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NASA NTRS2018

Tandem LA-ICP-MS & LIBS; A New Micro-Analytical Technique for the Measurement of Every Element in the Periodic Table

Laser Ablation ICP-MS (LA-ICP-MS) has been widely accepted as a microanalytical technique for in-situ trace (ppb) elemental analysis on the micron scale in a variety of geologic materials. LA-ICP-MS (single or multi-collector) provides both elemental and isotopic measurements critical for a wide range of geological research by generating a fine grained aerosol (nm scale) during the laser ablation event and delivering that aerosol to the ICP ion source of the mass spectrometer via an inert carrier gas. LA-ICP-MS, however, suffers from limitations in analyzing high ionization potential elements as well as elements subject to atmospheric and argon based interferences. LA-ICP-MS also has limitations in analyzing major elements due to detector saturation. An alternative laser ablation technique, Laser Induced Breakdown Spectroscopy (LIBS), employs an optical spectrometer integrated into the laser ablation system that analyzes the laser induced plasma at the sample surface across the entire optical spectrum for emission lines of every element in the periodic table. Elements that are difficult or impossible to measure with LA-ICP-MS are now possible to analyze with LIBS down to low ppm levels with CCD and/or ICCD detection. We introduce a new laser based technique, “Tandem LA-LIBS”, that combines LA for ICP-MS and LIBS into one integrated laser ablation system. This system has the effect of expanding the elemental coverage and the dynamic range of the laser ablation experiment as measurements from ppb to % level matrix elements can now be analyzed in a single ablation experiment. We present both femtosecond and nanosecond Tandem LA-LIBS quantitative and qualitative data on wide range of geological materials for those elements that are difficult or impossible by traditional LA-ICP-MS techniques such as F, H, O, N, C, S, halogens, etc. We also demonstrate that the simultaneous measurement of trace, minor and major elements are now possible in a single laser ablation experiment with Tandem LA-LIBS technology.

LA-ICP-MS↗

NASA NTRSDate not supplied

Comparison between Hayabusa 2 Spectral Measurements and Simulations

Recent Hayabusa 2 spectral measurements are compared with state-of-the-art shock-layer radiation simulations resulting from the LAURA/HARA code suite. These simulations include coupled ablation, which accounts for the injection of ablation products into the flowfield, and coupled radiation, which accounts for radiative energy loss in the flowfield. To enable the cou- pled ablation simulations, a best-estimate model is developed for Hayabusa’s carbon-phenolic ablator based on available published information. The comparison between the simulations and measurements focuses on three atomic nitrogen lines, two atomic oxygen lines, one atomic hydrogen line, and the CN Violet band system. Parametric uncertainties are evaluated for these features using a recently developed capability in LAURA/HARA. These uncertainties capture the impact of flowfield and radiation modeling uncertainties on the observed radiation simulations, and include the impact of coupled radiation and ablation product emission. The resulting parametric uncertainties range from 50 to over 100% for the atomic lines, and 80% for CN Violet. Uncertainties in flowfield kinetics and non-Boltzmann modeling provide the dominant contribution to these uncertainties. For the atomic nitrogen and oxygen lines, com- parisons between the nominal simulations and measurements show agreement well within these simulation uncertainty bounds, with the measurements within 20% of the nominal simulation over most of the trajectory, and within 10% at peak emission. This excellent agreement is un- precedented for observed radiation measurements of atomic lines, where previous Stardust and Hayabusa 1 comparisons were significantly worse. The improved agreement for these Hayabusa 2 comparisons is both the result of improved measurement quality and enhanced flowfield and radiation modeling. For the ablation product emission from the 656 nm atomic hydrogen line, the agreement between measurements and simulations is within ±20% for all trajectory points except one outlier. This good comparison is surprising considering this species depends on the ablation rate and is shown to have strong wake contribution, which both contribute to the large ≈±70% parametric uncertainty. For the ablation product emission from CN Violet, the simulations over-predict the measurements by over a factor of two early in the trajectory, with the comparison improving later in the trajectory. This disagreement for CN Violet is consistent with the ≈±80% parametric uncertainty bounds evaluated for the simulations. To utilize the excellent agreement between the atomic line simulations and measurements in the assessment of the radiative heating margin for a future flight vehicle, the flowfield property binning (FPB) approach is applied. Using the Mars Sample Return (MSR) Earth Entry System (EES) as an example, the combination of the excellent Hayabusa comparisons and the FPB analysis justify nearly a 10% reduction in the radiative heating margin from an assumed baseline margin of 30%. This direct quantitative link between observed radiation measurements and the radia- tive heating margin of future flight vehicles demonstrates the value of these observed radiation measurements.

