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Heatshield design for spacecraft entering the atmosphere of Mars may be affected by the presence of atmospheric dust. Particle impacts with sufficient kinetic energy can cause spallation damage to the heatshield that must be estim...
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Heatshield design for spacecraft entering the atmosphere of Mars may be affected by the presence of atmospheric dust. Particle impacts with sufficient kinetic energy can cause spallation damage to the heatshield that must be estimated. The dust environment in terms of particle size distribution and number density can be inferred from ground-based or atmospheric observations at Mars. Using a Lagrangian approach, the particle trajectories through the shock layer can be computed using a set of coupled ordinary differential equations. The dust particles are small enough that non-continuum effects must be accounted for when computing the drag coefficient and heat transfer to the particle surface. Surface damage correlations for impact crater diameter and penetration depth are presented for fused-silica, AVCOAT, Shuttle tiles, cork, and Norcoat Liege. The cork and Norcoat Liege correlations are new and were developed in this study. The modeling equations presented in this paper are applied to compute the heatshield erosion due to dust particle impacts on the ExoMars Schiaparelli entry capsule during dust storm conditions.
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摘要 :
Heatshield design for spacecraft entering the atmosphere of Mars may be affected by the presence of atmospheric dust. Particle impacts with sufficient kinetic energy can cause spallation damage to the heatshield that must be estim...
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Heatshield design for spacecraft entering the atmosphere of Mars may be affected by the presence of atmospheric dust. Particle impacts with sufficient kinetic energy can cause spallation damage to the heatshield that must be estimated. The dust environment in terms of particle size distribution and number density can be inferred from ground-based or atmospheric observations at Mars. Using a Lagrangian approach, the particle trajectories through the shock layer can be computed using a set of coupled ordinary differential equations. The dust particles are small enough that non-continuum effects must be accounted for when computing the drag coefficient and heat transfer to the particle surface. Surface damage correlations for impact crater diameter and penetration depth are presented for fused-silica, AVCOAT, Shuttle tiles, cork, and Norcoat Liege. The cork and Norcoat Liege correlations are new and were developed in this study. The modeling equations presented in this paper are applied to compute the heatshield erosion due to dust particle impacts on the ExoMars Schiaparelli entry capsule during dust storm conditions.
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A spacecraft entering the Martian atmosphere during a dust storm may experience recession to the heatshield due to dust particle impacts. Aerodynamic drag is the primary force that determines the trajectory of the dust particles t...
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A spacecraft entering the Martian atmosphere during a dust storm may experience recession to the heatshield due to dust particle impacts. Aerodynamic drag is the primary force that determines the trajectory of the dust particles through the shock layer. This paper examines the effect of particle drag model on the heatshield recession. Three particle drag models are assessed including two that are intended to be applicable over a wide range of particle flow conditions. Particle trajectories are computed in conditions measured during the 2007 major global dust storm. It was found that accounting for Knudsen number and compressibility effects made a large difference in the estimated particle impact velocity. The two drag models that were valid for transitional, compressible particle flow environments predicted only slightly different amounts of heatshield recession due to dust particle impacts. A brief description of the effects of non-spherical particles on drag coefficients is provided.
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The 2013-2022 Decaedal survey for planetary exploration has identified probe missions to Uranus and Saturn as high priorities. This work endeavors to examine the uncertainty for determining aeroheating in such entry environments. ...
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The 2013-2022 Decaedal survey for planetary exploration has identified probe missions to Uranus and Saturn as high priorities. This work endeavors to examine the uncertainty for determining aeroheating in such entry environments. Representative entry trajectories are constructed using the TRAJ software. Flowfields at selected points on the trajectories are then computed using the Data Parallel Line Relaxation (DPLR) Computational Fluid Dynamics Code. A Monte Carlo study is performed on the DPLR input parameters to determine the uncertainty in the predicted aeroheating, and correlation coefficients are examined to identify which input parameters show the most influence on the uncertainty. A review of the present best practices for input parameters (e.g. transport coefficient and vibrational relaxation time) is also conducted. It is found that the 2σ-uncertainty for heating on Uranus entry is no more than 2.1%, assuming an equilibrium catalytic wall, with the uncertainty being determined primarily by diffusion and H_2 recombination rate within the boundary layer. However, if the wall is assumed to be partially or non-catalytic, this uncertainty may increase to as large as 18%. The catalytic wall model can contribute over 3x change in heat flux and a 20% variation in film coefficient. Therefore, coupled material response/fluid dynamic models are recommended for this problem. It was also found that much of this variability is artificially suppressed when a constant Schmidt number approach is implemented. Because the boundary layer is reacting, it is necessary to employ self-consistent effective binary diffusion to obtain a correct thermal transport solution. For Saturn entries, the 2a uncertainty for convective heating was less than 3.7%. The major uncertainty driver was dependent on shock temperature/velocity, changing from boundary layer thermal conductivity to diffusivity and then to shock layer ionization rate as velocity increases. While radiative heating for Uranus entry was negligible, the nominal solution for Saturn computed up to 20% radiative heating at the highest velocity examined. The radiative heating followed a non-normal distribution, with up to a 3x variation in magnitude. This uncertainty is driven by the H_2 dissociation rate, as H_2 that persists in the hot non-equilibrium zone contributes significantly to radiation.
