Patent Description:
Gas turbine rotor assemblies include successive rows of blades that extend from respective rotor disks that may be arranged in an axially stacked configuration. The rotor stack may be assembled through a multitude of systems such as fasteners, fusion, tie-shafts and various combinations thereof.

Gas turbine rotor assemblies operate in an environment where significant pressure and temperature differentials exist across component boundaries that primarily separate a core gas flow path, a secondary cooling flow path, and relatively static cavities such as rotor bores located axially between rotor disks. For high-pressure and high-temperature applications, the components experience thermo-mechanical fatigue (TMF) across these boundaries. Although resistant to the effects of TMF, the components may be of a heavier-than-optimal weight to prevent TMF and achieve the desired performance requirements.

<CIT> discloses a spoked rotor for a gas turbine engine.

<CIT> discloses a coolant recovery type gas turbine.

A rotor assembly of a gas turbine engine according to one aspect of the invention is provided by claim <NUM>.

In the additionally thereto, in the foregoing embodiment, the passage is annular.

In the alternative or additionally thereto, in the foregoing embodiment, the rotor is a spoked rotor including a plurality of circumferentially spaced first spokes generally located radially between the platforms and the rotor disks, and the first channel is one of a plurality of first channels with each first channel defined circumferentially between adjacent first spokes.

In the alternative or additionally thereto, in the foregoing embodiment, the spacer includes radially inner and outer rings and a plurality of circumferentially spaced second spokes spanning radially between the inner and outer rings, and the second channel is one of a plurality of second channels with each second channel defined circumferentially between adjacent second spokes.

In the alternative or additionally thereto, in the foregoing embodiment, the rotor disk includes a rim connected to the plurality of first spokes and defining in-part the plurality of first channels.

In the alternative or additionally thereto, in the foregoing embodiment, cooling air flows through the plurality of second channels, then through the plurality of first channels, then into the passage for cooling of the plurality of the spacers, the plurality of platforms, the rim and the hub.

In the alternative or additionally thereto, in the foregoing embodiment, the assembly includes a structure extending axially and disposed radially inward of the rotor disk and the shell, wherein the structure defines at least in-part a supply conduit in fluid communication between the passage and a rotor bore defined at least in-part between adjacent rotor disks.

A method of operating a secondary flowpath system of a gas turbine engine according to another aspect of the invention is provided by claim <NUM>.

It should be understood, however, the following description and figures are intended to be exemplary in nature and non-limiting.

<FIG> schematically illustrates a gas turbine engine <NUM> disclosed as a two-spool turbo fan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. The fan section <NUM> drives air along a bypass flowpath while the compressor section <NUM> drives air along a core gas flowpath (see arrow C) for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engine architecture such as turbojets, turboshafts, three-spool turbofans, land-based turbine engines, and others.

The engine <NUM> generally includes a low spool <NUM> and a high spool <NUM> mounted for rotation about an engine axis A via several bearing structures <NUM> and relative to a static engine case <NUM>. The low spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM> of the fan section <NUM>, a low pressure compressor <NUM> (LPC) of the compressor section <NUM> and a low pressure turbine <NUM> (LPT) of the turbine section <NUM>. The inner shaft <NUM> drives the fan <NUM> directly, or, through a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low spool <NUM>. An exemplary reduction transmission may be an epicyclic transmission, namely a planetary or star gear system.

The high spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure compressor <NUM> (HPC) of the compressor section <NUM> and a high pressure turbine <NUM> (HPT) of the turbine section <NUM>. A combustor <NUM> of the combustor section <NUM> is arranged between the HPC <NUM> and the HPT <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate about the engine axis A. Core airflow is compressed by the LPC <NUM> then the HPC <NUM>, mixed with the fuel and burned in the combustor <NUM>, then expanded over the HPT <NUM> and the LPT <NUM>. The LPT <NUM> and HPT <NUM> rotationally drive the respective low spool <NUM> and high spool <NUM> in response to the expansion.

