Method and apparatus for turbine cooling

Claim 3

A method according to claim 1 wherein said compressor includes airfoils, and said stratifying step is effected by differently configuring aerodynamic profiles thereof from inner diameter to outer diameter.

Claim 4

A method according to claim 2 wherein said airfoils are rotor blade airfoils.

Claim 5

A method according to claim 3 wherein said airfoils are configured to effect said above average pressure near tips thereof and said below average pressure therebelow to roots thereof.

Claim 6

A gas turbine engine comprising: a compressor having a plurality of stages of rotor blades and stator vanes for compressing air, at least a last stage of said rotor blades having airfoils configured for stratifying at an exit of said compressor a pressure profile of said compressed air to provide a main flow having a below average pressure and a separate bleed flow having an above average pressure; a combustor disposed in flow communication with said compressor for receiving said main flow for mixing with fuel to generate hot combustion gases; a turbine disposed in flow communication with said combustor for receiving said combustion gases therefrom, and operatively joined to said compressor for powering said compressor; and a heat exchanger disposed in flow communication with said compressor for receiving and cooling said bleed flow, and disposed in flow communication with at least said turbine to provide cooled bleed flow thereto for cooling thereof.

Claim 7

A gas turbine engine comprising: a compressor having a plurality of stages of rotor blades and stator vanes for compressing air, at least a last stage of said rotor blades having airfoils configured for stratifying at an exit of said compressor a pressure profile of said compressed air to provide a main flow having a below average pressure and a separate bleed flow having an above average pressure; a combustor disposed in flow communication with said compressor for receiving said main flow for mixing with fuel to generate hot combustion gases; a turbine disposed in flow communication with said combustor for receiving said combustion gases therefrom, and operatively joined to said compressor for powering said compressor; and a heat exchanger disposed in flow communication with said compressor for receiving and cooling said bleed flow, and disposed in flow communication with at least said turbine to provide cooled bleed flow thereto for cooling thereof.

Claim 8

An engine according to claim 5 wherein said last stage blade airfoils have different aerodynamic profiles from root-to-tip to effect said main and bleed flows.

Claim 9

An engine according to claim 6 wherein said bleed flow is effected adjacent to one of said airfoil roots and tips over a radial extent of about 10% to about 15% of said compressor exit, and said main flow is effected in a remaining radial extent thereof.

Claim 10

An airfoil according to claim 7 wherein said above average pressure is up to about 6%.

Claim 11

An engine according to claim 7 wherein said bleed flow is effected adjacent to said airfoil tips, and said airfoils each further includes a mid-span section having a camber less than camber of said airfoil tip and root thereof.

Claim 12

An engine according to claim 7 further comprising a fan disposed upstream from said compressor and having a bypass duct extending downstream therefrom, and wherein said heat exchanger is an air-to-air heat exchanger disposed in said bypass duct for receiving fan bypass air to cool said bleed flow.