Patent Description:
A gas turbine engine includes a compressor section, a combustor section and a turbine section. Some gas turbine engines may be configured with an axial flow turbine rotor, where combustion product flow generally axially through the turbine section. Other typically smaller gas turbine engines may be configured with a radial flow turbine rotor, where combustion products flow radially into the turbine section, are turned by the radial flow turbine rotor, and flow generally axially out of the turbine section. While known radial flow turbine rotors have various advantages, there is still room in the art for improvement. There is a need in the art, for example, for a relatively small radial flow turbine rotor which can withstand relatively high turbine section temperatures and/or exposure to prolonged elevated turbine section temperatures.

A prior art assembly having the features of the preamble of claim <NUM> is disclosed in <CIT>. Further prior art assemblies for has turbine engines are disclosed in <CIT>, <CIT>, <CIT>, <CIT> and <CIT>.

According to an aspect of the present invention, an assembly is provided for a gas turbine engine, as claimed in claim <NUM>.

The assembly may also include a shaft connected to and extending axially between the compressor rotor and the turbine rotor. The cooling circuit may extend through the shaft from the compressor rotor to the turbine rotor.

The compressor rotor may be configured as a radial flow compressor rotor. The turbine rotor may also or alternatively be configured as a radial flow turbine rotor.

The turbine rotor may include a turbine rotor surface. The cooling circuit may also include an outlet in the turbine rotor surface. The cooling circuit may extend through the turbine rotor to the outlet fluidly coupling the inlet to the outlet.

The turbine rotor surface may be or otherwise include a turbine gas path surface.

The turbine rotor may include a hub and a plurality of blades. The blades may be arranged circumferentially about the hub. The blades may project out from the turbine rotor surface.

The turbine rotor surface may form a peripheral boundary of a bore in the turbine rotor.

The cooling circuit may at least or only include the inlet, the outlet and a passage extending from the inlet to the outlet.

The cooling circuit may also include a second inlet in the compressor rotor. The cooling circuit may extend from the second inlet, through the compressor rotor and into the turbine rotor fluidly coupling the second inlet to the outlet.

The second inlet may be in the gas path surface.

The second inlet may be axially spaced from the inlet along the rotational axis.

The second inlet may be circumferentially spaced from the inlet about the rotational axis.

The cooling circuit may include a first capillary, a second capillary and an artery. The first capillary may extend from the inlet to the artery. The second capillary may extend from the second inlet to the artery. The artery may extend to the outlet.

The first capillary may be or otherwise include a helical capillary. The second capillary may also or alternatively be or otherwise include a helical capillary. The artery may still also or alternatively be or otherwise include a helical artery.

At least a portion of the cooling circuit may spiral about the rotational axis as the cooling circuit extends from the inlet to the outlet.

The assembly may also include a second cooling circuit. The second cooling circuit may include a second inlet in the gas path surface. The second cooling circuit may extend from the second inlet, through the compressor rotor and into the turbine rotor.

<FIG> is a partial, sectional schematic illustration of a gas turbine engine <NUM>. This gas turbine engine <NUM> of <FIG> is a single spool, radial flow gas turbine engine. The gas turbine engine <NUM> may be configured as an auxiliary power unit (APU), a supplemental power unit (SPU) or a primary power unit (PPU) for generating shaft power, electrical power, bleed flow, or other uses for an aircraft. The gas turbine engine <NUM> may alternatively be configured as a turbojet gas turbine engine, a turboshaft gas turbine engine, a turboprop gas turbine engine or any other type of gas turbine engine that generates thrust for propelling the aircraft during flight. The present disclosure, however, is not limited to such an exemplary gas turbine engine nor to aircraft propulsion system applications. For example, the gas turbine engine <NUM> may alternatively include more than one spool and/or be configured in a land based gas turbine engine configured for electrical power generation, an air power generation unit for air mobility, a hybrid power architecture unit, etc..

The gas turbine engine <NUM> of <FIG> extends axially along an axial centerline <NUM> between a forward, upstream airflow inlet <NUM> and an aft, downstream airflow exhaust <NUM>. This axial centerline <NUM> may also be a rotational axis for various components within the gas turbine engine <NUM>.