In one non-limiting example, the gas turbine engine <NUM> is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> bypass ratio is greater than about six (<NUM>:<NUM>). The geared architecture <NUM> can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about <NUM>:<NUM>, and in another example is greater than about <NUM>:<NUM>. The geared turbofan enables operation of the low spool <NUM> at higher speeds that can increase the operational efficiency of the LPC <NUM> and LPT <NUM> and render increased pressure in a fewer number of stages.

A pressure ratio associated with the LPT <NUM> is pressure measured prior to the inlet of the LPT <NUM> as related to the pressure at the outlet of the LPT <NUM> prior to an exhaust nozzle of the gas turbine engine <NUM>. In one non-limiting example, the bypass ratio of the gas turbine engine <NUM> is greater than about ten (<NUM>:<NUM>); the fan diameter is significantly larger than the LPC <NUM>; and the LPT <NUM> has a pressure ratio that is greater than about five (<NUM>:<NUM>). It should be understood; however, that the above parameters are only exemplary of one example of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

In one non-limiting example, a significant amount of thrust is provided by the bypass flow path (see arrow B) due to the high bypass ratio. The fan section <NUM> of the gas turbine engine <NUM> is designed for a particular flight condition - typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>,<NUM> meters). This flight condition, with the gas turbine engine <NUM> at its best fuel consumption, is also known as Thrust Specific Fuel consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section <NUM> without the use of a fan exit guide vane system. The low Fan Pressure Ratio according to one, non-limiting, example of the gas turbine engine <NUM> is less than <NUM>: <NUM>. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (T/<NUM>)<NUM>, where "T" represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting example of the gas turbine engine <NUM> is less than about <NUM>,<NUM> feet per second (<NUM> meters per second).

Referring to <FIG>, the HPC <NUM> includes a rotor assembly <NUM>, assembled from a plurality of successive HPC rotors <NUM> that alternate with HPC, ring-shaped, spacers <NUM> of the assembly arranged in an axially stacked configuration. The rotor stack may be assembled in a compressed tie-shaft configuration, with a central shaft (not shown) assembled concentrically within the rotor stack and secured with a nut (not shown), to generate a preload that compresses and retains the HPC rotor <NUM> with the HPC spacers <NUM> together as a spool. Friction at the interfaces between the HPC rotor <NUM> and the HPC spacers <NUM> may be solely responsible to prevent rotation between adjacent rotor hardware. That is, the rotors <NUM> and the spacers <NUM> generally rotate as one piece.

Referring to <FIG>, each HPC rotor <NUM> generally includes a plurality of blades <NUM> circumferentially spaced around a rotor disk <NUM>. The rotor disk <NUM> generally includes a hub <NUM>, a rim <NUM>, and a web <NUM> that spans between the hub and rim. Each blade <NUM> may include an attachment section or root <NUM>, a platform <NUM> and an airfoil <NUM>. Although not shown, static vane rings or vane stages may be spaced radially outward from respective spacers <NUM> to redirect the core airflow upon the adjacent, aft, blades <NUM>.

The HPC rotor <NUM> may be a hybrid dual alloy, integrally bladed, rotor (IBR), with the blades <NUM> being manufactured of one type of material and the rotor disk <NUM> manufactured of a different material. Bi-metal construction provides material capability to separately address varying temperature requirements. For example, the blades <NUM> may be manufactured of a single crystal nickel alloy that are transient liquid phase bonded with the rotor disk <NUM> that is manufactured of a different material such as an extruded billet nickel alloy. Alternatively, or in addition to different materials, the blades <NUM> may be subject to a first type of heat treatment and the rotor disks <NUM> to a different heat treatment thereby providing different material characteristics. That is, the bi-metal construction may include different chemical compositions as well as different treatments of the same chemical compositions such as that provided by different heat treatments.

