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.
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
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.
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.
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
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.