Cupped contour for gas turbine engine blade assembly

An exemplary method of forming an endwall with a contour includes casting an endwall with at least one cooling channel having an opening from the endwall, and covering the opening with a cupped contour formed on the endwall. An exemplary gas turbine engine blade assembly includes an endwall with a plurality of cooling channels, an airfoil extending radially from the endwall to a tip, and a cupped contour formed on the endwall to provide a cooling chamber between the cupped contour and a radially facing surface of the endwall.

BACKGROUND

This disclosure relates to a contoured endwall of a gas turbine engine blade assembly.

Gas turbine engines typically include a compressor section, a combustor section, and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.

The compressor and turbine sections of a gas turbine engine typically include alternating rows of rotating blades and stationary vanes. The turbine blades rotate and extract energy from the hot combustion gases that are communicated through the gas turbine engine. The turbine vanes prepare the airflow for the next set of blades.

The vanes and blades extend from endwalls that may be contoured to manipulate flow. The outer casing of an engine static structure may include one or more blade outer air seals (BOAS) providing endwalls that are be contoured to manipulate flow by reducing secondary flow losses and flow migration.

SUMMARY

A method of forming an endwall with a contour according to an exemplary aspect of the present disclosure includes, among other things, casting an endwall with at least one cooling channel having an opening from the endwall, and covering the opening with a cupped contour that is formed on the endwall to provide a portion of a gas path surface.

In a further non-limiting embodiment of the foregoing method, the method includes forming the cupped contour on the endwall using additive manufacturing.

In a further non-limiting embodiment of any of the foregoing methods, the endwall has a first material composition, and the cupped contour has a second material composition different than the first material composition.

In a further non-limiting embodiment of any of the foregoing methods, the method includes forming the cupped contour on a radially facing surface of the endwall.

In a further non-limiting embodiment of any of the foregoing methods, the method includes covering an inlet to at least one cooling channel of the endwall with the cupped contour.

In a further non-limiting embodiment of any of the foregoing methods, an outlet and the inlet open to a cooling chamber provided by the cupped contour.

In a further non-limiting embodiment of any of the foregoing methods, the endwall is an endwall of a blade.

In a further non-limiting embodiment of any of the foregoing methods, the method includes providing at least one cooling channel in the endwall using a core held within a mold during the casting.

In a further non-limiting embodiment of any of the foregoing methods, a portion of the core providing the opening of the cooling channel provides an attachment point that secures the core within the mold.

In a further non-limiting embodiment of any of the foregoing methods, the endwall is an endwall of a blade assembly and the attachment point is a first attachment point, a second attachment point that secures the core within the mold is adjacent a blade tip of the blade assembly, and a third attachment point that secures the core within the mold is adjacent a root of the blade assembly.

A gas turbine engine assembly according to another exemplary aspect of the present disclosure includes, among other things, an endwall having a first material composition, and a cupped contour of a second material composition that is formed on the endwall. The first material composition is different than the second material composition.

In a further non-limiting embodiment of the foregoing assembly, at least one cooling channel of the endwall opens to a cooling chamber provided by the cupped contour.

In a further non-limiting embodiment of any of the foregoing assemblies, the endwall is entirely radially misaligned from the cooling chamber.

In a further non-limiting embodiment of any of the foregoing assemblies, the cupped contour covers one or more cooling channel openings.

In a further non-limiting embodiment of any of the foregoing assemblies, the cupped contour covers one or more cooling channel inlets.

In a further non-limiting embodiment of any of the foregoing assemblies, the endwall is an endwall of a blade.

In a further non-limiting embodiment of any of the foregoing assemblies, the endwall is a cast endwall, and the cupped contour is an additively manufactured cupped contour.

A gas turbine engine blade assembly according to an exemplary aspect of the present disclosure includes, among other things, an endwall with a plurality of cooling channels, an airfoil extending radially from the endwall to a tip, and a cupped contour formed on the endwall to provide a cooling chamber between the cupped contour and a radially facing surface of the endwall.

In a further non-limiting embodiment of the foregoing assembly, at least one opening of the plurality of cooling channels opens to the cooling chamber, and at least one inlet of the plurality of cooling channels opens to the cooling chamber. The cupped contour covers the at least one opening and the at least one inlet.

