WO2016133488A1 - Turbine airfoil cooling system with film cooling hole within protruded cooling hole support - Google Patents

Turbine airfoil cooling system with film cooling hole within protruded cooling hole support Download PDF

Info

Publication number
WO2016133488A1
WO2016133488A1 PCT/US2015/016004 US2015016004W WO2016133488A1 WO 2016133488 A1 WO2016133488 A1 WO 2016133488A1 US 2015016004 W US2015016004 W US 2015016004W WO 2016133488 A1 WO2016133488 A1 WO 2016133488A1
Authority
WO
WIPO (PCT)
Prior art keywords
wall
film cooling
cooling hole
elevated
sectional area
Prior art date
Application number
PCT/US2015/016004
Other languages
French (fr)
Inventor
Gary B. Merrill
Jr. Jonathan E. SHIPPER
Original Assignee
Siemens Aktiengesellschaft
Siemens Energy, Inc.
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Priority to PCT/US2015/016004 priority Critical patent/WO2016133488A1/en
Publication of WO2016133488A1 publication Critical patent/WO2016133488A1/en

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/90—Coating; Surface treatment

Definitions

  • This invention is directed generally to turbine airfoils, and more particularly to cooling systems in hollow turbine airfoils of turbine engines.
  • gas turbine engines typically include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power.
  • Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit.
  • Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures.
  • turbine blades must be made of materials capable of withstanding such high temperatures.
  • turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
  • Cooling systems often include a plurality of film cooling holes in the outer wall of an airfoil to exhaust cooling fluids to form a film of cooling air to protect and cool the outer wall at the outer surface. When cooling air is ejected through a hole in an airfoil wall, it combines with a much higher temperature gas flow which creates high thermal stresses locally where thermal barrier coating
  • temperatures can vary substantially over very small distances. Such temperature differences leads to locally high thermal stresses and then typically to cracks leading to local damage around the cooling feature that can ultimately lead to spalled thermal barrier coating and reduced component life. Thus, an improved film cooling hole is desired.
  • a turbine airfoil for a gas turbine engine having a more resilient film cooling hole configuration is disclosed.
  • the film cooling hole may be positioned in a protruded cooling hole support in an outer wall of the airfoil and may have an inlet in communication with a cooling system and an outlet in an elevated film cooling exhaust surface that extends outwardly further than an adjacent outer surface of the outer wall.
  • a thermal barrier coating may be positioned on the outer surface of the outer wall, and the outer surface of the elevated film cooling exhaust surface may be positioned between the outer surface of the outer wall and an outer surface of the thermal barrier coating, thereby positioning the elevated film cooling exhaust surface above the outer surface of the outer wall of the airfoil.
  • the turbine airfoil may include a generally elongated hollow airfoil formed from an outer wall, and having a leading edge, a trailing edge, a pressure side, a suction side, and a cooling system positioned within interior aspects of the generally elongated hollow airfoil and formed from at least one cavity.
  • the turbine airfoil may also include one or more film cooling holes positioned in the protruded cooling hole support in the outer wall and having an inlet in communication with the cooling system and an outlet in an elevated film cooling exhaust surface on the protruded cooling hole support that extends outwardly further than an adjacent outer surface of the outer wall.
  • a thermal barrier coating may be on the outer surface of the outer wall.
  • the outer surface of the elevated film cooling exhaust surface may be positioned between the outer surface of the outer wall and an outer surface of the thermal barrier coating.
  • the outer surface of the elevated film cooling exhaust surface may be positioned between a distance between 30 percent and 50 percent of a distance of thickness of the thermal barrier coating from the outer surface of the outer wall to the outer surface of the thermal barrier coating.
  • the elevated film cooling exhaust surface may be supported sidewalls extending between a base at the outer surface of the outer wall and the elevated film cooling exhaust surface that are tapered from the base at the outer surface of the outer wall to the elevated film cooling exhaust surface.
  • the sidewalls extending between the base of the outer surface of the outer wall and the elevated film cooling exhaust surface may be tapered nonlinearly from the base at the outer surface of the outer wall to the elevated film cooling exhaust surface.
  • a cross-sectional area of the outer surface of the outer wall may be greater than a cross-sectional area of the outlet of the film cooling hole.
