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NASA NTRS2019

Sustaining Phenolic Impregnated Carbon Ablator (PICA) for Future NASA Missions Including Discovery and New Frontiers

Phenolic Impregnated Carbon Ablator (PICA) was invented in the mid 1990's and due to its relatively low density and efficient performance has been the heat shield TPS of choice for a range of missions includ-ing, Stardust, OSIRIS-Rex, Mars Science Laboratry (MSL) and Mars 2020. PICA has also been the TPS solution on numerous Discovery and New Frontiers proposals, as both the heat shield and back shell TPS and is under consideration as both for the Mars Sample Return Earth Entry Vehicle (EEV) and the heat shield on the Sample Retrieval Lander (SRL). Recently NASA's Science Mission Directorate (SMD) has funded an activity to develop a more sus-tainable version of PICA and to expand the demon-strated capabilities of PICA both in manufacturing and aerothermal performance.

Thermal Protection Material↗

NASA NTRS2010

Microscopic Scale Simulation of the Ablation of Fibrous Materials

Ablation by oxidation of carbon-fiber preforms impregnated in carbonized phenolic matrix is modeled at microscopic scale. Direct numerical simulations show that the carbonized phenolic matrix ablates in volume leaving the carbon fibers exposed. This is due to the fact that the reactivity of carbonized phenolic is higher than the reactivity of carbon fibers. After the matrix is depleted, the fibers ablate showing progressive reduction of their diameter. The overall material recession occurs when the fibers are consumed. Two materials with the same carbon-fiber preform, density and chemical composition, but with different matrix distributions are studied. These studies show that at moderate temperatures (< 1000 K) the microstructure of the material influences its recession rate; a fact that is not captured by current models that are based on chemical composition only. Surprisingly, the response of these impregnated-fiber materials is weakly dependent on the microstructure at very high temperatures (e.g., Stardust peak heating conditions: 3360K).

Lachaud, Jean Romain↗

NASA NTRS2012

Nonequilibrium Ablation of Phenolic Impregnated Carbon Ablator

In previous work, an equilibrium ablation and thermal response model for Phenolic Impregnated Carbon Ablator was developed. In general, over a wide range of test conditions, model predictions compared well with arcjet data for surface recession, surface temperature, in-depth temperature at multiple thermocouples, and char depth. In this work, additional arcjet tests were conducted at stagnation conditions down to 40 W/sq cm and 1.6 kPa. The new data suggest that nonequilibrium effects become important for ablation predictions at heat flux or pressure below about 80 W/sq cm or 10 kPa, respectively. Modifications to the ablation model to account for nonequilibrium effects are investigated. Predictions of the equilibrium and nonequilibrium models are compared with the arcjet data.

Milos, Frank S.↗

NASA NTRS2003

Explicit Pore Pressure Material Model in Carbon-Cloth Phenolic

An explicit material model that uses predicted pressure in the pores of a carbon-cloth phenolic (CCP) composite has been developed. This model is intended to be used within a finite-element model to predict phenomena specific to CCP components of solid-fuel-rocket nozzles subjected to high operating temperatures and to mechanical stresses that can be great enough to cause structural failures. Phenomena that can be predicted with the help of this model include failures of specimens in restrained-thermal-growth (RTG) tests, pocketing erosion, and ply lifting

Gutierrez-Lemini, Danton↗

NASA NTRS2013

Fracture in Phenolic Impregnated Carbon Ablator

This paper describes the development of a novel technique to understand the failure mechanisms inside thermal protection materials. The focus of this research is on the class of materials known as phenolic impregnated carbon ablators. It has successfully flown on the Stardust spacecraft and is the thermal protection system material chosen for the Mars Science Laboratory and SpaceX Dragon spacecraft. Although it has good thermal properties, structurally, it is a weak material. To understand failure mechanisms in carbon ablators, fracture tests were performed on FiberForm(Registered TradeMark) (precursor), virgin, and charred ablator materials. Several samples of these materials were tested to investigate failure mechanisms at a microstructural scale. Stress-strain data were obtained simultaneously to estimate the tensile strength and toughness. It was observed that cracks initiated and grew in the FiberForm when a critical stress limit was reached such that the carbon fibers separated from the binder. However, both for virgin and charred carbon ablators, crack initiation and growth occurred in the matrix (phenolic) phase. Both virgin and charred carbon ablators showed greater strength values compared with FiberForm samples, confirming that the presence of the porous matrix helps in absorbing the fracture energy.