In a further non-limiting embodiment of any of the foregoing assemblies, the endwall is a cast endwall, and the cupped contour is and additively manufactured cupped contour.

DETAILED DESCRIPTION

This disclosure relates to a contoured endwall of an assembly of a gas turbine engine assembly. More particularly, this disclosure relates to a cupped contour that provides a cooling chamber. A cooling channel formed within an endwall of a gas turbine engine assembly opens to the cooling chamber within the cupped contour.

FIG. 1schematically illustrates a gas turbine engine20. The gas turbine engine20is disclosed herein as a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26, and a turbine section28.

The fan section22drives air along a bypass flow path B in a bypass duct defined within a nacelle, while the compressor section24drives air along a core flow path C for compression and communication into the combustor section26, and then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, the examples herein are not limited to use with two-spool turbofans and may be applied to other types of turbomachinery, including direct drive engine architectures, three-spool engine architectures, and ground-based turbines.

The low speed spool30generally includes an inner shaft40that interconnects a fan42, a first (or low) pressure compressor44and a first (or low) pressure turbine46. The inner shaft40is connected to the fan42through a speed change mechanism, which in exemplary gas turbine engine20is illustrated as a geared architecture48, to drive the fan42at a lower speed than the low speed spool30.

The high speed spool32includes an outer shaft50that interconnects a second (or high) pressure compressor52and a second (or high) pressure turbine54. A combustor56is arranged between the high pressure compressor52and the high pressure turbine54. A mid-turbine frame58of the engine static structure36is arranged generally between the high pressure turbine54and the low pressure turbine46. The mid-turbine frame58further supports the bearing systems38in the turbine section28. The inner shaft40and the outer shaft50are concentric and rotate via bearing systems38about the engine central longitudinal axis A, which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor44then the high pressure compressor52, mixed and burned with fuel in the combustor56, then expanded over the high pressure turbine54and low pressure turbine46. The mid-turbine frame58includes airfoils, which are in the core airflow path C.

The turbines46,54rotationally drive the respective low speed spool30and high speed spool32in response to the expansion. It will be appreciated that each of the positions of the fan section22, compressor section24, combustor section26, turbine section28, and fan drive gear system48may be varied. For example, gear system48may be located aft of combustor section26or even aft of turbine section28, and fan section22may be positioned forward or aft of the location of gear system48.

Referring now toFIG. 2with continuing reference toFIG. 1, an example gas turbine engine blade assembly60from the turbine section28of the gas turbine engine ofFIG. 1includes an endwall64, an airfoil68, and a root72or base. The airfoil68extends radially in a first direction from the endwall64to a tip region76. The root72extends radially in an opposite, second direction from the endwall64.

The blade assembly60includes a cupped contour80formed on the endwall64. The cupped contour80helps manipulate and direct flow over and near the endwall64when the gas turbine engine blade assembly60is used in the gas turbine engine20. In another example, an endwall of a vane includes the cupped contour.

Referring now toFIGS. 3-5with continuing reference toFIG. 2, the blade assembly60includes at least one cooling channel84. During operation, fluid, such as air from an air supply88, communicates through the cooling channel84to cool the gas turbine engine blade assembly60. The compressor section24of the gas turbine engine20(FIG. 1) provides the air supply88in some examples.

In this example, air enters the at least one cooling channel84near the root72. The air communicates through the cooling channel84and exits at an outlet near the tip region76of the airfoil68. The air communicates thermal energy from the blade assembly60.

For simplicity, a single cooling channel84is shown inFIG. 3. The blade assembly60could include, however, a network of several cooling channels84that are each separate and distinct from one another. The cooling channels84could have a serpentine configuration. The cooling channels84could have an outlet in a leading edge of the airfoil68, the trailing edge of the airfoil68, or some other area of the blade assembly60.

The example cooling channel84extends through a portion of the endwall64, such that the endwall64provides a portion of the at least one cooling channel84. The endwall64provides an opening92from the portion of the cooling channel84within the endwall64. Air communicates from the cooling channel84of the endwall64through the opening92into a cooling chamber96provided by the cupped contour80. Moving air into the cooling chamber96cools the cupped contour80, the endwall64, or both. The cooling chamber96is part of the cooling channel84in this example.