The gas turbine engine <NUM> includes a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. The gas turbine engine <NUM> also includes a static engine structure <NUM>. This static engine structure <NUM> houses the compressor section <NUM>, the combustor section <NUM> and the turbine section <NUM>. The static engine structure <NUM> of <FIG> also forms the airflow inlet <NUM> and the airflow exhaust <NUM>.

The engine sections <NUM>, <NUM> and <NUM> are arranged sequentially along a (e.g., annular) core flowpath <NUM> as the core flowpath <NUM> extends through the gas turbine engine <NUM> from the airflow inlet <NUM> to the airflow exhaust <NUM>. The compressor section <NUM> and the turbine section <NUM> each include a respective rotor <NUM>, <NUM>. The compressor rotor <NUM> may be configured as a radial flow compressor rotor, which may also be referred to as a radial outflow compressor rotor. The compressor rotor <NUM> of <FIG>, for example, is configured to receive an axial inflow and provide a radial outflow. The compressor rotor <NUM> of <FIG> thereby turns an axial flow radially outward. Similarly, the turbine rotor <NUM> may be configured as a radial flow turbine rotor, which may also be referred to as a radial inflow turbine rotor. The turbine rotor <NUM> of <FIG>, for example, is configured to receive a radial inflow and provide an axial outflow. The turbine rotor <NUM> of <FIG> thereby turns a radial flow axially aft.

The compressor rotor <NUM> is connected to the turbine rotor <NUM> through an engine shaft <NUM>. This shaft <NUM> is rotatably supported by the static engine structure <NUM> through a plurality of bearings <NUM>; e.g., rolling element bearings, thrust bearings, journal bearings, etc..

The combustor section <NUM> includes a (e.g., annular) combustor <NUM> with a (e.g., annular) combustion chamber <NUM>. The combustor <NUM> may be configured as a reverse flow combustor. Inlets ports into the combustion chamber <NUM>, for example, may be arranged at (e.g., on, adjacent or proximate) and/or towards an aft end <NUM> of the combustor <NUM>. An outlet <NUM> from the combustor <NUM> may be arranged axially aft of an inlet <NUM> to the turbine section <NUM>. The combustor <NUM> may also be arranged radially outboard of and/or axially overlap at least a (e.g., aft) portion of the turbine section <NUM>. With this arrangement, the core flowpath <NUM> of <FIG> reverses its directions (e.g., from a forward-to-aft direction to an aft-to-forward direction) a first time as the core flowpath <NUM> extends into the combustion chamber <NUM>. The core flowpath <NUM> of <FIG> then reverses its direction (e.g., from the aft-to-forward direction to the forward-to-aft direction) a second time as the core flowpath <NUM> extends from the combustion chamber <NUM> into the turbine section <NUM>. The present disclosure, however, is not limited to the foregoing exemplary combustor section arrangement.

During operation, air enters the gas turbine engine <NUM> and, more particularly, the core flowpath <NUM> through the airflow inlet <NUM>. The air within the core flowpath <NUM> may be referred to as core air. This core air is compressed by the compressor rotor <NUM> and directed into the combustion chamber <NUM>. Fuel is injected via one or more fuel injectors (not shown) and mixed with the compressed core air to provide a fuel-air mixture. This fuel-air mixture is ignited within the combustion chamber <NUM> via an igniter (not shown), and combustion products thereof flow through the turbine section <NUM> and cause the turbine rotor <NUM> to rotate. This rotation of the turbine rotor <NUM> drives rotation of the compressor rotor <NUM> and, thus, compression of the air received from the airflow inlet <NUM>. An exhaust section <NUM> of the gas turbine engine <NUM> receives the combustion products from the turbine section <NUM>. This exhaust section <NUM> directs the received combustion products out of the gas turbine engine <NUM> through the airflow exhaust <NUM>.

Cycle performance of the gas turbine engine <NUM> may be tied to temperature within the turbine section <NUM>. Generally speaking, increasing the turbine section <NUM> temperature may facilitate increasing gas turbine engine efficiency and/or power generation. However, typical turbine rotor materials may degrade when subject to relatively high turbine section temperatures. A compressor-turbine rotating assembly <NUM> (e.g., a spool) of the present disclosure therefore is configured with internal cooling to facilitate provision of higher turbine section temperatures and/or operation at elevated turbine section temperatures for longer durations.

