Patent ID: 12203386

DETAILED DESCRIPTION

FIG.1is a partial, sectional schematic illustration of a gas turbine engine20. This gas turbine engine20ofFIG.1is a single spool, radial flow gas turbine engine. The gas turbine engine20may 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 engine20may 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 engine20may 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 engine20ofFIG.1extends axially along an axial centerline22between a forward, upstream airflow inlet24and an aft, downstream airflow exhaust26. This axial centerline22may also be a rotational axis for various components within the gas turbine engine20.

The gas turbine engine20includes a compressor section28, a combustor section30and a turbine section32. The gas turbine engine20also includes a static engine structure34. This static engine structure34houses the compressor section28, the combustor section30and the turbine section32. The static engine structure34ofFIG.1also forms the airflow inlet24and the airflow exhaust26.

The engine sections28,30and32are arranged sequentially along a (e.g., annular) core flowpath36as the core flowpath36extends through the gas turbine engine20from the airflow inlet24to the airflow exhaust26. The compressor section28and the turbine section32each include a respective rotor38,40. The compressor rotor38may be configured as a radial flow compressor rotor, which may also be referred to as a radial outflow compressor rotor. The compressor rotor38ofFIG.1, for example, is configured to receive an axial inflow and provide a radial outflow. The compressor rotor38ofFIG.1thereby turns an axial flow radially outward. Similarly, the turbine rotor40may be configured as a radial flow turbine rotor, which may also be referred to as a radial inflow turbine rotor. The turbine rotor40ofFIG.1, for example, is configured to receive a radial inflow and provide an axial outflow. The turbine rotor40ofFIG.1thereby turns a radial flow axially aft.

The compressor rotor38is connected to the turbine rotor40through an engine shaft42. This shaft42is rotatably supported by the static engine structure34through a plurality of bearings44; e.g., rolling element bearings, thrust bearings, journal bearings, etc.

The combustor section30includes a (e.g., annular) combustor46with a (e.g., annular) combustion chamber48. The combustor46may be configured as a reverse flow combustor. Inlets ports into the combustion chamber48, for example, may be arranged at (e.g., on, adjacent or proximate) and/or towards an aft end50of the combustor46. An outlet52from the combustor46may be arranged axially aft of an inlet54to the turbine section32. The combustor46may also be arranged radially outboard of and/or axially overlap at least a (e.g., aft) portion of the turbine section32. With this arrangement, the core flowpath36ofFIG.1reverses its directions (e.g., from a forward-to-aft direction to an aft-to-forward direction) a first time as the core flowpath36extends into the combustion chamber48. The core flowpath36ofFIG.1then reverses its direction (e.g., from the aft-to-forward direction to the forward-to-aft direction) a second time as the core flowpath36extends from the combustion chamber48into the turbine section32. The present disclosure, however, is not limited to the foregoing exemplary combustor section arrangement.

During operation, air enters the gas turbine engine20and, more particularly, the core flowpath36through the airflow inlet24. The air within the core flowpath36may be referred to as core air. This core air is compressed by the compressor rotor38and directed into the combustion chamber48. 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 chamber48via an igniter (not shown), and combustion products thereof flow through the turbine section32and cause the turbine rotor40to rotate. This rotation of the turbine rotor40drives rotation of the compressor rotor38and, thus, compression of the air received from the airflow inlet24. An exhaust section56of the gas turbine engine20receives the combustion products from the turbine section32. This exhaust section56directs the received combustion products out of the gas turbine engine20through the airflow exhaust26.

Cycle performance of the gas turbine engine20may be tied to temperature within the turbine section32. Generally speaking, increasing the turbine section32temperature 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 assembly58(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 toFIG.2, the rotating assembly58includes the compressor rotor38, the turbine rotor40and the shaft42. This rotating assembly58also includes one or more internal cooling circuits60configured to provide the internal cooling to the turbine rotor40.