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The 2013-2022 Decaedal survey for planetary exploration has identified probe missions to Uranus and Saturn as high priorities. This work endeavors to examine the uncertainty for determining aeroheating in such entry environments. ...
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The 2013-2022 Decaedal survey for planetary exploration has identified probe missions to Uranus and Saturn as high priorities. This work endeavors to examine the uncertainty for determining aeroheating in such entry environments. Representative entry trajectories are constructed using the TRAJ software. Flowfields at selected points on the trajectories are then computed using the Data Parallel Line Relaxation (DPLR) Computational Fluid Dynamics Code. A Monte Carlo study is performed on the DPLR input parameters to determine the uncertainty in the predicted aeroheating, and correlation coefficients are examined to identify which input parameters show the most influence on the uncertainty. A review of the present best practices for input parameters (e.g. transport coefficient and vibrational relaxation time) is also conducted. It is found that the 2σ-uncertainty for heating on Uranus entry is no more than 2.1%, assuming an equilibrium catalytic wall, with the uncertainty being determined primarily by diffusion and H_2 recombination rate within the boundary layer. However, if the wall is assumed to be partially or non-catalytic, this uncertainty may increase to as large as 18%. The catalytic wall model can contribute over 3x change in heat flux and a 20% variation in film coefficient. Therefore, coupled material response/fluid dynamic models are recommended for this problem. It was also found that much of this variability is artificially suppressed when a constant Schmidt number approach is implemented. Because the boundary layer is reacting, it is necessary to employ self-consistent effective binary diffusion to obtain a correct thermal transport solution. For Saturn entries, the 2a uncertainty for convective heating was less than 3.7%. The major uncertainty driver was dependent on shock temperature/velocity, changing from boundary layer thermal conductivity to diffusivity and then to shock layer ionization rate as velocity increases. While radiative heating for Uranus entry was negligible, the nominal solution for Saturn computed up to 20% radiative heating at the highest velocity examined. The radiative heating followed a non-normal distribution, with up to a 3x variation in magnitude. This uncertainty is driven by the H_2 dissociation rate, as H_2 that persists in the hot non-equilibrium zone contributes significantly to radiation.
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The present paper explores the effect of surface catalytic efficiency of copper calorimeters, used to calibrate an arcjet test, on the predicted recession of an ablative material. The measured heating rate and pressure at the stag...
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The present paper explores the effect of surface catalytic efficiency of copper calorimeters, used to calibrate an arcjet test, on the predicted recession of an ablative material. The measured heating rate and pressure at the stagnation point of the calorimeter are used to infer the centerline enthalpy of the arc-heated flow assuming the exposed copper surface has a catalytic efficiency of unity to atom recombination. Since the inferred enthalpy is utilized in determining the thermal response of the ablative material, the influence of catalytic efficiency on the inferred centerline enthalpy, and its consequent impact on material response are explored in the present work. Comparisons are made of the results from computations performed with a modern CFD code against those from Goulard's engineering model. A parametric study of the effect of surface catalycity and inferred centerline enthalpy on the predicted surface recession of three TPS materials - Teflon, PICA, and molded AVCOAT? - is also performed. For catalytic efficiencies greater than 0.02, the engineering correlation of Goulard and CFD predictions for partially-catalytic heating rate and centerline enthalpy match closely. Over the expected range of copper catalytic efficiencies, the effect on predicted surface recession by assuming a partially-catalytic calorimeter surface was less than 12%.
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The present paper explores the effect of surface catalytic efficiency of copper calorimeters, used to calibrate an arcjet test, on the predicted recession of an ablative material. The measured heating rate and pressure at the stag...
