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At least 1,387 records · Page 77
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Validation of Loci-Stream for Autogenous Pressurization of Cryogenic Propellant Tank

Autogenous pressurization of cryogenic propellant tanks eliminates the need to have an additional pressurant tank on the space vehicle, which is highly advantageous due to reduced vehicle mass and design complexity. Autogenous pressurization therefore is one of the key technologies for deep space exploration and long-term space missions. The complex interaction of thermal gradients, turbulence and phase change near the interface make the problem a challenging one to model. Nodal analysis tools and reduced order models are unable to capture the necessary physics. 3-D CFD analyses are necessary to fully characterize autogenous pressurization. CFD analyses pose their own difficulties. The requisite CFD tool to tackle this problem need to be modular with the ability to incorporate various physics models, efficient, and computationally scalable for simulating flight size tanks. NASA MSFC's Loci-Stream CFD tool along with the VOF module is a great candidate to fit this mold. We demonstrate our modeling approach and validation of Loci-Stream for predicting autogenous pressurization of a flight scale propellant tank in order for the solver to serve as a reliable design and analysis tool for NASA's CFM application needs. Liquid hydrogen tank pressurization tests carried out at the MSFC test stand 300 facilities provide reliable validation data for this purpose. These tests were modeled using the Loci-Stream solver with a newly implemented two-phase sharp interface treatment. We show that our modeling approach and CFD solver predict the autogenous pressurization phenomena satisfactorily, and document challenging aspects of modeling this problem.

cryogenic fluid management↗
NASA NTRSDate not supplied

Evaluating Liftoff Debris for NASA’s Space Launch System (SLS) Prior to the Artemis I Launch

The SLS Artemis I launch vehicle is the first of several planned Artemis launch vehicles, with a number of design differences from earlier NASA missions that incur liftoff debris risk to the mission. As a test vehicle, the Artemis I hardware also endured environments and tests not planned for future missions, which led to several additional factors contributing to an evolving liftoff debris risk to the SLS vehicle. This paper will summarize these risk factors and address the processes used to evaluate and communicate the risks to support a successful Artemis I launch. It will discuss how the evolving risks that were quantified and evaluated by a Cross-Program team of debris Subject Matter Experts to mitigate liftoff debris hazards and communicate updated risk to the SLS vehicle. This process was performed through the inaugural use of an SLS debris day-of-launch (DOL) standard operating procedure that will be used for subsequent Artemis missions. This paper addresses the risk of liftoff debris, debris released by the vehicle or from the launch pad during liftoff through vehicle tower clear. Expected liftoff debris is well understood from previous NASA programs’ experience and from tests of materials, processes and functions that are known to release liftoff debris. These expected sources were assessed and cleared well ahead of launch day. However, given the ever-changing schedules and environments, processes were in place to evaluate any additional potential liftoff debris risks identified during launch countdown. Although many of the Artemis vehicle hardware components are similar to those on the NASA Shuttle Program, there are important differences in the architecture of the Artemis I vehicle which require new assessments of liftoff debris risk for the Artemis missions. The more favorable Artemis crew module location and surfaces are far less vulnerable to debris impacts; however, the longer vehicle can result in higher liftoff debris impact energies to those components on the aft end of the vehicle. Additionally, the positional change of the RS-25 liquid engines to nearer the Booster nozzle exit plane along with the change in Booster throat plug design is a disadvantage to the overall liftoff debris risk which resulted in additional test and analysis efforts for evaluating the integrated vehicle debris risk. In spite of the comprehensive tests and analyses of Artemis I expected liftoff debris, a number of additional tests/processes were completed prior to the Artemis I mission that were required to support a complete understanding of a new launch vehicle, but increased the risk of releasing liftoff debris. The hardware endured several additional cryogenic loading cycles, including the Green Run tests at Stennis Space Center, Wet Dress Rehearsals at Kennedy Space Center, and multiple launch attempts. Each of these cycles induced stresses in the thermal protection system (TPS) materials, increasing the risk of damage to and release of the TPS. Additionally, induced and weather environmental factors that could increase the likelihood of debris release were significant. Vibrations and stresses in the TPS were induced by a required roll-back to the Vehicle Assembly Building before Hurricane Ian to protect the vehicle from damage by high winds. Wind damage and potential internal stresses to several outer mold line materials on the integrated SLS vehicle and mobile launcher were caused by weathering Hurricane Nicole at Pad 39B the week before launch. A thorough imagery scan of the vehicle was performed after each event and the damage observed was repaired, removed, or assessed and the risk to the mission evaluated. Mitigation of debris risk can occur by tests and analyses to show debris impacted components as damage tolerant, by new/improved processes for prevention of debris availability, or redesign. Risk mitigation processes for Artemis I-specific liftoff debris events and the development and use of the SLS debris day of launch (DOL) procedures that will be used for subsequent Artemis missions will be described.

