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 airfoil with ribs having connector arms. This document was published after the priority date of the present application, and therefore can be considered with respect to novelty only.

<CIT> discloses a stator vane for an industrial turbine including a serpentine flow cooling circuit for cooling the airfoil.

<CIT> discloses a turbine airfoil including a central cavity defined by an outer wall including pressure and suction sides extending between and joined at leading and trailing edges. Rib structures located in the central cavity define radial central channels extending across the chordal axis and flow passes associated with each central channel are connected in series to form a serpentine cooling path extending in the direction of the chordal axis.

According to a first aspect of the invention, there is provided an airfoil as claimed in claim <NUM>.

In a further embodiment of any of the foregoing embodiments, the radial cooling passage is flow isolated from the first and second cooling channels.

In a further embodiment of any of the foregoing embodiments, the radial cooling passage narrows to a neck portion, and the neck portion extends through the region between the first and second channel legs.

In a further embodiment of any of the foregoing embodiments, the turn channel includes a middle section between first and second ears. The middle section splits into the first and second channel legs.

In a further embodiment of any of the foregoing embodiments, the first and second channel legs each have a cross-sectional area defined by a product of a thickness and width (width × thickness) of the channel leg at the cross-section, and from the first channel end of the first cooling channel, the first and second channel legs increase in thickness and decrease in width.

In a further embodiment of any of the foregoing embodiments, the first and second channel legs increase in thickness by diffusion angles of no greater than <NUM> degrees.

In a further embodiment of any of the foregoing embodiments, the radially-extending ribs comprise first, second, and third ribs each connect the first and second sides of the airfoil wall. Each of the first, second, and third ribs define a tube portion that circumscribes a rib passage, wherein the rib passage is the radial cooling passage. First and second connector arms solely join the tube portion to, respectively, the first and second sides of the airfoil wall. The first rib, the second rib, and the airfoil wall bound a first cooling channel there between. The second rib, the third rib, and the airfoil wall bound a second cooling channel there between.

In a further embodiment of any of the foregoing embodiments, the turn channel includes a middle section between first and second ear sections, and the middle section splits into the first and second legs.

In a further embodiment of any of the foregoing embodiments, each of the first and second channel legs has a mouth at the first channel end and a crest at an apex of the turn channel. The mouth defines a cross-sectional area. The crest defines a cross-sectional area, and the cross-sectional area of the crest is from <NUM>% to <NUM>% of the cross-sectional area of the mouth.

In a further embodiment of any of the foregoing embodiments, the first and second channel legs each have a cross-sectional area defined by a product of a thickness and width (width × thickness) of the channel leg at the cross-section, and the first and second channel legs have an inner turn radius that is greater than a thickness dimension of the cross-sectional area of the mouth by a factor of at least <NUM>. The inner turn radius is taken from a line at an intersection of planes associated with the cross-sectional area of the mouth and the cross-sectional area of the crest.

In a further embodiment of any of the foregoing embodiments, the first and second legs have diffusion angles of no greater than <NUM> degrees.

In a further embodiment of any of the foregoing embodiments, the rib passage is flow isolated from the first and second cooling channels.

A gas turbine engine according to a second aspect of the present invention is provided in claim <NUM>.

In a further embodiment of any of the foregoing embodiments, the turn channel includes a middle section between first and second ear sections, and the middle section splits into the first and second channel legs.

In a further embodiment of any of the foregoing embodiments, the first and second channel legs each have a cross-sectional area defined by a product of a thickness and width (width × thickness) of the channel leg at the cross-section, and the first and second channel legs have an inner turn radius that is greater than a thickness dimension of the cross-sectional area of the mouth by a factor of at least <NUM>. The inner turn radius is taken from an intersection of planes of the cross-sectional area of the mouth and the cross-sectional area of the crest.

In a further embodiment of any of the foregoing embodiments, the first and second channel legs each have a cross-sectional area defined by a product of a thickness and width (width × thickness) of the channel leg at the cross-section, and the first and second channel legs increase in thickness from the first channel end with diffusion angles of no greater than <NUM> degrees.

"Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]^<NUM>.

<FIG> illustrates a sectioned airfoil <NUM> 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 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 sides 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. For a blade, the airfoil wall <NUM> will typically span in a radial direction from an inner platform to a free tip end. In a vane, the airfoil wall <NUM> will typically span in a radial direction from an inner platform to an outer platform.

The airfoil <NUM> further includes ribs <NUM> that each connect the first and second sides 62c/62d of the airfoil wall <NUM>. In the illustrated example, the airfoil has four such ribs <NUM>, although the airfoil <NUM> in modified examples can include fewer or additional 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>.

