Patent Description:
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.

The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction.

<CIT> discloses an airtfoil with the features of the preamble of claim <NUM>.

<CIT> describes a turbine component comprising a root, a tip, and an airfoil portion having a leading and a trailing edge, an external suction side and pressure side wall between the leading and trailing edge. The walls enclose a central cavity for the passage of cooling air, said cavity being partitioned into a leading edge- and a trailing edge region by at least one longitudinally extending first web connecting the suction side wall with the pressure side wall and a second longitudinally extending web, connecting said first web with the suction side wall, thereby defining a first and second entry chamber. The first web is provided with at least one cross-over hole between the first entry chamber and the leading edge chamber, whereas the second web has no openings.

<CIT> describes airfoils which include a leading edge, a trailing edge, a pressure side exterior wall, and a suction side exterior wall. A plurality of cooling passages are formed within the airfoil. A plurality of first interior ribs extend from the pressure side exterior wall to the suction side exterior wall, and a plurality of second interior ribs extend from the suction side exterior wall toward the pressure side exterior wall and intersect with a first interior rib. At least one pressure side main body cooling passage is defined between the pressure side exterior wall and two first interior ribs of the plurality of first interior ribs and at least one suction side main body cooling passage is defined between the suction side exterior wall, a first interior rib, and a second interior rib.

<CIT> describes an airfoil which includes an airfoil outer wall having widthwise spaced apart pressure and suction sidewall sections extending chordally between leading and trailing edges of the airfoil and extending longitudinally from a base to a tip. Inside the airfoil is at least one internal cooling circuit having a plurality of longitudinally extending circuit channels between longitudinally extending internal ribs extending widthwise between the pressure and suction sidewall sections and a longitudinally extending first sidewall film cooling chamber positioned between one of the sidewall sections and a first inner wall bounding the cooling circuit. Sidewall film cooling holes extend through the pressure sidewall section from the first sidewall film cooling chamber. The internal ribs have corresponding rib angles with respect to a centerline, the first inner wall has a corresponding first wall angle with respect to the centerline, and each of the rib angles and the first wall angle are constant in a longitudinal direction from the base to the tip. In the preferred embodiment, all the inner walls are substantially parallel to each other and all the transverse ribs are substantially parallel to each other or are substantially parallel to the first and second inner walls.

<CIT> describes an internally cooled turbine blade which includes dynamic cool air flowing radial passageways with an inlet at the root and a discharge at the tip, the passageways feeding a plurality of radially spaced film cooling holes in the airfoil surface. Replenishment holes radially spaced in the inner wall of the radial passageways communicate with internal cooling passages in the blade to replenish the cooling air lost to the film cooling holes. The discharge orifice is sized to match the backflow margin to achieve a constant film hole coverage throughout the radial length. Trip strips may be employed to augment the pressure drop distribution.

<CIT> describes a turbine engine component, such as a blade or a vane. The turbine engine component has a pressure side and a suction side. Each of the pressure and suction sides has an external wall and an internal wall. A first set of fluid passageways is located on the pressure side between the external wall and the internal wall. A second set of fluid passageways is located on the suction side between the external wall and the internal wall. Each of the fluid passageways in the first set and in the second set has a wavy configuration. The turbine engine component may also have one or more wavy trailing edge cooling passageways for cooling a trailing edge portion of the component.

<CIT> describes an airfoil that includes a leading edge and a trailing edge. A first exterior wall extends between the leading edge and the trailing edge. A second exterior wall opposite the first exterior wall extends between the leading edge and the trailing edge. A first cavity is adjacent the first exterior wall and includes a central portion and at least one forward extending slot passage and at least one aft extending slot passage.

<CIT> describes an airfoil for a gas turbine engine that includes an airfoil section that has an internal wall and an external wall. The external wall defines pressure and suction sides that extend in a chordwise direction between a leading edge and a trailing edge, a first impingement cavity and a second impingement cavity bounded by the external wall at a leading edge region that defines the leading edge, a first feeding cavity separated from the first impingement cavity and from the second impingement cavity by the internal wall, and a first crossover passage within the internal wall that connects the first impingement cavity and the first feeding cavity. The first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity. A second crossover passage within the internal wall connects to the second impingement cavity. The second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

The present invention provides an airfoil according to the appended claims.

