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Smart book charts for TPSM: Heatshield for Extreme Entry Environment Technology (HEEET), Conformal Ablative TPS (CA-TPS), 3D Woven Multifunctional Ablative TPS (3D MAT), and Adaptable, Deployable, Entry and Placement Technology (ADEPT).
This presentation introduces a new sizing and margin methodology for dual-layer Thermal Protection Systems (TPS). The methodology has been tailored for application to a dual-layer 3D-woven TPS called Heat-shield for Extreme Entry Environments Technology (HEEET). Sizing is performed for a reference Saturn probe mission to show how uncertainties in trajectory, aerothermal modelling and TPS response impact the sizing of each layer.
Exploration of ballistic coefficient and entry flight path angle to determine loom requirements for weaving a dual-layer thermal protection material such as HEEET (Heatshield for Extreme Entry Environment Technology).
The Common Probe Study was funded by the NASA's Planetary Science Division in the Science Mission Directorate in 2018 to investigate the feasibility of a common aeroshell design for atmospheric probe missions at Venus, Jupiter, Saturn, Uranus, and Neptune. The study involved 4 NASA Centers: Ames Research Center, Goddard Space Flight Center, Langley Research Center, and the Jet Propulsion Laboratory. The common aeroshell design that was studied was a 400 kg, 1.5 m diameter, 45-degree sphere cone shape with a high density heatshield material (Heatshield for Extreme Entry Environments Technology, or HEEET) and a parachute system to extract the descent vehicle. This size of aeroshell could accommodate a descent vehicle of 0.75 m diameter, which could encompass both Tier 1 and Tier 2 science instruments at each of the 5 destinations. Study methodology: First, a notional payload of instruments for each destination was defined based on the top priority measurements indicated by the Planetary Science Decadal Survey. Steep and shallow entry flight path angles (EFPA) were defined for each planet based on qualification and operational g-load limits for current, state-of-the-art instruments. Interplanetary trajectories were then identified that bounded the EFPA range.Next, 3-DoF simulations for entry trajectories were run using the entry state vectors from the interplanetary trajectories. Conical ribbon parachutes were sized based on heatshield separation dynamics. Aero-heating correlations were used to generate stagnation point convective and radiative heat flux profiles. High fidelity thermal response models for various TPS materials were used to size stagnation point thicknesses, with margins based on previous studies. Backshell TPS masses were assumed based on scaled heat fluxes from the heatshield and also from previous mission concepts.Based on these analyses, we have found that the common design is applicable for atmospheric probe missions for 4 out of the 5 destinations. Because of the unique gravity well for Jupiter, the entry environments are more severe resulting in heat loads an order of magnitude higher than for the other destinations.The next step is to determine what follow-on activities NASA should engage in. A questionnaire for the atmospheric probe community has been developed, with a focus on what size of aeroshell should be further analyzed (smaller or same diameter), and what incentives would make using such an aeroshell, if assembled and available, desirable to mission proposers.Preliminary results from this questionnaire will be presented.
The Mars Sample Return (MSR) Earth Entry Vehicle (EEV) is currently planned on being released from its Micro-Meteorite/Orbital Debri (MM/OD) shielding housing about two days before the Earth entry phase. This leaves the EEV exposed to incoming MM/OD impacts, potentially damaging the heat shield and compromising its Entry, Decent, and Landing (EDL) integrity. Currently, two materials are proposed to comprise the MSR-EEV heat shield, a dual layer material Heat-shield for Extreme Entry Environment Technology (HEEET) and Phenolic Impregnated Carbon Ablator (PICA). PICA has been well characterized for OD class impacts, ~7km/s high mass impacts, from previous testing done in the Orion program, but hasn't under-gone extensive MM impact testing. HEEET is a rela-tively new material with minimal prior testing in re-gards to High Velocity Impacts (HVI). In order to in-form selectability of a material, it is crucial to under-stand the material performance when faced with an HVI, directly affecting mission success probability.The current measure of a Thermal Protection Sys-tem's (TPS) performance against an HVI is evaluating a thermally sized material against its derived Balistic Limit Equation (BLE). A BLE is generated empirically from multiple shots of HVI testing, and is used as a first order method in evaluating TPS's performance against the expected MM/OD environment. This method has proved useful for previous uniform densi-ty TPS materials, but has never been validated against a dual layer recession material such as HEEET. Testing in the MSR program for FY19 has a re-quirement to assess the effects of a dual layer TPS by testing various thicknesses of HEEET's Recession Layer, seen in Figure 1. From this data, BLE's will be derived for the individual thickness ratio samples, as well as the material as a whole to evaluate if a heritage form of the BLE can capture the complex physics associated with a dual layer system. Crater morpholo-gy will also be assessed with post-test Non-Destructive Evaluation (NDE) methods such as CT scanning to visualize if a BLE can well predict the associated pene-tration depths, since a BLE assumes full disinigration of the impacting particle and is generally only used to size a spherical crater ? disregarding any shrapenel effects from a high density impactor.To inform this analysis, expected MM environ-ments from the Meteoroid Engineering Model (MEM) and analytical equations for mass flux of incoming MM were evaluated against the notional MSR-EEV trajectory [1]. Using those dispersions, a monte-carlo was run to determine the most probable particle pa-rameters, as well as the riskiest in terms of full bondline penetration. From these probabilities, a test matrix was designed to test against bounding cases for the various parameters of the BLE: projectile density, projectile mass, projectile velocity, and the impact angle.This presentation will discuss the performance of the dual layer TPS material HEEET against a wide range of impact kinetic energies and densitites, as well as the comparison of HEEET to PICA in terms of MM/OD performance and selectability criteria.