The ribs <NUM> are generally radially elongated between an inner diameter and outer diameter to span 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> defines a tube portion <NUM> that circumscribes a rib passage <NUM>, and first and second connector arms 70a/70b that solely join the tube portion <NUM> to, respectively, the first and second sides 62c/62d of the airfoil wall <NUM>. As used herein, the phrase "solely join" or variations thereof refers to the arm 70a being the exclusive structural attachment of the tube portion <NUM> to the first side 62c and the arm 70b being the exclusive structural attachment of the tube portion <NUM> to the second side 62d. Such an attachment configuration permits the rib <NUM> to reinforce the sides 62c/62d to facilitate reduction in bulging from internal pressure, while still permitting the rib <NUM> to move and thermally expand and contract at a different rate than the sides 62c/62d during thermal cycling.

The ribs <NUM> partition the interior cavity of the airfoil <NUM> such that the airfoil wall <NUM> and the rib <NUM> bound cooling channels <NUM> there between. In the illustrated example, a forward one of the cooling channels <NUM> may be considered a first cooling channel and the next aft cooling channel <NUM> may be considered a second cooling channel. The terminology "first" and "second" is to differentiate that there are two distinct cooling channels. It is to be understood that the terms "first" and "second" are interchangeable and that the first cooling channel could alternatively be termed as the second cooling channel and that the second cooling channel could alternatively be termed as the first cooling channel, provided the cooling channels are consecutive (a single tube portion <NUM> in between). If the airfoil <NUM> includes additional cooling channels <NUM>, any two consecutive cooling channels are considered first and second cooling channels.

Due to the geometry of the tube portions <NUM> and the connector arms 70a/70b, the cooling channels <NUM> in the illustrated example have an I-shape. In the I-shape, a bottom leg 72a of the "I" extends along the first side 62c, a top leg 70b of the "I" extends along the second side 62d, and the middle leg 72c of the "I" extends between a forward side of one tube portion <NUM> and an aft side of another tube portion <NUM>. The top and bottom legs 72a/72b of the "I" are bound by the pressure/suction sides of the tube portions and the first and second sides 62c/62d of the airfoil wall <NUM>.

Cooling air, such as bleed air from the compressor section <NUM> of the engine <NUM>, can be provided to the cooling channels <NUM> and the rib passage <NUM>. The cooling air can be fed from a radially inner or radially outer location into the cooling channels <NUM> and rib passage <NUM>. For example, the tube portions <NUM> are continuous such that the cooling channels <NUM> are flow isolated from the rib passages <NUM>. As used herein, the phrase "flow isolated" or variations thereof refers to the cooling channels <NUM> not being fluidly connected to the rib passages <NUM> such that cooling air cannot flow there between. For instance, such flow isolation permits air in the cooling channels <NUM> and the rib passages <NUM> to be used at differential pressures. In this regard, cooling air in the cooling channels <NUM> can be discharged through cooling holes or the like in the side walls 62c/62d to serve for cooling the side walls 62c/62d, while cooling air in the rib passage <NUM> can serve to cool a blade tip or platform or be provided to other downstream structures.

Referring to <FIG>, the radial extents of the cooling channels <NUM> are shown. Each cooling channel <NUM> extends between first and second cooling channel radial ends <NUM>/<NUM> (hereafter "ends"). For instance, in this example, the end <NUM> is a radially outer end and the end <NUM> is a radially inner end. The examples herein, however, are applicable to the radially inner end, the radially outer end, or both.

The cooling channels <NUM> are connected in a serpentine flow pattern, indicated at cooling path circuit <NUM>. In this regard, the airfoil <NUM> includes turn channels <NUM> at the ends <NUM>/<NUM>. In <FIG> the turn channels <NUM> are only schematically shown. The turn channels <NUM> serve to transfer flow from the one cooling channel <NUM> to the next consecutive cooling channel <NUM> (or alternatively any two adjacent cooling channels <NUM>).

Traditional turn channels in serpentine configurations are typically designed for simple channel geometries to turn flow from one channel to the immediately neighboring channel. Such turns, however, are inapplicable to more complex channel geometries and channel configurations and do not permit advanced cooling flow configurations. In this regard, as will be described below, the turn channel <NUM> according to the present disclosure facilitates turning between complex channels that are not directly next to each other.

<FIG> shows a representation of the region identified in <FIG>. As channels and features inside of a solid object are difficult to view, the turn channel <NUM> in <FIG> is shown in a negative view, where solid and open regions in the actual turn channel <NUM> are shown in the inverse relation in the figure, i.e., solid in the figure is open in the actual and open in the figure is solid in the actual. <FIG>, <FIG>, and <FIG> are negative representations.

Referring also to <FIG>, the turn channel <NUM> splits at the end <NUM> of the first cooling channel <NUM> into first and second channel legs <NUM>/<NUM> such that there is a region <NUM> between the first and second channel legs <NUM>/<NUM>. On the other side of the region <NUM>, the first and second channel legs <NUM>/<NUM> merge at the end <NUM> of the receiving cooling channel <NUM>. A radial cooling passage, here the rib passage <NUM>, extends through the region <NUM> between the first and second channel legs <NUM>/<NUM>. As best viewed in <FIG>, the rib passage <NUM> narrows to a neck portion 68a, and the neck portion 68a extends through the region <NUM>. The relatively narrow neck portion 68a permits lower angle splits. The split and region <NUM> thereby enable a pass-through configuration in which the rib passage <NUM> can continue to extend radially, yet still permit turning of the cooling air in the cooling channels <NUM>.

