Patent Description:
In gas turbine engines, compressed air discharged from a compressor section and fuel introduced from a source of fuel are mixed together and burned in a combustion section, creating combustion products defining a high temperature and high pressure working gas. The working gas is directed through a hot gas path in a turbine section of the engine, where the working gas expands to provide rotation of a turbine rotor. The turbine rotor may be linked to an electric generator, wherein the rotation of the turbine rotor can be used to produce electricity in the generator.

In view of high pressure ratios and high engine firing temperatures implemented in modern engines, certain components, such as airfoils, e.g., stationary vanes and rotating blades within the turbine section, must be cooled with cooling fluid, such as air discharged from a compressor in the compressor section, to prevent overheating of the components. In order to push gas turbine efficiencies even higher, there is a continuing drive to reduce coolant consumption in the turbine. For example, it is known to form turbine blades and vanes of ceramic matrix composite (CMC) materials, which have higher temperature capabilities than conventional superalloys, which makes it possible to reduce consumption of compressor air for cooling purposes.

Effective cooling of turbine airfoils requires delivering the relatively cool air to critical regions such as along the trailing edge of a turbine blade or a stationary vane. The associated cooling apertures may, for example, extend between an upstream, relatively high pressure cavity within the airfoil and one of the exterior surfaces of the turbine blade. Blade cavities typically extend in a radial direction with respect to the rotor and stator of the machine. Achieving a high cooling efficiency based on the rate of heat transfer is a significant design consideration in order to minimize the volume of coolant air diverted from the compressor for cooling.

The trailing edge of a turbine airfoil is made relatively thin for aerodynamic efficiency. The relatively narrow trailing edge portion of a gas turbine airfoil may include, for example, up to about one third of the total airfoil external surface area. Turbine airfoils are often manufactured by a casting process involving a casting core, typically made of a ceramic material. The core material represents the hollow flow passages inside turbine airfoil. It is beneficial for the casting core to have sufficient structural strength to survive through the handling during the casting process. To this end, the coolant exit apertures at the airfoil trailing edge may be designed to have larger dimensions near the root and the tip of the airfoil, to form a stronger picture frame like configuration, which may result in higher coolant flow near the airfoil root and tip than desired.

In <CIT> a turbine nozzle for a gas turbine engine is disclosed which includes a hollow airfoil vane including a first wall, a second wall, and a trailing edge cavity with a pin bank formed of axially spaced rows of pins, with core strengtheners extending through the pin bank and with rows of turbulators extending between the pin bank and the trailing edge cooling slots. Further, in <CIT> a cast turbine blade has a trailing edge cooling circuit with chord-wise spaced apart radial rows of radially interspaced cylindrical pins extending from the pressure side wall to the suction side wall, whereby adjacent the platform pedestal stubs extending from both the pressure and the suction side wall are provided to reduce stress concentrations in that area. In <CIT> an airfoil for a gas turbine engine is disclosed which includes pressure and suction surfaces that are provided by pressure and suction walls extending in a radial direction and joined at a leading edge and a trailing edge. A cooling passage is arranged between the pressure and suction walls and extending to the trailing edge. In <CIT> a turbine blade is disclosed which has a leaf blade through which a hot gas is flowable. A throttle element is equipped with two projections at respective openings with respect to a flow direction of a channel. Each projection is attached to one of two surfaces arranged in an inner-facing manner. In <CIT> a blade or a vane component for a turbomachine is disclosed. Further, <CIT> discloses an airfoil for use in a turbomachine such as a stationary vane in a gas turbine. The airfoil has a plurality of longitudinally extending ribs in its trailing edge region that form first cooling fluid passages extending from the airfoil cavity to the trailing edge of the airfoil. The first cooling fluid passages are tapered so that their height and width decrease as they extend toward the trailing edge. In <CIT> a turbine blade is disclosed capable of being cooled by a coolant gas supplied to a hollow region. A plurality of meandering flow paths that guide the coolant gas between the suction wall surface and the pressure wall surface while causing the coolant gas to repeatedly meander are continuously arranged from the hub side toward the tip side of the turbine blade, and the meandering flow paths adjacent to each other cause the coolant gas to meander in different repetitive patterns.

It is desirable to have an improvement to achieve not only a strong casting core but also a limitation in the coolant flow.

Briefly, aspects of the present invention provide a turbine blade having an airfoil with trailing edge framing features.