The various features and advantages of the present invention will become apparent to those skilled in the art from the following detailed description.

Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including but not limited to three-spool architectures.

In a further example, the engine <NUM> bypass ratio is greater than about <NUM>:<NUM>, with an example embodiment being greater than about <NUM>:<NUM>, the geared architecture <NUM> is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM> and the low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>:<NUM>. In one disclosed embodiment, the engine <NUM> bypass ratio is greater than about <NUM>:<NUM>, the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about <NUM>:<NUM>. The low pressure turbine <NUM> pressure ratio is pressure measured prior to the inlet of low pressure turbine <NUM> as related to the pressure at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including but not limited to direct drive turbofans.

The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about <NUM>:<NUM>.

<FIG> shows a representation of a sectioned airfoil <NUM> not falling within the scope of the claims, used in the turbine engine <NUM> (see also <FIG>). The airfoil <NUM> is a turbine blade; however, it is to be understood that this disclosure is also applicable to cooled blades or vanes.

The airfoil <NUM> includes an (outer) airfoil wall <NUM> that spans in a radial direction and delimits the aerodynamic profile of the airfoil <NUM>. In this regard, the wall <NUM> defines a leading end 62a, a trailing end 62b, and first and second side walls 62c/62d that join the leading end 62a and the trailing end 62b. In this example, the first side 62c is a pressure side and the second side 62d is a suction side. A mean line (M) extends from the leading end 62a to the trailing end 62b. All locations on the mean line (M) are equidistant from the first and second side walls 62c/62d (in a perpendicular direction to the mean line at the location). For purposes of this disclosure, elements, regions, or portions thereof that are below the mean line (M) in <FIG> are considered to be on the pressure side, and elements, regions, or portions thereof that are above the mean line (M) in <FIG> are considered to be on the suction side.

The airfoil <NUM> further includes one or more ribs <NUM>. In the illustrated example, the airfoil has three such ribs <NUM>, although the airfoil <NUM> in modified examples can include a single rib <NUM>, two ribs <NUM>, or more than three ribs <NUM>. And although a single rib <NUM> is described in some instances herein, it is to be understood that each such rib <NUM> has the described attributes of the single rib <NUM>.

Of the three ribs <NUM> depicted, one is adjacent the leading end 62a, followed by two intermediate ribs <NUM>, followed by a final rib <NUM> adjacent the trailing end 62b. For purposes of describing the attributes and functions of the ribs <NUM>, the first of the intermediate ribs <NUM> (adjacent the leading end rib <NUM>) will be considered a first rib 64a and the next intermediate rib <NUM> will be considered a second rib 64b. The terminology "first" and "second" is to differentiate that there are two distinct ribs. It is to be understood that the terms "first" and "second" are interchangeable and that the first rib could alternatively be termed as the second rib and that the second rib could alternatively be termed as the first rib, provided the ribs are adjacent one another. If the airfoil <NUM> includes additional ribs <NUM>, any two adjacent ribs <NUM> are considered first and second ribs.

Each rib <NUM> connects the first and second sides 62c/62d of the airfoil wall <NUM>. Each rib <NUM> is generally longitudinally elongated between an inner diameter and outer diameter such that it spans the full or substantially full longitudinal distance of the airfoil wall <NUM>. The term substantially full refers to at least <NUM>% of the longitudinal distance between the inner diameter and outer diameter.

Each rib <NUM> includes a first arm <NUM> by which the rib <NUM> is solely connected to the first or second side wall 62c, and second and third arms <NUM>/<NUM> by which the rib <NUM> is solely connected to the other of the first or second side wall 62d. Exclusive of any cooling apertures, each arm <NUM>/<NUM>/<NUM> is a solid, continuous wall. In general, except for a later example of a cross arm, the arms <NUM>/<NUM>/<NUM> do not have any structural appendages.

