Paper Proceedings

Monday
9:30-11:30	Interdisciplinary-01	Abstract Presentation Paper	A Case Study of AI-assisted Creation of a Thermodynamics Model of Precipitation Formation During Rapid Depressurization of a Vented Container Dr. Jedediah Storey
	Interdisciplinary-02	Abstract Presentation Paper	AI/ML Frameworks for Microgravity Flow Boiling:  Identifying, Quantifying, and Generating Two-Phase Interfacial Features V.S. Devahdhanush and Arkadeep Paul
	Interdisciplinary-03	Abstract Presentation Paper	Techno-feasibility systems analysis of carbon dioxide and methane separation systems: From airborne carbon capture to ECLSS applications Dr. Ashwin Ravichandran, Pooja Santhamoorthy, Stephen Summits, Emmanuel Skountzos, Joakim H. Stenlid, David Mebane, Lyndsey McMillon-Brown, and John Lawson (VIRTUAL)
	Aerosciences-10	Abstract Presentation Paper	Euler-to-RANS Mapping of Shock-Interaction Regimes on Slender Biconic Forebodies at Mach 2–5 Aneesh Kukreti (VIRTUAL)
	Thermal Control and Protection-01	Abstract Presentation Paper	Design & Thermal Analysis of Hyperion: The University of Tennessee’s New 1 MW Continuous Arcjet Facility Zachary Colovos, Damiano Baccarella, Killan Samuels, Mitchell Trotsky, and Christian Isaacs
	Thermal Control and Protection-02	Abstract Presentation Paper	Advanced and Efficient Thermal Model Correlation of a Satellite Electronics Box Using HEEDS and Thermal Desktop Dr. John Pederson and Lina L. Maricic
	Thermal Control and Protection-03	Abstract Presentation Paper	Europa Clipper Thermal System Performance In-Flight vs. Predictions Hared Ochoa
	Thermal Control and Protection-04	Abstract Presentation Paper	FROSTE: Thermal Design and Thermal Vacuum Testing of Cryogenic Stowage for Lunar Sample Return Faiyaj Khan, Elijah Stewart, Claire Silaire, Kacy Vanden Bergh, and Erin Hayward
	Thermal Control and Protection-05	Abstract Presentation Paper	Development and Testing of Additively Manufactured Phase Change Material Thermal Storage Units Mike Pauken
1:00-3:00	Aerosciences-01	Abstract Presentation Paper	Towards Isentropic Design Using the Entropy Map Generated from 3-D CFD Results Dr. Bijay Sultanian
	Aerosciences-03	Abstract Presentation Paper	Investigation of Peltier-based Environmental Control Systems for Low-cost/Attritable EO/IR Pods on Transonic Platforms Mark Whittum
	Thermal Control and Protection-06	Abstract Presentation Paper	Structural Performance Evaluation of Titanium-Water Heat Pipe – GFRC Radiators for Fission Surface Power Sandeep Hatte, Zayed Ahmed, Mason Pratt, Calin Tarau, Jeffrey Diebold, Srujan Rokkam, Tyler Beach, and Jeffrey Hopkins
	Thermal Control and Protection-07	Abstract Presentation Paper	Fabrication and Testing of Additively Manufactured Titanium – Water Loop Heat Pipe for Space Applications Sandeep Hatte, Calin Tarau, Jeff Diebold, Srujan Rokkam, Linda Yan, and Ying Sun
	Thermal Control and Protection-08	Abstract Presentation Paper	Performance and Modeling Comparison Between Working Fluids for 3D Printed Loop Heat Pipes Jacob Sonnek, Lawrence Bradley, Philip Graybill, and Adam Shreve
	Thermal Control and Protection-09	Abstract Presentation Paper	Experimental Characterization of Additively Manufactured Nickel-Titanium Shape Memory Alloy Heat Pipes William Sixel, Bilal Bomani, Tomé S. Guenkax, Christopher Greer, and Ryan Overdorff
	Thermal Control and Protection-18	Abstract Presentation Paper	SRM Testbed for Thermal Testing of Internal Compnents Robert Frederick
	Interdisciplinary-21	Abstract Presentation Paper	Predicting Operational Performance of xEMU Boot at Lunar South Pole Temperatures using Thermal Desktop® Emma Quick and Noah Andersen (VIRTUAL)
	Interdisciplinary-06	Abstract Presentation Paper	Designing tank outlets to prevent vortex formation and minimize liquid residuals Mark Wollen
	Thermal Control and Protection-16	Abstract Presentation Paper	Thermal Design and Analysis of Green Propulsion Dual Mode CubeSat Savanna Lyles and Clark Teems
	Thermal Control and Protection-17	Abstract Presentation Paper	Variable Emittance Coating Systems Thermal Modeling for a 3U CubeSat Sarah Stewart, Sam Keller, Sydney Taylor, and Ognjen Ilic
3:30-5:00	Thermal Control and Protection-10	Abstract Presentation Paper	Variable Heat Rejection System for Space Habitats Dr. Calin Tarau, Ramy Abdelmaksoud, Jeff Diebold, Srujan Rokkam, William Johnson, and Justin Boyer
	Thermal Control and Protection-11	Abstract Presentation Paper	Multifunctional Pulsating Heat Pipes for Electric Motors Jeff Diebold, Jacob Sonnek, Parag Bajaj, and Arijit Banerjee
	Thermal Control and Protection-12	Abstract Presentation Paper	Recent Development of High-Temperature Alkali Metal Heat Pipes at ACT Jeff Diebold, Calin Tarau, Sndeep Hatte, Roopesh Kumar Nathan Van Velson, and David-Paul Schulze
	Aerosciences-04	Abstract Presentation Paper	Comparative Assessment of Hypersonic Air-Breathing Inlet Architectures Under Common Freestream Conditions Mookesh Dhanasar, Frederick Ferguson, Connor Ramaswamy, Leonard Uitenham, and Wisdom Calmday
	Aerosciences-05	Abstract Presentation Paper	Comparative Assessment of Inversely Designed Hypersonic Waverider Configurations Mookesh Dhansar, Frederick Ferguson, Connor Ramaswamy, Ajit Kelkar, and Donovan Berry
	Aerosciences-06	Abstract Presentation Paper	Three-dimensional, Multi-Phase Unsteady Flow of Water Undergoing  Multiple Flow Regimes (free surface, droplet, and fluid film) William Dziedzic
	Thermal Control and Protection-20	Abstract Presentation Paper	Thermal Design and Thermal Vacuum Testing of the StarBurst Instrument Stephanie Mauro
	Thermal Control and Protection-21	Abstract Presentation Paper	Developmental and Cryogenic Thermal Vacuum Testing Lessons Learned Mackenzie Byrnes and Heather Grimes
	Thermal Control and Protection-22	Abstract Presentation Paper	Thermal Performance of Flight Imagery Launch Monitoring Real-time System (FILMRS) and Enhance FILMRS (EFILMRS) camera systems for Artemis Debby Hernandez
	Interdisciplinary-07	Abstract Presentation Paper	Moon Base Transportation – Deliveries to the Lunar Surface Anthony Terracciano, Maxwell Martin, Charles J. Weyandt, John R. Campbell Jr., Jared Nardontonia, Richard Nederlander, Clyde O’Quinn, Ben Tutera, and Elaine Voll
	Interdisciplinary-08	Abstract Presentation Paper	Thermal Analysis of the Advanced Modular Power Systems(AMPS) Power Electronic Modules Hashmatullah Hasseeb and Firas G. Asfoor
	Interdisciplinary-09	Abstract Presentation Paper	Contact Line Pinning in Microgravity Joshua Whitehead, Lucas Hunt, and Ranga Narayanan

Tuesday
9:30-11:30	Cryogenics-01	Abstract Presentation Paper	A Scalable, Calibration-Free Multiphase Arbitrary Lagrangian Eulerian Solver for Sloshing Dynamics and Phase Change Dr. Antonio Cantiani
	Cryogenics-02	Abstract Presentation Paper	CFD-assisted nodal modeling of sloshing in a cryogenic propellant tank Dr. Alok Majumdar, Andre LeClair, Ramon L. Otero, and Jacob Brodnik
	Cryogenics-03	Abstract Presentation Paper	Cryogenic ullage pressure collapse via condensation and ullage heat loss Dr. Samuel Darr
	Thermal Control and Protection-23	Abstract Presentation Paper	SR-1 Freedom Thermal Architecture Challenges Scott Thomas
	Thermal Control and Protection-24	Abstract Presentation Paper	Power and Propulsion Element Steerable High Gain Antenna Lunar Transit Thermal Analysis Tracking Methodology Matthew Faykus
	Thermal Control and Protection-25	Abstract Presentation Paper	Thermal Design Considerations for Venus Orbiters: Lessons Learned from VenSAR Tyler Schmidt
	Thermal Control and Protection-26	Abstract Presentation Paper	Thermal Design of a FALCON Cloud Radar Instrument Concept Tyler Schmidt, Katarina Aguayo, and Rogelio Rosas
	Thermal Control and Protection-13	Abstract Presentation Paper	Architected Lattice Wicks for Monolithic Conformal Heat Pipes and Vapor Chambers—Characterization: Permeability, Capillary, Surface Area, Thermal, and Mechanical Dr. Robin Pham and Presented by Calin Tarau
	Thermal Control and Protection-14	Abstract Presentation Paper	Passive Thermal Switching Using Freeze–Thaw Behavior in Heat Pipes for Lunar Night Survival (Lunar Outpost) Joe Durante
	Aerosciences-07	Abstract Presentation Paper	Aerothermal Performance Enhancement of Supersonic TsAGI S-12 Airfoils Using Graphene Nanoplatelet Coatings: Experimental and CFD Investigations Lakshmanan Kasi, Rajalakshmi Padmanabhan, and Perarasu Thangavelu
	Aerosciences-09	Abstract Presentation Paper	High-Fidelity CFD Simulation of Liquid, Hybrid, and Solid Rocket Propulsion Systems Dr. Xiao Ren
	Aerosciences-11	Abstract Presentation Paper	SABRE: A Physics-Driven Framework For Satellite Re-Entry Demise Analysis Suyash Tandon
1:00-3:00	Cryogenics-04	Abstract Presentation Paper	Initial Computational Analysis of Helium Subsurface Pressurization Jacob Brodnick and Jason Hartwig
	Cryogenics-05	Abstract Presentation Paper	Modeling of Priming Events Using GFSSP in Liquid Propulsion Systems Alak Bandyopadhyay, Alok K. Majumdar, and Andre C. Leclair
	Cryogenics-06	Abstract Presentation Paper	Validation of accurate cryogenic fluid vapor-liquid boundary conditions via molecular simulations Daniel Vigil, Ashwin Ravichandran, and John Lawson (VIRTUAL)
	Cryogenics-07	Abstract Presentation Paper	Behavior of Ullage Bubbles during Blowdown in Low-g Experiment (BUBBLE): Overview of a Cryogenic Tank Depressurization Test Justin Pesich, Benjamin Nugent, and Steven Soriano
	Thermal Control and Protection-35	Abstract Presentation Paper	Spacecraft Radiator Protection from Ionizing Radiation, Dust, and Excessive Heat Loss Vijay Devarakonda, Michael D. Hogue, and Darnell Cowan
	Thermal Control and Protection-37	Abstract Presentation Paper	The Multi-Purpose Habitation Module (MPH): an overview of thermal modelling and lunar surface effects Lorenzo Strappato, Salvatore Lauretta, Davide Perrone, Simone Illiano, Roverto Bertacin, and Marilena Amoroso
	Thermal Control and Protection-27	Abstract Presentation Paper	Evaluation of AMSOIL®-ANT PGW Coolant Formula Change and Super Space AMSOIL Development Scott Hansen, Woody Beringer, Toni Griego, Lauren Foley, David Brockett, and Garret Gibeau
	Thermal Control and Protection-28	Abstract Presentation Paper	Orbit-Informed Thermal Modeling Workflow for Thermo-Mechanical Solder Fatigue Prediction of Electronics Ian Pond, Josh Akman, and Mina Karimaghaei
	Thermal Control and Protection-29	Abstract Presentation Paper	ONYX-Thermal – A New Concept for General Purpose Heat Transfer Software Dr. Dean Schrage
	Thermal Control and Protection-30	Abstract Presentation Paper	Intermediate temperature oscillating heat pipe radiators for space nuclear power Alex Miller and Scott Hayden
	Thermal Control and Protection-31	Abstract Presentation Paper	An Overview of the PULSE Space Thermal Control Pump Development at the Jet Propulsion Laboratory AJ Mastropietro, Michael R. Johnson, Warrick S. Leigh, Duval A. Johnson, Samuel G. Dupas, Talia R. Spitz, Dalia Raafat, Alejandro, Lopez Ortega, Allison Ayad, and Patrick Phelps
	Thermal Control and Protection-32	Abstract Presentation Paper	Artemis II Mission Performance of the Orion Active Thermal Control System Ellie Thurston, Michael Johnston, Thomas Chen, Eddi Uribe, Dongeun Lee, John Whitmore, Johnathan Hernandez, Maddie Haas, Matthew Gietzel, and Ian Anchondo
	Thermal Control and Protection-33	Abstract Presentation Paper	BubbleTrack: A Computer Vision-Based Framework for Bubble Detection and Tracking Sara Youssoufi
	Thermal Control and Protection-34	Abstract Presentation Paper	Heat Transfer and Aerodynamic Losses Caused By additively Manufactured GRX-810 Turbine Blades Enhanced With Micor-Machining and Chemical Polishing Phil Ligrani, Hallie Collopy, Mason Hancock, Jason Sheth, and Paul Gradl

Wednesday
9:30-11:30	Cryogenics-08	Abstract Presentation Paper	Recuperator and Heat Exchanger Development at NASA MSFC for Cryocooler Applications and Methane Liquefaction Demonstration Benjamin Williams, Sophia Angele, Mark Antonison, Aron Griffin, Mark Antonison, William Sixel, and Lonnie Webb
	Cryogenics-09	Abstract Presentation Paper	Materials for Cryogenic Propellant Dewetting Ethan Paulsen
	Cryogenics-10	Abstract Presentation Paper	Cryogenic Far-Infrared Instrument Performance Supported by Spectrally Extended Black Coatings Dr. Tamas Rev and Dina Katsir
	Cryogenics-11	Abstract Presentation Paper	Technology Integration Analysis of H2/N2 Gas Generator Use in Repressurization Systems for Human Lunar Landing Spacecraft Sydney Therien, Elaine Hatfield Tumlin, Adrian Soler, and Jonathan A. Bentley
	Thermal Control and Protection-40	Abstract Presentation Paper	Design of a Novel Radiant Heat Pump Thermal Control System for Space Conditioning of Lunar Surface Habitats Dr. Ardeshir Moftakhri
	Thermal Control and Protection-41	Abstract Presentation Paper	Additively Manufactured Porous Radiator For Static Thermal Management Kristen Ess and Rydge Mulford
	Thermal Control and Protection-42	Abstract Presentation Paper	Lunar Surface Crater Thermal Effects on Lander Radiator Performance Will Grier, Lisa Erickson, William Birmingham, and Tai Valdes
	Thermal Control and Protection-43	Abstract Presentation Paper	Ground Slope Effects on Lander Radiator Performance Lisa Erickson
	Thermal Control and Protection-39	Abstract Presentation Paper	Gravity-Independent Heat Exchangers and Oil-Free Compressors for Modular Refrigeration System Cara Martin and Dennis Nasuta
	Thermal Control and Protection-38	Abstract Presentation Paper	Development and Initial Validation of a High-Vacuum Thermal Conductivity Testbed for Aerospace Thermal Interface Materials Amy Chang and Deborah Hernandez
	Thermal Control and Protection-44	Abstract Presentation Paper	Thermal-Fluid Analysis of a Liquid-Cooled Battery Module for Electrified Aircraft Jarred Wilhite and Erik Stalcup
1:00-3:00	Cryogenics-12	Abstract Presentation Paper	Advanced Thermodynamic Modeling and Thermophysical Property Prediction of Cryogenic Fluids for Clean Energy Infrastructure Dr. Arash Pakravesh (VIRTUAL)
	Cryogenics-13	Abstract Presentation Paper	Additive Manufacturing for Cryogenic Fluid Management (AM4CFM) –Test Design for Liquid Acquisition Devices (LADs) Jack Parr
	Cryogenics-14	Abstract Presentation Paper	Additive Manufacturing for Cryogenic Fluid Management (AM4CFM) – Experimental Investigation for Liquid Acquisition Device Usage Dylan Foster
	Cryogenics-15	Abstract Presentation Paper	Development of Argon Loop Heat Pipe with Large Area Heat Collection Nathan Van Velson, Roopesh Kumar, and Calin Tarau
	Thermal Control and Protection-45	Abstract Presentation Paper	Thermal Analysis of Large Format Lithium-Ion Batteries in Thermal Runaway Ethan Sparks
	Thermal Control and Protection-46	Abstract Presentation Paper	Thermal Performance of a Lunar Rover RESS Konrad Brown, Trisha matthews, Peter Andruskiewicz, and Derek Lahr
	Thermal Control and Protection-47	Abstract Presentation Paper	Passive Thermal Runaway Propagation Prevention to Improve Safety of Li ion Batteries Vijay Devarakonda, Tanvi Gupta, Ranadip Saha, and Partha Mukherjee
	Thermal Control and Protection-48	Abstract Presentation Paper	Geometric Optimization of a Passive Ram-Airflow Cooling Duct for Li-ion Battery Thermal Management in Electric Semi-Truck Vehicles Rohan Jain and Bhargav Narayanan
	Interdisciplinary-10	Abstract Presentation Paper	SCIFLI Artemis II Launch Observation Overview Richard Schwartz, Jennifer Inman, Kyle Denny, Rufer Shann, Matt Boyda, Meaghan McCleary, and Carey Scott
	Interdisciplinary-12	Abstract Presentation Paper	SCIFLI Artemis II Reentry Airborne Operations Arianna Haven and Robert A. Conn
	CUI-06	Abstract Presentation Paper	SCIFLI Artemis II Orion Reentry Observation Summary Carey Scott Jr., Matthey T. Boyda, Kylel D. Scott, Arianna Haven, Meaghan M. McCleary, Shann J. Rufer, Alireza Mazaheri, Chris Johnston, and Andrew C. McCrea
3:30-5:00	Thermal Control and Protection-49	Abstract Presentation Paper	The ATA Technolgoy: Advanced Active Thermal Control for Space Applications Dr. Lucas Anderson, Charles Swenson, Miguel Nunes, and Robert Wright
	Thermal Control and Protection-50	Abstract Presentation Paper	A 2026 Update on the Active Mechanical Pumped Fluid Loop (MPFL) Projects at the Jet Propulsion Laboratory (JPL) Hared Ochoa
	Thermal Control and Protection-51	Abstract Presentation Paper	RANS-based assessment of transpiration cooling effectiveness in a curved porous nozzle wall exposed to high‑enthalpy supersonic flow Dr. Ram Adhikari
	Thermal Control and Protection-52	Abstract Presentation Paper	Parametric-Based Heat Rejection Trade Study for Lunar and Martian Surface Operations Noah Andersen and Thomas Chen
	Thermal Control and Protection-53	Abstract Presentation Paper	In-Situ Heat Rejection to the Martian Atmosphere Richard Schunk
	Thermal Control and Protection-54	Abstract Presentation Paper	An Update on Making Ultrasonic Additive Manufacturing (UAM) of Liquid Cold Plate Heat Exchangers a Reality for NASA and Aerospace Thermal Management Applications AJ Mastropietro, Gordy Cucullu, Nicholas Keyawa, Teri Juarez, Scott Innes, Scott Roberts, Jason Riley, Luke Walker, Lucas Anderson, and Charles Swenson

Thursday
9:30-11:30	Interdisciplinary-14	Abstract Presentation Paper	Participating Media Calculations in Ansys Thermal Desktop Douglas Bell
	Interdisciplinary-16	Abstract Presentation Paper	Thermal Modeling Philosophy for the Dragonfly Lander: Integrating Thermal Desktop and CFD for Titan Surface Operations Dr. Hui Liu, Evan Cosentino, Kurt Gonter, and Jane He
	Aerosciences-08	Abstract Presentation Paper	Space Launch System Artemis II Post-Flight Ascent Aerothermal Environments Summary Brandon Mobley and Samantha Summers
	Thermal Control and Protection-55	Abstract Presentation Paper	Coupled Thermal–Mechanical Modeling of a Deployable Radiator with Partial PCM Transformation and Runtime Radiation Articulation Joshua Taylor, Anthony Lococo, Franklin L. Robinson, Asher Leff, Darin Sharar, Rydge B. Mulford, and Brian D. Iverson
	Thermal Control and Protection-56	Abstract Presentation Paper	Development of a Thermal Radiator Optimization Tool with Alternate Coolants Richard Schunk
	Thermal Control and Protection-57	Abstract Presentation Paper	ASHRAE Climatic Design Methodology Applied to Diurnal Thermal Simulation Daniel Dannelley, Christopher Henry, and Vitaly Meyzler
	Thermal Control and Protection-58	Abstract Presentation Paper	Hot Moon: Designing a 20 kg Rover to Survive the Heat of Equatorial Lunar Noon (Lunar Outpost) Izzy Golemme
	CUI-02	Abstract Presentation Paper	Development of Self-Regulating Heater Prototype for Space Applications Dr. Jianjian Wang, Srujan Rokkam, and Nathan Van Velson
	CUI-03	Abstract Presentation Paper	Heat Pipe Thermal Management for High-Density Hall Thrusters Brett Leitherer, Quang Truong, Calin Tarau, Jeff Diebold, Benjamin Jorns, and William Hurley
	CUI-04	Abstract Presentation Paper	ISS Battery Charging  System Thermal Design and Development Dr. Siraj Jalali
	CUI-05	Abstract Presentation Paper	Space Launch System Core Stage Base Aerothermodynamics Post-Flight Reconstruction for Artemis II Dr. Manish Mehta, Brandon Mobley, and Samantha Summers
	Interdisciplinary-17	Abstract Presentation Paper	Machine-Learning Coupled Phase Change Closure Model  Based on Interfacial Heat Flux Discontinuity Dr. Sunjae Kim and Yusin Jeong
	Interdisciplinary-18	Abstract Presentation Paper	Physics-Based Postprocessing of 3-D CFD Results Dr. Bijay Sultanian
	Interdisciplinary-19	Abstract Presentation Paper	Closed-Cycle Transpiration-Cooled Flash Combustor with Cryogenically-Coupled Water Recovery for Thrust Augmentation in Pre-Cooled Airbreathing Vehicles Steven Schryba
	Interdisciplinary-20	Abstract Presentation Paper	Improved post-processing workflow for space thermal design insight Joel Gagnon and Jean-Frederic Ruel

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Abstracts

Aerosciences-01
Towards Isentropic Design Using the Entropy Map Generated from 3-D CFD Results
Dr. Bijay Sultanian

Using a CFD code, CFD engineers generate detailed distributions of the primitive variables like velocity, static pressure, and static temperature throughout the computational domain of a flow device. While the results are a means of quick flow visualization to identify regions of flow separation and recirculation (secondary flows), the primary goal of the CFD analysis is to minimize total pressure loss in the flow domain, thereby making the device more isentropic in its performance. Instead of postprocessing the 3-D CFD results to compute total pressure distribution in the flow domain, the CFD engineer should generate an entropy map to directly identify regions of excess entropy production for design improvements. This presentation shows how to generate an entropy map from 3-D CFD results, highlighting why this approach is better than the widely used method based on the postprocessed total pressure distribution. The proposed physics-based methodology is identical for both incompressible and compressible flows.

Aerosciences-03
Investigation of Peltier-based Environmental Control Systems for Low-cost/Attritable EO/IR Pods on Transonic Platforms
Mark Whittum

Airborne Electro-Optical and Infrared (EO/IR) pods are used extensively in sensing, national security, and surveillance applications. They are employed worldwide, frequently on fourth generation supersonic platforms such as the F-15 and F-16. The sophisticated EO/IR sensors at the heart of these podded systems are typically installed in sealed bays; such bays minimize ingress of contaminants but also present significant sensor thermal management challenges. Often, these pods rely on Vapor Cycle Machine (VCM) based Environmental Control Systems (ECSs) for sensor thermal management. VCM-based ECSs are heavy, complex, expensive and present maintenance and reliability problems.
The market / design space is opening up for simpler, lower-tier, lower-cost EO/IR sensor systems. Such systems are attractive to potential customers who employ either smaller, less capable fourth generation transonic platforms such as the L-159 or T-50, or attritable Unmanned Aerial Vehicles (UAVs). These lower-tier/cost sensor systems necessitate the need to examine lower cost/lower complexity ECS and thermal management systems.
One option for a lower-complexity ECS is a Peltier-based solid state heat pump, also known as a Thermo-Electric Cooler (TEC). A Peltier based system would eliminate the refrigerant and compressor at the heart of a VCM system and thereby offer customers significant advantages in both initial purchase price and maintainability/reliability costs.
This presentation will investigate the potential use of Peltier-based ECSs on lower-tier platforms, using preliminary system thermal design techniques with broad applicability. Basic flight mechanics, thermodynamics, reduced order modeling methods, and knowledge of applicable thermal environments will be employed to determine the cooling capacities and Concepts of Operations (CONOPS) available for Peltier-based ECSs on transonic airborne EO/IR pods.

