Thermal Protection System Design Research Articles

Chemical and structural degradation of room temperature vulcanizing (RTV) silicone at high temperatures

Room temperature vulcanizing (RTV) silicone is an insulating adhesive used on thermal protection systems of space capsules. RTV materials experience significant heating during entry into an atmosphere. To understand the response of RTV materials in entry environments, RTV 511 and RTV 560 are heated in a mass spectrometer, a thermogravimetric analyzer, and a high-temperature flow tube furnace to understand the underlying chemical and structural degradation processes occurring during the decomposition of the two RTV materials. In addition, infrared spectroscopy is used to identify the functional groups and the polymer backbone configuration. X-ray computed tomography is used to create three-dimensional volumes of the decomposed samples to examine the structure of the degraded RTV materials. The three-dimensional volumes are also used to evaluate the permeability of degraded RTV materials. It is observed that both RTV materials decompose in three stages. Analysis of the three-dimensional volumes generated through X-ray computed tomography indicates that the decomposition begins with the formation of small cavities in both RTV variants. However, as the materials are subjected to higher temperatures, RTV 511 evolves into an outer solid layer with cavities inside, while the solid material in RTV 560 is distributed more homogeneously throughout the material, resulting in a structure with better mechanical integrity. Although most regions of both RTV variants are impermeable, certain regions of RTV 511 have a permeability that is two orders of magnitude higher than the permeability of the base heat shield material. The overall analysis indicates that the response of RTV materials to heat must be considered in the design and analysis of thermal protection systems.

Thermal conductivity prediction of aerogel considering heat treatment effect and the application to a TPS of long endurance HFVs

Purpose For a thermal protection system (TPS) of long endurance hypersonic flight vehicle (HFV), its thermal insulation property not only determines by the manufactured morphology but also changes along time. A thermal conductivity prediction model for aerogel considering heat treatment effect is carried out and applied to solve the heat conduction problem of a TPS. The aim of this study is to provide theoretical and numerical references for further development of aerogels applying to TPSs. Design/methodology/approach A thermal conductivity prediction model for aerogel is established considering treatment effect. The heat conduction problem of a TPS is derived and solved by combining the differential quadrature method and the Runge–Kutta method. The prediction results of aerogel thermal conductivities are verified by comparing with those in literature, while the calculated temperature field of TPS is verified by comparing with that by ABAQUS. Findings Numerical results show that when applying the current prediction model, the calculated high temperature area in the aerogel layer is narrowed due to the decrease of the thermal conductivity during heat treatment process. Originality/value This study will be beneficial to carry out the precise design of TPS for long endurance HFVs.

Flow and heat transfer analysis of laminar flow in hypersonic aircraft with variable specific heat capacity at small attack angles

Purpose As to different angles of attack and nonlinear problems caused by high temperatures in coexisting hypersonic aircraft, people mainly rely on fluid software for research but lack analysis of flow mechanisms. Owing to computational difficulties, few people use numerical algorithms to combine them for discussion. Hence, this study aims to make a deep inquiry into the laminar flow and heat transfer of compressible Newtonian fluid in hypersonic aircraft with small attack angles. Design/methodology/approach In this paper, on the basis of mass, momentum and energy conservation laws, the governing equations of the hypersonic boundary layer are established. Viscosity, specific heat capacity and thermal conductivity are considered nonlinear functions concerning temperature. In virtue of the MacCormack finite difference method, the stationary numerical solutions are solved directly, and the validity of the algorithm is verified. Findings The results demonstrate that at Mach number 5, compared to the 0° attack angle, the maximum temperature near-wall at the 3° attack angle increases by about 25%. An enjoyable phenomenon is discovered, where the position corresponding to the maximum wall shear force shifts back as the attack angle and Mach number increase. The relationship between the near-wall maximum temperature versus attack angle and Mach number is fitted through numerical calculation results. Originality/value Empirical formulas can be used to estimate heat transfer characteristics at small attack angles, which will guide the design of aircraft thermal protection systems.