Thermal Protection System Materials & thermal control coatings were characterized by a variety of methodologies including heat flow meter method (HFM), dynamic scanning calorimetry (DSC), fourier transform infrared spectroscopy (FTIR), and Ultra-Violet & Visible light spectroscopy (UV-VIS). Physical properties were measured to aid in the construction of thermal response models for a number a projects including MEDLI and MEDLI2. Comparing pre-flight predictions with actual flight data can allow for a reduction of margins and improve designs. The models enable design of TPS for flight missions and analysis of flight data to understand aerothermal environments experienced on the mission. This presentation reveals critical information: heat capacity (cp vs T), thermal conductivity ( vs T), and emissivity (E vs T) for Super Lightweight Ablative (SLA), Phenolic Impregnated Carbon Ablative (PICA-D), & Heatshield for Extreme Entry Environment Technology (HEEET). These properties were also measured for various thermal control coatings manufactured by AZ Technology. A portion of this work is incomplete and highlights questions to investigate in the future.
This presentation is an overview of Heatshield for Extreme Entry Environment Technology (HEEET) providing the motivation, implementation (2014-2019), documentation, final assessment, and mission infusion.
No abstract available
Thermal Protection System (TPS) modeling requires accurate representation and prediction of the thermomechanical behavior of ablative materials. State-of-the-art TPS materials such as Phenolic Impregnated Carbon Ablator (PICA) have a proven flight record and demonstrate exceptional capabilities for handling extreme aerothermal heating conditions. The constant push for lightweight materials that are flexible in their design and performance, and hence allow for a wide range of mission profiles, has led NASA over the past years to develop its Heatshield for Extreme Entry Environment Technology (HEEET). HEEET is based primarily on a dual layer woven carbon fiber architecture and the technology has successfully been tested in arc-jet facilities. These recent developments have sparked interest in the accurate micro-scale modeling of composite weave architectures, to predict the structural response of macro-scale heatshields upon atmospheric entry. This effort can be extended to incorporate in-depth failure mechanics analyses as a result of local thermal gradients or high-velocity particle impact.
The effect of microcracking in the phenolic matrix of a three-dimensional woven thermal protection system (TPS) and the resulting material stiffness reduction was studied via a comparison of finite element analysis results from a linear analysis and an iterative linear analysis. A TPS is necessary to protect space vehicles from the aerodynamic heating of planetary entry. The Heatshield for Extreme Entry Environment Technology (HEEET) project has developed a TPS for use in high heat-flux and pressure missions. The material is a dual-layer continuous dry weave, which is then infiltrated with a low-density phenolic resin matrix to form a rigid ablator. The phenolic resin matrix does not have structural load transfer requirements, and testing has shown that the phenolic resin can fully satisfy thermal requirements when the matrix contains microcracks. Due to high stresses in the through-the-thickness direction of the material, phenolic microcracks may form in the matrix material, which would result in a reduction of stiffness. An exploratory study was conducted to determine if reduction in material stiffness would change the load paths and/or decrease the structural margins. A comparison was performed between a linear finite element analysis that did not take into account phenolic microcracking and an iterative linear finite element analysis that accounted for propagation of phenolic microcracking. Four subcases using varying assumptions were analyzed and the results indicate that the assumed strength at which the phenolic microcracking propagates was the critical parameter for determining the extent of microcracking in the phenolic matrix. Phenolic microcracking does not have an adverse effect on the structural response of the test article and is not a critical failure.
