Patent Description:
Gas turbine engines may include a fan for propulsion air. The fan may also deliver air into a core engine where it is compressed. The compressed air is then delivered into a combustion section, where it is mixed with fuel and ignited. The combustion gas expands downstream over and drives turbine blades. Static vanes may be positioned adjacent to the turbine blades to control the flow of the products of combustion.

Due to exposure to hot combustion gases, numerous components of a gas turbine engine, such as turbine blades and vanes, may include cooling schemes that circulate airflow to cool the component during engine operation. Thermal energy is transferred from the component to the airflow as the airflow circulates through the cooling scheme to cool the component.

<CIT> discloses a radially diffused tip flag.

<CIT> discloses a high-effectiveness cooled turbine vane or blade.

<CIT> discloses a turbine blade cooling system with a lower turning vane bank.

An airfoil for a gas turbine engine according to an example of the present invention includes an airfoil section extending in a radial direction from a root section to a tip portion. The airfoil section has an external wall and an internal wall. The external wall defines pressure and suction sides extending in a chordwise direction between a leading edge and a trailing edge, and the pressure and suction sides are spaced apart in a thickness direction between the leading edge and the trailing edge. The airfoil section establishes an internal cooling arrangement including a first cooling passage having a first section and a tip flag section. The first section extends in the radial direction from the root section. The tip flag section extends in the chordwise direction along the tip portion from the first section to the trailing edge. The first section includes a plurality of branched paths established by at least one turning vane that interconnects the internal wall and the external wall. The at least one turning vane has an arcuate profile and is arranged such that the plurality of branched paths join together along the tip flag section. The internal wall extends in the chordwise direction such that the plurality of branched paths are bounded in the thickness direction between the internal wall and the external wall adjacent the at least one turning vane.

Optionally, and in accordance with the above, the plurality of branched paths includes first, second and third branched paths, and the at least one turning vane includes a first turning vane and a second turning vane that cooperate to separate the first, second and third branched paths.

Optionally, and in accordance with any of the above, each of the first and second turning vanes extends between an upstream end and a downstream end. The upstream end of the first turning vane is aligned with a first rib relative to the chordwise direction. The upstream end of the second turning vane is aligned with a second rib relative to the chordwise direction, and the first and second ribs cooperate to separate the first, second and third branched paths.

Optionally, and in accordance with any of the above, the tip flag portion is established along a reference plane intersecting the leading and trailing edges and the pressure and suction sides, the tip flag portion expands outwardly in the thickness direction along the reference plane from the second branched path towards the trailing edge, and the internal wall follows along the second branched path in the reference plane.

Optionally, and in accordance with any of the above, the upstream end of the second turning vane establishes a first aspect ratio that is less than or equal to about <NUM>:<NUM>. The downstream end of the second turning vane establishes a second aspect ratio that is greater than or equal to about <NUM>: <NUM>. The airfoil section includes a radially inwardly facing wall and a radially outwardly facing wall extending in the chordwise direction to bound the tip flag section, and the downstream end of the second turning vane is aligned with the radially outwardly facing wall relative to the chordwise direction.

Optionally, and in accordance with any of the above, each of the first and second turning vanes is segmented between the upstream and downstream ends to establish at least one crossover passage interconnecting an adjacent pair of the branched paths.

Optionally, and in accordance with any of the above, the internal cooling arrangement includes a serpentine cooling passage including a first section, a second section and a third section. The second section interconnects the first section and the third section, and the first section extends outwardly from the root section and the third section extends inwardly from the tip portion relative to the radial direction. The tip flag section and the third section of the serpentine cooling passage are situated on opposite sides of the internal wall relative to the thickness direction.

Optionally, and in accordance with any of the above, the serpentine cooling passage is established between the internal wall and the pressure side, and the branched paths are established between the internal wall and the suction side.

Optionally, and in accordance with any of the above, the airfoil section extends in the radial direction from a platform section to the tip portion. The branched paths are dimensioned to branch outwardly from a trunk of the first section at a position inward of the platform section relative to the radial direction.

Optionally, and in accordance with any of the above, the tip flag portion includes a first set of exit ports along the trailing edge, the internal cooling arrangement includes a leading edge cooling passage bounded by the external wall along the leading edge, and a trailing edge cooling passage including a second set of exit ports along the trailing edge that are inward of the first set of exit ports relative to the radial direction.

Optionally, and in accordance with any of the above, the airfoil is a turbine blade.

A casting core assembly is provided in accordance with claim <NUM>.

Optionally, and in accordance with the above, the at least one arcuate slot extends between a first end and a second end, the first portion includes at least one elongated slot bounded by an adjacent pair of the branched sections, and the first end of the at least one arcuate slot is aligned with the at least one elongated slot relative to a chordwise direction.

Optionally, and in accordance with any of the above, the at least one arcuate slot extends between a first end and a second end. The skin core includes at least one bridge spanning between an adjacent pair of the branched sections such that the at least one arcuate slot is interrupted between the first and second ends, and the at least one bridge corresponds to at least one crossover passage interconnecting an adjacent pair of the branched paths.

Optionally, and in accordance with any of the above, the casting core assembly includes a serpentine core corresponding to a serpentine cooling passage. The skin core and the serpentine core are arranged in spaced relationship such that the first cooling passage and the serpentine cooling passage are opposite sides along an internal wall of the airfoil relative to a thickness direction.

Optionally, and in accordance with any of the above, the casting core assembly includes a leading edge core corresponding to a leading edge cooling passage bounded by an external wall along a leading edge of the airfoil. A trailing edge core corresponds to a trailing edge cooling passage including a second set of exit ports along the trailing edge of the airfoil. The tip flag portion of the skin core is at least partially aligned with the trailing edge core relative to the thickness direction, and the serpentine core is spaced apart from and forward of the trailing edge core relative to the chordwise direction.

Optionally, and in accordance with any of the above, the trailing edge core includes a second row of protrusions corresponding to a second row of exit ports along the trailing edge of the airfoil.

Optionally, and in accordance with any of the above, the casting core assembly includes at least one connector that joins the leading edge core and the serpentine core. The at least one connector corresponds to at least one crossover passage extending between the leading edge cooling passage and the serpentine cooling passage.

Optionally, and in accordance with any of the above, the skin core includes a protrusion extending from the first bend. The protrusion corresponds to a purge passage interconnecting the first cooling passage and an aperture along an external surface of the airfoil, and the purge passage is dimensioned to eject particulate from the first cooling passage in operation.