Aerosciences-04
Comparative Assessment of Hypersonic Air-Breathing Inlet Architectures Under Common Freestream Conditions
Mookesh Dhanasar, Frederick Ferguson, Connor Ramaswamy, Leonard Uitenham, and Wisdom Calmday

Air-breathing inlets are critical components of high-speed propulsion systems, serving as the interface between the external flowfield and downstream engine components. Their primary function is to capture, compress, and condition the incoming airstream while minimizing total pressure losses and flow distortions prior to combustion. As flight speeds increase into the supersonic and hypersonic regimes, inlet performance becomes increasingly dependent on the effective management of shock-wave structures, flow compression processes, and aerodynamic integration with the vehicle forebody.
This paper presents a comparative assessment of multiple hypersonic air-breathing inlet configurations developed under common reference freestream conditions. A unique aspect of the present work is the application of a consistent design methodology and operating environment to evaluate multiple inlet architectures, thereby enabling direct comparisons of geometric characteristics, flowfield behavior, and aerodynamic performance. The inlet concepts investigated include wedge-derived, star-shaped, and axisymmetric forebody-inlet configurations intended for future high-speed air-breathing propulsion applications.
The design methodology utilizes ideal oblique shock relations coupled with streamline marching techniques to generate three-dimensional inlet geometries from prescribed compressible flowfields. Resulting geometries are converted into solid models and analyzed using computational fluid dynamics (CFD) tools to evaluate flow compression characteristics, shock structures, pressure distributions, Mach number contours, and overall flow uniformity within the inlet flowpath.
Results currently available include completed inlet geometries and CFD-based evaluations for multiple inlet classes under common operating conditions. Comparative analyses highlight the influence of inlet architecture on compression efficiency, flowfield structure, and downstream flow quality. The results further demonstrate the utility of the design framework as a rapid assessment tool for the development and evaluation of future hypersonic air-breathing propulsion systems.
The work is currently at the development stage, with initial CFD analyses completed and additional studies underway. Future efforts will incorporate inward-turning (Busemann-type) inlet concepts and higher-fidelity simulations using advanced CFD solvers, including Ansys Fluent and OpenFOAM. The long-term objective is to establish an expandable inlet design and assessment framework capable of supporting the systematic evaluation of emerging high-speed propulsion concepts.

Aerosciences-05
Comparative Assessment of Inversely Designed Hypersonic Waverider Configurations
Mookesh Dhansar, Frederick Ferguson, Connor Ramaswamy, Ajit Kelkar, and Donovan Berry

The development of hypersonic platforms requires aerodynamic configurations capable of efficiently generating lift while maintaining favorable performance characteristics within highly compressible flow environments. This research paper presents a parametric family of inversely designed hypersonic waverider configurations developed for unpowered high-speed flight applications. Unlike conventional geometry-first design approaches, the inverse-design methodology derives aerodynamic geometries directly from prescribed supersonic and hypersonic flowfields, producing configurations inherently adapted to their operating environments. In this approach, the vehicle geometry is effectively extracted / ‘carved’ from the governing flowfield, resulting in lifting surfaces that remain closely aligned with the underlying compressible-flow physics.
The methodology begins with the generation of supersonic and hypersonic flowfields using idealized two-dimensional oblique shock relations. Inviscid two-dimensional geometries are extracted from these solutions and extended into three-dimensional vehicle configurations through streamline cross-marching techniques. A unique aspect of the present work is that all configurations are developed under common reference freestream conditions and design methodologies, enabling direct geometric and aerodynamic comparisons across multiple classes of hypersonic vehicles.
The study examines a family of wedge-derived, caret-derived, axisymmetric, and star-shaped waverider configurations generated using the same inverse-design framework. Comparative assessments are conducted to evaluate configuration geometry, volumetric characteristics, aerodynamic performance metrics, and flowfield behavior. Analytical performance estimates and computational fluid dynamics (CFD) analyses are used to investigate lift, drag, lift-to-drag ratio, and representative flow structures associated with each configuration class.
Results currently available include completed geometric development, aerodynamic performance assessments, and CFD-based evaluations for representative configurations. The analyses demonstrate the adaptability of the inverse-design methodology for rapidly generating diverse hypersonic vehicle concepts while providing a consistent basis for aerodynamic comparison and early-stage design trade studies. The work further establishes a scalable framework for the systematic development and evaluation of future hypersonic vehicle families derived from common flowfield principles.The research is currently at the development stage, with configuration generation and initial aerodynamic assessments completed and additional CFD analyses ongoing. Future efforts will expand the configuration library and investigate the application of the methodology to integrated high-speed air-breathing propulsion systems and powered hypersonic flight vehicles.

Aerosciences-06
Three-dimensional, Multi-Phase Unsteady Flow of Water Undergoing Multiple Flow Regimes (free surface, droplet, and fluid film)
William Dziedzic

Rockets produce deafening sound waves at ignition. Without a suppression system, these extreme sound waves would reflect off the pad and Mobile Launcher, shaking the rocket violently and potentially causing it to tear itself apart. NASA’s sound suppression system for the Artemis SLS launch is the Ignition Overpressure Protection and Sound Suppression (IOP/SS) located at Launch Pad 39B at the Kennedy Space Center. This system uses a water deluge to help reduce the extreme acoustic energy and heat generated by the Space Launch System (SLS) rocket. At launch, the SLS rocket will produce nearly nine million pounds of thrust. The purpose of the sound suppression system is to dampen sound and vibrations to keep the rocket and the launch pad safe at lift-off. During the launch of Artemis, 450,000 gallons of water will be released onto the mobile launcher and flame deflector. When activated, it achieves a peak flow rate of 1.1 million gallons per minute.
Recent upgrades to the Mobile Launcher Rainbird nozzles were implemented for the Artemis II launch to increase water flow rates and improve the mobile launcher deck water coverage while minimizing the water throw distance to avoid impacting the Artemis II rocket nozzles. This analysis describes the use of a completed study of a three-dimensional, multi-phase, free surface transient computational fluid dynamics (CFD) model to simulate the water flowing out through the Rainbird nozzle and ensure compliance with current Artemis requirements. Flow rates, pressures, velocities, nozzle throw distances and coverage areas were analyzed to ensure that system requirements are met. The completed results were used to make necessary modifications to the Artemis I IOP/SS Rainbird design by reshaping the nozzle used in the system.
To determine the effective performance for each of the different nozzle geometries CFD simulations were performed. The transient multiphase CFD flow analysis of water includes multiple flow regimes, including free surface, droplet, and fluid film on the zero deck. The Volume of Fluid (VOF) model, the Lagrangian Multiphase model, and the Fluid Film model were used. The interface between the water and air is tracked throughout the simulation. To capture the water free surface properly the Adaptive Mesh and Adaptive Time-Step methods were implemented to optimize the mesh size and total cpu time to solve.
The incompressible volume-of-fluid (VOF) method was used to characterize the water flowing through air. In this method, the interface between the water and air is tracked and in conjunction with the Navier-Stokes/Euler equations, the water and air were modelled as they interacted through the domain. The turbulence was modelled using a 2-equation cubic k-epsilon model in a statistically steady Reynolds-averaged Navier-Stokes (RANS) setup. The simulations were run in a “pseudo-transient” state.
The completed CFD simulation results were then compared to NASA’s Verification and Validation (V&V) water deluge testing for Artemis II verifying the IOP/SS system. Teams measure the pressure, vibration, and spray coverage, comparing it against safety requirements before launch clearance is granted.
The objective of the analysis and V&V testing was to ensure that the nozzle shapes provided adequate coverage across the blast deck between the north and south Rainbirds. The CFD simulation results and V&V test results indicated that the modified Rainbirds would provide adequate coverage between the north and south Rainbirds on the sides of the blasthole. Also, CFD analysis predicted that the deflection of water caused by the curved nozzle would allow the water to fall short of the vehicle at full flow, thus mitigating the vehicle impingement.
Another objective of the analysis and V&V testing was to examine the interaction of the Northwest Rainbird and the North Center Rainbird. It was expected that the water from the nozzles would flow between the two North Rainbirds and interact with one another. This interaction could potentially lead to vehicle impingement. The CFD simulation was able to provide the quantity referring to the flow rate of water and quality generally referring to the droplet size of the water, i.e., large droplets or mist in the vehicle area. The CFD simulation results and V&V test results compared well and predicted that this interaction is in the form of a mist.
The presentation will show the three-dimensional CFD Rainbird nozzle results of the volume fraction of the water, particle diameter size, fluid film thickness coverage on the zero deck. Both contour images and avi files will be shown.

Aerosciences-07
Aerothermal Performance Enhancement of Supersonic TsAGI S-12 Airfoils Using Graphene Nanoplatelet Coatings: Experimental and CFD Investigations
Lakshmanan Kasi, Rajalakshmi Padmanabhan, and Perarasu Thangavelu

Supersonic flight environments are characterized by strong shock-wave formation and intense aerodynamic heating, both of which adversely affect vehicle performance, structural integrity, and operational efficiency. This study investigates the effectiveness of graphene nanoplatelet (GNP) surface coatings as a passive aerothermal control strategy for mitigating shock-induced drag and reducing thermal loads on a TsAGI S-12 airfoil operating at Mach numbers ranging from 1.5 to 3.0. A combined experimental and computational approach was employed. Wind tunnel experiments were conducted to evaluate shock-wave behavior and surface temperature distributions, while high-fidelity computational fluid dynamics (CFD) simulations were performed using OpenFOAM to analyze pressure fields, velocity distributions, shock-wave characteristics, and thermal response.
The numerical model incorporated experimentally measured surface roughness values obtained from atomic force microscopy, representing uncoated and GNP-coated airfoil surfaces. Results demonstrate that the graphene coating reduces shock-wave angle by up to 18%, weakens adverse pressure gradients, and decreases aerodynamic drag by approximately 18% across the investigated Mach number range. Thermal analysis further revealed a reduction of nearly 21% in peak surface temperature, attributed to enhanced heat redistribution and reduced localized thermal loading. Good agreement between experimental measurements and CFD predictions confirms the reliability of the proposed modeling framework.
The findings establish graphene nanoplatelet coatings as a lightweight and scalable solution for simultaneous aerodynamic and thermal performance enhancement in high-speed aerospace vehicles. The study provides new insights into the coupling between surface engineering, shock-wave dynamics, and aerodynamic heating, offering a promising pathway for future aerothermal design optimization of supersonic platforms.

Aerosciences-08
Space Launch System Artemis II Post-Flight Ascent Aerothermal Environments Summary
Brandon L Mobley and Samantha Summers

Since 2011 the Aerosciences Branch/EV33 at NASA Marshall Space Flight Center has been involved with the development of ascent external aerothermal environments for the NASA Space Launch System (SLS) Block 1 launch vehicle for the purposes of supporting thermal analysis and the design of thermal protection systems. The SLS Block 1 Artemis I and II launch vehicles successfully launched from Pad39B at NASA Kennedy Space Center on November 16th, 2022 and April 1st, 2026, respectively. Over 70 aerothermal islands, consisting of over 265 operational instruments captured aerodynamic heating and plume induced environments throughout the launch vehicles. Gauges consisted of calorimeters, radiometers, gas temperature probes, pressure transducers, bi-directional pressure probes and thermocouples. Prior to launch, aerothermal design environment models were generated to predict ascent aerodynamic heating and plume induced environments over a design space that covered a range of vehicle trajectories that varied atmospheric, vehicle performance, and off-nominal, engine-out conditions. Post flight reconstruction models were developed for each flight island using the Day-of-Launch (DOL) Best Equivalent Trajectory (BET) that provided freestream conditions and propulsion system boundary conditions. This paper discusses a summary of the ascent aerothermal environments observed during the flights and the respective modelling approaches and the performance of them through comparisons of flight data and predictions.

Aerosciences-09
High-Fidelity CFD Simulation of Liquid, Hybrid, and Solid Rocket Propulsion Systems
Dr. Xiao Ren

Computational fluid dynamics (CFD) has become an increasingly important tool for understanding and optimizing rocket propulsion systems. However, accurately modeling rocket engines remains challenging due to the strong coupling between turbulence, combustion, multiphase flow, chemical kinetics, heat transfer, and compressible flow phenomena. This presentation highlights recent CFD applications spanning liquid, hybrid, and solid rocket propulsion systems.
For liquid rocket engines, simulations of single-element and multi-element combustors operating under both subcritical and supercritical conditions are presented. The effects of thermodynamic property models, detailed chemical kinetics, and combustion modeling approaches on combustion performance and chamber pressure oscillations are discussed.
For hybrid rocket motors, a case utilizing self-pressurizing liquid nitrous oxide oxidizer and an ABS-based fuel grain is examined. The simulations employ multiphase modeling to capture liquid-gas interface dynamics and oxidizer phase change, together with detailed chemistry to resolve combustion processes.
For solid rocket motors, a surface regression model is used to predict propellant burning behavior and grain recession. The analysis captures chamber pressure evolution, combustion-product transport, and nozzle flow characteristics, providing insight into the interaction between internal ballistics and nozzle performance.
Collectively, the presentation focuses on modeling approaches, validation against experimental data, and the physical insights that can be gained for propulsion design and performance prediction.

Aerosciences-10
Euler-to-RANS Mapping of Shock-Interaction Regimes on Slender Biconic Forebodies at Mach 2–5
Aneesh Kukreti

Slender biconic forebodies generate interacting shock systems that influence surface pressure loading, boundary-layer behavior, and thermal environments on supersonic and hypersonic vehicles. Inviscid Euler simulations are often used for rapid early screening, but their reliability for predicting shock-interaction regimes relative to viscous Reynolds-Averaged Navier-Stokes (RANS) solutions is strongly dependent on geometry, Reynolds number, and evaluation metrics. This work assesses where Euler-based screening is useful and where viscous RANS analysis is required for an idealized axisymmetric slender biconic forebody over Mach 2–5.
The primary contribution is a structured Euler-to-RANS screening workflow for shock-interaction regime targeting in slender biconic configurations. To reduce selection bias, the Euler case matrix, high-turn extension rule, topology classes, RANS selection strategy, and Euler adequacy criteria were all specified a priori prior to production runs. The study isolates Mach number and second-cone turn angle effects using a frozen geometry (forecone angle 10°, shoulder location x_c/L = 0.45) at a fixed Reynolds number Re_inf,x_c = 2.0 × 10^6. All viscous simulations assume a calorically perfect gas, axisymmetry, adiabatic walls, and fully turbulent SST closure.
The inviscid study produced a 20-case Euler topology map spanning Mach 2, 3, 4, and 5 and second-cone turn angles from 6° to 22°. Nineteen of twenty cases exhibited attached forecone and corner shock systems. The only near-boundary case occurred at Mach 2 and 22°, where the corner shock approached a predeclared near-detachment criterion. No Euler solution produced a detached shock system, Mach stem, or triple-point-like topology, suggesting that the inviscid detachment boundary may lie outside most of the tested envelope.
A targeted six-case RANS subset was used to evaluate whether attached Euler topology implies attached viscous behavior or quantitative agreement. This subset was a boundary-bracketing selection rather than a statistical sample of the full Euler matrix. Four of six RANS cases exhibited incipient corner separation at Mach 2 (14° and 22°), Mach 3 (14°), and Mach 4 (14°), defined by localized or near-zero wall skin friction without a sustained separated region. The remaining cases, including both Mach 5 conditions, remained attached.
Euler-to-RANS comparisons used fixed engineering tolerances of 10% in peak pressure coefficient and 7.5% in pressure drag coefficient. Under these criteria, Euler predictions were adequate for three of six cases, and inadequate for the Mach 2 near-boundary case and both Mach 5 cases. Sensitivity studies support these classifications: all incipient cases remained unchanged under Spalart–Allmaras turbulence modeling, Mach 5 isothermal-wall variations did not alter attachment behavior, and a repaired-grid solution confirmed attached flow for the Mach 5, 14° case with less than 5% variation in key force and pressure metrics.
The results show that Euler solutions are effective as a topology-screening and RANS case-selection tool for this restricted slender biconic configuration but are not a universal quantitative surrogate for viscous analysis. While Euler correctly captured broadly attached inviscid shock topology across the parameter space, RANS revealed incipient viscous separation in four cases and significant quantitative deviations in three of six comparisons. The resulting workflow provides a traceable and reproducible basis for determining when inviscid screening is sufficient and when higher-fidelity viscous simulation is required.

Aerosciences-11
SABRE: A PHYSICS-DRIVEN FRAMEWORK FOR SATELLITE RE-ENTRY DEMISE ANALYSIS
Suyash Tandon

Predicting whether satellite components survive atmospheric re-entry is critical for compliance with FCC and NASA casualty risk requirements ( ! < 1: 10,000). Industry-standard tools— NASA DAS and ESA DRAMA, approximate spacecraft using representative geometric shapes and empirical fragmentation rules, limiting fidelity for novel materials and complex component geometries. This presentation introduces SABRE (Spacecraft Atmospheric Breakup and Re-entry Evaluation), an in-house multi-fidelity simulation framework developed at Amazon Leo. SABRE couples panel-based aerodynamics (Newtonian/Fay-Riddell heating with Sentman free-molecular bridging), a 1D Crank-Nicolson transient thermal solver with char-layer state tracking, and physics-driven structural failure criteria (aerodynamic bending, thermal mismatch, and hoop stress) applied directly to CAD-derived mesh geometry. A 6-DOF trajectory propagator feeds perzone thermal loading as the spacecraft tumbles through descent. Validation against the Rochelle 1997 and Ostrom 2025 six-tool comparison benchmarks demonstrates agreement with community tools while capturing structural failure modes that melt-only models cannot predict. Application to a representative multi-component test satellite, including a magnetic iron ring, a pressurized titanium vessel, and CFRP structural panels demonstrates SABRE’s capability for componentlevel demise and casualty risk prediction across a range of materials and geometries.
References
[1] Rochelle, W.C., Marichalar, J.J., & Johnson, N.L. (1997). “Development and validation of
ORSAT.” AIAA-97-0475. NTRS 19970040121.
[2] Ostrom, C. et al. (2025). “Six-tool re-entry demise comparison study.” NTRS 20250002905.
[3] Fay, J.A. & Riddell, F.R. (1958). “Theory of stagnation point heat transfer in dissociated air.”
Journal of the Aeronautical Sciences, 25(2), 73–85.
[4] Sentman, L.H. (1961). “Free molecule flow theory and its application to the determination of
aerodynamic forces.” LMSC-448514

Cryogenics-01
A Scalable, Calibration-Free Multiphase Arbitrary Lagrangian Eulerian Solver for Sloshing Dynamics and Phase Change
Dr. Antonio Cantiani

Accurate simulation of multiphase flows involving sloshing dynamics and phase change remains a challenge for Computational Fluid Dynamics (CFD). Traditional methods, such as Volume of Fluid (VoF) approaches, typically require highly refined meshes and complex reconstruction algorithms, or empirical parameters that demand rigorous, case-specific calibration. This work introduces a novel, calibration-free multiphase Arbitrary Lagrangian-Eulerian (ALE) solver designed to fundamentally resolve these limitations. By explicitly separating the liquid and gas phases into distinct computational regions, the solver inherently maintains a perfectly sharp interface regardless of the mesh resolution used. This defining feature allows for the successful use of relatively coarse meshes without sacrificing accuracy. Furthermore, coupling this sharp interface tracking with appropriate thermal modelling completely eliminates the need for empirical tuning of the phase change model. The powerful combination of these two advantages, bypassing case-specific calibration and enabling coarse-mesh accuracy, makes the proposed methodology exceptionally well-suited and computationally efficient for large-scale simulations. In this work, the underlying mathematical framework is presented, algorithmic improvements are detailed, and the solver’s performance is demonstrated through rigorous benchmark validations. Specifically, the model has been validated using experimental data from liquid nitrogen (LN2) sloshing tests performed at the von Karman Institute for Fluid Dynamics, alongside literature data detailing the self-pressurization of cryogenic tanks. An assessment of these results demonstrates a highly accurate prediction of both pressure and temperature evolution under static (self pressurization) and dynamic (sloshing) conditions. Ultimately, this enhanced ALE CFD solver offers a highly predictive, physically accurate, and scalable tool for large-scale engineering applications where sloshing and evaporation/condensation are critical.

Cryogenics-02
CFD-assisted nodal modeling of sloshing in a cryogenic propellant tank
Dr. Alok Majumdar, Andre LeClair, Ramon Lopez Otero, and Jacob Brodnik

During autogenous pressurization, tank sloshing causes a significant increase in pressurant consumption to maintain constant ullage pressure during draining of the tank. This increase in pressurant consumption is caused by a significant increase in condensation at the liquid-vapor interface. Sloshing strongly affects the liquid side heat transfer coefficient and thereby the condensation rate. Traditionally, sloshing is modeled by CFD code using the VOF (Volume of Fluid) method to track the liquid vapor interface during sloshing. CFD calculations require a very fine grid to accurately compute the heat and mass transfer at the interface. Therefore, computations are time consuming and prohibit performing many parametric studies often needed during the design of a new system. This paper describes an alternative approach by developing a CFD-assisted nodal model to predict system parameters more economically with reasonable accuracy. In this approach, a multi-node model of tank pressurization was developed using GFSSP. A multi-node model was needed to account for stratification. The model computes heat and mass transfer at the interface to calculate the condensation rate. The liquid side heat transfer is computed using the parameters of sloshing dynamics such as frequency, wave amplitude, and interface area. The parameters of sloshing dynamics are computed by the CFD code LOCI-Stream. The model predictions were compared with test data for several cases.

Cryogenics-03
Cryogenic ullage pressure collapse via condensation and ullage heat loss
Dr. Samuel R Darr

Cryogenics-04
Initial Computational Analysis of Helium Subsurface Pressurization
Jacob Brodnick and Jason Hartwig

The dynamics of helium subsurface pressurization (HSP) in standard Earth gravity and low gravity were assessed using computational fluid dynamics (CFD) analysis. The intent is to help inform future trade studies of tank pressurization for spacecraft engine restarts and propellant transfer. Efficient methods of pressurizing directly into a cryogenic propellant are sought. Additionally, HSP has been shown to thermally condition cryogenic propellants which may be leveraged in future in-space operations. Initial results on replication of pressurant bubble flow regime are presented for an experiment conducted under a NASA Flight Opportunities grant, FO 244-P. Multiple pressurization rates were applied to an approximately 8 liter liquid nitrogen tank both on the ground and during a parabolic flight that periodically achieved low gravity. A viewport was used to directly image the bubble dynamics during pressurization. CFD simulation of the process used a volume of fluid method of distinct fluid phase representation. Qualitative comparison of bubble dynamics and classification of flow regime were made. Future work will include an expanded comparison of the tested parameter space and various quantitative comparisons. The present analysis is a part of a concerted effort to better understand if subsurface pressurization is viable or optimal for any applications or parameter spaces in future NASA missions.

Cryogenics-05
Modeling of Priming Events Using GFSSP in Liquid Propulsion Systems
Dr. Alak Bandyopadhyay, Alok K Majumdar, and Andre C Leclair

Water hammer analyses are crucial and critical in risk mitigation of aerospace propulsion systems as the analyses predict the transient pressure spikes caused by sudden change in fluid velocity. In this study, rigorous computational simulations have been performed using network flow solver code GFSSP to study the effect of pipe length, valve coefficient and downstream pressure on peak pressure evaluation and response time of the pressure surges in straight pipes. The results are compared with the test data conducted at Penn State by Moore et al. More than 50 different test cases by varying the pipe length and diameter, valve coefficient and exit pressure conditions (atmospheric. Sub-atmospheric and low pressure) and the inlet pressure are used in this simulation, and results are compared with the test data. For the simulation, a two fluid system model with water and air mixture has been used. Excellent comparison with test data was observed for relatively short pipes (0.51 m) with small valve coefficients (within 2%). The simulation results are compared reasonably well even for longer pipes (2 m) when the valve coefficients are small. However, there is huge discrepancy (>500%) of the simulation results from the test data for large valve coefficients. Hence an unsteady state friction factor model is used to compute the fluid friction coefficient to compute the viscous terms, especially for longer pipes with large valve coefficients and results improved a lot, and the difference between computed results and test data for these cases reduced significantly and, in some cases, it is less than 2%.

Cryogenics-06
Validation of accurate cryogenic fluid vapor-liquid boundary conditions via molecular simulations (VIRTUAL)
Daniel L Vigil, Ashwin Ravichandran, and John Lawson

Vapor-liquid interfaces drive many important phenomena in cryogenic fluid management, including heat transfer, evaporation, and capillary flow. Design of cryogenic fluid systems, such as propellant storage, requires accurate predictions of fluid behavior, including evaporation rates. Many models have been proposed for heat and mass transfer at vapor liquid interfaces, but the accuracy of these models in the context of cryogenic fluids has not been assessed. We use non-equilibrium molecular dynamics simulations, which allow for nanometer scale resolution of fluid phenomena, to evaluate the accuracy of a variety of vapor-liquid boundary conditions at evaporating and condensing interfaces. We find that an anisotropic temperature distribution is a critical ingredient for accurate prediction of intensive evaporation and condensation.