Ice Giant Missions launched in the next decade can have significantly reduced trip times by incorporating aerocapture as the primary method for orbit insertion instead of conventional propulsive techniques. Furthermore, a mass-efficient aerocapture system for a Neptune mission also enables the delivery of an orbiter, an atmospheric probe for in situ measurements to anchor global data collected by the orbiter, and possibly a lander at Triton. Aerocapture requires a Thermal Protection System (TPS) for the atmospheric entry, and it needs to be robust and mass efficient due to the size of the aeroshell required and the large heat load. Currently there are capable ablative thermal protection materials available, e.g., HEEET (Heatshield for Extreme Entry Environments Technology) and PICA (Phenolic-Impregnated Carbon Ablator) family. In this white paper, we show that the heat-shield TPS mass fraction can be as low as 5% to as high as 20% depending on total trip time reduction and other parameters. Further mass savings in the thermal protection system can be achieved with the use of conformal PICA, either standalone or in combination with HEEET with detailed design. We advocate the mission designers to consider aerocapture for Ice Giant missions so the science community can confidently consider missions to the Ice Giants to achieve unparalleled science in the coming decade.
Experiments were conducted in the NASA Ames Hypervelocity Free Flight Aerodynamic Facility (ballistic range) to quantify the effects on turbulent convective heat transfer of distributed surface roughness, and of isolated features, representative of thermal protection systems on atmospheric entry vehicles. The surface textures and features were applied on the conic frusta of 45o sphere-cone models having a nose-to-base radius ratio of 0.5, similar to the forebody geometry of the Galileo and Pioneer-Venus entry probes. Test conditions were selected to provide turbulent roughness Reynolds numbers, k+, in the ranges expected for outer planet entry missions. Turbulent flow on the conic frustum was achieved by tripping the flow on the sphere-segment nose cap with distributed surface roughness, created by sand-blasting the nose cap. Surface textures included distributed, acreage, roughness, as well as cavity and groove discrete features. The majority of the data to be presented are results for distributed roughness, which includes both random, sand-grain-like roughness, and regular pattern roughness. The pattern roughness was designed to represent the roughness on woven thermal protection system materials, such as NASA’s 3-D Medium Density Carbon Phenolic (3MDCP), also known as HEEET, developed by the Heatshield for Extreme Entry Environments Technology project. The pattern roughness tested is a 3-D wavy surface, and includes three different roughness element height-to-spacing ratios representative of two configurations of 3MDCP, and spanning ratios measured both before, and after, ablation in an arc jet test facility. The patterns tested in the ballistic range were laser-etched on metal models, and represented an idealized version of the real-world materials in which each roughness element was nearly identical. Additional tests were performed wherein the laser-etched patterns were degraded by sand-blasting with various sized grit media, to produce regular patterns with superimposed irregular roughness, more representative of flight materials. Results of each will be compared. The discrete features tested included cylindrical cavities and rectangular grooves of various width-to-depth ratios. Cavities represent either heatshield damage, such as from micro-meteoroid and orbital debris (MMOD) damage, or from designed penetrations, such as on the Genesis sample return capsule. The grooves were scaled representations of seams between segments of HEEET material in a notional tiled thermal protection system. The tests examined the effects on turbulent heating downstream of the isolated features. The tests were conducted at speeds between 3 km/s and 4 km/s in air between 0.15 atm and 0.25 atm (Mach numbers between 9 and 12). Roughness Reynolds numbers, k+, ranged from 12 to 70 for the sand roughness, and as high as 200 for the pattern roughness. Boundary-layer parameters required for calculating k+ were evaluated using computational fluid dynamics simulations using the DPLR (Data Parallel Line Relaxation) code. Each model included both rough- and smooth-wall segments, and heat transfer augmentation factors were determined as the ratio of the rough-wall to smooth-wall heat flux measured on each test.