Shock-Layer Radiative Heating↗

NASA NTRS2001

Roughness-Dominated Transition on Nosetips, Attachment Lines and Lifting-Entry Vehicles

Modeling of roughness-dominated transition is a critical design issue for both ablating and non-ablating thermal protection systems (TPS). Ablating TPS, used for planetary-entry and earth-return missions, first experience recession under high-altitude, low-Reynolds-number conditions. Such laminar-flow ablation causes the formation of a surface microroughness pattern characteristic of the TPS material composition and fabrication process. For non-ablating TPS, such as the overlapping-tile, metallic heatshields proposed for future reusable launch vehicles, the surface roughness pattern is established a priori by the engineering design and assembly procedure. In both cases, these distributed surface roughness patterns create disturbances within, and alter the mean velocity profile of, the laminar boundary layer flowing over the surface. As altitude decreases, Reynolds number increases, and flow field conditions capable of amplifying these roughness-induced perturbations are eventually achieved, i.e., transition onset occurs. Boundary layer transition to turbulence results in more severe heat-transfer rates. Ablating TPS experience increased recession rates, leading to potential bum-through, while non-ablating TPS experience accelerated temperature rise, leading to potential melting of key components.

Reda, Daniel C.↗

NASA NTRS1995

Ames Research Center Mars/Pathfinder Heat Shield Design Verification ARC-JET Test

Design verification tests were performed on samples representing the aerobrake of the Mars/Pathfinder vehicle. The test specimens consisted of the SLA-561V ablator bonded to the honeycomb structure. The primary objective was to evaluate the ablation materials performance and to measure temperatures within the ablator, at the structural bondline and at the back sheet of the honeycomb structure. Other objectives were to evaluate the effect of ablative repair plug material treatment and voids in the heat shield. A total of 29 models were provided for testing in the Ames 60MW arc-jet facility. Of these, 23 models were flat-faced and six remaining models were curved edge ones, intended to simulate the conditions on the curved rim of the forebody where the maximum shear occurred. Eight sets of test conditions were used. The stagnation point heating rates varied from 47 to 240 W/cm2 and the stagnation pressures from 0.15 to 0.27 atm. (The maximum flight values are 132 W/cm2 and 0.25 atm) The majority of these runs were made at a nominal stagnation pressure of 0.25 atm. Two higher pressure runs were made to check the current (denser) ablation material for spallation, or other forms of thermal stress failure. Over 60% of the flatfaced models yielded good thermocouple data and all produced useful surface recession information. Of the five curved-edge models that were tested, only one gave good data; the remaining ones experienced model-holder failure. The test results can be summarized by noting that no failure of the ablative material was observed on any model. Also, the bondline temperature design limit of 250 C was never reached within an equivalent flight time despite a stagnation point heat load that exceeded the maximum flight value by up to 130%. At heating rates of over 200W/cm2 and stagnation pressures of 0.25 atm, or greater, the average surface recessions exceeded 0.5 cm on some models. The surface roughness increased dramatically at pressures above 0.25 atm and was four times greater at 0.27 atm than at 0.25 atm. Procured repair plug material performed much better than room-temperature cured plugs, as observed in the previous tests. Voids in the ablator did not increase local temperatures and gaps did not grow during testing.

