Gas turbine engines with improved airfoil dust removal

• US 10,641,106 B2
• Filed: 11/13/2017
• Issued: 05/05/2020
• Est. Priority Date: 11/13/2017
• Status: Active Grant

• Current Assignee: Honeywell International Inc.
• Original Assignee: Honeywell International Inc.
• Inventors: Crites, Daniel C, Kahrs, Michael, Wakefield, Brandan, Riahi, Ardeshir
• Primary Examiner(s): Verdier, Christopher
• Assistant Examiner(s): Fountain, Jason A
• Application Number:
  US15/810,809
• Publication Number:
  US 20190145266A1
• Time in Patent Office: 904 Days
• Field of Search: None
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F05D 2220/32 in gas turbines
• CPC Class Codes: F05D 2230/50 Building or constructing in particular ways
• CPC Class Codes: F05D 2240/30 Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
• CPC Class Codes: F05D 2240/307 related to the tip of a rotor blade
• CPC Class Codes: F05D 2260/201 by impingement of a fluid
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/2212 by creating turbulence
• CPC Class Codes: F05D 2260/607 Preventing clogging or obstruction of flow paths by dirt, dust, or foreign particles
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

An airfoil for a rotor blade in a gas turbine engine includes a first side wall and a second side wall joined to the first side wall at a leading edge and a trailing edge. The airfoil further includes a tip cap extending between the first and second side walls such that the tip cap and at least portions of the first and second side walls form a blade tip and an internal cooling system. The internal cooling system includes a leading edge cooling circuit, a central cooling circuit, and a trailing edge cooling circuit. Each of the internal passages within the leading edge cooling circuit, the central cooling circuit, and the trailing edge cooling circuit is bounded in the radial outward direction with a surface that has at least one escape hole or that is positively angled in the radial outward direction relative to a chordwise axis.

Claims

1. An airfoil for a rotor blade in a gas turbine engine, comprising:

a first side wall;

a second side wall joined to the first side wall at a leading edge and a trailing edge, the first and second side walls extending in a radial outward direction from a base coupled to an airfoil platform;

a tip cap extending between the first and second side walls such that the tip cap and at least portions of the first and second side walls form a blade tip; and

an internal cooling system arranged within the first and second side walls configured to direct cooling air from one or more source passages, through the airfoil, and out of the airfoil, the internal cooling system includes a leading edge cooling circuit, a central cooling circuit, and a trailing edge cooling circuit, the internal cooling system further including tip slots formed in the blade tip and trailing edge slots formed in the trailing edge,

wherein the leading edge cooling circuit, the central cooling circuit, and the trailing edge cooling circuit are formed by a plurality of internal passages fluidly coupled to the one or more source passages, and each of the internal passages within the leading edge cooling circuit, the central cooling circuit, and the trailing edge cooling circuit is bounded in the radial outward direction with a surface that has at least one escape hole or that is positively angled in the radial outward direction relative to a chordwise axis, and the central cooling circuit includes;

a first radial passage to direct the cooling air received by the central cooling circuit from the one or more source passages in the radial outward direction, and the first radial passage is bounded in the radial outward direction by the tip cap such that at least a first portion of the cooling air in the first radial passage flows in the radial outward direction through the tip slots in the blade tip as the at least one escape hole;

a first chordwise passage fluidly coupled to receive a second portion of the cooling air from the first radial passage, the first chordwise passage extending from the first radial passage to at least a first portion of the trailing edge slots in the trailing edge such that the second portion of the cooling air flows out of the airfoil via the first portion of trailing edge slots; and

the leading edge cooling circuit includes a first leading radial passage extending in the radial outward direction and at least partially defined by a first leading radial wall in a chordwise aft direction, the tip cap in the radial outward direction and a second leading radial wall in a chordwise forward direction, and a second leading radial passage at least partially defined by the leading edge in the chordwise forward direction, the tip cap in the radial outward direction and the second leading radial wall in the chordwise aft direction,

wherein the trailing edge cooling circuit includes a first set of serpentine passages formed by a first trailing radial passage extending in the radial outward direction from the one or more source passages, a second trailing radial passage extending in a radial inward direction from the first trailing radial passage at a first transition, and a third trailing radial passage extending from the second trailing radial passage in the radial outward direction, and the first transition is at least partially formed by a first chordwise wall defining at least a first escape hole of the at least one escape hole and the third trailing radial passage is bounded in the radial outward direction by a second chordwise wall defining at least a second escape hole of the at least one escape hole.

