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NASA NTRSDate not supplied

Material Properties and Modeling of Room Temperature Vulcanizing Silicone

Room Temperature Vulcanizing silicone (RTV) is a high-temperature adhesive that has successfully been used as a gap-filler between Thermal Protection System (TPS) tiles for heatshields on numerous missions. It is also used to bond instrumentation plugs such as temperature and pressure sensors into the heatshields. While RTV has been traditionally assumed to be a non-porous and non-ablating material, numerous experiments have shown that RTV pyrolyzes and becomes highly porous as it is heated. Heating RTV has also shown swelling, or intumescence, which can pose unique problems that lead to roughness induced boundary-layer transition, surface oxide formation and contamination of heat shield sensors. Therefore, it is crucial to understand and model the intumescence phenomenon of RTV. As data for RTV material properties is limited, the first step in modeling RTV is to collect material properties such as pyrolysis mass-loss, microstructure change, virgin and char porosity, etc. which was performed in our initial study. Additionally, thermomechanical properties such as Young’s modulus and Poisson ratio are required for modeling the intumescence of RTV, which were taken from literature and the coefficient of thermal expansion was collected using in-situ heating and Micro Computed Tomography (µ-CT) in previous studies. Finally, numerous other properties such as pyrolysis gas properties, virgin and char thermal conductivity and specific heat were compiled from previous experiments and literature into a material database that can be used for simulations. In Porous Material Analysis Toolbox based on OpenFOAM (PATO) [4], structural mechanics coupled with material response was used for simulating the intumescence of RTV as it is heated. However, since the permeability of the material is very low, the pyrolysis gas creates an internal pressure build-up as the material is being heated, significantly contributing to the deformation of the material. To correctly characterize this phenomenon, additional physics models were implemented into PATO's stress analysis solver, and results were compared with RTV dilatometry test data as a preliminary verification case. Future work will include experiments of RTV at the Plasmatron X facility and the in-situ heating cell with µ-CT, and improvement of simulation tools to more accurately model RTV intumescence.

PATO↗

NASA NTRS2014

Screen Channel Liquid Acquisition Device Outflow Tests in Liquid Hydrogen

This paper presents experimental design and test results of the recently concluded 1-g inverted vertical outflow testing of two 325x2300 full scale liquid acquisition device (LAD) channels in liquid hydrogen (LH 2 ). One of the channels had a perforated plate and internal cooling from a thermodynamic vent system (TVS) to enhance performance. The LADs were mounted in a tank to simulate 1-g outflow over a wide range of LH 2 temperatures (20.3 – 24.2 K), pressures (100 – 350 kPa), and flow rates (0.010 – 0.055 kg/s). Results indicate that the breakdown point is dominated by liquid temperature, with a second order dependence on mass flow rate through the LAD. The best performance is always achieved in the coldest liquid states for both channels, consistent with bubble point theory. Higher flow rates cause the standard channel to break down relatively earlier than the TVS cooled channel. Both the internal TVS heat exchanger and subcooling the liquid in the propellant tank are shown to significantly improve LAD performance.

Fuel Depot↗

NASA NTRS1982

Improved multilayer insulation applications

Multilayer insulation blankets used for the attenuation of radiant heat transfer in spacecraft are addressed. Typically, blanket effectiveness is degraded by heat leaks in the joints between adjacent blankets and by heat leaks caused by the blanket fastener system. An approach to blanket design based upon modular sub-blankets with distributed seams and upon an associated fastener system that practically eliminates the through-the-blanket conductive path is described. Test results are discussed providing confirmation of the approach. The specific case of the thermal control system for the optical assembly of the Space Telescope is examined.

George Mikk↗

NASA NTRS2026

Helicopter and Urban Air Mobility (UAM) Laboratory Comparison (HULC) Psychoacoustic Test

The Helicopter and Urban Air Mobility (UAM) Laboratory Comparison (HULC) psychoacoustic test was conducted in July 2025 to compare human annoyance responses to noise from UAM vehicles and helicopters. This document provides a literature review, describes the HULC test methodology including stimuli selection and test administration procedures, and presents results from an initial analysis using multilevel statistical modeling. The test utilized 123 sound stimuli from four helicopters and six UAM vehicles across three flight phases (departure, cruise, and approach), presented to 40 subjects in the Exterior Effects Room (EER) at NASA's Langley Research Center. Subjects rated their annoyance on a 0 to 10 scale while imagining hearing these sounds multiple times daily near their homes. Key findings revealed that, at the same A-weighted sound exposure levels, UAM vehicles and helicopters produced similar annoyance responses across all flight phases, with mean differences of only approximately 1 dB and overlapping confidence intervals. However, at equivalent observer distances, UAM vehicles had inherently quieter operations and lower mean annoyance ratings. These results suggest that UAM vehicles and helicopters may be assessed in a similar manner for community noise impact. The quieter noise characteristics of UAM vehicles may provide advantages for operating in communities and integrating into urban environments.

psychoacoustics↗

NASA NTRSDate not supplied

Artemis II Mission Performance of the Orion Active Thermal Control System

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 those heat 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.

