Patent Description:
With each new engine design, gas temperatures increase and cooling flow requirements decrease. This requires cooling flow to be utilized in a more efficient manner and flow distribution to be tailored to prevent overcooling in certain regions. Because of radial gas temperature profiles, trailing edges of airfoils are a region where cooling flow distribution could be tailored and overall flow reduced. However, current cast trailing edges are already at the minimum area required to prevent core break during the casting process and thus do not allow a reduction or redistribution of flow. One option would be to replace cast trailing edge slots with drilled filmholes to allow a redistribution and reduction of cooling flow. However, this presents manufacturing challenges in the casting process resulting in an undesirable increase in variation of core position and wall thicknesses.

<CIT> and <CIT> disclose prior art gas turbine engine components as set forth in the preamble of claim <NUM>.

According to the invention, there is provided a gas turbine engine component as set forth in claim <NUM>.

In an embodiment, the cast slots extend in a generally axial direction and the drilled filmholes extend at an obtuse angle relative to the cast slots.

In another embodiment according to any of the previous embodiments, the body comprises a vane, blade or BOAS.

In another embodiment according to any of the previous embodiments, the body comprises a turbine component.

There is further provided a gas turbine engine according to claim <NUM>.

In an embodiment according to any of the previous embodiments, the cast slots having a length greater than a width, and wherein the length extends in the axial direction and the drilled filmholes extend along an obtuse angle relative to the cast slots.

In another embodiment according to any of the previous embodiments, the cast slots are configured to position a core during a casting process for the turbine component.

There is further provided a method of manufacturing a gas turbine engine component according to claim <NUM>.

In an embodiment, the cast slots are formed to extend in a generally axial direction and forming the drilled filmholes to extend at an obtuse angle relative to the cast slots.

In another embodiment according to any of the previous embodiments, filmhole sizes and spacing are tailored relative to the cast slots to meet cooling flow requirements for a specified engine configuration.

The foregoing features and elements may be combined in any combination without exclusivity, unless expressly indicated otherwise.

The low speed spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a first (or low) pressure compressor <NUM> and a second (or low) pressure turbine <NUM>. The high speed spool <NUM> includes an outer shaft <NUM> that interconnects a second (or high) pressure compressor <NUM> and a first (or high) pressure turbine <NUM>.

The geared architecture <NUM> may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>.

The fan section <NUM> of the engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>,<NUM>). The flight condition of <NUM> Mach and <NUM>,<NUM> ft (<NUM>,<NUM>), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about <NUM> ft / second (<NUM>/s).

<FIG> shows a vane <NUM> positioned between a first blade <NUM> and a second blade <NUM> in the turbine section <NUM>. In one example, the blades <NUM>, <NUM> are attached to a full hoop disk <NUM> with fore and aft coverplates <NUM>. The disks <NUM> are driven by a turbine rotor about the engine center axis A. Airfoil bodies <NUM> extend from a radially inward platform <NUM> of blades <NUM>,<NUM> to a tip <NUM>. Blade outer air seals (BOAS) <NUM> are mounted to a radially outer full hoop case <NUM> and are spaced from the tips <NUM> by a small clearance gap as known.

The vane <NUM> includes a radially inner platform <NUM>, a radially outward platform <NUM>, and an airfoil body <NUM> extending between the radially inner platform <NUM> and radially outer platform <NUM>. The vane <NUM> is coupled to the full hoop case <NUM> at the radially outer platform <NUM> and includes feather seals <NUM> between vane segments at the radially inner platform <NUM> and radially outer platform <NUM>. A full hoop inner air seal <NUM> is coupled to the radially inner platform <NUM> and cooperates with seals <NUM> on the coverplates <NUM> of the blades <NUM>, <NUM>.