Tran, Huy K.↗

NASA NTRS2007

Reentry Thermal Analysis of a Generic Crew Exploration Vehicle Structure

Comparative studies were performed on the heat-shielding characteristics of honeycomb-core sandwich panels fabricated with different materials for possible use as wall panels for the proposed crew exploration vehicle. Graphite/epoxy sandwich panel was found to outperform aluminum sandwich panel under the same geometry due to superior heat-shielding qualities and lower material density. Also, representative reentry heat-transfer analysis was performed on the windward wall structures of a generic crew exploration vehicle. The Apollo low Earth orbit reentry trajectory was used to calculate the reentry heating rates. The generic crew exploration vehicle has a graphite/epoxy composite honeycomb sandwich exterior wall and an aluminum honeycomb sandwich interior wall, and is protected with the Apollo thermal protection system ablative material. In the thermal analysis computer program used, the TPS ablation effect was not yet included; however, the results from the nonablation heat-transfer analyses were used to develop a "virtual ablation" method to estimate the ablation heat loads and the thermal protection system recession thicknesses. Depending on the severity of the heating-rate time history, the virtual ablation period was found to last for 87 to 107 seconds and the ablation heat load was estimated to be in the range of 86 to 88 percent of the total heat load for the ablation time period. The thermal protection system recession thickness was estimated to be in the range of 0.08 to 0.11 inches. For the crew exploration vehicle zero-tilt and 18-degree-tilt stagnation points, thermal protection system thicknesses of h = {0.717, 0.733} inches were found to be adequate to keep the substructural composite sandwich temperature below the limit of 300 F.

Ko, William L.↗

NASA NTRS2011

Processes to Open the Container and the Sample Catcher of the Hayabusa Returned Capsule in the Planetary Material Sample Curation Facility of JAXA

Japanese spacecraft Hayabusa, which returned from near-Earth-asteroid Itokawa, successfully returned its reentry capsule to the Earth, the Woomera Prohibited Area in Australia in Jun 13th, 2010, as detailed in another paper [1]. The capsule introduced into the Planetary Material Sample Curation Facility in the Sagamihara campus of JAXA in the early morning of June 18th. Hereafter, we describe a series of processes for the returned capsule and the container to recover gas and materials in there. A transportation box of the recovered capsule was cleaned up on its outer surface beforehand and introduced into the class 10,000 clean room of the facility. Then, the capsule was extracted from the box and its plastic bag was opened and checked and photographed the outer surface of the capsule. The capsule was composed of the container, a backside ablator, a side ablator, an electronic box and a supporting frame. The container consists of an outer lid, an inner lid, a frame for latches, a container and a sample catcher, which is composed of room A and B and a rotational cylinder. After the first check, the capsule was packed in a plastic bag with N2 again, and transferred to the Chofu campus in JAXA, where the X-ray CT instrument is situated. The first X-ray CT analysis was performed on the whole returned capsule for confirming the conditions of latches and O-ring seal of the container. The analysis showed that the latches of the container should have worked normally, and that the double Orings of the container seemed to be sealed its sample catcher with no problem. After the first X-ray CT, the capsule was sent back to Sagamihara and introduced in the clean room to exclude the electronic box and the side ablator from the container by hand tools. Then the container with the backside ablator was set firmly to special jigs to fix the lid of container tightly to the container and set to a milling machine. The backside ablator was drilled by the machine to expose heads of bolts, which combined the ablator to the outer lid of the container, and after the drilling had been finished, all the bolts were unscrewed and the backside ablator was removed from the container. Then, the container was sent to the Chofu X-ray facility again to examine in detail by a micro X-ray CT instrument in order to reconfirm that the condition of the latches of the lid of container was normal and that its double O-ring seemed to have been sealed after the last X-ray CT analysis.