Agrawal, Parul↗

NASA NTRS2004

Thermostructural Analysis of Carbon Cloth Phenolics "Ply Lifting" and Correlation to LHMEL Test Results

This paper provides a discussion of the history of Carbon Cloth Phenolic (CCP) ply lifting in the Redesigned Solid Rocket Motor (RSRM) Program, a brief presentation of theoretical methods used for analytical evaluation, and results of parametric analyses of CCP material subject to test conditions of the Laser Hardened Material Evaluation Laboratory. CCP ply lift can occur in regions of the RSRM nozzle where ply angle to flame surface is generally less than about 20 degrees. There is a heat rate dependence on likelihood and severity of the condition with the higher heating rates generally producing more ply lift. The event occurs in-depth, near the heated surface, where the load necessary to mechanically separate the CCP plies is produced by the initial stages of pyrolysis gas generation due to the thermal decomposition of the phenolic resin matrix. Due to the shallow lay-up angle of the composite, normal components of the indepth mechanical load, due to "pore pressure", are imparted primarily as a cross-ply tensile force on the interlaminar ply boundaries. Tensile capability in the cross-ply (out of plane) direction is solely determined by the matrix material capability. The elevated temperature matrix material capabilities are overcome by pressure induced mechanical normal stress and ply-lift occurs. A theoretical model used for CCP in-depth temperature, pressure, and normal stress prediction, based on first principles, is briefly discussed followed by a parametric evaluation of response variables subject to boundary conditions typical of on-going test programs at the LHMEL facility. Model response demonstrates general trends observed in test and provides insight into the interactivity of material properties and constitutive relationships.

Clayton, Louie↗

NASA NTRS2016

Investigation of Performance Envelope for Phenolic Impregnated Carbon Ablator (PICA)

The present work provides the results of a short exploratory study on the performance of Phenolic Impregnated Carbon Ablator, or PICA, at high heat flux and pressure in an arcjet facility at NASA Ames Research Center. The primary objective of the study was to explore the thermal response of PICA at cold-wall heat fluxes well in excess of 1500 W/cm (exp 2). Based on the results of a series of flow simulations, multiple PICA samples were tested at an estimated cold wall heat flux and stagnation pressure of 1800 W/cm (exp 2) and 130 kPa, respectively. All samples survived the test, and no failure was observed either during or after the exposure. The results indicate that PICA has a potential to perform well at environments with significantly higher heat flux and pressure than it has currently been flown.

Ablative Material↗

NASA NTRS2014

Thermal Testing of Woven TPS Materials in Extreme Entry Environments

NASAs future robotic missions to Venus and outer planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of current mid density ablators (PICA or Avcoat). Therefore mission planners assume the use of a fully dense carbon phenolic heatshield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic (CP) is a robust TPS however its high density and thermal conductivity constrain mission planners to steep entries, high heat fluxes, high pressures and short entry durations, in order for CP to be feasible from a mass perspective. In 2012 the Game Changing Development Program in NASAs Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System to meet the needs of NASAs most challenging entry missions. The high entry conditions pose certification challenges in existing ground based test facilities. Recent updates to NASAs IHF and AEDCs H3 high temperature arcjet test facilities enable higher heatflux (2000 Wcm2) and high pressure (5 atm) testing of TPS. Some recent thermal tests of woven TPS will be discussed in this paper. These upgrades have provided a way to test higher entry conditions of potential outer planet and Venus missions and provided a baseline against carbon phenolic material. The results of these tests have given preliminary insight to sample configuration and physical recession profile characteristics.

TPS↗

NASA NTRS2015

Pressure Venting Tests of Phenolic Impregnated Carbon Ablator (PICA)

A series of tests was devised to investigate the pressure venting behavior of one of the candidate ablators for the Orion capsule heat shield. Three different specimens of phenolic impregnated carbon ablator (PICA) were instrumented with internal pressure taps and subjected to rapid pressure changes from near vacuum to one atmosphere and simulated Orion ascent pressure histories. The specimens vented rapidly to ambient pressure and sustained no detectable damage during testing. Peak pressure differences through the thickness of a 3-inch-thick specimen were less than 1 psi during a simulated ascent pressure history.