A method of forming an airfoil for a gas turbine engine according to an example of the present invention includes forming a skin core, forming a serpentine core, forming a leading edge core, and forming a trailing edge core, and assembling the skin core, the serpentine core, the leading edge core and the trailing edge core together establish a core assembly. An airfoil is formed around the core assembly. The skin core corresponds to a first cooling passage of the airfoil. The first cooling passage includes a first section and a tip flag section. The first section extends in a radial direction, and the tip flag section extends in a chordwise direction from the first section to a trailing edge of the airfoil. The leading edge core corresponds to a leading edge cooling passage adjacent to a leading edge of the airfoil. The trailing edge core corresponds to a trailing edge cooling passage adjacent to the trailing edge of the airfoil. The skin core includes at least one arcuate slot corresponding to at least one turning vane outward of the trailing edge cooling passage relative to the radial direction. The skin core includes a plurality of branched sections corresponding to a plurality of branched paths along the first section of the first cooling passage that join along the tip flag section.

Optionally, and in accordance with the above, the step of forming the airfoil includes forming an airfoil section including an external wall and an internal wall. The external wall defines pressure and suction sides extending in a chordwise direction between the leading edge and the trailing edge. The pressure and suction sides are spaced apart in a thickness direction between the leading edge and the trailing edge. The tip flag section of the first cooling passage is established between the suction side of the airfoil and a first side of the internal wall relative to the thickness direction, and the serpentine cooling passage is established between the pressure side of the airfoil and a second side of the internal wall opposed to the first side relative to the thickness direction.

Optionally, and in accordance with any of the above, the step of forming the airfoil includes forming a platform section and a root section. The airfoil section extends outwardly from the platform section to a tip portion relative to the radial direction. The root section extends inwardly from the platform section relative to the radial direction and is dimensioned to mount the airfoil to a rotatable hub. The internal wall extends inwardly from the tip portion relative to the radial direction. The tip flag section is established along the tip portion of the airfoil. The branched paths are dimensioned to branch outwardly from a trunk of the first section at a position inward of the platform section relative to the radial direction. The assembly step includes coupling the skin core, the serpentine core, the leading edge core and the trailing edge core to each other at a position corresponding to the root section.

An airfoil for a gas turbine engine according to an example of the present invention includes an airfoil section extending in a radial direction from a root section to a tip portion. The airfoil section has an external wall defining pressure and suction sides extending in a chordwise direction between a leading edge and a trailing edge, and the pressure and suction sides are spaced apart in a thickness direction between the leading edge and the trailing edge. A platform section between the root section and the tip portion is relative to the radial direction. The airfoil section establishes an internal cooling arrangement including a first cooling passage having a first section and a tip flag section joined at a junction. The first section extends in the radial direction from the root section, and the tip flag section extends in the chordwise direction along the tip portion from the junction to the trailing edge. The first section includes a plurality of branched paths dimensioned to branch outwardly from a trunk of the first section at a position inward of the platform section relative to the radial direction, and the plurality of branched paths are dimensioned to join together along the junction.

Optionally, and in accordance with the above, the plurality of branched paths includes first, second and third branched paths separated by a plurality of turning vanes at the junction.

Optionally, and in accordance with any of the above, the position includes a first position and a second position. The third branched path branches from the trunk at the first position, and the trunk divides into the first and second branched paths at the second position radially outward of the first position relative to the radial direction.

Optionally, and in accordance with any of the above, the airfoil section includes an internal wall. The internal cooling arrangement includes a serpentine cooling passage including a first section, a second section and a third section. The second section interconnects the first section and the third section, and the first section extends outwardly from the root section and the third section extends inwardly from the tip portion relative to the radial direction. The tip flag section and the third section of the serpentine cooling passage are situated on opposite sides of the internal wall relative to the thickness direction.

The engine parameters described above and those in this paragraph are measured at this condition unless otherwise specified. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about <NUM>, or more narrowly greater than or equal to <NUM>. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about <NUM> ft / second (<NUM> meters/second), and can be greater than or equal to <NUM> ft / second (<NUM> meters/second).

<FIG> illustrates an exemplary section of a gas turbine engine, such as the turbine section <NUM> of <FIG>. Although the disclosure primarily refers to the turbine section <NUM>, it should be understood that other portions of the engine <NUM> can benefit from the teachings disclosed herein, including airfoils in the compressor section <NUM>, combustor panels or liners in the combustor section <NUM>, and other portions of the engine <NUM> that may be subject to elevated temperature conditions during engine operation. Other systems can benefit from the teachings disclosed herein, including gas turbine engines lacking a fan for propulsion. In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements.

The turbine section <NUM> includes a plurality of components <NUM> arranged relative to the engine axis A, including a rotor <NUM>, one or more airfoils <NUM>, and one or more blade outer air seals (BOAS) <NUM>. Example airfoils <NUM> include rotatable blades <NUM>-<NUM> and static vanes <NUM>-<NUM>. The rotor <NUM> is coupled to a rotatable shaft <NUM> (shown in dashed lines for illustrative purposes). The shaft <NUM> can be one of the shafts <NUM>, <NUM> of <FIG>, for example. The rotor <NUM> carries one or more blades <NUM>-<NUM> that are rotatable about the engine axis A in a gas path GP, such as the core flow path C.

Each airfoil <NUM> includes an airfoil section 62A extending in a spanwise or radial direction R from a first (e.g., inner) platform section 62B. Each blade <NUM>-<NUM> extends in the radial direction R from the platform section 62B to a tip portion 62T. Each vane <NUM>-<NUM> extends in the radial direction R from the first platform section 62B to a second (e.g., outer) platform section 62C. The platform sections 62B, 62C can bound or define a portion of the gas path GP. The airfoil section 62A generally extends in a chordwise or axial direction X between a leading edge 62LE and a trailing edge 62TE, and extends in a circumferential or thickness direction T between pressure and suction sides 62P, <NUM>. The pressure and suction sides 62P, <NUM> are joined at the leading and trailing edges 62LE, 62TE to establish an aerodynamic surface contour of the airfoil <NUM>. The root section 62R of the blade <NUM>-<NUM> can be mounted to, or can be integrally formed with, the rotor <NUM>. The vane <NUM>-<NUM> can be arranged to direct or guide flow in the gas path GP from and/or towards the adjacent blade(s) <NUM>-<NUM>.

Each BOAS <NUM> can be spaced radially outward from the tip portion 62T of the blade <NUM>-<NUM>. The BOAS <NUM> can include an array of seal arc segments that are circumferentially distributed or arranged in an annulus about an array of the airfoils <NUM> to bound the gas path GP.

The turbine section <NUM> can include at least one array of airfoils <NUM>, including at least one array of blades <NUM>-<NUM> and at least one array of vanes <NUM>-<NUM>, and can include at least one array of BOAS <NUM> arranged circumferentially about the engine axis A. The array of vanes <NUM>-<NUM> are spaced axially from the array of blades <NUM>-<NUM> relative to the engine axis A. The tip portions 62T of the blades <NUM>-<NUM> and adjacent BOAS <NUM> are arranged in close radial proximity to reduce the amount of gas flow that escapes around the tip portions 62T through a corresponding clearance gap.