Space Launch System↗
NASA NTRSDate not supplied

Validation of Loci-Stream for Autogenous Pressurization of Cryogenic Propellant Tank

Autogenous pressurization of cryogenic propellant tanks eliminates the need to have an additional pressurant tank on the space vehicle, which is highly advantageous due to reduced vehicle mass and design complexity. Autogenous pressurization therefore is one of the key technologies for deep space exploration and long-term space missions. The complex interaction of thermal gradients, turbulence and phase change near the interface make the problem a challenging one to model. Nodal analysis tools and reduced order models are unable to capture the necessary physics. 3-D CFD analyses are necessary to fully characterize autogenous pressurization. CFD analyses pose their own difficulties. The requisite CFD tool to tackle this problem need to be modular with the ability to incorporate various physics models, efficient, and computationally scalable for simulating flight size tanks. NASA MSFC's Loci-Stream CFD tool along with the VOF module is a great candidate to fit this mold. We demonstrate our modeling approach and validation of Loci-Stream for predicting autogenous pressurization of a flight scale propellant tank in order for the solver to serve as a reliable design and analysis tool for NASA's CFM application needs. Liquid hydrogen tank pressurization tests carried out at the MSFC test stand 300 facilities provide reliable validation data for this purpose. These tests were modeled using the Loci-Stream solver with a newly implemented two-phase sharp interface treatment. We show that our modeling approach and CFD solver predict the autogenous pressurization phenomena satisfactorily, and document challenging aspects of modeling this problem.

CFM↗
NASA NTRSDate not supplied

Strategies for Determining Safety of Fermented Foods Produced in Space

The BioNutrients experiments examine the feasibility of using microbially produced foods – yogurt, kefir, fermented food products– to supplement the current prepackaged food system to provide freshly synthesized nutrients and genetically engineered protein therapeutics to support crew health for sustainable and long duration crewed exploration. Safety of the products must be established as direct consumption of such products is a risk due to potential contamination with pathogenic organisms. Identifying pathogen contamination is particularly challenging as the fermenting cultures can mask the presence of contaminating organisms. Here we consider current methods of detecting coliforms, aerobic colonies, non-lactic acid bacteria, molds (yeasts), and specific pathogens including Staphylococcus aureus and salmonella. We also consider alternate and additional testing including sequencing and quantitative PCR. We investigate the use of the E-Nose developed by NASA Ames Research Center. This portable device senses volatile organic compounds in real time and can be trained to recognize “good” uncontaminated products from contaminated ones as a means of predicting food safety. Finally, we consider methods of in-flight pasteurization to increase safety and decrease overall microbial load of these products. This presentation will summarize testing status and outline the strategy that will be employed in the BioNutrients-3 ISS fermented foods experiment.

Food↗
NASA NTRSDate not supplied

Influence of Feed Rate on Microstructure and Hardness of Conventionally Spun-formed Al 6061-O Plate

The effect of feed rate on the hardness and microstructure of a 7.35-mm-thick section from a spin-formed Al 6061 cylinder is investigated via microhardness mapping and electron backscatter diffraction (EBSD) analysis. The through-thickness hardness and microtexture profiles, corresponding to a true thickness strain of ≈ 0.3, are utilized to evaluate microstructural variations that could influence mechanical anisotropy and spin formability. The microhardness data reveal distinct gradients from the outer to the inner mold line that vary with feed rate and may be a function of the level of forming strain in the material. The microtextural mapping data expose a minor influence from feed rate, but contain trends through the material cross-section that suggest variations in the location-dependent stress states. The results demonstrate for the first time that the relationship between spin forming deformation and hardness or texture variations can be exploited to optimize processing parameters and mechanical response.

