The optimization of the thermodynamic cycles of aero engines has always been in the main targets of engineering efforts for environmental and economic reasons. The aero engine thermodynamic cycle is significantly affected by the selection of the cycle maximum temperature, which at the same time, should be high enough to achieve increased cycle efficiency but also be always kept within turbine blades material temperature limits to ensure turbine blades endurance and integrity. For these reasons, cooling techniques are used where a part of the compressor discharge air is usually used as a cooling medium for the turbine blades. Since this part of air does not participate in the heat addition process inside the combustion chamber its amount should be carefully estimated in order to simultaneously provide cooling air protection for the turbine blades and avoid loss of potential turbine work. Towards this direction, numerical tools such as the one presented in this work are developed. The present work is focused on the development of a surrogate model network for the calculation of turbine blade cooling for aero engine applications. In the analysis the turbine blade is modelled as a heat exchange unit composed of a number of interconnected sub-units. In each sub-unit the effect of turbine blade local geometrical features, such as inner channels length and hydraulic diameter, on heat transfer and pressure losses was incorporated through the use of literature based correlations. The investigations were focused on typical turbine blade conditions for recuperative aero engine applications as presented in Salpingidou et al. (2017). For the cooling mass flow calculations the suggestions of Young and Wilcock (2002a and 2002b) and Wilcock et al. (2005) were taken into consideration. Furthermore, after the development of the surrogate model network, the effect of incorporating heat transfer enhancers on inner flow turbine blade geometry was numerically assessed for two heat transfer augmentation surfaces, such as the ones presented in the work of Alam and Kim (2018), using experimental data available in literature describing their heat transfer and pressure loss characteristics targeting the identification of configurations with promising effect on the aero engine.