Patent Description:
A gas turbine engine generally includes a compressor section, a combustion section, and a turbine section. The compressor section progressively increases the pressure of air entering the gas turbine engine and supplies this compressed air to the combustion section. The compressed air and a fuel (e.g., natural gas) mix within the combustion section and burn within one or more combustion chambers to generate high pressure and high temperature combustion gases. The combustion gases flow from the combustion section into the turbine section where they expand to produce work. For example, expansion of the combustion gases in the turbine section may rotate a rotor shaft connected to a generator to produce electricity.

The turbine section generally includes a plurality of rotor blades. Each rotor blade includes an airfoil positioned within the flow of the combustion gases. In this respect, the rotor blades extract kinetic energy and/or thermal energy from the combustion gases flowing through the turbine section. Some rotor blades may include a tip shroud coupled to the radially outer end of the airfoil. The tip shroud reduces the amount of combustion gases leaking past the rotor blade.

The rotor blades generally operate in extremely high temperature environments. As such, the tip shroud of each rotor blade may define various cooling passages through which a coolant may flow. Nevertheless, the cooling passages may only provide limited cooling to certain portions of the tip shroud, such as various fillets. This may limit the operating temperature of the rotor blade and/or the service life of the rotor blade.

<CIT> relates to an internally cooled tip blade shroud. <CIT> relates to a cooled turbine blade. <CIT> relates to a turbine blade tip shroud. <CIT> relates to a turbine blade. <CIT> discloses an insulated cooling passageway for cooling a shroud of a turbine blade and discloses the technical features of the preamble of independent claim <NUM>.

In one aspect, the present disclosure is directed to a rotor blade for a turbomachine as disclosed in independent claim <NUM>.

The rotor blade includes an airfoil and a tip shroud coupled to the airfoil. The tip shroud includes a side surface. The airfoil and the tip shroud define a first cooling passage. The tip shroud further defines a second passage in fluid communication with the first cooling passage. The second cooling passage extends from the first cooling passage to a first outlet defined by the side surface. The first outlet is configured to direct a flow of coolant onto a tip shroud fillet of a first adjacent rotor blade.

In another aspect, the present disclosure is directed to a turbomachine including a turbine section having a plurality of rotor blades. A first rotor blade of the plurality of rotor blades includes an airfoil and a tip shroud coupled to the airfoil. The tip shroud includes a side surface. The airfoil and the tip shroud define a first cooling passage. The tip shroud further defines a second passage in fluid communication with the first cooling passage. The second cooling passage extends from the first cooling passage to a first outlet defined by the side surface. The first outlet is configured to direct a flow of coolant onto a tip shroud fillet of a second rotor blade of the plurality of rotor blades.

Reference will now be made in detail to present embodiments of the technology, one or more examples of which are illustrated in the accompanying drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the technology.

Each example is provided by way of explanation of the technology, not limitation of the technology. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present technology without departing from the scope thereof.

Although an industrial or land-based gas turbine is shown and described herein, the present technology as shown and described herein is not limited to a land-based and/or industrial gas turbine unless otherwise specified in the claims. For example, the technology as described herein may be used in any type of turbomachine including, but not limited to, aviation gas turbines (e.g., turbofans, etc.), steam turbines, and marine gas turbines.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> schematically illustrates a gas turbine engine <NUM>. As shown, the gas turbine engine <NUM> may include an inlet section <NUM>, a compressor section <NUM>, a combustion section <NUM>, a turbine section <NUM>, and an exhaust section <NUM>. The compressor section <NUM> and turbine section <NUM> may be coupled by a shaft <NUM>. The shaft <NUM> may be a single shaft or a plurality of shaft segments coupled together to form the shaft <NUM>.

The turbine section <NUM> may include a rotor shaft <NUM> having a plurality of rotor disks <NUM> (one of which is shown) and a plurality of rotor blades <NUM>. Each rotor blade <NUM> extends radially outward from and interconnects to one of the rotor disks <NUM>. Each rotor disk <NUM>, in turn, may be coupled to a portion of the rotor shaft <NUM> that extends through the turbine section <NUM>. The turbine section <NUM> further includes an outer casing <NUM> that circumferentially surrounds the rotor shaft <NUM> and the rotor blades <NUM>, thereby at least partially defining a hot gas path <NUM> through the turbine section <NUM>.