Spin forming↗
NASA NTRSDate not supplied

Analysis of a Landing System for Planetary Payloads Utilizing Passive Energy Absorbing Composite Structure

Delivery of a payload from space to a planetary surface currently requires the development of an application specific landing system to protect the payload from forces imparted during impact with the planet surface. Often, active energy attenuating systems such as retro-rockets, deployable parachutes, and airbags are utilized within these landing systems to reduce landing impact energy. Unfortunately, these active systems come at a cost; active energy attenuating systems are susceptible to system faults which may limit or completely negate their energy attenuating capability. Additionally, components needing to be stowed such as fuel, parachutes, and airbags increase design complexity, cost, and weight. To overcome these limitations, this study examines the potential of passive energy attenuation through energy absorbing structural design and composite materials to mitigate landing loads for small payload planetary delivery. Researchers at the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) have conducted extensive research into developing energy absorbing structures and components for the attenuation of impact energy under various loading conditions including aircraft crash and spacecraft impact. The current study leverages this research to design a lightweight planetary delivery system which utilizes unique outer mold line (OML) geometry and passive energy absorbing structural design to limit landing loads across potential planetary surface environments. The OML geometry is designed to control impact orientation and provide self-righting capabilities for slopped impact surfaces. The internal structure is composed of composite material structures arranged to provide energy absorption which is robust to impact angle and impact velocity. The developed planetary delivery design concept will be evaluated using finite element (FE) model analysis. Simulations of landing impacts with representative soil surface environments will be used to characterize the energy absorbing capabilities of the landing system. Sensitivity of predicted impact force to landing environment, impact angle, and impact velocity will be assessed to identify capabilities and limitations of the initial structural design. Results will be used to determine the feasibility of a lightweight composite structure to passively absorb landing energy for robust planetary payload delivery.

Crashworthiness↗
NASA NTRSDate not supplied

NASA's X-57 High Lift Motor Controller: Detailed Design, Test Results, and Outcomes

NASA's X-57 all-electric aircraft was a research project aimed at investigating lightweight and efficient electric propulsion components. The general approach was to utilize a distributed electric propulsion (DEP) design. An essential component of this design was the High Lift Motor Controller (HLMC), a motor drive which provided power to the High Lift Motors (HLMs) and High Lift Propellers (HLPs) responsible for providing additional thrust for take-off and landing. This paper presents the detailed design, test results, and outcomes from the development of the HLMC, a 14 kW, 1kg, 98.3% efficient, outer mold line (OML) cooled, silicon carbide (SiC) MOSFET-based inverter and controller.

Electric Aircraft↗
NASA NTRSDate not supplied

Validation of Two Phase Flow Modeling Techniques in Loci-Stream using Axial Jet Pressure Control and Autogenous Pressurization Experiment

Autogenous pressurization of cryogenic propellant tanks eliminates the need to have an additional pressurant tank on the space vehicle, which is highly advantageous due to reduced vehicle mass and design complexity. Autogenous pressurization therefore is one of the key technologies for deep space exploration and long-term space missions. The complex interaction of thermal gradients, turbulence and phase change near the interface make the problem a challenging one to model. Nodal analysis tools and reduced order models are unable to capture the necessary physics. 3-D CFD analyses are necessary to fully characterize autogenous pressurization. CFD analyses pose their own difficulties. The requisite CFD tool to tackle this problem need to be modular with the ability to incorporate various physics models, efficient, and computationally scalable for simulating flight size tanks. NASA MSFC's Loci-Stream CFD tool along with the VOF module is a great candidate to fit this mold. We demonstrate our modeling approach and validation of Loci-Stream for predicting autogenous pressurization of a flight scale propellant tank in order for the solver to serve as a reliable design and analysis tool for NASA's CFM application needs. Liquid hydrogen tank pressurization tests carried out at the MSFC test stand 300 facilities provide reliable validation data for this purpose. These tests were modeled using the Loci-Stream solver with a newly implemented two-phase sharp interface treatment. We show that our modeling approach and CFD solver predict the autogenous pressurization phenomena satisfactorily, and document challenging aspects of modeling this problem.

cryogenic fluid management↗
NASA NTRSDate not supplied

Influence of Feed Rate on Microstructure and Hardness of Conventionally Spun-formed Al 6061-O Plate

The effect of feed rate on the mechanical properties and microstructure of a 7.35-mm-thick section from a spin-formed Al 6061 cylinder is investigated via microhardness mapping and electron backscatter diffraction (EBSD) analysis. The through-thickness hardness and microtexture profiles, corresponding to a true thickness strain of ≈ 0.3, are utilized to evaluate microstructural variations that could influence mechanical anisotropy and spin formability. The microhardness data reveal distinct gradients from the outer to the inner mold line that vary with feed rate and may be a function of the level of forming strain in the material. The microtextural mapping data expose a minor influence from feed rate, but contain trends through the material cross-section that suggest variations in the location-dependent stress states. The results demonstrate for the first time that the relationship between spin forming deformation and hardness or texture variations can be exploited to optimize processing parameters and mechanical response.