In the illustrated example, the first arm <NUM> is attached to the first side wall 62c and the second and third arms <NUM>/<NUM> are attached to the second side wall 62d. In later examples, the attachments of some or all of the ribs <NUM> are reversed and the first arm <NUM> is attached to the second side wall 62d and the second and third arms <NUM>/<NUM> are attached to the first side wall 62c. As used herein, the phrase "solely connected" or variations thereof refers to the arm or arms being the exclusive structural attachment(s) of the rib <NUM> to the first side wall 62c or second side wall 62d, without any intermediate structures in between the attachments.

The first arm <NUM> has first and second ends 66a/66b. The second and third arms <NUM>/<NUM> likewise have respective first ends 68a/70a and second ends 68b/70b. In this example, the first end 68a of the first arm <NUM> is attached to the first side wall 62c, the second end 68b of the first arm <NUM> is attached at a node <NUM> to the respective first ends 68a/70a of the second and third arms <NUM>/<NUM>, and the respective second ends 68b/70b of the second and third arms <NUM>/<NUM> are attached to the second side wall 62d.

Except for connection through the arms <NUM>/<NUM>/<NUM> to the airfoil wall <NUM>, the ribs <NUM> are disjoined from each other. As used herein, the term "disjoined" refers to the ribs <NUM> excluding any structural attachments to each other. Such an attachment configuration permits each rib <NUM> to reinforce the side walls 62c/62d and facilitate reduction in bulging from internal pressure, while still permitting the ribs <NUM> to move and thermally expand and contract at a different rate than the side walls 62c/62d during thermal cycling and without interference from adjacent ribs <NUM>.

The ribs <NUM> partition the interior cavity of the airfoil <NUM> such that the airfoil wall <NUM> and the first and second ribs 64a/64b bound a continuous cooling channel <NUM> there between. The cooling channel <NUM> is continuous in that it spans the adjacent first and second ribs 64a/64b and side walls 62c/62d without any partitions. Similarly, the second and third arms <NUM>/<NUM> and second side wall 62d bound a rib passage <NUM>.

In the illustrated example, relative to the mean line (M), the node <NUM> is located toward the first side wall 62c and is thus closer in distance to the first side wall 62c than to the second side wall 62d. In reverse orientations of the rib <NUM>, the node <NUM> is closer to the second side wall 62d than to the first side wall 62c. In additional examples, of the distance from the first side wall 62c to the mean line (M), the node <NUM> is located in the initial <NUM>% of the distance from the first side wall 62c (or alternatively the second side wall 62d), or in the initial <NUM>% of the distance from the first side wall 62c (or alternatively the second side wall 62d). The relatively close proximity of the node <NUM> to the first side wall 62c (or alternatively to the second side wall 62d) serves to create a relatively narrow leg of the cooling channel <NUM> along the first side wall 62c (or alternatively along the second side wall 62d). Thus, cooling air in the cooling channel <NUM> serves to cool a relatively long extent of the first side wall 62c (or alternatively the second side wall 62d), which will be further appreciated from the description below.

Cooling air, such as bleed air from the compressor section <NUM> of the engine <NUM>, is provided to the cooling channel <NUM> and the rib passage <NUM>. The cooling air can be fed from a radially inner or radially outer location into the cooling channel <NUM> and rib passage <NUM> for radial flow through the cooling channel <NUM> and rib passage <NUM>. For example, the cooling channel <NUM> and the rib passage <NUM> in the illustrated example are flow isolated from each other. As used herein, the phrase "flow isolated" or variations thereof refers to passages, channels, or both that are not fluidly connected to each other within the airfoil <NUM> such that air cannot flow within the airfoil <NUM> from one passage or channel to the other passage or channel. For instance, such flow isolation permits air in the cooling channel <NUM> and rib passage <NUM> to be used at differential pressures. In a modified example, the radial flow in the cooling channels <NUM> and/or rib passages <NUM> may turn in a platform or tip and flow in the opposite radial direction into another cooling channel <NUM> or rib passage <NUM>.

