Turbine shroud with axially separated pressure compartments

A turbine shroud for a gas turbine engine includes an annular metallic carrier and a blade track. One or more cavities are formed between the carrier and blade track to provide cooling air to the blade track.

FIELD OF THE DISCLOSURE

The present disclosure relates generally to gas turbine engines, and more specifically to turbine shrouds used in gas turbine engines.

BACKGROUND

Compressors and turbines typically include alternating stages of static vane assemblies and rotating wheel assemblies. The rotating wheel assemblies include disks carrying blades around their outer edges. When the rotating wheel assemblies turn, tips of the blades move along blade tracks included in static shrouds that are arranged around the rotating wheel assemblies. Such static shrouds may be coupled to an engine case that surrounds the compressor, the combustor, and the turbine.

Some shrouds positioned in the turbine may be exposed to high temperatures from products of the combustion reaction in the combustor. Such shrouds sometimes include components made from materials that have low torsional rigidity. Due to the low torsional rigidity, the components of some turbine shrouds experience high stress levels when differential pressures are applied along their length. For example, exposing such components to differential pressures may reduce their useful lives in service.

SUMMARY

According to one aspect of the present disclosure, a turbine shroud for use in a gas turbine engine includes a carrier, a blade track, and at least one bulkhead. The carrier may comprise metallic materials and be adapted to be coupled to a turbine case. The blade track may comprise ceramic-matrix composite materials and be coupled to the carrier. The blade track may extend at least a portion of the way around a central axis so that the blade track is adapted to block hot gasses from passing over a turbine wheel surrounded by the turbine shroud. The at least one bulkhead may radially interconnect the carrier and the blade track to divide a space between the carrier and the runner into at least a first cooling-air cavity and a second cooling-air cavity. The at least one bulkhead may be formed to include a plurality of pressure-control holes sized to cause the first cooling-air cavity to have a first pressure and the second cooling-air cavity to have a second pressure, lower than the first pressure.

In illustrative embodiments, each of the pressure-control holes may include a first end positioned to pass air from the first cooling-air cavity into the pressure-control hole and a second end positioned to pass air from the pressure-control hole into the second cooling-air cavity, and wherein the first end is positioned radially outward of the second end.

In illustrative embodiments, the blade track may include a plurality of blade track segments positioned to surround the turbine wheel. Each blade track segment may include a runner that extends a portion of the way around the central axis so that the runner is adapted to block hot gasses from passing over the turbine wheel surrounded by the turbine shroud, a forward hanger that extends radially outward from the runner, and an aft hanger spaced axially from the forward hanger that extends radially outward from the runner.

In illustrative embodiments, the at least one bulkhead may comprise a first bulkhead located axially between the forward hanger and the aft hanger and a second bulkhead located axially between the first bulkhead and the aft hanger. The second bulkhead may radially interconnect the carrier and the runner to divide a space between the carrier and the runner into the second cooling-air cavity and a third cooling-air cavity.

In illustrative embodiments, the first bulkhead may be formed to include a first portion of the plurality of pressure-control holes and the second bulkhead may be formed to include a second portion of the plurality of pressure-control holes to cause the second cooling-air cavity to have the second pressure and the third cooling-air cavity to have a third pressure, lower than the second pressure.

In illustrative embodiments, each of the first and second bulkheads may include a divider that extends radially inward from the carrier and be formed to include the pressure-control holes and a seal member that interconnects the divider and the runner of the blade track segment.

In illustrative embodiments, each of the pressure-control holes in the first portion may include a first end positioned to pass air from the first cooling-air cavity into the pressure-control hole and a second end positioned to pass air from the pressure-control hole into the second cooling-air cavity. The first end may be positioned radially outward of the second end.

In illustrative embodiments, each of the pressure-control holes in the second portion may include a third end positioned to pass air from the second cooling-air cavity into the pressure-control hole and a fourth end positioned to pass air from the pressure-control hole into the third cooling-air cavity. The third end may be positioned radially outward of the fourth end.

