This paper represents further development of Modified Modal Domain Analysis (MMDA) (Sinha, 2009), which is a breakthrough method for the reduced-order modeling of a bladed rotor with geometric mistuning. The bases vectors for model reduction in MMDA have been formed using the mode shapes of cyclic sectors with blades’ geometries perturbed along the POD (Proper Orthogonal Decomposition) features. The use of mode shapes from modal analyses of cyclic sectors perturbed along the POD features adds an additional step of creating the finite element models of artificially perturbed geometries. Here, an alternative formulation of MMDA is presented in which bases vectors are created from cyclic sectors with actual blades. Therefore, the additional step of creating artificial blades with geometries perturbed along POD features is avoided. The MMDA is also extended to a bladed rotor in which a few blades have extremely large mistuning; for example, blended airfoils. The validity of proposed approaches is shown by comparing with ANSYS results for full (360 degree) bladed rotor.

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A three-dimensional (3D) blade design method for an axial compressor transonic stage to optimize the aerodynamic performance is presented in this paper. The blade is defined by three profiles and a radial stacking line. Each profile is a multi circular arc (MCA), and the stacking line is defined as a B-spline curve with six design parameters. To ensure the off-design performance, a multi-objective genetic algorithm (MOGA) is applied. The objective functions are the efficiency, shock position and leading edge pressure difference at the design point. Because shock position and leading edge pressure difference can evaluate the potential for stalling, the method can generate blades with a wide operating range with just one performance prediction. This method is applied in a transonic blade design. The result shows that the efficiency of the optimized blade at the design point is increased and the operation range is expanded compared with the original blade.

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Exhaust nozzle research was conducted to develop design technologies for propulsion system of Hypersonic transport. According to engine cycle study, nozzle cross sectional area should be variable for nozzle pressure ratio up to 270 level. Trust coefficient should be 0.95 to meet 2.0kg/hr/kg of SFC at Mach 5. And nozzle liner have to endure 1900C level gas temperature. So exhaust nozzle to have two dimensional (2-D) variable geometry and cooling structure. Aerodynamic study was conducted by model test and CFD. As result, target thrust coefficient was achieved. Variable geometry mechanism is designed to have convergent/divergent flaps and side walls. Convergent/divergent flaps are separated to introduce ambient air from throat into inside of nozzle at take-off. Integrated film and impingement cooling structure is designed using composite for liner based on heat transfer coefficient distribution and cooling efficiency, acquired by model test. This paper describes 2-D Variable Exhaust Nozzle Research.

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Flexible Fuel is used for land-based power generation gas turbine to meet the environmental requirements, which bring high temperature gases with variable composition to the turbine downstream. in this paper, a high-order and high-resolution in-house CFD code (named START) was developed to capture more features of the real multi-species flow in the turbine. The code was validated with four samples, ranging from subsonic to transonic and supersonic flows. The validation showed that the code had high accuracy and fidelity to simulate the flows, which included both shock wave and species diffusion. Then the solver was used to simulate the unsteady multi-species flow in a 1.5-stage turbine, which included CO2 coolant injection at the 1st stator trailing edge. The cooling effect on the rotor downstream and associated losses were studied in terms of the coolant-to-free stream velocity ratio.

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Hot streaks can cause local hot spots on the blade surfaces of high-pressure turbine stages, resulting in locally higher thermal loads. These local loads represent a potential source of blade life reduction and blade failure. The blade regions exposed to higher thermal loads are determined by the effect of the unsteady blade row interaction on the migration path of the hot streaks. In order to improve understanding of these effects an experimental study on the effect of shaping the inlet temperature distortion has been undertaken.1 The experimental investigations have been performed in the axial turbine facility “LISA” at ETH Zurich. The test configuration consists of a one-and-1/2 stage, unshrouded, highly loaded axial turbine with a hot streak generator placed upstream of the first vane row. The latter is designed to provide different shapes of the inlet temperature distortion, as well as different circumferential and spanwise positions. The steady and unsteady aerodynamic effects are measured respectively with pneumatic probes and the in-house developed Fast Response Aerodynamic Probe (FRAP) technology. The unsteady thermodynamic effects are measured in a time resolved manner with the in-house developed Fast Response Entropy Probe (FENT). The time resolved measurements are made in planes at the inlet to the first vane row as well as downstream of it and downstream of the rotor. The current paper presents the results of the first shaped hot streak injection and analyzes the mechanisms involved in the convection and the migration of the hot streak through the bladed rows. The effect of the first stationary blade row on the path of the hot streak is explained by an analysis of the flow field and temperature field at the exit of the first nozzle guide vane row. Mixing and heat conduction as well as the unsteady effect of the downstream rotor cause the total temperature distortion to diminish thus generating a more uniform distribution. The effect of the rotating blade row is shown with the flow field and the temperature field at the exit of the rotor. The measurements reveal a radial migration of the hot streak which is confined in circumferential direction by the pressure side of the rotor wake causing the fluid to partially go into the tip leakage vortex. Furthermore, at the suction side of the rotor blade the hot gases are confined in between the passage vortices of the row. The root mean square of the unsteady pressure signal acquired can be used for tracing the mixing process and losses showing the interaction of the hot streak with the secondary flow structures.

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