Fujimura, A.↗

NASA NTRS2013

Thermal Protection and Control

During all phases of a spacecraft's mission, a Thermal Protection System (TPS) is needed to protect the vehicle and structure from extreme temperatures and heating. When designing TPS, low weight and cost while ensuring the protection of the vehicle is highly desired. There are two main types of TPS, ablative and reusable. The Apollo missions needed ablators due to the high heat loads from lunar reentry. However, when the desire for a reusable space vehicle emerged, the resultant_ Space Shuttle program propelled a push for the development of reusable TPS. With the growth of reqsable TPS, the need for ablators declined, triggering a drop off of the ablator industry. As a result, the expertise was not heavily maintained within NASA or the industry. When the Orion Program initiated a few years back, a need. for an ablator reemerged. Yet, due to of the lack of industry capability, redeveloping the ablator material took several years and came at a high cost. As NASA looks towards the future with both the Orion and Commercial Crew Programs, a need to preserve reusable, ablative, and other TPS technologies is essential. Research of the different TPS materials alongside their properties, capabilities, and manufacturing process was performed, and the benefits of the materials were analyzed alongside the future of TPS. Knowledge of the different technologies has the ability to help us know what expertise to maintain and ensure a lack in the industry does not occur again.

Greene, Effie E.↗

NASA NTRS2018

Radiative Heating of Large Meteoroids During Atmospheric Entry

A high-fidelity approach for simulating the aerothermodynamic environments of meteor entries was developed, which allows the commonly assumed heat transfer coefficient of 0.1 to be assessed. This model uses chemically reacting computational fluid dynamics (CFD), coupled with radiation transport and surface ablation. Coupled radiation accounts for the impact of radiation on the flowfield energy equations, while coupled ablation explicitly models the injection of ablation products within the flowfield and radiation simulations. For a meteoroid with a velocity of 20 km/s, coupled radiation is shown to reduce the stagnation point radiative heating by over 60%. The impact of coupled ablation (with coupled radiation) is shown to provide at least a 70% reduction in the radiative heating relative to cases with only coupled radiation. This large reduction is partially the result of the low ionization energies of meteoric ablation products relative to air species. The low ionization energies of ablation products, such as Mg and Ca, provide strong photoionization and atomic line absorption in regions of the spectrum that air species do not. MgO and CaO are also shown to provide significant absorption. Turbulence is shown to impact the distribution of ablation products through the shock-layer, which results in up to a 100% increase in the radiative heating downstream of the stagnation point. To create a database of heat transfer coefficients, the developed model was applied to a range of cases. This database considered velocities ranging from 14 to 20 km/s, altitudes ranging from 20 to 50 km, and nose radii ranging from 1 to 100 m. The heat transfer coefficients from these simulations are below 0.045 for the range of cases, for both laminar and turbulent, which is significantly lower than the canonical value of 0:1. When the new heat transfer model is applied to a Tunguska-like 15 Mt entry, the effect of the new model is to lower the height of burst by up to 2 km, depending on assumed entry angle. This, in turn, results in a significantly larger ground damage footprint than when the canonical heating assumption is used.