Introduction: The three Venus missions that were recently selected for upcoming flight (VERITAS, DAVINCI+, and EnVision) will be incredibly valuable to our understanding of Venus’ history, geology, and atmosphere. However, even once completed, key gaps in our knowledge of Venus, and more generally the formation and active processes on rocky, Earth-like planets, will still persist. Remaining questions include 1) how global intrinsic magnetic fields might be maintained on rocky worlds, and how they could then go extinct, and 2) what role atmospheric sulfur chemistry plays in climates of Earth-like planets, which is an increasingly timely subject as Earth’s own atmospheric sulfur content is climbing due to human activity. These questions require in-situ observations from Venus’ cloud deck, at the altitudes at which the UV absorber exists. The Venus Environment Research and Novel Exploration (VERNE) mission will address these questions with an aerial platform that will drift around the equatorial region of the planet for 9 days. VERNE will collect data to determine the identity of the mysterious UV absorber, while also taking magnetic field measurements over the tesserae, the regions on Venus that are most likely to retain remanent crustal magnetization, in order to understand the potential role of a past intrinsically-generated global magnetic field on Venus. Mission Objectives: The two major science objectives that drive the VERNE mission are: 1) Determine the identity of the Venusian unknown ultraviolet absorber(s). First observed approximately a century ago [1], the composition of Venus’ ultraviolet (UV) absorber is one of the oldest mysteries in Venus atmospheric chemistry [2,3]. While several candidate UV absorbers (mostly sulfur species) have been proposed, no consensus has been reached on its composition and its specific interactions with the atmosphere. Determining the identity of the UV absorber will aid climate models by showing how and where incident solar energy is absorbed by the atmosphere and make progress towards understanding the chemical and energetic processes taking place above Venus’ upper cloud deck [4]. 2) Determine if Venus retains evidence of a past, internally-generated magnetic field. While Venus does not currently have an intrinsically-generated magnetic field, evidence for the existence of a past field on Venus and a timeline of its decay will fill in a more holistic picture of the evolution of Venus’ geological record and atmosphere. As the oldest geologic units on the surface, Venus’ tesserae may still have remanent crustal magnetization signatures within the rocky composition [5]. In any case, the signatures detected will provide insight into Venus’ past geological and core dynamo activity. Mission Summary: VERNE includes a 3-part flight system made up of 1) an entry system with a HEEET (Heatshield for Extreme Entry Environment Technology) aeroshell, 2) an orbiter for data relay, and 3) an in-situ balloon and gondola. After entry into Venus’ atmosphere, the balloon will be deployed within the upper cloud deck, at an altitude of 62 km, over a tesserae region. The in-situ data collection will last for the duration of 2 full circumnavigations of the planet, which will take ~9 days. Instrumentation Suite: The four instruments that comprise VERNE’s instrument payload will enable the identification of the unknown UV absorber and the detection of remanent crustal magnetization if it exists in the tesserae. The proposed instrument suite cycle during mission operations is shown in Fig. 1. The four instruments are described below: 1) Adams (ion neutral mass spectrometer) has a range of 18-257 AMU and will detect and distinguish the mixing ratios of O₂, H₂O, H₂SO₄, S, S₂, S₃, S₄, S₅, S₆, S₇, S₈, SO, SO₂, OSSO, SO₃, Cl₂, FeCl₃ and other trace sulfur and organic species. The spatial (longitudinal) and temporal (day/night) variations will be observed throughout 2 circumnavigations with a sample cadence of 12 minutes. 2) Shelley (nephelometer) will determine the size distribution of the aerosols (0.4 to 36 um) in the atmosphere. With a size resolution of <0.7 um, it can determine which mode of H2SO4 is present and characterize the large (>30 um) organic particles previously detected by the Venera and Galileo missions [6,7]. 3) Herbert (UV imager) will measure UV radiance at 283 nm (the wavelength of SO2 absorption) and 365 nm (the unknown part of the absorber). UV images will be taken concurrently with the INMS and nephelometer to correlate UV absorption with abundances of the UV absorber species. 4) Vonnegut (magnetometer) has a range of >600 nT and a precision and accuracy of 1 nT. If magnetized by a past field, the crust may have retained a magnetization of up to 3 A/m2 [5]. With a noise floor of 10 nT, the magnetometer will be able to detect RCM from an altitude of 62 km, even if the thickness of the magnetized crust is just 1 km (Fig. 2). Mission Concept Design: VERNE will be launched with a mass of 3300 kg in an intermediate-high performance class vehicle with a 4-m fairing. The 475-day mission includes 466 days for the cruise, coasting, and orbit initialization phases before entry, descent, and balloon deployment. During the 9-day science phase, the aerial platform will make 2 circumnavigations of the planet at an altitude of 62 km. The INMS and the nephelometer will acquire data for 2 hours during the