Spin forming↗
NASA NTRSDate not supplied

A Preliminary Assessment of Step Effects on the BOLT Geometry in Mach 6 Flow

An infrared thermography measurement technique has been installed in the Langley Aerothermodynamic Laboratory. The system is calibrated for use with commercially available thermoplastics and ceramic materials, and to the surface temperature ranges when exposed to the facility hypersonic flow. Herein an uncertainty quantification is performed, revealing 5% surface temperature and 20% heat transfer uncertainty bands to two-sigma confidence. As a first application of the test measurement technique, an investigation was conducted to explore the correlation between the height of forward facing (FF) and rearward facing (RF) two-dimensional steps, free stream Reynolds number, and the onset of boundary layer transition from laminar to turbulent in Mach 6 flow on a 50% scale Boundary Layer Transition (BOLT) flight test vehicle geometry. Computational predictions were generated to identify potential x/L locations of transition onset for various flow conditions with a smooth outer mold line geometry and are compared to experimental results. The BOLT Step Test was conducted in the NASA Langley Aerothermodynamic Laboratory 20-inch Mach 6 wind tunnel at Reynolds numbers between 1.4 and 11.3 million conditions for various step heights. Data were collected using a developmental infrared thermography test technique and reduced to heat transfer for comparison with CFD solutions. This is a preliminary assessment of the developmental infrared thermography test technique used to acquire and reduce the experimental data. insight into the ongoing comparative analysis of the effects of FF and RF steps on boundary layer transition for the BOLT geometry in Mach 6 flow.

BOLT↗
NASA NTRS2024

Pull-off Behavior of Stitched Composite T-Joints

T-joints are key structural elements that connect opposing surfaces, thereby providing the load path between flat or curved panels (i.e., upper and lower wing skins) and transverse components (i.e., stiffeners). Due to the low interlaminar strength of polymer matrix composites and their geometrical discontinuities, these joints are vulnerable to pull-off loads. To address these issues, through-thickness reinforcements can be employed to enhance the interlaminar capability of these type of joints. In this study, T-joints were manufactured using through-thickness stitching in dry carbon preforms and cured using the vacuum-assisted resin transfer molding (VARTM) process. Stitched and unstitched T-joints were tested under pull-off loading conditions, and surface strain fields were obtained using a 3D digital image correlation system. The ultimate load, displacement, and absorbed energy of the stitched T-joints were greater than their unstitched counterparts by approximately 16%, 34%, and 58%, respectively. Failure mechanisms were identified by examining fracture surfaces using optical microscopy. Results demonstrate that through-thickness stitching significantly improves the damage tolerance of T-joints, which highlights the effectiveness of stitching to enhance the structural integrity of large aerospace components.

VARTM↗
NASA NTRS2024

NASA's X-57 High Lift Motor Controller: Detailed Design, Test Results, and Outcomes

NASA's X-57 all-electric aircraft was a research project aimed at investigating lightweight and efficient electric propulsion components. The general approach was to utilize a distributed electric propulsion (DEP) design. An essential component of this design was the High Lift Motor Controller (HLMC), a motor drive which provided power to the High Lift Motors (HLMs) and High Lift Propellers (HLPs) responsible for providing additional thrust for take-off and landing. This paper presents the detailed design, test results, and outcomes from the development of the HLMC, a 14 kW, 1kg, 98.3% efficient, outer mold line (OML) cooled, silicon carbide (SiC) MOSFET-based inverter and controller.

Electric Aircraft↗
NASA NTRSDate not supplied

A Preliminary Assessment of Step Effects on the BOLT Geometry in Mach 6 Flow

An infrared thermography measurement technique has been installed in the Langley Aerothermodynamic Laboratory. The system is calibrated for use with commercially available thermoplastics and ceramic materials, and to the surface temperature ranges when exposed to the facility hypersonic flow. Herein an uncertainty quantification is performed, revealing 5% surface temperature and 20% heat transfer uncertainty bands to two-sigma confidence. As a first application of the test measurement technique, an investigation was conducted to explore the correlation between the height of forward facing (FF) and rearward facing (RF) two-dimensional steps, free stream Reynolds number, and the onset of boundary layer transition from laminar to turbulent in Mach 6 flow on a 50% scale Boundary Layer Transition (BOLT) flight test vehicle geometry. Computational predictions were generated to identify potential x/L locations of transition onset for various flow conditions with a smooth outer mold line geometry and are compared to experimental results. The BOLT Step Test was conducted in the NASA Langley Aerothermodynamic Laboratory 20-inch Mach 6 wind tunnel at Reynolds numbers between 1.4 and 11.3 million conditions for various step heights. Data were collected using a developmental infrared thermography test technique and reduced to heat transfer for comparison with CFD solutions. This is a preliminary assessment of the developmental infrared thermography test technique used to acquire and reduce the experimental data. insight into the ongoing comparative analysis of the effects of FF and RF steps on boundary layer transition for the BOLT geometry in Mach 6 flow.

