GAS TURBINE ENGINE COMPONENT HAVING FOAM CORE AND COMPOSITE SKIN WITH COOLING SLOT

• US 20150040570A1
• Filed: 12/23/2013
• Published: 02/12/2015
• Est. Priority Date: 03/03/2013
• Status: Active Grant

• Current Assignee: Rolls-Royce North American Technologies, Inc. (Rolls-Royce Holdings Plc)
• Original Assignee: Rolls-Royce North American Technologies, Inc. (Rolls-Royce Holdings Plc)
• Inventors: Uskert, Richard C., Freeman, Ted J., Thomas, David J.
• Granted Patent: US 9,759,090 B2
• Time in Patent Office: Days
• US Class Current: 60/754
• CPC Class Codes: F01D 25/08 Cooling of machines or engines in general F01P; Heating; Heat-insulation of blade-carrying members, of blades F01D5/00
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 5/147 Construction, i.e. structural features, e.g. of weight-saving hollow blades F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18
• CPC Class Codes: F01D 5/185 Liquid cooling F01D5/181 takes precedence
• CPC Class Codes: F01D 5/282 Selecting composite materials, e.g. blades with reinforcing filaments
• CPC Class Codes: F01D 9/02 Nozzles; Nozzle boxes; Stator blades; Guide conduits , e.g. individual nozzles nozzle boxes F01D9/047
• CPC Class Codes: F05D 2240/35 Combustors or associated equipment
• CPC Class Codes: F05D 2260/203 by transpiration cooling
• CPC Class Codes: F05D 2300/514 Porosity
• CPC Class Codes: F05D 2300/603 Composites; e.g. fibre-reinforced
• CPC Class Codes: F05D 2300/6033 Ceramic matrix composites [CMC]
• CPC Class Codes: F05D 2300/612 Foam
• CPC Class Codes: F23R 2900/03041 Effusion cooled combustion chamber walls or domes
• CPC Class Codes: F23R 2900/03042 Film cooled combustion chamber walls or domes
• CPC Class Codes: F23R 2900/03043 Convection cooled combustion chamber walls with means for guiding the cooling air flow means for creating turbulence F23R2900/03045
• CPC Class Codes: F23R 3/002 Wall structures F23R3/02 and F23R3/007 take precedence
• CPC Class Codes: F23R 3/007 constructed mainly of ceramic components
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft
• CPC Class Codes: Y10T 29/49316 Impeller making

Abstract

In one embodiment, a gas turbine engine component includes a foam based core and a composite skin member. Both the foam based core and the composite skin member can be used to structurally support the gas turbine engine component. The composite skin member can be a CMC material and is used to partially encapsulate the foam core. The gas turbine engine component can take the form of an airfoil member such as a blade or a vane, a combustor liner, etc. A first portion of the composite skin member includes a first surface extending past an edge of the component creating a step approximate an edge section. In another embodiment, composite skin members can be used to form a continuous shape for the edge section such that the foam core forms part of a gas path surface.

Claims

1. An apparatus comprising:

a cooled gas turbine engine airfoil member having an outer skin that forms a gas path surface of the cooled gas turbine engine airfoil member, the outer skin including a composite material having a matrix material;

a foam core also having the matrix material and disposed internal to the composite material of the outer skin, the foam core occupying substantially an entirety of space internal to the composite material of the outer skin; and

a foam based discharge slot having the matrix material also located at a trailing edge portion of the cooled gas turbine engine airfoil member wherein a width of the slot along a span of the airfoil member is greater than a height of the airfoil member in a thickness direction of the airfoil member, the foam core defining the foam based discharge slot and having a cooling passage through which cooling fluid exits the cooled gas turbine engine airfoil member.

2. The apparatus of claim 1, wherein the outer skin includes a plurality of composite layers.

3. The apparatus of claim 2, wherein the foam of the foam based discharge slot forms part of a pressure side of the cooled gas turbine engine airfoil member.