Christopher O Johnston↗

NASA NTRS2015

Radiant Heating Tests on Meteoritic Material

A new effort geared toward modeling the physics of meteor entry and break-up is underway at NASA Ames Research Center. This is part of a broader interdisciplinary effort on providing physics-based risk assessment models for potentially hazardous objects. As part of the entry modeling task we are seeking to improve our understanding of, among other things, the ablation of meteoric material during high speed entry into earths atmosphere. Meteoroid entry differs greatly in some key respects from spacecraft entry modeling. First, the aerothermal environment at these high velocities (18 km/s) is dominated by radiation. Second, meteoroids less than, say, 50m in size, will likely lose a significant portion of their mass during the high-speed atmospheric entry process due to vaporization as well as melting and spallation. The mass of the object, in turn, directly affects the amount of energy deposited in the atmosphere, and therefore the amount of damage done. Thus it is important for us to understand and be able to model this process in greater detail in order to assess the hazard posed by these objects.In this presentation, we first give an overview of the simple ablation models that are typically used in meteor entry calculations. Additionally, we will present a new model which utilizes a similar approach to what is typically done for spacecraft TPS response modeling. This uses an equilibrium assumption near the surface to compute the ablation rate, as is done in heritage material response codes. Next we describe the radiant heating experiment which uses a high-powered laser to emulate the radiation dominated heating environment experienced by meteoroids during atmospheric entry. The facility the Laser Hardened Material Evaluation Laboratory (LHMEL) permits us to expose samples of meteoritic material to heating rates in excess of 100kW/sq.cm. An overview of the experimental set-up and test plan for the initial exploratory campaign at this facility will be given. Then we present both qualitative and quantitative results from this initial test series. Comparisons between predicted and measured ablation rates suggest that there may be significant blockage of the incident beam by the ablation plume. Furthermore, melt is shown to be a significant ablation mechanics, even at high heating rates (16 kW/sq.cm). Finally, comparisons between the phenomenology of the ablation of terrestrial rocks namely, basalt -- to that of meteorites show very different behavior. This is shown to likely be due in part to the effect of composition on the melt viscosity.

Stern, Eric↗

NASA NTRS2015

TPSAS-NF1676L-21180-DND

Ablative materials are used as the Thermal Protection System (TPS) for space vehicles that enter a planetary atmosphere. During the entry phase, the vehicle experiences high rates of aerodynamic heating. These ablative materials, acting as a heat shield, protect the vehicle from the harsh environment and ensure a safe entry. Ablative materials reduce the incoming heat flux through various phenomena such as near-surface oxidation, pyrolysis chemical reaction, vaporization, and other erosive processes, collectively known as ablation. Spallation is one such ablative process, in which solid particles are ejected from the material into the flow field. The many conditions leading to spallation, however, are not well understood and cannot be described accurately by simple phenomenological models. However, the spallation phenomenon might impact the aerodynamic heating rates of re-entry vehicles using ablative heat shields. For example, the solid particles ejected from the surface can modify the near surface chemical processes and flow field. To investigate spallation effects, a code was developed to compute the dynamics of spalled particles. The code uses a finite-rate chemistry model to study the chemical interactions of the particles with the flow field. The spallation code is dynamically coupled to a CFD solver that models the time accurate hypersonic flow field around an ablative sample. An experimental investigation was also performed at the NASA Langley HYMETS facility. HYMETS is a 400 kW arcjet that can simulate hypersonic flight and Earth or Martian entry conditions. Using high speed cameras, an IR camera, NIR and UV-VIS spectrometers, and other diagnostic instruments, test data on various graphite, LI-900, FiberForm, and PICA test samples were obtained, using target heat flux conditions of 100, 200, and 400 W/cm 2 . From the test campaign, measurements of particle velocity, acceleration and trajectory are obtained.