During operation, the gas turbine engine <NUM> produces mechanical rotational energy, which may, e.g., be used to generate electricity. More specifically, air enters the inlet section <NUM> of the gas turbine engine <NUM>. From the inlet section <NUM>, the air flows into the compressor <NUM>, where it is progressively compressed to provide compressed air to the combustion section <NUM>. The compressed air in the combustion section <NUM> mixes with a fuel to form an air-fuel mixture, which combusts to produce high temperature and high pressure combustion gases <NUM>. The combustion gases <NUM> then flow through the turbine <NUM>, which extracts kinetic and/or thermal energy from the combustion gases <NUM>. This energy extraction rotates the rotor shaft <NUM>, thereby creating mechanical rotational energy for powering the compressor section <NUM> and/or generating electricity. The combustion gases <NUM> exit the gas turbine engine <NUM> through the exhaust section <NUM>.

<FIG> is a side view of an exemplary rotor blade <NUM>, which may be incorporated into the turbine section <NUM> of the gas turbine engine <NUM> in place of the rotor blade <NUM>. As shown, the rotor blade <NUM> defines an axial direction A, a radial direction R, and a circumferential direction C. In general, the axial direction A extends parallel to an axial centerline <NUM> of the shaft <NUM> (<FIG>), the radial direction R extends generally orthogonal to the axial centerline <NUM>, and the circumferential direction C extends generally concentrically around the axial centerline <NUM>. The rotor blade <NUM> may also be incorporated into the compressor section <NUM> of the gas turbine engine <NUM> (<FIG>).

As illustrated in <FIG>, the rotor blade <NUM> may include a dovetail <NUM>, a shank portion <NUM>, and a platform <NUM>. More specifically, the dovetail <NUM> secures the rotor blade <NUM> to the rotor disk <NUM> (<FIG>). The shank portion <NUM> couples to and extends radially outward from the dovetail <NUM>. The platform <NUM> couples to and extends radially outward from the shank portion <NUM>. The platform <NUM> includes a radially outer surface <NUM>, which generally serves as a radially inward flow boundary for the combustion gases <NUM> flowing through the hot gas path <NUM> of the turbine section <NUM> (<FIG>). The dovetail <NUM>, the shank portion <NUM>, and the platform <NUM> may define an intake port <NUM>, which permits a coolant (e.g., bleed air from the compressor section <NUM>) to enter the rotor blade <NUM>. In the embodiment shown in <FIG>, the dovetail <NUM> is an axial entry fir tree-type dovetail. Alternately, the dovetail <NUM> may be any suitable type of dovetail. In fact, the dovetail <NUM>, shank portion <NUM>, and/or platform <NUM> may have any suitable configurations.

Referring now to <FIG> and <FIG>, the rotor blade <NUM> further includes an airfoil <NUM>. In particular, the airfoil <NUM> extends radially outward from the radially outer surface <NUM> of the platform <NUM> to a tip shroud <NUM>. The airfoil <NUM> couples to the platform <NUM> at a root <NUM> (i.e., the intersection between the airfoil <NUM> and the platform <NUM>). In this respect, the airfoil <NUM> defines an airfoil span <NUM> extending between the root <NUM> and the tip shroud <NUM>. The airfoil <NUM> also includes a pressure side surface <NUM> and an opposing suction side surface <NUM> (<FIG>). The pressure side surface <NUM> and the suction side surface <NUM> are joined together or interconnected at a leading edge <NUM> of the airfoil <NUM> and a trailing edge <NUM> of the airfoil <NUM>. As shown, the leading edge <NUM> is oriented into the flow of combustion gases <NUM> (<FIG>), while the trailing edge <NUM> is spaced apart from and positioned downstream of the leading edge <NUM>. The pressure side surface <NUM> and the suction side surface <NUM> are continuous about the leading edge <NUM> and the trailing edge <NUM>. Furthermore, the pressure side surface <NUM> is generally concave, and the suction side surface <NUM> is generally convex.

