV's geometry - long and linear - is naturally compatible with AG operation. By rotating the vehicle about its center-of-mass and perpendicular to its flight vector at approximately 3.0 - 5.2 rpm, a centrifugal force and AG environment corresponding to approximately 0.38 - 1.0 g can be established to help maintain crew fitness out to Mars and back. Vehicles using NTP/BNTP can more readily accommodate the heavier payload mass and increased RCS propellant loading associated with AG operation, and can travel faster to and from Mars thereby reducing the crew's exposure to galactic cosmic radiation and solar flares. Mission scenario descriptions, key vehicle features and operational characteristics for each propulsion options are presented using the lift capability and payload volumes estimated for the SLS-1A and HLV.

Nuclear Propulsion↗
NASA NTRS2017

VIII. New Zr IV-VII, Xe IV-V, and Xe VII Oscillator Strengths and the Al, Zr, and Xe Abundances in the Hot White Dwarfs G191-B2B and RE0503-289

Context: For the spectral analysis of high-resolution and high-signal-to-noise spectra of hot stars, state-of-the-art non-local thermodynamic equilibrium (NLTE) model atmospheres are mandatory. These are strongly dependent on the reliability of the atomic data that is used for their calculation. Aims: To search for zirconium and xenon lines in the ultraviolet (UV) spectra of G191−B2B and RE 0503−289, new Zr iv-vii, Xe iv-v, and Xe vii oscillator strengths were calculated. This allows, for the first time, determination of the Zr abundance in white dwarf (WD) stars and improvement of the Xe abundance determinations.Methods: We calculated Zr iv-vii, Xe iv-v, and Xe vii oscillator strengths to consider radiative and collisional bound-bound transitions of Zr and Xe in our NLTE stellar-atmosphere models for the analysis of their lines exhibited in UV observations of the hot WDs G191−B2B and RE 0503−289. Results: We identified one new Zr iv, 14 new Zr v, and ten new Zr vi lines in the spectrum of RE 0503−289. Zr was detected for the first time in a WD. We measured a Zr abundance of −3.5 +/- 0.2 (logarithmic mass fraction, approx. 11 500 times solar). We identified five new Xe vi lines and determined a Xe abundance of −3.9 +/- 0.2 (approx. 7500 times solar). We determined a preliminary photospheric Al abundance of −4.3 +/- 0.2 (solar) in RE 0503−289. In the spectra of G191−B2B, no Zr line was identified. The strongest Zr iv line (1598.948 Å) in our model gave an upper limit of −5.6 +/- 0.3 (approx. 100 times solar). No Xe line was identified in the UV spectrum of G191−B2B and we confirmed the previously determined upper limit of −6.8 +/- 0.3 (ten times solar). Conclusions: Precise measurements and calculations of atomic data are a prerequisite for advanced NLTE stellar-atmosphere modeling. Observed Zr iv-vi and Xe vi-vii line profiles in the UV spectrum of RE 0503−289 were simultaneously well reproduced with our newly calculated oscillator strengths.

Rauch, T.↗
NASA NTRS2016

Conventional and Bimodal Nuclear Thermal Rocket (NTR) Artificial Gravity Mars Transfer Vehicle Concepts