Referring to <FIG>, the rotating assembly <NUM> includes the compressor rotor <NUM>, the turbine rotor <NUM> and the shaft <NUM>. This rotating assembly <NUM> also includes one or more internal cooling circuits <NUM> configured to provide the internal cooling to the turbine rotor <NUM>.

The compressor rotor <NUM> includes a compressor hub <NUM> and a plurality of compressor blades <NUM>; e.g., compressor vanes. The compressor hub <NUM> of <FIG> extends radially between and to an inner surface <NUM> (e.g., bore surface) of the compressor hub <NUM> and a gas path surface <NUM> of the compressor hub <NUM>. The compressor inner surface <NUM> may form an outer peripheral boundary of an internal bore <NUM> within the compressor rotor <NUM>, which internal compressor bore <NUM> of <FIG> extends axially through the compressor rotor <NUM>. Referring to <FIG>, the compressor gas path surface <NUM> may form a (e.g., radial and/or axial) peripheral boundary of the core flowpath <NUM> within the compressor section <NUM>. Referring again to <FIG>, the compressor hub <NUM> also extends axially between and to the compressor gas path surface <NUM> and an aft, downstream side surface <NUM> of the compressor hub <NUM>.

The compressor blades <NUM> are arranged circumferentially about the compressor hub <NUM> and the axial centerline <NUM> in an annular array. The compressor blades <NUM> are connected to (e.g., formed integral with) the compressor hub <NUM>. Each of the compressor blades <NUM> of <FIG> projects (e.g., axially forward) from the compressor hub <NUM> and its compressor gas path surface <NUM> to a leading edge <NUM> of the respective compressor blade <NUM>, as well as a (e.g., unsupported, unshrouded) side <NUM> of the respective compressor blade <NUM>. Each of the compressor blades <NUM> of <FIG> also projects (e.g., radially outward) from the compressor hub <NUM> and its compressor gas path surface <NUM> to a trailing edge <NUM> of the respective compressor blade <NUM>, as well as the respective compressor blade side <NUM>.

The turbine rotor <NUM> includes a turbine hub <NUM> and a plurality of turbine blades <NUM>; e.g., turbine vanes. The turbine hub <NUM> of <FIG> extends radially between and to an inner surface <NUM> (e.g., bore surface) of the turbine hub <NUM> and a gas path surface <NUM> of the turbine hub <NUM>. The turbine inner surface <NUM> may form an outer peripheral boundary of an internal bore <NUM> within the turbine rotor <NUM>, which internal turbine bore <NUM> of <FIG> extends axially through the turbine rotor <NUM>. This internal turbine bore <NUM> and the internal compressor bore <NUM> may be parts of a common bore internal to the rotating assembly <NUM>, which internal rotating assembly bore may extend axially along the axial centerline <NUM> through the rotating assembly <NUM>. Alternatively, the internal turbine bore <NUM> may be discrete from the internal compressor bore <NUM>. Referring to <FIG>, the turbine gas path surface <NUM> may form a (e.g., radial and/or axial) peripheral boundary of the core flowpath <NUM> within the turbine section <NUM>. Referring again to <FIG>, the turbine hub <NUM> also extends axially between and to the turbine gas path surface <NUM> and a forward, upstream side surface <NUM> of the turbine hub <NUM>.

The turbine blades <NUM> are arranged circumferentially about the turbine hub <NUM> and the axial centerline <NUM> in an annular array. The turbine blades <NUM> are connected to (e.g., formed integral with) the turbine hub <NUM>. Each of the turbine blades <NUM> of <FIG> projects (e.g., radially outward) from the turbine hub <NUM> and its turbine gas path surface <NUM> to a leading edge <NUM> of the respective turbine blade <NUM>, as well as a (e.g., unsupported, unshrouded) side <NUM> of the respective turbine blade <NUM>. Each of the turbine blades <NUM> of <FIG> also projects (e.g., axially aft) from the turbine hub <NUM> and its turbine gas path surface <NUM> to a trailing edge <NUM> of the respective turbine blade <NUM>, as well as the respective turbine blade side <NUM>.