  • the outer surface of the outer wall may be at least fifty percent larger than a cross-sectional area of the outlet of the film cooling hole.
  • the material forming the film cooling hole and the elevated film cooling exhaust surface may also be defined by a cross-sectional area of a base at an intersection of sidewalls.
  • the outer surface of the outer wall may be larger than a cross-sectional area of the elevated film cooling exhaust surface.
  • the cross- sectional area of the base at the intersection of sidewalls and the outer surface of the outer wall may be greater than 150 percent of a cross-sectional area of the outlet of the film cooling hole.
  • protruded cooling hole support is that the protruded cooling hole support forming the film cooling hole may itself be cooled more effectively so that the protruded cooling hole support will not experience locally higher temperatures due to its raised location and thermal exposure.
  • Another advantage of the protruded cooling hole support is that the cooling fluid flowing through the protruded cooling hole support is marginally heated up thereby reducing the overall temperature difference between the inlet and the outlet.
  • protruded cooling hole support is that thermal stresses within the thermal barrier coating may be reduced because of the protruded cooling hole support forming the film cooling hole and the elevated film cooling exhaust surface being positioned within the thermal barrier coating.
  • the outlet of the film cooling hole may be configured to reduce local stress and bonding adjacent to the hole outlet.
  • Figure 1 is a perspective view of a turbine airfoil with an internal cooling system including film cooling holes in protruding cooling hole supports.
  • Figure 2 is a cross-sectional view of the turbine airfoil and internal cooling system taken along section line 2-2 in Figure 1 .
  • Figure 3 is a perspective view of a film cooling hole in a protruding cooling hole support.
  • a turbine airfoil 1 0 for a gas turbine engine having a more resilient film cooling hole 14 configuration supported by a protruded cooling hole support 8 is disclosed.
  • the film cooling hole 14 may be positioned in a protruded cooling hole support 8 in an outer wall 16 of the airfoil 10 and may have an inlet 18 in communication with a cooling system 20 and an outlet 22 in an elevated film cooling exhaust surface 24 that extends outwardly further than an adjacent outer surface 26 of the outer wall 16.
  • a thermal barrier coating 28 may be positioned on the outer surface 26 of the outer wall 16, and an outer surface 30 of the elevated film cooling exhaust surface 24 may be positioned between the outer surface 26 of the outer wall 16 and an outer surface 32 of the thermal barrier coating 28, thereby positioning the elevated film cooling exhaust surface 24 above the outer surface 26 of the outer wall 16 of the airfoil 10. Such position more effectively cools material forming the film cooling hole 14 and heats the cooling air flowing through the film cooling hole 14, thereby reducing thermal stress at the film cooling hole 14 and improving local stress and bonding adjacent to the outlet 22 of the film cooling hole 14.
  • the turbine airfoil 10 may be formed from a generally elongated hollow airfoil 34 formed from an outer wall 16 and having a leading edge 36, a trailing edge 38, a pressure side 40, a suction side 42, and a cooling system 20 positioned within interior aspects of the generally elongated hollow airfoil 34 and formed from at least one cavity 44.
  • the turbine airfoil 10 is not limited to any particular configuration, and may be a rotatable turbine blade, stationary turbine vane or other airfoil.
  • the turbine airfoil 10 may include one or more film cooling holes 14 positioned in the protruded cooling hole support 8 in the outer wall 16.
  • the film cooling hole 14 may include an inlet 18 in communication with the cooling system 20 and an outlet 22 in an elevated film cooling exhaust surface 24 that extends outwardly further than an adjacent outer surface 26 of the outer wall 16.
  • the turbine airfoil 10 may include a thermal barrier coating 28 on the outer surface 26 of the outer wall 16.
  • the thermal barrier coating 28 may be any appropriate coating such as a conventional thermal barrier coating or a coating yet to be conceived.
  • the outer surface 30 of the elevated film cooling exhaust surface 24 may be positioned between the outer surface 30 of the outer wall 16 and an outer surface 32 of the thermal barrier coating 28.
  • the outer surface 30 of the elevated film cooling exhaust surface 24 may be positioned between a distance between 30 percent and 50 percent of a distance of thickness of the thermal barrier coating 28 from the outer surface 26 of the outer wall 16 to the outer surface 32 of the thermal barrier coating 28.
  • the thermal barrier coating 28 may be between about 650 microns and about 1 ,000 microns.
  • the elevated film cooling exhaust surface 24 may be positioned between about 300 and 800 microns from the outer surface 26 of the outer wall 16.