Between the full hoop case <NUM> and the radially outer platform <NUM> is a vane outer diameter cavity <NUM>, and between the radially inner platform <NUM> and the inner air seal <NUM> is a vane inner diameter cavity <NUM>. Radially inward of the inner air seal <NUM> and between the disks <NUM> is a rotor cavity <NUM>. One or more orifices <NUM> are formed in the inner air seal <NUM> to direct flow into the vane inner diameter cavity <NUM>. Cooling flow is also directed into the vane outer diameter cavity <NUM> as known.

Cooling channels <NUM> (<FIG>) are formed within the vane <NUM> to receive the cooling air flow. The airfoil body <NUM> of the vane <NUM> extends from a leading edge <NUM> to a trailing edge <NUM>. In the example shown in <FIG>, the air flow from the vane inner diameter cavity <NUM> is directed into a fore inner cooling channel 100a and is directed out from the airfoil body <NUM> via the leading edge <NUM> as indicated by arrow <NUM>. The air from the vane outer diameter cavity <NUM> is directed into an aft inner cooling channel 100b and is directed out from the airfoil body <NUM> via the trailing edge <NUM> as indicated by arrows <NUM>.

With each new engine design, gas temperatures increase and cooling flow requirements decrease. This requires cooling flow to be utilized in a more efficient manner and flow distribution to be tailored to prevent overcooling in certain regions. Because of radial gas temperature profiles, trailing edges of airfoils is an area where cooling flow distribution could be tailored and overall flow reduced. However, current cast trailing edges are already at the minimum area required to prevent core break during the casting process and thus do not allow a reduction or redistribution of flow.

A casting process uses a core to form open internal areas within a component. The core is positioned in a die and material is supplied to the die to flow around the core to cast the component. Once the component is cast the core is removed to provide the open areas within the component. As known, the casting process requires a certain core area to make sure the core remains intact during the process. For a cast airfoil body, the trailing edge includes a plurality of cast slots that extend from the inner diameter to the outer diameter. The slots are formed to be at the minimum size that is required to prevent core break during the casting process. Further, the structure on the core used to form the slots helps position the core during the casting process. The resulting component includes cast slots for the entire radial span which results in excess cooling flow. Further, as the cast slots are the same along the radial span there is uniform cooling flow that is not tailorable to address hotter areas of the airfoil body. Replacing the cast trailing edge slots with drilled filmholes provides more tailoring to allow a redistribution and reduction of cooling flow. However, as discussed above, the cast trailing edge slots help position the core during the casting process and removing them would result in larger variation of the core position and in wall thicknesses.

The subject invention provides a configuration where the airfoil body <NUM> includes one or more cast slots <NUM> formed in the trailing edge <NUM> and one or more drilled cooling holes <NUM>. The cast slots <NUM> and drilled filmholes <NUM> direct flow from the internal channel 100b to an external location of the body <NUM>. The cast slots <NUM> are formed by core structures that are used to position the core during casting; however, drilled filmholes <NUM> are machined subsequent to casting to reduce excess cooling flow and to tailor flow to cool hotter areas of the body <NUM>.

As shown in <FIG>, which falls outside the scope of the claims, the airfoil body <NUM> extends from a radially inner end <NUM> to a radially outer end <NUM>. In one example, the cast slots <NUM> are positioned radially outward of the drilled filmholes <NUM>. A section of a cast slot <NUM> is shown in <FIG> and a section of a drilled filmhole <NUM> is shown in <FIG>. The cast slot <NUM> extends from the internal channel 100b to an external surface <NUM> of the airfoil body <NUM>. The cast slots <NUM> have a length that is greater than a width and extend from an inner slot end <NUM> to an outer slot end <NUM> at the external surface <NUM>. The slots <NUM> can be located anywhere along the radial span of the airfoil body <NUM>. The slots <NUM> can be the same length or variable lengths relative to each other.

The drilled filmhole <NUM> extends from the internal channel 100b to the external surface <NUM>. The filmholes <NUM> can have different shapes and/or sizes and can be located anywhere along the radial span of the airfoil body <NUM>. Further, the filmholes <NUM> can be orientated at different angles relative to the slots <NUM> and/or body surface.

