Patent Description:
Conventional turbine airfoils used in blades and vanes of gas turbine engines have a trailing edge that is thin for aerodynamic efficiency. However, a lack of cooling surface area on the interior makes it difficult to cool the thin trailing edge. The trailing edge is typically cast integrally with the entire blade by using a ceramic core. The features and size of the ceramic core are reflected in the trailing edge. However, core design considerations must be weighed against trailing edge design considerations. For example, larger core features that create impingement channels in the trailing edge are better for core strength, but larger impingement channels mean reduced flow metering. Hence, a well designed core that balances core considerations with trailing edge cooling requirements is a key aspect of a well designed trailing edge design.

Impingement cooling along the mean camber line in a turbine airfoil trailing edge is known. In this arrangement orifices are cast as part of the trailing edge and are oriented with the mean camber line create high-speed impingement jets of cooling fluid. These impingement jets may impinge a surface between adjacent downstream impingement orifices, and this results in an increased heat transfer rate. Single, double, or triple impingement may occur before the spent cooling fluid is exhausted from the trailing edge into the combustion gas path. The series of impingement orifices also act to meter the flow and this provides a more efficient use of the cooling fluid.

By virtue of their location on the mean camber line the impingement orifices are located between the concave interior surface on the suction side and the convex interior surface on the pressure side of the airfoil. Prior cooling schemes have improved heat transfer by angling the impingement orifices such that they produce impingement jets that impinge the concave and convex interior surfaces. This, in turn, cools the respective exterior surfaces of the trailing edge. Other prior cooling schemes place various surface features on the interior surfaces coincident with the impingement jets. However, operating temperatures of gas turbine engines continue to increase. This leaves room in the art for improvements to cooling of the trailing edge.

<CIT> discloses a prior art gas turbine engine component according to the preamble of claim <NUM>.

<CIT>, which is prior art according to Art <NUM>(<NUM>) EPC only, discloses a prior art gas turbine engine component having trip strips.

<CIT> discloses a prior art crossover cooled airfoil trailing edge.

<CIT> discloses a prior art turbine blade outer air seal with optimized cooling. The invention is disclosed in claim <NUM>.

The invention is explained in the following description in view of the drawings that show:.

The present inventors have developed a cooling arrangement for a turbine airfoil's trailing edge, where the trailing edge and all elements of the cooling arrangement are integrally cast with the airfoil using a ceramic casting core. The invention capitalizes on advances in casting core technology to form an arrangement where the elements harmonize to form an unexpectedly extremely efficient cooling arrangement. Specifically, the airfoil and trailing edge are cast around a ceramic casting core configured to form impingement orifices and chevrons within the trailing edge. Some of the impingement orifices direct impingement jets toward chevrons disposed on an interior surface on the pressure side of the airfoil. Other impingement orifices direct impingement jets toward chevrons disposed on an interior surface on the suction side. There may be one or several rows of impingement orifices. Spent impingement air exhausts from the trailing edge into the flow of combustion gases. Compared to impingement jets all pointing in the same direction, alternating the target of the impingement jest from suction side to pressure side not only helps to increase the surface area being cooled, but it also serves to strengthen the trailing edge section of the ceramic core, thereby increasing production yield while allowing the diameter of the impingement jets to be smaller. Moreover, the use of chevrons not only increases surface area, but it also serves to spread the cooling air in order to more evenly cool the surface and to increase the area being cooled effectively when compared to traditional turbulators or parallel grooves.