Notably, in this example, the cupped contour80covers the opening92so that air does not exit directly from the cooling chamber96. Moving air through the cooling chamber96of the cooling channel84cools the cupped contour80. Prior art contours have lacked cooling chambers, and thus were prone to experience undesirable temperature extremes.

The endwall64is, in this example, a cast metallic component. The endwall64is cast together with the airfoil68and the root72. The cupped contour80is then applied to the endwall64via an additive manufacturing process.

Referring now toFIG. 6, with continuing reference toFIG. 3, a mold100and a core104provide an open area108that receives molten material when casting the endwall64, the airfoil68, and the root72. The open area108is filled with a wax that melts when molten material is introduced into the open area108.

The core104is a silica-based ceramic core in this example, but other materials such as alumina based ceramics or refractory metals are possible. The core104blocks the molten material from entering areas that will form the cooling chamber84within the endwall64, the airfoil68, and the root72.

The core104includes a first attachment point112, a second attachment point116, and a third attachment point120. The attachment points112,116, and120provide an at least 3-point physical datum nest on the core104surface retaining the core104to design intent relative to the mold100when casting the endwall64, the airfoil68, and the root72.

The first attachment point112extends through an area that will provide the opening92to the cooling chamber96. The second attachment point116is adjacent an area that will form the tip region76. The third attachment point120is adjacent an area that will provide the root72.

Prior art cores used to create cooling channels within gas turbine engine blades have lacked at least an attachment point near the endwall, which can complicate the casting process as the core was more difficult to stabilize within a mold. With a lack of a third contact retention feature, outside of attachment points116and120, the core exhibits a rotational degree of freedom revolving around an axis created by the centroid points of116and120. This rotation, during casting, allows the core104to misalign and can negatively impact the desire casting wall thickness distribution.

Referring now toFIG. 7, when the molten material has hardened within the open area108, the endwall64, the airfoil68, and the root72are removed from the mold100. The cooling chamber84extends through these portions of the gas turbine engine blade assembly60.

The opening92within the endwall64extends through a surface98of the endwall64that will face radially outward when positioned within the gas turbine engine20ofFIG. 1. The cupped contour80is radially outside the surface of the endwall64such that the endwall64is radially misaligned from the cooling chamber96.

Referring now toFIG. 8, an additive manufacturing process is used, in this example, to form the cupped contour80on the endwall64, which will cover the opening92. Adding the cupped contour80after casting, rather than casting the cupped contour80, can reduce manufacturing complexity associated with casting components of complex geometries.

The cupped contour80has a material composition that can be different than a material composition of the remaining portions of the blade assembly60. Typical materials are classed as superalloys of the nickel or cobalt base element variety with typical materials being PW1484, CMSX-4, Mar-M-247, Mar-M-509, Rene 80, Rene N5, Haynes Alloy X, Inconel 625, Inconel 723, and others whose material properties are suitable to high temperature applications.

Additive manufacturing uses a wand124to stack successive layers L1and L2of material, and additional layers, until the desired cupped contour80is formed and covers the opening92. The endwall64and remaining portions of the blade assembly60can be preheated from 500 to 1900 degrees Fahrenheit to facilitate adhesion between the cupped contour80and the remaining portion of the blade assembly60, for example. The preheating temperature is typically 500 degrees Fahrenheit below the sintering temperature of the deposited material. Upon initial deposit, secondary heating operations with the laser can be used to further consolidate the material and provide heat treatment to further solution the alloy.

Air can move from the cooling channel84of the endwall64through the opening92into the cooling chamber96, and from the cooling chamber96through the opening back to the cooling channel84of the endwall64. Thus, air can move to the cooling chamber96and from the cooling chamber96through the singular opening92of the cooling channel84.

In another example, one or more cooling channels in the endwall64open to the cooling chamber96to permit air air into the cooling chamber96, and one or more other cooling channels open to the cooling chamber96to permit air to move from the cooling chamber96back to the network of cooling channels84in the endwall64.

Air could move from to or from the cooling chamber96through an opening in the airfoil68, or another portion of the blade assembly60in other examples.