BOLT↗
NASA NTRS2024

Capsule_Grid User Guide

Software utility CAPSULE_GRID and several related procedures that automate grid generation for computational fluid dynamics calculations about axisymmetric atmospheric entry vehicles are described as a supplement to the documentation within the source codes. This application program and its ancillary software utilities are all in the public domain at the web site https://software.nasa.gov/software/category/all/arc/1/cfdtools with the exception of the 2D hyperbolic volume gridding step for which scripts driving either GRIDGEN1 or HYPGEN2 are available. CAPSULE_GRID was developed at NASA Ames Research Center in support of the DPLR3 real gas flow solver. Both forebodies (only) and full bodies are treated, for axisymmetric 2D flow calculations or for 3D calculations at angle of attack. The expected application is to the smooth outer mold line of a hypersonic atmospheric entry vehicle, for which a single layer of point-matched structured grid blocks normally suffices. Control files for a variety of representative cases are included with illustrations of results, and some insights into the subtleties of curvature-based grid point distribution are provided. Gridding of a capsule on a sting is supported. Additional procedures for treating arbitrarily long 2D wakes, prompted by asteroid studies and relevant to spacecraft such as Stardust, are also documented. Most recently, treating an asymmetric forebody (including an off-center nose, with or without an aft body) has been incorporated as an extra step.

spacecraft↗
NASA NTRSDate not supplied

Carbon Fiber Composite Processing Using Isothermal Resins

To meet future demand for single-aisle composite aircraft production, manufacturing rates are expected to increase by up to six times current production rates. Resin infusion is a manufacturing method that has potential to enable fabrication of structures at the desired rates. However, production cycle times with currently available commercial infusion resins cannot support the rates needed for an estimated 80 aircraft per month. Recently, rapid-curing isothermal resins were developed that can reduce the overall processing time. These resins can be infused, cured, and demolded at a single temperature below 100 °C, thereby eliminating the need for time consuming temperature ramps and shortening the overall cure time. In addition, the low temperatures enable use of low-cost tooling during production. This presentation will discuss the use of resin transfer molding to produce composites with rapid-curing isothermal resins, manufacturing considerations, and performance characteristics of the composites.

John M Gardner↗
NASA NTRS2026

Evaluating Energy Absorption Methods for Integrated Composite Seat Designs

Composite materials have become ubiquitous in the aerospace industry due to their exceptionally light weight and high strength characteristics, as well as their unique ability to be engineered and tailored to meet specific loading conditions and performance requirements. These advanced materials offer superior strength-to-weight ratios compared to traditional metallic materials, making them particularly valuable in weight-critical aerospace applications where every pound saved translates to improved efficiency and performance. In currently operating fleets of commercial and military aircraft, composite materials have been successfully applied to critical structural components, including primary load-bearing elements such as the fuselage sections and flooring structures, which must withstand significant in-flight loads and provide passenger safety. Additionally, these materials have been specifically tailored and optimized for aerodynamic components such as wings and tail assemblies, where their ability to be molded into complex shapes while maintaining structural integrity is particularly advantageous. The application of composite materials extends beyond primary structural elements into the realm of internal cabin components, most notably in innovative seat designs where weight reduction and structural integration are paramount concerns. Modern composite seat structures can be designed to integrate multiple functions, including structural support, comfort features, and safety systems, all while maintaining the lightweight characteristics essential for aircraft performance.