At least a segment (or an entirety) of the shaft <NUM> extends axially along the axial centerline <NUM> between the compressor rotor <NUM> and its compressor hub <NUM> and the turbine rotor <NUM> and its turbine hub <NUM>. The shaft <NUM> is connected to (e.g., formed integral with) the compressor rotor <NUM> and its compressor hub <NUM> and the turbine rotor <NUM> and its turbine hub <NUM>. The shaft <NUM> thereby rotationally couples / links the turbine rotor <NUM> to the compressor rotor <NUM>.

Referring to <FIG>, the internal cooling circuits <NUM> are arranged circumferentially about the axial centerline <NUM> in an annular array. Referring to <FIG> and <FIG>, each of the internal cooling circuits <NUM> may include one or more cooling circuit inlet passages 98A-C (generally referred to as <NUM>) (e.g., capillaries), at least (or only) one cooling circuit outlet passage <NUM> (e.g., artery), one or more cooling circuit inlets 102A-C (generally referred to as <NUM>), and at least (or only) one cooling circuit outlet <NUM>.

Each of the inlet passages <NUM> extends longitudinally between and to a respective one of the circuit inlets <NUM> and the outlet passage <NUM>. Each inlet passage <NUM> may be configured into a forward, upstream portion of the rotating assembly <NUM>, which rotating assembly portion includes one or more of the rotating assembly components <NUM>, <NUM> and <NUM>. Each inlet passage <NUM> of <FIG>, for example, extends from its circuit inlet <NUM> - through the compressor rotor <NUM> and its hub <NUM>, through the shaft <NUM>, and to or into the turbine rotor <NUM> and its hub <NUM> - to the outlet passage <NUM>. The outlet passage <NUM> extends longitudinally between and to the inlet passages <NUM> and the circuit outlet <NUM>. The outlet passage <NUM> may be configured into an aft, downstream portion of the rotating assembly <NUM>, which rotating assembly portion includes at least the turbine rotor <NUM>. The outlet passage <NUM> of <FIG>, for example, extends from the inlet passages <NUM> - within or through the turbine rotor <NUM> and its hub <NUM> - to the circuit outlet <NUM>. The respective internal cooling circuit <NUM> and its passages <NUM> and <NUM> thereby extend through (or within) the rotating assembly <NUM>, and may fluidly couple the circuit inlets <NUM> to the circuit outlet <NUM> in parallel.

Referring to <FIG>, at least a portion or an entirety of one or more or all of the internal cooling circuits <NUM> may each spiral about the axial centerline <NUM> as the respective internal cooling circuit <NUM> extends, for example, from one or more or all of its circuit inlets <NUM> to its circuit outlet <NUM>. More particularly, one or more or all of the inlet passages <NUM> may each have a helical geometry. Each inlet passage <NUM> of <FIG>, for example, extends circumferentially about (e.g., partially or completely around) the axial centerline <NUM> as the respective inlet passage <NUM> extends longitudinally and axially from its circuit inlet <NUM> to the outlet passage <NUM>. The outlet passage <NUM> may also or alternatively have a helical geometry. The outlet passage <NUM> of <FIG>, for example, extends circumferentially about (e.g., partially or completely around) the axial centerline <NUM> as the outlet passage <NUM> extends longitudinally and axially from the inlet passages <NUM> to the outlet passage <NUM>.

A pitch of the outlet passage helical geometry may be selected based on cooling requirements for the turbine rotor <NUM> (see <FIG>). For example, to increase surface area of the outlet passage <NUM> within the turbine rotor <NUM> and, thus, increase turbine rotor cooling, the pitch of the outlet passage helical geometry may be decreased. By contrast, to decrease the surface area of the outlet passage <NUM> within the turbine rotor <NUM> and, thus, decrease turbine rotor cooling, the pitch of the outlet passage helical geometry may be increased. The pitch of the outlet passage helical geometry may be the same as or different (e.g., less) than a pitch of each inlet passage helical geometry.