<FIG> shows an exemplary embodiment of an airfoil <NUM> having a pressure side <NUM>, a suction side <NUM>, a leading edge <NUM>, a trailing edge <NUM>, a mean camber line <NUM>, a first row <NUM> of impingement orifices <NUM>, a second row <NUM> of impingement orifices <NUM>, and a third row <NUM> of impingement orifices <NUM>, where each row <NUM>, <NUM>, <NUM> is oriented radially in the gas turbine engine from a base of the airfoil <NUM> to a tip of the airfoil <NUM>. The rows <NUM>, <NUM>, <NUM> are disposed in a trailing edge portion <NUM> of the airfoil <NUM>. Cooling fluid enters the first row <NUM> of impingement orifices and exits the trailing edge portion <NUM> via exhaust orifices <NUM>. The pressure side <NUM> is cooled via cooling of a pressure side interior surface <NUM>, and the suction side <NUM> is cooled via a cooling of a suction side interior surface <NUM>.

<FIG> shows a close-up of the trailing edge portion <NUM> of <FIG>. Within this cross section fresh cooling fluid <NUM> from an upstream cavity <NUM> within the airfoil <NUM> enters an impingement orifice inlet <NUM> of an impingement orifice <NUM> of the first row <NUM>, and travels through and exhausts from the impingement orifice <NUM> via an impingement orifice outlet <NUM> in the form of an impingement jet <NUM>. A center of the impingement orifice outlet <NUM> of the impingement orifice <NUM> of the first row <NUM> is disposed on the suction side <NUM> of the mean camber line <NUM>, although it need not necessarily be so long as the respective impingement jet <NUM> is directed at an angle <NUM> to the mean camber line <NUM>. In this cross section the impingement jet <NUM> of the first row <NUM> is directed toward a target area <NUM> on a concave first-row interior surface <NUM> on the suction side <NUM> at an angle of impingement <NUM>. A surface feature (not shown) or plural surfaces features are positioned such that at least a portion of the surface feature is within the target area <NUM>.

Spent cooling fluid <NUM> from the impingement jet <NUM> of the first row <NUM> becomes fresh cooling fluid <NUM> for the second row <NUM>. The fresh cooling fluid <NUM> enters the impingement orifice inlet <NUM> of an impingement orifice <NUM> of the second row <NUM>, and travels through and exhausts from the impingement orifice <NUM> via an impingement orifice outlet <NUM> in the form of an impingement jet <NUM>. The impingement orifice inlet <NUM> of the impingement orifice <NUM> of the second row <NUM> may be at a different elevation than the impingement orifice outlet <NUM> of the impingement orifice <NUM> of the first row <NUM>, and hence the impingement orifice <NUM> of the second row <NUM> is represented using dotted lines. A center of the impingement orifice outlet <NUM> of the impingement orifice <NUM> of the second row <NUM> is disposed on the pressure side <NUM> of the mean camber line <NUM>, although it need not necessarily be so long as the respective impingement jet <NUM> is directed at an angle to the mean camber line <NUM>. In this cross section the impingement jet <NUM> of the second row <NUM> is directed toward a target area <NUM> on a convex second-row interior surface <NUM> on the pressure side <NUM>. A surface feature (not shown) or plural surfaces features are positioned such that at least a portion of the surface feature is within the target area <NUM>.

Spent cooling fluid from the impingement jet <NUM> of the second row <NUM> becomes fresh cooling fluid <NUM> for the third row <NUM>. The fresh cooling fluid <NUM> enters the impingement orifice inlet <NUM> of an impingement orifice <NUM> of the third row <NUM>, and travels through and exhausts from the impingement orifice <NUM> via an impingement orifice outlet <NUM> in the form of an impingement jet <NUM>. The impingement orifice inlet <NUM> of the impingement orifice <NUM> of the third row <NUM> may be at a same elevation as the impingement orifice outlet <NUM> of the impingement orifice <NUM> of the first row <NUM>, and hence the impingement orifice <NUM> of the third row <NUM> is represented using solid lines. A center of the impingement orifice outlet <NUM> of the impingement orifice <NUM> of the third row <NUM> is disposed on the suction side <NUM> of the mean camber line <NUM>, although it need not necessarily be so long as the respective impingement jet <NUM> is directed at an angle to the mean camber line <NUM>. In this cross section the impingement jet <NUM> of the third row <NUM> is directed toward a target area <NUM> on a concave third-row interior surface <NUM> on the suction side <NUM>. A surface feature (not shown) or plural surfaces features are positioned such that at least a portion of the surface feature is within the target area <NUM>. Spent cooling fluid <NUM> from the impingement jet <NUM> of the third row <NUM> exhausts from the trailing edge portion <NUM> via the exhaust orifices <NUM>.