In illustrative embodiments, the carrier may include a diffuser and a carrier hanger coupled to the diffuser. The carrier hanger may be positioned to support the diffuser. The diffuser may be positioned to provide cooling air to the first cooling-air cavity.

In illustrative embodiments, the diffuser may define a cooling air diffuser plenum in fluid communication with the first cooling-air cavity by a cavity supply aperture.

In illustrative embodiments, the diffuser may be formed to include at least one diffuser inlet aperture positioned to pass air into the cooling air diffuser plenum.

In illustrative embodiments, the at least one bulkhead may be coupled to a radially inner portion of the cooling air diffuser plenum.

In illustrative embodiments, the carrier may further include a forward bracket coupled to the diffuser and positioned to support the forward hanger of the blade track to couple the blade track with the carrier.

In illustrative embodiments, the turbine shroud may further include a seal member positioned between the forward bracket and the blade track.

In illustrative embodiments, the turbine shroud may further include a retainer coupled to the carrier. The retainer may include a diffuser mount and a retainer hanger coupled to the diffuser mount. The retainer hanger may be positioned to support the diffuser mount. The diffuser mount may be positioned to engage the diffuser of the carrier.

In illustrative embodiments, the retainer may further include an aft bracket coupled to the diffuser mount and positioned to support the aft hanger of the blade track to couple the blade track with the retainer.

In illustrative embodiments, the turbine shroud may further include a seal member positioned between the aft bracket and the blade track.

According to another aspect of the present disclosure, a turbine shroud for use in a gas turbine engine includes a carrier, a retainer, a blade track, and at least one bulkhead. The carrier may comprise metallic materials and be adapted to be coupled to a turbine case. The retainer may comprise metallic materials and be adapted to be coupled to the turbine case. The retainer may be coupled to the carrier. The blade track may comprise ceramic-matrix composite materials and extend around a central axis so that the blade track is adapted to block hot gasses from passing over a turbine wheel. The blade track may be positioned to engage with the carrier and retainer to maintain alignment of the blade track with the turbine wheel.

In illustrative embodiments, the at least one bulkhead may radially interconnect the carrier and the blade track to divide a space between the carrier and the runner into at least a first cooling-air cavity and a second cooling-air cavity. The at least one bulkhead may be configured to provide means for creating a pressure differential between subsequent cooling-air cavities such that the each subsequent cooling-air cavity has a pressure lower than the upstream cooling-air cavity.

In illustrative embodiments, the at least one bulkhead may include a divider extending radially inward from the carrier and a seal member positioned between the divider and the blade track.

In illustrative embodiments, the dividers may be formed to include a plurality of pressure-control holes and each of the pressure-control holes extends between an axially forward first end and an axially aft second end. The axially forward first end may be positioned radially outward of the axially aft second end.

In illustrative embodiments, the blade track may be formed to include grooves to allow cooling air located in the first cooling-air cavity to pass between the blade track and the seal member into the second cooling-air cavity.

In illustrative embodiments, the seal member may be configured to allow cooling air located in the first cooling-air cavity to pass between the blade track and the seal member into the second cooling-air cavity.

In illustrative embodiments, the seal member may be formed to include one or more apertures extending between the first cooling-air cavity and the second cooling-air cavity.

According to another aspect of the present disclosure, a method of assembling a turbine shroud for use in a gas turbine engine may include engaging a forward hanger of a blade track with a forward bracket of a carrier, engaging a runner of the blade track with a bulkhead of the carrier, the runner being coupled to the forward hanger, and engaging an aft bracket of a retainer with an aft hanger coupled to the runner of the blade track. The bulkhead may divide a spaced defined by the carrier, blade track, and retainer into an axially forward first cavity and an axially aft second cavity. The first and second cavities may be in fluid communication with one another by a plurality of pressure-control holes formed through the bulkhead and sized to cause the first cavity to have a first pressure and the second cavity to have a second pressure, lower than the first pressure.

