This paper presents the preliminary results of using artificial neural networks in the prediction of gas side convective heat transfer coefficients on a high pressure turbine blade. The artificial neural network approach which has three hidden layers was developed and trained by nine inputs and it generates one output. Input and output data were taken from an experimental research program performed at the von Karman Institute for Fluid Dynamics by Camci and Arts [5,6] and Camci [7]. Inlet total pressure, inlet total temperature, inlet turbulence intensity, inlet and exit Mach numbers, blade wall temperature, incidence angle, specific location of measurement and suction/pressure side specification of the blade were used as input parameters and calculated heat transfer coefficient around a rotor blade used as output. After the network is trained with experimental data, heat transfer coefficients are interpolated for similar experimental conditions and compared with both experimental measurements and CFD solutions. CFD analysis was carried out to validate the algorithm and to determine heat transfer coefficients for a closely related test case. Good agreement was obtained between CFD results and neural network predictions.

This manuscript discusses the numerical (finite element) and analytical modelling of structural interactions between gas turbine components in case of excessive axial movement and overspeed. Excessive axial movement, which may occur after a shaft failure, results in contact between rotating and static turbine components under high forces. These forces create friction which can act as a counter torque, potentially retarding the ‘free-rotating’ components. The study is based on a shaft failure scenario of a ‘three-shaft’, high ‘bypass’ ratio, civil ‘large-fan’ engine. Coupled analytical performance and friction methods are used as stand-alone tools to investigate the effect of rubbing between rotating and stationary components. The method is supported by ‘high-fidelity’, ‘three-dimensional’, thermomechanical finite element simulations using LS-DYNA software. The novelty of the work reported herein lies in the development of a generalised modelling approach that can produce useful engine design guidelines to minimise the terminal speed of a free running turbine after an unlocated shaft failure. The study demonstrates the advantage of using a fast analytical formulation in a design space exploration, after verifying the analytical model against finite element simulation results. The radius and the area of a stationary seal platform in the turbine assembly are changed systematically and the design space is explored in terms of turbine acceleration, turbine dislocation rate and stationary component mass. The radius of the friction interface increases due to the increasing radius of the nozzle guide vane flow path and stationary seal platform. This increases the frictional torque generated at the interface. It was found that if the axial dislocation rate of the free running turbine wheel is high, the resulting friction torque becomes more effective as an overspeed prevention mechanism. Reduced contact area results in a higher axial dislocation rate and this condition leads to a design compromise between available friction capacity, during shaft failure contact and seal platform structural integrity.

The geared turbofan engines bring the potential to rotate the fan at lower speed and allow an increase in diameter, which in turn leads to an increase in propulsive efficiency through high by-pass ratio. The low-pressure turbine stages driving the fan can also rotate at high speed resulting in fewer stages when compared to traditional turbofans. However, when operating at high speed, pressure fluctuations due to self-excited vibrations increase and may provoke flutter instabilities. In a geared architecture, to deliver the high power required by the fan and the intermediate-pressure compressor, the low-pressure turbine system operates at higher temperatures compared to its predecessors. This phenomenon requires structural materials with higher heat resistance, which carries the inconvenience of poor welding suitability. That is the reason why alternative non-welded blade shroud joint techniques are so important, techniques as the blade interlock mechanism studied in this work. This manuscript examines the effects of different design parameters of a low-pressure turbine blade shroud interlock on flutter stability, to make future recommendations for geared engines. The shrouded turbine rotor blades feature blade interlocks, which enhances the dynamic stability by providing stiffness to the rotor blade row. To assess the stability of the system, a parametric design of a turbine blade-disk assembly was prepared. In the parametric model the design variables that define the blade interlock are the interlock angle, interlock axial position, interlock contact length and height, knife seal position and pre-twist angle. After parametrization, a finite element model of the turbine blade and disk assembly was prepared with cyclic symmetry boundary condition. The stresses caused by rotation were calculated in a static structural analysis and these were used as pre-stress boundary conditions in modal analysis. The modal results were afterwards exchanged with an aerodynamic model to obtain the aerodynamic damping for different blade interlock design configurations. In the present work, the dynamic response of the first three excitation modes was analyzed. It was found that the third mode was stable for all the design points, whereas first and second modes were unstable at least for the reference design point. Among the considered six different parameters that define the blade interlock geometry, the interlock contact position turned to be the most influential parameter for modal response and for flutter stability. Moving the interlock contact position towards the trailing edge gave the most beneficial results. On the other hand, the interlock angle showed the least influence on both, the modal analysis and flutter behavior. The accomplished Design of Experiments and subsequent optimization process also conclude that there exists an interdependency between the studied parameters.

This manuscript discusses the operation of an electrically driven fan for a hybrid-electric propulsion system for BAe-146 aircraft. The thrust requirement is fed into an integrated cycle and aerodynamic design tool for the sizing of a ducted fan as one of the main propulsors, podded under the wing as a replacement for a turbofan engine. The electric motor design is initiated with the torque and speed requirements and with the dimensional constraints arising from the driven fan geometry. The fan operation and aerodynamic design are derived by changing the fan pressure ratio and hub-to-tip ratio to obtain a 2-D design space. Surface-mounted permanent magnet electric motor designs are mapped on the 2-D fan design space. The design and operational flexibility of the system is investigated through three scenarios. In the first scenario, the aircraft rate of climb is changed to downsize the electric motor. In the second scenario, the electric motor rated frequency is changed to increase the power density and in the third scenario the electric motor current density is changed for the same purpose. The investigated three scenarios provide design and operational guidelines for reducing the weight of the electric motor for a direct drive application.

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