Method and apparatus for cooling an airfoil

An airfoil for use in a gas turbine engine, wherein said airfoil is coolable by cooling air and is operable under operating conditions within said gas turbine engine, said coolable airfoil comprising:

• an internal cavity;
• an external wall;
• at least one cooling air passage disposed in said external wall, said passage having a plurality of segments connected in series by at least one turn, wherein at least one of said passage segments includes an inlet aperture connecting said passage to said internal cavity, and another of said passage segments includes an exit aperture connecting said passage to a region outside said airfoil;
• and
• wherein each said passage segment has a length, and said length is limited such that and at least fifty percent of said length is subject to a cooling air velocity profile which includes entrance effects when said airfoil is operated under said operating conditions.

A coolable wall, comprising:

• a first external surface;
• a second external surface;
• and
• at least one cooling air passage disposed in said wall between said first and second external surfaces, said passage having a plurality of segments connected in series by at least one turn, wherein each said passage segment has a length over diameter ratio equal to or less than 20;
• wherein a single one of said passage segments includes an inlet aperture extending between said passage and said first external surface at an upstream end of said passage, and a single other of said passage segments includes an exit aperture extending between said passage and said second external surface at a downstream end of said passage;
• and
• wherein cooling air can enter said passage through said inlet aperture and exit said passage through said exit aperture.

A wall for use in an apparatus within a gas turbine engine, wherein said wall is coolable by cooling air and said apparatus is operable under operating conditions within said gas turbine engine, said coolable wall comprising:

• an internal surface exposed to said cooling air;
• an external surface exposed to core gas;
• and
• at least one cooling air passage disposed in said wall between said internal and external surfaces, said passage having a plurality of segments connected in series by at least one turn, wherein one of said passage segments includes an inlet aperture extending between said passage and said internal surface, and another of said passage segments includes an exit aperture extending between said passage and said external surface;
• and
• wherein each said passage segment has a length, and said length is limited such that and at least fifty percent of said length is subject to a cooling air velocity profile which includes entrance effects when said airfoil is operated under said operating conditions.

A method for cooling a wall within a gas turbine engine, comprising the steps of:

• providing a wall having an first surface and a second surface, wherein a source of cooling air is contiguous with said first surface and a source of core gas is contiguous with said second surface;
• providing a set of operating conditions for said gas turbine engine;
• providing a passage disposed within said wall between said first and second surfaces, said passage including a plurality of segments connected to one another by at least one turn, wherein an inlet aperture extends between one of said segments and said first surface, and an exit aperture extends between another of said segments and said second surface, and wherein each of said segments has a length;
• sizing said length of each said segment such that under said operating conditions cooling air passing through any of said passage segments will have a velocity profile with entrance region effect characteristics for at least fifty percent of said length.

A method for cooling a wall within a gas turbine engine, comprising the steps of:

• providing a wall having an first surface and a second surface, wherein a source of cooling air is contiguous with said first surface and a source of core gas is contiguous with said second surface;
• providing a set of operating conditions for said gas turbine engine;
• providing a passage disposed within said wall between said first and second surfaces, said passage including a plurality of segments connected to one another by at least one turn, wherein an inlet aperture extends between one of said segments and said first surface, and an exit aperture extends between another of said segments and said second surface, and wherein each of said segments has a length;
• sizing said lengths of said passage segments such that all of said passage segments have a length over diameter ratio equal to or less than 20.

A method for cooling a wall within a gas turbine engine, said method comprising the steps of:

• providing a wall having a first surface and a second surface, wherein cooling air is contiguous with said first surface and core gas is contiguous with said second surface;
• providing a plurality of passages within said wall, each said passage including a plurality of segments connected to one another by at least one turn, wherein a first aperture extends between one of said segments and said first surface, and a second aperture extends between another of said segments and said second surface;
• determining an expected thermal load under a predetermined set of operating conditions in each of a plurality of regions along said wall;
• selectively tuning each said passage to provide a particular amount of heat transfer performance for said set of operating conditions;
• and
• positioning said passages in said regions such that said heat transfer performance of said passages substantially equals said expected thermal load in said region.

A method for cooling a wall within a gas turbine engine, comprising the steps of:

• providing a wall having an first surface and a second surface, wherein a source of cooling air is contiguous with said first surface and a source of core gas is contiguous with said second surface;
• providing a passage disposed within said wall between said first and second surfaces, said passage including a plurality of segments connected to one another in series by at least one turn, wherein an inlet aperture extends between one of said segments and said first surface, and an exit aperture extends between another of said segments and said second surface;
• providing a set of operating conditions for said gas turbine engine, said operating conditions including a pressure difference across said wall, and a core gas pressure value adjacent said exit aperture;
• determining a desired difference in pressure across said exit aperture;
• determining a cooling gas pressure inside said passage adjacent said exit aperture using said desired difference in pressure across said exit aperture;
• determining a desired pressure difference across said plurality of segments using said pressure difference across said wall and said cooling gas pressure inside said passage adjacent said exit aperture;
• and
• sizing said inlet aperture to provide said desired pressure difference across said plurality of segments.