Ellie Thurston↗

NASA NTRSDate not supplied

Extending mARC II Arc-Jet Test Duration via Design and Implementation of Model System Cooling Sleeve

The mARC II is a 30 kW arc-jet facility at NASA Ames Research Center used to generate high-enthalpy flows for low-cost thermal protection system (TPS) technology development. Sustained operation of downstream instrumentation and material samples is constrained by thermal loading transmitted through the arc-jet test environment, limiting achievable run times and experimental throughput. This work presents the design, integration, and validation of a cooling sleeve implemented on the sweep arm drive motor feedthrough to mitigate thermal accumulation during testing. The addition of the cooling sleeve is a simple, robust upgrade that translates directly into enhanced facility capability by supporting longer run durations, reduced turnaround time, and higher throughput.

numerical simulations↗

NASA NTRSDate not supplied

Extending mARC II Arc-Jet Test Duration via Design and Implementation of Model System Cooling Sleeve

sting arm↗

NASA NTRS2015

Numerical Modeling of the Transient Chilldown Process of a Cryogenic Propellant Transfer Line

Before cryogenic fuel depots can be fully realized, efficient methods with which to chill down the spacecraft transfer line and receiver tank are required. This paper presents numerical modeling of the chilldown of a liquid hydrogen tank-to-tank propellant transfer line using the Generalized Fluid System Simulation Program (GFSSP). To compare with data from recently concluded turbulent LH2 chill down experiments, seven different cases were run across a range of inlet liquid temperatures and mass flow rates. Both trickle and pulse chill down methods were simulated. The GFSSP model qualitatively matches external skin mounted temperature readings, but large differences are shown between measured and predicted internal stream temperatures. Discrepancies are attributed to the simplified model correlation used to compute two-phase flow boiling heat transfer. Flow visualization from testing shows that the initial bottoming out of skin mounted sensors corresponds to annular flow, but that considerable time is required for the stream sensor to achieve steady state as the system moves through annular, churn, and bubbly flow. The GFSSP model does adequately well in tracking trends in the data but further work is needed to refine the two-phase flow modeling to better match observed test data.

Temperature Sensors↗

NASA NTRS2006

Effect of Thermomechanical Processing on the Microstructure, Properties, and Work Behavior of a Ti50.5 Ni29.5 Pt20 High-Temperature Shape Memory Alloy

TiNiPt shape memory alloys are particularly promising for use as solid state actuators in environments up to 300 °C, due to a reasonable balance of proper ties, including acceptable work output. However, one of the challenges to commercializing a viable high-temperature shape memory alloy (HTSMA) is to establish the appropriate primary and secondary processing techniques for fabrication of the material in a required product form such as rod and wire. Consequently, a Ti 50.5 Ni 29.5 Pt 20 alloy was processed using several techniques including single- pass high-temperature extrusion, multiple-pass high-temperature extrusion, and cold drawing to produce bar stock, thin rod, and fine wire, respectively. The effects of heat treatment on the hardness, grain size, room temperature tensile properties, and transformation temperatures of hot- and cold-worked material were examined. Basic tensile properties as a function of temperature and the strain-temperature response of the alloy under constant load, for the determination of work output, were also investigated for various forms of the Ti 50.5 Ni 29.5 Pt 20 alloy, including fine wire.

