Handbook of Turbomachinery Second Edition Revised - Ventech!

Existing challenges associated with cooling of the nozzle end walls and
combustor liner exit transition walls require close interaction between
combustor and turbine section designers. Maintaining a thin boundary layer
at the combustor exit by constantly converging the liner walls toward the
nozzles, and then converging the nozzle end walls via contouring, help
reduce secondary flow losses originated at the nozzle end wall. An earlier
study [55] has shown that introduction of end-wall film cooling upstream of
the leading edge in combination with end-wall contouring suppresses the
formation of horseshoe vortex at the nozzle leading edge, providing marked
reduction in aerodynamic losses and preventing the cooler film layer from
being diverted from the end wall. The studies also show that higher film
blowing ratios utilizing maximum available pressure head, equivalent to the
pressure drop through combustor, lead to significant improvement in the
end-wall film coverage.
As a result of the flatter radial temperature profile mentioned earlier,
turbine component cooling design strategy has to change beginning with the
stage 1 nozzle, which typically consumes nearly 50% of the total turbine
cooling budget in high-temperature engines. Tightening emission control
requirements justify efforts toward reduction of nozzle-cooling flows by
using spent air from combustor liner cooling and also by applying advanced
thermal barrier coatings in combination with nozzle internal convective
cooling, instead of full-coverage film cooling. These options represent a
tradeoff between product cost and engine efficiency to satisfy emission and
life requirements. Reduced amount of the turbine nozzle-cooling flow also
assists in reducing combustor exit temperature for a fixed rotor inlet
temperature.
Similar considerations regarding emission, flatter radial temperature
profiles, and product cost must be applied during the selection of a bladecooling
design. Recent advances in internal blade cooling, particularly for
highly thermally loaded leading edges of blades, including a technique based
on swirling flow [86, 140, 141], as well as increased confidence in durability
of blade thermal barrier coatings, improve the potential for a nonfilmcooled
blade leading-edge option for turbine inlet temperatures up to
2,370 8F (1,300 8C).
Afour-quadrants chart shown in Fig. 36 illustrates alogical sequence
for preliminary cross-disciplinary selection of the blade leading-edge cooling
[142]. Two applications are considered in parallel: one for 15,000 h of
operation, representing typical aero-engine blade life, and another for
60,000 h of operation, representing industrial engine blade life expectancy.
Both engines are assumed to have similar thermodynamic cycle, design and
materials. For both specified applications, quadrant I shows a correlation
between initially assumed blade stresses s and maximum allowable metal
Copyright © 2003 Marcel Dekker, Inc.