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Conjugate Heat Transfer in Participating Media
Published in John R. Howell, M. Pinar Mengüç, Kyle Daun, Robert Siegel, Thermal Radiation Heat Transfer, 2020
John R. Howell, M. Pinar Mengüç, Kyle Daun, Robert Siegel
An application that incorporates all of these difficulties is high-speed atmospheric re-entry of space vehicles. Here, the shock layer incorporates non-LTE effects near the shock, hypersonic flow, highly spectral line radiation emission and absorption (and possibly scattering in the boundary layer near the ablating heat shield), and strongly coupled chemical reactions. At velocities expected for lunar mission returns to Earth, radiation and convection to the spacecraft thermal protection system are roughly equal. For a Mars mission return to Earth, it is expected that over 90% of the heat flux incident on an ablating heat shield will be from radiation. The flux levels in this case are predicted to be so high that a new generation of heat shield materials is required (Section 19.4.2).
Understanding Your Accident Model
Published in Sidney Dekker, The Field Guide to Understanding ‘Human Error’, 2017
A barrier model can faithfully explain the last few minutes (or seconds, or perhaps hours, depending on the time constants in your domain) before an accident. One reason for this is that it might make sense, in many domains, to see risk during those final minutes or seconds as energy that is not contained (a drug with a high therapeutic index, for example, or two aircraft that come too close to each other in the sky). The “physical cause” of the loss of the Space Shuttle Columbia in February 2003 was “a breach in the Thermal Protection System on the leading edgeof the left wing. The breach was initiated by a piece of insulating foam that separated from the left bipod ramp of the External Tank and struck the wing in the vicinity of the lower half of Reinforced Carbon-Carbon panel 8 at 81.9 seconds after launch. During re-entry, this breach in the Thermal Protection System allowed superheated air to penetrate the leading-edge insulation and progressively melt the aluminum structure of the left wing, resulting in a weakening of the structure until increasing aerodynamic forces caused loss of control, failure of the wing, and breakup of the Orbiter.”5
What is Your Accident Model?
Published in Sidney Dekker, The Field Guide to Understanding Human Error, 2017
The “physical cause” of the loss of the Space Shuttle Columbia in February 2003 was “a breach in the Thermal Protection System on the leading edge of the left wing. The breach was initiated by a piece of insulating foam that separated from the left bipod ramp of the External Tank and struck the wing in the vicinity of the lower half of Reinforced Carbon-Carbon panel 8 at 81.9 seconds after launch. During re-entry, this breach in the Thermal Protection System allowed superheated air to penetrate the leading-edge insulation and progressively melt the aluminum structure of the left wing, resulting in a weakening of the structure until increasing aerodynamic forces caused loss of control, failure of the wing, and breakup of the Orbiter.”4
Effect of Thermal Ablation at the Fluid-Solid Interface of a Hypersonic Reentry Vehicle in Rarefied Flow Regime
Published in International Journal of Computational Fluid Dynamics, 2021
In a reentry vehicle, the heat flux impinging on a vehicle surface is high enough to cause thermo-structural failure. Thermal Protection System (TPS) protects internal structure, payload, probes and aerodynamic surface by shielding them from such high heat fluxes during a hypersonic flight. TPS is a single point of failure optimally designed to trade-off between weight, cost considerations and vehicle safety. Generally, TPS can be classified into two categories: non-ablative (reusable) material, which does not undergo any significant change in the mass/surface profile, such as the one used in the tiles of the space shuttle, and ablative (non-reusable) material, which accommodates heat load by sacrificing its mass through the processes of phase change, such as the one used in space capsules during a hypersonic flight. The idea of using an ablative TPS material is based on the assumption that heat used to ablate the layers of a material surface is prevented from penetrating further into the material, and therefore does not allow the internal material temperature to rise significantly. Conducting experiments in a reentry-like environment has practical limitations. The need for material thermal response (MTR) solver is thus essential in studying problems like TPS sizing and aerothermodynamic analysis of reentry vehicles, where the thermal solver is coupled with a non-equilibrium flow solver to study reentry problems realistically.