At present reusable cooling techniques for hypersonic vehicles utilize a passive heat shield as it is applied for the Space Shuttle for example. This technique does not allow to adjust the heat protection performance during flight and therefore the peak heat loads are dimensioning for the entire flight path. An alternative to such a system can be the film cooling technique which is already state of art for cooling of turbine blades. Thereby mostly air is led through discrete holes in the blade to establish a cooling film on the surface that reduces the penetrating heat loads. Most of the investigations in the past were done for quasi two dimensional turbulent flow at low Mach numbers. An overview of more recent investigations about three dimensional film cooling is given in Ref. [1]. Beside of this the active cooling technique has also been investigated in some works concerning hypersonic flows [2]. Thereby porous materials were used to inject the cooling gas through the surface and therefore a cooling effect could already be achieved in the wall. This so called transpiration cooling technique is also applied for cooling of rocket combustion chambers. All of these studies show that an active cooling system is an effective method to lower the heat loads into a structure. This gives the motivation to investigate the feasibility of an active cooling in laminar hypersonic flow as it mainly appear during the critical reentry phase.


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    Title :

    Experimental and numerical investigation of film cooling in hypersonic flows


    Contributors:
    Heufer, K.A. (author) / Olivier, H. (author)


    Publication date :

    2009


    Size :

    6 Seiten, 7 Bilder, 8 Quellen




    Type of media :

    Conference paper


    Type of material :

    Print


    Language :

    English




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