2. The airfoil of claim 1, wherein the first chordwise passage decreases in cross-sectional area between the first radial passage and the trailing edge slots.

3. The airfoil of claim 1, wherein the central cooling circuit includes the first chordwise passage defined at least in part by the tip cap, the first chordwise wall, the second chordwise wall, and the trailing edge, and

wherein the first transition is fluidly coupled to the first chordwise passage by the at least the first escape hole of the at least one escape hole and the third radial passage is fluidly coupled to the first chordwise passage by the at least the second escape hole of the at least one escape hole.

4. The airfoil of claim 1, wherein the trailing edge cooling circuit includes

the third trailing radial passage extending in the radial outward direction and being formed by a first radial wall in a chordwise forward direction and a second radial wall in a chordwise aft direction;

and

a trailing edge chamber being at least partially formed the second radial wall in the chordwise forward direction and the trailing edge in the chordwise aft direction, the second radial wall defining holes that fluidly couple the first radial passage to the trailing edge chamber,

wherein the trailing edge chamber is fluidly coupled to the trailing edge slots.

5. The airfoil of claim 1, wherein each of the tip slots have a converging cross-sectional area from an inlet to an outlet.

6. The airfoil of claim 1, wherein each of the trailing edge slots have a converging cross-sectional area from an inlet to an outlet.

7. A method for producing an airfoil of a rotor blade for a gas turbine engine, comprising the steps of:

providing a baseline airfoil with an internal cooling system defined by internal walls that define internal passages configured to direct cooling air from one or more source passages, through the airfoil, and out of the airfoil, wherein the internal cooling system includes a leading edge cooling circuit, a central cooling circuit, and a trailing edge cooling circuit, the internal cooling system further including tip slots formed in a blade tip and trailing edge slots formed in a trailing edge, each of the internal passages within the leading edge cooling circuit, the central cooling circuit, and the trailing edge cooling circuit bounded in a radial outward direction with a surface, the central cooling circuit includes a first radial passage to direct the cooling air received by the central cooling circuit from the one or more source passages in the radial outward direction, and the first radial passage is bounded in the radial outward direction by the tip cap, and the central cooling circuit includes a first chordwise passage fluidly coupled to receive a second portion of the cooling air from the first radial passage, the first chordwise passage extending from the first radial passage to at least a first portion of the trailing edge slots in the trailing edge such that the second portion of the cooling air flows out of the airfoil via the first portion of trailing edge slots, the leading edge cooling circuit includes a first leading radial passage extending in the radial outward direction and at least partially defined by a first leading radial wall in a chordwise aft direction, the tip cap in the radial outward direction and a second leading radial wall in a chordwise forward direction, and a second leading radial passage at least partially defined by a leading edge in the chordwise forward direction, the tip cap in the radial outward direction and the second leading radial wall in the chordwise aft direction, and the trailing edge cooling circuit includes a first set of serpentine passages formed by a first trailing radial passage extending in the radial outward direction from the one or more source passages, a second trailing radial passage extending in a radial inward direction from the first trailing radial passage at a first transition, and a third trailing radial passage extending from the second trailing radial passage in the radial outward direction;

evaluating the internal walls and identifying stagnation areas including identifying uninterrupted chordwise walls that oppose cooling air flow in the radial outward direction;

providing a modified airfoil from the baseline airfoil by adding escape holes through the chordwise walls in the stagnation areas, wherein the surface has one of the escape holes or is positively angled in the radial outward direction relative to a chordwise axis, the first transition is at least partially formed by a first chordwise wall defining at least one of the escape holes, the third trailing radial passage is bounded in the radial outward direction by a second chordwise wall defining at least one of the escape holes, and at least a first portion of the cooling air in the first radial passage flows in the radial outward direction through the tip slots in the blade tip as one of the escape holes; and

producing the modified airfoil.

8. The method of claim 7, wherein the identifying the stagnation areas include identifying the internal walls that redirect cooling air flow in the radial inward direction.