daytime and nighttime during each circumnavigation, while the UV imager will be on for the duration of the daytime, and the magnetometer will be operational throughout the entirety of the science phase. Data will be stored and processed with the JPL-designed Sphinx command and data handling system. Data will be sent from the balloon to the orbiter using an S-band relay link, stored on the orbiter, and then forwarded to Earth where it will be received by the DSN. Conclusion: VERNE will fill key gaps in our understanding of the history and ongoing processes related to the geology and atmosphere of Venus and rocky worlds in general. Even with adequate flight system contingencies and expected costs below the $900M New Frontiers cost cap, VERNE is not without its risks and challenges. Further trade spaces to explore include 1) using solely battery power vs. including solar panels to increase the mission duration and 2) investigating the use of lightweight materials and 3D-printed structures to reduce the gondola mass, among others. Acknowledgments: We would like to thank the JPL Planetary Science Summer School, especially our mentors Troy Hudson, Karl Mitchell, and Leslie Lowes, as well as our Team-X study lead Al Nash and the members of Team-X. We’d additionally like to thank our review panel for asking insightful questions and providing valuable feedback. References: [1] Ross, F. E. (1928) Astrophysical J., 68, 57-92. [2] Rossow, W. B. et al. (1980) J. Geophysical Research, 85, 8107-8128. [3] Pinto, J. P. et al. (2021) Nature Communications, 12, 175. [4] Titov, D. V. et al. (2007) Cosmic Research, 21, 401. [5] O’Rourke, J. et al. (2019) Geophysical Research Lett., 46, 5768–5777. [6] Limaye, S. S. et al. (2018) Astrobiology, 18(9), 1181-1198. [7] Grinspoon, D. H. et al. (2013) Planetary and Space Sci., 41(7), 515-542. [8] Parker, R. L. (2003) J. Geophysical Research, 108, 5006. *The cost information contained in this document is of a budgetary and planning nature and is intended for informational purposes only. It does not constitute a commitment on the part of JPL and/or Caltech.
Previous investigation of direct delivery of a probe into Neptune’s atmosphere showed said approach to be challenging due to high stagnation pressures and heat loads. Entering at a high entry flight path angle (EFPA) results in stagnation pressure beyond current Heatshield for Extreme Entry Environment Technology (HEEET) qualification, while entering at a lower EFPA results in higher heat loads, leading to required HEEET thicknesses beyond manufacturing capabilities. Aerocapture is an enticing alternative entry scenario because it enables a fast trip and the placement of a payload in orbit around Neptune that can also reach Triton. Here, we investigate entry conditions associated with delivering a probe from orbit to assess feasibility. We consider whether 1) this entry scenario results in reasonable conditions, and 2) HEEET single insulating layer (mid-density carbon phenolic, 3MDCP) can be used to decrease mass, risk, and cost.
Introduction: Uranus was identified as the third highest priority flagship mission in the 2012-2022 Planetary Science Decadal Survey. This latest concept study was requested by the Decadal Survey panel to determine NASA’s planetary science priorities from 2022-2032. This study focused on the probe’s entry and descent aspects and associated trades for viable trajectory options. Uranus Mission and Descent Probe: The proposed Uranus Orbiter and Probe (UOP) Flagship mission will investigate Uranus and its surrounding moons using an orbiting spacecraft with a Uranus descent probe. Unlike previous studies the probe release will occur after orbit insertion allowing sufficient separation of critical events during the orbit insertion burn. The probe will be released at an altitude that allows one hour of in situ atmospheric readings that will be relayed to the orbiter. Afterwards the orbiter will transition to the moon tour phase of the mission. The configuration chosen for the entry aeroshell was a 45° sphere-cone. This shape has been used in the past in the Pioneer Venus Galileo missions. However, the nose radius considered in the present study differed from the values used in either of the previous configurations, primarily to reduce the heat flux at the stagnation point. A two-step approach was used in the development of flight trajectories for the chosen configuration. In the first step, the trajectory code POST2 was used to screen the thousands of entry states provided by interplanetary trajectory simulations, which were terminated at an altitude of 2000 km from the reference surface (1 bar) of Uranus. The screening criteria were: (i) optimization of the communication geometry between the entry probe and orbiter to ensure at least 1 hour of science measurements, (ii) peak stagnation point pressures to be less than six bar, (iii) peak heat fluxes to be less than 5 kW/cm2; the latter two constraints being the limits of ground-test capabilities of the arc jets at NASA Ames Research Center. The entry team investigated two trajectories that met the criteria above, a shallow entry (high heat load ~44 kJ/cm2) and a steeper entry (high heat rate ~1950 W/cm2). In the second step, the two bounding candidate entry states from the POST2 screening process were used in developing flight trajectories using TRAJ coupled with FIAT (a materials