Pyrolysis Behavior, Pyrolysis Kinetics, Heat Transfer and Pyrolysis Simulation of Almandine/Boron Phenolic Resin Ceramizable Composites

Ceramizable composites have attracted great interest for the thermal protection of hypersonic vehicles. In our research described here the pyrolysis process of almandine/boron phenolic resin ceramizable composites was investigated; we note it can be divided into three stages (room temperature to 300 °C, 300 °C to 700 °C and 700 °C to 1300 °C). In addition, the pyrolysis products were identified by an analytical pyrolyzer coupled with a gas chromatography/mass spectrometry set-up (Py-GC/MS), and they were composed of phenols (69.82%), aromatic hydrocarbons (23.89%), and aromatic ethers and aliphatic esters (6.29%). Furthermore, the possible pyrolysis routes were proposed. Moreover, the pyrolysis kinetic parameters were calculated, and the activation energy, order of the reaction and pre-exponential factor were 54.18 KJ/mol, 0.869 and 2.828 s−1, respectively. In addition, Comsol Multiphysics finite element software was used to simulate the heat transfer and pyrolysis behavior. After the composite (with a thickness of 10 mm) was heated for 200 s at a heat flux of 480 KW/m2, the surface and back surface temperatures were 1200.6 °C and 163.7 °C, respectively. The thicknesses of the carbonized, pyrolysis and original layers were 2.11, 3.67 and 4.22 mm, respectively, which reflects the good thermal insulation behavior of the composite. We suggest the combination of calculation and simulation results can provide a theoretical basis for the design of thermal protection systems for hypersonic vehicles which are able to meet long-term and large-area thermal protection requirements.

Bilayer lattice structure integrated with phase change material for innovative thermal protection system design

In recent years, ceramic matrix composite (CMC) lattice structure has become a promising candidate for a new generation of integrated thermal protection systems (ITPSs) used in the ultrahigh-temperature environment. However, the thermal short circuit effect caused by lattice cores limits their application. In this work, we propose a bilayer CMC lattice structure filled with phase change material (PCM) and thermal insulation material for integrated thermal protection systems (ITPSs) to mitigate the thermal short circuit effect. The numerical simulation method is used to verify the effectiveness of the bilayer lattice structure in mitigating the thermal short circuit effect compared with one-layer ITPS. By inserting the middle face sheet, the thermal insulation effect enhances about 16%. Furthermore, the effects of thermal insulation materials, thermal conductivity and thickness of paraffin/expanded graphite composite PCM, and the diameter of the lattice core strut on the thermal insulation performance are discussed. The results show that the thickness of PCM has an optimal solution. Optimizing results show that when the aerogel insulation fiber is used as the thermal insulation material, the thermal conductivity k=0.82 W/(m⋅K) and the thickness t3=10 mm of PCM, the diameter of the core strut d=1.5 mm, the bilayer PCM-ITPS has the best thermal insulation performance. When the ultrahigh-temperature load is continuously loaded, the highest temperature of Top Face Sheet is more than 2300°C, but the highest temperature of Bottom Face Sheet is only 30.12°C, which thermal insulation effect is far better than the traditional ITPS. This work gives a novel scheme to solve the thermal short circuit effect, improving the thermal protection performance and accelerate the application of lattice structure in ITPS.

Deep-Learning Strategy Based on Convolutional Neural Network for Wall Heat Flux Prediction

Aerodynamic thermal prediction plays an important role in the design of hypersonic aircraft, especially in the design of the aircraft’s thermal protection system. The main challenges of the aerothermal prediction lie in the slow converging speed and the strict requirements of the computational grid. In this paper, a convolutional-neural-network-based hybrid-features deep-learning strategy is constructed to efficiently predict aerodynamic heating, which is named the convolutional neural network/hybrid-feature method. The hybrid features of this strategy consist of the normal distribution of physical quantities from the wall and the flow parameters at the extreme temperature point. The strategy, which extends through the multilayer perceptron regression layer method, constructs the relationship between the hybrid features and the wall heat flux to obtain a high-precision model trained by the flowfield data without gradient convergence. It is demonstrated that the model has a better inflow generalization ability to predict wall heat flux with different inflow conditions and angles of attack by zero-angle-of-attack training data, which has great potential in aircraft thermal protection system design and shape optimization.