Referring to <FIG>, a spoke <NUM> may be generally carried between the rim <NUM> of the disk <NUM> and the root <NUM> of the blade <NUM>. The spokes <NUM> may be circumferentially spaced from one another, thereby defining axial or semi-axial channels or slots <NUM> between adjacent spokes <NUM>, and as part of a secondary air flowpath (see arrow S). The spokes <NUM> may be machined, cut with a wire EDM or other processes to provide the desired shape. An interface <NUM> of each spoke <NUM> generally defines the transient liquid phase bond and/or heat transition between the blades <NUM> and the rotor disk <NUM>. That is, the spoke <NUM> contains the interface <NUM>, and 'heat treat transition' is generally the transition between different heat treatments.

Different from traditional HPC designs that lack spokes, the spokes <NUM> of the present disclosure provides a reduced area that is not subject to thermo-mechanical fatigue (TMF) seen in more traditional outer diameter walls (i.e. blade platforms and rotor rim) due to the relatively high temperature gradient. In the present disclosure, the blade platforms <NUM> are exposed to the relatively hot core gas flowpath C; however, the radially inward spoked configuration acts to segment the hot outer diameter wall thus allowing for thermal growth between the platforms <NUM> and the rotor rim <NUM> of the rotor disk <NUM>. The spoked configuration further provides the cooling channel <NUM> of the secondary flowpath S that thermally isolates the rotor rim <NUM> from the core gas flowpath C, thereby minimizing the thermal gradient between the rim <NUM> and the disk hub <NUM>.

Referring to <FIG> and <FIG>, the HPC spacers <NUM> may have a similar architecture to adjacent portions of the HPC rotors <NUM>. That is, each spacer <NUM> may include an outer ring <NUM> that spans axially between and seals to adjacent platforms <NUM> of the blades <NUM>, an inner ring <NUM> spaced radially inward of the outer ring <NUM> and spanning axially between the adjacent rims <NUM> of the rotor disks <NUM>, and a plurality of spokes <NUM> spaced circumferentially from one-another and each spanning radially between and generally engaged to the outer and inner rings <NUM>, <NUM>. A plurality of channels or slots <NUM> may each be defined circumferentially between respective and adjacent spokes <NUM> and radially between the outer and inner rings <NUM>, <NUM>. Each spoke <NUM> may have an interface <NUM> that generally bisects each spoke and provides the bond between the outer and inner rings <NUM>, <NUM>.

In one, non-limiting, example, the outer rings <NUM> may be manufactured of the same material as the blades <NUM>, and the inner ring <NUM> may be manufactured of the same material as the rotor disks <NUM>. Alternatively, the HPC spacers <NUM> may be manufactured of a single material but subjected to the different heat treatments that transition within the spokes <NUM>. In another example, a relatively low temperature configuration will benefit from usage of a single material such that the spokes <NUM> facilitate a weight reduction. In another example, low-temperature bi-metal designs may further benefit from dissimilar materials for weight reduction where, for example, low density materials may be utilized when load carrying capability is less of a concern.

The rotor geometry provided by the spokes <NUM>, <NUM> reduces the conduction of core gas flowpath C heat to the rotor disk <NUM> and the sealing inner ring <NUM> of the HPC spacer <NUM>. Furthermore, the spokes <NUM>, <NUM>, respective channels <NUM>, <NUM>, and cooling airflow therein; enable an IBR rotor to withstand increased exit temperatures of the high pressure compressor (T3 levels) with currently available materials. Rim cooling may also be reduced from conventional allocation. In addition, the overall configuration provides weight reduction at similar stress levels to more traditional configurations.