As shown in <FIG>, the rotor blade <NUM>, and, more particularly, the airfoil <NUM> and the tip shroud <NUM>, may define one or more radially-extending cooling passages <NUM> extending therethrough. More specifically, the radially-extending cooling passages <NUM> may extend from the intake port <NUM> through the airfoil <NUM> to the tip shroud <NUM>. In this respect, coolant may flow through the radially-extending cooling passages <NUM> from the intake port <NUM> to the tip shroud <NUM>. In the embodiment shown in <FIG>, for example, the airfoil <NUM> defines seven radially-extending cooling passages <NUM>. In alternate embodiments, however, the airfoil <NUM> may define more or fewer radially-extending cooling passages <NUM>.

As mentioned above, the rotor blade <NUM> includes the tip shroud <NUM>. As illustrated in <FIG> and <FIG>, the tip shroud <NUM> couples to the radially outer end of the airfoil <NUM> and generally defines the radially outermost portion of the rotor blade <NUM>. In this respect, the tip shroud <NUM> reduces the amount of the combustion gases <NUM> (<FIG>) that escape past the rotor blade <NUM>. As shown, the tip shroud <NUM> may include a seal rail <NUM>. Alternate embodiments, however, may include more seal rails <NUM> (e.g., two seal rails <NUM>, three seal rails <NUM>, etc.) or no seal rails <NUM>.

Referring now to <FIG>, the tip shroud <NUM> includes various surfaces. For example, the tip shroud <NUM> may include a forward side surface <NUM> positioned at a forward end <NUM> of the tip shroud <NUM> and an aft side surface <NUM> positioned at an aft end <NUM> of the tip shroud <NUM>. The tip shroud <NUM> may also include a first pressure side surface <NUM>, a second pressure side surface <NUM>, and a third pressure side surface <NUM> positioned on a pressure side <NUM> of the tip shroud <NUM>. Similarly, the tip shroud <NUM> may also include a first suction side surface <NUM>, a second suction side surface <NUM>, and a third suction side surface <NUM> positioned on a suction side <NUM> of the tip shroud <NUM>. The surfaces <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> maybe collectively referred to as a side surface <NUM>. Furthermore, the tip shroud <NUM> also includes a radially outer surface <NUM> from which the seal rail <NUM> may extend outward. As shown, in some embodiments, the seal rail <NUM> may extend between the second pressure side surface <NUM> and the second suction side surface <NUM>. In alternate embodiments, however, the tip shroud <NUM> may have any suitable combination and/or configuration of surfaces.

In the embodiment shown in <FIG>, the tip shroud <NUM> has a Z-notch configuration. More specifically, the first, second, and third pressure side walls <NUM>, <NUM>, <NUM> define a Z-shape. In this respect, a pressure side convex fillet <NUM> transitions between the first and second pressure side walls <NUM>, <NUM>, while a pressure side concave fillet <NUM> transitions between the second and third pressure side walls <NUM>, <NUM>. The first, second, and third suction side walls <NUM>, <NUM>, <NUM> define a Z-shape that is complementary to the Z-shape of the pressure side walls <NUM>, <NUM>, <NUM>. As such, a suction side concave fillet <NUM> transitions between the first and second suction side walls <NUM>, <NUM>, while a suction side convex fillet <NUM> transitions between the second and third suction side walls <NUM>, <NUM>. In alternate embodiments, however, the tip shroud <NUM> may have any suitable shape and/or configuration.