Grain Size↗

NASA NTRSDate not supplied

The Radsetta Stone – Establishing a Common Environment Taxonomy Across Space Nuclear Technologies

Qualitative descriptions, such as “high temperature” or “long exposure” vary substantially according to the context and can obfuscate messaging. A quantitative taxonomy for describing space nuclear environments - encompassing fission, radioisotope, and natural - has been established to enable clear communication across technology applications. Zones have been defined to partition environments with an intent to achieve a useful level of granularity. These zones enable a more detailed assessment than the Technology Readiness Level and can also be used to identify gaps and overlaps spanning space nuclear technologies. This framework has been named the Radsetta Stone – a paronomasia of the Rosetta Stone. There are two directions in which the Radsetta Stone can be used. The user may be a technology developer or a technology customer. The developer would use the Radsetta Stone to communicate the current state of the technology they are developing. The maximum zone for each parameter would then be used to state the current limit of their technology. The customer could use the Radsetta Stone to communicate the environmental conditions of their system. Zones would be associated with different locations within the system. The Radsetta Stone is intended to enable clear communication regarding space nuclear technology. The following legend and tables show the levels of the Radsetta Stone. These levels show success criteria gates in a similar manner as Technology Readiness Levels.

taxonomy↗

NASA NTRS2026

Developing a Pyrolysis Gas Thermal Blocking Model for Reentry Demise

In NASA’s Object Reentry Survival Analysis Tool (ORSAT), aerodynamic drag and aerothermal heating coefficients are computed for each of the free-molecular, continuum, and transitional flow regimes using analytical and semi-analytical methods. These heating coefficients were derived for typical metallic materials that melt and do not have a strong gas-phase contribution to the flow in the boundary layer. Modern satellites typically feature fiber-reinforced polymer (FRP) components, such as solar array booms, facesheets of sandwich panels, or overwraps for composite-overwrapped pressure vessels (COPV). These FRP materials do not behave the same as metals in the reentry environment, but instead will pyrolyze and develop significant volumes of gas into the boundary layer. Accurately predicting the reentry demise of FRP components is critical to assessing the reentry casualty risk for modern spacecraft. Research in recent years has shown that this demisability can depend heavily on how the expulsion of gaseous pyrolysis products through the outer surface of the material affects the heat flux at the surface. The ODPO has been developing a reduced-order model of the effect of pyrolysis gas blowing on the heat flux based on correlations between a blowing factor and a non-dimensional heat flux to be incorporated in the upcoming version 7.3 of the Object Reentry Survivability Analysis Tool (ORSAT). This presentation discusses the progress of this development project and the challenges remaining for generalizing the model across families of FRP materials.

Benton Greene↗

NASA NTRS1977

Calculations of the Moon's Thermal History at Different Concentrations of Radioactive Elements, Taking Into Account Differentiation on Melting

Calculations of the thermal history of the moon were carried out by solving the thermal conductivity equation for the case in which the heat sources are the long-lived radioactive elements Th, U, and K 40 . The concentrations of these elements were adjusted to give four variations of the heat flow: 1.35 x 10 -8 and 0.91 x 10 -6 cal cm -2 s -1 (I and I**), 0.61 x 10 -6 (variant II, the terrestrial mixture of Lyubimova and Starkova), and 0.236 x 10 -6 (variant III, the chondrite model of Urey and MacDonald). In the same calculations, we considered layering of the differentiated material with transport to the surface of the radioactive elements after the temperature of the layer rose to 200 K above the melting temperature, which is considered in five variants that differ in the amount of transported radioactive elements: 100 percent ( n = 1), 80 percent n = 0.8), 40 percent, 20 percent ( n = 0.6, 0.4, 0.2). During fusion the heat capacity and heat conductivity were changed. We considered two variants of an initially cold (273 K) and hot (900 K) Moon. Calculations show that the interior of the Moon was heated to melting during the first 0.7 to 2.3 x 10 9 years. The maximum fusion involved practically the entire Moon to a distance from 15 to 45 km beneath the surface, and started 3.5 to 4.0 x 10 9 years ago (I, I**), or 2.5 to 3.0 x 10 9 years ago (II, III) and continued for 1 to 2 x 10 9 years. Today the Moon is cooling. The current thickness of the solid crust is from 150 to 200 km and the heat flow exceeds the stationary value 1.5 fold. Apparently the most realistic variant is II (terrestrial mixture) for an initially hot Moon, and gives, regardless of the moderate concentration of radioactive elements, a heat flow of 0.9 to 0.95 x 10 -6 cal cm -2 s -1 , which agrees with the radioastronomical measurements of Troitsky and Krotikov and with the Apollo 15 data.

O I Ornatskaya↗

NASA NTRSDate not supplied

Thermal Design and Thermal Vacuum Testing of the StarBurst Instrument

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.