The turbine section <NUM> includes a cooling arrangement <NUM> for providing cooling augmentation to the components <NUM> during engine operation. The cooling arrangement <NUM> include one or more cooling cavities or plenums P1, P2 defined by a portion of the engine static structure <NUM> such as the engine case <NUM>. The plenum P2 can be at least partially defined or bounded by a rotatable portion of the engine <NUM>, such as the rotor <NUM>. One or more coolant sources CS (one shown) are configured to provide cooling air to the plenums P1, P2. The plenums P1, P2 are configured to receive pressurized cooling flow from the coolant source(s) CS to cool portions of the components <NUM> including the airfoils <NUM> and/or BOAS <NUM>. Coolant sources CS can include bleed air from an upstream stage of the compressor section <NUM> (<FIG>), bypass air, or a secondary cooling system aboard the aircraft, for example. Each of the plenums P1, P2 can extend in the circumferential direction T between adjacent airfoils <NUM> and/or BOAS <NUM>.

<FIG> and <FIG> illustrate an exemplary gas turbine engine component <NUM> including an internal cooling arrangement <NUM>. The component <NUM> can be any of the components disclosed herein, including a combustion liner or panel incorporated into the combustor section <NUM>, and the BOAS <NUM> and airfoils <NUM> such as the blades <NUM>-<NUM> and vanes <NUM>-<NUM> of the turbine section <NUM>. In the illustrative example of <FIG>, the component <NUM> is an airfoil <NUM> shown as a blade <NUM>-<NUM>. The blade <NUM>-<NUM> can be a rotatable turbine blade incorporated into one or more rows of the turbine section <NUM>.

Referring to <FIG>, the airfoil <NUM> includes an airfoil section 162A extending outwardly in a radial (e.g., first) direction R from a root section 162R to a tip portion 162T. The tip portion 162T can establish a terminal end of the airfoil section 162A. The airfoil section 162A can extend outwardly in the radial direction R from a platform section 162B to the tip portion 162T. In other examples, the airfoil <NUM> is a vane including inner and outer platform sections, as illustrated by the platform sections 62B, 62C of the vane <NUM>-<NUM> (<FIG>). The root section 162R can be dimensioned to extend inwardly from the platform section 162B relative to the radial direction R. The root section 162R can be dimensioned to mount the airfoil <NUM> to a rotatable hub.

The airfoil section 162A includes an external wall <NUM> and at least one internal wall (or rib) <NUM>, as illustrated in <FIG>. The external wall <NUM> can define pressure and suction sides 162P, <NUM> extending in a chordwise (e.g., second) direction X between a leading edge 162LE and a trailing edge 162TE. The pressure side 162P and suction side <NUM> can be spaced apart in a thickness (e.g., third) direction T between the leading edge 162LE and trailing edge 162TE.

The internal cooling arrangement <NUM> can include one or more cooling passages dimensioned to convey cooling flow to adjacent portions of the component <NUM>. The cooling arrangement <NUM> can include a skin core (e.g., first) cooling passage <NUM>, a serpentine (e.g., second) cooling passage <NUM>, a leading edge (e.g., third) cooling passage <NUM>, and a trailing edge (e.g., fourth) cooling passage <NUM>. The cooling passages <NUM>, <NUM>, <NUM>, <NUM> can be coupled to a coolant source CS (shown in dashed lines in <FIG> for illustrative purposes) to convey cooling flow to adjacent portions of the component <NUM>. It should be understood that one or more of the cooling passages <NUM>, <NUM>, <NUM> and/or <NUM> can be omitted and/or combined, and fewer or more than four cooling passages may be utilized in accordance with the teachings disclosed herein. The cooling arrangement <NUM> can be established by the airfoil section 162A, platform section 162B and/or root section 162R. The internal wall <NUM> can extend in the chordwise direction X to establish a double wall arrangement that separates portions of the cooling arrangement <NUM>, as illustrated in <FIG>.

The leading edge cooling passage <NUM> can be established adjacent to the leading edge 162LE of the airfoil <NUM>. The trailing edge cooling passage <NUM> can be established adjacent to the trailing edge 162TE of the airfoil <NUM>. The leading edge cooling passage <NUM> can be bounded by the external wall <NUM> along the leading edge 162LE (see also <FIG>). The trailing edge cooling passage <NUM> can be bounded by the external wall <NUM> along the trailing edge 162TE. The serpentine cooling passage <NUM> can be situated between the leading edge cooling passage <NUM> and trailing edge cooling passage <NUM> relative to the chordwise direction X. The skin core cooling passage <NUM> can extend aft of the leading edge cooling passage <NUM> relative to the chordwise direction X. Portions of the skin core cooling passage <NUM> can be aligned with the serpentine cooling passage <NUM>, leading edge cooling passage <NUM> and/or trailing edge cooling passage <NUM> relative to the chordwise direction X. The cooling arrangement <NUM> can include at least one crossover passage <NUM> extending between and interconnecting the serpentine cooling passage <NUM> and leading edge cooling passage <NUM>, as illustrated in <FIG> and <FIG>. In other examples, each crossover passage <NUM> is omitted.

Referring to <FIG>, with continuing reference to <FIG>, the skin core cooling passage <NUM> can include a radial (e.g., first) section 170A and a tip flag (e.g., second) section 170B joined at a first bend (or junction) 170C. The first section 170A can extend in the radial direction R from the root section 162R. The tip flag section 170B can be established along or can otherwise be adjacent to the tip portion 162T of the airfoil <NUM>. The tip flag section 170B can be dimensioned to extend in the chordwise direction X along the tip portion 162T from the first section 170A at the first bend 170C to the trailing edge 162TE of the airfoil <NUM> (see also <FIG>). The tip flag section 170B can extend transversely, such as at an approximately <NUM> degree angle, from the first section 170A at the first bend 170C. For the purposes of this disclosure, the terms "approximately," "about" and "substantially" mean ±<NUM> percent of the stated value or relationship unless otherwise indicated. Arrangement of the tip flag section 170B in a substantially axial or chordwise direction X can be utilized to maximize or otherwise increase internal convective heat transfer adjacent the tip portion 162T, which may have a relative lesser thickness than other portions of the airfoil <NUM>. The tip flag section 170B can be utilized in combination with various configurations along the tip portion 162T, including a pressure side tip shelf and/or a tip squealer pocket (see, e.g., <FIG>).