Cryogenics-07
Behavior of Ullage Bubbles during Blowdown in Low-g Experiment (BUBBLE): Overview of a Cryogenic Tank Depressurization Test
Justin Pesich, Benjamin Nugent, and Steven Soriano

Behavior of Ullage Bubbles during Blowdown in Low-g Experiment (BUBBLE) is a cryogenic tank depressurization drop tower experiment currently being designed at NASA Glenn Research Center. There is a need for liquid level rise data while venting a cryogenic propellant tank below the liquid saturation pressure in a reduced gravity environment. When a cryogenic tank is vented to a vapor pressure below the liquid saturation pressure, bubbles become entrained in the liquid causing the bulk liquid-vapor interface to rise. In reduced gravity, bubble rise velocity is diminished leading to a larger liquid level rise compared to a 1g environment. The purpose of the experiment is to gain further understanding of tank depressurization fluid physics and obtain high-fidelity data for model validation in 1g and reduced gravity environments. On-orbit cryogenic propellant tanks operating at high fill levels must efficiently manage venting operations to avoid the risk of liquid entrainment in the vent line, which could lead to asymmetric control thruster loads, freezing and clogging, and loss of liquid propellant. Validated models could be used to design settling and venting profiles to reduce risk and increase efficiency for cryogenic storage and transfer operations.

Cryogenics-08
Recuperator and Heat Exchanger Development at NASA MSFC for Cryocooler Applications and Methane Liquefaction Demonstration
Benjamin Williams, Sophia Angele, Mark Antonison, Aron Griffin, Mark Antonison, William Sixel, and Lonnie Webb

NASA MSFC is investigating advancements for highly efficient cryocooler systems for cryogenic fluid management (CFM) of liquid propellant systems in multi-month to multi-year space missions. Current developmental cryocooler designs employ one working fluid and are composed of several critical components including a counter-flow heat exchanger known as a recuperator. Heat exchanger and recuperator designs are typical in industrial applications but there are limited examples of recuperators in cryocoolers for space mission applications. Most of these space mission cryocoolers involve cooling refrigerant for instruments on satellites or telescopes. State of the art examples of cryocoolers for long duration management of rocket engine cryogenic propellants have only advanced to ground-based system testing. Therefore, NASA MSFC has undertaken the task of designing a recuperator for cryogenic propellant management to mature the technology and buy down critical technology gaps such as cryogenic fluid storage and cryogenic fluid transfer. Current iterations of the component for cryocooler applications utilize a counter-flow Printed Circuit Heat Exchanger (PCHE) design with Ti6Al4V plate stacks containing chemically etched channels. The design is currently in the analysis stage but there is funding and support to manufacture and test the full cryocooler system in the near future. Additionally, as a risk mitigation step, NASA is pursuing a technology demonstration to liquefy gaseous methane using several subscale components from the aforementioned cryocooler system. This requires the PCHE recuperator design to be scaled down and manufactured in SS316 to adhere to liquefaction system requirements and manufacturing constraints. Due to the immaturity of this technology, there have been numerous lessons learned throughout the design-analysis cycle processes for both the cryocooler and liquefaction demonstration designs. This paper will summarize the recuperator development progress that has been completed at NASA MSFC including design and manufacturing as well as emphasizing lessons learned.

Cryogenics-09
Materials for Cryogenic Propellant Dewetting
Ethan Paulsen

Cryogenics-10
Cryogenic Far-Infrared Instrument Performance Supported by Spectrally Extended Black Coatings
Dr. Tamas Rev and Dina Katsir

Cryogenic Far-Infrared Instrument Performance Supported by Spectrally Extended Black Coatings
Cryogenic far-infrared space instruments are highly sensitive to internally generated straylight, unwanted reflections, and instrument self-emission. In thermal-infrared systems, reflections from internal baffles, housings, calibration cavities, and other non-optical surfaces may contribute to additive background, spatial artifacts, and radiometric uncertainty. For instruments designed to observe cold targets and small temperature differences, the optical and thermal behavior of internal surfaces becomes an important part of the overall instrument design.
A relevant example is NASA’s Europa Thermal Emission Imaging System, E-THEMIS, on the Europa Clipper mission. E-THEMIS is designed to map Europa’s surface temperature and thermal behavior in infrared bands. These temperature measurements may help scientists identify thermally anomalous regions that, when combined with other mission data, could provide clues about ice-shell structure, recent geological activity, and possible exchange between the icy surface and the suspected subsurface ocean. This type of measurement requires careful control of both reflected straylight and instrument-generated radiative background.
This presentation discusses spectrally extended black coatings as candidate surface treatments for cryogenic mid- and far-infrared instruments. These coatings are designed to provide low reflectance and high infrared emissivity over an extended spectral range, supporting reduced multiple-reflection straylight and more predictable radiative behavior inside instrument cavities. Potential application areas include internal baffles, housings, calibration cavities, and non-optical structural surfaces where straylight suppression, emissivity control, cleanliness, vacuum compatibility, and space environmental durability are required.
The unique aspect of this work is its focus on black coating performance beyond conventional visible and near-infrared straylight suppression, extending the discussion into the mid- and far-infrared regime where coating morphology, thickness, emissivity, cleanliness, and integration constraints become especially important. Rather than treating black coatings only as optical absorbers, the presentation considers their combined optical, thermal-radiometric, and implementation roles in cryogenic infrared instrument design.
The presentation will focus on coating-related considerations, including spectral behavior, cryogenic and vacuum compatibility, cleanliness, and integration on representative instrument surfaces. The work highlights how spectrally extended black coating design can contribute to the optical and thermal-radiometric toolbox available for future cryogenic far-infrared space missions.

Cryogenics-11
Technology Integration Analysis of H2/N2 Gas Generator Use in Repressurization Systems for Human Lunar Landing Spacecraft
Sydney E Therien, Elaine Hatfield Tumlin, Adrian Soler, and Jonathan A Bentley

Solid compounds that react to form large volumes of gas are currently being investigated for gas generator applications as a potential mechanism to quickly recover a human landing system (HLS) spacecraft from sudden losses in tank pressure. Slosh-driven ullage collapse, where tank pressure drops in response to cold liquid splashing into warmer ullage and therefore cooling and condensing the ullage gas, may be encountered in a landing maneuver. If tank pressure drops outside of the allowable range, hazards like tip-over pose a severe risk as typical pressurization systems are not built to react quickly to these events. Swift repressurizations could be made possible by GGs, but only if the technology is compatible with the cryogenic propellants used on HLS landers (LH2, LCH4, LOX). This work will limit its scope to GGs that create H2 and N2 since these solids have the most researched chemistry and their gases already have many uses in spacecraft.
To determine if H2/N2 GGs are suitable for this application, analysis was performed to answer questions of sizing, propellant compatibility, and engine impacts. The volume of gas required for repressurizing an agnostic HLS lander was calculated to see if GGs could produce on the same order of magnitude as might be required. Information on H2 and N2 solubility in the cryogenic propellants was gathered. Departure from the nominal Isp due to propellant displacement and contamination from GG byproducts of a LCH4/LOX engine was calculated using Cequel (CEA Excel plug-in). These results were used to find the maximum displacement of propellants by dissolved gas that ensured mission delta Vs are still achievable. Results of this work will inform design of GG repressurization systems and help advance them to a technology readiness level that could see them implemented on an HLS spacecraft.
Currently, the bulk of research and analyses have been conducted. The final report will synthesize these data and critically evaluate each sublimated gas/propellant combination in terms of sizing, compatibility, and performance. It will contain a thermodynamic model for sizing a GG system for an HLS spacecraft. It will also contain a detailed summary of the analysis performed to understand risks of H2/N2 pressurant gas diffusion in selected propellant systems and the contingent effects on engine performance.

Cryogenics-12
Advanced Thermodynamic Modeling and Thermophysical Property Prediction of Cryogenic Fluids for Clean Energy Infrastructure (VIRTUAL)
Dr. Arash Pakravesh

Accurate prediction of thermophysical properties and phase behavior for cryogenic fluids—specifically liquid hydrogen (LH2), liquefied natural gas (LNG), and hydrogen-enriched natural gas (HENG) blends—is critical for the design, thermal management, and safe operation of modern aerospace and ground-based energy systems. In low-temperature regimes (sub-100 K) and high-pressure processing environments, traditional cubic equations of state (EoS) often fail to capture strong density fluctuations, non-ideal mixing behaviors, and precise derivative properties near saturation boundaries.
This paper presents an integrated thermal-fluids modeling framework based on advanced Statistical Associating Fluid Theory (specifically PρT-SAFT) alongside reference multiparameter equations (GERG-2008, AGA8, and modified cubic EoS). The model evaluates volumetric properties (PρT), vapor-liquid equilibria (VLE), speed of sound, and second-order derivative properties across wide operational envelopes relevant to storage, transfer, and liquefaction systems.
Special emphasis is placed on the precise computation of second-order thermodynamic derivatives and derivative-dependent properties across pure components and multicomponent mixtures of hydrogen, methane, nitrogen, and carbon dioxide. Accurate modeling of speed of sound, isobaric and isochoric heat capacities, isothermal compressibility, and the Joule-Thomson coefficient is essential for capturing subtle thermal shifts. These derivatives directly govern transient expansion dynamics in throttle valves, Boil-Off Gas (BOG) accumulation in sealed cryogenic tanks, acoustic shock formation in fluid lines, and local cooling or unintended heating during rapid depressurization.
The presented framework equips thermal and fluids engineers with highly accurate, derivative-consistent thermodynamic inputs required for robust boundary conditions in CFD solvers, thermal network codes, and the optimization of low-loss cryogenic transfer systems.

Cryogenics-13
Additive Manufacturing for Cryogenic Fluid Management (AM4CFM) –Test Design for Liquid Acquisition Devices (LADs)
Jack T Parr

In-space cryogenic propellant management is one of the greatest short-term tasks concerning future spaceflight. Behavior of cryogenic propellants is not intuitive in microgravity environments and controlling them is of the utmost importance. Liquid Acquisition Devices (LADs) are a necessary way of controlling propellants in these environments. Engines must reliably receive single phase liquid propellant to function while concerns of cryogenic fluid boiling in the tank are also present. This project will test the functionality of additively manufactured (AM) lattices and porous metal as a replacement for traditionally manufactured screens in gallery channel LADs and impact on future unique capabilities of AM to incorporate active cooling into a gallery channel LAD. LADs exist as a suite of tools that engineers can use for cryogenic fluid management, and this study is the first to characterize gallery channel LADs made with additive manufacturing. Metal additive manufacturing techniques have progressed much over the years, and their application towards complex lattice and porous material construction is seen as advantageous towards LAD design and manufacturing. The test articles were manufactured using Laser Powder Bed Fusion (L-PBF) and the lattices contain repeating geometric formations designed to promote capillary flow channel formation. Several different lattice unit cells were tested including Body Centered Cubic (BCC), Octet, HexaProfile, and PentaCube with unit cell sizes ranging from 1-1.5 mm. In addition to lattice test articles, we tested permeable additive manufactured samples as well. The wicking testing will help characterize LADs by correlating their capillary parameters like permeability and effective wicking diameter. Characterizing LADs enables predictive models to be generated based on the lattice attributes and helps with future designs. The test coupons were tested both vertically and horizontally. The design, procedures, and results of a wicking rate test for several different additively manufactured test coupons are discussed. Key performance parameters for the wicking rate test were the weight of fluid absorbed by the LAD sample as well as the wicking front of the coupon. The more weight that the sample wicked indicates a sample that can wick more fluid, and the wicking front being higher shows a better ability of the sample to wick fluid. Mass absorbed by the coupons was measured with a high precision Secura scale while the wicking front was monitored with an infrared camera.

Cryogenics-14
Additive Manufacturing for Cryogenic Fluid Management (AM4CFM) – Experimental Investigation for Liquid Acquisition Device Usage
Dylan Foster, Travis Belcher, Carly Romnes, Lonnie Webb, Jack Parr, Ethan White, and Cassidy Brozovich

In-space propellant management remains a high-risk area for Cryogenic Fluid Management (CFM). With growing mission complexity and advanced vehicle maneuvering necessary to reliably meet objectives in microgravity environments, it will be necessary to advance Liquid Acquisition Device (LAD) technologies to ensure single-phase liquid propellant delivery to the engine on-time and at scale. This work is the first investigation into the use of additive manufacturing for novel LAD design methodologies. Additive manufacturing enables unique geometries and methods for achieving the porous capillary behaviors necessary for LAD performance. Porous geometries to be assessed are lattice structures and permeable additives, both manufactured using Laser Powder Bed Fusion (L-PBF). Lattices are repeating geometric formations which, when designed intentionally, have been shown to promote formation of capillary flow channels. Lattice unit cells investigated include Body-Centered-Cubic (BCC), Octet, HexaProfile, and PentaCube, with unit cell sizes of 1 mm and 1.5 mm. Permeable additives are manufactured by adjusting build parameters such as laser power, scan speed, and hatch spacing to reduce the volumetric energy density, the total energy going into a part, and intentionally creates porosity which allows fluid to permeate through a part. Samples were created to characterize the relevant parameters necessary for predicting performance in a LAD use-case. This includes experiments to capture wicking rates, bubble point, and liquid flow-through pressure drop. The aforementioned tests will demonstrate the capacity for additive LADs to rewet dry portions of a screen, prevent vapor ingestion, and quantify flow resistances during propellant delivery. This testing will culminate in the design and subscale system testing of an additively manufactured gallery channel which incorporates complex porous features. Preliminary results will be presented from each aspect of the characterization work, as well as a discussion on modeling efforts for development of generalizable predictive methods for additive LAD designers.

Cryogenics-15
Development of Argon Loop Heat Pipe with Large Area Heat Collection
Nathan Van Velson, Roopesh Kumar, and Calin Tarau

Loop heat pipes (LHPs) are passive, two-phase thermal management devices capable of transporting large heat loads over long distances without the use of mechanical pumps. They are widely employed in spacecraft thermal control systems to transfer waste heat from payloads to radiators. This work presents the development and experimental evaluation of an argon loop heat pipe designed to operate at cryogenic temperatures near 120 K, investigated in two distinct configurations. The first configuration corresponds to a conventional LHP architecture commonly used in space applications, while the second employs a modified architecture intended to reject heat loads distributed over a large surface area. Thermal performance testing was conducted over input power levels ranging from 50 W to 100 W at sink temperatures near 120 K. Experimental results demonstrate stable operation of the LHP across the tested conditions, with the conventional configuration successfully transporting heat loads of up to 100 W. In the modified configuration, the LHP provided effective cooling of a large-area heat source up to 40 W applied on the large area when 15 W was applied to the capillary pump. This work was conducted under the DOE SBIR Phase II program (DE-SC0022896).

Interdisciplinary-01
A Case Study of AI-assisted Creation of a Thermodynamics Model of Precipitation Formation During Rapid Depressurization of a Vented Container
Dr. Jedediah Storey

Precipitation may form in humid containers undergoing rapid depressurization. This precipitation may be liquid, i.e. fog, if the dewpoint is crossed above the freezing point of water, or direct snow crystallization if the dewpoint is crossed below the freezing point. Accurate modeling of this effect is potentially important for rapidly ascending vented compartments in aircraft or launch vehicles, as well as rapid depressurization of vacuum chambers [1]. A thermodynamics model of precipitation formation during the rapid depressurization of a container was developed in python. The model is written for a generic container and includes an optional liquid pool. Modeled physics include polytropic expansion and condensation in the container air, and film convection, evaporation, boiling, freezing, sublimation, structure heat transfer, freezing while evaporating, triple point boiling & freezing, flash vaporization, and splash enhancement from rigorous boiling for the pool. Scipy solve_ivp ‘BDF’ is used to integrate the ODE system in time. Transitions between pool regimes are smeared and attempt to conserve mass and energy. Results from several example cases spanning the full capabilities of the code will be presented. Numerous prior publications detail the thermodynamics of vacuum freezing of liquids [2-5, non-exhaustive]. However, unlike these prior works, this code was created as a case study of the use of AI Large Language Models (LLMs) to assist modeling of physical processes. Two different LLMs were utilized via NASA ChatGSFC [6]: the relatively inexpensive Claude Haiku 4.5, and the more expensive and capable Claude Opus 4.7, both by Anthropic. The attempts to use each LLM will be discussed, and estimates of performance, time, and cost of using the LLMs will be compared to the same for a thermodynamics expert, with little software development training, writing a similar code unassisted. The LLMs’ strengths and weaknesses in this context will be emphasized.
AI was not used in the writing of this paper, only in the creation of the model and python code.
[1] https://www.youtube.com/watch?v=8SOCni8JfiU
[2] doi:10.1016/j.vacuum.2023.112040
[3] doi: 10.1038/srep35324
[4] doi: 10.3969/j.issn.1672-7126.2009.06.08
[5] doi: 10.1016/0260-8774(96)00003-9
[6] NASA GSFC, ChatGSFC, https://chat.gsfc.nasa.gov

Interdisciplinary-02
AI/ML Frameworks for Microgravity Flow Boiling: Identifying, Quantifying, and Generating Two-Phase Interfacial Features
V.S. Devahdhanush and Arkadeep Paul

Interdisciplinary-03
Techno-feasibility systems analysis of carbon dioxide and methane separation systems: From airborne carbon capture to ECLSS applications (VIRTUAL)
Dr. Ashwin Ravichandran, Pooja Santhamoorthy, Stephen Summits, Emmanuel Skountzos, Joakim Halldin Stenlid, David Mebane, Lyndsey McMillon-Brown, and John Lawson

Separating chemical species and extracting or recycling resources from low‑concentration sources is a central challenge across many space applications, including Environmental Control and Life Support Systems (ECLSS), and in‑situ resource utilization (ISRU). Depending on the target chemical species, local environmental conditions, and system‑level constraints, different technologies and architectures may be viable. Selecting the optimal among these options or developing new ones for improved performance requires system-level, physics‑based process models capable of simulating end‑to‑end behavior under realistic operating conditions.
To demonstrate such a framework, we present a feasibility analysis of CO2‑capture and methane‑conversion concepts implemented on mobile airborne capture platforms. Using detailed modeling and optimization, we evaluate the efficiency of these processes across a range of scenarios incorporating different capture and conversion technologies and uncertainties in operating conditions. The technologies studied include liquid‑amine and porous‑adsorbent systems (e.g., zeolites and MOFs) for CO2 capture, as well as catalytic conversion systems for methane, spanning a wide range of relevant concentrations. For each case, the processes were modeled end‑to‑end, including associated equipment and storage units, and system sizing was optimized to meet platform footprint constraints. We also examine future technology‑improvement scenarios to highlight opportunities for increased efficiency and to inform technology development priorities. Overall, this work establishes a framework for techno‑feasibility analysis, considering systems-level constraint, for diverse systems in resource capture, recycle, and utilization.

Interdisciplinary-06
Designing tank outlets to prevent vortex formation and minimize liquid residuals
Mark A Wollen

Reducing liquid residuals when draining propellant tanks is critical for many applications. In launch vehicle boosters and upper stages, propellant flow rates are very high, and the engine’s pumps generally can’t tolerate vapor ingestion. In larger vehicles, insufficient attention to tank outlet design can result in 1000s of kg of residual liquid remaining in a tank when vapor ingestion occurs, and residual liquid mass is even worse than unneeded structural mass, because it is mass that could have been used to make additional impulse. Unfortunately, proper tank outlet and sump design is often overlooked early in the vehicle design process, leading to poorer-than-expected overall vehicle performance. There are two primary causes of vapor ingestion in a draining tank that must be considered. The first is vortex formation, or vortexing, which can cause vapor to be ingested at very high liquid levels, and is usually addressed with some configuration of baffles to dissipate angular momentum as liquid is drawn towards the tank outlet. The second is normal vapor ingestion, or pullthrough, which occurs independent of vortexing, and is primarily minimized by proper outlet sizing and sump design. Normal pullthrough can be predicted with a variety of analytical tools or models, but vortexing is inherently difficult to address with these methods. Fortunately, simple scale model testing has been found to be a reliable and cost-effective means of assessing risk of vortex formation, and for developing baffle and outlet configurations to prevent it. This presentation provides a straightforward analytical methodology for developing an effective tank sump and outlet design. Design and demonstration of an experimental procedure for developing effective anti-vortex baffles is also discussed. An additional benefit of outlets designed to minimize tank liquid residuals is that they generally have minimal flow separation and low inlet pressure loss, which can be particularly advantageous for cryogenic propellants.

framework for techno‑feasibility analysis, considering systems-level constraint, for diverse systems in resource capture, recycle, and utilization.

Interdisciplinary-07
Moon Base Transportation – Deliveries to the Lunar Surface
Dr. Anthony C Terracciano, Maxwell Martin, Charles J Weyandt, John R Campbell Jr, Jared Nardontonia, Richard Nederlander, Clyde O’Quinn, Ben Tutera, and Elaine Voll

Development of the Moon Base will enable a home away from home for astronauts who will live and work at humanity’s first lunar outpost. In this effort, NASA’s Moon Transportation Office is responsible for enabling the transformational missions required to deliver habitats, supplies, science payloads, and all other elements needed to cultivate a permanent presence on the Lunar Surface. The Mission Concept (MC) is characterized through evaluation of an end-to-end architecture that can successfully deliver a generalized heavy large-volume payload, in excess of 4000 kg, to a precision landing and touchdown on the lunar surface.
The mission architecture utilizes a single launch configuration of a Lunar Lander (LL) with a unique propellant system. The LL has an integral orbital transfer capability and features jettisonable elements. The design circumvents the need for prop transfer on orbit and multiple launch configurations. The launch vehicle (LV) for this work will assume the capability to deliver a payload in excess of 40,000 kg to orbit, affording multiple LV solutions. Considerations for the LL and payload deployment from the fairing are assumed to be handled through compliance with a launch providers’ Interface Requirements Document (IRD). The MC will span from launch at Kennedy Space Center (KSC) to terminal descent and touchdown on the lunar surface, requiring a total ΔV on the order of 6 km/s beyond what is required to get the vehicle stack to a 200 km circular Low Earth Orbit (LEO). Major mission phases include: launch and launch vehicle separation, transfer operations, pre-landing navigation, and lunar descent and touchdown.
A Concept of Operations (ConOps) is used as the primary design driver for defining architecture of the vehicles necessary to achieve final payload delivery. Numerous ground rules and assumptions will be provided for each phase of the mission. Concept designs for the LL is presented. An emphasis of the design maximizes a feasible path for maturation, manufacturing, and operation. A self-imposed practical consideration for this effort is the incorporation of legacy designed hardware to minimize expensive, time-intensive, and high-risk hardware development cycles. The propulsion system of the LL adopts a conventional storable bipropellant configuration of monomethyl hydrazine (MMH) and mixed oxides of nitrogen (MON3). This effort will showcase a unique propellant delivery system to minimize the reliance on propellant management devices (PMDs) during descent.
Numerous key constraints have been considered, across the multiple segments of the mission. These include the unique aspects of center of gravity (CG) management, thruster plume effects including self-impingement, propulsion system hardware limitations, navigation during multiple mission phases, and landing gear geometry for uneven terrain. These constraints shape the trades necessary for precision landing of heavy cargo and ensure compatibility with broader Moon Base Transportation concepts.
The resulting insights inform future transportation strategies for the Moon and beyond; directly contributing to the development of cargo‑delivery standards that will support the long‑term buildup of a sustainable, continuously inhabited Moon Base.

Interdisciplinary-08
Thermal Analysis of the Advanced Modular Power Systems(AMPS) Power Electronic Modules
Hashmatullah Hasseeb and Firas G Asfoor

In an effort to transform future space power system architectures and operations, the Advanced Modular Power Systems (AMPS) project is currently developing an interface standard for power electronic modules to enable modular architectures for space power systems. Each architecture would consist of one or more modular electronic units (MEU) that would comprise of a tailored combination of modules. The standard currently comprises modules enabling switchgear, power conversion, and data interfacing. The objective of the AMPS project is to enable the modularity and interchangeability of different architectures using these standardized modules in hopes of reducing complexity and cost, increasing inherent redundancy and reliability, and minimizing power system redesign for future programs.
A key enabler to these future modular architectures will be the thermal management system, which has to contend with acquiring heat from the electronics modules at the card level and then transporting that heat to the radiator of the larger system. The focus of this paper, therefore, will be the thermal design and analysis of multiple AMPS power electronics modules and the development of their thermal-vacuum test.

Interdisciplinary-09
Contact Line Pinning in Microgravity
Joshua L Whitehead, Lucas Hunt, and Ranga Narayanan

Preventing contact line motion is essential for many hydrodynamic systems in microgravity, including systems vulnerable to dry-out, multiphase thermal control systems, and semiconductor manufacturing. Without pinning the contact line, the fluids may spontaneously spread around each other, rather than holding a nominally flat interface. Fluid configurations in microgravity are governed by the system’s free energy, which is minimized at equilibrium. The free energy minimum may be shifted by altering the interfacial energies in the system, e.g., with a chemical coating on the wall. With this method, the system’s free energy minimum can be made to correspond to a favorable contact line position. The effectiveness of this technique is demonstrated experimentally using the Rayleigh-Taylor instability. This instability occurs in a heavy-over-light fluid system when the Bond number exceeds a critical value. This instability is manifested by interfacial wave modes. When the contact line is pinned, the allowable wave modes are restricted, and the critical Bond number is increased relative to the free-slip case. Ongoing experiments have resulted in the interface having the expected wave mode at the onset of instability, and being stable for larger Bond numbers than the free-slip case. The novelty of this work lies in introducing a free energy-based strategy for passive contact line control in microgravity and in providing experimental validation that such a technique can suppress contact line motion. In this talk, I will show the contact line’s dependence on interfacial energies in the limit of negligible gravity and demonstrate its application using a pinned edge, Rayleigh-Taylor experiment.
Funding from: NSF 2422919 is acknowledged.