Thermal Modeling of Umbilical Tower During the Lift-off Of a Launch Vehicle

Umbilical Tower (UT) forms a critical structure along with mobile service tower, launch pedestal and jet deflectors of a rocket launch pad. UT will be exposed to severe thermal environments during lift off of the launch vehicle owing to its close proximity. UT experiences thermal loads until the launch vehicle lifts off and clears it. Hence thermal environments need to be quantified to assess the survivability of UT for subsequent missions. This paper brings out the methodology developed for estimating different thermal environments experienced by UT structure during a typical rocket launch. The engineering method result in a quick estimate of flow properties of the plume from the rocket exhaust compared to CFD, which will be computationally intensive and time consuming to arrive a converged accurate solution for heat flux estimation. ISRO’s launch vehicles have first stage with Solid Rocket Motor (SRM) as core and solid motors/liquid engines as multiple strap-ons. Thermal loads to launch pad structures include radiative and convective heating from exhaust plumes. The aluminum particle laden exhaust plumes of solid rockets are dominating radiative heat source. Convective heating occurs from hot jets and direct plume impingement. The relative movement of launch vehicle with respect to UT structure is an important parameter affecting thermal environments. It varies for each launch depending on the vehicle variant and its lift-off trajectory. This calls for a detailed assessment of thermal loads for thermal adequacy studies. A finite element based thermal model for umbilical tower is also generated for thermal response studies and Thermal Protection System (TPS) design. The computed temperature levels at various location of UT structure are compared with measured values during typical lift-off. Results indicate that temperature levels of UT structure are within allowable limits and the adopted methodology can predict temperature within a maximum difference of ±5%.

Design of Aerospace Vehicles’ Thermal Protection Based on Heat-Insulating Materials with Optimal Structure

Highly porous open-cell carbon materials have great potential for use as high-temperature thermal insulation for space vehicles due to a unique combination of properties: low density, high rigidity, sufficient compressive strength, and low thermal conductivity. The physical properties of these materials essentially depend on their microstructure. This implies the possibility of constructing a new advanced technique for the optimal design of multilayer thermal protection systems for aerospace vehicles, taking into account the dependence of materials’ thermal properties on microstructure. The formulation of the optimization problem traditional to thermal design implies the determination of the layer thicknesses that provide a minimum specific mass of the thermal protection, subject to the specified constraints on the maximum temperatures in the layers. The novelty of this work lies in the fact that, along with the thickness of the layers, the design parameters include the cell diameter and porosity, which characterize the structure of highly porous cellular materials. The innovative part of the presented paper lies in the determination of cell diameter and the porosity of open-cell carbon foam together with the thickness of the layers for multilayer thermal insulation, ensuring the required operational temperature on the boundaries of the layers and a minimum of the total mass of the system. This article reveals new possibilities for using the numerical optimization method to determine the geometric parameters of the thermal protection system and the morphology of the materials used. A new methodology for designing heat-loaded structures based on the simultaneous selection of macro- and micro-parameters of the system is proposed. The basic principles of constructing an algorithm for designing a multilayer thermal protection system are outlined, taking into account the possibility of choosing the parameters of the highly porous materials’ structure. The reliability of the developed optimization method was verified by comparing the results of mathematical modeling with experimental data obtained for highly porous cellular materials with known microstructure parameters.

Mesoscale Simulation of Polymer Pyrolysis by Coarse-Grained Molecular Dynamics: A Parametric Study.

Full comprehension of the pyrolysis of polymer materials is crucial for the design and application of thermal protection systems; however, it involves complex phenomena at different spatial and temporal scales. To bridge the gap between the abundant atomistic simulations and continuum modeling in the literature, we perform a novel mesoscale study of the pyrolysis process using coarse-grained molecular dynamics (CG MD) simulations. Polyethylene (PE) consisting of united atoms including implicit hydrogen is considered a model polymer, and the configurational change of PE in thermal degradation is modeled by applying the bond-breaking phenomenon based on bond energy or bond length criteria. A cook-off simulation is implemented to optimize the heuristic protocol of bond dissociation by comparing the reaction products with a ReaxFF simulation. The aerobic hyperthermal pyrolysis under oxygen bombardment is simulated at a large scale of hundreds of nanometers to observe the intricate phenomena occurring from the surface to the depth inside the material. The intrinsic thermal durability of the model polymer at extreme conditions with and without oxygen environment can be effectively simulated from the proposed mesoscale simulation to predict important thermal degradation properties required for continuum-scale pyrolysis and ablation simulations. This work serves as an initial investigation of polymer pyrolysis at the mesoscale and helps understand the concept at a larger scale.