The tip shroud <NUM> may define pressure side and suction side cooling passages <NUM>, <NUM>. As shown, the cooling passages <NUM>, <NUM> respectively extend from different radially-extending cooling passages <NUM> to pressure side and suction side outlets <NUM>, <NUM> defined by the side surface <NUM>. For example, the pressure side cooling passage <NUM> is fluidly coupled to one of the radially-extending cooling passages <NUM>, such as one of the cooling passages <NUM> positioned forward of the seal rail <NUM>. As such, the pressure side cooling passage <NUM> extends through the tip shroud <NUM> to the pressure side outlet <NUM>. As shown, the cooling passage <NUM> may be positioned forward of the seal rail <NUM> in some embodiments. Similarly, the suction side cooling passage <NUM> is fluidly coupled to another of the radially-extending cooling passages <NUM>, such as one of the cooling passages <NUM> positioned aft of the seal rail <NUM>. In this respect, the suction side cooling passage <NUM> extends through the tip shroud <NUM> to the suction side outlet <NUM>. As shown, the cooling passage <NUM> maybe positioned aft of the seal rail <NUM> in some embodiments. In alternate embodiments, the pressure side cooling passage <NUM> and outlet <NUM> may positioned aft of the seal rail <NUM> and the suction side cooling passage <NUM> and outlet <NUM> may be positioned forward of the seal rail <NUM>. The pressure side and suction side cooling passages <NUM>, <NUM> extend toward the side surface <NUM> in opposite directions and may generally be parallel or substantially parallel to each other as shown in <FIG>. In certain embodiments, the cooling passages <NUM>, <NUM> may extend along the seal rail <NUM>, such as parallel or substantially parallel to the seal rail <NUM>. Although, in other embodiments, the positioning of the cooling passage <NUM>, <NUM> may be independent of the seal rail <NUM>. The tip shroud <NUM> may entirely define the cooling passages <NUM>, <NUM>. Alternatively, the cooling passages <NUM>, <NUM> may extend through the tip shroud <NUM> in any suitable manner. In further embodiments, the tip shroud <NUM> may define only one of the pressure side or suction side cooling passages <NUM>, <NUM>.

As mentioned above, the pressure side and suction side cooling passages <NUM>, <NUM> respectively have pressure side and suction side outlets <NUM>, <NUM> defined by the side surface <NUM>. In the embodiment shown in <FIG>, the pressure side convex fillet <NUM> defines the pressure side outlet <NUM> and the suction side convex fillet <NUM> defines the suction side outlet <NUM>. In this respect, and as will be described in greater detail below, the outlets <NUM>, <NUM> are configured to direct a flow of coolant onto the concave fillets <NUM>, <NUM> of the adjacent rotor blades. The outlets <NUM>, <NUM> are configured to expel the coolant at a sufficient velocity to traverse a gap <NUM> (<FIG>) between the outlet <NUM>, <NUM> and the corresponding concave fillet <NUM>, <NUM> of the adjacent rotor blade such that the coolant impinges on the corresponding concave fillet <NUM>, <NUM> of the adjacent rotor blade. The outlets <NUM>, <NUM> have the same diameter as the corresponding cooling passage <NUM>, <NUM>. In alternate embodiments, however, any suitable portion of the side surface <NUM> may define the outlets <NUM>, <NUM> so long as the outlets <NUM>, <NUM> may be configured to direct the flow of coolant onto suitable tip shroud fillets of the adjacent rotor blades.

Referring now to <FIG>, the tip shroud <NUM> may include a plug <NUM> positioned within a radially outer portion of the radially-extending cooling passages <NUM> to which the pressure side and/or suction side cooling passages <NUM>, <NUM> fluidly couple. As shown, the plug <NUM> may direct coolant <NUM> flowing through the cooling passage <NUM> into the corresponding cooling passage <NUM>, <NUM>. In particular embodiments, the plug <NUM> may direct all of the coolant <NUM> flowing through the cooling passage <NUM> into the corresponding cooling passage <NUM>, <NUM>. The plug <NUM> may be a weld or other suitable structure that occludes the radially outer portion of the corresponding cooling passages <NUM>.

<FIG> illustrates a plurality of adjacent rotor blades <NUM>. As shown, first, second, and third rotor blades 100A, 100B, 100C are axially aligned and circumferentially spaced apart. In this respect, the pressure side outlet <NUM> of the first rotor blade 100A is axially aligned with the suction side concave fillet <NUM> of the second rotor blade 100B. Similarly, the pressure side outlet <NUM> of the second rotor blade 100B is axially aligned with the suction side concave fillet <NUM> of the third rotor blade 100C. Furthermore, the suction side outlet <NUM> of the second rotor blade 100B is axially aligned with the pressure side concave fillet <NUM> of the first rotor blade 100A. Similarly, the suction side outlet <NUM> of the third rotor blade 100A is axially aligned
with the pressure side concave fillet <NUM> of the second rotor blade 100B. In alternate embodiments, the outlets <NUM>, <NUM> of the rotor blades 100A-C may be aligned with any suitable concave tip shroud fillets of the corresponding adjacent rotor blades.