Blosser, Max L.↗

NASA NTRS2011

Fracture in Phenolic Impregnated Carbon Ablator

The thermal protection materials used for spacecraft heat shields are subjected to various thermal-mechanical loads during an atmospheric reentry which can threaten the structural integrity of the system. This paper discusses the development of a novel technique to understand the failure mechanisms inside the thermal protection material, Phenolic Impregnated Carbon Ablator (PICA). PICA has successfully flown on the Stardust spacecraft and was the TPS material chosen for the Mars Science Laboratory (MSL), that will fly in 2011. Although PICA has good thermal properties, structurally, it is a weak material. To thoroughly understand failure in PICA, experiments were performed using FiberForm(Registered TradeMark) (precursor of PICA), virgin and furnace-charred PICA. Several small samples were tested inside an electron microscope to investigate the failure mechanisms. Micrographs were obtained before and after the failure in order to study crack initiation and growth. Videos were obtained to capture failure mechanisms in real time. Stress-strain data was obtained simultaneously for all the samples with the help of a data acquisition system, integrated to the mechanical stages. It was found that cracks initiated and grew in the FiberForm when a critical stress limit was reached such that the carbon fibers separated from the binder. However, both for virgin and charred PICA, crack initiation and growth occurred in the matrix (phenolic) phase. Both virgin and charred PICA showed greater strength values compared to FiberForm coupons, confirming that the presence of the porous matrix helps in absorbing the fracture energy.

Agrawal, Parul↗

NASA NTRS2011

Fracture in Phenolic Impregnated Carbon Ablator

The thermal protection materials used for spacecraft heat shields are subjected to various thermal-mechanical loads during an atmospheric entry which can threaten the structural integrity of the system. This paper discusses the development of a novel technique to understand the failure mechanisms inside thermal protection materials. The focus of research is Phenolic Impregnated Carbon Ablator (PICA). It has successfully flown on the Stardust spacecraft and is the TPS material chosen for the Mars Science Laboratory (MSL) and Dragon spacecraft. Although PICA has good thermal properties, structurally, it is a weak material. In order to thoroughly understand failure in PICA, fracture tests were performed on FiberForm* (precursor of PICA), virgin and charred PICA materials. Several samples of these materials were tested to investigate failure mechanisms at a microstructural scale. Stress-strain data were obtained simultaneously to estimate the fracture toughness. It was found that cracks initiated and grew in the FiberForm when a critical stress limit was reached such that the carbon fibers separated from the binder. However, both for virgin and charred PICA, crack initiation and growth occurred in the matrix (phenolic) phase. Both virgin and charred PICA showed greater strength values compared to FiberForm coupons, confirming that the presence of the porous matrix helps in absorbing the fracture energy.

Agrawal, Parul↗

NASA NTRSDate not supplied

Characterizing Porous and Nonporous Phenolic Resins from Molecular Dynamics Simulations

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.

phenolic↗

NASA NTRS2002

Thermostructural Analysis of Carbon Cloth Phenolic Material Tested at the Laser Hardened Material Evaluation Laboratory

RSRM nozzle liner components have been analyzed and tested to explore the occurrence of anomalous material performance known as pocketing erosion. Primary physical factors that contribute to pocketing seem to include the geometric permeability, which governs pore pressure magnitudes and hence load, and carbon fiber high temperature tensile strength, which defines a material limiting capability. The study reports on the results of a coupled thermostructural finite element analysis of Carbon Cloth Phenolic (CCP) material tested at the Laser Hardened Material Evaluation Laboratory (the LHMEL facility). Modeled test configurations will be limited to the special case of where temperature gradients are oriented perpendicular to the composite material ply angle. Analyses were conducted using a transient, one-dimensional flow/thermal finite element code that models pore pressure and temperature distributions and in an explicitly coupled formulation, passes this information to a 2-dimensional finite element structural model for determination of the stress/deformation behavior of the orthotropic fiber/matrix CCP. Pore pressures are generated by thermal decomposition of the phenolic resin which evolve as a multi-component gas phase which is partially trapped in the porous microstructure of the composite. The nature of resultant pressures are described by using the Darcy relationships which have been modified to permit a multi-specie mass and momentum balance including water vapor condensation. Solution to the conjugate flow/thermal equations were performed using the SINDA code. Of particular importance to this problem was the implementation of a char and deformation state dependent (geometric) permeability as describing a first order interaction between the flow/thermal and structural models. Material property models are used to characterize the solid phase mechanical stiffness and failure. Structural calculations were performed using the ABAQUS code. Iterations were made between the two codes involving the dependent variables temperature, pressure and across-ply strain level. Model results comparisons are made for three different surface heat rates and dependent variable sensitivities discussed for the various cases.

