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DOE OSTI2017

Multi-species impurity granule injection and mass deposition projections in NSTX-U discharges

Here, by employing a neutral gas shielding (NGS) model to characterize impurity granule injection, the ablation rates for three different species of granule: lithium, boron, and carbon, are determined. Utilizing the duration of ablation events recorded on experiments performed at DIII-D to calibrate the NGS model, we quantify the ablation rate with respect to the plasma density profile. The species-specific granule shielding constant is then used to model granule ablation within NSTX-U discharges. Simulations of 300, 500 and 700 micron diameter granules injected at 50 m s –1 are presented for NSTX-U L-mode type plasmas, as well as H-mode discharges with low natural ELM frequency. Additionally, ablation calculations of 500 micron granules of each species are presented at velocities ranging from 50–150 m s –1 . In H-mode discharges these simulations show that the majority of the injected granule is ablated within or just past the edge steep gradient region. At this radial position, the perturbation to the background plasma generated by the ablating granule can lead to conditions advantageous for the rapid triggering of ELM crashes.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

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↗

DOE OSTI2019

Does Drug-Eluting Bead TACE Enhance the Local Effect of IRE? Imaging and Histopathological Evaluation in a Porcine Model

Objectives: We conducted an in vivo trial on swine to compare the ablation volumes of irreversible electroporation (IRE) followed by drug-eluting beads transarterial chemoembolization (DEB-TACE) versus IRE only. Materials and Methods: Nine swine underwent CT-guided IRE in one liver lobe and IRE immediately followed by DEB-TACE in a different liver lobe. For DEB-TACE, 100–300 µm beads (DC-Beads{sup ®}) were loaded with 50 mg doxorubicin. For IRE, the NanoKnife{sup ®} was used employing two electrodes according to the vendor’s protocol. Imaging follow-up was performed including CT-based lesion volume assessment using contrast-enhanced CT (venous phase) on days 1, 3, and 7 after the procedure. Three animals were killed for histopathological analysis after each follow-up. Results: Ablation volumes in CT in the IRE + DEB-TACE group were 15.4 ± 10.5 ml on day 1, 8.7 ± 5.6 ml on day 3, and 1.6 ± 0.7 ml on day 7. In the IRE group, the corresponding values were 5.2 ± 5.2 ml on day 1, 1.0 ± 1.2 ml on day 3, and 0.1 ± 0.1 ml on day 7. On day 1 and day 3, ablation volumes of IRE + TACE group were significantly larger than in the IRE group (p < 0.05). 96% of beads were depicted in or around ablative lesions. 69% of these beads were found in the surrounding hemorrhagic infiltration and 31% within the ablative lesion itself. Conclusions: Combination of IRE immediately followed by DEB-TACE resulted in larger ablation volumes compared to IRE alone, suggesting that local efficacy of IRE can be enhanced by post-IRE DEB-TACE.

62 RADIOLOGY AND NUCLEAR MEDICINE↗

DOE OSTI2015

Spherical strong-shock generation for shock-ignition inertial fusion

Recent experiments on the Laboratory for Laser Energetics' OMEGA laser have been carried out to produce strong shocks in solid spherical targets with direct laser illumination. The shocks are launched at pressures of several hundred Mbars and reach Gbar upon convergence. The results are relevant to the validation of the shock-ignition scheme and to the development of an OMEGA experimental platform to study material properties at Gbar pressures. The experiments investigate the strength of the ablation pressure and the hot-electron production at incident laser intensities of ∼2 to 6 × 10{sup 15 }W/cm{sup 2} and demonstrate ablation pressures exceeding 300 Mbar, which is crucial to developing a shock-ignition target design for the National Ignition Facility. The timing of the x-ray flash from shock convergence in the center of the solid plastic target is used to infer the ablation and shock pressures. Laser–plasma instabilities produce hot-electrons with a moderate temperature (<100 keV). The instantaneous conversion efficiencies of laser power into hot-electron power reached up to ∼15% in the intensity spike. The large amount of hot electrons is correlated with an earlier x-ray flash and a strong increase in its magnitude. This suggests that hot electrons contribute to the augmentation of the shock strength.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

DOE OSTI2016

Performance of indirectly driven capsule implosions on the National Ignition Facility using adiabat-shaping

A series of indirectly driven capsule implosions has been performed on the National Ignition Facility to assess the relative contributions of ablation-front instability growth vs. fuel compression on implosion performance. Laser pulse shapes for both low and high-foot pulses were modified to vary ablation-front growth and fuel adiabat, separately and controllably. Three principal conclusions are drawn from this study: (1) It is shown that reducing ablation-front instability growth in low-foot implosions results in a substantial (3-10X) increase in neutron yield with no loss of fuel compression. (2) It is shown that reducing the fuel adiabat in high-foot implosions results in a significant (36%) increase in fuel compression together with a small (10%) increase in neutron yield. (3) Increased electron preheat at higher laser power in high-foot implosions, however, appears to offset the gain in compression achieved by adiabat-shaping at lower power. These results taken collectively bridge the space between the higher compression low-foot results and the higher yield high-foot results.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