  • Sidewalls 46 extending between a base 48 at the outer surface 26 of the outer wall 16 and the elevated film cooling exhaust surface 24 may be tapered from the base 48 at the outer surface 26 of the outer wall 16 to the elevated film cooling exhaust surface 24. There may exist a single sidewall 46 curved around the perimeter of the elevated film cooling exhaust surface 24 or a plurality of sidewalls 46. The sidewalls 46 increases the available bond surface are and bond
  • the sidewalls 46 extending between the base 48 of the outer surface 26 of the outer wall 16 and the elevated film cooling exhaust surface 24 may be tapered nonlinearly from the base 48 at the outer surface 26 of the outer wall 16 to the elevated film cooling exhaust surface 24.
  • a cross-sectional area of the outer surface 26 of the outer wall 16 may be greater than a cross-sectional area of the outlet 22 of the film cooling hole 14.
  • the cross-sectional area of the outer surface 26 of the outer wall 16 may be generally orthogonal to a longitudinal axis 52 of the film cooling hole 14.
  • the cross-sectional area of the outer surface 26 of the outer wall 16 may be at least fifty percent larger than a cross-sectional area of the outlet 22 of the film cooling hole 14.
  • a cross- sectional area of a base 48 at an intersection 50 of sidewalls 46 and the outer surface 26 of the outer wall 16 may be larger than a cross-sectional area of the elevated film cooling exhaust surface 24.
  • the cross-sectional area of the base 48 at an intersection 50 of sidewalls 46 and the outer surface 26 of the outer wall 16 may be generally orthogonal to a longitudinal axis 52 of the film cooling hole 14.
  • the cross-sectional area of the base 48 at an intersection 50 of sidewalls 46 and the outer surface 26 of the outer wall 16 may be greater than 150 percent of a cross- sectional area of the outlet 22 of the film cooling hole 14.
  • the film cooling hole 14 and elevated film cooling exhaust surface 24 may be formed via a casting process, such as an investment casting using a lost wax process. Because the configuration of the film cooling hole 14 and elevated film cooling exhaust surface 24 are relatively small, a shell insert may be used.
  • a ceramic core may be formed. The method may include placing the ceramic core within an inner cavity formed by an inner surface of a wax die and injecting wax into one or more openings formed between the ceramic core and the inner surface of the wax die. The method may also include removing the wax die to reveal a wax component, coating the wax component with a ceramic coating to form an external ceramic shell with a ceramic core positioned therein and removing the wax component within the ceramic coating leaving one or more cavities within the ceramic coating.
  • a shell insert may be attached to the external ceramic shell at each location of a film cooling hole.
  • the shell insert may include protruded core features to create an internal cast hole feature in the airfoil surface.
  • the method may also include filling the cavity within the ceramic coating with a molten metal and removing the ceramic shell and the ceramic core to form a cast component.
  • the protruded cooling hole support 8 forming the film cooling hole 14 may itself be cooled more effectively so the protruded cooling hole support 8 will not experience locally higher temperatures due to its raised location and thermal exposure.
  • the cooling fluid flowing through the surface feature is marginally heated up thereby reducing the overall temperature difference between the inlet 18 and outlet 22.
  • Thermal stresses within the thermal barrier coating 28 may be reduced because of the protruded cooling hole support 8 forming the film cooling hole 14 and the elevated film cooling exhaust surface 24 being positioned within the thermal barrier coating 28.
  • the outlet 22 of the film cooling hole 14 may be configured to reduce local stress and bonding adjacent to the hole outlet.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A turbine airfoil (10) for a gas turbine engine (12) having a more resilient film cooling hole (14) configuration is disclosed. The film cooling hole (14) may be positioned in an outer wall (16) of the airfoil (10) and may have an inlet (18) in communication with a cooling system (20) and an outlet (22) in an elevated film cooling exhaust surface (24) that extends outwardly further than an adjacent outer surface (26) of the outer wall (16). A thermal barrier coating (28) may be positioned on the outer wall (16), and the outer surface (30) of the elevated film cooling exhaust surface (24) may be positioned between the outer surface (26) of the outer wall (16) and an outer surface (32) of the thermal barrier coating (28), thereby positioning the elevated film cooling exhaust surface (24) above the outer surface (26) of the outer wall (16) of the airfoil (10). Such position reduces thermal stress at the film cooling hole (14) and improving local stress and bonding adjacent to the outlet (22).

