Handbook of Turbomachinery Second Edition Revised - Ventech!

them. The larger gas-path divergence angle is associated with a reduced
number of stages. This should be taken into consideration regarding control
of the blade-tip clearances that becomes more difficult due to a greater effect
of relative transient axial displacements in a rotor–stator system.
In cases where the turbine rotor speed is limited by the tip speed of the
larger diameter of the downstream stages, an option of a shorter stage 1
blades at larger tip diameter (retaining the equivalent gas-path area) can be
more beneficial for the turbine thermodynamic efficiency. This option
typically requires a higher blade count and larger cooling flow for the stage
blade–disc system with increased disc pumping, which might negatively
affect gains in the turbine thermal efficiency. The larger blade-tip diameter
also results in a larger and usually heavier and more costly engine package.
Optimization between these options has to involve a cross-disciplinary team
that primarily addresses aerodynamic, stress, heat-transfer, and cost factors.
The optimized blade height and tip diameter determine the height of stage 1
nozzle exit, consequently strongly affecting the height of the nozzle leading
edge and the shape of combustor liner exit transition. Establishment of
correlations between these various influencing factors and development of
physically and/or statistically proven algorithms are necessary steps toward
an optimized turbine multidisciplinary design system. A number of
multivariable design optimization tools using deterministic or stochastic
approaches (Goel et al.; Egorov et al.; Tappeta et al.) [137–139] have been
developed and applied in certain areas of gas turbine engine design.
However, development of algorithms that may be unique for each turbine
manufacturer due to the differences in design criteria for each company
continues to be a major obstacle in applying these optimizers.
A proper combustor design cannot be performed in isolation either.
Higher compressed air-to-fuel flow ratios required to control nitrogen oxide
emission by lowering the flame temperature demand a larger amount of
compressor discharge air to be introduced in the combustor primary zone.
This limits the cooling air budget available for the combustor liner and for
the turbine components, particularly when the line-cooling air circuit is in
parallel (not in series) with the primary air for fuel injectors. Back-side liner
convective cooling methods are becoming more preferred to avoid relatively
cold carbon oxide formation zones resulting from a more traditional filmcooling
method. Particularly interesting for this application are the
techniques based on a low-pressure-drop convective back-side cooling,
allowing the use of the spent liner cooling air in series with the primary
combustion air. With growing demands for air in the primary combustion
zone, to control nitrogen oxide emission, the amount of air available for
dilution decreases. This leads to a flatter combustor exit radial temperature
profile and results in higher gas temperature near the end walls.
Copyright © 2003 Marcel Dekker, Inc.