In one example, an axial direction is defined in a direction extending from the leading edge <NUM> to the trailing edge <NUM>, which is a direction that is common with the engine center axis A. A radial direction is defined as a direction that is perpendicular to the axial direction, and which extends radially outward from the center axis A. The cast slots <NUM> extend generally in the axial direction and the drilled filmholes <NUM> extend at an obtuse angle relative to the axially extending cast slots <NUM>. In other words, the filmholes <NUM> are orientated to be neither perpendicular nor parallel to the cast slots <NUM>. The filmholes <NUM> can be orientated at any of various angles to tailor cooling flow as needed.

In the example shown in <FIG>, the cast slots <NUM> are positioned radially outward of the drilled filmholes <NUM>. The cast slots <NUM> are spaced along the upper half of the radial span and the filmholes <NUM> are spaced along the lower half of the radial span. The cast slots <NUM> extend in an axial direction and the filmholes <NUM> are drilled at desired locations to tailor the flow to meet cooling flow requirements. In the example shown, the filmholes <NUM> extend from a fore hole end at a radial outward angle to an aft hole end. Further, filmholes <NUM> are more closely spaced together at the center of the body than at the radially inner end <NUM> to tailor cooling flow.

In the example shown in <FIG>, which falls outside the scope of the claims, the cast slots <NUM> are positioned radially inward of the drilled filmholes <NUM>. The cast slots <NUM> are spaced along the lower half of the radial span and the filmholes <NUM> are spaced along the upper half of the radial span. The cast slots <NUM> extend in an axial direction and the filmholes <NUM> are drilled at desired locations to tailor the flow to meet cooling flow requirements. In the example shown, some filmholes <NUM> extend from a fore hole end in a radial outward angle to an aft hole end, while other filmholes <NUM> extend from the fore hole end at a radial inward angle to an aft hole end. Further, filmholes <NUM> are more closely spaced together at the radially outer end <NUM> than at a center of the body.

In the embodiments shown in <FIG>, the cast slots <NUM> are interspersed with the drilled filmholes <NUM> in the radial direction along the trailing edge <NUM> in various patterns. Further, the filmholes <NUM> have various shapes/sizes and are orientated at various angles relative to the axially extending cast slots <NUM>.

<FIG> shows an example where the cast slots <NUM> and filmholes <NUM> are formed in a turbine blade <NUM>, <NUM>. Optionally, the cast slots <NUM> and filmholes <NUM> could be formed in a trailing edge of a BOAS <NUM> as shown in <FIG>.

The invention uses both cast trailing edge slots and drilled trailing edge filmholes to achieve a beneficial arrangement. The trailing edge slots are used where core positioning is required and drilled filmholes are used in areas where cooling flow can be reduced. The filmholes can have various sizes, shapes, and spacing in order to achieve the necessary requirements.

Claim 1:
A gas turbine engine component (<NUM>) comprising:
a body (<NUM>) having a leading edge (<NUM>) and a trailing edge (<NUM>);
at least one internal channel (<NUM>) formed within the body (<NUM>), wherein the channel (<NUM>) includes an inlet to direct cooling flow into the body (<NUM>);
at least one cast slot (<NUM>) formed in the trailing edge (<NUM>); and
at least one drilled filmhole (<NUM>) formed in the trailing edge (<NUM>), wherein the cast slot (<NUM>) and drilled filmhole (<NUM>) direct flow from the internal channel (<NUM>) to an external location from the body (<NUM>), the at least one cast slot (<NUM>) comprises a plurality of cast slots (<NUM>), the at least one drilled filmhole (<NUM>) comprises a plurality of drilled filmholes (<NUM>), an axial direction is defined in a direction extending from the leading edge (<NUM>) to the trailing edge (<NUM>), and a radial direction is defined as a direction that is perpendicular to the axial direction, characterised in that:
the cast slots (<NUM>) are interspersed with the drilled filmholes (<NUM>) in the radial direction along the trailing edge (<NUM>).