Description

TURBINE AIRFOIL COOLING SYSTEM WITH FILM COOLING HOLE WITHIN PROTRUDED COOLING HOLE SUPPORT
FIELD OF THE INVENTION
This invention is directed generally to turbine airfoils, and more particularly to cooling systems in hollow turbine airfoils of turbine engines.
BACKGROUND
Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures. Cooling systems often include a plurality of film cooling holes in the outer wall of an airfoil to exhaust cooling fluids to form a film of cooling air to protect and cool the outer wall at the outer surface. When cooling air is ejected through a hole in an airfoil wall, it combines with a much higher temperature gas flow which creates high thermal stresses locally where thermal barrier coating
temperatures can vary substantially over very small distances. Such temperature differences leads to locally high thermal stresses and then typically to cracks leading to local damage around the cooling feature that can ultimately lead to spalled thermal barrier coating and reduced component life. Thus, an improved film cooling hole is desired.
SUMMARY OF THE INVENTION
A turbine airfoil for a gas turbine engine having a more resilient film cooling hole configuration is disclosed. The film cooling hole may be positioned in a protruded cooling hole support in an outer wall of the airfoil and may have an inlet in communication with a cooling system and an outlet in an elevated film cooling exhaust surface that extends outwardly further than an adjacent outer surface of the outer wall. A thermal barrier coating may be positioned on the outer surface of the outer wall, and the outer surface of the elevated film cooling exhaust surface may be positioned between the outer surface of the outer wall and an outer surface of the thermal barrier coating, thereby positioning the elevated film cooling exhaust surface above the outer surface of the outer wall of the airfoil. Such position more effectively cools material forming the film cooling hole and heats the cooling air flowing through the film cooling hole, thereby reducing thermal stress at the film cooling hole and improving local stress and bonding adjacent to the outlet of the film cooling hole.
In at least one embodiment, the turbine airfoil may include a generally elongated hollow airfoil formed from an outer wall, and having a leading edge, a trailing edge, a pressure side, a suction side, and a cooling system positioned within interior aspects of the generally elongated hollow airfoil and formed from at least one cavity. The turbine airfoil may also include one or more film cooling holes positioned in the protruded cooling hole support in the outer wall and having an inlet in communication with the cooling system and an outlet in an elevated film cooling exhaust surface on the protruded cooling hole support that extends outwardly further than an adjacent outer surface of the outer wall. A thermal barrier coating may be on the outer surface of the outer wall. The outer surface of the elevated film cooling exhaust surface may be positioned between the outer surface of the outer wall and an outer surface of the thermal barrier coating. The outer surface of the elevated film cooling exhaust surface may be positioned between a distance between 30 percent and 50 percent of a distance of thickness of the thermal barrier coating from the outer surface of the outer wall to the outer surface of the thermal barrier coating.
The elevated film cooling exhaust surface may be supported sidewalls extending between a base at the outer surface of the outer wall and the elevated film cooling exhaust surface that are tapered from the base at the outer surface of the outer wall to the elevated film cooling exhaust surface. In another embodiment, the sidewalls extending between the base of the outer surface of the outer wall and the elevated film cooling exhaust surface may be tapered nonlinearly from the base at the outer surface of the outer wall to the elevated film cooling exhaust surface. A cross-sectional area of the outer surface of the outer wall may be greater than a cross-sectional area of the outlet of the film cooling hole. In at least one
embodiment, the outer surface of the outer wall may be at least fifty percent larger than a cross-sectional area of the outlet of the film cooling hole. The material forming the film cooling hole and the elevated film cooling exhaust surface may also be defined by a cross-sectional area of a base at an intersection of sidewalls. The outer surface of the outer wall may be larger than a cross-sectional area of the elevated film cooling exhaust surface. In at least one embodiment, the cross- sectional area of the base at the intersection of sidewalls and the outer surface of the outer wall may be greater than 150 percent of a cross-sectional area of the outlet of the film cooling hole.