A Martin↗

NASA NTRS2016

Launch Pad in a Box

NASA Kennedy Space Center (KSC) is developing a new deployable launch system capability to support a small class of launch vehicles for NASA and commercial space companies to test and launch their vehicles. The deployable launch pad concept was first demonstrated on a smaller scale at KSC in 2012 in support of NASA Johnson Space Center's Morpheus Lander Project. The main objective of the Morpheus Project was to test a prototype planetary lander as a vertical takeoff and landing test-bed for advanced spacecraft technologies using a hazard field that KSC had constructed at the Shuttle Landing Facility (SLF). A steel pad for launch or landing was constructed using a modular design that allowed it to be reconfigurable and expandable. A steel flame trench was designed as an optional module that could be easily inserted in place of any modular steel plate component. The concept of a transportable modular launch and landing pad may also be applicable to planetary surfaces where the effects of rocket exhaust plume on surface regolith is problematic for hardware on the surface that may either be damaged by direct impact of high speed dust particles, or impaired by the accumulation of dust (e.g., solar array panels and thermal radiators). During the Morpheus free flight campaign in 2013-14, KSC performed two studies related to rocket plume effects. One study compared four different thermal ablatives that were applied to the interior of a steel flame trench that KSC had designed and built. The second study monitored the erosion of a concrete landing pad following each landing of the Morpheus vehicle on the same pad located in the hazard field. All surfaces of a portable flame trench that could be directly exposed to hot gas during launch of the Morpheus vehicle were coated with four types of ablatives. All ablative products had been tested by NASA KSC and/or the manufacturer. The ablative thicknesses were measured periodically following the twelve Morpheus free flight tests. The thermal energy from the Morpheus rocket exhaust plume was only found to be sufficient to cause appreciable ablation of one of the four ablatives that were tested. The rocket exhaust plume did cause spalling of concrete during each descent and landing on a landing pad in the hazard field. The Extended Abstract ASE Earth and Space Conference April, 2016 - Orlando, FL concrete surface was laser scanned following each Morpheus landing, and the total volume of spalled concrete that eroded between the first and final landings of the Morpheus Project's test campaign was estimated. This paper will also describe a new deployable launch system (DLS) capability that is being developed at KSC and was publicly announced in May 2015 (KSC Partnerships, 2015). The DLS is a set of multi-user Ground Support Equipment that will be used to test and launch small class launch vehicles. The system is comprised of four main elements: the Launch Stand, the Flame Deflector, the Pad Apron and the KAMAG transporter. The system elements are designed to be deployed at launch or test sites within the KSC/CCAFS boundaries. The DLS is intended to be used together with the Fluid and Electrical System of the Universal Propellant Servicing Systems and Mobile Power Data and Communications Unit.

Ablatives Testing↗

NASA NTRS2016

Launch Pad in a Box

Mobile Launch Pad Development↗

NASA NTRSDate not supplied

A Parallelized Oxidation-Driven Surface Recession Framework in DSMC Code, SPARTA

Spacecrafts rely on ablative thermal protection systems (TPS) made of composites consisting of a carbon-based reinforcement and a polymeric matrix. These materials are designed to withstand high-temperature oxidation and surface recession during re-entry into the Earth's atmosphere. However, ablation occurs due to a complex interplay of thermal, mechanical, and chemical factors, making it challenging to determine the individual impact of each on the TPS's overall degradation. In this study, we have developed an ablation model that can leverage a finite rate carbon oxidation model to predict material recession and surface states more accurately. Stochastic PArallel Rarified-gas Time-accurate Analyzer (SPARTA), a direct-simulation Monte Carlo (DSMC) code, is modified to allow oxidation-driven ablation of implicitly defined carbon surfaces. In SPARTA, implicit surfaces are generated from the grid corner point values via a marching cubes algorithm, therefore creating a new set of surface elements every time ablation is performed. The finite-rate oxidation model developed by Gopalan et. al can perform both gas-surface and pure-surface reactions and is now adapted to tally surface data on a per grid cell basis. The ablation functionality was also adjusted so once the reactions have occurred, the number of reactions leading to CO formation can be converted to corner point reduction values; therefore, carbon removal is directly proportional to surface recession. We also briefly discuss some unique challenges associated with parallelizing this dynamic surface state and geometry. Finally, we analyze the performance of this parallelized implicit chemistry model with simple 2D and 3D benchmark cases by producing surface state statistics, area changes over time, and visualization across a range of surface temperatures and processors with and without load-balancing.