An IPID Control Method for Aerodynamic Heating Ground Simulation of Hypersonic Vehicles

In order to accurately predict the thermal environment of hypersonic vehicles, many scholars have devoted themselves to the research related to the aerodynamic heating characteristics of hypersonic vehicles, which provides important solutions for the design of thermal protection systems. An IPID controller is designed for the aerothermal ground simulation system of hypersonic vehicles to complete the target tracking.

A Parametric Physics-Informed Deep Learning Method for Probabilistic Design of Thermal Protection Systems

Precise and efficient calculations are necessary to accurately assess the effects of thermal protection system (TPS) uncertainties on aerospacecrafts. This paper presents a probabilistic design methodology for TPSs based on physics-informed neural networks (PINNs) with parametric uncertainty. A typical thermal coating system is used to investigate the impact of uncertainty on the thermal properties of insulation materials and to evaluate the resulting temperature distribution. A sensitivity analysis is conducted to identify the influence of the parameters on the thermal response. The results show that PINNs can produce quick and accurate predictions of the temperature of insulation materials. The accuracy of the PINN model is comparable to that of a response surface surrogate model. Still, the computational time required by the PINN model is only a fraction of the latter. Considering both computational efficiency and accuracy, the PINN model can be used as a high-precision surrogate model to guide the TPS design effectively.

A Markov chain Monte Carlo analysis of carbon-phenolic ablation through a lumped physical model

Payloads and other rocket components use Thermal Protection Systems (TPS) to avoid the effects of high heat fluxes generated by aerodynamic heating or propellant burning. The most used type of TPS is the ablative thermal shield, employing polymeric composite ablators. Ablative and thermophysical properties of such materials must be known for TPS design. These properties are often difficult to access experimentally, and vary according to the manufacturing process, ratio between reinforcement and matrix and the fiber direction in the case of composites, and other process variables. Inverse solution methods were initially applied in computational models to solve the heat transfer problems in space vehicles. This work proposes an inverse methodology, using the Markov chain Monte Carlo (MCMC) method, to determine the ablative properties of carbon-phenolic materials manufactured by biased tape wrapping, starting from arc-jet ablation results of samples of this material. A simplified but robust lumped physical model is used to represent the direct problem. The obtained values of these properties were used in the direct problem in a more refined mesh, and the simulation was compared to the experiments. Good agreement between numerical and experimental data was obtained, allowing the use of this methodology in estimation of ablation properties.

Experimental and numerical study of the thermal response of a diesel fuel tank exposed to fire impingement

Fire exposure of diesel fuel tanks can lead to catastrophic accidents. In this study, a fire test was conducted with an 80 L cuboid fuel tank filled to 18% capacity with diesel. More importantly, an innovative three-dimensional computational fluid dynamics (CFD) model was developed aiming at predicting fuel tanks pressurization rate and temperature distributions in the inner fluid. The CFD model was verified using the experimental data and the result showed close agreement. Then, the attention was focused on the influence of different fire engulfment zones and filling levels on the thermal response of the diesel fuel tank. The results indicated that, when fire engulfed the bottom and sides of the diesel fuel tank, temperature stratification appeared in the vapor space and gradually diminished with increasing filling level; however, the liquid phase could be considered isothermal, which was dominated by the convection at the bottom. Additionally, the high filling level caused the pressurization of the fuel tank to slow down due to the absence of temperature stratification in the liquid phase. The combined experimental and numerical analysis can offer valuable guidance for improve the design of thermal protection systems and the fire risk management of fuel tank explosions.

Numerical Simulation of an Electrothermal Ice Protection System in Anti-Icing and Deicing Mode

The design of efficient thermal ice protection systems is a challenging task as these systems operate in complex environments involving several coupled physical phenomena such as phase change, boundary-layer flow, and heat transfer. Moreover, certification rules are becoming more stringent, and there is a strong incentive for the reduction of fuel consumption. In this context, numerical tools provide a powerful asset during the design phase but also to gain insight into the physical mechanisms at play. This article presents modeling and simulation strategies for thermal ice protection systems. First, the model describing the behavior of the thermal protection system is presented. Second, a model and associated numerical method is presented for unsteady ice accretion. Third, the coupling methodology between the ice accretion solver and the heat conduction solver is described. In the fourth part, different methods to simulate the boundary-layer flow are described. Finally, some relevant examples are presented, both in steady and unsteady configurations.