An advantage of the protruded cooling hole support is that the protruded cooling hole support forming the film cooling hole may itself be cooled more effectively so that the protruded cooling hole support will not experience locally higher temperatures due to its raised location and thermal exposure.
Another advantage of the protruded cooling hole support is that the cooling fluid flowing through the protruded cooling hole support is marginally heated up thereby reducing the overall temperature difference between the inlet and the outlet.
Yet another advantage of the protruded cooling hole support is that thermal stresses within the thermal barrier coating may be reduced because of the protruded cooling hole support forming the film cooling hole and the elevated film cooling exhaust surface being positioned within the thermal barrier coating. The outlet of the film cooling hole may be configured to reduce local stress and bonding adjacent to the hole outlet.
These and other embodiments are described in more detail below.
BRIEF DESCRIPTION OF THE DRAWINGS
The accompanying drawings, which are incorporated in and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention.
Figure 1 is a perspective view of a turbine airfoil with an internal cooling system including film cooling holes in protruding cooling hole supports. Figure 2 is a cross-sectional view of the turbine airfoil and internal cooling system taken along section line 2-2 in Figure 1 .
Figure 3 is a perspective view of a film cooling hole in a protruding cooling hole support.
DETAILED DESCRIPTION OF THE INVENTION
As shown in Figures 1-3, a turbine airfoil 1 0 for a gas turbine engine having a more resilient film cooling hole 14 configuration supported by a protruded cooling hole support 8 is disclosed. The film cooling hole 14 may be positioned in a protruded cooling hole support 8 in an outer wall 16 of the airfoil 10 and may have an inlet 18 in communication with a cooling system 20 and an outlet 22 in an elevated film cooling exhaust surface 24 that extends outwardly further than an adjacent outer surface 26 of the outer wall 16. A thermal barrier coating 28 may be positioned on the outer surface 26 of the outer wall 16, and an outer surface 30 of the elevated film cooling exhaust surface 24 may be positioned between the outer surface 26 of the outer wall 16 and an outer surface 32 of the thermal barrier coating 28, thereby positioning the elevated film cooling exhaust surface 24 above the outer surface 26 of the outer wall 16 of the airfoil 10. Such position more effectively cools material forming the film cooling hole 14 and heats the cooling air flowing through the film cooling hole 14, thereby reducing thermal stress at the film cooling hole 14 and improving local stress and bonding adjacent to the outlet 22 of the film cooling hole 14.
In at least one embodiment, as shown in Figures 1 and 2, the turbine airfoil 10 may be formed from a generally elongated hollow airfoil 34 formed from an outer wall 16 and having a leading edge 36, a trailing edge 38, a pressure side 40, a suction side 42, and a cooling system 20 positioned within interior aspects of the generally elongated hollow airfoil 34 and formed from at least one cavity 44. The turbine airfoil 10 is not limited to any particular configuration, and may be a rotatable turbine blade, stationary turbine vane or other airfoil. As shown in Figure 3, the turbine airfoil 10 may include one or more film cooling holes 14 positioned in the protruded cooling hole support 8 in the outer wall 16. The film cooling hole 14 may include an inlet 18 in communication with the cooling system 20 and an outlet 22 in an elevated film cooling exhaust surface 24 that extends outwardly further than an adjacent outer surface 26 of the outer wall 16.
The turbine airfoil 10 may include a thermal barrier coating 28 on the outer surface 26 of the outer wall 16. The thermal barrier coating 28 may be any appropriate coating such as a conventional thermal barrier coating or a coating yet to be conceived. The outer surface 30 of the elevated film cooling exhaust surface 24 may be positioned between the outer surface 30 of the outer wall 16 and an outer surface 32 of the thermal barrier coating 28. In at least one embodiment, the outer surface 30 of the elevated film cooling exhaust surface 24 may be positioned between a distance between 30 percent and 50 percent of a distance of thickness of the thermal barrier coating 28 from the outer surface 26 of the outer wall 16 to the outer surface 32 of the thermal barrier coating 28. In at least one embodiment, the thermal barrier coating 28 may be between about 650 microns and about 1 ,000 microns. The elevated film cooling exhaust surface 24 may be positioned between about 300 and 800 microns from the outer surface 26 of the outer wall 16.