StarBurst↗

NASA NTRS1992

Nickel hydrogen common pressure vessel battery development

Our present design for a common pressure vessel (CPV) battery, a nickel hydrogen battery system to combine all of the cells into a common pressure vessel, uses an open disk which allows the cell to be set into a shallow cavity; subsequent cells are stacked on each other with the total number based on the battery voltage required. This approach not only eliminates the assembly error threat, but also more readily assures equal contact pressure to the heat fin between each cell, which further assures balanced heat transfer. These heat fin dishes with their appropriate cell stacks are held together with tie bars which in turn are connected to the pressure vessel weld rings at each end of the tube.

Jones, Kenneth R.↗

NASA NTRSDate not supplied

Aerothermodynamic Analysis of a Flexible Thermal Protection System under Reentry Loads

The Carryall Block 1 reentry vehicle being developed by Outpost Space utilizes a strut supported semirigid deployable heatshield. This consists of a flexible thermal protection system, a heat-resistant fabric stack, stretched over actuated spars. The advantages of a deployable heatshield include reduced heat loading and earlier deceleration in the trajectory. However, the nature of the flexible thermal protection system necessitates considering the loaded shape of the heat shield. The flexible thermal protection system will deflect under reentry loads leading to areas of higher heating rates as well as a reduced axial coefficient as compared to the nominal shape. The Carryall Block 1 is analyzed using NASA’s FUN3D and DPLR CFD solvers with a deflected shape based on the catenary equations. The aerodynamic results are found to be within a percent for both solvers and both structured and unstructured mesh types. Fluid Structure Interaction (FSI) analysis is currently a work in progress, using file I/O to communicate between FUN3D and LS-DYNA, a commercial nonlinear structural solver. Challenges in deforming the geometry, mesh, and initial results are presented here.

thermal protection system↗

NASA NTRS2002

Development Status of 3 Battery Systems for the X-38 Crew Return Vehicle

This viewgraph presentation gives an overview of the development status of three battery systems for the X-38 crew return vehicle. Details are given on the design features, the lithium battery module, PCM composite heat sinks, carbon fibercore blocks for Qual battery, battery module base housing, heat sink characteristics, and battery qualifications.

Darcy, Eric↗

NASA NTRS1995

Computation of three-dimensional mixed convective boundary layer flow

The paper presents the numerical solution of heat and mass transfer during cross-flow (orthogonal) mixed convection. In this class of flow, a buoyancy-driven transport in the vertical direction and a forced convective flow in the horizontal direction results in a three-dimensional boundary layer structure adjacent to the plate. The rates of heat and mass transfer are determined by a combined influence of the two transport processes. The equations for the conservation of mass, momentum, energy, and species concentration were solved along with appropriate boundary conditions to determine the distributions of velocity components, temperature, and concentration across the thickness of the boundary layer at different locations on the plate. Results were expressed in dimensionless form using Reynolds number, Richardson number for heat transfer, Richardson number for mass transfer, Prandtl number, and Schmidt number as parameters. It was found that the transport is dominated by buoyancy at smaller vertical locations and at larger distances away from the forced convection leading edge. Effects of forced convection appeared to be very strong at smaller horizontal distances from the leading edge. The cross stream forced convection enhanced the rate of heat and mass transfer by a very significant amount.

Prashandt Gadepalli↗

NASA NTRS1962

Mechanical Design of Spacecraft

In the spring of 1962, engineers from the Engineering Mechanics Division of the Jet Propulsion Laboratory gave a series of lectures on spacecraft design at the Engineering Design seminars conducted at the California Institute of Technology. Several of these lectures were subsequently given at Stanford University as part of the Space Technology seminar series sponsored by the Department of Aeronautics and Astronautics. Presented here are notes taken from these lectures. The lectures were conceived with the intent of providing the audience with a glimpse of the activities of a few mechanical engineers who are involved in designing, building, and testing spacecraft. Engineering courses generally consist of heavily idealized problems in order to allow the more efficient teaching of mathematical technique. Students, therefore, receive a somewhat limited exposure to actual engineering problems, which are typified by more unknowns than equations. For this reason it was considered valuable to demonstrate some of the problems faced by spacecraft designers, the processes used to arrive at solutions, and the interactions between the engineer and the remainder of the organization in which he is constrained to operate. These lecture notes are not so much a compilation of sophisticated techniques of analysis as they are a collection of examples of spacecraft hardware and associated problems. They will be of interest not so much to the experienced spacecraft designer as to those who wonder what part the mechanical engineer plays in an effort such as the exploration of space.

MARINER SPACE PROBE↗