Frequently, the mechanical integrity of gas turbine components is designed for a hot day, sea level take-off, where the maximum values are encountered for critical temperatures, such as the ones at the compressor and combustor outlet and the turbine rotor inlet stations. Turbine cooling flow rates are then defined taking into consideration maximum allowable metal temperatures, stresses, component life expectancy and heat transfer technology. Remaining unchanged as a percent of the core engine mass flow through the rest of the flight envelope, excessive cooling mass flows are actually being used during the cruise and the descent segment, since these operating points are characterized by significantly reduced temperatures. The main objective of the current work is the preliminary evaluation of the performance benefits, which can be achieved during a long range civil flight when decreasing the cooling bleed fraction during cruise. This is considered an essential step before any study concerning the consequences upon lifting is conducted. A conventional engine is optimized to meet the respective flight requirements, operating under constant cooling fraction throughout the mission. Reduction in cooling mass flow is applied, changing in such a way its off-design performance. Changes in typical engine parameters are identified and are graphically presented versus bleed flow reduction. Moreover, making use of a model providing for the drag polar of an airframe, while taking into account of the continuous weight reduction due to fuel burn, the variation of fuel consumption during cruise is also calculated. Fuel benefits are identified; a 40% reduction of the cooling fraction results in cruise fuel dropping by 0.75%. This can be justified on the basis of decreasing the cooling of the mainstream and increasing the mass flow, which is expanded through the turbine stages upstream. Although a metal temperature increase is also expected, it is accompanied by a Combustor Outlet and Turbine Entry temperature reduction.
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