According a first aspect of the present invention, the present invention provides a turbine blade having an airfoil having the features of claim <NUM>.

According a second aspect of the present invention, the present invention provides a casting core for forming a turbine blade having an airfoil having the features of claim <NUM>.

The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention.

In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the scope of the present invention as defined by the claims.

In the drawings, the direction X denotes an axial direction parallel to an axis of the turbine engine, while the directions R and T respectively denote a radial direction and a tangential (or circumferential) direction with respect to said axis of the turbine engine.

Referring now to <FIG>, a turbine airfoil <NUM> is illustrated according to one embodiment. As illustrated, the airfoil <NUM> is a turbine blade for a gas turbine engine. It should however be noted that aspects of the invention could additionally be incorporated into stationary vanes in a gas turbine engine. The airfoil <NUM> includes an outer wall <NUM> adapted for use, for example, in a high pressure stage of an axial flow gas turbine engine. The outer wall <NUM> delimits a hollow interior <NUM> (see <FIG>). The outer wall <NUM> extends span-wise along a radial direction R of the turbine engine and includes a generally concave shaped pressure sidewall <NUM> and a generally convex shaped suction sidewall <NUM>. The pressure sidewall <NUM> and the suction sidewall <NUM> are joined at a leading edge <NUM> and at a trailing edge <NUM>. The outer wall <NUM> may be coupled to a root <NUM> at a platform <NUM>. The root <NUM> may couple the turbine airfoil <NUM> to a disc (not shown) of the turbine engine. The outer wall <NUM> is delimited in the radial direction by a radially outer airfoil end face (airfoil tip cap) <NUM> and a radially inner airfoil end face <NUM> coupled to the platform <NUM>.

Referring to <FIG>, a chordal axis <NUM> may be defined extending centrally between the pressure sidewall <NUM> and the suction sidewall <NUM>. In this description, the relative term "forward" refers to a direction along the chordal axis <NUM> toward the leading edge <NUM>, while the relative term "aft" refers to a direction along the chordal axis <NUM> toward the trailing edge <NUM>. As shown, internal passages and cooling circuits are formed by radial coolant cavities 41a-f that are created by internal partition walls or ribs 40a-e which connect the pressure and suction sidewalls <NUM> and <NUM> along a radial extent. In the present example, coolant may enter one or more of the radial cavities 41a-f via openings provided in the root of the blade <NUM>, from which the coolant may traverse into adjacent radial coolant cavities, for example, via one or more serpentine cooling circuits. Examples of such cooling schemes are known in the art and will not be further discussed herein. Having traversed the radial coolant cavities, the coolant may be discharged from the airfoil <NUM> into the hot gas path, for example via exhaust orifices <NUM>, <NUM> located along the leading edge <NUM> and the trailing edge <NUM> respectively. Although not shown in the drawings, exhaust orifices may be provided at multiple locations, including anywhere on the pressure sidewall <NUM>, suction sidewall <NUM>, and the airfoil tip <NUM>.

The aft-most radial coolant cavity 41f, which is adjacent to the trailing edge <NUM>, is referred to herein as the trailing edge coolant cavity 41f. Upon reaching the trailing edge coolant cavity 41f, the coolant may traverse axially through an internal arrangement <NUM> of trailing edge cooling features, located in the trailing edge coolant cavity 41e, before leaving the airfoil <NUM> via coolant exit slots <NUM> arranged along the trailing edge <NUM>. Conventional trailing edge cooling features included a series of impingement plates, typically two or three in number, arranged next to each other along the chordal axis. However, this arrangement provides that the coolant travels only a short distance before exiting the airfoil at the trailing edge. It may be desirable to have a longer coolant flow path along the trailing edge portion to have more surface area for transfer of heat, to improve cooling efficiency and reduce coolant flow requirement.

The present embodiment, as particularly illustrated in <FIG>, provides an improved arrangement of trailing edge cooling features. In this case, the impingement plates are replaced by an array of cooling features embodied as pins <NUM>. Each feature or pin <NUM> extends all the way from the pressure sidewall <NUM> to the suction sidewall <NUM> as shown in <FIG>. The features <NUM> are arranged in radial rows as shown in <FIG>. The features <NUM> in each row are interspaced to define axial coolant passages <NUM>, with each coolant passage <NUM> extending all the way from the pressure sidewall <NUM> to the suction sidewall <NUM>. The rows, in this case fourteen in number, are spaced along the chordal axis <NUM> to define radial coolant passages <NUM>.