Referring to <FIG>, the channels <NUM>, <NUM> that flank the respective spokes <NUM>, <NUM> may receive airflow from an upstream, inlet, HPC airflow supply ring <NUM> that may be a spacer similar to spacer <NUM>. The supply ring <NUM> includes opposite outer and inner surfaces <NUM>, <NUM> that may both be substantially cylindrical. A plurality of circumferentially spaced flow ducts <NUM> are in and defined by the supply ring <NUM> and may be ramped. Each duct <NUM> may include an inlet <NUM> generally defined by the outer surface <NUM> and an outlet <NUM> in direct fluid communication with respective channels <NUM> of the adjacent HPC rotor <NUM>; which, in-turn, are in direct fluid communication with the channels <NUM> of the adjacent, downstream, spacer <NUM>. The outer surface <NUM> may define in-part the core gas flowpath C, thus the inlet <NUM> may be in direct fluid communication with the core gas flowpath C at an upstream pressure stage location.

It is further contemplated and understood that various flow paths may be defined through various configurations of the inlet ring <NUM>. For example, the inlet ring <NUM> may draw cooling flow from the core gas flowpath C flow, secondary cooling flow, or combinations thereof. The cooling airflow may be communicated not only forward to aft toward the turbine section <NUM>, but also aft to forward within the engine <NUM>. Further, the airflow may be drawn from adjacent static structure such as vanes to effect boundary flow turbulence as well as other flow conditions. That is, the HPC spacers <NUM> and the inlet ring <NUM> may facilitate through-flow for use in rim cooling, purge air for use down-stream in the compressor, and turbine or bearing compartment operation.

Referring to <FIG>, a second, non-limiting embodiment of an inlet ring is illustrated wherein like elements have like identifying numerals except with the addition of a prime symbol as a suffix. An inlet ring <NUM>' of the second embodiment may have flow ducts <NUM>' having an inlet <NUM>' that is carried and defined by an inner surface <NUM>' of the ring <NUM>'.

Referring to <FIG>, each inner ring <NUM> of the spacers <NUM> may include an upstream circumferential flange <NUM> and a downstream circumferential flange <NUM> that may be located radially inboard of, and thereby captured radially, by the respective, adjacent, rotor rims <NUM>. That is, each inner ring <NUM> is engaged through the stacked configuration of the HPC <NUM>. In the disclosed tie-shaft configuration with multi-metal rotors, the stacked configuration is arranged to accommodate the relatively lower load capability alloys on the core gas flowpath C side of the rotor hardware, while maintaining the load carrying capability between the inner rings <NUM> and the rims <NUM> of the rotor disks <NUM> to transmit rotor torque and carry the centrifugal loads of the blades and segments <NUM>.

The alternating rotor rim <NUM> to inner ring <NUM> configuration may carry the rotor stack preload that may be upward of <NUM>,<NUM> pounds (<NUM>,<NUM> kilograms) through the high load capability material of the rotor rim <NUM> to inner ring <NUM> interface, while permitting the usage of a high temperature resistant, yet lower load capability materials in the blades <NUM> and the sealing outer rings <NUM> of the spacers <NUM> that are exposed to the high temperature core gas flowpath C. The axial rotor stack load path may facilitate the use of a disk specific alloy to carry the stack load and allows for the high temperature resistant material to seal the rotor from the core gas flow path. That is, the inner diameter loading and outer diameter sealing permits a segmented airfoil and seal platform design that facilitates relatively inexpensive manufacture and high temperature capability.

Referring to <FIG> and <FIG>, the HPC rotor assembly <NUM> further includes a rear shell <NUM> attached to and projecting in a downstream direction from the final, downstream, rotor <NUM> of the rotor assembly <NUM>. The rear shell <NUM> is substantially concentric to axis A, generally conical in shape, and converges in a downstream direction. The shell <NUM> includes radially inner and outer walls <NUM>, <NUM> with a passage <NUM> defined there-between. The passage <NUM> is generally part of the secondary air flowpath S, is generally annular, and is in direct fluid communication with the channels <NUM> in the adjacent, upstream, rotor <NUM>. The shell <NUM> is constructed and arranged to rotate in unison with the rotors <NUM> and spacers <NUM>.