Digital image correlation↗
NASA NTRSDate not supplied

Thermo-Poro-Mechanical Modeling of RTV Intumescence

Room temperature vulcanizing (RTV) silicone is a high-temperature adhesive used as a gap-filler between heatshield tiles in numerous entry missions. Its propensity to intumesce, or swell upon exposure to heat, is a well-known effect that needs to be carefully quantified during design. At tile interfaces of charring ablators, intumescence, combined with differential recession, could cause the gap filler to protrude past the ablator outer mold line, forming a “fence”. Fencing can in turn cause transition to turbulence of the flow wetting the heat shield, leading to augmented surface heating. Recent experiments conducted at the Plasmatron X facility, the high enthalpy wind tunnel of the Center for Hypersonics and Entry Systems Studies, have shown prominent fencing of RTV gap fillers in PICA, under both nitrogen and air plasmas. Similar observations are well known in the arcjet literature. Further experiments under controlled environment, performed using in situ X-ray micro-computed tomography (micro-CT) at the Advanced Light Source (ALS), have shown heating rate-dependent swelling and shrinkage of RTV during pyrolysis. To simulate RTV intumescence, a novel model was introduced in the Porous Materials Analysis Toolbox based on OpenFOAM, PATO, to account for pore-pressure buildup within both closed- and open-pores. The governing equation for the thermo-poro-mechanical response were developed, assuming linear elasticity for the charring silicone. A new multi-pyrolysis model that tracks non-monotonic advancement of material properties with pyrolysis was proposed. This model addresses the limitations of state-of-the-art ablator models to capture the different stages of thermal degradation and coupled thermomechanics. Swelling of RTV was simulated using the new thermo-poro-mechanical model and compared against in situ micro-CT data. Results showed good agreement in intumescence height and temperature profiles at all heating rates, indicating that the key factor contributing to RTV swelling is the internal pressure build-up within closed- and open-pores. As RTV is cured into a soft (rubbery) compound with low-porosity and permeability, initial temperature increase and pyrolysis gas production cause a significant increase of internal pressure, causing a pronounced volume growth. As thermal degradation progresses, rigidization of the silicone occurs due to char hardening which counteract volume shrinkage after gas pressure relief. Overall, our model shows that accounting for changes in properties such porosity, permeability and key thermomechanical coefficients is crucial for capturing the RTV volume change during ablation and enable a predictive capability for heatshield tile interface response. A plan for future calibration of thermomechanical properties evolution during degradation is discussed, as a key next step to close the new model.

RTV↗
NASA NTRSDate not supplied

Thermo-Poro-Mechanical Modeling of RTV Intumescence

Room temperature vulcanizing (RTV) silicone is a high-temperature adhesive used as a gap-filler between heatshield tiles in numerous entry missions. Its propensity to intumesce, or swell upon exposure to heat, is a well-known effect that needs to be carefully quantified during design. At tile interfaces of charring ablators, intumescence, combined with differential recession, could cause the gap filler to protrude past the ablator outer mold line, forming a “fence”. Fencing can in turn cause transition to turbulence of the flow wetting the heat shield, leading to augmented surface heating. Recent experiments conducted at the Plasmatron X facility, the high enthalpy wind tunnel of the Center for Hypersonics and Entry Systems Studies, have shown prominent fencing of RTV gap fillers in PICA, under both nitrogen and air plasmas. Similar observations are well known in the arcjet literature. Further experiments under controlled environment, performed using in situ X-ray micro-computed tomography (micro-CT) at the Advanced Light Source (ALS), have shown heating rate-dependent swelling and shrinkage of RTV during pyrolysis. To simulate RTV intumescence, a novel model was introduced in the Porous Materials Analysis Toolbox based on OpenFOAM, PATO, to account for pore-pressure buildup within both closed- and open-pores. The governing equation for the thermo-poro-mechanical response were developed, assuming linear elasticity for the charring silicone. A new multi-pyrolysis model that tracks non-monotonic advancement of material properties with pyrolysis was proposed. This model addresses the limitations of state-of-the-art ablator models to capture the different stages of thermal degradation and coupled thermomechanics. Swelling of RTV was simulated using the new thermo-poro-mechanical model and compared against in situ micro-CT data. Results showed good agreement in intumescence height and temperature profiles at all heating rates, indicating that the key factor contributing to RTV swelling is the internal pressure build-up within closed- and open-pores. As RTV is cured into a soft (rubbery) compound with low-porosity and permeability, initial temperature increase and pyrolysis gas production cause a significant increase of internal pressure, causing a pronounced volume growth. As thermal degradation progresses, rigidization of the silicone occurs due to char hardening which counteract volume shrinkage after gas pressure relief. Overall, our model shows that accounting for changes in properties such porosity, permeability and key thermomechanical coefficients is crucial for capturing the RTV volume change during ablation and enable a predictive capability for heatshield tile interface response. A plan for future calibration of thermomechanical properties evolution during degradation is discussed, as a key next step to close the new model.

silicone intumescence↗