thermal response and sizing code) and a margins policy to determine a margined uniform thickness (hence mass) of the forward heatshield material based on the aerothermal environments at the stagnation point. Since the combination of TRAJ and FIAT size the TPS based on stagnation point environments only, flow field computations using DPLR were necessary to determine turbulent aerothermal environments on the conical flank, and the augmentation of these environments due to surface roughness. The environments at select locations on the forward heatshield were then used to size the thermal protection material, with the largest thickness value then used to estimate the mass. The newly developed woven thermal protection material –HEEET (Heatshield for Extreme Entry Environments Technology) –was considered for the forward heatshield and PICA (Phenolic-Impregnated Carbon Ablator) was considered for the backshell. These NASA-developed materials are at TRL 6 and TRL 9, respectively. Furthermore, two options were considered for the HEEET material: (i) a dual-layer option with a denser recession layer on top and an insulative layer underneath it, and (ii) a single layer option consisting of the insulative layer alone. Results: It is clear that probe entry states are feasible and the selected TPS options are able to perform in the predicted aerothermal environments thus enabling the mission to meet of the descent probe portion of this flagship mission
Previous investigation of direct delivery of a probe into Neptune’s atmosphere showed said approach to be challenging due to high stagnation pressures and heat loads. Entering at a high entry flight path angle (EFPA) results in stagnation pressure beyond current Heatshield for Extreme Entry Environment Technology (HEEET) qualification, while entering at a lower EFPA results in higher heat loads, leading to required HEEET thicknesses beyond manufacturing capabilities. Aerocapture is an enticing alternative entry scenario because it enables a fast trip and the placement of a payload in orbit around Neptune that can also reach Triton. Here, we investigate entry conditions associated with delivering a probe from orbit to assess feasibility. We consider whether 1) this entry scenario results in reasonable conditions, and 2) HEEET single insulating layer (mid-density carbon phenolic, 3MDCP) can be used to decrease mass, risk, and cost.
Interplanetary spacecraft are exposed to meteoroid fluxes that range in speeds from 10 to 72 km/sec, far above the capability of today’s test facilities to provide predictions for the likelihood of spacecraft critical penetration. Of special interest are sample return missions, which (though protected by shielding) must often survive years of exposure to the meteoroid environment in order to re-enter Earth’s atmosphere with their scientific cargo. This paper describes the simulation of meteoric material damage to thermal protection systems (TPS) housed beneath protective “garage” (shielding) enclosures using the Smooth Particle Hydrodynamics Code (SPHC) operated by the Institute for Defense Analyses and Stellingwerf Consulting in support of ongoing NASA tasks. The study outlined in this paper considered the impact effect of both meteoric materials such as iron, ice, and chondrites (dunnite), and non-meteoric materials such as aluminum and nylon against both external shielding materials (single and dual aluminum bumpers) and Heat shield for Extreme Entry Environments Technology (HEEET) TPS materials, used alone and in conjunction with shielding. A general predictive damage equation to HEEET TPS is developed from these SPHC simulations for velocities up to 70 km/sec.
Phenolic resins are an important component of many ablative heat shield materials, which protect spacecrafts from the extreme temperatures reached during atmospheric entry. Examples include the high-density Heritage Carbon Phenolic (HCP) used in the Pioneer-Venus and Galileo missions, as well as the low-density Phenolic Impregnated Carbon Ablator (PICA) used in the Mars Science Laboratory and Mars 2020 missions. Additionally, recent developments within NASA have produced the mid-density Heatshield for Extreme Entry Environment Technology (HEEET) and its derivative 3D Woven Mid-Density Carbon Phenolic (3MDCP). Unlike the nonporous phenolic in HCP, PICA and HEEET/3MDCP are fabricated by infusing preforms with diluted phenolic formulations to obtain a lower density porous matrix. Despite the importance of the phenolic phase to the material response during entry, the variation in properties of porous and nonporous phenolic is not well understood. Here, we present an investigation of porous and nonporous phenolic resins using molecular dynamics (MD) simulations. Resin cure is mimicked in the simulations through the inclusion of representative reaction templates to generate accurate models of the complex crosslinked structures. To create porous models, explicit solvent molecules are included during the cure simulations. We observe nanoscale separation of the phenolic and solvent phases, which results in significant differences in the final structures of porous and nonporous models. In addition to a quantitative assessment of the network structure and porosity, we elucidate the effects of the phenolic formulation on the final material properties. These results are compared with experimental data as appropriate.