The features <NUM> in adjacent rows are staggered in the radial direction. The axial coolant passages <NUM> of the array are fluidically interconnected via the radial flow passages <NUM>, to lead a pressurized coolant in the trailing edge coolant cavity 41f toward the coolant exit slots <NUM> at the trailing edge <NUM> via a serial impingement scheme. In particular, the pressurized coolant flowing generally forward-to-aft impinges serially on to the rows of features <NUM>, leading to a transfer of heat to the coolant accompanied by a drop in pressure of the coolant. Heat may be transferred from the outer wall <NUM> to the coolant by way of convection and/or impingement cooling, usually a combination of both.

According to the invention, each feature <NUM> is elongated along the radial direction. That is to say, each feature <NUM> has a length in the radial direction which is greater than a width in the chord-wise direction. A higher aspect ratio provides a longer flow path for the coolant in the passages <NUM>, leading to increased cooling surface area and thereby higher convective heat transfer. In relation to the double or triple impingement plates, the described arrangement provides a longer flow path for the coolant and has been shown to increase both heat transfer and pressure drop to restrict the coolant flow rate. Such an arrangement may thus be suitable in advanced turbine blade applications which require smaller amounts of cooling air.

The exemplary turbine airfoil <NUM> is manufactured by a casting process involving a casting core, typically made of a ceramic material. The core material represents the hollow coolant flow passages inside turbine airfoil <NUM>. It is beneficial for the casting core to have sufficient structural strength to survive through the handling during the casting process. To this end, the coolant exit slots <NUM> at the trailing edge <NUM> may be designed to have larger dimensions at the span-wise ends of the airfoil, i.e., adjacent to the root and the tip of the airfoil <NUM>, to form a stronger picture frame like configuration. However, such a configuration may result in higher coolant flow near the airfoil root and tip than desired. Embodiments of the present invention provide an improvement to achieve not only a strong casting core but also a limitation in the coolant flow.

<FIG>, <FIG> and <FIG> illustrate portion of an exemplary casting core for manufacturing the inventive turbine airfoil <NUM>. The illustrated core element 141f represents the trailing edge coolant cavity 41f of the turbine airfoil <NUM>. The core element 141f has a core pressure side <NUM> and a core suction side <NUM> extending in a span-wise direction, and further extending chord-wise toward a core trailing edge <NUM>. <FIG> illustrate a views looking from the core suction side <NUM>, with <FIG> illustrating a first span-wise end portion which is adjacent to the radially outer airfoil end face <NUM> (airfoil tip cap), and <FIG> illustrating a second span-wise end portion which is adjacent to the radially inner airfoil end face <NUM> coupled to the platform <NUM>. <FIG> illustrate views looking from the core pressure side <NUM>, with <FIG> illustrating a first span-wise end portion which is adjacent to the radially outer airfoil end face <NUM> (airfoil tip cap), and <FIG> illustrating a second span-wise end portion which is adjacent to the radially inner airfoil end face <NUM> coupled to the platform <NUM>. As shown, the core element 141f comprises an array of perforations <NUM> there-through, located between span-wise ends of the core element 141f. Each perforation <NUM> extends all the way from the core pressure side <NUM> to the core suction side <NUM>. The perforations <NUM> form the cooling features the <NUM> in the trailing edge coolant cavity 41f (see <FIG>). Each perforation <NUM> is correspondingly elongated in the radial or span-wise direction. The array comprises multiple radial rows of said perforations <NUM> with the perforations <NUM> in each row being interspaced radially by interstitial core elements <NUM> that form the coolant passages <NUM> in the turbine airfoil <NUM>. The core elements <NUM> form the trailing edge coolant exit slots <NUM> of the turbine airfoil <NUM>.

As shown in <FIG> and <FIG>, the array of perforations <NUM> is located between the span-wise ends of the core element 141f, but does not extend all the way up to the span-wise ends thereof. As per embodiments of the present invention, at the span-wise ends of the core element 141f, indentations are provided on the core pressure side <NUM> and/or the core suction side <NUM>. In the non-limiting example as illustrated herein, at the radially outer span-wise end, indentations are provided at a chord-wise upstream location of the core element 141f, which is generally thicker. At the relatively narrow chord-wise downstream location, perforations may formed through the core element 141f along the radially outer span-wise end thereof. At the radially inner span-wise end, perforations are eliminated altogether. According to the invention, chord-wise spaced indentations 172A and 182A are provided on the first and second span-wise ends of the core pressure side <NUM> respectively (<FIG>) and chord-wise spaced indentations 172B and 182B are provided on the first and second span-wise ends of the core suction side <NUM> respectively (<FIG>).