The first section 170A of the skin core cooling passage <NUM> can include one or more branched paths <NUM>. The branched paths <NUM> can include first, second, and third branched paths (indicated at <NUM>-<NUM> to <NUM>-<NUM>) that establish a trifurcation. Although three branched paths <NUM> are illustrated, it should be understood that fewer or more than three branched paths <NUM> can be utilized in accordance with the teachings disclosed herein. The branched paths <NUM> can be dimensioned to branch outwardly from a trunk <NUM> of the first section 170A of the skin core cooling passage <NUM> at a position inward of the platform section 162B relative to the radial direction R (see also <FIG>). The branched paths <NUM> serve to divide cooling flow conveyed by the trunk <NUM> to downstream portions of the skin core cooling passage <NUM>. The third branched path <NUM>-<NUM> can be dimensioned to branch or divide from the trunk <NUM> at a first position, and the trunk <NUM> can be dimensioned to divide into the first and second branched paths <NUM>-<NUM>, <NUM>-<NUM> at a second position radially outward of the first position relative to the radial direction R, as illustrated in <FIG>.

Referring back to <FIG>, the serpentine cooling passage <NUM> can include a first section <NUM>-<NUM>, second section <NUM>-<NUM>, and third section <NUM>-<NUM>. The second section <NUM>-<NUM> can be dimensioned to interconnect the first section <NUM>-<NUM> and third section <NUM>-<NUM>. The first section <NUM>-<NUM> can be dimensioned to extend outwardly from the root section 162R relative to the radial direction R. The third section <NUM>-<NUM> can be dimensioned to extend inwardly from the tip portion 162T relative to the radial direction R. The third section <NUM>-<NUM> can be dimensioned to substantially span between the platform section 162B and tip portion 162T. The first section <NUM>-<NUM> can be joined to the second section <NUM>-<NUM> at a bend <NUM>. The third section <NUM>-<NUM> can be joined to the second section <NUM>-<NUM> at another bend <NUM>. Each of the bends <NUM> can be dimensioned to turn approximately <NUM> degrees such that the bends <NUM> have a generally C-shaped geometry. The third section <NUM>-<NUM> can be forward of the first section <NUM>-<NUM> relative to the chordwise direction X, as illustrated in <FIG>, although the opposite arrangement can be utilized.

The tip flag section 170B can include a first set of ports <NUM> established along the trailing edge 162TE of the airfoil <NUM> (see also <FIG> and <FIG>). The trailing edge cooling passage <NUM> can include a second set of ports <NUM> along the trailing edge 162TE of the airfoil <NUM>. The first set of ports <NUM> and second set of ports <NUM> can be dimensioned to eject cooling flow from the respective cooling passages <NUM>, <NUM> to provide film cooling to external surfaces of the airfoil <NUM> adjacent the trailing edge 162TE. The second set of ports <NUM> can be established radially inward of the first set of ports <NUM> relative to the radial direction R. The first set of ports <NUM> and second set of ports <NUM> can be at least partially aligned in the thickness direction T (see also <FIG>).

The double wall arrangement established by the skin core cooling passage <NUM> relative to the internal wall <NUM> can serve to at least partially thermally isolate or shield adjacent portions of the serpentine cooling passage <NUM> and/or leading edge cooling passage <NUM> from elevated temperatures caused by hot gases communicated along exposed surfaces of the airfoil <NUM>, such as along the suction side <NUM> of the airfoil <NUM>. For example, the internal wall <NUM> can extend inwardly from the tip portion 162T relative to the radial direction R, as illustrated in <FIG>. The tip flag section 170B of the skin core cooling passage <NUM> can be situated on an opposite side of the internal wall <NUM> from portions of the serpentine cooling passage <NUM> and/or leading edge cooling passage <NUM> relative to the thickness direction T, as illustrated in <FIG> and <FIG>. At least the third section <NUM>-<NUM> of the serpentine cooling passage <NUM> and/or portions of the leading edge cooling passage <NUM> can be established between the inner wall <NUM> and the pressure side 162P, and the branched paths <NUM> of the first cooling passage <NUM> and/or tip flag section 170B can be established between the internal wall <NUM> and the suction side <NUM>, as illustrated in <FIG>, although the opposite arrangement can be utilized. The shielding can reduce heat pickup of cooling flow in the serpentine cooling passages <NUM> and/or leading edge cooling passage <NUM> by reducing net heat flux from the relatively hot external wall <NUM> along adjacent surfaces of the suction side <NUM>. Reduction in heat pickup of the cooling flow can improve cooling effectiveness by maximizing or otherwise increasing the potential temperature gradient between the external gasses and the relatively cooler internal cooling flow through the cooling arrangement <NUM>.

Referring to <FIG>, with continuing reference to <FIG>, the branched paths <NUM> can be established by one or more elongated ribs <NUM>. The cooling passages <NUM>, <NUM>, <NUM> are omitted from <FIG> for illustrative purposes. The ribs <NUM> can include a first rib <NUM>-<NUM> and a second rib <NUM>-<NUM> opposed to the first rib <NUM>-<NUM>. The ribs <NUM>-<NUM>, <NUM>-<NUM> can be dimensioned to extend along the first section 170A of the first cooling passage <NUM>. One or more of the ribs <NUM> can be aligned with the platform section 162B with respect to the radial direction R. The ribs <NUM>-<NUM>, <NUM>-<NUM> can be dimensioned to separate adjacent pairs of the branched paths <NUM>-<NUM>, <NUM>-<NUM>, <NUM>-<NUM>. The ribs <NUM> can serve to provide convective cooling and improve rigidity of adjacent portions of the airfoil <NUM>, including reduced compressive strain and improved distribution of shearing loads between the external wall <NUM> and internal wall <NUM>.

The branched paths <NUM> can be established by at least one or more turning vanes <NUM>. The turning vanes <NUM> can be dimensioned to convey cooling flow from the first section 170A to the tip flag section 170B of the skin core cooling passage <NUM>. The turning vanes <NUM> can serve as heat augmentation features that provide convective cooling to adjacent portions of the airfoil <NUM>.

The turning vanes <NUM> can include a first turning vane <NUM>-<NUM> and a second turning vane <NUM>-<NUM> that opposes the first turning vane <NUM>-<NUM>. It should be understood that fewer or more than two turning vanes <NUM> can be utilized in accordance with the teachings disclosed herein. The component <NUM> can include other heat augmentation features at various positions along the cooling arrangement <NUM>, such as pedestals, trip strips, fins, dimples, raised protrusions, etc., to meter flow and/or provide convective cooling to adjacent portions of the component <NUM>. For example, the skin core cooling passage <NUM> can include one or more rows of pedestals <NUM>. The pedestals <NUM> can be dimensioned to span between opposed surfaces of the tip flag section 170B (see <FIG>). The component <NUM> can also include film cooling holes coupled to various positions along the cooling arrangement <NUM> to provide film cooling augmentation.

Each of the turning vanes <NUM> can be dimensioned to span between opposed walls bounding the skin core cooling passage <NUM>. For example, each of the turning vanes <NUM> can be dimensioned to interconnect the internal wall <NUM> and external wall <NUM>, as illustrated in <FIG>, which can improve rigidity of adjacent portions of the airfoil <NUM>, including along the tip portion 162T adjacent the bend 170C.

