Patent Description:
Gas turbine engines typically include at least a compressor section, a combustor section, and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.

The compressor section and the turbine section may each include alternating rows of rotor and stator assemblies. The rotor assemblies carry rotating blades that create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine. The stator assemblies include stationary structures called stators that direct the core airflow to the blades to either add or extract energy.

Some rotor assemblies employ cover plates that retain the blades to disks of the rotor assemblies and seal between adjacent sets of blades and stators. A limited amount of space may be available for mounting the cover plates. These space limitations may complicate the installation and removal of the cover plates.

<CIT>, which is prior art under Art. <NUM>(<NUM>) EPC, discloses prior art bayoneted anti-rotation turbine seals.

<CIT>, which is prior art under Art. <NUM>(<NUM>) EPC, discloses a prior art double snapped cover plate for a rotor disk.

<CIT> discloses a prior art turbine disk cooling system.

<CIT> discloses a prior art gas turbine engine with integrated abradable seal and mount plate.

<CIT> discloses a prior art rotor blade retaining apparatus. <CIT> discloses a rotor disk with two cover plates positioned respectively on the front and the rear of the disk.

According to the present invention, there is provided a rotor assembly as set forth in claim <NUM>.

In an embodiment, an outer face of the first tab and the second tab is offset from an outer face of the body.

In a further non-limiting embodiment of any of the foregoing embodiments, a seal land extends from the body.

In a further non-limiting embodiment of any of the foregoing embodiments, the seal land includes at least one seal that seals against a static structure adjacent to the body.

In a further non-limiting embodiment of any of the foregoing embodiments, the slot extends radially outward from a base of the first tab and the second tab.

In a further non-limiting embodiment of any of the foregoing embodiments, each of the first tab and the second tab include a gradually decreasing thickness in a direction toward a tip of each of the first tab and the second tab.

In a further non-limiting embodiment of any of the foregoing embodiments, an inner surface of the first tab and the second tab extends at the angle.

This disclosure relates to rotor assembly cover plates that retain blades to disks of the rotor assemblies and seal between adjacent sets of blades and stators. As detailed herein, among other features, the cover plates described in this disclosure are radially and circumferentially retained without reducing the effectiveness of the cover plate bores. The exemplary cover plates may be installed and/or removed from relatively tight spaces of a rotor assembly. In other embodiments, the cover plates described in this disclosure may include one or more bumpers that limit deflection of portions of the cover plate toward a rotor disk rim, thereby reducing stresses and increasing part life.

The exemplary gas turbine engine <NUM> is a two-spool turbofan engine that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. Alternative engines might include an augmenter section (not shown) among other systems for features. The fan section <NUM> drives air along a bypass flow path B, while the compressor section <NUM> drives air along a core flow path C for compression and communication into the combustor section <NUM>. The hot combustion gases generated in the combustor section <NUM> are expanded through the turbine section <NUM>. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures.

The gas turbine engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine centerline longitudinal axis A. The low speed spool <NUM> and the high speed spool <NUM> may be mounted relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that other bearing systems <NUM> may alternatively or additionally be provided.

The inner shaft <NUM> can be connected to the fan <NUM> through a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The high speed spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure compressor <NUM> and a high pressure turbine <NUM>. In this embodiment, the inner shaft <NUM> and the outer shaft <NUM> are supported at various axial locations by bearing systems <NUM> positioned within the engine static structure <NUM>.

A combustor <NUM> is arranged between the high pressure compressor <NUM> and the high pressure turbine <NUM>. A mid-turbine frame <NUM> may be arranged generally between the high pressure turbine <NUM> and the low pressure turbine <NUM>. The mid-turbine frame <NUM> can support one or more bearing systems <NUM> of the turbine section <NUM>. The mid-turbine frame <NUM> may include one or more airfoils <NUM> that extend within the core flow path C.

The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor <NUM> and the high pressure compressor <NUM>, is mixed with fuel and burned in the combustor <NUM>, and is then expanded over the high pressure turbine <NUM> and the low pressure turbine <NUM>. The high pressure turbine <NUM> and the low pressure turbine <NUM> rotationally drive the respective high speed spool <NUM> and the low speed spool <NUM> in response to the expansion.