As shown in <FIG>, the indentations 172A-B and 182A-B (shown in <FIG> and <FIG>) form framing features 72A-B, 82A-B in a respective framing passage <NUM>, <NUM> in the trailing edge coolant cavity 41f of the turbine airfoil <NUM>. The framing passages <NUM> and <NUM> are located at first and second span-wise ends respectively of the trailing edge coolant cavity 41f. In particular, the respective framing passage <NUM>, <NUM> is located between the cooling features <NUM> and a respective airfoil radial end face <NUM>, <NUM>. The framing features 72A-B, 82A-B are configured as ribs. As can be seen, the ribs 72A, 82A protrude from the pressure sidewall <NUM> of the airfoil <NUM>, and the ribs 72B, 82B protrude from the suction sidewall <NUM> of the airfoil <NUM>. Each of the ribs 72A-B, 82A-B extends only partially between the pressure sidewall <NUM> and the suction sidewall <NUM>.

The indentations 172A-B, 182A-B maintain strength of the ceramic core at the root and the tip, as opposed to complete perforations through the core pressure and suction sides. In the illustrated embodiment, as shown in the radial top view in <FIG>, the indentations 172A on the core pressure side <NUM> and the indentations 172B on the core suction side <NUM> are alternately positioned along the chord-wise direction. Like-wise, as shown in the radial bottom view in <FIG>, the indentations 182A on the core pressure side <NUM> and the indentations 182B on the core suction side <NUM> are alternately positioned along the chord-wise direction.

Claim 1:
A turbine blade (<NUM>) having an airfoil comprising:
an outer wall (<NUM>) delimiting an airfoil interior (<NUM>), the outer wall (<NUM>) extending span-wise along a radial direction of a turbine engine and being formed of a pressure sidewall (<NUM>) and a suction sidewall (<NUM>) joined at a leading edge (<NUM>) and at a trailing edge (<NUM>),
a trailing edge coolant cavity (41f) located in the airfoil interior (<NUM>) between the pressure sidewall (<NUM>) and the suction sidewall (<NUM>), the trailing edge coolant cavity (41f) being positioned adjacent to the trailing edge (<NUM>) and in fluid communication with a plurality of coolant exit slots (<NUM>) positioned along the trailing edge (<NUM>),
wherein a plurality of cooling features (<NUM>) are located in the trailing edge coolant cavity (41f) and are disposed in a flow path of the coolant flowing toward the coolant exit slots (<NUM>),
the cooling features (<NUM>) being located between the radially outer span-wise end of the trailing edge coolant cavity (41f) and the radially inner span-wise end of the trailing edge coolant cavity,
wherein each cooling feature (<NUM>) has a length in the radial direction which is greater than a width in the chord-wise direction,
wherein the cooling features comprise an array of pins (<NUM>), each pin (<NUM>) extending from the pressure sidewall (<NUM>) to the suction sidewall (<NUM>), the array comprising multiple chord-wise spaced apart radial rows of said pins (<NUM>) with the pins (<NUM>) in each row being interspaced radially to define coolant passages (<NUM>) therebetween,
wherein at least one framing passage (<NUM>, <NUM>) is formed at a span-wise end of the trailing edge coolant cavity (41f), wherein the at least one framing passage (<NUM>, <NUM>) comprises a first framing passage (<NUM>) and a second framing passage (<NUM>) formed at span-wise opposite ends of the trailing edge coolant cavity (41f), and framing features (72A-B, 82A-B) located in both the first and second framing passages (<NUM>, <NUM>), the framing features configured as ribs (72A-B, 82A-B) arranged chord-wise spaced apparat on the pressure sidewall (<NUM>) and/or the suction sidewall (<NUM>) and protruding from the pressure sidewall (<NUM>) and/or the suction sidewall (<NUM>), the ribs (72A-B, 82A-B) extending partially between the pressure sidewall (<NUM>) and the suction sidewall (<NUM>), wherein each rib (72A-B, 82A-B) is aligned with a respective row of said pins (<NUM>) in the radial direction.