The turning vanes <NUM> can have various geometries to direct flow along the skin core cooling passage <NUM>. Each of the turning vanes <NUM> can include a main body 174A extending between a first (e.g., upstream) end 174B and a second (e.g., downstream) end 174C. The turning vane <NUM> can be dimensioned such that a length of the main body 174A between the ends 174B, 174C has a substantially arcuate shaped profile. The arcuate profile of the turning vanes <NUM> can reduce turbulence and separation, improve filling of the cooling passage <NUM> with cooling flow, and reduce dead zones or stagnation through adjacent portions of the skin core cooling passage <NUM>, including through the first bend 170C, which can improve cooling effectiveness and component durability. The arcuate profile can be a simple curve or compound curve established by one or more radii. The turning vanes <NUM> can be dimensioned to extend between approximately <NUM> degrees and approximately <NUM> degrees about a respective point to establish the arcuate profile. Other geometries can be utilized, such as one or more linear segments joined at an angle.

The turning vanes <NUM> can be arranged at various positions and orientations relative to each other, the cooling arrangement <NUM> and the airfoil <NUM>. The turning vanes <NUM> can be dimensioned such that the upstream ends 174B are axially forward of the downstream ends 174C relative to the chordwise direction X. The turning vanes <NUM> can be dimensioned such that the upstream ends 174B are inward of the downstream ends 174C with respect to the radial direction R. The turning vanes <NUM> and bend 170C can be outward of the trailing edge cooling passage <NUM> relative to the radial direction R (see <FIG>).

The turning vanes <NUM> can cooperate to substantially or completely fluidly separate the branched paths <NUM>. Each of the turning vanes <NUM> can be arranged such that the branched paths <NUM> extend along the respective turning vanes <NUM> and then join together along the tip flag section 170B to diffuse the cooling flow. The internal wall <NUM> can be dimensioned to extend in the chordwise direction X such that portions of the branched paths <NUM> adjacent to the turning vanes <NUM> are bounded in the thickness direction T between the internal wall <NUM> and the external wall <NUM>, as illustrated in <FIG>.

The upstream ends 174B of the turning vanes <NUM> can be aligned with a respective one of the ribs <NUM> relative to the chordwise direction X, which can serve to reduce turbulence through the skin core cooling passage <NUM>. For example, the upstream end 174B of the first turning vane <NUM>-<NUM> can be at least partially aligned with the first rib <NUM>-<NUM> relative to the chordwise direction X. The upstream end 174B of the second turning vane <NUM>-<NUM> can be at least partially aligned with the second rib <NUM>-<NUM> relative to the chordwise direction X. Aligning the turning vanes <NUM> and ribs <NUM> can reduce losses that may otherwise be caused by turbulence.

The main body 174A of each turning vane <NUM> can be continuous between the first and second ends 174B, 174C to fluidly isolate adjacent portions of the branched paths <NUM>. In the illustrative example of <FIG>, each turning vane <NUM> of component <NUM> can be segmented or interrupted between upstream and downstream ends 274B, 274C to establish at least one crossover passage <NUM>. Each crossover passage <NUM> can be dimensioned to interconnect an adjacent pair of the branched paths <NUM>. The crossover passages <NUM> can serve to increase flow from a radially inward one of the branched paths <NUM> to a radially outward one of the branched path <NUM>, which may be assisted by centrifugal forces caused by rotation of the airfoil <NUM> during engine operation.

The skin core cooling passage <NUM> can include at least one purge passage <NUM>. The serpentine cooling passage <NUM> can include at least one purge passage <NUM> (<FIG>). Each purge passage <NUM>, <NUM> can interconnect the respective cooling passage <NUM>, <NUM> and a respective aperture along an external surface of the airfoil <NUM>. For example, the purge passage <NUM> can extend from the first bend 170C. The purge passage <NUM> can extend from the third section <NUM>-<NUM> of the serpentine cooling passage <NUM>. The purge passages <NUM>, <NUM> can be dimensioned to eject particulate from the cooling passages <NUM>, <NUM> in operation, which can reduce a likelihood of blockage such as through the first bend 170C of the skin core cooling passage <NUM>. The purge passage <NUM> can also reduce flow separation and recirculating flows in the first bend 170C of the skin core cooling passage <NUM>. The purge passages <NUM>, <NUM> can be formed utilizing various techniques, including a casting or drilling operation.

The tip flag section 170B of the cooling passage <NUM> can be dimensioned to expand or flair outwardly from the branched paths <NUM> to diffuse cooling flow communicated from the branched paths <NUM>. For example, the tip flag section 170B can be established along a reference plane REF (illustrated in dashed lines in <FIG> for illustrative purposes). The reference plane REF can intersect the leading and trailing edges 162LE, 162TE and the pressure and suction sides 162P, <NUM> of the airfoil <NUM>. The tip flag section 170B can be dimensioned to expand or flair outwardly in the thickness direction T along the reference plane REF from the first, second and/or third branched paths <NUM>-<NUM>, <NUM>-<NUM>, <NUM>-<NUM> towards the trailing edge 162TE, as illustrated by the contouring of the branched paths <NUM>-<NUM>, <NUM>-<NUM> of <FIG>. The internal wall <NUM> can be dimensioned to follow at least the first and/or second branched paths <NUM>-<NUM> in the reference plane REF, as illustrated in <FIG>.

Referring to <FIG>, with continuing reference to <FIG>, the turning vanes <NUM> can be dimensioned to diffuse or otherwise communicate cooling flow through the branched paths <NUM> and towards the tip flag section 170B. Each of the turning vanes <NUM> can be dimensioned to establish one or more aspect ratios to establish the cooling arrangement <NUM>. For the purposes of this disclosure, the term "aspect ratio" means a ratio of a width to a height at a position along the turning vane <NUM>. The height dimension can have a major component in the radial direction R or axial direction X. The width dimension can have a major component in the thickness direction T. The width can be established by a span of the turning vane <NUM> between the external wall <NUM> and internal wall <NUM>. The aspect ratio may be the same or may differ at positions along a length of the main body 174A of the turning vane <NUM>. For example, the aspect ratio of the first turning vane <NUM>-<NUM> can be substantially constant between the upstream end 174B and downstream end 174C.

