Abstract
The atmospheric hypersonic flight of sub-orbital and space vehicles generates aerodynamic heating and high wall heat fluxes, inducing high temperatures on the vehicle's structures and affecting their mechanical behavior, besides degrading the operation of board equipment. Furthermore, since payload preservation is always mandatory, the use of efficient Thermal Protection Systems (TPS) is a key-requirement for any spacecraft design. As an outcome, designing the TPS is a critical aspect of any rocket development program, since an undersized system may result in catastrophic failure, and an oversized one implies increased mass and cost. Sub-orbital platforms are a low-cost alternative for microgravity research. A sub-orbital platform (SARA) is being developed by Instituto de Aeronautica e Espaco (IAE) for such an application, and its current design uses a conventional layer of cork as TPS to protect its lateral surface, with the trade-off of large mass. Alternatively, a Thermally Integrated Structural Sandwich Core (TISSC), which consists of a structural sandwich panel in a three-layer plate with two face sheets and the core, presents advantages such as lightweight, low maintenance, insulation as well as load bearing capabilities, and low life-cycle cost. In this work, a TISSC is proposed to replace SARA's current TPS. The main contribution of the presented methodology is to couple the aerodynamic heating, heat transfer in porous insulation and thermo-structural analyses of the proposed configuration in order to evaluate the TISSC TPS perfor - mance. The results are compared with those obtained for the current SARA TPS design, showing improvements in thermal insulation and structural strength, as well as a remarkable mass reduction.
Highlights
Space vehicles reach very high velocities during atmospheric flight, inducing aerodynamic heating (Anderson, 1989) due to a supersonic shock wave in the vicinity of the vehicle’s wall and due to friction against the air molecules
A careful design of an efficient thermal protection system is paramount to the spacecraft’s mission completion, since an undersized system may result in catastrophic failure, and an oversized one implies increased mass and cost (Moraes Jr., 1998)
This paper aims at developing a methodology for modeling the thermo-structural behavior of structural thermal protection systems and the application of such systems
Summary
Space vehicles reach very high velocities during atmospheric flight, inducing aerodynamic heating (Anderson, 1989) due to a supersonic shock wave in the vicinity of the vehicle’s wall and due to friction against the air molecules. In the regions where heat loads and air temperature are lower, the surface warming is less critical, and a thermal insulation layer can be employed Such thermal protection should be able to prevent or reduce the two main heat transfer modes involved, convection and radiation. For the generation of recoverable space vehicles, the use of metallic thermal protection systems has been considered It comprises a high-temperature resistant metallic alloy integrated into the spacecraft structure, filled with a light weight fibrous material. Aims at applying an integrated methodology for modeling the thermostructural behavior of TISSC, where the aerodynamic heating, heat transfer in porous insulation and thermo-structural analyses are coupled in sequence, which allows to obtain the temperature and the stress distribution for the whole trajectory of a vehicle, for particular instants or conditions like in the previously mentioned studies. By taking the limit of Eq 7 as R → 0, the following expression is obtained: 0.332(ρ*e μ*e) 2
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