DOE OSTI2013

Real-time x-ray fluoroscopy-based catheter detection and tracking for cardiac electrophysiology interventions

Purpose: X-ray fluoroscopically guided cardiac electrophysiology (EP) procedures are commonly carried out to treat patients with arrhythmias. X-ray images have poor soft tissue contrast and, for this reason, overlay of a three-dimensional (3D) roadmap derived from preprocedural volumetric images can be used to add anatomical information. It is useful to know the position of the catheter electrodes relative to the cardiac anatomy, for example, to record ablation therapy locations during atrial fibrillation therapy. Also, the electrode positions of the coronary sinus (CS) catheter or lasso catheter can be used for road map motion correction.Methods: In this paper, the authors present a novel unified computational framework for image-based catheter detection and tracking without any user interaction. The proposed framework includes fast blob detection, shape-constrained searching and model-based detection. In addition, catheter tracking methods were designed based on the customized catheter models input from the detection method. Three real-time detection and tracking methods are derived from the computational framework to detect or track the three most common types of catheters in EP procedures: the ablation catheter, the CS catheter, and the lasso catheter. Since the proposed methods use the same blob detection method to extract key information from x-ray images, the ablation, CS, and lasso catheters can be detected and tracked simultaneously in real-time.Results: The catheter detection methods were tested on 105 different clinical fluoroscopy sequences taken from 31 clinical procedures. Two-dimensional (2D) detection errors of 0.50 {+-} 0.29, 0.92 {+-} 0.61, and 0.63 {+-} 0.45 mm as well as success rates of 99.4%, 97.2%, and 88.9% were achieved for the CS catheter, ablation catheter, and lasso catheter, respectively. With the tracking method, accuracies were increased to 0.45 {+-} 0.28, 0.64 {+-} 0.37, and 0.53 {+-} 0.38 mm and success rates increased to 100%, 99.2%, and 96.5% for the CS, ablation, and lasso catheters, respectively. Subjective clinical evaluation by three experienced electrophysiologists showed that the detection and tracking results were clinically acceptable.Conclusions: The proposed detection and tracking methods are automatic and can detect and track CS, ablation, and lasso catheters simultaneously and in real-time. The accuracy of the proposed methods is sub-mm and the methods are robust toward low-dose x-ray fluoroscopic images, which are mainly used during EP procedures to maintain low radiation dose.

62 RADIOLOGY AND NUCLEAR MEDICINE↗

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 OSTI2019

Percutaneous Irreversible Electroporation (IRE) of Hepatic Malignancy: A Bi-institutional Analysis of Safety and Outcomes

Aim: Irreversible electroporation (IRE) is a non-thermal ablative option in patients unsuitable for standard thermal ablation, due to its potential to preserve collagenous structures (vessels and ducts) and a reduced susceptibility to heat sink effects. In this series from two large tertiary referral hepatobiliary centres, we aim to assess the safety/outcomes of hepatic IRE. Materials and Methods: Bi-institutional retrospective, longitudinal follow-up series of IRE for primary hepatic malignancy; [hepatocellular carcinoma (n = 20), cholangiocarcinoma (n = 3)] and secondary metastatic disease; colorectal (n = 28), neuroendocrine (n = 1), pancreatic (n = 1), breast (n = 1), gastrointestinal stromal tumour (GIST, n = 1) and malignant thymoma (n = 1). Outcome measures included procedural safety/effectiveness, time to progression and time to death. Results: Between 2013 and 2017, 52 patients underwent percutaneous IRE of 59 liver tumours in 53 sessions. All tumours were deemed unsuitable for thermal ablation. Cases were performed using ultrasound (US) or computed tomography (CT) guidance. A complete ablation was achieved in n = 44, (75%) of cases with an overall complication rate of 17% (n = 9). Of the complete ablation group, median time to progression was 8 months. At 12 months, 44% were progression-free (95% CI 30–66%). The data suggest that larger lesion size (> 2 cm) is associated with shorter time to progression and there is highly significant difference with faster time to progression in mCRC compared with HCC. Median survival time was 38 months. Conclusion: This bi-institutional review is the largest UK series of IRE and suggests this ablative technology can be a useful tool, but appears to mainly induce local tumour control rather than cure with HCC having better outcomes than mCRC.

62 RADIOLOGY AND NUCLEAR MEDICINE↗