Clayton, J. Louie↗

NASA NTRS2019

Assessment of a One-Dimensional Finite Element Charring Ablation Material Response Model for Phenolic-Impregnated Carbon Ablator

In this study, mathematical formulations to model charring ablation problems were numerically implemented using finite element analysis (FEA) with ABAQUS, which account for the material decomposition and progressive surface removal in the heat conduction and the surface energy balance equations. FEA was performed for a one-dimensional model to predict the temperature and ablation histories of a phenolic-impregnated carbon ablator sample (i.e., a common heat shield material for hypersonic vehicles and spacecraft) subjected to oxy-acetylene torch flame (i.e., 0.8 SLPM acetylene gas to 2.7 SLPM oxygen gas). The recovery enthalpy and convective heat transfer coefficient for the ablation model were calculated based on gas compositions and two assumed surface conditions (i.e., equilibrium and frozen). Simulations using the calculated recovery enthalpy and convective heat transfer coefficient resulted in a recession rate of 6.38 times (equilibrium) and 14.08 times (frozen) higher than the experimental data, despite fair agreement of the surface temperature. In addition, the effect of the heat transfer coefficient was investigated through a steady-state ablation analysis. The results of the analysis indicate that there is not one single value for the heat transfer coefficient that would allow the prediction to match both measured recession rate and surface temperature. Possible reasons for such an inconsistency are provided and discussed.

Thermal Protection System↗

NASA NTRS2019

A Discussion of the Need to Sustain Mission Ready TPS and for Continued Development of Innovative Entry System Technologies

Flight proven entry system and TPS technologies are critical for the successful execution of in-situ science missions at Venus. Emerging new technologies point to new possibilities and offer innovative approaches to delivering small satellites for orbital science. Venus entry can be very demanding and there are only a few flight proven TPS, some developed by Industry and others by NASA, capable of meeting the mission needs. NASA developed TPS has predominately been transferred to Industry and it is assumed industry will maintain the fabrication capability. However, lack of mission needs may result in obsolence of TSP fabrication capability if there is no money and no motivation. Even within NASA, its' expertise could be diverted to higher priority objectives and thereby the readiness for particular material systems can be impacted or lost. Atrophy of capabilities can come about in other ways as well such as changes to raw materials. Even small manufacturing process changes can demand requalification and TRL may be degraded. Carbon-Phenolic is a text book example. After a long period of absence of US Venus missions, VEXG and the Science community is making the case for future missions. It is insufficient to assume the TSP technologies will be there in 5 or 10 years without active and continual planning and assessment. After Galileo, Carbon-Phenolic materials and fabrication skills were allowed to atrophy. Then when missions needed it, in early 2000, it was no longer possible to make the heritage Carbon-Phenolic. What do we need to do? The first step is to advocate for the establishment of TPS readiness assess-ment. The assessment will involve understanding threats and opportunities, and the development of risk mitigation strategies. VEXAG needs to advocate for such an active monitoring of the needed capabilities, assessment of emerging risks and development of risk mitigation strategies with implementation plans. Such an approach reduces the threat of material obsolence and helps maintain the availability of entry system and TPS technology capabilities, both old and new. Venus probes, landers, balloons and other variable altitude missions, and skimmer missions such as "Cu-pid's Arrow" as well as aerocapture missions to deliver small spacecraft require qualified entry systems and ablative TPS. VEXAG advocated for HEEET in 2013/2014 and the community is well versed with the need to sustain it. But, other TPS that need to be sustained may not be apparent to VEXAG community. The following figure summarizes the ablative TPS capabilities vs Venus mission needs for both primary heatshield and backshell.