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The present paper explores the effect of surface catalytic efficiency of copper calorimeters, used to calibrate an arcjet test, on the predicted recession of an ablative material. The measured heating rate and pressure at the stagnation point of the calorimeter are used to infer the centerline enthalpy of the arc-heated flow assuming the exposed copper surface has a catalytic efficiency of unity to atom recombination. Since the inferred enthalpy is utilized in determining the thermal response of the ablative material, the influence of catalytic efficiency on the inferred centerline enthalpy, and its consequent impact on material response are explored in the present work. Comparisons are made of the results from computations performed with a modern CFD code against those from Goulard's engineering model. A parametric study of the effect of surface catalycity and inferred centerline enthalpy on the predicted surface recession of three TPS materials - Teflon, PICA, and molded AVCOAT™ - is also performed. For catalytic efficiencies greater than 0.02, the engineering correlation of Goulard and CFD predictions for partially-catalytic heating rate and centerline enthalpy match closely. Over the expected range of copper catalytic efficiencies, the effect on predicted surface recession by assuming a partially-catalytic calorimeter surface was less than 12%.
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The NEQAIR line-by-line radiation code has been incorporated into the DPLR Navier-Stokes flow solver such that the NEQAIR subroutines are now callable functions of DPLR. The coupled DPLR-NEQA1R code was applied to compute the conv...
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The NEQAIR line-by-line radiation code has been incorporated into the DPLR Navier-Stokes flow solver such that the NEQAIR subroutines are now callable functions of DPLR. The coupled DPLR-NEQA1R code was applied to compute the convective and radiative heating rates over high-mass Mars entry vehicles. Two vehicle geometries were considered - a IS m diameter 70-degree sphere cone configuration and a slender, mid-L/D vehicle with a diameter of 5 m called an Ellipsled. The entry masses ranged from 100 to 165 metric tons. Solutions were generated for entry velocities ranging from 6.5 to 9.1 km/s. The coupled fluids-radiation solutions were performed at the peak heating location along trajectories generated by the Traj trajectory analysis code. The impact of fluids-radiation coupling is a function of the level of radiative heating and the freestream density and velocity. For the high-mass Mars vehicles examined in this study, coupling effects were greatest for entry velocities above 8.5 km/s where the surface radiative heating was reduced by up 17%. Generally speaking, the Ellipsled geometry experiences a lower peak radiative heating rate but a higher peak turbulent convective heating rate than the 70-degree sphere cone vehicle.
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The NEQAIR line-by-line radiation code has been incorporated into the DPLR Navier-Stokes flow solver such that the NEQAIR subroutines are now callable functions of DPLR. The coupled DPLR-NEQA1R code was applied to compute the conv...
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The NEQAIR line-by-line radiation code has been incorporated into the DPLR Navier-Stokes flow solver such that the NEQAIR subroutines are now callable functions of DPLR. The coupled DPLR-NEQA1R code was applied to compute the convective and radiative heating rates over high-mass Mars entry vehicles. Two vehicle geometries were considered - a IS m diameter 70-degree sphere cone configuration and a slender, mid-L/D vehicle with a diameter of 5 m called an Ellipsled. The entry masses ranged from 100 to 165 metric tons. Solutions were generated for entry velocities ranging from 6.5 to 9.1 km/s. The coupled fluids-radiation solutions were performed at the peak heating location along trajectories generated by the Traj trajectory analysis code. The impact of fluids-radiation coupling is a function of the level of radiative heating and the freestream density and velocity. For the high-mass Mars vehicles examined in this study, coupling effects were greatest for entry velocities above 8.5 km/s where the surface radiative heating was reduced by up 17%. Generally speaking, the Ellipsled geometry experiences a lower peak radiative heating rate but a higher peak turbulent convective heating rate than the 70-degree sphere cone vehicle.
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The present paper describes preliminary development of an in-depth material response model within NASA's DPLR software. A complete description of the formulation of a fluid-coupled, ablative material response model is provided. Th...
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The present paper describes preliminary development of an in-depth material response model within NASA's DPLR software. A complete description of the formulation of a fluid-coupled, ablative material response model is provided. The model accounts for finite-rate chemistry and convection of pyrolysis gases through a porous medium. The material response equations are implemented in DPLR to produce a structured solver, which is implemented on a parallel computing platform. An independent coupling interface allows the new DPLR-ARM code to communicate boundary data with DPLR, enabling coupled simulations of fluid and material response. A preliminary demonstration of the feasibility of this approach is presented by analysis of two conduction-only problems: a titanium ballistic range projectile, and a carbon ablator subject to low heat flux in an arc jet. Results of the simulations show that proper characterization of material properties, such as surface emissivity, and the non-isotropy and temperature dependence of thermal conductivity, have a strong influence on the predicted temperature field. Additionally, it is found that the solution is sensitive to frequency of data exchange in the coupling procedure.
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