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

In this embodiment of the exemplary gas turbine engine <NUM>, a significant amount of thrust is provided by the bypass flow path 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> (<NUM>,<NUM> feet). This flight condition, with the gas turbine engine <NUM> at its best fuel consumption, is also known as bucket cruise 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 embodiment of the example gas turbine engine <NUM> is less than <NUM>. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram°K)/(<NUM>]<NUM> (or [(Tram°R)/(<NUM> °R)]<NUM>). The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine <NUM> is less than about <NUM>/s (<NUM> fps).

The compressor section <NUM> and the turbine section <NUM> may include alternating rows of rotor assemblies and stator assemblies (shown schematically) that carry airfoils. For example, rotor assemblies carry a plurality of rotating blades <NUM>, while stator assemblies carry stationary stators <NUM> (or vanes) that extend into the core flow path C to influence the hot combustion gases. The blades <NUM> create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine <NUM> along the core flow path C. The stators <NUM> direct the core airflow to the blades <NUM> to either add or extract energy.

<FIG> illustrates a portion <NUM> of a gas turbine engine, such as the gas turbine engine <NUM> of <FIG>. In this embodiment, the portion <NUM> is part of a turbine section <NUM> of the gas turbine engine <NUM>. However, this disclosure is not limited to the turbine section <NUM>, and the various features of this disclosure could extend to other sections of the gas turbine engine <NUM>, including but not limited to the compressor section <NUM>. The portion <NUM> is not necessarily drawn to scale and has been enlarged to better illustrate its various features and components.

In one embodiment, the portion <NUM> includes a rotating rotor assembly <NUM> and a stationary stator assembly <NUM>. Of course, additional stages of rotor and stator assemblies than are shown may be employed within the portion <NUM>. The rotor assemblies <NUM> carry blades <NUM>, while the stator assemblies <NUM> carry stators <NUM>. Each row of blades <NUM> and stators <NUM> is circumferentially disposed about the engine centerline longitudinal axis A.

The blades <NUM> of the rotor assembly <NUM> are circumferentially disposed about a rotor disk <NUM> that rotates about the engine centerline longitudinal axis A to move the blades <NUM>. The rotor disk <NUM> includes a rim <NUM>, a bore <NUM> and a web <NUM> that extends between the rim <NUM> and the bore <NUM>. The blades <NUM> extend outwardly from the rim <NUM> of the rotor disk <NUM> toward an engine casing <NUM>.

A cover plate <NUM> (shown schematically in <FIG>) may be positioned on one or both of a first axial side <NUM> (i.e., an upstream side) and a second axial side <NUM> (i.e., a downstream side) of the rotor disk <NUM>. The cover plates <NUM> partially extend along a root <NUM> of each blade <NUM>, in one embodiment. The cover plates <NUM> axially retain the blades <NUM> to the rotor disk <NUM>, such as within slots (not shown) formed in the rim <NUM> of the rotor disk <NUM>.

In addition to providing blade retention, the cover plates <NUM> may form an annular seal between the core flow path C and a secondary cooling flow path F that is radially inward from the core flow path C. The secondary cooling flow path F communicates cooling fluid to cool portions of the rotor assembly <NUM>, including but not limited to the rim <NUM>, the bore <NUM>, and the web <NUM> of the rotor disk <NUM>.

<FIG> illustrates one exemplary cover plate <NUM> that may be incorporated into a rotor assembly <NUM>. The cover plate <NUM> includes a body <NUM> that radially extends between a radially outer portion <NUM> and a bore <NUM>. In one embodiment, the body <NUM> is an annular structure (i.e., a full ring hoop). The bore <NUM> is generally opposite the radially outer portion <NUM> (i.e., at a radially inner section of the body <NUM>). The bore <NUM> may include a thickness T that is a greater thickness than the remaining portions of the body <NUM> of the cover plate <NUM>.

The body <NUM> axially extends between an inner face <NUM> (which faces toward the blade <NUM> and the rotor assembly <NUM>) and an outer face <NUM> (which faces away from the rotor assembly <NUM>). Cavities <NUM> may extend between the inner face <NUM> of the cover plate <NUM> and a root <NUM> of a blade <NUM> or a rotor disk <NUM> of the rotor assembly <NUM>.

The cover plate <NUM> includes one or more radial retention features <NUM> that limit radial deflection between the cover plate <NUM> and the rotor disk <NUM> of the rotor assembly <NUM>. The cover plate <NUM> could include additional retention features. The radial retention feature <NUM> extends from the inner face <NUM> and engages inner diameter surface <NUM> of the rotor disk <NUM> to provide radial interference between the cover plate <NUM> and the rotor disk <NUM>.