In this exemplary embodiment, the rows <NUM>, <NUM>, <NUM> within this cross section alternate from suction side <NUM> to pressure side <NUM> to suction side. It is possible that they may point to different sides but not necessarily in an alternating pattern as shown. For example, in an alternate exemplary embodiment the first row <NUM> and the second row may point to the pressure side <NUM> while the third row may point to the suction side <NUM>. Likewise, the arrangement seen may vary as the location of the cross section is varied from base to tip of the airfoil <NUM>.

<FIG> is a side view of a trailing edge portion <NUM> of a casting core <NUM> used to form an alternate exemplary embodiment of a trailing edge portion <NUM> configured for double impingement. The casting core may be made of a ceramic material. <FIG> is a close up of a region within the trailing edge portion <NUM> of <FIG> having impingement- orifice-forming structures <NUM> that form the impingement orifices <NUM> within the airfoil <NUM> when the casting core <NUM> is removed.

<FIG> schematically depicts in a perspective view a portion of the casting core ending at line <NUM>-<NUM> of <FIG>, with a line representing a path that a flow of cooling fluid may take at the same location within the airfoil <NUM>. In other words, an outer surface <NUM> of the casting core <NUM> is being modeled as an inner surface of the airfoil <NUM> it forms. <FIG> is similar to <FIG>, but taken along line <NUM>-<NUM> of an impingement-orifice-forming structure <NUM> that is immediately adjacent the impingement-orifice-forming structure <NUM> of line <NUM>-<NUM>. When comparing <FIG>, it is apparent that in this exemplary embodiment, adjacent impingement-orifice-forming structures <NUM> within a first row <NUM> of impingement-orifice-forming structures <NUM> alternate which side they point to. They may alternate every-other as shown, or they may alternate in other groups, such as two pointing to one side, and then two pointing to another side etc..

Also visible in <FIG> is a second row <NUM> of impingement-orifice-forming structures <NUM>. In this view it can be seen that the first row <NUM> and the second row are offset vertically (i.e. from the base of the airfoil to the tip of the airfoil). This makes the path taken by the cooling fluid more tortuous and hence more efficient.

<FIG> shows a casting core coupon <NUM> that demonstrates the cooling arrangement to be used in the trailing edge portion <NUM>. This exemplary embodiment employs a triple impingement cooling arrangement that includes the first row <NUM>, the second row <NUM>, and a third row <NUM> of impingement-orifice-forming structures <NUM>. Chevron arrangement forming structures <NUM> can be seen formed in an outer surface <NUM> of the casting core coupon <NUM> and are configured to form chevron arrangements in the inner surfaces of the trailing edge portion <NUM>. It can be seen that within each row <NUM>, <NUM>, <NUM> the impingement-orifice-forming structures <NUM> alternate their direction. It can also be seen that the chevron arrangement forming structures <NUM> are coordinated with the impingement-orifice-forming structures <NUM> so that impingement jets will direct cooling fluid onto respective chevron arrangements.