Referring to <FIG>, the rotor assembly <NUM> may further include a structure <NUM> that may be generally located radially inward of, and axially aligned to, the hubs <NUM> of the rotor disks <NUM> and the shell <NUM>. The structure <NUM>, or portion thereof: may generally rotate with the rotors <NUM>, the spacers <NUM> and the shell <NUM>; may extend axially; and, may be substantially concentric to axis A. With the rotor assembly <NUM> being part of the HPC <NUM>, the structure <NUM> may generally include the low spool <NUM> and a tie-shaft or bore tube <NUM> that may be part of the high spool <NUM>. Another example of a tie-shaft <NUM> is disclosed in <CIT>, assigned to Pratt & Whitney Canada Corporation of the United Technologies Corporation in Hartford, Connecticut.

The rotor assembly <NUM> further includes a plurality of forward rotor bores <NUM> and a rearward rotor bore <NUM>. Each rotor bore <NUM> is defined axially between respective, adjacent, rotor disks <NUM> and radially between the inner ring <NUM> of the spacer <NUM> and the structure <NUM>. Rotor bore <NUM> may be located immediately rearward of the plurality of forward rotor bores <NUM> and may generally be defined axially between the rearward-most rotor disk <NUM> and the inner wall <NUM> of the shell <NUM>. Selected forward rotor bores <NUM> and the rearward rotor bore <NUM> may generally be in fluid communication with one-another through a plurality of supply conduits <NUM> that may be defined, in-part, by the tie-shaft <NUM> of the structure <NUM>. More specifically, the supply conduits <NUM> may extend through the hubs <NUM> of a selected number of rotor disks <NUM> and generally adjacent to the tie-shaft <NUM>. Each supply conduit <NUM> at each respective rotor disk <NUM> may further be a plurality of circumferentially spaced conduits with the number of conduits and size generally dictated by the thermal gradient of the respective rotor disk.

The rotor assembly <NUM> may further include an axially extending discharge conduit <NUM> that may be defined, at least in-part, radially between the tie-shaft <NUM> and the low spool <NUM>. The discharge conduit <NUM> may include at least one inlet <NUM> extending radially through the tie-shaft <NUM>, and being in fluid communication between the discharge conduit <NUM> and at least the forward-most rotor bore <NUM> that is predetermined to require a warm airflow due to an excessive temperature gradient. It is further contemplated and understood that each rotor bore <NUM> may communicate with a respective inlet <NUM>. The distribution of inlets <NUM> and individual flow cross section areas of the inlets may generally be dependent upon the temperature gradients of the adjacent disks <NUM> of each respective rotor bore <NUM>.

In operation, the secondary airflow first flows through the channels <NUM>, <NUM> thus cooling and picking up heat from the adjacent rims <NUM> of the disks <NUM>, roots <NUM> of the blades <NUM>, and outer and inner rings <NUM>, <NUM> of the spacers <NUM>. From the rearward-most channels <NUM>, the secondary airflow enters the passage <NUM> in the shell <NUM> and may pick up further heat from the surrounding walls <NUM>, <NUM>. In a warmed or heated state, the secondary airflow may then enter the rearward-most rotor bore <NUM>, flow in an axial forward direction, through the conduit <NUM>, and into the upstream rotor bore <NUM>. This flow may continue in the forward direction and into successive rotor bores <NUM> through successive conduits <NUM> as dictated by the thermal needs of the rotor assembly <NUM>. As the heated secondary airflow enters the conduits <NUM> and bores <NUM>, the air heats the rotor hubs <NUM> and may heat a portion of the webs <NUM> of the rotor disks <NUM>, thereby reducing temperature gradients that reduces thermal stress. A reduction in thermal stress increases rim life, facilitates a reduction in bore size and rotor weight, and may realize achievement of a full life HPC rotor.