The aspect ratio of the second turning vane <NUM>-<NUM> can be substantially constant or can vary between the upstream end 174B and downstream end 174C. For example, the upstream end 174B of the second turning vane <NUM>-<NUM> can be dimensioned to establish a first aspect ratio W1:H1 defined by a first width W1 (<FIG>) and a first height H1 (<FIG>). The downstream end 174C of the second turning vane <NUM>-<NUM> can be dimensioned to establish a second aspect ratio W2:H2 defined by a second width W2 (<FIG>) and a second height H2 (<FIG>). The widths W1, W2 and heights H1, H2 can be defined as a maximum dimension at the respective upstream and downstream ends 174B, 174C excluding any radiusing. The second aspect ratio W2:H2 can be equal to or greater than the first aspect ratio W1:H1. For example, the first aspect ratio W1:H1 can be less than or equal to about <NUM>:<NUM>, or more narrowly greater or equal to about <NUM>:<NUM>. The second aspect ratio W2:H2 can be greater than or equal to about <NUM>:<NUM>, or more narrowly greater than or equal to about <NUM>: <NUM>. The first height H1 and second height H2 can be approximately equal. A ratio W2:W1 of the second width W2 to the first width W1 can be greater than or equal to <NUM>:<NUM>, or more narrowly can be greater than or equal to <NUM>:<NUM>, such that the second turning vane <NUM>-<NUM> flairs or extends outwardly in a direction along a length of the main body 174A from the first end 174B towards the second end 174C, as illustrated by <FIG>. The diffusion scheme can serve to reduce the expansion ratio through the skin core cooling passage <NUM> with respect to a first position (e.g., inlets) immediately upstream of the turning vanes <NUM> and a second position (e.g., outlets) immediately downstream of the turning vanes <NUM>.

The airfoil section 162A includes a radially inwardly facing wall <NUM> and a radially outwardly facing wall <NUM> that opposes the radially inward facing wall <NUM>, as illustrated in <FIG> and <FIG>. The walls <NUM>, <NUM> are dimensioned to extend in the chordwise direction X to bound the skin core cooling passage <NUM> along the tip flag section 170B. The turning vanes <NUM> can be spaced apart from the walls <NUM>, <NUM>. The downstream end 174C of the second turning vane <NUM>-<NUM> can be aligned with the radially outward facing wall <NUM> relative to the chordwise direction X, which can serve to improve flow attachment through the third branched path <NUM>-<NUM> along the radially outward facing wall <NUM> and can reduce a likelihood of dead zones through the skin core cooling passage <NUM>.

The downstream end 174C of the first turning vane <NUM>-<NUM> can be radially outward of, or can otherwise be offset from, the downstream end 174C of the second turning vane <NUM>-<NUM> relative to the radial direction R, which can reduce a likelihood of dead zones in the skin core cooling passage <NUM>. The downstream end 174C of the first turning vane <NUM>-<NUM> can be offset from the downstream end 174C of the second turning vane <NUM>-<NUM> relative to the chordwise direction X. Offsetting or staggering the downstream ends 174C of the turning vanes <NUM>-<NUM>, <NUM>-<NUM> utilizing the techniques disclosed herein can improve load distribution by reducing an area in which the airfoil <NUM> is unsupported across the skin core cooling passage <NUM>, including adjacent the tip portion 162T which can have a relatively lesser thickness than other portions of the airfoil section 162A.

The downstream end 174C of the second turning vane <NUM>-<NUM> can be arranged relative to the radially inwardly facing wall <NUM> and radially outwardly facing wall <NUM> relative to the radial direction R. Referring to <FIG>, a first distance D1 can be established between the downstream end 174C of the second turning vane <NUM>-<NUM> and the radially inwardly facing wall <NUM>. A second distance D2 can be established between the downstream end 174C of the second turning vane <NUM>-<NUM> and the radially outwardly facing wall <NUM>. The first and second distances D1, D2 can be established as the minimum distances between the downstream end 174C of the second turning vane <NUM>-<NUM> and the respective walls <NUM>, <NUM>. A ratio D2:D1 of the second distance D2 divided by the first distance D1 can be between approximately <NUM>:<NUM> and <NUM>:<NUM>, such as about <NUM>:<NUM>. The disclosed ratio D2:D1 can be utilized to reduce flow separation through the third branched path <NUM>-<NUM> and along the radially outwardly facing wall <NUM>.

The turning vanes <NUM> can establish a set of exits of the branched paths <NUM>, as illustrated by exits E1, E2, E3 in <FIG> (shown in dashed lines for illustrative purposes). A first exit E1 of the first branched path <NUM>-<NUM> can be established by a minimum distance between the downstream end 174C of the first turning vane <NUM>-<NUM> and the radially inwardly facing wall <NUM>. A second exit E2 of the second branched path <NUM>-<NUM> can be established by a minimum distance between the downstream end 174C of the first turning vane <NUM>-<NUM> and an adjacent portion of the second turning vane <NUM>-<NUM>. A third exit E3 of the third branched path <NUM>-<NUM> can be established by a minimum distance between the downstream end 174C of the second turning vane <NUM>-<NUM> and the radially outwardly facing wall <NUM>. The turning vanes <NUM> can be dimensioned such that the cross sectional areas of the exits E1-E3 are within <NUM> percent or <NUM> percent of each other, or more narrowly can be substantially equal to each other. Utilizing the disclosed cross sectional areas of the exits E1-E3 disclosed herein, losses can be reduced by limiting an amount of acceleration of cooling flow through the branched paths <NUM> prior to diffusing the cooling flow along the tip flag section 170B.

In operation, cooling flow can be conveyed by the coolant source CS (<FIG>) to the cooling passages <NUM>, <NUM>, <NUM>, <NUM> at one or more inlets or plenums established in the root section 162R. The cooling flow can be communicated to downstream portions of the cooling passages <NUM>, <NUM>, <NUM>, <NUM>, including through the branched paths <NUM> and tip flag section 170B of the skin core cooling passage <NUM>, to provide cooling augmentation to adjacent portions of the component <NUM>. At least one of the serpentine cooling passage <NUM> and/or leading edge cooling passage <NUM> can be at least partially thermally shielded by the skin core cooling passage <NUM> and internal wall <NUM>. After picking up heat due to convective heat transfer the cooling airflow can be ejected from the component <NUM> into the adjacent gas path and/or can be communicated to another portion of the engine.

<FIG> illustrate a casting core assembly <NUM> for a gas turbine engine component. The casting core assembly <NUM> may be utilized to establish any of the cooling arrangements or schemes disclosed herein, including the internal cooling arrangement <NUM>. The casting core assembly <NUM> can include a skin (e.g., first) core <NUM>, serpentine (e.g., second) core <NUM>, leading edge (e.g., third) core <NUM>, and/or (e.g., fourth) trailing edge core <NUM>. It should be understood that one or more of the cores <NUM>, <NUM>, <NUM> and/or <NUM> can be omitted and/or combined, and fewer or more than four cores may be utilized in accordance with the teachings disclosed herein. The cores <NUM>, <NUM>, <NUM> and/or <NUM> can be utilized to form one or more cooling passages in the gas turbine engine component to convey cooling flow during operation.