<FIG> shows a cross section of a portion of a casting <NUM> made using the casting core coupon <NUM> of <FIG>. A void <NUM> exists where the casting core coupon <NUM> was formerly present and this void <NUM> represents the upstream cavity <NUM> that supplies fresh cooling fluid <NUM> to the first row <NUM> of impingement orifices <NUM>. Also visible are the second row <NUM> and the third row <NUM> of impingement orifices <NUM>, and a first row <NUM>, a second row <NUM>, and a third row <NUM> of chevron arrangements <NUM>. Each chevron arrangement <NUM> may have one or more than one individual chevrons <NUM>. Using the first row <NUM> of impingement orifices <NUM> to explain a configuration of each row, it can be seen that there are first group impingement orifices <NUM> oriented into the page and second group impingement orifices <NUM> oriented out of the page. The first group impingement orifices <NUM> can be likened to impingement orifices that direct an impingement jet <NUM> toward a target area <NUM> on the concave first-row interior surface <NUM> on the suction side <NUM>. Similarly, the second group impingement orifices <NUM> can be likened to impingement orifices that direct an impingement jet <NUM> toward a target area <NUM> on a convex first-row interior surface on the pressure side <NUM> (not visible in this view).

The impingement orifices <NUM> may be circular in cross section, but because they are angled toward the first-row interior surface <NUM> impingement orifices <NUM> with a circular cross section form an oval-shaped target area <NUM>. The target area <NUM> may range in size, and may include smaller <NUM>, mid-range <NUM>, and larger <NUM> target areas, where the size is relative to how much of the chevron arrangement <NUM> lies within the target area <NUM>. A shape of a perimeter <NUM> of the target area <NUM> may be varied by varying a shape of the cross section of the impingement orifice <NUM> itself. For example, if the cross section of the impingement orifice were oval with a longer axis oriented in and out of the page, the ovality of the perimeter <NUM> would be increased from that produced by the impingement orifice having the circular cross section. Conversely, if the ovality of the cross section were oriented such that the longer axis was more parallel to the first-row interior surface <NUM>, then the shape of the perimeter <NUM> would be more circular. Likewise, by changing the angle of impingement <NUM> the ovality of the perimeter <NUM> can be changed. The shape of the cross section of the impingement orifice <NUM> and the angle of impingement <NUM> can be manipulated as necessary to achieve whatever shape is desired for the perimeter <NUM> of the target area <NUM>. In addition, the shape of the perimeter <NUM> may be the same for all target areas <NUM>, or some or all of the target areas <NUM> may have their own, unique perimeter shape. These perimeter shapes may be selected to accommodate local cooling requirements and local geometries etc..

Each chevron <NUM> includes a tip <NUM> and two wings <NUM>. Adjacent chevrons <NUM> form a groove <NUM> there between that may be used to guide the cooling fluid. The tip <NUM> may be a closed tip <NUM> or an open tip <NUM>. The wings <NUM> may be continuous <NUM> or discontinuous <NUM>. The configuration of chevron arrangements <NUM> may vary from one chevron arrangement <NUM> to the next and may be selected to accommodate local cooling requirements and local geometries etc. The chevrons <NUM> may span an entirety of its target area <NUM>, or the target area <NUM> may be larger than the span of the chevron <NUM>. The tip <NUM> of one or all chevrons <NUM> in chevron arrangement <NUM> may be disposed within the target area <NUM>.

Spent cooling fluid <NUM> may flow in the grooves <NUM> formed by the wings <NUM> of the chevrons <NUM>. These grooves <NUM> may be oriented so that they guide the spent cooling fluid <NUM> along a same path the spent cooling fluid <NUM> would have taken if the chevron arrangement <NUM> were not present. In other words, streamlines <NUM> present in the spent cooling fluid <NUM> would naturally follow a course if the chevron arrangement <NUM> were not present. The chevron arrangement <NUM> can be configured so that the wings <NUM> and/or the grooves <NUM> follow the same streamlines as shown in chevron arrangement <NUM>. The result is that the spent cooling fluid <NUM> will lose little or no energy as a result of the presence of the chevron arrangement <NUM>, but will benefit from the increased surface area created by the chevron arrangement.