Interdisciplinary-10
SCIFLI Artemis II Launch Observation Overview
Richard J Schwartz, Jennifer Inman, Kyle Denny, Rufer Shann, Boyda Matt, Meaghan McCleary, and Carey Scott

Ground-based and airborne visual and infrared observations of NASA’s Space Launch System (SLS) were conducted by the Scientifically Calibrated In-Flight Imagery (SCIFLI) team in collaboration with MARS Scientific, the NASA WB-57 aircraft program and a USAF NT-43a operated by Denmar Technical Services during the Artemis II launch. Artemis II was the first crewed flight into deep space in over five decades where the crew aboard the Orion crew module performed a lunar flyby before returning to Earth. Spatially resolved visual and infrared imagery was collected by all the airborne and ground-based assets during the launch and ascent phases of flight while solid rocket booster separation and the jettison of the launch abort system were effectively captured and resolved by a subset of the airborne and ground-based imaging assets. This abstract will provide an overview of the mission operations required to plan, deploy for, and execute the successful observation of the Artemis II launch using three aircraft and six ground-based imaging platforms.

Interdisciplinary-12
SCIFLI Artemis II Reentry Airborne Operations
Arianna R Haven and Robert A Conn

The Scientifically Calibrated In‑Flight Imagery (SCIFLI) team was tasked with developing, coordinating, and executing a multi‑aircraft, long‑range imaging campaign to acquire high‑fidelity, spatially resolved engineering imagery of the Artemis II Orion Crew Module (CM) during atmospheric reentry. The objective of this effort was to provide continuous, calibrated optical measurements from entry interface through splashdown to support post‑flight assessment of the Orion Thermal Protection System (TPS) and the Landing and Recovery System (LRS).
To meet these requirements, SCIFLI designed an airborne architecture consisting of at least four high‑altitude, long‑range imaging aircraft positioned at strategic points along the predicted ground track of the CM. Aircraft selection was driven by operational altitude, endurance, sensor‑accommodation capability, live‑streaming availability, and field of regard relative to CM reentry geometry. Ultimately, six high‑altitude aircraft were selected to participate in the mission to meet all performance requirements and to ensure coverage for early‑return and off‑nominal reentry scenarios. The SCIFLI mission assets included the NASA Gulfstream V, Gulfstream III, and WB‑57 equipped with advanced optics and precision tracking systems, alongside a U.S. Navy (USN) P‑3, U.S. Air Force (USAF) NT-43A, and Missile Defense Agency (MDA) Gulfstream G550.
Aircraft placement was determined through iterative trajectory modeling and line‑of‑sight analysis using mission‑specific reentry predictions. The primary driver was the requirement for continuous observational coverage; however, placement was refined to meet additional time‑specific imaging needs provided by the Orion TPS and LRS teams. These refinements incorporated constraints such as desired viewing angles for char‑loss detection, shock‑layer luminosity characterization, temperature retrieval on the ablative heat shield, and parachute deployment observations. SCIFLI conducted multiple rounds of geometry trades, optimizing altitude, standoff distance, and sensor pointing to preserve image quality while accommodating trajectory uncertainties. During the campaign, each aircraft was assigned to a specific key event along the reentry trajectory, ensuring overlapping coverage and providing redundancy across assets.
Airspace coordination constituted a major portion of the operational planning effort. SCIFLI led the integration of airspace‑reservation requests, flight‑path optimization, and multi‑agency deconfliction across military and civilian airspace authorities. This process included coordination with the FAA and DoD range activities, as well as the execution of high‑fidelity rehearsal flights to validate time‑on‑station predictions and identify potential airspace conflicts or communications issues. Dynamic coordination procedures were established to ensure aircraft could respond to late‑breaking trajectory updates, maintain safe separation, and sustain required viewing geometry under uncontrolled operational constraints such as weather.
While the USN, USAF, MDA, and WB‑57 assets were provided as pre‑integrated, mission‑ready imaging systems—meaning SCIFLI was not involved in payload engineering or aircraft‑level integration—the NASA Gulfstream assets required substantial engineering integration efforts. SCIFLI deployed multiple multispectral payloads paired with specialized single-pane optical windows across the two NASA Gulfstream aircraft, each requiring platform‑specific mechanical and electrical interfacing. Payloads included stabilized tracking gimbals, calibrated multispectral sensors, live-streaming architecture, and data‑recording subsystems. Integration activities required close coordination with aircraft program offices to complete structural analyses, airworthiness documentation, equipment rack analyses, wiring plans, and related tasks. These efforts were followed by installation, sensor‑alignment procedures, and boresight calibrations prior to the first test flight.
The extensive coordination, system preparation, and the operational discipline applied throughout the campaign resulted in the successful collection of expected data from all SCIFLI airborne imaging assets. These data will help NASA better characterize Orion’s in‑flight performance and inform the design of future spacecraft and missions. In this presentation, I will present the underlying analyses, trades, and coordination strategies that informed aircraft selection and positioning, along with the engineering design and integration processes required to field SCIFLI’s multispectral payloads.

Interdisciplinary-14
Participating Media Calculations in Ansys Thermal Desktop
Douglas Bell

Participating media—materials that absorb, emit, and potentially scatter radiation—are increasingly important in modeling high-fidelity thermal environments encountered in space systems, including propulsion plumes, habitat viewports, and in-space manufacturing processes. Accurate representation of radiative transport in these media is essential for predicting heat transfer in environments where traditional surface-to-surface radiation assumptions are insufficient. To address this need, participating media radiation modeling capabilities have been implemented in Ansys Thermal Desktop, extending its applicability to a broader class of aerospace thermal problems.
This paper will present a comprehensive validation and verification effort for the implementation. Comparisons against analytical closed-form solutions will be used to establish baseline accuracy, while systematic grid convergence studies will quantify numerical error and solution stability across varying discretizations. The sensitivity of results to optical properties and model resolution will also be discussed. Additionally, representative application cases relevant to NASA missions will be presented, illustrating the impact of participating media on predicted temperature distributions and heat fluxes.
The results of this evaluation should demonstrate strong agreement with analytical solutions and expected convergence behavior, providing confidence in the robustness and accuracy of the methodology. Best practices will be identified to guide analysts in incorporating participating media effects into spacecraft and ground test thermal models.

Interdisciplinary-16
Thermal Modeling Philosophy for the Dragonfly Lander: Integrating Thermal Desktop and CFD for Titan Surface Operations
Dr. Hui Liu, Evan Cosentino, Kurt Gonter, and Jane He

NASA’s Dragonfly mission will deliver a rotorcraft lander to the surface of Titan, an environment that presents thermal challenges fundamentally different from those of traditional spacecraft. Unlike vacuum-based missions, Dragonfly operates within Titan’s dense atmosphere, where convection dominates both internal and external heat transfer. The lander uses MMRTG waste heat, active fan-driven air circulation, cold duct thermal trim device, and foam insulation to maintain component temperatures during both operational and hibernation conditions. These characteristics require a thermal modeling approach that accurately captures airflow-driven heat transfer while remaining computationally efficient for mission-level thermal analysis.
This presentation describes the thermal modeling philosophy developed for Dragonfly and the integration of Thermal Desktop (TD) and Computational Fluid Dynamics (CFD) models for flight thermal analysis. Thermal Desktop serves as the primary thermal prediction tool due to its efficiency in evaluating large mission case matrices and its spacecraft thermal analysis heritage. High-fidelity CFD models developed in ANSYS Fluent are used to explicitly resolve airflow, convection, and thermal-fluid interactions within the lander. The two modeling approaches are cross-anchored using common geometry, material properties, heat loads, and environmental conditions, allowing CFD-derived airflow physics to inform TD convection correlations while TD provides rapid system-level thermal assessments.
A unique aspect of this work is the development of a structured TD/CFD integration methodology for a convection-dominated planetary spacecraft. Rather than treating CFD solely as a one-time validation, TD and CFD models are benchmarked against each other, with ongoing validation against system-level test data to ensure consistent predictions of heat transfer, temperature distributions, and transient thermal response. Results to date show strong agreement between the two models, increasing confidence in Dragonfly’s Titan surface thermal predictions. The comparisons also highlight the importance of accurately characterizing convection-driven heat transfer and airflow distribution in the lander’s thermal control system.
This approach combines the physics fidelity of CFD with the efficiency, heritage, and partner accessibility of TD. The integrated TD/CFD methodology may also provide a practical framework for thermal analysis of future atmospheric planetary missions where convection plays a dominant role.

Interdisciplinary-17
Machine-Learning Coupled Phase Change Closure Model Based on Interfacial Heat Flux Discontinuity
Dr. Sunjae Kim and Yusin Jeong

Two-phase boiling heat transfer underpins some of the most demanding thermal management technologies in modern aerospace engineering, such as spacecraft thermal control loops, cryogenic propellant management, and high-power electronics cooling. Despite this broad relevance, its predictive simulation has remained one of the most long-standing open problems in Computational Fluid Dynamics (CFD). The difficulty originates in the nature of the liquid-vapor interface itself, a freely deforming and continuously moving surface whose position, shape, and topology are part of the unknown flow field. Imposing a physically faithful phase-change condition on such an interface is inherently demanding, and existing closures can be classified precisely by how they attempt to do so.
Conventional closures fall into two families, distinguished by how they represent the thin thermal boundary layer (TBL) adjacent to the interface whose internal temperature gradient drives the phase change. Scalar-based models, such as the Lee and Hertz-Knudsen-Schrage (HKS) families, compute the evaporation rate from a single scalar deviation field multiplied by an empirical relaxation coefficient, and although computationally cheap, they bypass the TBL physics entirely and yield strongly case-dependent, grid-sensitive predictions. Vector-based models instead evaluate the evaporation rate directly from the heat-flux-jump condition, a physically rigorous formulation that nonetheless demands TBL-resolving meshes around every interface element together with one-sided gradient stencils that cross the interface, making it prohibitively expensive. Researchers are forced to choose between an inexpensive but unreliable scalar closure and an accurate but unaffordable vector closure. Thus, the present work aims to dissolve this long-standing trade-off by training a feed-forward neural network on data produced by a high-fidelity vector-based model and coupling the trained network in real time to the running CFD solver. The resulting phase-change closure operates at the per-step cost of a scalar-based model while retaining the accuracy of a vector-based one.
The vector-based baseline is built around a coupled level-set volume of fluid (CLSVOF) interface representation, with the incompressible Navier-Stokes, energy, level-set, and volume-fraction transport equations solved through a PISO velocity-pressure coupling and all interfacial source terms supplied by an in-house UDF. The principal numerical obstacle of one-fluid formulations is the accurate evaluation of the interfacial heat flux, because the temperature stored at a mixture cell containing the interface is a volume-fraction-weighted average of the local liquid and vapor temperatures rather than the temperature of either pure phase, so that its gradient reflects only this averaged field and cannot be decomposed into the distinct one-sided liquid and vapor gradients required by the heat-flux jump.
This obstacle is overcome by a normal probe method coupled with a pathfinding algorithm. The probe samples one-sided temperatures along the interface normal into pure-liquid and pure-vapor regions, while the pathfinding algorithm locates valid stencil cells across complex interface topologies, producing clean estimates of the one-sided liquid and vapor temperature gradients required by the heat-flux jump. The solver is validated against the analytical Stefan solution in both vapor-driven and liquid-driven configurations. Expanding the validation effort, the solver was evaluated for film boiling simulations and successfully reproduced bubble-growth data of Phan et al. (2025) and the experiment of Kilimenko et al. (1981) with qualitative agreement in bubble morphology and quantitative agreement in growth-rate evolution and averaged Nusselt number.
Despite this accuracy, the per-cell pathfinding imposes two practical limitations. First, because the pathfinding stencils cross MPI partition boundaries, the heat-flux evaluation cannot be parallelized through domain decomposition and is restricted to single-process execution. Second, the pathfinding, repeated at every interface cell and time step, dominates the overall wall-clock cost. A data-driven closure that supplies the interfacial heat flux without per-cell search at run time is therefore required.
The trained network replaces the explicit TBL-resolving computation of the interfacial heat flux. Coupled in real time to the CFD solver, it supplies the interfacial source term at every time step without the explicit, per-cell heat-flux evaluation that the vector-based baseline requires. This formulation simultaneously eliminates two practical bottlenecks of the vector-based approach, namely the prohibitive cost of TBL-resolving meshes around every interface element and the loss of parallel scalability when normal-probe stencils repeatedly cross MPI partition boundaries. The ML-coupled solver delivers a 5.5x wall-clock speed-up over the physics-based baseline at comparable accuracy, and its predictions are shown to be independent of the grid resolution, the imposed wall thermal boundary condition, and the gravitational configuration, indicating that the network has learned a generalizable representation of TBL physics rather than memorizing a single operating point.
This work delivers three principal contributions. First, the physically exact heat-flux-jump condition is used as the evaporation closure in place of empirical mass-transfer coefficient. Second, a Normal Probe Method with pathfinding renders the vector-based approach robust on arbitrary interface topologies within a one-fluid framework. Third, a data-driven phase-change closure model is accomplished which not only accelerates phase change computation time but also provides independency against grid resolution, wall boundary conditions, and gravity.

Interdisciplinary-18
Physics-Based Postprocessing of 3-D CFD Results
Dr. Bijay K. Sultanian

Using modern, high-power computing resources and user-friendly commercial and in-house CFD codes, CFD design engineers can quickly compute 3-D CFD results in a complex, high-fidelity computational domain representing a flow device. These results consist of detailed distributions of the primitive variables like velocity, static pressure, and static temperature throughout the computational domain. To compute the section-average value of any of these quantities, the question arises whether one should compute the area-weighted-average value or the mass-weighted-average value of the quantity. This presentation discusses a physics-based method to compute an accurate section-average value of any variable in the 3-D CFD results. Using a power-generation gas turbine exhaust diffuser as an example, the presentation shows how to accurately compute its pressure-rise coefficient when the diffuser has non-uniform flow properties at its inlet and outlet. The proposed methodology is presented for both incompressible and compressible flows.

Interdisciplinary-19
Closed-Cycle Transpiration-Cooled Flash Combustor with Cryogenically-Coupled Water Recovery for Thrust Augmentation in Pre-Cooled Airbreathing Vehicles
Steven T Schryba

This presentation introduces a Thrust Augmentation Flash Combustor (TAFC) architecture developed for a cryogenically-precooled, multi-domain hybrid vehicle. The TAFC integrates three coupled thermal-fluid loops into a single closed-cycle propulsion subsystem: (1) a liquid hydrogen (LH₂) transpiration-cooling circuit through a dual-layer porous wall, (2) a cryogenically-coupled steam condenser that recovers combustion product water using LH₂ as the cold sink, and (3) a recirculation manifold that returns condensed H₂O to either secondary cooling, thrust re-injection, or vehicle reserve storage. The flash chamber wall is a bonded composite of porous ceramic matrix composite (C/SiC) inner liner and porous nickel-superalloy outer structure joined by a refractory metal interlayer, with controlled-porosity micro-perforations sized to deliver LH₂ transpiration cooling at peak heat fluxes approaching 50 MW/m². The same LH₂ that cools the wall is then injected as fuel — the coolant and the fuel are the same fluid, exploiting hydrogen’s wide flammability limits and high specific heat.
The architecture is intended for phase-selectable operation: short-takeoff/landing, supersonic dash, and high-altitude transition modes. It is not engaged during cruise or vacuum-phase operation, where the parent vehicle’s baseline cryogenic propulsion cycle is more efficient.
Unique Aspects of the Work
Three elements distinguish this concept from existing pre-cooled airbreathing engine work (e.g., precooled turbojets, rocket-based combined cycle architectures, and published synergetic air-breathing rocket studies):
Fuel-as-coolant transpiration through a hybrid CMC/superalloy porous wall. Prior transpiration-cooled combustor work has generally used monolithic porous materials. The proposed dual-layer construction places C/SiC at the gas-side interface to handle the 2,200 K combustion temperature and resist hydrogen embrittlement, with porous Inconel as a structural backing that manages coolant manifolding and pressure loads. The refractory interlayer accommodates CTE mismatch over the 79 K to 2,200 K thermal cycling range.
Cryogenically-coupled water recovery. The TAFC condenses combustion-product steam back to liquid using the same LH₂ supply that feeds the combustor, capturing latent heat of condensation as useful pre-warming of the fuel. Recovery rates between 55% and 85% are projected depending on mission phase, with the highest recovery occurring at high altitude where ambient conditions favor condensation.
Closed three-loop integration. The combustor coolant loop, the air precooler loop, and the water recovery loop share a single LH₂ source and a coordinated thermal management network. This integration is the central novel claim — it has not been demonstrated in any precooled airbreathing engine architecture to date.
Available Data, Results, and Conclusions
First-order analytical results to be presented include:
Three operating-point performance maps (STOL, supersonic dash, high-altitude transition) with chamber pressures of 28–45 bar, mixture ratios of 5.8–7.2 (air/LH₂ mass), and thrust augmentation factors between 1.35× and 2.10× the baseline cryogenic cycle.
A peak wall heat flux estimate of approximately 52 MW/m² in dash mode, with a transpiration coolant flow fraction of 1.8–2.4% of core mass flow.
A water recovery accounting showing 2.8–4.5 kg/s recovered H₂O across operating points, against an available LH₂ heat-sink capacity of approximately 11 MW.
A subsystem mass estimate of approximately 1,232 kg, against a parent-vehicle mass margin of 4,275 kg.
A failure-mode survey covering transpiration channel blockage, condenser frost-up, and ignition reliability under fuel-rich start conditions.
Conclusions to date: the cycle closes thermodynamically at all three design points, mass margin is preserved, and the principal technical risks are concentrated in (a) demonstrating cyclic durability of the porous CMC/superalloy wall and (b) achieving stable condenser operation across the full ambient envelope.
Status of the Work
The work is currently in the concept and analytical development phase. Activities completed: cycle analysis, mass budgeting, performance-point characterization, materials trade study, failure-mode identification, and patent novelty assessment. Activities planned but not yet started: detailed CFD of the transpiration wall, sub-scale porous wall coupon testing, condenser heat-exchanger sizing CFD, and ignition demonstration testing. No hardware has been fabricated. No test data exists. All numerical results presented are first-order analytical estimates and are presented as such.
CUI/ITAR/EAR Status
This abstract contains no CUI, ITAR, or EAR material. The author requests consideration for either the open or the restricted session as the program permits.

Interdisciplinary-20
Improved post-processing workflow for space thermal design insight
Joel Gagnon, Jean-Frederic Ruel, and Florian Sanchez

The renewed global interest in large satellite constellations for high-bandwidth communication has accelerated the demand for accurate, efficient, and traceable thermal engineering tools for space programs. This paper presents recent developments in the TMG solver for thermal analysis workflows supporting the ongoing satellite communication space race, with emphasis on constellations operating in Very-Low Earth Orbit (VLEO). The study focuses on thermal design assessment during critical mission phases, including launch and early orbit, where rapid environmental transitions and coupled conductive-radiative effects can drive design margins and hardware qualification decisions.
A key contribution of this work is the development of a built-in reporting tool for allowable temperature margins within Simcenter 3D Space Systems Thermal. The tool enables thermal analysts to rapidly evaluate component-level temperature compliance across steady-state and transient load cases, identify hot and cold margin drivers, and communicate results in a structured and repeatable format. By consolidating margin calculations, allowable temperature limits, and case-by-case thermal results, the reporting capability reduces manual post-processing effort and improves consistency across design reviews, subsystem assessments, and program-level reporting.
The paper also presents a new graphical heat-flow representation capability that supports interpretation of conductive and radiative energy exchanges within the thermal model. Used together, the temperature-margin reporting tool and heat-flow visualization provide analysts with a more efficient way to diagnose critical thermal paths, understand the impact of boundary conditions, and prioritize design changes early in the development cycle. These capabilities are particularly valuable for spacecraft programs requiring frequent design iterations, large numbers of analysis cases, and clear traceability between simulation results and thermal requirements.
Representative steady-state and transient analyses are presented for launch and early-orbit scenarios relevant to VLEO communication spacecraft. The results demonstrate how automated margin reporting and heat-flow visualization can support rapid assessment of thermal compliance, improve communication between thermal analysts and system stakeholders, and strengthen confidence in early thermal design decisions. These developments enhance the efficiency, reliability, and scalability of spacecraft thermal control system analysis, contributing to high-throughput workflows needed for future satellite constellation programs.

Interdisciplinary-21
Predicting Operational Performance of xEMU Boot at Lunar South Pole Temperatures using Thermal Desktop® (VIRTUAL)
Emma J Quick and Noah Andersen

The spacesuit boots that will be used on Artemis lunar south pole surface missions will be exposed to extremely cold temperatures (down to ~50 K). To assess the performance of the government’s Exploration Extravehicular Mobility Unit (xEMU) lunar boot in these permanently shadowed regions, testing was performed at the Jet Propulsion Lab (JPL) in the Cryogenic Ice Transfer, Acquisition Development, and Excavation Laboratory (CITADEL) thermal vacuum (TVAC) chamber. This paper documents the data analysis, thermal boot model correlation, and operational predictions conducted using data from the xEMU CITADEL TVAC test. Expected thermal conductances within the boot and between the boot and environment were calculated from test data, which was then used as an initial guess for conductances within a Thermal Desktop (TD) model. Correlation of the TD model using the internal SOLVER feature was performed across 10 different test points which varied external temperature, internal boot ventilation flowrate, and contact pressure. Operational performance at the lunar south pole was then predicted using results from the correlated model. While the predictions provide evidence for acceptable performance of the boots at the 100K environment test point, there is still substantial uncertainty in performance, especially at the 48K test point. This uncertainty is due in part to testing limitations such as contacting the foot to a hard metal plate rather than granular regolith, and model limitations such as the lack of a realistic foot model. These limitations and their impacts are addressed in detail in this paper. The results of this test series and model correlation underscore the importance of additional improved testing and modeling for characterizing the expected thermal resistance between the outside of the boot and the lunar surface.

Thermal Control and Protection-01
Design & Thermal Analysis of Hyperion: The University of Tennessee’s New 1 MW Continuous Arcjet Facility
Zachary Colovos, Damiano Baccarella, Killian Samuels, Mitchell Trotsky, and Christian Isaacs

The University of Tennessee’s new 1 MW segmented arcjet facility, Hyperion, has been fully designed and is currently under construction as major components are sequentially delivered. Hyperion incorporates the largest and most powerful segmented arc heater in academia and is intended to operate as a continuous, high-enthalpy ground-test facility capable of multi-minute run times. This capability directly addresses a persistent gap between shorter-duration, small-scale university facilities and large-scale arcjet wind tunnels, significantly expanding the range of accessible test conditions for hypersonic research. In particular, Hyperion will ultimately enable high-fidelity aerothermal characterization, advanced diagnostic development, and testing of TPS materials in conditions near planetary entry enthalpies for the industry. Our presentation introduces Hyperion’s architecture, including its major subsystems such as the gas, power, and active thermal management subsystems, as well as its overall testing envelope. Design and thermal performance analyses of each tunnel component is then detailed, including the new segmented arc heater, test section, supersonic diffuser, and heat exchanger. This overview is followed by an update on the manufacturing and assembly status of these novel components and certain commercially available systems. Hyperion is currently in the midst of component fabrication, subsystem installation, infrastructure renovations, and preparations for the extensive water routing for thermal management across the facility. Overall system integration and initial operational testing will occur later this year.

Thermal Control and Protection-02
Advanced and Efficient Thermal Model Correlation of a Satellite Electronics Box Using HEEDS and Thermal Desktop
Dr. John M Pederson, Lina Li Maricic

Correlating thermal models to data obtained from thermal vacuum (TVAC) testing, an essential step in model validation, is often time-consuming to complete by hand, and much more so with large numbers of parameters and temperature measurements. This study introduces a method leveraging the Siemens Simcenter HEEDS multidisplinary design optimization (MDO) software, integrated with Ansys Thermal Desktop, to perform such a parameter correlation in an automated and efficient manner. In this method, the problem is posed as a HEEDS optimization design study in which the inputs are the thermal parameters relevant to the simulation, the outputs are the temperature responses yielded from solving the model, and the constraints are the ranges allowed for each input and output (i.e., the temperature tolerance to which the model must be correlated). The optimization objective function minimizes the differences between simulated temperatures and test data. The HEEDS workflow leverages a MATLAB script that directly modifies model parameters and extracts simulated temperatures via OpenTD, the application programming interface (API) for Thermal Desktop; this process allows HEEDS to rapidly define and run thousands of simulations to search the parameter space, converge on one or multiple feasible solutions that match the collected experimental data to the specified tolerance (typically less than 3–5 °C), and collect the simulated responses to inform the next design candidate. HEEDS then employs adaptive global and local search algorithms to generate new design candidates, optimizing a weighted sum of all objectives until the specified number of simulations is reached. This approach was tested on a thermal model of a communication satellite’s electronics box (E-box), the technical details of which are omitted to protect customer-proprietary information. Thirteen temperature sensors were installed on and inside the chassis, with initial model-to-test temperature differences of more than 10 °C. A total of nineteen model parameters were adjusted to match the model predictions to the thirteen thermocouple responses across the two (hot and cold) TVAC thermal-balance conditions, meeting the correlation criterion in one week; this was significantly faster than the typical manual process (requiring an estimated three to four weeks) and required much less human-in-the-loop time. This method thus demonstrates substantial efficiency gains and provides a repeatable framework for thermal model correlation that is applicable to a wide range of spacecraft thermal validation efforts. The efficiency of this method was further demonstrated through a successful correlation of a large system-level payload model to a TVAC test across three thermal-balance conditions, completed within a two-week timeframe. An additional advancement was made to recycle previous HEEDS runs that recorded different temperature outputs, significantly reducing the number of HEEDS evaluations and thus the computation time required. This capability enables the optimization to begin running ahead of time regardless of the final thermocouple locations, which allows the method to integrate into a more realistic engineering environment where temperature measurement locations are not always known in advance and test schedules are often subject to delays.