The compressor rotor38includes a compressor hub62and a plurality of compressor blades64; e.g., compressor vanes. The compressor hub62ofFIG.2extends radially between and to an inner surface66(e.g., bore surface) of the compressor hub62and a gas path surface68of the compressor hub62. The compressor inner surface66may form an outer peripheral boundary of an internal bore70within the compressor rotor38, which internal compressor bore70ofFIG.2extends axially through the compressor rotor38. Referring toFIG.1, the compressor gas path surface68may form a (e.g., radial and/or axial) peripheral boundary of the core flowpath36within the compressor section28. Referring again toFIG.2, the compressor hub62also extends axially between and to the compressor gas path surface68and an aft, downstream side surface72of the compressor hub62.

The compressor blades64are arranged circumferentially about the compressor hub62and the axial centerline22in an annular array. The compressor blades64are connected to (e.g., formed integral with) the compressor hub62. Each of the compressor blades64ofFIG.2projects (e.g., axially forward) from the compressor hub62and its compressor gas path surface68to a leading edge74of the respective compressor blade64, as well as a (e.g., unsupported, unshrouded) side76of the respective compressor blade64. Each of the compressor blades64ofFIG.2also projects (e.g., radially outward) from the compressor hub62and its compressor gas path surface68to a trailing edge78of the respective compressor blade64, as well as the respective compressor blade side76.

The turbine rotor40includes a turbine hub80and a plurality of turbine blades82; e.g., turbine vanes. The turbine hub80ofFIG.2extends radially between and to an inner surface84(e.g., bore surface) of the turbine hub80and a gas path surface86of the turbine hub80. The turbine inner surface84may form an outer peripheral boundary of an internal bore88within the turbine rotor40, which internal turbine bore88ofFIG.2extends axially through the turbine rotor40. This internal turbine bore88and the internal compressor bore70may be parts of a common bore internal to the rotating assembly58, which internal rotating assembly bore may extend axially along the axial centerline22through the rotating assembly58. Alternatively, the internal turbine bore88may be discrete from the internal compressor bore70. Referring toFIG.1, the turbine gas path surface86may form a (e.g., radial and/or axial) peripheral boundary of the core flowpath36within the turbine section32. Referring again toFIG.2, the turbine hub80also extends axially between and to the turbine gas path surface86and a forward, upstream side surface90of the turbine hub80.

The turbine blades82are arranged circumferentially about the turbine hub80and the axial centerline22in an annular array. The turbine blades82are connected to (e.g., formed integral with) the turbine hub80. Each of the turbine blades82ofFIG.2projects (e.g., radially outward) from the turbine hub80and its turbine gas path surface86to a leading edge92of the respective turbine blade82, as well as a (e.g., unsupported, unshrouded) side94of the respective turbine blade82. Each of the turbine blades82ofFIG.2also projects (e.g., axially aft) from the turbine hub80and its turbine gas path surface86to a trailing edge96of the respective turbine blade82, as well as the respective turbine blade side94.

At least a segment (or an entirety) of the shaft42extends axially along the axial centerline22between the compressor rotor38and its compressor hub62and the turbine rotor40and its turbine hub80. The shaft42is connected to (e.g., formed integral with) the compressor rotor38and its compressor hub62and the turbine rotor40and its turbine hub80. The shaft42thereby rotationally couples/links the turbine rotor40to the compressor rotor38.

Referring toFIG.3, the internal cooling circuits60are arranged circumferentially about the axial centerline22in an annular array. Referring toFIGS.2and4, each of the internal cooling circuits60may include one or more cooling circuit inlet passages98A-C (generally referred to as98) (e.g., capillaries), at least (or only) one cooling circuit outlet passage100(e.g., artery), one or more cooling circuit inlets102A-C (generally referred to as102), and at least (or only) one cooling circuit outlet104.