The cover plate <NUM> may additionally include a seal land <NUM> that axially extends from the outer face <NUM> of the body <NUM>. The seal land <NUM> includes one or more seals <NUM>, such as knife edge seals, that seal relative to a static structure <NUM>. In one embodiment, the static structure <NUM> is part of an adjacent stator assembly (see, for example, the stator assembly <NUM> of <FIG>). The seal land <NUM> is radially outward of the radial retention feature <NUM>, in one embodiment.

Referring to <FIG>, a plurality of tabs <NUM> are circumferentially spaced about the bore <NUM> of the cover plate <NUM>. For example, the bore <NUM> may include a first tab 100A, a second tab 100B circumferentially spaced from the first tab 100A, and a slot <NUM> defined between the tabs 100A, 100B (best shown in <FIG>). In one embodiment, the cover plate <NUM> includes twenty-two slots <NUM>. However, the number of tabs and slots of the cover plate are not intended to limit this disclosure and may vary depending upon the size and configuration of the rotor assembly <NUM>, among other factors.

The tabs <NUM> and the slots <NUM> extend at an angle α relative to a slot axis <NUM> that extends through the bore <NUM> (see <FIG>). In one embodiment, the angle α extends between the slot axis <NUM> and a radial axis <NUM> of the bore <NUM>. An inner surface <NUM> of the tabs <NUM> may also be angled. The angle α could be any angle. The tabs <NUM> and slots <NUM> are angled so that the cover plate <NUM>, and in particular the outer face <NUM> of the body <NUM>, can clear disk tabs <NUM> that extend from the rotor disk <NUM> during installation and removal of the cover plate <NUM> relative to the rotor assembly <NUM>. In one embodiment, an outer face <NUM> of the tabs <NUM> is offset from the outer face <NUM> of the body <NUM>.

The tabs <NUM> may include a gradually decreasing thickness T2 in a direction toward a tip <NUM> of each tab <NUM>. The gradually decreasing thickness T2 is established, at least in part, by the angled inner surface <NUM> of the tabs <NUM>.

In one embodiment, the slots <NUM> extend radially into the bore <NUM> of the cover plate <NUM> (see <FIG>). A portion of the slot <NUM> may extend radially outward of the tabs <NUM>.

<FIG> illustrates an exemplary mounting scheme of the cover plate <NUM> relative to a first rotor disk 56A and a second rotor disk 56B. The angled tabs <NUM> provide clearance for bayonetting the cover plate <NUM> onto the rotor disk 56A over the disk tabs <NUM>. The tabs <NUM> of the cover plate <NUM> engage the disk tabs <NUM> to axially retain the cover plate <NUM>.

Disk tabs <NUM> of the second rotor disk 56B extend through the first rotor disk 56A and into the slots <NUM> defined between the tabs <NUM> of the cover plate <NUM>. In one embodiment, the disk tabs <NUM> extend through slots <NUM> between the disk tabs <NUM> of the first rotor disk 56A. Extension of the disk tabs <NUM> into the slots <NUM> circumferentially retains the cover plate <NUM> relative to the rotor assembly <NUM>. In other words, the cover plate <NUM> is prevented from rotating relative to the rotor assembly <NUM> during engine operation.

<FIG> schematically illustrates the use of a tool <NUM> for installing a cover plate <NUM> to a rotor assembly <NUM>. The angled slots <NUM> of the cover plate <NUM> allow the tool <NUM> to be inserted from the side of the outer surface <NUM> of the cover plate <NUM> without blocking the disk tabs <NUM>. The tool <NUM> is insertable between the tabs <NUM> and can be used to rotate the cover plate <NUM> during installation or removal.

<FIG> illustrates a cover plate <NUM> arrangement, outside of the scope of the present invention, that may be incorporated into a rotor assembly <NUM>. The cover plate <NUM> includes a body <NUM> having a mid-section <NUM> that extends between a radially outer portion <NUM> and a retaining leg <NUM>. In one embodiment, the body <NUM> is an annular structure (i.e., a full ring hoop).