Alternately, the wings <NUM> and/or the grooves <NUM> can be disposed at an angle to the streamlines <NUM> that the spent cooling fluid <NUM> would naturally form, as shown in chevron arrangement <NUM>. This configuration forces the spent cooling fluid <NUM> to flow over the wings <NUM> and this creates turbulence, thereby increasing a cooling effect. A length of the wings <NUM> and a wing angle <NUM> of the wing <NUM> to the natural streamline <NUM> need to be designed to strike a balance between a desire to increase turbulence, and hence increase a cooling efficiency, and a desire to reduce a boundary layer that may form on a downstream side of the wing <NUM>, which forms at longer wing <NUM> lengths and greater wing angles <NUM>. The wing angle <NUM> will also determine how well the created turbulence follows the wings <NUM> and/or grooves, which also affects heat transfer. Similarly, if the wing <NUM> is discontinuous, a length between gaps <NUM> needs to be selected to maximize cooling effectiveness by balancing turbulence creation with boundary layer formation.

In an exemplary embodiment the wings <NUM> and/or the grooves <NUM> can be configured to guide spend cooling fluid <NUM> toward the impingement orifice inlet <NUM> of a subsequent impingement orifice <NUM>. For example, chevron arrangement <NUM> guides spent cooling fluid <NUM> from second row impingement orifice <NUM> toward impingement orifice inlets <NUM> of third row impingement orifices <NUM>, <NUM>. This may be done to improve flow efficiency through the trailing edge portion <NUM>. Alternately, the wings <NUM> and/or the grooves <NUM> can be configured to guide spent cooling fluid <NUM> toward interstitial structure <NUM> between impingement orifice <NUM> should greater fluidic chaos be desired at that location.

In an exemplary embodiment the cooling arrangement may be configured such that a stagnation point <NUM> within a target area <NUM> is arranged upstream of the tip <NUM> of one or all chevrons <NUM> in the chevron arrangement <NUM>. Doing so ensures spent cooling fluid <NUM> flows along the wings <NUM> and/or the grooves <NUM> away from the tips <NUM> as opposed to flowing upstream toward the tips <NUM>, which ensures a more uniform flow.

Various cooling arrangements with other surface features and flow paths were considered but this combination of multiple rows of angled impingement on respective chevron ribs provided the greatest heat transfer rate, resulting at least from the increased surface area and increased turbulence, while allowing formation of the trailing edge portion <NUM> integral to the airfoil <NUM> using a casting core <NUM> (which may be made of a ceramic material). The improved heat transfer test results using the arrangement disclosed herein can be seen in <FIG>. When compared to results from other configurations, all of which are below the results from the configuration disclosed herein, it become apparent that this increased cooling represents an improvement in the art.

Claim 1:
A gas turbine engine component, comprising:
an internal surface (<NUM>) comprising a row (<NUM>, <NUM>, <NUM>) of chevron arrangements (<NUM>);
an opposing internal surface (<NUM>) comprising a row of chevron arrangements; and
first and second rows (<NUM>, <NUM>) of impingement orifices, the first row (<NUM>) of impingement orifices including an impingement orifice (<NUM>) configured to direct an impingement jet of cooling fluid onto the internal surface (<NUM>) at an angle other than perpendicular and the second row of impingement orifices (<NUM>) including an impingement orifice (<NUM>) configured to direct an impingement jet of cooling fluid onto the opposing internal surface (<NUM>) at an angle other than perpendicular, wherein a target area (<NUM>) of each of the impingement jets encompasses a tip (<NUM>) of a chevron (<NUM>) within a corresponding one of the chevron arrangements (<NUM>), and wherein wings (<NUM>) of the chevron (<NUM>) diverge in conjunction with a divergence of spent impingement air,
characterised in that:
each of the chevron arrangements (<NUM>) comprises a plurality of chevrons (<NUM>), wherein wings (<NUM>) of adjacent chevrons (<NUM>) of each of the chevron arrangements (<NUM>) form a groove (<NUM>), and the component is configured such that spent impingement air flows into the groove (<NUM>).