Dynamical behaviors of aluminum droplets impinging on horizontal and inclined surfaces

Aluminum (Al) droplets collide with the wall of the combustion chamber in a solid rocket motor at different angles of incidence and velocities, which exhibit rebounding, splashing, and other behaviors. It increases the loss of two-phase flow to degrade the performance of the engine and poses a challenge to the thermal protection system. In this study, an experimental system in which molten Al droplets impact the walls of the chamber is designed. The size of the Al droplets produced by the generator ranged from 120 μm to 2000 μm while their speeds of impact ranged from 0.85 m/s to 9 m/s. This was used to examine the process of impact of Al droplets on flat and inclined walls. The residence time and spreading diameter of the droplets were quantified to reveal their mechanism of collision. Their empirical correlations were obtained by regressing the experimental data, and the predictive error of the model was within 20% of the experimental values. Furthermore, a theoretical model was developed to analyze energy distribution during the collision between the droplets and the wall. It showed that the effect of viscous dissipation energy on droplet flow was negligibly small, but the adhesion energy of the droplet increased significantly with its initial energy. In addition, the dimensionless sliding distance was S≈8 when the tangential Weber number was greater than 90, where this was related to the high surface tension of the Al droplets. The results here provide a theoretical basis for the design of thermal protection system of solid rocket motors.

Heat transfer and pyrolysis gas flow characteristics of light-weight ablative thermal protection system in the blunt body

Thermal protection system (TPS) is the critical system to protect the structure and electronic components of hypersonic vehicles from aerodynamic heating. Since the blunt body suffers the most severe thermal environment in a flight vehicle, the design of TPS for the blunt body is crucial. To investigate the thermal performance of the light-weight ablative material in the blunt body, a multi-dimensional thermal response model is proposed considering heat transfer, pyrolysis gas flow, and gas blowing effect. The characteristics of heat diffusion, pyrolysis process, gas flow, and the effect of pyrolysis gas blowing are investigated. It shows that the developed model can accurately predict temperature and gas flow. The temperature and charred thickness at the stagnation point are the largest, decrease along the surface of the sphere nose, and keep almost the same at the cone. Due to the surface pressure distribution, a large part of the pyrolysis gas in the sphere nose flows through the ablative material and injects into the air at the sphere-cone junction. In contrast, the gas at the cone flows vertically to the cone surface. The gas flow rate and pressure decrease as time elapses. Moreover, gas blowing mainly influences the thermal response near the sphere-cone junction, leading to a minimum blowing factor of 0.965 and a max temperature reduction of 7.4% at 100s. The effect of gas blowing is enhanced with the increase of phenolic resin content. This study deepens the understanding of the thermal performance of light-weight ablative material in the blunt body and guides the ablative thermal protection design.

Kinetic Monte Carlo modeling of the interface heterogeneous catalysis effect based on machine learning

<p indent="0mm">Heterogeneous catalysis in a high-enthalpy environment has a remarkable influence on the aerodynamic heating of hypersonic aircraft, and modeling this multiscale effect is an important prerequisite for the lightweight and low-redundancy design of an aircraft thermal protection system. A mesoscale interface heterogeneous catalysis model applicable to macroscale flow field calculation was established by the Monte Carlo method. By the radial basis function neural network algorithm, a machine learning-based reaction kinetics analytical surrogate model was constructed, by which fast and accurate prediction of the catalytic recombination coefficient was realized. The results showed that the machine learning model predicted a recombination coefficient similar to the kinetic Monte Carlo method while reducing the magnitude of the calculation time to that of the traditional macroscale method. Coupled with the macroscale CFD solver, the proposed model can efficiently predict surface heat flux with a multiscale heterogeneous catalysis effect. Modeling of microscale results by the machine learning method and its application in macroscale simulation not only covers the heterogeneous catalytic kinetics but also improves computational efficiency. This modeling provides a theoretical basis for multiscale modeling of complicated high-temperature effects on hypersonic aircraft surfaces.