A variety of countermeasures have been developed to address the debilitating physiological effects of zero-gravity (0-g) experienced by cosmonauts and astronauts during their approximately 0.5 to 1.2 year long stays in low Earth orbit (LEO). Longer interplanetary flights, combined with possible prolonged stays in Mars orbit, could subject crewmembers to up to approximately 2.5 years of weightlessness. In view of known and recently diagnosed problems associated with 0-g, an artificial gravity (AG) spacecraft offers many advantages and may indeed be an enabling technology for human flights to Mars. A number of important human factors must be taken into account in selecting the rotation radius, rotation rate, and orientation of the habitation module or modules. These factors include the gravity gradient effect, radial and tangential Coriolis forces, along with cross-coupled acceleration effects. Artificial gravity Mars transfer vehicle (MTV) concepts are presented that utilize both conventional NTR, as well as, enhanced bimodal nuclear thermal rocket (BNTR) propulsion. The NTR is a proven technology that generates high thrust and has a specific impulse (Isp) capability of approximately 900 s-twice that of today's best chemical rockets. The AG/MTV concepts using conventional Nuclear Thermal Propulsion (NTP) carry twin cylindrical International Space Station (ISS)- type habitation modules with their long axes oriented either perpendicular or parallel to the longitudinal spin axis of the MTV and utilize photovoltaic arrays (PVAs) for spacecraft power. The twin habitat modules are connected to a central operations hub located at the front of the MTV via two pressurized tunnels that provide the rotation radius for the habitat modules. For the BNTR AG/MTV option, each engine has its own closed secondary helium(He)-xenon (Xe) gas loop and Brayton Rotating Unit (BRU) that can generate 10s of kilowatts (kWe) of spacecraft electrical power during the mission coast phase eliminating the need for large PVAs. A single inflatable TransHab-type habitation module is also used with multiple vertical floors oriented radial to the MTV spin axis. The BNTR MTV's geometry-long and linear-is naturally compatible with AG operation. By rotating the vehicle about its center-of-mass (CM) and perpendicular to its flight vector at approximately 3.0 to 5.2 rpm, a centrifugal force and AG environment corresponding to approximately 0.38 to 1.0 g can be established to help maintain crew fitness out to Mars and back. Vehicles using NTP/ Bimodal Nuclear Thermal Propulsion (BNTP) can more readily accommodate the heavier payload mass and increased RCS propellant loading associated with AG operation, and can travel faster to and from Mars thereby reducing the crew's exposure to galactic cosmic radiation and solar flares. Mission scenario descriptions, key vehicle features and operational characteristics for each propulsion option are presented using the lift capability and payload volumes estimated for the Space Launch System (SLS)-1B and followon Heavy Lift Vehicle (HLV).

Nuclear Propulsion↗
NASA NTRS1992

Electric propulsion system technology

The work performed in fiscal year (FY) 1991 under the Propulsion Technology Program RTOP (Research and Technology Objectives and Plans) No. (55) 506-42-31 for Low-Thrust Primary and Auxiliary Propulsion technology development is described. The objectives of this work fall under two broad categories. The first of these deals with the development of ion engines for primary propulsion in support of solar system exploration. The second with the advancement of steady-state magnetoplasmadynamic (MPD) thruster technology at 100 kW to multimegawatt input power levels. The major technology issues for ion propulsion are demonstration of adequate engine life at the 5 to 10 kW power level and scaling ion engines to power levels of tens to hundreds of kilowatts. Tests of a new technique in which the decelerator grid of a three-grid ion accelerator system is biased negative of neutralizer common potential in order to collect facility induced charge-exchange ions are described. These tests indicate that this SAND (Screen, Accelerator, Negative Decelerator) configuration may enable long duration ion engine endurance tests to be performed at vacuum chamber pressures an order of magnitude higher than previously possible. The corresponding reduction in pumping speed requirements enables endurance tests of 10 kW class ion engines to be performed within the resources of existing technology programs. The results of a successful 5,000-hr endurance of a xenon hollow cathode operating at an emission current of 25 A are described, as well as the initial tests of hollow cathodes operating on a mixture of argon and 3 percent nitrogen. Work performed on the development of carbon/carbon grids, a multi-orifice hollow cathode, and discharge chamber erosion reduction through the addition of nitrogen are also described. Critical applied-field MPD thruster technical issues remain to be resolved, including demonstration of reliable steady-state operation at input powers of hundreds to thousands of kilowatts, achievement of thruster efficiency and specific impulse levels required for missions of interest, and demonstration of adequate engine life at these input power, efficiency, and specific impulse levels. To address these issues we have designed, built, and tested a 100 kW class, radiation-cooled applied-field MPD thruster and a unique dual-beam thrust stand that enables separate measurements of the applied- and self-field thrust components. We have also initiated the development of cathode thermal and plasma sheath models that will eventually be used to guide the experimental program. In conjunction with the cathode modeling, a new cathode test facility is being constructed. This facility will support the study of cathode thermal behavior and erosion mechanisms, the diagnosis of the near-cathode plasma and the development and endurance testing of new, high-current cathode designs. To facilitate understanding of electrode surface phenomenon, we have implemented a telephoto technique to obtain photographs of the electrodes during engine operation. In order to reduce the background vacuum tank pressure during steady-state engine operation in order to obtain high fidelity anode thermal data, we have developed and are evaluating a gas-dynamic diffuser. A review of experience with alkali metal propellants for MPD thrusters led to the conclusion that alkali metals, particularly lithium, offer the potential for significant engine performance and lifetime improvements. These propellants are also condensible at room temperature, substantially reducing test facility pumping requirements. The most significant systems-level issue is the potential for spacecraft contamination. Subsequent experimental and theoretical efforts should be directed toward verifying the performance and lifetime gains and characterizing the thruster flow field to assess its impact on spacecraft surfaces. Consequently, we have begun the design and development of a new facility to study engine operation with alkali metal propellants.