In the illustrated example, the turn channel <NUM> includes several sections to receive the cooling air flow from the different legs of the cooling channels <NUM>. The turn channel <NUM> includes first and second ears 86a/86b and a middle section 86c that connects the ears 86a/86b. The middle section 86c receives cooling air flow from the middle leg 72c of the cooling channel <NUM>, and the ears 86a/86b receive cooling air flow from, respectively, the legs 72a/72b of the channel <NUM>. In this example, it is the middle section 86c of the turn channel <NUM> that splits into the first and second channel legs <NUM>/<NUM> at wedge portion <NUM>. The wedge portion <NUM> defines a wedge half-angle WA between the split sides of the cooling channel legs <NUM>/<NUM>. For example, the wedge half-angle WA is from <NUM> degrees to <NUM> degrees. In one example, the sides of the cooling channel legs <NUM>/<NUM> are straight, however, in other examples the sides have a beta-spline shape, in which case, the half-angle is taken from the tangent lines.

The turn channel <NUM> is designed to facilitate smooth turning flow of the cooling air from one cooling channel <NUM> to the next. In further examples, the turn channel <NUM> has the features described below with reference also to <FIG> (the ears 86a/86b are excluded in these views). Each of the first and second channel legs <NUM>/<NUM> has a mouth 88a at the end <NUM> of the cooling channel <NUM> (<FIG>) and a crest 88b at an apex of the turn channel <NUM>. The mouth 88a defines a cross-sectional area of W1×T1 (width × thickness), the crest defines a cross-sectional area of W2 × T2 (width × thickness), and the cross-sectional area of the crest 88b is from <NUM>% to <NUM>% of the cross-sectional area of the mouth 88a. Additionally, W2 is from <NUM>% to <NUM>% of W1, and T2 is from <NUM>% to <NUM>% of T1. In one further example, from the channel end <NUM> at the mouth 88a, the channel legs <NUM>/<NUM> increase in thickness and decrease in width up to the crest 88b.

The mouth 88a and crest 88b represent planes, which if extended intersect and a line L. The first and second channel legs <NUM>/<NUM> have an inner turn radius R taken with regard to the line L. The inner turn radius R is greater than the thickness T1 of the cross-sectional area of the mouth 88a by a factor of at least <NUM> and up to <NUM>. Additionally, the first and second channel legs <NUM>/<NUM> have diffusion angles DA taken between radially inner and outer surfaces of each channel leg <NUM> or <NUM>. The diffusion angle is not greater than <NUM> degrees. The attributes above facilitate smooth transition into the turn channel <NUM> and smooth turning of the flow, to reduce flow detachment and avoid pressure loss therefrom.

Additionally, the design of the turn channel <NUM> facilitates maintaining the cooling air flow in the same legs of the cooling channels <NUM> through the turn. For instance, the ears 86a/86b are radially thicker than the channel legs <NUM>/<NUM>. Flow in the ears 86a/86b thus would have to turn to flow out of the ears 86a/86b into the channel legs <NUM>/<NUM>. As a result, the ears 86a/86b contain the flow therein and serve as flow tracks for the cooling air flow from the legs 72a/72b of the cooling channel <NUM>, to turn that flow into the corresponding legs 72a/72b of the receiving cooling channel <NUM> after the turn. Moreover, the cross-sectional areas of the mouth 88a and the crest 88b are substantially equal. As a result, although the flow in the middle section 88c splits into the channel legs <NUM>/<NUM>, the flow is not substantially constricted or diffused. This permits the cooling air to continue flowing at substantially the same flow rate, thereby reducing rate increases or decreases that tend to disrupt flow and cause pressure loss. As will be appreciated, however, the examples according to this disclosure will also find use in other serpentine configurations that have other channel shapes. In that regard, the channel legs <NUM>/<NUM> will be configured to correspond to the cooling channel geometry. Accordingly, the pass-through design herein can be applied to any number of channel configurations.

Claim 1:
An airfoil (<NUM>) for a gas turbine engine comprising:
an airfoil wall (<NUM>) defining a leading end (62a), a trailing end (62b), a first side (62c), and a second side (62d), the airfoil wall circumscribing an interior cavity; and
radially-extending ribs (<NUM>) that partition the interior cavity into first and second cooling channels (<NUM>) and a radial cooling passage (<NUM>) situated between the first and second cooling channels, the first and second cooling channels each extending from a first channel end (<NUM>) to a second channel end (<NUM>); and
a turn channel (<NUM>) connecting the first channel end of the first cooling channel to the first channel end of the second cooling channel, characterised by
the turn channel splitting at the first channel end of the first cooling channel into first and second channel legs (<NUM>; <NUM>) such that there is a region (<NUM>) between the first and second channel legs, the first and second channel legs merging at the first channel end of the second cooling channel, the radial cooling passage extending through the region between the first and second channel legs.