Gas turbine engine having improved core system

A method of cooling a combustion system of a gas turbine engine including a booster compressor having a plurality of stages, wherein said combustion system produces pulses of gas having increased pressure and temperature from a fluid flow provided thereto, comprising the following steps:

• (a) providing a first source of compressed air from said booster compressor to said combustion system;
• and
• (b) providing a second source of compressed air from said booster compressor to cool said combustion system;
• wherein pressure of compressed air from said second source is greater than pressure of compressed air from said first source by a predetermined amount.

A gas turbine engine comprising:

• (a) a compressor positioned at a forward end of said gas turbine engine having a plurality of stages, each said stage including a stationary compressor blade row and a rotatable blade row connected to a first drive shaft and interdigitated with said first compressor blade row;
• (b) a combustion system for producing pulses of gas having increased pressure and temperature from a fluid supplied to an inlet thereof;
• (c) a load connected to a second drive shaft;
• and,
• (d) a turbine downstream of and in flow communication with said combustion system for powering said first and second drive shafts;
• wherein a first source of compressed air from said compressor is provided to said combustion system inlet and a second source of compressed air from said compressor is provided to cool said combustion system.

A gas turbine engine having a longitudinal centerline therethrough, comprising:

• (a) a fan section at a forward end of said gas turbine engine including at least a first fan blade row connected to a drive shaft;
• (b) a booster compressor positioned downstream of said fan section including a plurality of stages, each said stage including a stationary compressor blade row and a rotating compressor Wade row connected to said drive shaft and interdigitated with said stationary compressor blade row;
• (c) a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flaw provided to an inlet thereof;
• and,
• (d) a low pressure turbine downstream of and in flow communication with said combustion system which powers said drive shaft;
• wherein a first source of compressed air from said booster compressor is provided to said combustion system inlet and a second source of compressed air from said booster compressor is provided at a forward end of said low pressure turbine so as to mitigate effects of said gas pulses thereon.