Sidewalls 46 extending between a base 48 at the outer surface 26 of the outer wall 16 and the elevated film cooling exhaust surface 24 may be tapered from the base 48 at the outer surface 26 of the outer wall 16 to the elevated film cooling exhaust surface 24. There may exist a single sidewall 46 curved around the perimeter of the elevated film cooling exhaust surface 24 or a plurality of sidewalls 46. The sidewalls 46 increases the available bond surface are and bond
effectiveness through mechanical interlock to provide an improved local adhesion capability. The sidewalls 46 extending between the base 48 of the outer surface 26 of the outer wall 16 and the elevated film cooling exhaust surface 24 may be tapered nonlinearly from the base 48 at the outer surface 26 of the outer wall 16 to the elevated film cooling exhaust surface 24.
A cross-sectional area of the outer surface 26 of the outer wall 16 may be greater than a cross-sectional area of the outlet 22 of the film cooling hole 14. The cross-sectional area of the outer surface 26 of the outer wall 16 may be generally orthogonal to a longitudinal axis 52 of the film cooling hole 14. The cross-sectional area of the outer surface 26 of the outer wall 16 may be at least fifty percent larger than a cross-sectional area of the outlet 22 of the film cooling hole 14. A cross- sectional area of a base 48 at an intersection 50 of sidewalls 46 and the outer surface 26 of the outer wall 16 may be larger than a cross-sectional area of the elevated film cooling exhaust surface 24. The cross-sectional area of the base 48 at an intersection 50 of sidewalls 46 and the outer surface 26 of the outer wall 16 may be generally orthogonal to a longitudinal axis 52 of the film cooling hole 14. The cross-sectional area of the base 48 at an intersection 50 of sidewalls 46 and the outer surface 26 of the outer wall 16 may be greater than 150 percent of a cross- sectional area of the outlet 22 of the film cooling hole 14.
The film cooling hole 14 and elevated film cooling exhaust surface 24 may be formed via a casting process, such as an investment casting using a lost wax process. Because the configuration of the film cooling hole 14 and elevated film cooling exhaust surface 24 are relatively small, a shell insert may be used. In particular, a ceramic core may be formed. The method may include placing the ceramic core within an inner cavity formed by an inner surface of a wax die and injecting wax into one or more openings formed between the ceramic core and the inner surface of the wax die. The method may also include removing the wax die to reveal a wax component, coating the wax component with a ceramic coating to form an external ceramic shell with a ceramic core positioned therein and removing the wax component within the ceramic coating leaving one or more cavities within the ceramic coating. A shell insert may be attached to the external ceramic shell at each location of a film cooling hole. The shell insert may include protruded core features to create an internal cast hole feature in the airfoil surface. The method may also include filling the cavity within the ceramic coating with a molten metal and removing the ceramic shell and the ceramic core to form a cast component.
The protruded cooling hole support 8 forming the film cooling hole 14 may itself be cooled more effectively so the protruded cooling hole support 8 will not experience locally higher temperatures due to its raised location and thermal exposure. The cooling fluid flowing through the surface feature is marginally heated up thereby reducing the overall temperature difference between the inlet 18 and outlet 22. Thermal stresses within the thermal barrier coating 28 may be reduced because of the protruded cooling hole support 8 forming the film cooling hole 14 and the elevated film cooling exhaust surface 24 being positioned within the thermal barrier coating 28. The outlet 22 of the film cooling hole 14 may be configured to reduce local stress and bonding adjacent to the hole outlet.
The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.

