The overall concept remains unchanged
Even though heat shielding materials have evolved into higher performance ablative materials and reusable shielding materials, the overall concept of the re-entry heat shield has never fundamentally changed from the original blunt cone design. Today, the size of the decelerator and associated heat shield is still limited to the internal diameter of the launcher fairing.
Several game-changing solutions have since emerged
Major game-changing solutions have since emerged, such as ballute, inflatable shapes, deployable devices to increase the re-entry section to reduce aerothermal stresses (with reduction of the ballistic coefficient) or re-entry mass (with iso-ballistic coefficient). The inflatable heat shield has several advantages over existing designs. Not only can it be folded into a very small package, but it is also lightweight and cost-effective.
Several game-changing solutions have since emerged
From a thermal point of view, the main environmental constraint is constituted by the aerodynamic heat flux applied to the protective layers of the Flexible Thermal Protection System. The protective layers (first layer) must withstand a potential temperature varying from 1500 to 1800 °C for a low ballistic coefficient. The temperature of the insulation layers (second layer) should not exceed its operating temperature (i.e., in the temperature range from 600 to 1000 °C approximately). The temperature of the gas barrier (third layer) should not exceed 200 °C to protect critical inflatable systems located in close contact just below this layer.
Results
Benefits: Known approach.
Limitations: Limited to the internal diameter of the launcher's fairing, heavyweight, does not allow the approach to landing zones in the Martian highlands, limits the landed mass in the Martian lowlands.
Why it matters.
Making standard Flexible Thermal Protection System will reduce both the mass budget of re-entry vehicles and their development time.
We’re thinking:
We will model the Flexible Thermal Protection System as a solid element using the Finite Element Method (FEM) to automatically calculate the conduction through the three layers of material and the radiation, and fully parametrise all the geometrical and material properties such as construction points, thermal conductivities, thermo-optical properties, densities, etc. in a way that the same Flexible Thermal Protection System thermal model can be quickly reused and adapted for another mission.
Furthermore, We will validate our Flexible Thermal Protection System thermal model with the test and flight data and make it accessible to the entire space thermal engineers’ community via our Digital Engineer® Try it now
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