4. The apparatus of claim 2, wherein the foam based discharge slot is in the form of a gill slot having an aft portion that extends away from a pressure side of the airfoil member toward a suction side of the airfoil member.

5. The apparatus of claim 1, wherein the foam is a reticulated foam construction structured to convey a cooling fluid therethrough and deliver the fluid to the discharge slot.

6. The apparatus of claim 5, wherein the matrix material of the outer skin is a ceramic based composite material.

7. An apparatus comprising:

a composite layer forming an outer portion of a gas turbine engine component having a cooling passage disposed therein, the gas turbine engine component having a first side of the composite layer and a second side of the composite layer, the first and second sides of the composite layer extending between an upstream portion of the gas turbine engine component and a downstream portion of the gas turbine engine component;

a structural foam core disposed internal to the composite layer and substantially extending upstream of both the first side of composite layer and second side of the composite layer, the structural foam core having an open cell construction to permit a cooling fluid to pass; and

a discharge slot having an elongate shape that includes a greater dimension in width than a dimension in a thickness direction, the discharge slot defined by the composite layer of the first side and second side to form a barrier edge to discourage bulk transfer of cooling fluid through the first side and second side, and wherein the discharge slot is also defined by the structural foam core that is disposed at an exit of the discharge slot and is sandwiched between the first side of composite layer and second side of composite layer.

8. The apparatus of claim 7, wherein the structural foam core is structured to bear greater than 50% of a load imparted to the gas turbine engine component.

9. The apparatus of claim 8, wherein the gas turbine engine component includes a plurality of discharge slots, wherein the gas turbine engine component is one of a blade and a vane, and which further includes a plurality of composite layers in the first side and the second side.

10. The apparatus of claim 9, wherein the exit forms a gas path surface of the component, and wherein the composite layer is a plurality of composite layers.

11. The apparatus of claim 9, wherein the discharge slot is a gill slot.

12. The apparatus of claim 8, which further includes a plurality of composite layers in the first side and second side, wherein the plurality of composite layers are formed from a ceramic matrix composite material, and wherein the gas turbine engine component is a combustor liner.

21. The apparatus of claim 7, wherein the gas turbine engine component forms one of a blade and a vane, the component having a plurality of composite plies arranged in an airfoil shape;

and further comprising a core means for structurally supporting the gas turbine engine component, forming the internal cooling passage, and defining the discharge slot of the component.

22. The apparatus of claim 21, wherein the core means is a ceramic matrix composite material.

13. (canceled)

14. (canceled)

15. A method comprising:

providing a foam core of a gas turbine engine component; and

enveloping the foam core of the gas turbine engine component with a layer of fiber ply material;

infiltrating the layer of fiber ply material and the foam core with a matrix material to form a composite outer layer and a foam core having the matrix material; and

forming a cooling slot at a trailing edge of the gas turbine engine component such that a fluid passage internal to the foam core extends to an exit of the gas turbine engine composite at the cooling slot.

16. The method of claim 15, wherein the forming includes removing a portion of at least one of the layer of fiber ply material and the foam core, and wherein the gas turbine engine component is one of a blade and a vane.

17. The method of claim 15, wherein the providing includes receiving the foam core from an earlier manufacturing activity of forming the foam core, wherein the enveloping happens after the providing.

18. The method of claim 15, wherein the extending includes wrapping the layer of fiber ply material from a first side of the trailing edge of the gas turbine engine component, around a leading edge of the component, and to a second side of the trailing edge of the component.

19. The method of claim 18, wherein the enveloping and the extending occur as part of a composite ply lay-up operation, wherein the forming occurs no earlier than the laying, and which further includes laying a first portion of a second layer of fiber ply material upon the first portion of the layer of fiber ply material.

20. The method of claim 15, wherein the forming includes defining a cooling slot that extends along a span of the gas turbine engine component and that has a height smaller than a length of the extension along the span, and which further includes co-curing the foam core and the layer of fiber ply material during a rigidization process.