Each of the inlet passages98extends longitudinally between and to a respective one of the circuit inlets102and the outlet passage100. Each inlet passage98may be configured into a forward, upstream portion of the rotating assembly58, which rotating assembly portion includes one or more of the rotating assembly components38,42and62. Each inlet passage98ofFIG.2, for example, extends from its circuit inlet102—through the compressor rotor38and its hub62, through the shaft42, and to or into the turbine rotor40and its hub80—to the outlet passage100. The outlet passage100extends longitudinally between and to the inlet passages98and the circuit outlet104. The outlet passage100may be configured into an aft, downstream portion of the rotating assembly58, which rotating assembly portion includes at least the turbine rotor40. The outlet passage100ofFIG.2, for example, extends from the inlet passages98—within or through the turbine rotor40and its hub80—to the circuit outlet104. The respective internal cooling circuit60and its passages98and100thereby extend through (or within) the rotating assembly58, and may fluidly couple the circuit inlets102to the circuit outlet104in parallel.

Referring toFIG.4, at least a portion or an entirety of one or more or all of the internal cooling circuits60may each spiral about the axial centerline22as the respective internal cooling circuit60extends, for example, from one or more or all of its circuit inlets102to its circuit outlet104. More particularly, one or more or all of the inlet passages98may each have a helical geometry. Each inlet passage98ofFIG.4, for example, extends circumferentially about (e.g., partially or completely around) the axial centerline22as the respective inlet passage98extends longitudinally and axially from its circuit inlet102to the outlet passage100. The outlet passage100may also or alternatively have a helical geometry. The outlet passage100ofFIG.4, for example, extends circumferentially about (e.g., partially or completely around) the axial centerline22as the outlet passage100extends longitudinally and axially from the inlet passages98to the outlet passage100.

A pitch of the outlet passage helical geometry may be selected based on cooling requirements for the turbine rotor40(seeFIG.2). For example, to increase surface area of the outlet passage100within the turbine rotor40and, 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 passage100within the turbine rotor40and, 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 toFIG.5, one or more or all of the circuit inlets102for a respective internal cooling circuit60(seeFIGS.2and4) may be disposed in the compressor gas path surface68. With this arrangement, the respective internal cooling circuit60(seeFIGS.2and4) may draw a quantity of the (e.g., relatively cool and pressurized) core air from the core flowpath36within the compressor section28for cooling the rotating assembly58and its turbine rotor40(seeFIG.2). The circuit inlets102may be distributed along and/or to a (e.g., concave, pressure) side106of a respective one of the compressor blades64. The circuit inlet102A, for example, may be disposed at a forward, upstream location108A; e.g., at or about the compressor blade leading edge74. The circuit inlet102B may be disposed at an intermediate location108B. This intermediate location108B may be axially spaced aft, downstream from the upstream location108A along the axial centerline22. The intermediate location108B may also or alternatively be circumferentially spaced from the upstream location108A about the axial centerline22. The circuit inlet102C may be disposed at an aft, downstream location108C such that, for example, the intermediate location108B is between the upstream location108A and the downstream location108C. The downstream location108C may be axially spaced aft, downstream from the intermediate location108B along the axial centerline22. The downstream location108C may also or alternatively be circumferentially spaced from the intermediate location108B about the axial centerline22.

Referring toFIG.6, the circuit outlet104for a respective internal cooling circuit60(seeFIGS.2and4) may be disposed in the turbine inner surface84. The circuit outlet104may be disposed at an intermediate location along the turbine rotor40leaving, for example, an aft, downstream portion of the turbine rotor40and its turbine hub80substantially uncooled. Of course, in other embodiments, the one or more of the internal cooling circuits60(seeFIGS.2and4) may extend further aft, downstream along the turbine rotor40and its turbine hub80.

Referring toFIG.7, the circuit outlet104and the outlet passage100each have an outlet size110; e.g., a diameter, a maximum width, etc. This outlet size110may be different (e.g., greater) than an inlet size (e.g., a diameter, a maximum width, etc.) of each circuit inlet102and each inlet passage98. The outlet size110, for example, may be selected such that a cross-sectional area (outlet area) of the circuit outlet104and/or the outlet passage100is exactly or approximately (e.g., +/−5%) equal to a cross-sectional area (inlet area) of each circuit inlet102and/or each inlet passage98. 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 circuit60.