Venkatapathy, Ethiraj↗

NASA NTRS2014

Woven Thermal Protection System (WTPS) a Novel Approach to Meet NASA's Most Demanding Reentry Missions

NASA's future robotic missions to Venus and outer planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of current mid-density ablators (PICA or Avcoat). Therefore mission planners assume the use of a fully dense carbon phenolic heat shield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic (CP) is a robust Thermal Protection System (TPS) however its high density and thermal conductivity constrain mission planners to steep entries, high heat fluxes, pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose certification challenges in existing ground based test facilities. In 2012 the Game Changing Development Program in NASA's Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System (WTPS) to meet the needs of NASA's most challenging entry missions. This presentation will summarize maturation of the WTPS project.

carbon phenolic↗

NASA NTRS2014

Woven Thermal Protection System (WTPS) - a Novel Approach to Meet NASA's Most Demanding Missions

NASAs future robotic missions utilizing an entry system into Venus and the outer planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of state of the art low to mid density ablators such as PICA or Avcoat. Therefore mission planners typically assume the use of a fully dense carbon phenolic heat shield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic is a robust TPS material however its high density and relatively high thermal conductivity constrain mission planners to steep entries, with high heat fluxes and pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose challenges for certification in existing ground based test facilities and the longer-term sustainability of CP will continue to pose challenges. In 2012 the Game Changing Development Program (GCDP) in NASAs Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System (WTPS) to meet the needs of NASAs most challenging entry missions. This project was highly successful demonstrating that a Woven TPS solution compares favorably to CP in performance in simulated reentry environments and provides the opportunity to manufacture graded materials that should result in overall reduced mass solutions and enable a much broader set of missions than does CP. Building off the success of the WTPS project GCDP has funded a follow on project to further mature and scale up the WTPS concept for insertion into future NASA robotic missions. The matured WTPS will address the CP concerns associated with ground based test limitations and sustainability. This presentation will discuss results from the WTPS heat-shield for extreme entry environment technology (HEEET) project

woven TPS↗

NASA NTRSDate not supplied

Rocket Lab Venus - Enabling Low-Cost Interplanetary Missions

Rocket Lab’s mission to Venus [1], launching in January 2025, aims to demonstrate that small launch vehicles, such as Electron, and high-energy small spacecraft, such as Photon, can enable a new paradigm of regular, low-cost interplanetary missions. The Electron launch vehicle will deliver the Photon spacecraft carrying the entry probe into low earth orbit, at which point Photon will take over and put both itself and the attached probe on an Earth-Venus transfer trajectory, with assistance from a Lunar flyby. The primary science objective of the mission is to sample the Venusian cloud layers between atitudes of 45 and 60 km above the surface with an autofluorescing nephelometer. The ~1kg scientific instrument will be carried by a small ~17 kg direct entry probe which will separate from the Photon spacecraft approximately 30 minutes prior to atmospheric entry. The probe outer mould line is a scaled down version of the Deep Space 2 probes [2]. NASA’s newly developed 3D Woven Carbon Phenolic, a derivative of the Heatshield for Extreme Entry Environment Technology (HEEET) material [3], is an enabling technology for this mission, and it will comprise the probe forebody heat shield.The probe will communicate directly back to the Earth during its descent via a small on-board radio frequency (RF) antenna. The hyperbolic Venus encounter will see the probe enter the Venusia atmosphere at approximately 11 km/s with an entry flight path angle (EFPA) of about 10°. This relatively low entry angle was selected to increase the duration of the descent phase, reduce the peak deceleration, and reduce the sensitivity of the nominal entry environments to navigational errors. For reference, the Pioneer Venus probes entered at EFPAs varying from about 25° to 70°, resulting in peak decelerations between around 220 and 450 g’s and peak stagnation point heat fluxes ranging from about 40 to 75 MW/m2 [4]. In comparison, the Rocket Lab Venus probe will experience a lower preak deceleration of around 60 g’s and a lower peak stagnation point heat flux of about 15 MW/m2, at the cost of a greater total heat load and correspondingly higher thermal protection system (TPS) mass fraction. This paper recounts the development history of the TPS for the Rocket Lab Venus probe, with a focus on the unique requirements of the mission. In particular, the effect on the TPS design of the mission’s tight budget, short development timeline, slim mass and RF link margins, and mild but long duration entry heating. Rocket Lab’s initial investigations into a ‘hybrid’ TPS [5] based on a scaled version of the Pioneer Venus full-density carbon phenolic forebody shield, and the abandonment of this approach in favour of NASA’s 3D Woven Carbon Phenolic is presented. Additionally, the leveraging of previous testing under the HEEET project at higher conditions and the associated TPS margining strategy that was adopted to avoid the need for extensive mission-specific arc jet testing is discussed. Finally, this paper concludes with a brief consideration of the feasibility of a future low-cost aerocapture technology demonstrator mission based on the 2025 Rocket Lab Venus mission and entry probe.

Lyle Campbell↗