The cores <NUM>, <NUM>, <NUM>, <NUM> can be arranged at various positions and orientations relative to each other. For example, the serpentine core <NUM> and/or skin core <NUM> can be aft of the leading edge core <NUM> relative to a chordwise (or first) direction X. The serpentine core <NUM> can be spaced apart from, and can be forward of, the trailing edge core <NUM> relative to the chordwise direction X.

The skin core <NUM> can correspond to a first cooling passage of a gas turbine engine component, such as the skin core cooling passage <NUM>. The serpentine core <NUM> can correspond to a second cooling passage of the gas turbine engine component, such as the serpentine cooling passage <NUM>. The leading edge core <NUM> can correspond to a third cooling passage of the gas turbine engine component, such as the leading edge cooling passage <NUM>. An isolated view of the serpentine core <NUM> and leading edge core <NUM> is illustrated in <FIG>. The trailing edge core <NUM> can correspond to a fourth cooling passage of the gas turbine engine component, such as the trailing cooling passage <NUM>. An isolated view of the skin core <NUM> and trailing edge core <NUM> is illustrated in <FIG>.

Referring to <FIG>, with continuing reference to <FIG>, the skin core <NUM> can include a first portion 386A and a tip flag portion 386B extending from the first portion 386A at a first bend 386F. The first portion 386A can correspond to the first section 170A, and the tip flag portion 386B can correspond to the tip flag section 170B of the skin core cooling passage <NUM>. The tip flag portion 386B can be dimensioned to expand outwardly from the first bend 386F to at least partially wrap about an adjacent portion of the serpentine core <NUM>, as illustrated by <FIG>.

The tip flag portion 386B can be at least partially aligned with the trailing edge core <NUM> relative to the thickness direction T. For example, the tip flag portion 386B can include a first row of protrusions 386C. The first row of protrusions 386C can correspond to the first row of exit ports <NUM>. The trailing edge core <NUM> can include a second row of protrusions 392C. The second row of protrusions 392C can correspond to the second row of exit ports <NUM>. The skin core <NUM> and trailing edge core <NUM> can be arranged such that the first row of protrusions 386C are substantially aligned with the second row of protrusions 392C relative to a thickness direction T (see also <FIG>).

The skin core <NUM> can include at least one protrusion <NUM> extending from the first bend 386F. Each protrusion <NUM> can correspond to a respective purge passage <NUM>.

The tip flag portion 386B of the skin core <NUM> and the trailing edge core <NUM> can be coupled to each other by a connector <NUM>. In other examples, the connector <NUM> is omitted.

The skin core <NUM> can include at least one or more arcuate slots <NUM>. The arcuate slots <NUM> can correspond to one or more of the turning vanes <NUM>. Each of the arcuate slots <NUM> can be dimensioned to extend between a first end 386SB and a second end 386SC to establish a substantially arcuate shaped profile. The skin core <NUM> can include a plurality of branched sections 386D along the first section 386A. Each of the branched sections 386D can correspond to a respective one of the branched paths <NUM> along the first section 170A of the skin core cooling passage <NUM>. The branched sections 386D can be dimensioned to bound one or more of the arcuate slots <NUM> and then join along the tip flag portion 386B.

The arcuate slot <NUM> can be continuous or uninterrupted between the first end 386SB and second end 386SC. In the illustrative example of <FIG>, skin core <NUM> includes at least one or more bridges 486E spanning between an adjacent pair of branched sections 486D such that an arcuate slot <NUM> is interrupted between the first and second ends 486SB, 486SC (see also <FIG>). Each of the bridges 486E can correspond to a respective one of the crossover passages <NUM> (<FIG>).

Still referring to <FIG>, the first portion 386A of the skin core <NUM> can include at least one or more elongated slots 386R bounded by an adjacent pair of the branched sections 386D. Each of the elongated slots 386R can correspond to respective one of the ribs <NUM>. The first end 386SB of each of the arcuate slots <NUM> can be aligned with a respective one of the elongated slots 386R relative to the chordwise direction X.

The skin core <NUM> and serpentine core <NUM> can be arranged in a spaced relationship such that the first cooling passage <NUM> and serpentine cooling passage <NUM> established by the respective skin core <NUM> and serpentine core <NUM> are established on, and extend along, opposite sides of the internal wall <NUM> of the airfoil <NUM> relative to the thickness direction T (see e.g., <FIG>).

Referring to <FIG>, with continuing reference to <FIG>, the casting core assembly <NUM> can include at least one connector <NUM> that joins the leading edge core <NUM> and serpentine core <NUM>. Each connector <NUM> can correspond to a respective crossover passage <NUM> (<FIG>). The leading edge core <NUM> and serpentine core <NUM> can be coupled to each other by a second connector <NUM>, which can correspond to a position external to the resultant gas turbine engine component. The connectors <NUM>, <NUM> can serve to fix or otherwise limit relative movement between the leading edge core <NUM> and serpentine core <NUM> during formation of the respective component and can simplify positioning of the cores <NUM>, <NUM> as a unit.

Referring to <FIG>, cores <NUM>, <NUM>, <NUM>, <NUM> can be joined together by a third connector <NUM> to establish a core assembly <NUM>. The third connector <NUM> can be dimensioned to establish a plenum in the root section 162R of the airfoil <NUM>, which can be coupled to the coolant source CS to convey cooling flow to the corresponding cooling passages <NUM>, <NUM>, <NUM>, <NUM> (see <FIG>). The leading edge core <NUM> and serpentine core <NUM> can be coupled to each other by a second connector <NUM>. The connectors <NUM>, <NUM> can serve to improve positioning of the core assembly <NUM> during formation of the respective gas turbine engine component. Joining the cores <NUM>/<NUM>, <NUM>/<NUM>, <NUM>/<NUM> and/or <NUM>/<NUM> together utilizing the techniques disclosed herein, including during a core injection and manufacturing process, can improve casting process capability by improving internal and external wall control, relative core displacement and core true position tolerance during wax injection and subsequent metal pour operations during the investment casting process.

<FIG> illustrates a method in a flowchart <NUM> for forming a component for a gas turbine engine. Method <NUM> can be utilized to form any of the gas turbine engine components disclosed herein, including the components <NUM>, <NUM> (e.g., airfoil <NUM>) and cooling arrangements <NUM>, <NUM>. Method <NUM> can be utilized with any of the core assemblies disclosed herein, including core assembly <NUM>, <NUM>. Fewer or additional steps than are recited below can be performed within the scope of this disclosure, and the recited order of steps is not intended to limit this disclosure. Reference is made to the component <NUM> and core assembly <NUM> for illustrative purposes.