From at least the forward-most rotor bore <NUM>, the secondary airflow is discharged from the rotor bore, through the inlet <NUM>, and into the discharge conduit <NUM>. The discharge conduit <NUM> may generally channel the airflow (as one, non-limiting, example) to the LPT <NUM>. This expelled airflow may generally be dumped into a pre-specified stage of the LPT <NUM> to achieve the desired pressure differentials through the entire secondary air flowpath S, wherein the channels <NUM>, <NUM>, the passage <NUM>, the supply conduits <NUM> and the discharge conduit <NUM> may be considered part or all of the secondary flowpath S.

Each inner ring <NUM> may further have at least one aperture <NUM> for flowing a portion of secondary airflow from at least one of the channels <NUM> of the spacers <NUM> and into the respective bores <NUM> for ventilation. This ventilating airflow (see arrow <NUM> in <FIG>) may be minimal and may amount to one to two percent of the airflow flowing through the passage <NUM>. The ventilating airflow <NUM> may be discharged from the bores <NUM> via the discharge conduit <NUM>.

TMF may occur in traditional rotors because of large temperature differences between the full hoop ring (flowpath side) and the inner full hoop ring (i.e. hubs <NUM>). These rings may tend to grow or thermally expand in different amounts due to the different temperatures that each are exposed to; thus, generally causing a fight or interference between the rings. The spoked rotor configuration of the present disclosure may substantially eliminate the TMF on the hot flowpath side and the spokes <NUM> because the hot ring is divided into segments thereby eliminating any interference. In more traditional engine designs, another location for TMF would be the inner rings <NUM>, <NUM>; however, through application the present disclosure, the rings <NUM>, <NUM> are now cooler than the hot ring due to the cooling of the present disclosure making the temperature delta between the rings and hubs <NUM> much less. This essentially eliminates or greatly reduces the TMF.

Referring to <FIG>, a third embodiment of a rotor assembly is illustrated wherein like elements to the first embodiment have the same identifying numerals except with the addition of a "double prime" suffix. The rotor assembly <NUM>" of the third embodiment may not heat the rotor hubs <NUM>" and instead, the secondary airflow may cool the shell <NUM>" and may then flow directly to an LPT <NUM>". That is, secondary airflow is routed through a passage <NUM>" in the shell <NUM>", then directly into an inlet <NUM>" and into a discharge conduit <NUM>". The inlet <NUM>" may be in and communicate through a high spool <NUM>" and the discharge conduit <NUM>" may be defined radially between the low and high spools <NUM>", <NUM>". Although not illustrated, it is further contemplated and understood that the secondary airflow, once serving the HPC, may be routed to the HPT or elsewhere within the engine.

It is understood that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," and the like are with reference to the normal operational attitude and should not be considered otherwise limiting. It is also understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will also benefit. Although particular step sequences may be shown, described, and claimed, it is understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.

Claim 1:
A rotor assembly of a gas turbine engine comprising:
a HPC rotor (<NUM>) for rotation about an engine axis (A) and including a rotor disk (<NUM>) and a plurality of blades (<NUM>) each including a platform (<NUM>) attached to the rotor disk (<NUM>) with a first channel (<NUM>) defined radially between the platforms (<NUM>) and the rotor disk (<NUM>); and
a rear shell (<NUM>) attached to and projecting rearward from the final, downstream, HPC rotor (<NUM>) of the rotor assembly (<NUM>), the rear shell (<NUM>) being concentric to the engine axis (A), conical in shape, converging in a downstream direction and characterised in that the rear shell (<NUM>) includes a radially inner wall (<NUM>) and a radially outer wall (<NUM>) with a passage (<NUM>) defined between the inner (<NUM>) and outer (<NUM>) walls in fluid communication with the first channel (<NUM>), wherein the shell (<NUM>) is constructed and arranged to rotate with the adjacent rotor (<NUM>); and further characterised by
a ring-shaped spacer (<NUM>) disposed adjacent to and upstream of the HPC rotor (<NUM>) with a second channel (<NUM>) in the spacer (<NUM>) being in fluid communication with the first channel (<NUM>);
wherein the first channel (<NUM>), the second channel and the passage (<NUM>) are part of a secondary cooling air flowpath (<NUM>).