In <FIG>, the cooling channel <NUM> primarily serves to cool the first side wall 62c and the rib passage <NUM> primarily serves to cool the second side wall 62d. For instance, the cooling channel <NUM> as shown generally has a T-shape, with the top leg of the "T" extending along the first side wall 62c. The bottom of the stem of the "T" borders the second side wall 62d. The cooling channel <NUM> is thus exposed to a larger area of the first side wall 62c than the second side wall 62d to provide more cooling to the first side wall 62c than to the second side wall 62d. On the other hand, the rib passage <NUM> is bound by the second side wall 62d but has no exposure at the first side wall 62c and thus serves primarily to cool the second side wall 62d.

With such a configuration, the cooling of the first side wall 62c is in essence segregated from the cooling of the second side wall 62d. This enables the cooling to be controlled and optimized for each side wall 62c/62d. For instance, different cooling air pressures are utilized in the cooling channels <NUM> versus the rib passages <NUM>. Such pressures can be controlled, for example, by metering orifices or the like at or near the inlets of the cooling air into the cooling channels <NUM> and rib passages <NUM>.

The orientation of the rib or ribs <NUM> provides additional cooling configurations. For instance, as shown in the airfoil <NUM> in <FIG>, not falling within the scope of the claims, the orientation of the ribs <NUM> are reversed from that of <FIG> such that the first arms <NUM> are attached to the second side wall 62d and the second and third arms <NUM>/<NUM> are attached to the first side wall 62c. For viewability, not all of the elements described for <FIG> are numbered in <FIG>. However, it is to be understood that but for the reverse orientation, the ribs <NUM> in <FIG> or in any of the figures herein have the same attributes as the ribs <NUM> in <FIG>.

In the airfoil <NUM> of <FIG>, according to the invention, the ribs <NUM> have alternating orientations. Such a configuration alters the geometry of the cooling channel such that the cooling channel <NUM> in airfoil <NUM> has a Z-shape. In this case, since the top and bottom legs of the "Z" extend along the first side wall 62c and the second side wall 62d, the cooling channel <NUM> provides cooling to both side walls 62c/62d. The rib passages <NUM> of two of the ribs <NUM> cool the second side 62d, and the rib passage <NUM> of the other rib <NUM> cools the first side wall 62c. The rib passages <NUM> thus also serve in this configuration to cool both side walls 62c/62d.

<FIG> illustrate further example airfoils <NUM>/<NUM>/<NUM> that are the same, respectively, as the airfoils <NUM>/<NUM>/<NUM> of <FIG> except that the airfoils <NUM>/<NUM>/<NUM> include cooling apertures <NUM>. Flow arrows of the cooling air are shown and a depiction of a flow arrow that extends through a wall indicates that there is a cooling hole or aperture <NUM> at that location (not all of which are numbered).

The cooling apertures <NUM> provide additional cooling schemes to further enhance cooling. For instance, some of the cooling apertures <NUM> may serve as impingement holes to concentrate flow onto the inside surface of the adjacent portion of the wall <NUM>. Other of the cooling apertures that are on the wall <NUM> serve as film cooling holes for the exterior surfaces of the airfoil wall <NUM>. Other of the cooling apertures <NUM> that are on the long portions of the second and third arms <NUM>/<NUM> serve as feed holes to feed cooling air from the rib passage <NUM> into the cooling channel <NUM>. Therefore, various configurations of the cooling apertures <NUM> can be used to control cooling air flow in the respective airfoils. In these examples, although cooling air may flow radially, the cooling apertures <NUM> provide for impingement cooling and axial flow of the cooling air.

<FIG> illustrates another example orientation of the ribs <NUM> and a cooling configuration. In this example, two consecutive ribs <NUM> are oriented the same and the third rib <NUM> has the reverse orientation. This provides a T-shaped cooling channel <NUM> and a Z-shaped cooling channel <NUM>. Such a configuration may be used to provide more cooling to an area of one of the side walls 62c or 62d using the top of the "T" of the cooling channel <NUM>, while providing cooling to both side walls 62c/62d at other areas using the Z-shape of the cooling channel <NUM>. As in prior examples, the airfoil <NUM> may also include cooling apertures <NUM>, although in a further example the cooling apertures <NUM> are excluded for a radial flow scheme.