Brophy, John R.↗
NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100hrs to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne (AR) is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗
NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100 kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5 kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100 hours to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗
NASA NTRS2017

Fusion-Enabled Pluto Orbiter and Lander

The Pluto orbiter mission proposed here is credible and exciting. The benefits to this and all outer-planet and interstellar-probe missions are difficult to overstate. The enabling technology, Direct Fusion Drive, is a unique fusion engine concept based on the Princeton Field-Reversed Configuration (PFRC) fusion reactor under development at the Princeton Plasma Physics Laboratory. The truly game-changing levels of thrust and power in a modestly sized package could integrate with our current launch infrastructure while radically expanding the science capability of these missions. During this Phase I effort, we made great strides in modeling the engine efficiency, thrust, and specific impulse and analyzing feasible trajectories. Based on 2D fluid modeling of the fusion reactors outer stratum, its scrape-off-layer (SOL), we estimate achieving 2.5 to 5 N of thrust for each megawatt of fusion power, reaching a specific impulse, Isp, of about 10,000 s. Supporting this model are particle-in-cell calculations of energy transfer from the fusion products to the SOL electrons. Subsequently, this energy is transferred to the ions as they expand through the magnetic nozzle and beyond. Our point solution for the Pluto mission now delivers 1000 kg of payload to Pluto orbit in 3.75 years using 7.5 N constant thrust. This could potentially be achieved with a single 1 MW engine. The departure spiral from Earth orbit and insertion spiral to Pluto orbit require only a small portion of the total delta-V. Departing from low Earth orbit reduces mission cost while increasing available mission mass. The payload includes a lander, which utilizes a standard green propellant engine for the landing sequence. The lander has about 4 square meters of solar panels mounted on a gimbal that allows it to track the orbiter, which beams 30 to 50 kW of power using a 1080 nm laser. Optical communication provides dramatically high data rates back to Earth. Our mass modeling investigations revealed that if current high-temperature superconductors are utilized at liquid nitrogen temperatures, they drive the mass of the engine, partly because of the shielding required to maintain their critical temperature. Second generation materials are thinner but the superconductor is a very thin layer deposited on a substrate with additional layers of metallic classing. Tremendous research is being performed on a variety of these superconducting materials, and new irradiation data is now available. This raises the possibility of operating nearfuture high-temperature superconductors at a moderately low temperature to dramatically reduce the amount of shielding required. At the same time, a first generation space engine may require low-temperature superconductors, which are higher TRL and have been designed for space coils before (AMS-02 experiment for the ISS). We performed detailed analysis of the startup system and thermal conversion system components. The ideal working fluid was determined to be a blend of Helium and Xenon. No significant problems were identified with these subsystems. For the RF system, we conceived of a new, more efficient design using state-of-the-art switch amplifiers, which have the potential for 100% efficiency. This report presents details of our engine and trajectory analyses, mass modeling efforts, and updated vehicle designs.