At step 598A, a core assembly <NUM> is formed. Step 598A can include forming one or more cores to establish the core assembly <NUM>. For example, a skin (or first) core <NUM> can be formed at step 598AA. The skin core <NUM> can correspond to a first cooling passage of the gas turbine engine component, such as the skin core cooling passage <NUM>. A serpentine (or second) core <NUM> can be formed at step 598AB. The serpentine core <NUM> can correspond to a second cooling passage of the gas turbine engine component, such as the serpentine cooling passage <NUM>. A leading edge (or third) core <NUM> can be formed at step 598AC. The leading edge core <NUM> can correspond to a third cooling passage of the gas turbine engine component, such as the leading edge cooling passage <NUM>. In examples, steps 598AB and 598AC are performed concurrently such that the serpentine core <NUM> and leading edge core <NUM> are coupled with one or more connectors <NUM>, <NUM> to establish a unitary component (<FIG> and <FIG>). A trailing edge (or fourth) core <NUM> can be formed at step 598AD. The trailing edge core <NUM> can correspond to a fourth cooling passage of the gas turbine engine component, such as the trailing edge cooling passage <NUM>.

Various techniques can be utilized to form the casting core assembly <NUM> including each of the cores <NUM>, <NUM>, <NUM>, <NUM>. Exemplary techniques can include core die tooling, injection molding, flexible tooling, fugitive core, lithographic tooling, and/or advanced additive manufacturing processes. Other techniques can include laser powder bed metal fusion additive manufacturing techniques such as direct metal laser sintering (DMLS) and selective laser sintering (SLS) processes. Various materials can be utilized to form the casting core assembly <NUM> including the cores <NUM>, <NUM>, <NUM>, <NUM>. Exemplary materials include non-metallic materials such as ceramics and metallic materials such as refractory metals. Materials forming the respective cores <NUM>, <NUM>, <NUM>, <NUM> can be the same or can differ.

Step 598A can include assembling the cores <NUM>, <NUM>, <NUM> and/or <NUM> together to establish the core assembly <NUM> at step 598AE. The cores <NUM>, <NUM>, <NUM> and/or <NUM> can be formed as separate and distinct components prior to assembling the cores <NUM>, <NUM>, <NUM>, <NUM> at step 598AE. Step 598AE can include coupling the cores <NUM>, <NUM>, <NUM> and/or <NUM> to each other at a position corresponding to a root section 162R of the airfoil R, as illustrated by the core assembly <NUM> (<FIG>). In examples, two or more of the cores <NUM>, <NUM>, <NUM>, <NUM> can be injection molded or otherwise formed in a single die.

At step 598B, the component <NUM> (e.g., airfoil <NUM>) is fabricated or otherwise formed around the core assembly <NUM>. Step 598B can utilize an investment casting technique in which the core assembly <NUM> is situated in a mold. The core assembly <NUM> can be coated with a wax material to establish a predetermined component geometry. The wax material can be coated with another material, such as a metallic or ceramic slurry that can be hardened into a shell. The wax material can be melted out of the shell and molten material such as a metal or metal alloy can be deposited into the resultant cavity. Various materials can be utilized to form the component <NUM>, including metallic materials. Exemplary metallic materials can include metal and metal alloys such as a high temperature nickel alloy. The deposited metal material can solidify to form the component <NUM>. The core assembly <NUM> can be leached out or otherwise removed to establish the cooling arrangement <NUM> within the component <NUM>, and the shell can be removed. Investment casting techniques are generally known, but utilizing investment casting techniques to form the components and cooling arrangements disclosed herein is not known.

Step 598B can include forming the airfoil section 162A of the airfoil <NUM> at step 598BA. Step 598BA can include forming the airfoil section 162A including the external wall <NUM> and internal wall <NUM> to establish a double wall arrangement. The double wall arrangement can improve local thermal cooling effectiveness of the component <NUM>. Step 598B can occur such that the tip flag section 170B of the skin core cooling passage <NUM> is established between the suction side <NUM> of the airfoil <NUM> and a first side of the internal wall <NUM> relative to the thickness direction T, and such that the serpentine cooling passage <NUM> is established between the pressure side 162P of the airfoil <NUM> and a second side of the internal wall <NUM> opposed to the first side of the internal wall <NUM> relative to the thickness direction T, as illustrated in <FIG>.

Step 598B can include forming the root section 162R of the airfoil <NUM> at step 598BB and/or forming one or more platform sections of the airfoil <NUM> at step 598BC. The platform sections can include inner and/or outer platform sections, such as the platform section 162B of the airfoil <NUM>.

One or more finishing operations may be performed at step 598C. Exemplary finishing operations can include heat treating the component <NUM>, milling or grinding operation to establish a predetermined geometry of the component <NUM>, electrical discharge machining (EDM) and/or laser drilling cooling holes in the component, and depositing one or more coatings onto internal and/or external surfaces of the component <NUM> such as a thermal barrier coating (TBC).

It should be understood that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," and the like are with reference to the normal operational altitude of the engine and should not be considered otherwise limiting.

Claim 1:
An airfoil (<NUM>; <NUM>; <NUM>), optionally a turbine blade (<NUM>-<NUM>), for a gas turbine engine (<NUM>) comprising:
an airfoil section (62A; 162A; 262A) extending in a radial direction (R) from a root section (62R; 162R) to a tip portion (62T; 162T; 262T), the airfoil section (62A; 162A; 262A) having an external wall (<NUM>) and an internal wall (<NUM>), the external wall (<NUM>) defining pressure and suction sides (62P, <NUM>; 162P, <NUM>) extending in a chordwise direction (X) between a leading edge (62LE; 162LE; 262LE) and a trailing edge (62TE; 162TE; 262TE), and the pressure and suction sides (62P, <NUM>;162P, <NUM>) spaced apart in a thickness direction (T) between the leading edge (62LE; 162LE; 262LE) and the trailing edge (62TE; 162TE; 262TE), wherein the airfoil section (62A; 162A; 262A) establishes an internal cooling arrangement (<NUM>; <NUM>) comprising:
a first cooling passage (<NUM>) including a first section (170A; 270A) and a tip flag section (170B; 270B), the first section (170A; 270A) extending in the radial direction (R) from the root section (62R; 162R), and the tip flag section (170B; 270B) extending in the chordwise direction (X) along the tip portion (62T; 162T; 262T) from the first section (170A; 270A) to the trailing edge (62TE; 162TE; 262TE), characterised in that:
the first section (170A; 270A) includes a plurality of branched paths (<NUM>) established by at least one turning vane (<NUM>; <NUM>) that interconnects the internal wall (<NUM>) and the external wall (<NUM>), the at least one turning vane (<NUM>; <NUM>) having an arcuate profile and arranged such that the plurality of branched paths (<NUM>; <NUM>) join together along the tip flag section (170B; 270B); and
the internal wall (<NUM>) extends in the chordwise direction (X) such that the plurality of branched paths (<NUM>; <NUM>) are bounded in the thickness direction (T) between the internal wall (<NUM>) and the external wall (<NUM>) adjacent the at least one turning vane (<NUM>; <NUM>).