<FIG> illustrates a variation of the airfoil <NUM>. In the variation, the airfoil <NUM> includes a cross arm <NUM> connecting the second and third arms <NUM>/<NUM>. The cross arm <NUM> partitions the rib passage <NUM> to create an additional cooling passage <NUM> along either the first side wall 62c or the second side wall 62d. The cooling passage <NUM> is aligned with the rib <NUM> in that the cooling passage <NUM> is between the second and third arms <NUM>/<NUM>. The rib passages <NUM> thus do not border a hot wall. As a result, there is a reduced need or no need for cooling air in the rib passages <NUM>. However, the segregation of cooling air is maintained in that the cooling passages <NUM> can be fed separately from the cooling channels <NUM>/<NUM>. In such a cooling scheme, the rib passages <NUM> in essence become space-eaters that reduce cooling flow area inside the airfoil <NUM>/<NUM>.

The cooling passages <NUM> may be minicores or minicore passages. A "minicore" or "minicore passage" is a reference to the small investment casting core that is typically used to make an embedded cooling passage, as opposed to a main core that is used to form a main or central core cavity in an airfoil. Such minicore passages are typically high aspect ratio passages that are much thinner than passages in the central portion of the airfoil. The cooling passages <NUM> are fed cooling air radially, separately from the rib passages <NUM> and cooling channels <NUM>/<NUM> but alternatively or additionally are fed from the cooling channels <NUM>/<NUM> and/or rib passages <NUM>.

<FIG> illustrate, respectively, further example airfoils <NUM> and <NUM>. The airfoils <NUM>/<NUM> are the same, respectively, as airfoils <NUM>/<NUM> except that the side wall 162c (<FIG>) and the side wall 162d (<FIG>) include a cooling passage <NUM> embedded therein. For example, each side wall 162c/162d includes an inner portion 82a and an outer portion 82b between which the cooling passage <NUM> is embedded. The cooling passages <NUM> may also be high aspect ratio minicore or minicore passages, as discussed above. In these examples, the cooling passages <NUM> are offset from the ribs <NUM> in that portions of the cooling passages <NUM> are not between the second and third arms <NUM>/<NUM>. The cooling passages <NUM> are fed cooling air radially, separately from the rib passages <NUM> and cooling channels <NUM>/<NUM> but alternatively or additionally are fed from the cooling channels <NUM>/<NUM> and/or rib passages <NUM>.

The cooling passages <NUM> cool the respective side walls 162c/162d such that the rib passages <NUM> do not border a hot wall. As a result, there is a reduced need or no need for cooling air in the rib passages <NUM>. However, the segregation of cooling air is maintained in that the cooling passages <NUM> can be fed separately from the cooling channels <NUM>/<NUM>. In such a cooling scheme, the rib passages <NUM> in essence become space-eaters that reduce cooling flow area inside the airfoil <NUM>/<NUM>.

As also demonstrated in the examples herein, the airfoils may also include a cooling passage <NUM> toward the trailing end 62b. Like the cooling passages <NUM>, the cooling passages <NUM> are high aspect ratio minicore passages. As shown in <FIG>, the cooling passages <NUM> are generally V-shaped, with one leg of the "V" on the pressure side and the other leg on the suction side. The cooling passages <NUM> may include cooling apertures as discussed above for discharging cooling air to the exterior of the airfoil, however, in the examples shown the cooling passages <NUM> at least have an internal aperture <NUM> that directly connects the cooling passage <NUM> to a trailing end passage <NUM>. Thus, rather than being dumped overboard to the pressure or suction side, at least a portion of the cooling air in the cooling passages <NUM> is fed into the trailing end passage <NUM> for discharge from the trailing end 62b. In other examples, as shown in <FIG>, <FIG>, <FIG>, <FIG>, <FIG>, and <FIG>, the cooling passage <NUM> is exclusively on the suction side, while in <FIG> and <FIG> the cooling passages <NUM> is exclusively on the pressure side. As will be appreciated, the cooling passages <NUM> are used as shown in combination with the ribs and rib passages described herein but may alternatively be used in other configurations that do not have such ribs and rib passages.