Application of improved k-ω-γ model to deformed hypersonic Inflatable Aerodynamic Decelerator aeroshell

The scalloping surface of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) promotes flow transition. Precise prediction of flow transition and surface aeroheating over the HIAD is imperative to the design of thermal protection system. In this paper, the improved k-ω-γ model considering the separation instability is first applied to the deformed HIAD aeroshell under various angles of attack to assess and verify its performance on such complex configuration, where multiple transition mechanisms coexist and interact. The results of the original k-ω-γ model are also provided for comparison and the prediction mechanisms of the improved model are further dissected. The results reveal that the flow separation and crossflow caused by the scalloping surface are highly sensitive to the angle of attack (α). With the α increasing, an opposite tendency between them can be found, that is, larger α causes stronger and earlier crossflow, but much later flow separation, which leads to the alternation of dominant instabilities for transition in turn. At the relatively high angle of attack, the crossflow becomes critical, at a small angle of attack, however, the separation instability is dominant. The first-mode disturbance can also accelerate the flow instability, while the influence of second-mode disturbance can be neglected due to relatively low post-shock Mach numbers. As a result, the original k-ω-γ model, not attempting to incorporate separated flow transition, predicts an opposite tendency to the experiments. While the transition onset, extent and shape simulated by the improved k-ω-γ model compare quite well with the experiments.

Transient Heat Transfer Analysis on Composite Solids for Hypersonic Flow Applications

Surface heating rates have the importance in high speed flow due to viscous or aerodynamic heating which leads to enormous rise in surface temperature. Prediction of exact heat transfer rates is essential for design of protection systems for aerodynamic bodies aimed. For the thermal protection system, the aerodynamic Surfaces are generally configured with multiple solids. Hence present study is planned to develop the conjugate heat transfer (CHT) solver for composite solids. A CHT solver is in hypersonic flow environment. Finite volume framework and explicit time stepping have been implemented for both the compressible Navier-Stokes solver and conduction solver. Loose coupling has been established between the above mentioned fluid and solid domain solvers. The following observations were made in the CHT simulations. In the case of hypersonic flows over finite thickness flat plate, considerable difference in surface heat transfer rate has been noticed along the stream wise direction at the solid-fluid interface and also the Effect of wall material on heat diffusion for hypersonic application has been analyzed using loosely coupled CHT solver. The CHT solver has also been implemented for hypersonic flow over composite cylinder. Necessity of insulating material in the region of stagnation point has been realized as the outcome of these studies.&#x0D; HIGHLIGHTS&#x0D; &#x0D; Accurate prediction of Convective heating rates is essential for the design of thermal protection system for hypersonic vehicles&#x0D; In The thermal protection system, the aerodynamic surfaces are generally configured with multiple solids. Hence the present study is planned to develop the conjugate heat transfer (CHT) solver for composite solids&#x0D; The CHT solver has been implemented for hypersonic flow over composite flat plate and cylinder. It has been observed flows over finite thickness flat plate, a considerable difference in surface heat transfer rate has been noticed along the stream wise direction at the solid-fluid interface and the necessity of insulating material in the region of stagnation point has been realized as the outcome of these studies&#x0D; &#x0D; GRAPHICAL ABSTRACT

Assessment of a 3D ablation material response model for lightweight quartz fiber reinforced phenolic composite

AbstractTo understand the ablation mechanism of materials is crucial for the design of thermal protection system (TPS) for manned reentry vehicles. To investigate the thermal behavior of lightweight quartz fiber reinforced phenolic (LQP) composite subjected to an aerodynamic hyper‐thermal environment, a 3D ablation material response model for the LQP composite is developed with surface and volume ablation in consideration. It is used to predict surface and back temperature, linear ablation velocity, and quality loss rate of the material. The 3D temperature field, pyrolysis degree, gas flow, and pressure field are analyzed as well. To evaluate the accuracy of the model, an ablation test of oxyacetylene flame is adopted on the LQP composite. In the ablation process, the pyrolysis gases propagate primarily through the surface primarily and through the sidewall afterwards. The gas pressure peaks at the early stage rather than increasing constantly. The analysis of thermal insulation mechanism of the LQP composite indicates that the radiation and heat blockage effect are the primary factors to improve the ablation resistance, accounting for 50.0% and 37.3% of the improvement, respectively. The low thermal conductivity determines the admirable thermal insulation performance. The surface temperature of material can be reduced by introducing high emissivity powder into the material, preparing coating on the surface, and adjusting the content of phenolic and silica fiber.