DSMC↗

NASA NTRS1994

Trajectory-based heating analysis for the European Space Agency/Rosetta Earth Return Vehicle

A coupled, trajectory-based flowfield and material thermal-response analysis is presented for the European Space Agency proposed Rosetta comet nucleus sample return vehicle. The probe returns to earth along a hyperbolic trajectory with an entry velocity of 16.5 km/s and requires an ablative heat shield on the forebody. Combined radiative and convective ablating flowfield analyses were performed for the significant heating portion of the shallow ballistic entry trajectory. Both quasisteady ablation and fully transient analyses were performed for a heat shield composed of carbon-phenolic ablative material. Quasisteady analysis was performed using the two-dimensional axisymmetric codes RASLE and BLIMPK. Transient computational results were obtained from the one-dimensional ablation/conduction code CMA. Results are presented for heating, temperature, and ablation rate distributions over the probe forebody for various trajectory points. Comparison of transient and quasisteady results indicates that, for the heating pulse encountered by this probe, the quasisteady approach is conservative from the standpoint of predicted surface recession.

RADIATIVE ABLATING FLOWFIELD A↗

DOE OSTI2020

Runaway electron seed formation at reactor-relevant temperature

Systematic variation of the pre-disruption core electron temperature (T e ) from 1 to 12 keV using an internal transport barrier scenario reveals a dramatic increase in the production of ‘seed’ runaway electrons (REs), ultimately accessing near-complete conversion of the pre-disruption current into sub-MeV RE current. Injected Ar pellets are observed to ablate more intensely and promptly as T e rises. At high T e , the observed ablation exceeds predictions from published thermal ablation models. Simultaneously, the thermal quench (TQ) is observed to significantly shorten with increasing T e —a surprising result. While the reason for the shorter TQ is not yet understood, candidate mechanisms include: insufficiently accurate thermal ablation models, enhanced ablation driven by the seed RE population, or significant parallel heat transport along stochastic fields. Kinetic modeling that self-consistently treats the plasma cooling via radiation, the induced electric field, and the formation of the seed RE is performed. Including the combined effect of the inherent dependence of hot-tail RE seeding on T e together with the shortened TQ, modeling recovers the progression towards near-complete conversion of the pre-disruption current to RE current as T e rises. Measurement of the HXR spectrum during the early current quench (CQ) reveals a trend of decreasing energy with pre-disruption T e . At the very highest T e (≈ 12 keV), ≈ 100% conversion of the thermal current to runaway current is found. The energy of this peculiar RE beam is inferred to be sub-MeV as it emits vanishingly few MeV hard x-rays (HXRs). These measurements demonstrate novel TQ dynamics as Te is varied and illustrate the limitations of treating the RE seed formation problem without considering the inter-related dependencies of the pellet ablation, radiative energy loss, and resultant variations of the TQ duration. Finally, if the observed shortening of the TQ with increasing T e extends to fusion-grade plasmas, than their propensity to form large quantities of RE seeds at high T e may be far worse than previously thought. Positively, the high T e scenario in DIII-D produces REs so prodigiously that it can serve as a meaningful new platform for demonstrating RE avoidance techniques.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

NASA NTRS1990

Conditioned taste aversion and motion sickness in cats and squirrel monkeys

The relationship between vomiting and conditioned taste aversion was studied in intact cats and squirrel monkeys and in cats and squirrel monkeys in which the area postrema was ablated by thermal cautery. In cats conditioned 7-12 months after ablation of the area postrema, three successive treatments with xylazine failed to produce either vomiting or conditioned taste aversion to a novel fluid. Intact cats, however, vomited and formed a conditioned aversion. In squirrel monkeys conditioned 6 months after ablation of the area postrema, three treatments with lithium chloride failed to produce conditioned taste aversion. Intact monkeys did condition with these treatments. Neither intact nor ablated monkeys vomited or evidenced other signs of illness when injected with lithium chloride. When the same ablated cats and monkeys were exposed to a form of motion that produced vomiting prior to surgery, conditioned taste aversion can be produced after ablation of the area postrema. The utility of conditioned taste aversion as a measure of subemetic motion sickness is discussed by examining agreement and disagreement between identifications of motion sickness by conditioned taste aversion and vomiting. It is suggested that a convincing demonstration of the utility of conditioned taste aversion as a measure of nausea requires the identification of physiological correlates of nausea, and caution should be exercised when attempting to interpret conditioned taste aversion as a measure of nausea.