As will be appreciated from the examples herein, various cooling configurations are available via the ribs <NUM>. Furthermore, such configurations facilitate segregating cooling air flow used to cool the pressure side from cooling air flow used to cool the suction side. As a result, cooling air flows to the pressure and suction sides can be optimized, thereby facilitating a reduction in the use of cooling air and increase in efficiency.

The airfoils described herein may be fabricated from superalloys using such processes as investment casting or additive manufacturing. For example, in an investment casting process, one or more investment cores are fabricated and then used in the casting of the superalloy to define internal features in the airfoil. Depending on the geometry of the investment core, it may be formed from a ceramic or other suitable material in a molding process. The molding process involves injecting the material into the cavity of a molding die to form the shape of the core. In particular, for single pull direction molding in which two molding die halves are closed and opened along a single axial direction, there must be a suitable parting line along the molded component which allows the die halves to open without destroying the molded component and/or equipment. In other words, if the cavity were designed such that the molded component interlocked with the cavity in one or both die halves, the dies would not open, would destroy the molded component when opened, and/or would damage the equipment. In this regard, the geometry of the cores required to form the channels <NUM>/<NUM> described herein also facilitates core molding in that the geometries have suitable parting lines for injection molding. For instance, the channels <NUM>/<NUM> herein have the T-shape or Z-shape such that the respective corresponding T- and Z-shaped cores can be injection molded in a die having a single pull direction. <FIG> illustrate, respectively, a T-shape core <NUM> and a Z-shaped core <NUM>. The molding die halves are indicated at <NUM> and <NUM> and form parting lines <NUM> along the cores <NUM>/<NUM>. Although shown schematically, it is to be understood that the cores <NUM>/<NUM> have the geometric attributes of the channels <NUM>/<NUM> shown and described herein so as to form the ribs as shown and described. Moreover, the cores and channels herein also exemplify a method of fabricating the airfoil in an investment casting process by using the core or cores to form the disclosed airfoils.

Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this invention. In other words, a system designed according to an embodiment of this invention will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures.

Claim 1:
An airfoil (<NUM>) comprising:
an airfoil wall (<NUM>) defining a leading end (62a), a trailing end (62b), a first side wall (62c), and a second side wall (62d); and
first and second ribs (64a, 64b) each connect the first and second side walls (62c, 62d) of the airfoil wall (<NUM>), each of the ribs (64a, 64b) including:
a first arm (<NUM>) by which the respective first or second rib (64a, 64b) is solely joined to one of the first or second side walls (62c, 62d), and second and third arms (<NUM>, <NUM>) by which the respective first or second rib (64a, 64b) is solely connected to the other of the first or second side walls (62c, 62d); and
wherein for each rib the first arm (<NUM>), the second arm (<NUM>), and the third arm (<NUM>) each include a first end (66a, 68a 70a) and a second end (66b, 68b, 70b), the first end (66a) of the first arm (<NUM>) is attached to one of the first and second side walls (62c, 62d), the second end (66b) of the first arm (<NUM>) is attached at a node (<NUM>) to the respective first ends (68a, 70a) of the second and third arms (<NUM>, <NUM>), the second ends (68b, 70b) of the second and third arms (<NUM>, <NUM>) are attached to the other of the first and second side walls (62c, 62d), wherein
the first end (66a) of the first arm (<NUM>) of the first rib (64a) is attached to one of the side walls (62c, 62d), and the first end (66a) of the first arm (<NUM>) of the second rib (64b) is attached to the other of the side walls (62c, 62d), characterized in that:
there is a cooling channel <NUM> between adjacent first and second ribs (64a, 64b), and the cooling channel (<NUM>) has a Z-shape.