Thermal Control and Protection-03
Europa Clipper Thermal System Performance In-Flight vs. Predictions
Hared A Ochoa

The Europa Clipper Mission Launched on October 14, 2024 and begun its 5.5 year cruise towards Jupiter. Clipper will be the first flight system to execute dedicated science investigations of the Jovian moon Europa. Through its mission, Clipper will be exposed to a wide range of thermal environments, including ~0.82AU perihelion, 5.6AU maximum sun range, and a 9.2 hour Jupiter Eclipse. Additionally, due to the large sun range during its science campaign and its use of Solar Arrays as the power source, the thermal design is intended to be power efficient. Before Launch, the Clipper spacecraft went over a thorough environmental test campaign including system TVAC (STV) for thermal system verification and validation. While the test was designed to be as flight-like as possible, certain test constraints (facility size, hardware availability, etc.), led to some difference from the flight configuration. This paper provides a summary of the STV campaign, the pre-flight expected performance of the thermal subsystem, and comparison to early performance in flight. Findings are summarized along with future work and major lessons learned for the program.

Thermal Control and Protection-04
FROSTE: Thermal Design and Thermal Vacuum Testing of Cryogenic Stowage for Lunar Sample Return
Faiyaj Khan, Elijah Stewart, Claire Silaire, Kacy Vanden Bergh, and Erin Hayward

The Frozen Return of Samples to Earth (FROSTE) project is developing a capability to return lunar samples from the south pole targeting to maintain cryogenic temperatures for lunar regolith from collection, transit, re-entry, to final delivery to the lunar surface sample procurement facility. The target is to maintain the frozen volatiles within the regolith in their collected state. The system must operate in several different environments including but not limited to within crewed vehicles, lunar surface locations in the sun, permanently shaded regions (PSRs), transit, and re-entry into Earth’s atmosphere. The thermal design must be robust to maintain sample temperatures around 80 K both with active thermal management using Stirling cycle cryocoolers, and passively in periods of limited available power using a phase change material (PCM), multilayer insulation (MLI), and a structural design approach to limit conduction into the sample. The project is currently testing a sub-scale version of the stowage container which will be put under various thermal environments including a proxy lunar PSR environment and cabin environment. Unique hardware was used for this test such as integrated MLI, Stirling cycle cryocoolers, and a fully additively manufactured test article. This paper will discuss the thermal analysis done on the sub-scale container, preliminary results from the testing, explanation of the test hardware and observations in using it, lessons learned from handling the hardware, and future goals such as how the test data will feed into a full-scale thermal model.

Thermal Control and Protection-05
Development and Testing of Additively Manufactured Phase Change Material Thermal Storage Units
Dr. Michael Pauken, Jonathan Lam, Ryan Watkins, and Djuna Copley-woods

Phase Change Material (PCM) Thermal Storage Units (TSUs) are commonly used in spacecraft thermal control systems to absorb transient heat loads while maintaining tight temperature limits. Traditional TSU designs often rely on complex assemblies consisting of machined housings and high-conductivity internal structures, resulting in long fabrication schedules, high cost, and limited design flexibility. This work investigates the feasibility of using additive manufacturing (AM) to produce an integrated PCM TSU with embedded lattice structures that provide both structural support and enhanced thermal transport.The additively manufactured TSUs were fabricated from Al6061-RAM2 and designed to replicate the envelope and interface geometry of the flight-proven SHERLOC TSU developed for the Mars 2020 mission. Several internal lattice geometries were investigated, including cross-lattice, tapered cross-lattice, porous cross-lattice, and diamond lattice configurations, with the objective of maximizing heat transfer surface area while minimizing mass and maintaining manufacturability. Six prototype TSUs were fabricated and filled with n-decane PCM using a vacuum-assisted fill process.The units were heated with film heaters, instrumented with thermocouples, and tested in a thermal vacuum chamber over multiple thermal cycles between approximately -50°C and +20°C. The testing characterized latent heat storage performance, thermal gradients, repeatability, and the effects of PCM retention during cycling. Results demonstrated that the AM approach is feasible and capable of achieving thermal energy storage performance comparable to or exceeding the heritage design while significantly simplifying manufacturing. Several of the tested lattice configurations achieved thermal storage capacity greater than the SHERLOC reference unit. The testing also identified important lessons learned related to PCM fill methodology and sealing approaches for additively manufactured hardware. In parallel with testing, detailed thermal models were developed to predict transient thermal behavior and latent heat storage capacity. Correlation between analytical predictions and experimental results is discussed, along with the influence of lattice geometry on thermal performance. The results demonstrate the potential for additively manufactured PCM TSUs to reduce cost and lead time while enabling scalable thermal energy storage solutions for future planetary and spaceflight applications.

Thermal Control and Protection-06
Structural Performance Evaluation of Titanium-Water Heat Pipe – GFRC Radiators for Fission Surface Power
Dr. Sandeep Hatte, Zayed Ahmed, Mason Pratt, Calin Tarau, Jeffrey Diebold, Srujan Rokkam, Tyler Beach, and Jeffrey Hopkins

Advanced Cooling Technologies, Inc. (ACT), under a SBIR Phase III program conducted structural and thermal assessment of titanium-water (Ti-H2O) heat pipe and Graphite Fiber Reinforced Composite (GFRC) face sheet-based radiators for Fission Surface Power (FSP) applications. ACT Previously developed and fabricated radiator panels by directly bonding GFRC face sheets to titanium-water heat pipes. The heat pipes incorporated a proprietary wick structure within the evaporator and non-condensable gas (NCG) to provide freeze-thaw tolerance. The GFRC face sheets featured a unique alignment of carbon fibers to cancel the Coefficient of Thermal Expansion (CTE) mismatch with titanium in the axial direction.
Recently, ACT performed numerical analysis and experimental demonstration of structural integrity of the radiators under launch vehicle shock and vibration environments. The numerical analysis identified optimized structural reinforcements configurations that exhibited successful structural response to random vibration loads corresponding to launch conditions. Following the fabrication and assembly of structural reinforcements, radiators will undergo experimental shock and vibration testing. Thermal testing will be conducted before and after the shock and vibration testing and will be used to evaluate the response of radiators to launch vehicle loads. In addition, face sheet material and its bonding with titanium heat pipes will be characterized for any structural damage. Success of shock and vibration testing will be evaluated from the structural integrity of the radiators, and their power rejection capability. In addition, ACT conducted detailed freeze-thaw tolerance performance testing of the heat pipe using two different freezing methodologies. First, in the case of freezing only the condenser; whereas the evaporator remained warm, the heat pipe showed 12 successful freeze-thaw tolerant cycles of testing without failure. In the second case of freezing the entire heat pipe, a total of 10 successful freeze-thaw tolerance tests were conducted without any failure. Following each freeze-thaw cycle, the heat pipe regained its performance, and the radiator module rejected same amount of power at similar operating conditions. Finally, ACT performed a theoretical assessment of the probability of heat pipe failure and face sheet area loss resulting from micrometeoroid impacts over a 10-year operational lifetime in the lunar environment.

Thermal Control and Protection-07
Fabrication and Testing of Additively Manufactured Titanium – Water Loop Heat Pipe for Space Applications
Dr. Sandeep Hatte, Calin Tarau, Jeff Diebold, Srujan Rokkam, Lida Yan, and Ying Sun

Next-generation space missions require efficient, lightweight, and highly reliable passive thermal management systems to support extended operational lifetime. Loop Heat Pipes (LHPs) are a key thermal management technology for satellites, planetary orbiters, and surface exploration missions due to their ability to passively transport heat over long distance with minimum temperature drop. In this work, Advanced Cooling Technologies (ACT) and its collaborators have developed additively manufactured titanium-water loop heat pipes. LHP evaporators have been fabricated using Laser Power Bed Fusion (LPBF) technique. ACT carried out titanium LHP fabrication using additively manufactured titanium evaporator. The titanium evaporator consists of an integrated fine-pore primary wick fabricated directly through additive manufacturing, whereas a secondary wick was fabricated using a coarse-pore screen of titanium. Additionally, ACT conducted theoretical analysis to determine the LHP power carrying capability for a range of operating temperatures from 100 °C to 125 °C. Following the fabrication, ACT will conduct thermal testing to demonstrate the working of a one-of-a-kind additively manufactured titanium-water LHP. Initial testing will focus on transporting nominal heat loads on the order of several hundred watts to demonstrate stable LHP operation. Additional testing will investigate startup behavior at low powers, thermal transport characteristics, and effective thermal conductivity, while experimental results will be compared against analytical predictions. The results will provide insight into the fabrication feasibility and performance of additively manufactured titanium-water LHPs for next-generation space applications.

Thermal Control and Protection-09
Experimental Characterization of Additively Manufactured Nickel-Titanium Shape Memory Alloy Heat Pipes
William R Sixel, Bilal Bomani, Tomé Seichi da Nóbrega Guenka, Christopher Greer, and Ryan Overdorff

Shape memory alloys (SMA) have been identified for use in spacecraft components as replacement for conventional deployment mechanisms. They may be used in thermal management components such as radiators to create self-deploying radiators. One SMA, NiTi, has also been developed for additive manufacturing processes. Heat pipes are a common way to create highly effective and lightweight spaceflight radiators, and heat pipes can also be made from NiTi and related alloys. The wick is the critical element of a functioning heat pipe, and recent progress over the past years has led to the development of additively manufactured heat pipe wicks in various materials. The combination of these efforts is the focus of this project: creating an additively manufactured, shape memory alloy self-deploying heat pipe radiator. This paper will focus on the experimental characterization of these additively manufactured NiTi heat pipes. The heat pipe coupons were additively manufactured by direct metal laser sintering (DMLS), with an integral liquid cooled condenser. Heat is input to the heat pipe via a thin film heater. Thermocouples were spot welded to the heat pipes to measure temperature at several axial locations. The heat pipes were tested with two working fluids: water and ethanol. Ethanol is not an ideal working fluid for heat pipes but is useful in characterizing them because it wets well to a wide variety of surfaces. Water is in general a superior working fluid for heat pipes, but its contact angle and therefore wicking performance strongly depends on the surface chemistry of the surface it is in contact with. A particular measurement of interest in this test is the evaporator to condenser thermal conductance, which will be compared in the full paper to recently published correlations for additively manufactured heat pipes. Experimental results for two straight geometry and one bellows geometry heat pipe will be presented. The bellows geometry is of interest for condenser of the self-deploying radiator design.

Thermal Control and Protection-10
Variable Heat Rejection System for Space Habitat
Dr. Calin Tarau, Ramy Abdelmaksoud, Jeff Diebold, Srujan Rokkam, William Johnson, and Justin Boyer

NASA seeks new thermal control technologies that will enable crew habitats to survive and operate through the lunar night for extended periods of time on the lunar surface, where temperatures range from -193°C or lower in shadowed regions (including night) to 120°C at the equator. In addition, the solicited technologies would allow habitats or pressurized rovers to operate in all these environments, as well as deep space and transit from Earth to the Moon. This presentation reports on the development of a high turndown ratio, freeze-tolerant, and passive Thermal Control System for Planetary Surface and Space Habitats based on multiple Non-Integrated Hot Reservoir Variable Conductance Heat Pipes that cool and control thermally a benign fluid-based single-phase loop that collects waste heat from the habitat. In addition, the developed system is capable of active set point maintenance by using severely minimized control power applied to the reservoirs. The feasibility of the system was demonstrated during this work, both mathematically and experimentally, where an experimental turndown ratio of 52.6 was measured.

Thermal Control and Protection-11
Multifunctional Pulsating Heat Pipes for Electric Motors
Dr. Jeff Diebold, Jacob Sonnek, Parag Bajaj, and Arijit Banerjee

This presentation discusses an innovative new approach to electric motor thermal management. Pulsating heat pipes, passive two-phase heat transfer devices made from copper capillary tubing, replace the conventional electrical windings within the motor. The multifunctional pulsating heat pipe simultaneously carries electric current and serves as its own thermal management device. The motion of the two-phase working fluid within the winding effectively transports heat, dissipated in the copper envelope, with an effective thermal conductivity >10x that of copper. This presentation presents a preliminary electric motor design utilizing hollow conductors, thermal characterization of a pulsating heat pipe winding prototype, preliminary thermal analysis of a motor utilizing the new approach, and an experimental demonstration of a multifunctional pulsating heat pipe carrying alternating current.

Thermal Control and Protection-12
Recent Development of High-Temperature Alkali Metal Heat Pipes at ACT
Dr. Jeff Diebold, Calin Tarau, Sandeep Hatte, Roopesh Kumar, Nathan Van Velson, and David-Paul Schulze

Alkali metal heat pipes are passive two-phase heat transfer devices that utilize alkali metals such as potassium and sodium to transfer heat at high temperatures (>300°C). Advanced Cooling Technologies (ACT) has a long history of developing unique high-temperature alkali metal heat pipes for wide variety of applications such as nuclear power, solar power, and thermal energy storage. This presentation will highlight several recent RD programs at ACT to develop high-temperature and high-power alkali metal heat pipe systems. Highlighted programs include:
• The development of a unique high-temperature multi-material heat pipe system for transporting heat from a nuclear reactor to a Stirling engine as part of the NASA/DOE FSP program. The system used sodium-Kovar thermosyphons and a sodium-Haynes 230 loop thermosyphon.
• Active development of high-power wicks for alkali metal heat pipes for cooling nuclear reactors. Wicks have shown >10x improvement over conventional wick design.
• Large high-power heat pipes and thermosyphons, power > 10kW.
• A high-temperature (>600°C) thermal storage system using alkali metal heat pipe technology.
• An alkali metal heat-pipe heat spreader and pulsating heat pipe for a particle receiver board in concentrated solar power that reduced peak temperature by 130°C.
For each program, the application, design, prototype, and experimental results will be highlighted.

Thermal Control and Protection-13
Architected Lattice Wicks for Monolithic Conformal Heat Pipes and Vapor Chambers—Characterization: Permeability, Capillary, Surface Area, Thermal, and Mechanical
Dr. Robin Pham and Presented by Calin Tarau

Conventional heat pipes and vapor chambers rely heavily on sintered, screen, grooved, or otherwise stochastic wick structures whose capillary and hydraulic behavior is often coupled to manufacturing constraints rather than independently designed. Additively manufactured lattice wicks provide a path toward architected capillary media in which permeability, capillary pressure, effective pore size, phase-change surface area, structural contribution, and printability can be co-designed for the thermal-fluid requirements of the device. Unlike inserted sintered wicks, printed lattice wicks being materially dense components can also function as load-bearing porous structures, enabling thinner, lighter, and more integrated two-phase thermal devices – capable of forming into any complicated shape and size, as previously shown in our past hardware demonstration. This presentation summarizes ongoing characterization of lattice wick structures for monolithic heat pipes, vapor chambers, and conformal two-phase thermal spreaders.
The work focuses on lattice wicks as a functional replacement and extension of traditional sintered wicks. Candidate lattice topologies are evaluated using wick-relevant metrics, including permeability, capillary pressure, effective pore radius, liquid-return capability, manufacturability, and mechanical contribution to the pressure-containing structure. Particular emphasis is placed on the difference between nominal lattice geometry and as-printed hydraulic behavior. Metal additive manufacturing enables highly ordered wick morphologies, but overhang severity, incomplete fusion, sagging, pore occlusion, and local print defects can significantly affect the resulting capillary network. The presentation will discuss how these effects influence topology selection and how print-aware lattice design can improve wick performance and repeatability.
The characterization effort is connected to a demonstrated monolithic, additively manufactured titanium-water vapor chamber with integrated lattice wick structures and a fully connected three-dimensional flow network. This hardware serves as a system-level proof case for the lattice wick approach. The device demonstrates how the wick, vapor space, pressure envelope, and thermal-control surface can be integrated into a single printed two-phase component rather than assembled from separate heat pipes, spreaders, inserted wicks, or joined thermal interfaces. In this architecture, the lattice wick is not only a capillary return structure; it is also a designable phase-change surface and a potential structural element within the printed pressure boundary.
The broader implication is that lattice wicks can enable conformal, geometry-specific two-phase thermal devices for spacecraft systems where planar heat spreaders or one-dimensional heat pipes are difficult to integrate. In the longer term, print-aware lattice design workflows and higher-resolution metal additive manufacturing processes may support direct integration of thermal-control structures with compact devices and miniature spacecraft platforms. The objective of this work is to establish architected lattice wicks as a practical design platform for next-generation monolithic two-phase thermal hardware, while identifying the characterization data and design rules needed to mature the technology for mission-specific applications.
What Is Unique About the Work
This work treats the wick as an engineered, architected thermal-fluid-structural component rather than as an inserted porous material. The novelty is the combined characterization of lattice wick hydraulic performance, capillary behavior, phase-change surface area, manufacturability, and structural contribution, followed by demonstration in a monolithic three-dimensional two-phase thermal device.
Available Data, Results, and Conclusions
Available results include printed lattice wick coupon screening, comparison against traditional sintered-wick performance trends, permeability and effective-pore-radius characterization, manufacturability observations for multiple lattice topologies, pressure qualification of a printed titanium-water vapor chamber, thermal testing, orientation-dependent surface-temperature uniformity data, and repeated thermal cycling of the demonstrated hardware. Additional planned work will expand characterization of capillary pressure, phase-change surface area, mechanical strength, and device-level performance implications.
Status of Work
Development and testing. Coupon-level characterization and first hardware demonstration have been completed. Additional wick characterization and expanded application analysis are ongoing.

Thermal Control and Protection-14
Passive Thermal Switching Using Freeze–Thaw Behavior in Heat Pipes for Lunar Night Survival (Lunar Outpost)
Joe Durante

Thermal Control and Protection-15
Topology-Driven Thermofluidic Modulation of Pulsating Heat Pipes
Dr. Md Zishan Akhter and Philip Richard Hart

Pulsating heat pipes are compact, wickless, passive two-phase devices with strong potential for high-heat-flux thermal management in aerospace, power electronics, battery, and renewable-energy systems. Their performance is governed by coupled thermo-hydrodynamic processes, including evaporation-driven vapor expansion, condensation-induced pressure variation, capillary confinement, liquid-slug inertia, vapor-plug coalescence, and intermittent flow reversal. Despite high effective thermal conductance and structural simplicity, PHPs remain constrained by delayed start-up, unstable oscillations, evaporator dry-out, vapor stagnation, and elevated thermal resistance. These limitations are strongly influenced by internal channel topology, which regulates hydraulic impedance, capillary-pressure imbalance, interfacial curvature, phase clustering, and vapor–liquid residence time.
This work proposes a topology-modulated single-loop pulsating heat pipe in which controlled local and global diameter variations passively regulate oscillatory two-phase transport. A uniform 3 mm internal-diameter configuration is used as the reference geometry. The device consists of a copper wall structure with 1 mm wall thickness, a 60 mm evaporator section, a 60 mm adiabatic section, and a 30 mm condenser section. Water is used as the working fluid at a 50% filling ratio. The heat pipe is operated vertically at 90°. A heat input of 90 W is applied to the evaporator, the condenser wall is maintained at 20 °C, and the adiabatic section is thermally insulated.
Five topology-modified configurations are investigated under identical operating conditions. The first introduces localized periodic contraction on one side of the loop, reducing the local hydraulic diameter from 3 mm to 2 mm. The second introduces localized periodic expansion on one side, increasing the local hydraulic diameter from 3 mm to 4 mm. The third combines contraction and expansion on opposite sides of the loop, producing a mixed-leg topology with asymmetric capillary and inertial response. The fourth employs alternating expansion and contraction regions, imposing periodic acceleration, relaxation, and reorganization of the oscillating slug–plug train. The fifth adopts a straight dual-diameter architecture, with one leg maintained at 4 mm and the opposite leg maintained at 2 mm. In all modulated geometries, the minimum and maximum internal diameters are fixed at 2 mm and 4 mm, while the global evaporator, adiabatic, and condenser dimensions are preserved.
High-fidelity transient Volume-of-Fluid simulations are performed to resolve coupled vapor–liquid interface dynamics, phase-fraction evolution, pressure–velocity coupling, and phase-change heat transfer throughout the closed loop. The numerical framework incorporates evaporation and condensation source terms, surface-tension effects, capillary forces, no-slip wall conditions, conjugate thermal transport through the copper wall, constant evaporator heat input, isothermal condenser cooling, and adiabatic thermal isolation. Identical working fluid, filling ratio, orientation, heat input, and thermal boundary conditions are maintained across all cases. This isolates the influence of geometric topology on the underlying thermo-hydrodynamic response.
The novelty of this work is the systematic decomposition of diameter-modulated PHP behavior into individual topological mechanisms. Channel diameter is not treated as a single uniform design variable. Instead, contraction-dominated, expansion-dominated, mixed-leg, alternating, and straight dual-diameter architectures are evaluated as distinct mechanisms governing capillary-pressure redistribution, vapor generation, liquid-slug inertia, and condenser-side heat rejection. Internal geometry is therefore interpreted as a passive flow-control mechanism capable of biasing circulation, modifying oscillation persistence, redistributing vapor–liquid phases, and suppressing adverse dry-out behavior without external actuation.
The analysis extends beyond conventional thermal performance indicators through a dynamical-systems-informed assessment of PHP operation. Start-up behavior is classified according to bubble nucleation, growth, coalescence, and preferential propagation pathways. This enables topology-induced shifts in initiation mechanism and circulation bias to be identified. Oscillatory transport is quantified using kinematic and hydrodynamic metrics, including circulation velocity, reversal frequency, stroke amplitude, pressure-fluctuation intensity, and vapor–liquid phase-distribution asymmetry. High-order temporal analyses are further used to characterize transient oscillatory states, dominant frequency content, multi-scale intermittency, flow-regime transitions, and the distinction between stochastic, quasi-periodic, and circulation-dominated behavior.
Preliminary results indicate that topology modulation restructures the oscillatory response of the pulsating heat pipe by redistributing capillary-pressure imbalance and phase clustering across the loop. The modified-topology PHPs exhibit increased condenser-wall heat flux, (q”), and a corresponding reduction in equivalent thermal resistance, (Rth), of up to 66% relative to the uniform reference configuration. Constriction-dominated configurations intensify pressure-driven oscillations and promote liquid-slug acceleration through enhanced capillary confinement. Expansion-dominated geometries increase vapor-plug deformation, interfacial redistribution, and phase mixing, improving thermal communication with the condenser. The mixed-leg expansion–contraction configuration provides a favorable balance between pressure-driven circulation and phase redistribution, with improved circulation persistence and reduced dry-out susceptibility. Alternating expansion–contraction topology introduces repeated compression and relaxation of the two-phase flow, demonstrating direct coupling between local geometric forcing and global oscillation dynamics.
The research outcome will quantify topology-induced changes in start-up response, equivalent thermal resistance, condenser heat flux, oscillation persistence, phase-distribution asymmetry, and dry-out tendency across all five configurations. The integration of channel-topology engineering into PHP design establishes a scalable, passive enhancement strategy for next-generation thermal devices. The numerical insights obtained from this work will support subsequent experimental validation and future optimization of PHP geometries for high-efficiency thermal regulation. The research is currently ongoing and will be concluded as a full research article and presentation for the TFAWS conference.

Thermal Control and Protection-16
Thermal Design and Analysis of Green Propulsion Dual Mode CubeSat
Savanna Lyles and Clark Teems

Green Propulsion Dual Mode (GPDM) is a collaboration between NASA Marshall Space Flight Center (MSFC) and Georgia Tech to create a 6U CubeSat demonstrating the use of the green propellant Advanced Spacecraft Energetic Non-Toxic (ASCENT) in two propulsion systems: using a traditional chemical thruster and an electrospray thruster system. The driving thermal considerations include a planned orbit always in Sun view without any time in Earth’s eclipse, and a chemical thruster reaching 1300°C during firing. Thermal design and analysis on CubeSats are not always emphasized, particularly in experimental or academic environments. This paper will describe the thermal design and analysis performed on GPDM and the mitigations implemented for this mission, which will highlight some of the thermal challenges unique to experimental small satellites.