Claims

CLAIMS We claim:
1 . A turbine airfoil (10), characterized in that:
a generally elongated hollow airfoil (34) formed from an outer wall (16), and having a leading edge (36), a trailing edge (38), a pressure side (40), a suction side (42), and a cooling system (20) positioned within interior aspects of the generally elongated hollow airfoil (34) and formed from at least one cavity; and
at least one film cooling hole (14) positioned in the outer wall (16) and having an inlet (18) in communication with the cooling system (20) and an outlet (22) in an elevated film cooling exhaust surface (24) that extends outwardly further than an adjacent outer surface (26) of the outer wall (16).
2. The turbine airfoil (10) of claim 1 , further characterized in that a thermal barrier coating on the outer surface (26) of the outer wall (16), and wherein the outer surface (30) of the elevated film cooling exhaust surface (24) is positioned between the outer surface (26) of the outer wall (16) and an outer surface (32) of the thermal barrier coating (28).
3. The turbine airfoil (10) of claim 2, characterized in that the outer surface (30) of the elevated film cooling exhaust surface (24) is positioned between a distance between 30 percent and 50 percent of a distance of thickness of the thermal barrier coating (28) from the outer surface (26) of the outer wall (16) to the outer surface (32) of the thermal barrier coating (28).
4. The turbine airfoil (1 0) of claim 1 , characterized in that sidewalls (46) extending between a base (48) at the outer surface (26) of the outer wall (16) and the elevated film cooling exhaust surface (24) are tapered from the base (48) at the outer surface (26) of the outer wall (16) to the elevated film cooling exhaust surface (24).
5. The turbine airfoil (10) of claim 4, characterized in that the sidewalls (46) extending between the base (48) of the outer surface (26) of the outer wall (16) and the elevated film cooling exhaust surface (24) are tapered nonlinearly from the base (48) at the outer surface (26) of the outer wall (16) to the elevated film cooling exhaust surface (24).
6. The turbine airfoil (10) of claim 1 , characterized in that a cross- sectional area of the outer surface (26) of the outer wall (16) is greater than a cross- sectional area of the outlet (22) of the at least one film cooling hole (14).
7. The turbine airfoil (1 0) of claim 6, characterized in that the cross- sectional area of the outer surface (26) of the outer wall (16) is at least fifty percent larger than a cross-sectional area of the outlet (22) of the at least one film cooling hole (14).
8. The turbine airfoil (10) of claim 1 , characterized in that a cross- sectional area of a base (48) at an intersection (50) of sidewalls (46) and the outer surface (26) of the outer wall (16) is larger than a cross-sectional area of the elevated film cooling exhaust surface (24).
9. The turbine airfoil (1 0) of claim 8, characterized in that the cross- sectional area of the base (48) at the intersection (50) of sidewalls (46) and the outer surface (26) of the outer wall (16) is greater than 150 percent of a cross-sectional area of the outlet (22) of the at least one film cooling hole (14).
PCT/US2015/016004 2015-02-16 2015-02-16 Turbine airfoil cooling system with film cooling hole within protruded cooling hole support WO2016133488A1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PCT/US2015/016004 WO2016133488A1 (en) 2015-02-16 2015-02-16 Turbine airfoil cooling system with film cooling hole within protruded cooling hole support