area postrema↗

NASA NTRS2007

Lunar Return Reentry Thermal Analysis of a Generic Crew Exploration Vehicle Wall Structures

Thermostructural analysis was performed on generic crew exploration vehicle (GCEV) heat shielded wall structures subjected to reentry heating rates based on five potential lunar return reentry trajectories. The GCEV windward outer wall is fabricated with a graphite/epoxy composite honeycomb sandwich panel and the inner wall with an aluminum honeycomb sandwich panel. The outer wall is protected with an ablative Avcoat-5026-39H/CG thermal protection system (TPS). A virtual ablation method (a graphical approximation) developed earlier was further extended, and was used to estimate the ablation periods, ablation heat loads, and the TPS recession layer depths. It was found that up to 83 95 percent of the total reentry heat load was dissipated in the TPS ablation process, leaving a small amount (3-15 percent) of the remaining total reentry heat load to heat the virgin TPS and maintain the TPS surface at the ablation temperature, 1,200 F. The GCEV stagnation point TPS recession layer depths were estimated to be in the range of 0.280-0.910 in, and the allowable minimum stagnation point TPS thicknesses that could maintain the substructural composite sandwich wall at the limit temperature of 300 F were found to be in the range of 0.767-1.538 in. Based on results from the present analyses, the lunar return abort ballistic reentry was found to be quite attractive because it required less TPS weight than the lunar return direct, the lunar return skipping, or the low Earth orbit guided reentry, and only 11.6 percent more TPS weight than the low Earth orbit ballistic reentry that will encounter a considerable weight penalty to obtain the Earth orbit. The analysis also showed that the TPS weight required for the lunar return skipping reentry was much more than the TPS weight necessary for any of the other reentry trajectories considered.

Ko, William L.↗

NASA NTRS2014

High-Precision Measurement of Eu/Eu* in Geological Glasses via LA-ICP-MS Analysis

Elemental fractionation during laser ablation inductively coupled plasma mass spectrometry (LA-ICP-MS) analysis has been historically documented between refractory and volatile elements. In this work, however, we observed fractionation between light rare earth elements (LREEs) and heavy rare earth elements (HREEs) when using ablation strategies involving large spot sizes (greater than 100 millimeters) and line scanning mode. In addition: (1) ion yields decrease when using spot sizes above 100 millimeters; (2) (Eu/Eu*)(sub raw) (i.e. Europium anomaly) positively correlates with carrier gas (He) flow rate, which provides control over the particle size distribution of the aerosol reaching the ICP; (3) (Eu/Eu*)(sub raw) shows a positive correlation with spot size, and (4) the changes in REE signal intensity, induced by the He flow rate change, roughly correlate with REE condensation temperatures. The REE fractionation is likely driven by the slight but significant difference in their condensation temperatures. Large particles may not be completely dissociated in the ICP and result in preferential evaporation of the less refractory LREEs and thus non-stoichiometric particle-ion conversion. This mechanism may also be responsible for Sm-Eu-Gd fractionation as Eu is less refractory than Sm and Gd. The extent of fractionation depends upon the particle size distribution of the aerosol, which in turn is influenced by the laser parameters and matrix. Ablation pits and lines defined by low aspect ratios produce a higher proportion of large particles than high aspect ratio ablation, as confirmed by measurements of particle size distribution in the laser induced aerosol. Therefore, low aspect ratio ablation introduces particles that cannot be decomposed and/or atomized by the ICP and thus results in exacerbated elemental fractionation. Accurate quantification of REE concentrations and Eu/Eu* requires reduction of large particle production during laser ablation. For the reference materials analyzed in this work, the 100 millimeters spot measurements of Eu/Eu* agreed with GeoRem preferred values within 3 percent. Our long-term analyses of Eu/Eu* in MPI-DING glass KL-2G and USGS glass BIR-1G were reproducible at 3 percent (2 RSD).

geological glasses↗