Thermal Control and Protection-17
Variable Emittance Coating Systems Thermal Modeling for a 3U CubeSat
Sarah G Stewart, Sam Keller, Sydney Taylor, and Ognjen Ilic

Thermochromic variable emittance coatings (VECs) can passively change their heat rejection as a function of temperature. This makes them an attractive option for CubeSat and small sat thermal control, where mass and power are at a premium. Recent advances in the state of the art by VEC developers have made VECs promising for robotic spacecraft, notably low solar absorptance and large-scale manufacturing capability. This study is focused on demonstrating the benefit of VECs for a small satellite in Low Earth Orbit (LEO). The survival heater power is compared between a static emissivity coating and VECs for nadir and sun facing attitude configurations. The effect of VEC transition temperature, transition width, and hysteresis is also examined to determine the optimal properties of VECs for small CubeSat LEO applications. These trade studies are undertaken in both Thermal Desktop and COMSOL, and the modeling methods for VECs in both analysis tools are outlined.

Thermal Control and Protection-18
SRM Testbed for Thermal Testing of Internal Components
Dr. Robert Frederick

New internal insulator, nozzle, and seal materials must be screened to determine their response to the hostile internal environments of a solid rocket motor. This project details a solid rocket motor that provides 15 to 25 seconds of test time at pressures from 1,000 to 3,000 psi. The scope includes interior ballistic design, design and demonstration of a water-cooled zero-erosion nozzle, and results from initial demonstration tests. The motor burns a conventional AP/HTPB propellant, is ignited with a solid squib, and has pressure and thrust instrumentation to monitor the performance. Calculations and data are shown for the next phase of the project, in which the pressure will be modulated from 1,000 to 3,000 psi over two cycles during a single test. The motor demonstrated successful operation, the successful performance of the zero-erosion, water-cooled nozzle, and steady operating pressure at 1,000 psi for 30 seconds. Future work will include using the test section to evaluate SRM internal materials under controlled conditions.

Thermal Control and Protection-20
Thermal Design and Thermal Vacuum Testing of the StarBurst Instrument
Stephanie Mauro

The StarBurst Multimessenger Pioneer is a small satellite mission serving as a wide-field gamma-ray observatory designed to capture the initial emissions of short gamma-ray bursts, electromagnetic signatures of neutron star mergers. This paper presents the final thermal design and analysis of the StarBurst Instrument, comprising the bus-to-instrument interface plate, control electronics, and twelve crystal detector units, which form the core of the mission’s science capability. The passive thermal control system design requires consideration of restrictive keep-out zones, unknown orbital parameters, and narrow temperature limits of the detectors. Also summarized is the instrument level thermal vacuum cycle test, correlated model refinements, and updated model results. Following successful completion of the instrument test campaign, the hardware was integrated with the spacecraft bus for spacecraft level testing, including additional thermal vacuum testing. The results from the spacecraft level thermal vacuum test will further inform the instrument thermal model, ensuring accurate flight temperature predictions. StarBurst launches as a secondary payload in 2027 and has a mission duration of at least one year.

Thermal Control and Protection-21
Developmental and Cryogenic Thermal Vacuum Testing Lessons Learned
Mackenzie Byrnes and Heather Grimes

Developmental thermal vacuum (TVAC) testing is a critical step in maturing hardware designs and validating performance prior to flight qualification. Unlike qualification or acceptance testing, developmental testing provides flexibility to explore design margins, uncover integration challenges, and refine test approaches before formal verification activities begin. This presentation highlights the value of developmental testing while sharing common pitfalls encountered during developmental TVAC campaigns. Lessons learned from hands-on testing experience, including extensive developmental testing at cryogenic temperatures, will be shared in this presentation. Topics include test planning and preparation, instrumentation strategies, contamination control considerations, troubleshooting unexpected anomalies, and approaches for staying on schedule while meeting test objectives. The audience will gain practical insights and best practices that can improve test efficiency, reduce risk, and enhance the overall success of future developmental TVAC efforts.

Thermal Control and Protection-22
Thermal Performance of Flight Imagery Launch Monitoring Real-time System (FILMRS) and Enhance FILMRS (EFILMRS) camera systems for Artemis
Deborah Hernandez

NASA’s Flight Imagery Launch Monitoring Real-time System (FILMRS) and Enhanced FILMRS (EFILMRS) camera systems were developed for Artemis Core Stage. FILMRS has provided valuable performance data during Artemis I and II flights. These camera systems are based on Commercial Off the Shelf (COTS) cameras and image processing hardware, thermally controlled through application of various passive thermal management techniques. The EFILMRS system consists of a controller unit and multiple cameras employing three operational modes for use on future Artemis flights. The redesigned EFILMRS upgrades thermal performance and reduces mass, volume, and power. Though not mission or safety critical extensive thermal qualification and verification testing were performed, testing COTS components to their capability limits as opposed to expected environments. EFILMRS successfully completed qualification thermal testing and progressed to flight build. Flight units are currently in the building and acceptance test cycle. This paper summarizes FILMRS flight performance, and describes the associated design and verification test efforts.

Thermal Control and Protection-23
SR-1 Freedom Thermal Architecture Challenges
Scott Thomas

The SR-1 Freedom mission aims to demonstrate nuclear electric propulsion (NEP) and deliver the Skyfall helicopter payload to Mars. Repurposing the existing Gateway Power and Propulsion Element (PPE) spacecraft and combining it with a 20 kWe-class nuclear power module (NPM) presents a variety of thermal architecture challenges. These include unprecedented waste heat rejection requirements, integration with a spacecraft bus originally designed for a different mission profile, and novel packaging constraints. This presentation will describe the simplified concept of operations as it relates to on-orbit thermal environments, early-phase architecture trades and supporting analyses, and planned forward work. We will also discuss our strategy for integrated thermal modeling of the complete spacecraft and examples of interface challenges necessitated by this ambitious effort.

Thermal Control and Protection-24
Power and Propulsion Element Steerable High Gain Antenna Lunar Transit Thermal Analysis Tracking Methodology
Matthew Faykus

The Power and Propulsion Element (PPE) is an ion thruster propulsion spacecraft developed as an element of Space Reactor (SR-1) Freedom to provide propulsion, communications and power for the spacecraft. PPE was originally being developed for the use with the lunar orbiting space station Gateway as one of the first two planned elements. PPE was to be launched with the Habitation and Logistics Outpost (HALO) element in a configuration called the Co-Manifested Vehicle (CMV) that would arrive at a Near-Rectilinear Halo Orbit (NRHO) around the Moon via a lunar transit spiral trajectory phase [1]. The PPE communication system is equipped with two Steerable High Gain Antennas (SHGA) each steered by a two-axis gimbal (TAG) mechanism [2]. A thermal analysis was conducted during the near-Earth spiral phase of the mission using Thermal Desktop (TD). This analysis utilized multiple axis Earth tracking articulators in combination with SINDA system internal environmental heating symbols to produce accurate Earth ground station tracking communication system temperatures. The presentation submitted for TFAWS 2026 will provide an overview of the communication system thermal model and the analysis methodology. The work completed in combination with the published presentation will provide substantive information to the thermal community of the modeling capabilities of TD for future spacecraft missions. This work has been concluded as of April 2026.
References
[1] McGuire, M. L., McCarty, S. L., Hack, K., and Karn, S. N., “Application of Solar Electric Propulsion to the Low Thrust Lunar Transit of the Gateway Power and Propulsion Element”, 38th International Electric Propulsion Conference, IEPC, 2024.
[2] Dendy, R., Zeleznikar, D. J., and Zemba, M. J., “NASA LUNAR EXPLORATION – GATEWAY’S POWER AND PROPULSION ELEMENT COMMUNICATIONS LINKS”, 38th International Communications Satellite Systems Conference (ICSSC), AIAA, 2021.

Thermal Control and Protection-25
Thermal Design Considerations for Venus Orbiters: Lessons Learned from VenSAR
Tyler M Schmidt

The Venus Synthetic Aperture Radar (VenSAR) is a planned JPL contribution to the European Space Agency’s planned EnVision mission. The planned VenSAR instrument would be an S-band radar with RF electronics and an antenna consisting of a feed and a reflectarray. The planned mission would include multiple instruments to provide a comprehensive survey of the Venus atmosphere, surface, and subsurface. The primary objective is to better understand why Venus and Earth have evolved to be different.
After cruising from Earth to Venus, the spacecraft would perform aerobraking to gradually reduce the apoapsis altitude to a low eccentricity polar orbit for the science mission. The VenSAR reflectarray would nominally deploy several weeks after launch and thus participate in the aerobraking drag passes. Therefore, the reflectarray must survive the aerothermal flux in addition to the solar and planetary thermal loads. Once the science phase of the mission begins, the spacecraft attitude would need to be carefully managed to keep the reflectarray and feed within allowable flight temperatures. The antenna system would also need to stay within its pointing error budget to meet science requirements.
Recent work to advance the reflectarray to NASA technology readiness level six has focused on preparation for thermal and structural testing while finding solutions for competing requirements among the materials and processes, mechanical, RF, and thermal disciplines. The thermal environments pose significant thermal challenges regarding survivability of the composite panels over a single orbit and of fatigue-induced effects over thousands of orbits. The thermal team has considered a large scope of possible analysis cases to characterize the reflectarray temperatures and transient response. Since the reflectarray has a low thermal mass and is necessarily exposed to space for RF transmission, it is sensitive to the magnitude and direction of incident solar and aerothermal loads. Thermal model predictions for temperature, gradient, and ramp rate have been used for thermal cycling fatigue testing, structural analysis, and structural-thermal-electrical-performance analysis. Furthermore, there has been a concerted effort to resolve predicted temperature violations from the preliminary design review and identify parallel mitigation paths for a detailed design phase. This presentation will summarize the VenSAR thermal design and analysis effort since the preliminary design review. It will also share lessons learned about designing for a dynamic Venus orbital thermal environment in the context of a radar instrument.

Thermal Control and Protection-26
Thermal Design of a FALCON Cloud Radar Instrument Concept
Tyler M Schmidt, Katarina Aguayo, and Rogelio Rosas

JPL is developing a W-band cloud radar instrument concept that would be part of the Fleet for the Atmosphere Linking Commercial Observations with NASA (FALCON). The mission would launch no later than early 2029 and support cloud science objectives identified in the Earth science decadal survey. The instrument could be potentially repurposed for a future mission to support surface mapping of Earth’s moon or Mars. The instrument concept would consist of two Cassegrain antennas, three pallets of electronics that perform radio frequency functions, and an instrument deck that mechanically accommodates everything. The instrument concept would fly in a high inclination sun synchronous orbit. The instrument concept would be the lone payload on a spacecraft and would operate almost continuously on orbit for a nominal lifetime of two years.
There are two primary thermal challenges on FALCON-Radar. The first is that the large 2.1-meter antennas must remain within their allowable flight temperatures while maintaining sufficient co-boresight pointing alignment. The difference between day and night thermal environments introduces dynamic thermal elastic distortion that alters pointing. On orbit antenna pointing is predicted by an integrated structural-thermal-electrical-performance analysis, and the predictions are used to iterate on the design. The second challenge is that the high, steady state heat loads generated by the RF electronics must be managed to keep electronics from overheating. To accommodate this, heat pipes would transfer heat away from the electronics pallets to radiators where heat is rejected to space.
FALCON-Radar is a cost-capped NASA Class D mission, which grants a higher risk tolerance compared to classes A, B, or C. The planned mission inception to launch timeline is also under 36 months. Therefore, the design requires thoughtful balance of robustness and resources. Although more capable solutions may exist, the design must usually be tailored to meet (and not exceed) requirements. This presentation will provide an overview of the preliminary thermal design while including considerations for a fast-paced class D mission environment.

Thermal Control and Protection-27
Evaluation of AMSOIL®-ANT PGW Coolant Formula Change and Super Space AMSOIL Development
Scott Hansen, Woody Beringer, Toni Griego, Lauren Foley, David Brockett, and Garret Gibeau

AMSOIL® Propylene Glycol Water (PGW), is the baselined internal thermal control fluid for multiple manned spacecraft including Gateway (HALO, IHAB, and ALM modules) and Orion. The use of AMSOIL-ANT PGW was originally validated through extensive NASA/Collins materials-compatibility testing (circa 2008–2015). However, in late 2019, AMSOIL implemented an unannounced change in its PG inhibitor package. Because of this reformulation, concerns around materials compatibility arose. Therefore, NASA initiated corrosion rate and coolant stability testing in 2024. Results showed that the reformulated AMSOIL PG produced substantially elevated corrosion rates in aluminum alloys (about two times when compared to the original AMSOIL PG formula) accompanied by visible surface degradation. Coolant stability testing indicated accelerated propylene-glycol breakdown as well, with glycolate levels about nine times higher than the original formulation. Based on these findings, NASA Materials & Processes and Thermal Control Systems groups have baselined the use of the original PGW chemistry for all spacecraft use. In response, AMSOIL produced a reformulated variant, “Super Space AMSOIL” (ANT-SSA), compliant with MPCV 70156.This paper documents the reformulation testing completed, details on the formulation change, and the development of Super Space AMSOIL.

Thermal Control and Protection-28
Orbit-Informed Thermal Modeling Workflow for Thermo-Mechanical Solder Fatigue Prediction of Electronics
Ian Pond, Josh Akman, and Mina Karimaghaei

Thermomechanical fatigue life of solder joints is often a concern for electronics in space applications. However, because physical testing is constrained by cost, schedule, and test coverage, assessing fatigue life can be challenging. This presentation introduces a workflow for evaluating solder fatigue in electronics using temperature data generated from simulated orbital missions or, where available, on-orbit sensor data. In this workflow, Ansys Thermal Desktop is used to generate transient thermal time histories for a satellite based on planned mission profiles. Transient PCB time-temperature histories are then extracted and provided to Ansys Sherlock’s modified rainflow cycle counting algorithm to generate thermal cycle profiles. Ansys Sherlock is subsequently used to assess solder fatigue for components on the PCB. The presentation also reviews script-based automation of the workflow and sensitivity studies on the rainflow extraction parameters. Although the primary focus is a proactive design-for-reliability approach based on predictive thermal modeling, extensions to operational monitoring and remaining useful life assessments using real-time temperature sensor data are also discussed.

Thermal Control and Protection-29
ONYX-Thermal – A New Concept for General Purpose Heat Transfer Software
Dr. Dean S Schrage

The evolution of the modern computer aided engineering (CAE) thermal analysis tools has produced software that is capable of highly complex multidisciplinary simulations, featuring integrated CAD geometry, automated meshing, parametric modeling, optimization and topology evolution, etc. In our first paper on ONYX-Thermal in 2024, it was argued that this modern CAE toolset is overkill and that the majority of users can accomplish the bulk of their analyses with a focused and simplified toolset. Since that time, Mainstream has launched the development of ONYX with the primary goals of creating a general utility tool that is designed to prioritize 1) fast model construction, 2) fast solver speed, 3) insensitivity to user input error, and 4) a streamlined user interface (UI) to enhance overall user experience (UX). The present paper will detail the key features of our software development as they currently exist, highlighting enabling features that will bring about these software goals. The most important of these is the ability to import CAD geometry and convert to a control volume conductor network that can support an immediate heat transfer analysis, thus skipping the intrinsic meshing step in conventional CAE tooling. This CAD-to-simulation step is made possible by automatically converting shell geometry into thermal solids. Second, we demonstrate an analytical treatment to formulate an accurate conductance network on highly skewed triangle and quadrilateral elements, promoting fast analysis on very coarse meshes. We will demonstrate the speed at which model geometry can be modified, delivering a drag and drop feel with near instantaneous updates to element equivalencing and conductor calculation after each edit. The ONYX solver has been revamped from our earlier developmental code and now achieves a per node per iteration CPU grind of 1E-7 s. This fast solver speed facilitates a new feature called Thermal-Intern, an agent that allows the model to solve and display results continuously while the user is actively building and modifying a model. The paper will close with a discussion of the development roadmap showing how the current framework will accommodate advanced features such as model motion and phase change heat transfer.

Thermal Control and Protection-30
Intermediate temperature oscillating heat pipe radiators for space nuclear power
Alex Miller and Scott Hayden

Oscillating heat pipes (OHPs) are innovative heat transfer devices that have gained attention for their potential in intermediate temperature applications, particularly in support of lunar fission surface power (FSP) and nuclear electric propulsion (NEP) missions. OHPs function through the oscillating or pulsating motion of a saturated two-phase working fluid within a series of interconnected channels, enabling efficient passive heat transfer.
A recent NASA Small Business Innovation Research (SBIR) phase III project focused on advancing OHP technology for NEP heat rejection subsystem (HRS) applications. Lightweight, meter-scale, titanium-water oscillating heat pipe radiators were developed, exhibiting high performance: areal density < 3kg/m2, and specific power >1kW/kg. These advancements underscore the significant progress being made toward utilizing OHP technology in next-generation space missions.

Thermal Control and Protection-31
An Overview of the PULSE Space Thermal Control Pump Development at the Jet Propulsion Laboratory
A J Mastropietro, Michael R. Johnson, Warrick S. Leigh, Duval A. Johnson, Samuel G. Dupas, Talia R. Spitz, Dalia Raafat, Alejandro Lopez Ortega, Allison Ayad, and Patrick Phelps

The Jet Propulsion Laboratory (JPL) has a long history of flying successful missions that rely upon Mechanical Pumped Fluid Loops (MPFLs) as an integral part of their thermal control. MPFL thermal architectures enable many diverse space mission configurations that need to operate in a narrow temperature range despite widely varying external space or planetary environments. By utilizing either single-phase or two-phase working fluids, they are readily scalable from very low-power systems that need to efficiently utilize waste heat up to very high-power output systems, such as fission reactors that need to dissipate kilowatts of power. The operational reliability of these loops is predicated on the integration of robust long-life pumps.
Given the recent NASA and industry announcements regarding plans for fission reactor powered spacecraft, permanent lunar outposts, and orbital data centers, the forecasted need for long-life (i.e., greater than 10 years) maintenance-free space-rated thermal control pumps is anticipated to only increase. A more rapid cadence of pump hardware deliveries for various space missions is also expected.
Meeting this rising demand calls for a modern space pump design that can be developed at significantly lower cost and on shorter schedules. For this reason, JPL is investing in the development of PULSE (Pump Unit with Long-life for Space Environments), a scalable, high-reliability thermal control pump design. PULSE builds on decades of JPL’s MPFL experience and incorporates new technologies to guarantee long-term performance.
At the core of PULSE is a fully hydrodynamic fluid-bearing system that ensures near zero-wear operation post-assembly. To mature the design, JPL is pursuing a multi-pronged development approach that includes expanding in-house hydrodynamic bearing analysis capabilities, carrying out extensive bearing-material compatibility and wear-testing campaigns, and developing a dedicated small scale journal-bearing test stand to validate analytical models. This presentation will summarize the JPL PULSE development work that is being conducted in partnership with industry and NASA experts.

Thermal Control and Protection-32
Artemis II Mission Performance of the Orion Active Thermal Control System
Ellie Thurston, Michael Johnston, Thomas Chen, Eddie Uribe, Dongeun Lee, John Whitmore, Johnathan Hernandez, Maddie Haas, Matthew Gietzel, and Ian Anchondo

The Orion spacecraft has recently completed Artemis II, the first crewed mission around the moon in 54 years. This test flight demonstrated vehicle capability as the foundation for all future missions under the Artemis program. As a major system of Orion, the Active Thermal Control System (ATCS) successfully managed crew and vehicle heat loads to provide adequate cooling and thermal comfort for the duration of the Artemis II mission. Thermal control on the Orion vehicle is managed through two redundant cooling loops in the Crew Module (CM) that absorb heat loads from the cabin and Air Revitalization System (ARS) heat exchangers, the Liquid Cooling Garment (LCG) heat exchanger, and the avionics cold plates, and then transfer thoseheat loads to the Service Module (SM) radiator loops and/or the Phase Change Material (PCM) and ammonia boilers when required through thermal configuration or mission phase. During Artemis II, Orion did not necessitate any supplemental cooling from ammonia boilers until service module separation, allowing the maximum availability of ammonia cooling prior to entry. This presentation will provide an overview of the Orion ATCS major subsystems, highlight the performance during key mission phases, and comment on the applicability of the observed thermal response to future Artemis missions.

Thermal Control and Protection-33
BubbleTrack: A Computer Vision-Based Framework for Bubble Detection and Tracking
Dr. Sara Youssoufi

Direct Numerical Simulations (DNS) of multiple bubbles in nucleate pool boiling have been extensively utilized to investigate the complex phenomena governing phase-change heat transfer. Such studies are particularly relevant to advanced thermal management systems for satellites and spacecraft, where efficient heat dissipation and cooling mechanisms are essential for reliable operation. The simulations considered in this work are performed using an in-house DNS solver that resolves the incompressible Navier–Stokes equations coupled with mass, momentum, and energy conservation equations. The liquid–vapor interface is captured using a level-set method, enabling accurate representation of bubble nucleation, growth, deformation, coalescence, and departure.
Monitoring bubble dynamics is essential for understanding vortex structures, bubble growth, departure, deformation, and their impact on heat transfer. However, extracting bubble statistics from large DNS datasets can be computationally demanding and time-intensive. To address this challenge, this work presents BubbleTrack, a computer-vision-inspired framework for automated bubble detection and tracking. The developed algorithm identifies individual bubbles and extracts key properties, including centroid coordinates, equivalent bubble diameter, area, volume, and probability density functions.
The performance of BubbleTrack is validated through comparison with a VTK-based post-processing methodology implemented within the ParaView framework. Results show excellent agreement between the two approaches across all monitored bubble characteristics, demonstrating the accuracy and reliability of the proposed computer-vision algorithm. The development, implementation, and validation of BubbleTrack have been completed, and the framework is now available for large-scale analysis of DNS boiling simulations. The developed framework provides an efficient and automated tool for generating high-fidelity bubble statistics and supports advanced boiling research and thermal management applications.

Thermal Control and Protection-34
HEAT TRANSFER AND AERODYNAMIC LOSSES CAUSED BY ADDITIVELY MANUFACTURED GRX-810 TURBINE BLADES ENHANCED WITH MICRO-MACHINING AND CHEMICAL POLISHING
Phil Ligrani, Hallie Collopy, Mason Hancock, Jason Sheth, and Paul Gradl

With rising turbine inlet temperatures and higher pressure ratios, the performance demands on rocket engine turbine blades have become increasingly more demanding. Additive manufacturing (AM) offers a significant advantage for such environments by enabling the production of complex geometries that are difficult or impossible to achieve using conventional methods. Additive manufacturing allows the utilization of advanced high-temperature alloys like GRX-810, which offers greater strength and durability than currently used alloys. GRX-810 is a NiCoCr-based alloy that leverages the AM process to achieve enhanced properties. During laser powder bed fusion (L-PBF), yttrium oxide nanoparticles, pre-applied to the feedstock powder, are uniformly dispersed throughout the material. This microstructural refinement results in a substantial increase in ultimate tensile strength at high temperatures, making GRX-810 components well suited for the extreme thermal conditions present in rocket engine turbines. Its name derives from the Glenn Research Center Extreme Temperature designation for service above 810 °C.
However, an inherent byproduct of AM processes is a relatively rough surface texture, which varies with the post-processing method employed. In the present investigation, several surface enhancement conditions are evaluated, including chemical polishing (CP), and micro machining processing (MMP). Results obtained with these arrangements are compared to data obtained with an as built (AB) surface with no post processing. Within the investigation, the blade with the AB surface is denoted AM0, the blade with the CP surface is denoted AM2, and the blade with the MMP surface is denoted AM3. Within the present study, the different surface arrangements are evaluated for their effects on surface roughness, as well as their effects on turbine blade aerodynamic losses and tip surface heat transfer coefficient distributions.
To obtain the necessary aerodynamic and thermal data, experiments are conducted in a transonic linear cascade located within a transonic/supersonic wind tunnel. The cascade comprises five turbine blades, one of which is made of GRX-810 material, which is additively manufactured and centrally instrumented for aerodynamic loss and surface heat transfer measurements. The configuration includes four flow passages, with cascade inlet static and stagnation pressures, as well as recovery temperature, measured upstream of the blade row using a static pressure port, a Kiel probe, and two calibrated thermocouples, respectively. Surface static pressures around the central instrumented blade are obtained from 22 static pressure taps located circumferentially and axially at 50% span. Downstream of the blade row, at 0.25 Cx from the cascade exit plane, a one-dimensional traverse equipped with a total-static pressure probe and a calibrated thermocouple sweeps across the central blade’s wake to characterize aerodynamic loss parameters.
Recalling that the AM0 blade is associated with an AB surface, the AM2 blade is associated with a CP surface, and the AM3 blade is associated with an MMP surface, results show that the wake of the AM0 blade exhibits reduced stagnation pressure losses despite its higher mean surface roughness. AM2 produces the greatest aerodynamic losses on the pressure side, while AM3 shows the largest losses on the suction side. AM3 trends closer to AM0 on the pressure side but aligns with AM2 on the suction side. Overall, both AM2 and AM3 generate higher stagnation pressure losses than AM0, largely due to differences in surface roughness scale. Larger roughness features enhance near-wall mixing and increase boundary-layer energy, which delays separation and reduces wake intensity. These findings are consistent with normalized kinetic energy and Mach number distributions, while total pressure loss coefficients, ω pressure loss coefficients, ELC energy loss coefficients, and entropy changes exhibit inverted profiles of wake total pressure.
Spatially resolved surface temperature measurements along the blade tips are obtained using an infrared camera mounted atop the cascade casing. Optical access to the tip region is provided through a zinc selenide window integrated into the casing wall. Local surface heat transfer is evaluated using the impulse response method, in which time varying surface heat flux is derived from the corresponding time varying surface temperature at each camera pixel location. This approach yields spatially resolved distributions of both adiabatic wall temperature and surface heat transfer coefficients. Local Mach numbers within the tip gap passage are also determined from these spatially resolved adiabatic wall temperature distributions.
Dimensional heat transfer coefficients along the blade tips are generally higher for blades with rougher surface textures. AM0 is observed to have the largest line averaged heat transfer coefficients. The elevated heat transfer observed for AM0 and AM2 is attributed to their greater surface roughness, which enhances turbulent transport and mixing within the tip gap region. In contrast, AM3 displays the highest Mach numbers due to its smoother surface finish. The reduced roughness leads to lower surface pressures, allowing the flow to accelerate more readily and producing higher local velocities and Mach numbers along the blade tip passage. AM0 and AM2 exhibit very similar distributions and magnitudes of heat transfer coefficients, whereas AM3 shows a comparable qualitative distribution but with overall lower magnitudes. With all three blades (AM0, AM2, AM3), with all three surface processing approaches (AB, CP, MMP), the largest differences between suction and pressure sides occur near the trailing edge, reflecting the combined effects of flow acceleration, boundary-layer development, and local blade geometry. Additionally, lower Mach number regions appear near the leading edge and again toward the trailing edge, while higher Mach numbers are observed in regions of significant flow turning.