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2015/016004 WO2016133488A1 (en) 2015-02-16 2015-02-16 Turbine airfoil cooling system with film cooling hole within protruded cooling hole support

Publications (1)

Publication Number Publication Date
WO2016133488A1 true WO2016133488A1 (en) 2016-08-25

Family

ID=52589829

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2015/016004 WO2016133488A1 (en) 2015-02-16 2015-02-16 Turbine airfoil cooling system with film cooling hole within protruded cooling hole support

Country Status (1)

Country Link
WO (1) WO2016133488A1 (en)

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2641993A2 (en) * 2012-03-22 2013-09-25 Rolls-Royce plc A method of manufacturing a thermal barrier coated article
US20140102684A1 (en) * 2012-10-15 2014-04-17 General Electric Company Hot gas path component cooling film hole plateau

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2641993A2 (en) * 2012-03-22 2013-09-25 Rolls-Royce plc A method of manufacturing a thermal barrier coated article
US20140102684A1 (en) * 2012-10-15 2014-04-17 General Electric Company Hot gas path component cooling film hole plateau

Similar Documents

Publication Publication Date Title
JP6431737B2 (en) Method and system for cooling turbine components
EP3068975B1 (en) Gas turbine engine component and corresponding methods of manufacturing
US7270515B2 (en) Turbine airfoil trailing edge cooling system with segmented impingement ribs
US10415409B2 (en) Nozzle guide vane and method for forming such nozzle guide vane
EP2060745B1 (en) Gas turbine sealing segment
US6915840B2 (en) Methods and apparatus for fabricating turbine engine airfoils
US20120034101A1 (en) Turbine blade squealer tip
EP2614902B1 (en) Core for a casting process
US20100226788A1 (en) Turbine blade with incremental serpentine cooling channels beneath a thermal skin
EP2159375B1 (en) A turbine engine airfoil with convective cooling, the corresponding core and the method for manufacturing this airfoil
JP2018515343A (en) Ceramic core for multi-cavity turbine blades
US10704394B2 (en) Flow passage forming plate, flow passage forming member assembly and vane including the same, gas turbine, manufacturing method of flow passage forming plate, and modification method of flow passage forming plate
US7918647B1 (en) Turbine airfoil with flow blocking insert
US10364683B2 (en) Gas turbine engine component cooling passage turbulator
US20170089207A1 (en) Turbine airfoil cooling system with leading edge impingement cooling system and nearwall impingement system
EP2385216B1 (en) Turbine airfoil with body microcircuits terminating in platform
US8366393B2 (en) Rotor blade
EP2775101B1 (en) Gas turbine blade
US10683762B2 (en) Gas engine component with cooling passages in wall
US20150184518A1 (en) Turbine airfoil cooling system with nonlinear trailing edge exit slots
JP2016540150A (en) Investment casting for the vane segment of gas turbine engines.
WO2015195088A1 (en) Turbine airfoil cooling system with leading edge impingement cooling system
EP4028643B1 (en) Turbine blade, method of manufacturing a turbine blade and method of refurbishing a turbine blade
WO2016133488A1 (en) Turbine airfoil cooling system with film cooling hole within protruded cooling hole support
WO2016148695A1 (en) Turbine airfoil with internal cooling system having nearwall cooling channels formed from an inner wall formed separately from an outer wall forming the turbine airfoil

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 15706644

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase

Ref country code: DE

122 Ep: pct application non-entry in european phase

Ref document number: 15706644

Country of ref document: EP

Kind code of ref document: A1