During operation of the rotating assembly58ofFIG.2, some of the core air is bled from the core flowpath36within the compressor section28(seeFIG.1) and directed into the internal cooling circuits60through the circuit inlets102to provide cooling air. This cooling air is directed to the outlet passages100. As this cooling air flows through the outlet passages100, heat energy is transferred from the turbine rotor40and its turbine hub80into the cooling air. The heated cooling air is exhausted from the internal cooling circuits60(e.g., into the internal turbine bore88) through the circuit outlets104. The internal cooling circuits60may thereby utilize some of the relatively cool core air from the compressor section28(seeFIG.1) to cool the turbine rotor40.

In some embodiments, referring toFIG.8A, one or more or all of the internal cooling circuit elements98,100,102and104may each have a circular cross-sectional geometry. In other embodiments, referring toFIGS.8B-D, one or more or all of the internal cooling circuit elements98,100,102and104may 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., seeFIG.8B), a teardrop shaped cross-sectional geometry (e.g., seeFIG.8C), and a polygonal (e.g., diamond shaped, triangular, rectangular, etc.) cross-sectional geometry (e.g., seeFIG.8D). The present disclosure, however, is not limited to the foregoing exemplary cross-sectional geometries. In some embodiments, all of the internal cooling circuit elements98,100,102and104may have a common (e.g., the same) cross-sectional geometry. In other embodiments, some or all of the internal cooling circuit elements98,100,102and104may have different, unique cross-sectional geometries. Furthermore, in some embodiments, each internal cooling circuit element98,100,102,104may maintain a uniform cross-sectional geometry along its length. In other embodiments, one or more or all of the internal cooling circuit elements98,100,102and104may have a cross-sectional geometry that changes along at least a portion of its length.

In some embodiments, referring toFIGS.9A and9B, one or more or all of the internal cooling circuits60may each include a manifold114between the one or more inlet passages98and the outlet passage100. Referring toFIG.9A, this manifold114may have a constant cross-sectional geometry along its longitudinal length. Referring toFIG.9B, the manifold114may alternatively be tapered such that an area of its cross-sectional geometry increased as more inlet passages98are connected. This manifold114may be discrete from the outlet passage100, or configured as an upstream section of the outlet passage100.

In some embodiments, referring toFIGS.10and11, one or more or all of the circuit outlets104may each be disposed in the turbine gas path surface86.

In some embodiments, referring toFIG.11, one or more of the internal cooling circuits60A,60B,60C (generally referred to as60) may include a single one of the inlet passages98A,98B,98C and a single outlet passage100A,100B,100C (generally referred to as100). In such embodiments, the circuit outlet104A,104B,104C (generally referred to as104) may be disposed at a similar, but opposite (e.g., mirror image) location on the turbine rotor40as the circuit inlet102A,102B,102C is disposed on the compressor rotor38. For example, where the circuit inlet102A is disposed at or about a respective compressor blade leading edge74, the respective circuit outlet104A for the same internal cooling circuit60may be disposed at or about a respective turbine blade trailing edge96, and so on. With such an arrangement, relatively low pressure core air may be used for cooling a relatively cool portion of the turbine rotor40and its turbine hub80. By contrast, relatively high pressure core air may be used for cooling a relatively hot portion of the turbine rotor40and its turbine hub80.

In some embodiments, referring toFIGS.1,2and11, the rotating assembly58and its components38,40and42may be formed as a monolithic body. The rotating assembly58and its components38,40and42, 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. The present disclosure, however, is not limited to the foregoing exemplary formation processes.

While various embodiments of the present disclosure have been described, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the disclosure. For example, the present disclosure as described herein includes several aspects and embodiments that include particular features. Although these features may be described individually, it is within the scope of the present disclosure that some or all of these features may be combined with any one of the aspects and remain within the scope of the disclosure. Accordingly, the present disclosure is not to be restricted except in light of the attached claims and their equivalents.