During operation of the gas turbine engine <NUM>, the coolant <NUM> flows through the pressure side and suction side cooling passages <NUM>, <NUM> to respectively cool the pressure side concave fillet <NUM> and the suction side concave fillet <NUM> of the corresponding adjacent rotor blades 100A-C. More specifically, the coolant <NUM> (e.g., bleed air from the compressor section <NUM>) enters the rotor blade <NUM> through the intake port <NUM> (<FIG>). At least a portion of the coolant <NUM> flows through the cooling passages <NUM> in the airfoil <NUM> and into the pressure side and/or suction side cooling passages <NUM>, <NUM>. The coolant <NUM> exits the cooling passages <NUM>, <NUM> respectively through the outlets <NUM>, <NUM> and impinges on the adjacent concave fillets <NUM>, <NUM>, thereby cooling the concave fillets <NUM>, <NUM>. In particular, the coolant <NUM> exiting the outlet <NUM> of the first rotor blade 100A impinges on the concave fillet <NUM> of the second rotor blade 100B. Similarly, the coolant <NUM> exiting the outlet <NUM> of the second rotor blade 100B impinges on the concave fillet <NUM> of the third rotor blade 100C. Furthermore, the coolant <NUM> exiting the outlet <NUM> of the second rotor blade impinges on the concave fillet <NUM> of the first rotor blade 100A. Similarly, the coolant <NUM> exiting the outlet <NUM> of the third rotor blade 100A impinges on the concave fillet <NUM> of the second rotor blade 100B. In this respect, the outlets <NUM>, <NUM> expel the coolant at a sufficient velocity to traverse the gaps <NUM> between the outlets <NUM>, <NUM> and the corresponding concave fillets <NUM>, <NUM> to facilitate such impingement cooling.

As described in greater detail, above the rotor blade <NUM> includes a tip shroud <NUM> that defines a pressure side and/or suction side cooling passage <NUM>, <NUM>, which direct the coolant <NUM> onto the fillets <NUM>, <NUM> of adjacent rotor blades. In this respect, the rotor blade <NUM> provides greater cooling to the fillets <NUM>, <NUM> of the tip shroud <NUM> than conventional rotor blades. As such, the rotor blade <NUM> may be able to withstand higher operating temperatures and/or have a longer service life than conventional rotor blades.

Claim 1:
A rotor blade (<NUM>) for a turbomachine (<NUM>), the rotor blade (<NUM>) comprising:
an airfoil (<NUM>); and
a tip shroud (<NUM>) coupled to the airfoil (<NUM>), the tip shroud (<NUM>) including a side surface (<NUM>), the airfoil (<NUM>) and the tip shroud (<NUM>) defining a first cooling passage (<NUM>), the tip shroud (<NUM>) further defining a second cooling passage (<NUM>, <NUM>) in fluid communication with the first cooling passage (<NUM>), the second cooling passage (<NUM>, <NUM>) extending from the first cooling passage (<NUM>) to a first outlet (<NUM>, <NUM>) defined by the side surface (<NUM>), the first outlet (<NUM>, <NUM>) being configured such that in use a coolant (<NUM>) is directable onto a concave tip shroud fillet (<NUM>, <NUM>) of a first adjacent rotor blade (<NUM>);
characterized in that;
the first outlet (<NUM>, <NUM>) has the same diameter as the corresponding cooling passage (<NUM>, <NUM>) and wherein the first outlet (<NUM>, <NUM>) is configured to expel the coolant (<NUM>) at a sufficient velocity to traverse a gap (<NUM>) between the first outlet (<NUM>, <NUM>) and the concave tip shroud fillet (<NUM>, <NUM>) such that the coolant (<NUM>) impinges on the concave tip shroud fillet (<NUM>, <NUM>) of the first adjacent rotor blade (<NUM>).