Referring to <FIG>, one or more or all of the circuit inlets <NUM> for a respective internal cooling circuit <NUM> (see <FIG> and <FIG>) may be disposed in the compressor gas path surface <NUM>. With this arrangement, the respective internal cooling circuit <NUM> (see <FIG> and <FIG>) may draw a quantity of the (e.g., relatively cool and pressurized) core air from the core flowpath <NUM> within the compressor section <NUM> for cooling the rotating assembly <NUM> and its turbine rotor <NUM> (see <FIG>). The circuit inlets <NUM> may be distributed along and/or to a (e.g., concave, pressure) side <NUM> of a respective one of the compressor blades <NUM>. The circuit inlet 102A is disposed at a forward, upstream location 108A at the compressor blade leading edge <NUM>. The circuit inlet 102B may be disposed at an intermediate location 108B. This intermediate location 108B may be axially spaced aft, downstream from the upstream location 108A along the axial centerline <NUM>. The intermediate location 108B may also or alternatively be circumferentially spaced from the upstream location 108A about the axial centerline <NUM>. The circuit inlet 102C may be disposed at an aft, downstream location 108C such that, for example, the intermediate location 108B is between the upstream location 108A and the downstream location 108C. The downstream location 108C may be axially spaced aft, downstream from the intermediate location 108B along the axial centerline <NUM>. The downstream location 108C may also or alternatively be circumferentially spaced from the intermediate location 108B about the axial centerline <NUM>.

Referring to <FIG>, the circuit outlet <NUM> for a respective internal cooling circuit <NUM> (see <FIG> and <FIG>) may be disposed in the turbine inner surface <NUM>. The circuit outlet <NUM> may be disposed at an intermediate location along the turbine rotor <NUM> leaving, for example, an aft, downstream portion of the turbine rotor <NUM> and its turbine hub <NUM> substantially uncooled. Of course, in other embodiments, the one or more of the internal cooling circuits <NUM> (see <FIG> and <FIG>) may extend further aft, downstream along the turbine rotor <NUM> and its turbine hub <NUM>.

Referring to <FIG>, the circuit outlet <NUM> and the outlet passage <NUM> each have an outlet size <NUM>; e.g., a diameter, a maximum width, etc. This outlet size <NUM> may be different (e.g., greater) than an inlet size (e.g., a diameter, a maximum width, etc.) of each circuit inlet <NUM> and each inlet passage <NUM>. The outlet size <NUM>, for example, may be selected such that a cross-sectional area (outlet area) of the circuit outlet <NUM> and/or the outlet passage <NUM> is exactly or approximately (e.g., +/- <NUM>%) equal to a cross-sectional area (inlet area) of each circuit inlet <NUM> and/or each inlet passage <NUM>. Of course, in other embodiments, the outlet area may be different (e.g., greater or less) than the inlet area to decelerate or accelerate cooling air flowing through the respective internal cooling circuit <NUM>.

During operation of the rotating assembly <NUM> of <FIG>, some of the core air is bled from the core flowpath <NUM> within the compressor section <NUM> (see <FIG>) and directed into the internal cooling circuits <NUM> through the circuit inlets <NUM> to provide cooling air. This cooling air is directed to the outlet passages <NUM>. As this cooling air flows through the outlet passages <NUM>, heat energy is transferred from the turbine rotor <NUM> and its turbine hub <NUM> into the cooling air. The heated cooling air is exhausted from the internal cooling circuits <NUM> (e.g., into the internal turbine bore <NUM>) through the circuit outlets <NUM>. The internal cooling circuits <NUM> may thereby utilize some of the relatively cool core air from the compressor section <NUM> (see <FIG>) to cool the turbine rotor <NUM>.