temperature↗
NASA NTRS2004

Chondrites and the Protoplanetary Disk, Part 3

Contents include the following: Ca-, Al-Rich Inclusions and Ameoboid Olivine Aggregates: What We Know and Don t Know About Their Origin. Aluminium-26 and Oxygen Isotopic Distributions of Ca-Al-rich Inclusions from Acfer 214 CH Chondrite. The Trapping Efficiency of Helium in Fullerene and Its Implicatiion to the Planetary Science. Constraints on the Origin of Chondritic Components from Oxygen Isotopic Compositions. Role of Planetary Impacts in Thermal Processing of Chondrite Materials. Formation of the Melilite Mantle of the Type B1 CAIs: Flash Heating or Transport? The Iodine-Xenon System in Outer and Inner Portions of Chondrules from the Unnamed Antarctic LL3 Chondrite. Nucleosynthesis of Short-lived Radioactivities in Massive Stars. The Two-Fluid Analysis of the Kelvin-Helmholtz Instability in the Dust Layer of a Protoplanetary Disk: A Possible Path to the Planetesimal Formation Through the Gravitational Instability. Shock-Wave Heating Model for Chonodrule Formation: Heating Rate and Cooling Rate Constraints. Glycine Amide Hydrolysis with Water and OH Radical: A Comparative DFT Study. Micron-sized Sample Preparation for AFM and SEM. AFM, FE-SEM and Optical Imaging of a Shocked L/LL Chondrite: Implications for Martensite Formation and Wave Propagation. Infrared Spectroscopy of Chondrites and Their Components: A Link Between Meteoritics and Astronomy? Mid-Infrared Spectroscopy of CAI and Their Mineral Components. The Origin of Iron Isotope Fractionation in Chondrules, CAIs and Matrix from Allende (CV3) and Chainpur (LL3) Chondrites. Protoplanetary Disk Evolution: Early Results from Spitzer. Kinetics of Evaporation-Condensation in a Melt-Solid System and Its Role on the Chemical Composition and Evolution of Chondrules. Oxygen Isotope Exchange Recorded Within Anorthite Single Crystal in Vigarano CAI: Evidence for Remelting by High Temperature Process in the Solar Nebula. Chondrule Forming Shock Waves in Solar Nebula by X-Ray Flares. Organic Globules with Anormalous Nitrogen Isotopic Compositions in the Tagish Lake Meteorite: Products of Primitive Organic Reactions. Yet Another Chondrule Formation Scenario. CAIs are Not Supernova Condensates. Microcrystals and Amorphous Material in Comets and Primitive Meteorites: Keys to Understanding Processes in the Early Solar System. A Nearby Supernova Injected Short-lived Radionuclides into Our Protoplanetary Disk. REE+Y Systematics in CC and UOC Chondrules. Meteoritic Constraints on Temperatures, Pressures, Cooling Rates, Chemical Compositions, and Modes of Condensation in the Solar Nebula. The I-Xe Record of Long Equilibration in Chondrules from the Unnamed Antarctic Meteorite L3/LL3. Early Stellar Evolution.

Source record↗
NASA NTRS2004

Lunar and Planetary Science XXXV: Achondrite Mishmash

The session "Achondrite Mishmash" included teh following presentations:Constraining the Formation and Evolution of IAB Irons - High Precision; 40Ar/39Ar Ages on Plagioclase Separates from Silicate Inclusions of the Campo Del Cielo Meteorite; Radiogenic 129-Xenon in Silicate Inclusions in the Campo Del Cielo Iron Meteorite; Significance of Iron Isotope Mineral Fractionation in Pallasite and Iron Meteorites; An Experimental Study of Phosphoran Olivine and Its Significance in Main Group Pallasites; Characterization of the Distribution of Siderophile and Highly Siderophile Elements in the Milton and Eagle Station Pallasites; Relationships Between HED and Mesosiderite Meteorites: An Iron Isotope Perspective; Production Rates for Noble-Gas Isotopes in Eucrites; Evidence for Subsolidus Metasomatism in the Eucrite Parent Body; Possible Contact Metamorphism of the Polymict Eucrite Petersburg; Early and Late Stage Metals and Sulfides in Diogenites; Constraints on the Lithological Variation near the Surface of the HED Planetoid from the Petrology of 91 & 92 Series Antarctic Achondrites; Cosmic Ray Exposure Ages, Ar-Ar Ages, and the Origin and History of Eucrites; Trace Elements Abundances Vs Closure Temperature in Orthopyroxenes from Howardites; The First Mesosiderite-like Clast in a Howardite; 39Ar-40Ar Dating of Unusual Eucrite NWA 011: Is it from Vesta?; Trace Element Systematics of Northwest Africa 011: A "Eucritic" Basalt from a Non-Eucrite Parent Body ; 33S Anomaly in Acapulcoites and Lodranites; Magnetic Force Microscopy of Primitive Achondrites; Experimental Constraints on Ureilite Petrogenesis; Reflectance Spectra of Ureilites: Nature of the Mafic Silicate Absorption Features; Nitrogen and Noble Gases in Two Monomict Ureilites Acfer 277 and FRO 90036 from Hot and Cold Deserts; Solar Noble Gases in the Angrite Parent Body - Evidence from Volcanic Volatiles Trapped in D'Orbigny Glass; Trace Element Distribution Between Phases of the D'Orbigny Angrite ; Petrological Study of Six Angrites Including New Type; and Manganese-Chromium Isotope Systematics of Basaltic Achondrite Northwest Africa 011.