Thermal Control and Protection-35
Spacecraft Radiator Protection from Ionizing Radiation, Dust, and Excessive Heat Loss
Dr.Vijay V Devarakonda, Michael D Hogue, and Darnell Cowan

Under a Phase II SBIR project funded by NASA Johnson Space Center (Contract No. 80NSSC25C0088), Analytical Scientific Products LLC (ASP) has been developing an actively controlled louver to protect spacecraft radiators from degradation due to exposure to various types of environmental conditions. Of particular interest are ionizing radiation during spacecraft transit through the Van Allen belts, dust during spacecraft landing and surface operations on the moon, and excessive heat loss during the long lunar night especially near the poles where the local ambient temperatures can drop below -200°C. Exposure to ionizing radiation and dust can degrade the optical properties of the radiator coating that in turn reduces its ability to reject excess heat from the spacecraft into the environment. Exposure to the extremely low temperature conditions during the lunar night near the poles can freeze the radiator fluids that can compromise the integrity of the radiator.
Passive louvers constructed using thick aluminum vanes are used currently to protect spacecraft radiators from some of the above effects, but they have several drawbacks: (i) their high aerial density makes it impractical to scale them to protect the much larger spacecraft radiators needed for future manned missions to the moon and beyond, (ii) the bimetallic actuators used to open and close the vanes in passive louvers rely on external temperature alone and so the louver cannot offer protection from dust and ionizing radiation when the ambient temperatures are high, and (iii) the bimetallic actuators need time scales of the order of hours to open and close.
We have addressed the above limitations of passive louvers through the development of a low aerial density and rapidly actuating actively controlled louver. It is constructed from a custom alloy that offers much higher levels of protection against the ionizing radiation, dust and excessive heat loss at a fraction of weight compared to passive louvers while providing opening and closing time scales of the order of a second. Our modular design makes it easy to scale the system up or down depending on the application. We have recently constructed a 31 inch × 31 inch module of this louver and tested its functionality and effectiveness. These tests have shown that the louver can be opened and closed in less than 2 s. It reduces the transmission of ionizing radiation by 78% to 100% (depending on the radiation source), dust transmission by more than 93%, and heat loss by more than 97%. We are currently preparing to evaluate this device at the Johnson Space Center’s cryogenic vacuum chamber where it can be subjected to simulated lunar surface conditions.
This paper provides a detailed discussion of the test designs as well as the data from tests that demonstrate the effectiveness in reducing the transmission of ionizing radiation, dust, and heat under laboratory conditions.

Thermal Control and Protection-37
The Multi-Purpose Habitation Module (MPH): an overview of thermal modelling and lunar surface effects
Lorenzo Strappato, Salvatore Lauretta, Davide Perrone, Simone Illiano, Roberto Bertacin, and Marilena Amoroso

Lunar surface assets deployed on the lunar surface shall withstand extreme and highly variable environmental conditions compared to orbital ones. When such assets are both mobile and crewed, the reliability requirements and hence the design complexity are further amplified: external boundary conditions change continuously, uncertainties rise, and operational limits become narrower, even when astronauts are not present in the cabin.
In such context, the Multi-Purpose Habitation module (MPH) developed by Thales Alenia Space Italia S.p.A. under a contract signed with the Italian Space Agency (ASI) and in collaboration with the National Aeronautics and Space Administration (NASA) in the frame of the Artemis Accords, is designed to address these crucial challenges being the precursor shelter for the first human surface missions on the Moon.
In particular, the design boundaries imposed by several subsystems (like mobility) combined with the life support requirements, during crewed phases, impose very stringent constraints on the module thermal design and greatly increase the complexity of system-level thermal analyses.
This paper presents the latest progress in the system-level thermal modelling for MPH during the ongoing preliminary design phase.
Particular emphasis is placed on the adopted techniques for the simulation of the lunar surface environment and its interaction with the module, explaining current assumptions and constraints, and highlighting preliminary outcomes useful for future design refinements and to strengthen the consolidation of some enabling technologies under development in the frame of the ASI contract.
At this early design stage, a simplified “flat” model of the lunar surface has been used, to characterize primary thermal sensitivities, the main technical issues of such environment; a detailed 3D terrain model is planned to be implemented in the following phase of the project.
Ultimately, the objective of this paper is not only to assess and compare modelling methodologies, but to define an incremental approach to reduce uncertainties in the frame of the system-level thermal analyses and to support the design choices requested by such a pioneering and challenging project.

Thermal Control and Protection-38
Development and Initial Validation of a High-Vacuum Thermal Conductivity Testbed for Aerospace Thermal Interface Materials
Amy Chang and Deborah Hernandez

Thermal Interface Materials (TIMs) are critical components in spacecraft thermal management systems, where thermal performance is strongly influenced by vacuum conditions, interface contact resistance, and layered metallic joint behavior. However, manufacturer-reported thermal conductivity values are often derived under idealized conditions and may not accurately represent performance within operational aerospace applications. To address this limitation, the Testbed for Advanced Interface Materials in Vacuum (TAIMV) was developed as a modular vacuum-compatible thermal conductivity characterization platform capable of evaluating aerospace-relevant TIM configurations under both ambient and high-vacuum environments. The testbed was derived from the ASTM C1044-16 guarded hot plate methodology and incorporates interchangeable layers of stainless steel coupon geometries, independently controlled main and guard heaters, embedded resistance temperature detectors (RTDs), thermocouples, multi-layer insulation (MLI), and a temperature-controlled cold plate to characterize through-thickness thermal gradients across layered interfaces. In the current configuration, interface compression is limited to the nominal contact pressure generated by the experimental stack assembly. Initial experimental campaigns were conducted at ambient pressure and below 1×10-5 Torr for vacuum cases using multiple interface materials including Braycote 601EF and Krytox Low Vapor Pressure (LVP) vacuum greases across a range of thermal operating conditions. In parallel, a coupled numerical Python thermal model was developed to predict temperature distribution throughout the stack while accounting for conduction, radiation, and parasitic heat transfer pathways and effective interface resistance effects. Experimental measurements and numerical predictions showed consistent thermal trends across multiple operating conditions and environmental states. Results also revealed measurable differences between ambient and vacuum thermal behavior, demonstrating the importance of interface resistance, parasitic heat transfer mechanisms, and stack geometry in determining effective thermal performance within layered thermal interfaces. The presented work establishes a foundation for future thermal model correlation efforts and expanded characterization of aerospace thermal interface materials under representative environmental conditions. Future work will focus on the integration of a load cell system to enable controlled pressure-dependent characterization of thermal interface materials under compressive loading. This capability will allow investigation of the influence of contact pressure on effective thermal conductivity, interface resistance, and thermal performance within layered aerospace thermal interfaces under representative operational conditions.

Thermal Control and Protection-39
Gravity-Independent Heat Exchangers and Oil-Free Compressors for Modular Refrigeration System
Cara Martin and Dennis Nasuta

A modular refrigeration system using a vapor compression cycle (VCC) with gravity-independent heat exchangers and oil-free compressors was designed to meet targeted space refrigeration needs. The system is capable of scaling to meet capacities ranging from 200 W to 1000 W and could deliver an efficiency equivalent to 45% of Carnot COP. Linear compressor technology was tested and confirmed to be suitable for use, providing an oil-free solution that enables gravity-independent operation. Together with an optimized evaporator and coaxial condenser, the system will be capable of providing robust operation in microgravity conditions.

Thermal Control and Protection-40
Design of a Novel Radiant Heat Pump Thermal Control System for Space Conditioning of Lunar Surface Habitats
Dr. Ardeshir Moftakhari

NASA’s Artemis campaign requires compact, lightweight, and reliable thermal control systems capable of maintaining habitable indoor conditions in lunar surface habitats while operating under extreme environmental constraints. Lunar habitats experience severe thermal challenges, including near-vacuum conditions, large day/night temperature swings from approximately –170°C to +120°C, strong solar variability, limited water resources, stringent mass and volume constraints, and performance degradation from lunar dust accumulation. Current habitat thermal-control concepts can be bulky, difficult to reconfigure, and vulnerable to long-duration environmental exposure. This work presents a novel radiant heat pump thermal control system architecture for space conditioning of lunar surface habitats.
The proposed concept uses a packaged vapor-compression heat pump coupled with deployable condenser radiators to reject habitat heat directly to the lunar environment through radiation. The architecture is designed to support indoor space conditioning, heat rejection, dehumidification, and integration with life-support functions while reducing system mass, refrigerant charge, and operational complexity. A two-phase condenser radiator concept is emphasized because phase-change heat transfer can maintain more uniform refrigerant and radiator-panel temperatures, improving radiator utilization and reducing required radiator area. Low-GWP, non-toxic refrigerants and oil-free compression are considered to support environmentally responsible, gravity-independent, and reliable long-duration operation.
The work includes conceptual system design, thermodynamic cycle analysis, working-fluid screening, deployable radiator sizing, and segment-by-segment radiation-only condenser modeling. The radiator model marches along the refrigerant flow path and updates local refrigerant state, heat transfer, tube/panel conduction, panel temperature, and radiative heat rejection under lunar-vacuum boundary conditions. This approach captures the spatial variation in quality, temperature, and heat flux that cannot be represented by a single lumped radiator model. Key design variables include condenser saturation temperature, radiator geometry, surface emissivity, panel material, tube-panel conduction resistance, inlet superheat, and lunar boundary conditions.
The unique contribution of this work is the integration of a radiant vapor-compression heat pump with a deployable radiation-only two-phase condenser radiator for lunar habitat space conditioning. Unlike conventional terrestrial HVAC systems, the proposed system must reject heat without convection and remain functional across extreme lunar environmental conditions. Preliminary modeling is being used to evaluate the feasibility of rejecting approximately 15 kW of peak habitat heat load, assess radiator size and material tradeoffs, and identify operating strategies that improve heat rejection efficiency while maintaining compactness and reliability.
The work is currently in the concept development and modeling stage, with planned validation using NASA thermal analysis tools and future scaled testing under representative thermal-vacuum conditions. The expected outcome is a physics-based design framework for evaluating radiant heat pump thermal control systems for sustained lunar habitation. The results will support future development of compact, modular, and resilient thermal-control architectures for Artemis lunar surface habitats and long-duration Moon-to-Mars missions.

Thermal Control and Protection-41
Additively Manufactured Porous Radiator For Static Thermal Management
Kristen Ess and Rydge Mulford

Effective thermal management is essential for spacecraft operating in the harsh and unpredictable space environment, where radiation is the main form of heat transfer. Radiators are commonly utilized to reject excess heat; however, many designs are limited by their fixed radiative properties and often require additional technology to meet mission requirements. Additionally, additive manufacturing processes in the space environment have been proposed to manufacture in-situ components for spacecraft. Radiator coatings, however, require tailored materials which cannot be produced in the space environment without additional equipment. The ability to tailor radiative properties of additively manufactured components would enable production of surfaces with designed emissivity and absorptivity values. This work explores three separate methods for tailoring of surface emissivity via additive manufacturing processes, including femto-second laser etching of the finished surface, production of meso-scale surface features to generate cavities of varying aspect ratios, and introduction of porous layers that penetrate several millimeters into the printed surface which generate spherical, linked cavities that likewise utilize the cavity effect. For the porous layers approach, the proposed radiator transitions from a fully dense base to an increasing porosity gradient that extends outward. Thermal radiation is transferred from the spacecraft into the radiator, where the pores act as spherical cavities that will reflect and absorb the radiation, thus increasing the emissivity and absorptivity of the material when compared to a fully dense structure. Additive manufacturing allows for porosity control through the laser’s energy density which allows for controllable porosity gradients. While porous materials have been extensively studied for mechanical and conductive thermal properties, their radiative properties, particularly in additively manufactured systems remain unstudied. By utilizing Monte Carolo ray tracing, an analytical model has been developed to predict the viability of these designs, and the model has been verified using the view factor of a capped sphere. Based on the results of this model, a high porosity of 80% is expected to increase the surface emissivity by up to 3.3x the base emissivity. This work aims to numerically characterize the emissivity and absorptivity of surfaces produced via these three methods, including laser etching, meso-surfaces and porous additively manufactured material and assess its suitability for spacecraft thermal control. Prototypes of the surfaces have been produced and tested via an emissometer. Laser etched surfaces show a 3.5 increase in emissivity values compared to a base printed surface. Likewise, meso-surface features increase the emissivity of the surface 2.5x above the base emissivity of the printed material. The anticipated outcome of this work is a tailored emissivity surface produced via additive manufacturing, offering a simple and scalable alternative to conventional spacecraft thermal management systems which could be produced via additive manufacturing in a space environment.

Thermal Control and Protection-42
Lunar Surface Crater Thermal Effects on Lander Radiator Performance
Will Grier, Lisa Erickson, William Birmingham, and Tai Valdes

Lunar surface craters smaller than the spatial resolution of surface meshes used in typical Lunar surface thermal models (10 to 60 meters per pixel) may impact the accuracy of thermal model extrema predictions. The goal of this study is to investigate the thermal sensitivity of representative lander systems with realistic thermal surface orientations in bare and cratered terrain environments at relevant Artemis mission locations. This thermal analysis task investigates the impact of lunar surface craters on lander radiator performance by comparing heat rejection capability results between bare and cratered terrain environments. This study examines external body-mounted lander radiator thermal performance across varying lander heights (5.5m, 20m, 50m) and radiator orientations (horizontal, 45º tilted, and vertical) at two representative Artemis mission latitudes (-89.5ºS and -82.5ºS), spanning from the Shackleton Connecting Ridge to Mons Mouton Plateau.

Thermal Control and Protection-43
Ground Slope Effects on Lander Radiator Performance
Lisa Erickson

Near and far-field slopes on the lunar south pole can reduce a lunar vehicle’s heat rejection. However, it is common for thermal analysts to not want to include the additional complexity of slopes in their lunar surface models. This presentation describes sensitivity studies designed to provide insight into how much radiator performance degrades due to nearby slopes for rover and lander-sized assets near the south pole. It aims to answer the following questions: (1) What is the worst-case distance for vehicles near local slopes? (2) Can steady state analysis be used to identify the worst case? (3) Are there cases where we can ignore a nearby slope?
To answer these questions, two sensitivity studies were performed that examined how compared to a flat model radiator heat rejection drops from slopes. Study #1 used a 5.5m tall reduced-order asset to represent a small-lander/rover, with positions ranging from 25m onto the slope to 25m away from its base. Study #2 used a 20m asset to represent a tall lander, with positions ranging from 100m onto the slope to 1km away from its base. The assets had body-mounted horizontal, 45° inward-tilted, and vertical radiator orientations. The impact of slopes on different radiator orientations and on radiators pointing in different directions (i.e., slope versus sun) was investigated. The studies were conducted at a polar location of -82.5°S to capture the incident solar flux at the northmost, and therefore the hottest, range of potential Artemis landing sites (Mons Mouton Plateau).
Results show that minimum total heat rejection of all radiator panels occurs just before the maximum slope the asset can encounter. The per-panel worst case depends on the panel’s view of the slope or sun. For example, panels that view the slope had the worst case just before the slope, where the worst case for horizontal panels was on the slope itself. Study #1 showed that steady state cases can be used to predict worst case radiator heat rejections. Study #2 showed that it is possible for a slope to be far enough away that its impact on radiator heat rejection is negligible or is accounted for in design margin. However, it is also possible to experience a >10% drop in radiator heat rejection from slopes outside a typical landing radius of 100m. Therefore, the presence of slopes should be considered when attempting to predict worst-case hot scenarios.

Thermal Control and Protection-44
Thermal-Fluid Analysis of a Liquid-Cooled Battery Module for Electrified Aircraft
Jarred Wilhite and Erik Stalcup

The development of safe, energy-dense batteries is critical to advancing hybrid electric and fully electrified aircraft propulsion. Achieving this capability requires a thermal management system that can maintain battery performance and safety under demanding operational conditions. The objective of this project is to support the maturation of next-generation lithium-ion batteries for electrified aircraft by conducting performance testing on integrated battery modules, specifically a 2-cell series configuration module housed within an aluminum enclosure. Designed to operate at a nominal 7.2 V with charge rates up to 9.5 A and target discharge rates up to 40 A, the module powers an electric motor and DC-DC converter, generating substantial thermal loads that must be effectively managed to increase the usable energy and power density of electrified aircraft.
To address these thermal challenges, this study presents the development and thermal-fluid analysis of a liquid-cooled thermal management system (TMS). While the full aircraft architecture utilizes an eight-string configuration, the present work evaluates a representative single-string water coolant loop to characterize baseline performance. The active cooling loop circulates water through a reservoir, pump, leading-edge heat exchanger, the battery module, and a duct heat exchanger. Key performance metrics including component-level temperatures, mass flow rates, and pressure drops are quantified across the loop. These results are further extrapolated to assess the system level impacts such as TMS weight and power requirements.
Preliminary testing has successfully validated the single-string coolant loop at discharge rates up to 12 A, with ongoing evaluations aimed at characterizing thermal performance as the system scales toward the 40 A target load. Overall, this analysis provides foundational insights into the viability and optimization of liquid-cooled thermal management architectures for future electrified aircraft.

Thermal Control and Protection-45
Thermal Analysis of Large Format Lithium-Ion Batteries in Thermal Runaway
Ethan Sparks

Lithium-ion (Li-ion) batteries can deliver electrical power across a large range of applications; large format versions of these batteries are common in aerospace applications for their low mass and energy dense properties. Li-ion cell usage has increased because they are more power dense than fuel cells and have overall higher performance over lead-acid/NiCd batteries. However, Li-ion cells can experience Thermal Runaway (TR) via electro-chemical, mechanical, or thermal abuse. TR is a phenomenon where the stored energy in a cell is rapidly released along with vented gases and other effluents when the cell’s internal heat generation has surpassed its ability to dissipate that heat. If a Li-ion cell enters TR, a chain reaction (propagation) due to heat transfer within the battery can occur, potentially causing TR within other cells. Understanding a battery’s susceptibility to propagation is necessary to evaluate risks associated with TR phenomena. Thermal modeling anchored to battery testing is crucial for evaluating a battery design and its ability to prevent propagation when experiencing single cell TR. The presentation for TFAWS 2026 will cover considerations for battery TR thermal analysis, such as testing methods, influences on the cells, failure mechanisms, and TR parameters, as well as a walkthrough of a TR thermal analysis example.

Thermal Control and Protection-46
Thermal Performance of a Lunar Rover RESS
Konrad M Brown, Trisha Matthews, Peter Andruskiewicz, and Derek Lahr

This work summarizes the design, simulation, and test of thermal performance for a lunar rover battery module using 21700 cylindrical lithium-ion cells. The module uses an integrated heat spreader and potting material to modulate the temperature of the cells, ensure temperature uniformity, and provide protection against thermal runaway (TR) propagation and damage. The heat spreader has a novel design to minimize its mass, using both a three-cell grouping and a layered construction along the cell height. The resulting arrangement provides heat-conducting ribbon features to convey heat to the cold plate and heaters for temperature control, while also permitting two-dimensional heat conduction across the module during a thermal event.
To select a design for the conducting ribbons, nominal operating conditions were analyzed virtually on a two-dimensional section through the battery module. Parametrical studies of this geometry included ribbon thickness, the number of cells along a ribbon, and the number of cells between ribbons. Designs were simulated at variety of charging rates and durations to evaluate the cell temperature distribution and uniformity. A subset of “best” designs was identified that met the temperature criteria while minimizing the entire module mass and ribbon material.
Once preliminary “best” designs were chosen, TR propagation was simulated virtually on a conduction-only “mini-module” representative of the design. A coarse 3D FE mesh allowed heat transfer to be comprehended in both axial and radial directions while minimizing computational cost. The geometry was highly parameterized to enable study of potting, heatsink dimensions, and material properties. A conservative assumption was made to apply nominal cell energy to the desired trigger cell in a triangular-shaped, temporally varying profile. Propagation of thermal runaway to other cells was determined by thermal runaway onset temperature criteria. The model predicted that TR propagation would not occur.
As final verification, a physical thermal runaway test was performed with a mini-module containing 22 cells bused in parallel to demonstrate that TR propagation between cells would not occur and that the mini-module would still be operational despite losing a cell. The test was performed in near-vacuum and ambient temperature using single-cell heating as the initiation method. The triggered cell vented and experienced TR, ejecting most of the jellyroll from the can. No TR propagation to other cells was observed. Post-test the module was functional and retained >95% electrical capacity, performing consistently with expectations, analytical predictions, and the project goals.

Thermal Control and Protection-47
Passive Thermal Runaway Propagation Prevention to Improve Safety of Li ion Batteries
Dr. Vijay Devarakonda, Tanvi Gupta, Ranadip Saha, and Partha Mukherjee

Lithium-Ion batteries are vital to powering key technologies across aerospace, defense, and the automobile industry, often chosen due to their high energy density, long cycle-life, and affordability. As these industries evolve to become increasingly robust, battery failures due to mechanical or thermal abuse, overcharging, overdischarging, and prolonged exposure to extreme weather conditions must be addressed. Of particular interest is their potential escalation towards uncontrolled exothermic reactions within individual battery cells that encase many flammable materials, a condition that is referred to as thermal runaway. Abuse, abnormal use, or mechanical defects in any one cell within a large battery pack could lead to thermal runaway that generates enough heat, flammable content, and debris to trigger chain-reactions in neighboring cells, causing thermal runaway to propagate within the pack and leading to large-scale battery fires in industrial applications.
Increased regulatory oversight brought in by such fires has led to the development of several safety mechanisms, such as positive temperature coefficient elements, shutdown separators, and vents at the cell level; current interrupt devices, thermal management systems including cell cooling technologies, and modularized designs with partitions and insulation layers at the pack level. Even though these systems have led to marginal improvements in battery performance and safety, they are generally incapable of preventing fires once a cell begins thermal runaway. For example, liquid coolant systems integrated within electrical vehicles can cool individual cells and reallocate heat throughout the vehicle, yet they are unable to handle the large heating rates (>1°C/sec) that occur as a cell approaches the onset of thermal runaway. Currently, the resultant battery fires are fought by pouring copious amounts of water from outside the host system in ground-based applications like electric vehicles. This approach is resource and time inefficient, as the battery pack cannot be used after extinguishment, and is not applicable for aerospace applications due to the mass penalties associated with the fire extinguishment system.
In collaboration with Purdue University and with support from AFRL, Analytical Scientific Products has developed a low-cost, cell-agnostic technology to prevent thermal runaway propagation in battery packs, improving the cycle life and safety of energy-dense Li ion batteries. This is a complimentary technology to battery thermal management systems that employ water based liquid coolants to cool individual cells in battery packs. The product works with an inside-out approach, where the energy from one or more cells that experience excessive heating enroute to thermal runaway is used to release liquid coolant from the thermal management system on those cells. Upon contacting the failed cell’s surface, the liquid coolant undergoes local boiling to quench the overheated cells preventing further exothermic reactions and thermal runaway, while neighboring cells are cooled to prevent any propagation. By doing so, this technology ensures that battery failure is confined to a few cells in a single module, and that overall functionality is sustained. The simplicity of this concept also allows this cooling system to be implemented in packs of any size, shape, or form. This paper overviews the single-cell and component-level tests that demonstrate our technology’s modularized approach of preventing thermal runaway propagation.