The retaining leg <NUM> is generally opposite the radially outer portion <NUM> and extends to an inner diameter portion <NUM>. A retaining ring <NUM> may engage the inner diameter portion <NUM> of the cover plate <NUM> to axially secure the cover plate <NUM> to the rotor assembly <NUM>. In one arrangement, the retaining ring <NUM> engages both the inner diameter portion <NUM> of the cover plate <NUM> and a flange <NUM> of the rotor disk <NUM>.

The body <NUM> axially extends between an inner face <NUM> (which faces toward the blade <NUM> and the rotor disk <NUM>) and an outer face <NUM> (which faces away from the blade <NUM> and rotor disk <NUM>). Cavities <NUM> may extend between the inner face <NUM> of the cover plate <NUM> and a root <NUM> of a blade <NUM> or rotor disk <NUM> of the rotor assembly <NUM>.

The retaining leg <NUM> may include one or more radial retention features <NUM> that limit radial deflection between the cover plate <NUM> and the rotor disk <NUM>. In one arrangement, the retaining leg <NUM> extends from the body <NUM> such that the retention feature <NUM> engages an inner diameter surface <NUM> of the rotor disk <NUM> to provide radial interference between the cover plate <NUM> and the rotor disk <NUM>.

The cover plate <NUM> may additionally include a seal land <NUM> that axially extends from the outer face <NUM> of the body <NUM>. The seal land <NUM> includes one or more seals <NUM>, such as knife edge seals, that seal relative to a static structure <NUM>. In one arrangement, the static structure <NUM> is part of an adjacent stator assembly (see for example, the stator assembly <NUM> of <FIG>). The seal land <NUM> is radially outward of the retaining leg <NUM>, in one arrangement.

A fillet <NUM> connects the mid-section <NUM> of the body <NUM> to the retaining leg <NUM>. A bumper <NUM> extends from the inner face <NUM> of the body <NUM> of the cover plate <NUM> in a direction away from the outer face <NUM>. In one arrangement, the bumper <NUM> extends from the mid-section <NUM> of the body <NUM>. The bumper <NUM> may contact the rotor disk <NUM> (or root <NUM> of blade <NUM>) to limit a deflection D of the body <NUM> toward the rotor disk <NUM> (i.e., axial movement of the body <NUM> in a direction that extends from the outer face <NUM> toward the inner face <NUM>), thereby reducing stresses of the fillet <NUM>. The cover plate <NUM> could include additional bumpers than are shown in <FIG>.

In one arrangement, the bumper <NUM> is located radially outward of the fillet <NUM>. The fillet <NUM> and the bumper <NUM> may be radially offset by a distance <NUM>. The distance <NUM> may vary depending on certain design criteria, such as the size of the fillet <NUM>, among other factors. The bumper <NUM> may be positioned anywhere between the fillet <NUM> and the radially outer portion <NUM>.

In another arrangement, the bumper <NUM> is radially between the seals <NUM> of the seal land <NUM>. For example, a plane <NUM> that extends axially through a middle of the bumper <NUM> may extend radially between planes <NUM> that axially extend across radially outer surfaces <NUM> of the seals <NUM>.

Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.

Claim 1:
A rotor assembly (<NUM>) of a gas turbine engine (<NUM>), comprising:
a first rotor disk (56A) having disk tabs (<NUM>);
a second rotor disk (56B) having disk tabs (<NUM>);
at least one blade (<NUM>) carried by each of said rotor disks (56A, 56B); and
a cover plate (<NUM>) comprising:
a body (<NUM>) including at least one radial retention feature (<NUM>) which extends from an inner face (<NUM>) of said body (<NUM>);
a first tab (<NUM> A) near a bore (<NUM>) of said body (<NUM>);
a second tab (100B) circumferentially spaced from said first tab (100A); and
a slot (<NUM>) defined between said first tab (100A) and said second tab (100B), said first tab (100A), said second tab (100B) and said slot (<NUM>) extending at an angle (α) relative to a slot axis (<NUM>) that extends through said bore (<NUM>), the first tab (100A) and the second tab (100B) extending away from the first rotor disk (56A);
the cover plate positioned on a first axial side of said first rotor disk blade and a second axial side of said second rotor disk blade (<NUM>), wherein a disk tab of said disk tabs (<NUM>) of the second rotor disk (56B) extends through a slot (<NUM>) between the disk tabs (<NUM>) of the first rotor disk (56A) and is received by said slot (<NUM>) defined between said first tab (100A) and said second tab (100B).