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NASA NTRS1999

Cathodes Delivered for Space Station Plasma Contactor System

The International Space Station's (ISS) power system is designed with high-voltage solar arrays that typically operate at output voltages of 140 to 160 volts (V). The ISS grounding scheme electrically ties the habitat modules, structure, and radiators to the negative tap of the solar arrays. Without some active charge control method, this electrical configuration and the plasma current balance would cause the habitat modules, structure, and radiators to float to voltages as large as -120 V with respect to the ambient space plasma. With such large negative floating potentials, the ISS could have deleterious interactions with the space plasma. These interactions could include arcing through insulating surfaces and sputtering of conductive surfaces as ions are accelerated by the spacecraft plasma sheath. A plasma contactor system was baselined on the ISS to prevent arcing and sputtering. The sole requirement for the system is contained within a single directive (SSP 30000, paragraph 3.1.3.2.1.8): "The Space Station structure floating potential at all points on the Space Station shall be controlled to within 40 V of the ionospheric plasma potential using a plasma contactor." NASA is developing this plasma contactor as part of the ISS electrical power system. For ISS, efficient and rapid emission of high electron currents is required from the plasma contactor system under conditions of variable and uncertain current demand. A hollow cathode plasma source is well suited for this application and was, therefore, selected as the design approach for the station plasma contactor system. In addition to the plasma source, which is referred to as a hollow cathode assembly, or HCA, the plasma contactor system includes two other subsystems. These are the power electronics unit and the xenon gas feed system. The Rocketdyne Division of Boeing North American is responsible for the design, fabrication, assembly, test, and integration of the plasma contactor system. Because of technical and schedule considerations, the NASA Lewis Research Center was asked to manufacture and deliver the engineering model, the qualification model, and the flight HCA units for the plasma contactor system as government furnished equipment. To date, multiple units have been built. One cathode has demonstrated approximately 28 000-hr lifetime, two development HCA units have demonstrated over 15 000-hr lifetime, and one HCA unit has demonstrated more than 38 000 ignitions. All eight flight HCA's have been manufactured, acceptance tested, and are ready for delivery to the flight contractor.