Thermal Control and Protection-48
Geometric Optimization of a Passive Ram-Airflow Cooling Duct for Li-ion Battery Thermal Management in Electric Semi-Truck Vehicles
Rohan Jain and Bhargav Narayanan

Battery-based electric semi-trucks offer a promising alternative to traditional diesel trucks for highway freight due to reduced pollution and environmental impact. However, overheating of the Lithium-ion battery packs may cause severe safety problems and efficiency loss. This work presents an experimental and simulations-based investigation of passive ram air based cooling of the battery pack of an electric semi-truck. This approach utilizes the sustained high-speed airflow available during highway operation without the need for a dedicated fan/blower. This work seeks to optimize duct geometry for passive systems, something not found in current literature. With the goal of maximizing throat airspeed while minimizing pressure loss, nine candidate duct designs with various convergence angles and throat heights are 3D printed and aerodynamically tested in a low-speed wind tunnel. Duct exit velocities are measured with a pitot probe at multiple freestream speeds. Using continuity theorem, throat velocities and massflow capture rates are calculated. Computational Fluid Dynamics (CFD) simulations are carried out in ANSYS. Measurements and simulations identify a key tradeoff between massflow capture rate and speedup ratio of ducts. Aggressively converging ducts are found to offer high throat velocities but at low capture ratios (eg. 20% capture rate with 53% speedup), whereas mildly converging ducts demonstrate the opposite (eg. 98% capture rate with 6% speedup). This tradeoff addresses the varying thermal management needs for different electric semi-trucks. CFD simulation results are found to agree well with experimental measurements, which establishes the simulations developed in this work as a useful tool for evaluating aerothermodynamic performance of candidate designs without expensive time-consuming measurements. Measurements and simulations presented here may contribute towards the development of a promising thermal management technique for electric semi-trucks.

Thermal Control and Protection-49
The ATA Technolgoy: Advanced Active Thermal Control for Space Applications
Dr. Lucas S Anderson, Charles Swenson, Miguel Nunes, and Robert Wright

The Active Thermal Architecture (ATA) is an advanced active thermal control system based on a novel single-phase mechanically pumped fluid loop architecture. The ATA has been developed by Utah State University, Orion Space Solutions, and the Hawaii Spaceflight Laboratory for the upcoming Active Cooling for Multispectral Earth Sensors (ACMES) mission. The ATA is designed to solve a fundamental problem facing modern CubeSats and small satellites: how to manage, transfer, and fundamentally control large, concentrated heat loads within severely constrained space platforms. As payloads such as hyperspectral imagers, onboard edge processors, and cryogenically-cooled sensors push power densities to levels far beyond those of even traditional, large satellite classes.
The ATA addresses this challenge through a highly integrated, miniaturized single-phase mechanically pumped fluid loop (MPFL), capable of rejecting up to 150W of thermal load while maintaining precise cold plate temperature stability to within ±2.5° C. Key innovations include 3D printed Ultrasonic Additive Manufacturing (UAM) heat exchangers and radiators, deployable tracking radiators that offer variable area and view factor capability, and rotationally flexible fluid joints. Redundant micropumps with triple-seal designs offer long-duration spaceflight reliability. As an additive technology, the ATA reduces manufacturing costs and schedule.
At a system level, the ATA offers a scalable counterpart to power generation, enabling a power-in-power-out approach to spacecraft design. Ultimately, the ATA will serve as a critical enabling technology for the next era of space exploration.
This presentation will detail the design and operation of the ATA technology as well as its integration with the Hyperspectral Thermal Imaging 2.0 (HyTI 2.0) instrument on ACMES. In addition, performance modeling results and available ground-based performance metrics will be discussed. The ATA technology is currently in production with a ground qualification campaign scheduled for the 3rd quarter of 2026 and a demonstration flight mission in 2027.

Thermal Control and Protection-50
A 2026 Update on the Active Mechanical Pumped Fluid Loop (MPFL) Projects at the Jet Propulsion Laboratory (JPL)
Hared A Ochoa

Over the last decade, several Mechanical Pumped Fluid Loop (MPFL) systems for thermal control that JPL has been responsible for have continued to progress. Two flagship missions, Mars 2020 (M2020) and Europa Clipper, and three International Space Station (ISS) instruments (ECOSTRESS, OCO-3, and CAL) have all successfully advanced from design to implementation, launch, and operations. Meanwhile, the Mars Science Laboratory (MSL) Curiosity rover has continued to operate its MPFL flawlessly on the surface of Mars for over 13 years. Earlier this year, JPL also completed the Preliminary Design Review (PDR) of a comprehensive dual MPFL layout for the thermally demanding Mars Sample Return Lander (SRL) mission, resulting in more capable MPFL architecture that will most certainly benefit future missions.
Launched in 2018, the Ecosystem Spaceborne Thermal Radiometer Experiment on Space Station (ECOSTRESS) has since operated as part of the ISS measuring high resolution temperature and emissivity of Earth’s surface and monitoring the transpiration of plants. Similarly, the Orbiting Carbon Observatory (OCO-3) launched and joined the ISS in 2019, to monitor carbon dioxide and solar-induced fluorescence. Both instruments utilized the Japanese Experiment Module Exposed Facility (JEM-EF) active thermal control system onboard the ISS. The Cold Atom Laboratory (CAL), also launched in 2018, utilizes a series of custom aluminum heat exchangers that comprise a fluid loop which carries waste heat out of the payload. The payload itself uses lasers to cool temperatures to near absolute zero.
M2020 and Europa Clipper utilized their own unique MPFL systems as part of their thermal control architecture. M2020 reutilized the MSL fluid loop design to harvest waste heat from its avionics and its Radioisotope Thermoelectric Generator (RTG) and redistributed it as necessary. Since its launch in July of 2020, the M2020 Perseverance rover has continued to look for signs of conditions for life on Mars, along with collecting samples of rock and soil for a future lander to collect. Europa Clipper also capitalized on the success of the MSL MPFL system and utilized its MPFL to thermally control a large propulsion system with the power efficiency necessary to enable a deep space solar array mission to the Jovian system. Launched in 2024, Europa Clipper has successfully completed 1.5 years of cruise and is scheduled to arrive at Jupiter in 2030.
The more recent program, SRL, has continued to push the design and versatility of the MPFL architecture. The challenging SRL thermal requirements also motivated JPL to improve upon the heritage pump and passive thermal control valve designs implemented on previous missions. Two in-house efforts to develop a Pump Unit with Long-life for Space Environments (PULSE) and an Active Control Thermal Valve (ACTV) were initiated to further improve the capability and long-term reliability of spacecraft MPFL systems.
This presentation will summarize the status of each of these JPL MPFL projects, including the observed in-flight performance of the currently operating fluid loops.

Thermal Control and Protection-51
RANS-based assessment of transpiration cooling effectiveness in a curved porous nozzle wall exposed to high‑enthalpy supersonic flow
Dr. Ram Adhikari

Transpiration cooling has demonstrated strong potential for next‑generation launch systems, offering more uniform thermal protection while avoiding the structural penalties associated with film cooling. Foundational theoretical works, supported by recent experimental and numerical studies, have shown that mass injection through a porous wall can significantly reduce convective heat flux by thickening the boundary layer and decreasing wall shear. Cooling effectiveness, however, is highly sensitive to parameters such as coolant distribution and wall curvature. The present study focuses on curved wall geometries, which introduce additional complexity due to curvature‑induced modifications of the boundary layer and pressure gradients. For example, concave curvature can promote the formation of Görtler vortices, enhancing heat transfer, whereas convex curvature tends to stabilize the boundary layer and reduce thermal loads. Despite the relevance of these effects to nozzle designs, only a limited number of studies have examined transpiration cooling within a curved rocket nozzle environment. This gap highlights the need for a detailed numerical investigation to better understand how curvature influences cooling effectiveness and local heat‑flux distributions. The gas‑phase flow is modeled using compressible RANS equations with an SST- turbulence closure, while the porous wall is represented through a volume‑averaged energy equation. The numerical model is validated for a case with flat geometry by comparing the numerical results with the experimental data available in the literature, at Mach 5 and film temperature of 500K. The simulation results agree well with the experimental data. Then the numerical simulation is extended to curved wall geometries. Simulations conducted on six candidate designs of curve wall geometry show that wall curvature significantly alters cooling effectiveness, with concave regions exhibiting elevated heat flux and stronger sensitivity to blowing ratio. Transpiration cooling reduces peak thermal loads substantially, but its performance depends strongly on local wall curvature. The findings of this ongoing study provide new insight into the design of transpiration‑cooled thermal protection systems, particularly for rocket nozzle design problems.

Thermal Control and Protection-52
Parametric-Based Heat Rejection Trade Study for Lunar and Martian Surface Operations
Noah Andersen and Thomas Chen

Establishing and maintaining a sustained presence on the lunar and/or Martian surfaces will require a diverse portfolio of surface elements (e.g., habitation, mobility, power generation, etc.). Many of these systems generate excess heat that must be rejected across a wide range of magnitudes, temperatures, and duty cycles and under variable environmental conditions. To identify the most promising heat rejection approaches for this diverse portfolio, a heat rejection trade study was conducted to evaluate the performance of different technology approaches across a spectrum of surface environments and heat-load requirements. The trade study consisted of three stages: (1) development of a parametric-based modeling framework, (2) creation of a database of heat rejection technologies, surface elements, and environmental conditions for the Moon and Mars, and (3) execution of a quantitative analysis of various heat rejection technologies across different operating conditions and surface elements.
The modeling framework is developed in Python and Excel to prioritize small model size and hence low computational cost to enable large parametric sweeps while avoiding the reliance on proprietary software. Individual heat rejection processes are represented as simple Excel models, and a centralized Python script interfaces with the models to coordinate the parametric study. These simple sizing models were developed to take heat load requirements and environmental parameters as inputs and compute mass, power, and volume as outputs. Rather than assess each heat rejection technology separately for each surface element, a unified parametric space was developed to evaluate all technologies across all elements. This parametric space includes factors related to heat load (e.g., magnitude or temperature) and environment (e.g., surface temperature, sky temperature, solar flux).
This effort generated a database containing information on over 60 heat rejection technologies and 30 surface elements. For each surface element, the expected heat rejection requirements were documented and analyzed to determine the most common needs shared across all elements. Environmental conditions at various lunar and Martian latitudes were also established for worst-case hot and worst-case cold scenarios.
High-fidelity heat rejection models are currently under development. Preliminary trades between heat rejection technologies including radiators, venting technologies, convective coolers, and more have been conducted to identify promising options. This presentation will summarize the preliminary trade results and provide an overview and discussion of the expected heat loads and thermal environments for sustained surface operations on the Moon and Mars.

Thermal Control and Protection-53
In-Situ Heat Rejection to the Martian Atmosphere
Richard Gregory Schunk

In order to provide in-situ heat rejection to the thin Martian atmosphere, approaches relying on forced convection and buoyancy flow are conceptualized. A convective radiator concept utilizing a fan to draw local atmosphere through an expanded metal foam heat exchanger panel is developed. To provide a pumped fluid loop interface, flow tubes are embedded in the heat exchanger panel and heat is transferred from the tubes to the foam via conduction and subsequently convectively to the environment. To forgo the use of a specialized high throughput fan, a concept utilizing a buoyant chimney tower to draw atmospheric flow through a low pressure drop heat exchanger is also considered. Both concepts may best align with moderate- to high-temperature heat loads that provide a significant delta temperature relative to the ambient atmosphere temperature. While eliminating dust accumulation on radiator surfaces, some type of serviceable filtration or screening may be required to prevent accumulation of dust inside the expanded metal foam heat exchanger for both concepts.

Thermal Control and Protection-54
An Update on Making Ultrasonic Additive Manufacturing (UAM) of Liquid Cold Plate Heat Exchangers a Reality for NASA and Aerospace Thermal Management Applications
A J Mastropietro, Gordy Cucullu, Nicholas Keyawa, Teri Juarez, Scott Innes, Scott Roberts, Jason Riley, Luke Walker, Lucas Anderson, and Charles Swenson

Since 2013, the Jet Propulsion Laboratory (JPL) and NASA have been working jointly with Fabrisonic LLC to develop Ultrasonic Additive Manufacturing (UAM) technology to more readily fabricate both small and large liquid cold plate aluminum heat exchangers and radiators with embedded cooling channels without having to separately manufacture and bond on complex tubing assemblies. Through several NASA SBIRs and University SmallSat Technology Partnerships, the UAM technology has matured to the point that hermetic leak rates, burst pressures, and thermal efficiencies for applications that require channel sizes from ~0.2 inches up to ~0.3 inches in diameter are extremely compelling. However, more work is admittedly required around characterizing the material property allowables for aluminum structures that are built in this manner, especially if they are to serve a dual function as primary structure on a spacecraft.
Throughout the last decade, some partial measurements of material property allowables have been made on UAM aluminum structures, but the results have been statistically underpowered to draw firm conclusions. To achieve spaceflight qualified parts and encourage a more ready posture of wider adoption, UAM technology requires an infusion of funds to finally develop consistent A-Basis material property allowables in much the same way that they have been defined for titanium aerospace parts that are now routinely additively manufactured using Direct Metal Laser Sintering (DMLS).
Despite this restricted availability of material property allowables data, JPL recently performed an exhaustive trade on the UAM method of liquid cold plate heat exchanger fabrication for the Mars Sample Return Lander project. Ultimately the JPL-led team selected UAM as the baseline fabrication method for several planned sizeable (~1.8-meter x ~0.8-meter and ~0.8 meter x ~0.7meter) avionics mounting panels, some of which were also planned to be primary structure. As part of a risk reduction activity surrounding the decision, an in-house campaign to develop a limited S-Basis material property allowables dataset was initiated.
This presentation will present a status on the JPL-led S-Basis material property allowables effort, as well as provide an update on recent ongoing relevant work at Fabrisonic LLC including a collaborative effort with Orion Space Solutions that pertains to the upcoming Active Cooling for Multispectral Earth Sensors (ACMES) Mission.

Thermal Control and Protection-55
Coupled Thermal–Mechanical Modeling of a Deployable Radiator with Partial PCM Transformation and Runtime Radiation Articulation
Joshua H. Taylor, Anthony Lococo, Franklin L. Robinson, Asher Leff, Darin Sharar, Rydge B. Mulford, and Brian D. Iverson

Effective thermal management is critical for CubeSat missions to ensure reliable operation of onboard electronics within the variable thermal environment of space. This work presents the Thermal Desktop modeling methodology developed for a passive thermal control concept consisting of redeployable radiator fins containing solid–solid phase change material (PCM) actuated by shape memory alloy (SMA) elements. When internal temperatures rise, the PCM absorbs heat through a solid–solid phase transition before the SMA actuator deploys the fins radially outward, increasing the effective radiative area and establishing a thermal pathway for heat rejection to deep space. As the system cools, the PCM reverses phase and the fins return to their stowed configuration, enabling autonomous deployment without electrical power. The present work focuses on two modeling challenges that arise when simulating this coupled thermal–mechanical system in a SINDA/FLUINT-based environment: (1) accounting for partial PCM phase transitions with hysteresis, and (2) dynamically updating radiation exchange factors to reflect fin articulation during a transient solution.
Standard PCM modeling in Thermal Desktop applies a single specific heat curve representing either full heating or full cooling behavior. For the case when the PCM reaches a partial transformation in phase during heating (or vice versa during cooling), a hysteresis gap of 10 °C or more can occur between heating onset and cooling onset (departure from sensible heating/cooling cp values) and can introduce energy conservation errors when employing a single heating and cooling specific heat dependency on temperature. To address this, a node-level state machine algorithm was implemented entirely within the Thermal Desktop Logic Manager using SINDA/FLUINT FORTRAN. The algorithm tracks each PCM node independently through four states governing whether the node follows the heating curve, cooling curve, or a sensible-only bridging state for partial heating or cooling following reversal at partial transformation. The percent transformation is tracked in addition to temperature, ensuring conservation of energy regardless of the number of heating or cooling changes that occur. The implementation spans two PCM submodels, with state data stored in User Defined FORTRAN Arrays (UDFAs) to circumvent SINDA preprocessor restrictions on variable-indexed node access. Validation against an independent MATLAB finite difference model implementing identical logic showed maximum absolute errors below ~0.01% in temperature, effective specific heat, and tracked percent transformation.
To address heat exchange that changes as function of position, the articulation of radiation conductors between nodes was implemented as a runtime system that switches between pre-computed RadCAD datasets, one for each fin position. FORTRAN subroutines called from the Logic Manager were used to load Radks datasets at startup and overwrite the full set of radiation conductors when actuation position changes. Validation against a sequential baseline of several fixed-geometry RadCAD cases showed deployment timing and temperature agreement within ≲0.1% across all thermal submodels, confirming that the runtime approach reproduces results consistent with the standard RadCAD workflow while adding approximately 20% to solver runtime.
Together, these two modeling approaches enable fully coupled, transient simulation of temperature-dependent fin deployment, radiative view factor changes, and partial PCM energy storage within a single Thermal Desktop model. Ongoing work with the modeling framework is to reproduce experimentally observed deployment behavior from TVAC testing and to evaluate system performance across a range of orbital conditions and heat loads.

Thermal Control and Protection-56
Development of a Thermal Radiator Optimization Tool with Alternate Coolants
Richard Gregory Schunk

The development of a modeling tool to optimize the design of a thermal radiator based on capacity, rejection temperature, geometry, type and coolant is presented. The radiator size is determined from the desired capacity, environmental conditions and rejection temperature. Flow through the radiator is derived from the overall heat load and prescribed inlet to outlet delta temperature for a given coolant. The tool considers radiator geometry (i.e., tube spacing and diameter, face-sheet thickness, overall size, etc.) and general type (i.e., parallel/manifold versus serpentine) in the optimization. Selection from a handful of potential coolants is also available in the tool which primarily affects necessary tube diameter to maintain turbulent flow and minimize pressure drop. An Equivalent System Mass (ESM) approach is utilized to include pump power in the optimization. Results from the tool and potential future enhancements are also discussed.

Thermal Control and Protection-57
ASHRAE Climatic Design Methodology Applied to Diurnal Thermal Simulation
Dr. Daniel Dannelley, Christopher Henry, and Vitaly Meyzler

This study reviews methods for simulating hot ground diurnal environments and their impact on thermal design. Elevated ground temperatures can degrade performance in transient thermal systems by increasing thermal mass, which decreases range, or increase cooling requirements impacting power and infrastructure complexity. A new approach is introduced that integrates MIL HDBK 310 and ASHRAE global climatic data, applying building external load calculation techniques to efficiently predict in situ RTX system temperatures. A case study compares missile thermal response during MIL HDBK 310 Extreme Hot A1 environmental exposure with the response at specific geographic locations within the A1 envelope using ASHRAE data. Both methods incorporate spatial and temporal calculations for solar irradiation, ambient convection, sky radiation, albedo, and ground surface heating. Key advantages of ASHRAE methods include precise characterization of direct and diffuse solar loads, additional site specific climate detail, statistical trends that support future predictions, and high quality weather information used by the professional engineering community. An analytical tool was created that interprets ASHRAE weather data to calculate solar irradiance on surfaces of any orientation and generates beam, diffuse, ground reflected, and convective loads for use in advanced tools such as Thermal Desktop. Additionally, the tool includes a closed form transient surface temperature solution based on ASHRAE Design methodology supporting rapid sensitivity studies and design trades, enabling more efficient cycles of learning. These benefits improve missile thermal requirements definition, enhancing system performance and quality. The methodology is broadly applicable to other ground based systems such as shelters, radar systems, and additional fielded RTX products, representing a significant advancement that supports optimized solution development.

Thermal Control and Protection-58
Hot Moon: Designing a 20 kg Rover to Survive the Heat of Equatorial Lunar Noon (Lunar Outpost)
Izzy Golemme

CUI-02
Development of Self-Regulating Heater Prototype for Space Applications
Dr. Jianjian Wang, Srujan Rokkam, and Nathan Van Velson

Self-regulating heaters (SRHs) incorporate a temperature setpoint directly into the resistive heating material, enabling autonomous temperature control without separate temperature sensors. As the heater reaches its designed setpoint, the resistance of the SRH material increases sharply, limiting further heat generation and preventing overheating. SRHs are widely used in the petrochemical and automotive industries for applications such as pipe freeze protection and seat warmers. However, existing SRH technologies are generally not suitable for space applications because of their large form factors, limited flexibility, lack of reusability, or inability to withstand the space environment.
This work presents a lightweight, flexible self-regulating heater prototype specifically designed for spacecraft thermal control applications. The heater is based on a positive temperature coefficient (PTC) polymer composite consisting of conductive nanoparticles dispersed within a polymer matrix possessing a high thermal expansion coefficient. As the temperature rises above the designed setpoint, thermal expansion of the polymer matrix separates the conductive particles, disrupting conductive pathways and causing the composite film to become highly resistive. The demonstrated SRH prototype exhibits a setpoint temperature of approximately 1.5 °C, with the heater resistance changes by orders of magnitude when the temperature varies from -40 °C to 80 °C. This technology has the potential to significantly reduce the mass, power, and system complexity associated with conventional spacecraft propellant thermal control systems. The heater responds instantaneously to temperature variation and automatically shuts down or turns on the heating based on the environmental temperature, without the need for external temperature sensors.

CUI-03
Heat Pipe Thermal Management for High-Density Hall Thrusters
Brett Leitherer, Quang Truong, Calin Tarau, Jeff Diebold, Benjamin Jorns, and William Hurley

NASA is interested in developing higher power-density Hall thrusters (>20 kW) for future deep space exploration missions. This will pose significant thermal challenges as the heat generated inside the thruster is currently dissipated by conduction through multiple low thermally conductive materials in addition to radiative cooling. The elevated temperature of the Hall thruster components locally induces high thermal stress that can limit the operating power of the thruster. Under a NASA Small Business Innovation Research (SBIR) program, Advanced Cooling Technologies, Inc. (ACT) in collaboration with Plasmadynamics and Electric Propulsion Laboratory at University of Michigan (UM) explores a novel passive two-phase thermal management solution for high power-density Hall thrusters. The proposed design consists of embedding multiple high temperature heat pipes (HPs) within the configuration of the Hall thruster which will directly transport heat from the high heat load and high temperature regions to a back radiator panel extending out of the Hall thruster. Key benefits of the proposed solution include (i) enabling the Hall thruster to carry significantly higher overall power for the same footprint, (ii) introducing a new heat transfer path having lower thermal resistance to protect temperature-sensitive components from thermal failure, (iii) independently controlling the heat pipe operating temperature, (iv) maximizing radiator performance to minimize required radiator size and mass. This presentation summarizes the feasibility study activities of the proposed concept, including thermal modeling and proof-of-concept high temperature 3D structure HPs prototype demonstration testing within a 9kW Hall Thruster.

CUI-04
ISS Battery Charging System Thermal Design and Development
Dr. Siraj Jalali

The thermal model of the ISS Battery Chargers was developed and test correlated. The chargers were analyzed in ISS Airlock at 14.7 psia and 10.2 psia 113°F environments, and the heat flow to the cold plate was determined to check against the allowable limits. All major electronic components were modeled and analyzed to determine if the extreme temperatures of electronic components are within their operating limits. The analysis also determined the Battery Chargers extreme temperatures to compare with bare hand touch temperature limit. There are two modes of operations for the EBOT Chargers, Charging and Discharging modes, the chargers were analyzed for both modes.

CUI-05
Space Launch System Core Stage Base Aerothermodynamics Post-Flight Reconstruction for Artemis II
Dr. Manish Mehta, Brandon Mobley, and Samantha Summers

Artemis II was the first crewed integrated test flight of the NASA heavy-lift, human-rated, exploration-class launch vehicle, Space Launch System (SLS), and Orion spacecraft. Artemis II successfully launched from Pad39B at NASA Kennedy Space Center on April 1st, 2026. The integrated test flight was composed of launch and ascent of SLS vehicle from lift-off to RS-25 main engine cut-off (MECO), interim cryogenic propulsion stage (ICPS) in-space flight and Orion’s trajectory around the moon and landing in the Pacific Ocean which occurred on April 10th, 2026. The SLS total thrust of 8,800,000 lbf was powered by four LOX/LH2 RS-25 engines and two 5-segment solid rocket boosters. As a result, the base flow field for this vehicle were highly complex and observed the highest ascent heating environments on the launch vehicle. SLS Core Stage base aerothermodynamics covers rocket plume-induced convection and radiation of the vehicle’s aft region during powered flight from lift-off to MECO. This work discusses the SLS base flow physics observed during Artemis II flight and prelaunch and comparisons of post-flight reconstruction with Artemis I flight data, pre-flight heating models and Space Shuttle data.

CUI-06
SCIFLI Artemis II Orion Reentry Observation Summary
Carey Scott Jr., Matthey T. Boyda, Kylel D. Scott, Arianna Haven, Meaghan M. McCleary, Shann J. Rufer, Alireza Mazaheri, Chris Johnston, and Andrew C. McCrea

Artemis II was the first crewed flight into deep space in over five decades, during which the crew aboard the Integrity spacecraft performed a lunar flyby before returning to Earth. Airborne imaging systems were deployed by the NASA Scientifically Calibrated In-Flight Imagery (SCIFLI) team to produce visual, spectral, and infrared observations of the Orion crew module (CM) during the reentry phase of the mission. The SCIFLI team coordinated a multi-platform airborne observation using six fixed wing aircraft outfitted with a variety of imaging payloads to enable visual tracking and measurement of the CM over a wide range of flight conditions, ranging from Mach 32 through splashdown. Spatially resolved & radiometrically calibrated thermal and data were collected by each airborne imaging platform to support engineering analysis, verification, and validation of thermal protection system (TPS) and landing & recovery system (LRS) performance. The SCIFLI assets observed the Orion service module (SM) breakup, and maintained visual track of the crew module shortly after entry interface, providing confirmation of Integrity’s status through the plasma blackout until Integrity splashed down near San Diego, CA. This presentation will provide an overview of the mission including a summary of technical planning, engineering design, flight operations, and mission execution activities required for the successful observation of the Artemis II reentry using airborne imaging platforms.