In some embodiments, referring to <FIG>, one or more or all of the internal cooling circuit elements <NUM>, <NUM>, <NUM> and <NUM> may each have a circular cross-sectional geometry. In other embodiments, referring to <FIG>, one or more or all of the internal cooling circuit elements <NUM>, <NUM>, <NUM> and <NUM> may each have a non-circular cross-sectional geometry. Examples of the non-circular cross-sectional geometry include, but are not limited to, an oval cross-sectional geometry (e.g., see <FIG>), a teardrop shaped cross-sectional geometry (e.g., see <FIG>), and a polygonal (e.g., diamond shaped, triangular, rectangular, etc.) cross-sectional geometry (e.g., see <FIG>). The present disclosure, however, is not limited to the foregoing exemplary cross-sectional geometries. In some embodiments, all of the internal cooling circuit elements <NUM>, <NUM>, <NUM> and <NUM> may have a common (e.g., the same) cross-sectional geometry. In other embodiments, some or all of the internal cooling circuit elements <NUM>, <NUM>, <NUM> and <NUM> may have different, unique cross-sectional geometries. Furthermore, in some embodiments, each internal cooling circuit element <NUM>, <NUM>, <NUM>, <NUM> may maintain a uniform cross-sectional geometry along its length. In other embodiments, one or more or all of the internal cooling circuit elements <NUM>, <NUM>, <NUM> and <NUM> may have a cross-sectional geometry that changes along at least a portion of its length.

In some embodiments, referring to <FIG>, one or more or all of the internal cooling circuits <NUM> may each include a manifold <NUM> between the one or more inlet passages <NUM> and the outlet passage <NUM>. Referring to <FIG>, this manifold <NUM> may have a constant cross-sectional geometry along its longitudinal length. Referring to <FIG>, the manifold <NUM> may alternatively be tapered such that an aera of its cross-sectional geometry increased as more inlet passages <NUM> are connected. This manifold <NUM> may be discrete from the outlet passage <NUM>, or configured as an upstream section of the outlet passage <NUM>.

In some embodiments, referring to <FIG>, one or more or all of the circuit outlets <NUM> may each be disposed in the turbine gas path surface <NUM>.

In some embodiments, referring to <FIG>, one or more of the internal cooling circuits 60A, 60B, 60C (generally referred to as <NUM>) may include a single one of the inlet passages 98A, 98B, 98C and a single outlet passage 100A, 100B, 100C (generally referred to as <NUM>). In such embodiments, the circuit outlet 104A, 104B, 104C (generally referred to as <NUM>) may be disposed at a similar, but opposite (e.g., mirror image) location on the turbine rotor <NUM> as the circuit inlet 102A, 102B, 102C is disposed on the compressor rotor <NUM>. For example, where the circuit inlet 102A is disposed at or about a respective compressor blade leading edge <NUM>, the respective circuit outlet 104A for the same internal cooling circuit <NUM> may be disposed at or about a respective turbine blade trailing edge <NUM>, and so on. With such an arrangement, relatively low pressure core air may be used for cooling a relatively cool portion of the turbine rotor <NUM> and its turbine hub <NUM>. By contrast, relatively high pressure core air may be used for cooling a relatively hot portion of the turbine rotor <NUM> and its turbine hub <NUM>.

Referring to <FIG>, <FIG> and <FIG>, the rotating assembly <NUM> and its components <NUM>, <NUM> and <NUM> are formed as a monolithic body. The rotating assembly <NUM> and its components <NUM>, <NUM> and <NUM>, for example, may be additively manufactured, cast, machined and/or otherwise forms as a single, unitary body. By contrast, a non-monolithic body includes components that are discretely formed and subsequently attached together.

Claim 1:
An assembly (<NUM>) for a gas turbine engine (<NUM>), comprising:
a compressor rotor (<NUM>) comprising a gas path surface (<NUM>) wherein the compressor rotor (<NUM>) comprises a hub (<NUM>) and a plurality of blades (<NUM>), the hub (<NUM>) comprises the gas path surface (<NUM>), the plurality of blades (<NUM>) are arranged circumferentially about the hub (<NUM>), and the plurality of blades (<NUM>) project out from the gas path surface (<NUM>);
a turbine rotor (<NUM>) rotatable with the compressor rotor (<NUM>) about a rotational axis (<NUM>); and
a cooling circuit (<NUM>) comprising an inlet (<NUM>) in the gas path surface (<NUM>), the cooling circuit (<NUM>) extending from the inlet (<NUM>), through the compressor rotor (<NUM>) and into the turbine rotor (<NUM>),
characterized in that:
the inlet (<NUM>) is disposed at a leading edge (<NUM>) of a blade (<NUM>) of the plurality of blades (<NUM>) of the compressor rotor (<NUM>); and
at least the compressor rotor (<NUM>) and the turbine rotor (<NUM>) are formed together as a monolithic body.