Patterson, Michael J.↗
NASA NTRS2011

Extra-Zodiacal-Cloud Astronomy via Solar Electric Propulsion

Solar electric propulsion (SEP) is often considered as primary propulsion for robotic planetary missions, providing the opportunity to deliver more payload mass to difficult, high-delta-velocity destinations. However, SEP application to astrophysics has not been well studied. This research identifies and assesses a new application of SEP as primary propulsion for low-cost high-performance robotic astrophysics missions. The performance of an optical/infrared space observatory in Earth orbit or at the Sun-Earth L2 point (SEL2) is limited by background emission from the Zodiacal dust cloud that has a disk morphology along the ecliptic plane. By delivering an observatory to a inclined heliocentric orbit, most of this background emission can be avoided, resulting in a very substantial increase in science performance. This advantage enabled by SEP allows a small-aperture telescope to rival the performance of much larger telescopes located at SEL2. In this paper, we describe a novel mission architecture in which SEP technology is used to enable unprecedented telescope sensitivity performance per unit collecting area. This extra-zodiacal mission architecture will enable a new class of high-performance, short-development time, Explorer missions whose sensitivity and survey speed can rival flagship-class SEL2 facilities, thus providing new programmatic flexibility for NASA's astronomy mission portfolio. A mission concept study was conducted to evaluate this application of SEP. Trajectory analyses determined that a 700 kg-class science payload could be delivered in just over 2 years to a 2 AU mission orbit inclined 15 to the ecliptic using a 13 kW-class NASA's Evolutionary Xenon Thruster (NEXT) SEP system. A mission architecture trade resulted in a SEP stage architecture, in which the science spacecraft separates from the stage after delivery to the mission orbit. The SEP stage and science spacecraft concepts were defined in collaborative engineering environment studies. The SEP stage architecture approach offers benefits beyond a single astrophysics mission. A variety of low-cost astrophysics missions could employ a standard SEP stage to achieve substantial science benefit. This paper describes the results of this study in detail, including trajectory analysis, spacecraft concept definition, description of telescope/instrument benefits, and application of the resulting SEP stage to other missions. In addition, the benefits of cooperative development and use of the SEP stage, in conjunction with a SEP flight demonstration mission currently in definition at NASA, are considered.

Benson, Scott W.↗
DOE OSTI2022

A dry ancient plume mantle from noble gas isotopes

Primordial volatiles were delivered to terrestrial reservoirs during Earth’s accretion, and the mantle plume source is thought to have retained a greater proportion of primordial volatiles compared with the upper mantle. This study shows that mantle He, Ne, and Xe isotopes require that the plume mantle had low concentrations of volatiles like Xe and H 2 O at the end of accretion compared with the upper mantle. A lower extent of mantle processing alone is not sufficient to explain plume noble gas signatures. Ratios of primordial isotopes are used to determine proportions of solar, chondritic, and regassed atmospheric volatiles in the plume mantle and upper mantle. The regassed Ne flux exceeds the regassed Xe flux but has a small impact on the mantle Ne budget. Pairing primordial isotopes with radiogenic systems gives an absolute concentration of 130 Xe in the plume source of ∼1.5 × 10 7 atoms 130 Xe/g at the end of accretion, ∼4 times less than that determined for the ancient upper mantle. A record of limited accretion of volatile-rich solids thus survives in the He-Ne-Xe signatures of mantle rocks today. A primordial viscosity contrast originating from a factor of ∼4 to ∼250 times lower H 2 O concentration in the plume mantle compared with the upper mantle may explain (a) why giant impacts that triggered whole mantle magma oceans did not homogenize the growing planet, (b) why the plume mantle has experienced less processing by partial melting over Earth’s history, and (c) how early-formed isotopic heterogeneities may have survived ∼4.5 Gy of solid-state mantle convection.

58 GEOSCIENCES↗
DOE OSTI2024

New constraints on ultraheavy dark matter from the LZ experiment

Searches for dark matter with liquid xenon time projection chamber experiments have traditionally focused on the region of the parameter space that is characteristic of weakly interacting massive particles, ranging from a few GeV / c 2 to a few TeV / c 2 . Models of dark matter with a mass much heavier than this are well motivated by early production mechanisms different from the standard thermal freeze-out, but they have generally been less explored experimentally. In this work, we present a reanalysis of the first science run of the LZ experiment, with an exposure of 0.9 tonne × yr , to search for ultraheavy particle dark matter. The signal topology consists of multiple energy deposits in the active region of the detector forming a straight line, from which the velocity of the incoming particle can be reconstructed on an event-by-event basis. Zero events with this topology were observed after applying the data selection calibrated on a simulated sample of signal-like events. New experimental constraints are derived, which rule out previously unexplored regions of the dark matter parameter space of spin-independent interactions beyond a mass of 10 17 GeV / c 2 . Published by the American Physical Society 2024

72 PHYSICS OF ELEMENTARY PARTICLES AND FIELDS↗