DETAILED DESCRIPTION OF THE DRAWINGS

FIG. 1is an illustrative aerospace gas turbine engine10cut-away to show that the engine10includes a fan assembly12, a compressor14, a combustor16, and a turbine18. The fan assembly12pushes air through the engine10to propel an aircraft. The compressor14compresses and delivers air to the combustor16. The combustor16mixes fuel with the compressed air received from the compressor14and ignites the fuel. The hot, high pressure products of the combustion reaction in the combustor16are directed into the turbine18and the turbine18extracts work to drive the compressor14and the fan assembly12.

The turbine18illustratively includes at least one turbine wheel assembly30and a turbine shroud20positioned to surround the turbine wheel assembly30as shown inFIG. 2. The turbine wheel assembly30includes a plurality of blades32coupled to a rotor disk for rotation therewith. The hot, high pressure combustion products from the combustor16are directed toward the blades32of the turbine wheel assemblies30. The blades32are in turn pushed by the combustion products to cause the turbine wheel assembly30to rotate; thereby, driving the rotating components of the compressor14and/or the fan assembly12.

The turbine shroud20extends around the turbine wheel assembly30to block combustion products from passing over the blades32without pushing the blades32to rotate as suggested inFIGS. 2 and 4. The turbine shroud20illustratively includes a carrier22, a blade track24, and a retainer26. The carrier22is an annular, round metallic component and is configured to support the blade track24in position adjacent to the blades32of the turbine wheel assembly30. The illustrative blade track24is generally concentric with and nested into the carrier22along a rotational axis11of the engine10. The retainer26engages both the carrier22and the blade track24to position the carrier22and the blade track24relative to other static turbine components. The turbine shroud20is coupled to an outer case19of the gas turbine engine10.

In some embodiments, the turbine shroud20is made up of a number of segments that extend only part-way around the axis11which cooperate to surround the turbine wheel assembly30. In other embodiments, the turbine shroud20is annular and non-segmented to extend fully around the axis11and surround the turbine wheel assembly30. In yet other embodiments, portions of the turbine shroud20are segmented while other portions are annular and non-segmented. In one example, the blade track24is a unitary, one-piece hoop while the carrier22and retainer26are segmented and positioned to circumferentially surround the blade track24. In such an example, the blade track24is cross-keyed with the carrier22and retainer26or includes one or more hangers to maintain position of the blade track24relative to the carrier22and retainer26. In another example, the blade track24is segmented while the carrier22and retainer26are annular, one-piece structures. In such an example, the blade track24includes one or more hangers to maintain position of the blade track24relative to the carrier22and retainer26.

The pressure of the combustion products flowing from the combustor16through the turbine18decreases as work is done on the blades32by the combustion products rotating the wheel assembly30as suggested inFIGS. 2 and 3. For example, the pressure of the combustion products is higher at a leading end LE of the turbine wheel assembly30than at a trailing end TE of the turbine wheel assembly30. As such, higher pressures are exerted on a radially inner surface23of the blade track24at a leading end LE than at a trailing end TE. This pressure differential can create an undesired torque on the blade track24.

To counteract the differential pressures acting on the inner surface23of the blade track24, a plurality of differentially pressurized cavities42,44,46are illustratively formed along a radially outer surface25of the blade track24as shown inFIG. 2. The cavities42,44,46are pressurized with air flowing from the compressor14. In the illustrative embodiment, a first cavity42positioned at the leading end LE of the blade track24has a higher pressure than a second cavity44positioned substantially at a center of the blade track24. Similarly, the second cavity44has a higher pressure than a third cavity46positioned at the trailing end TE of the blade track24.

The varying pressures of the cavities42,44,46apply forces to the outer surface25of the blade track24and reduce the relative pressure differential acting on the leading end LE and trailing end TE of the blade track24as suggested inFIGS. 2 and 3. In one embodiment, the pressure within each cavity42,44,46is determined based on the number of cavities used and position of the cavity42,44,46. For example, the difference in pressure between each cavity will be smaller if four cavities are used as compared to if three cavities are used. Additionally, cavities positioned toward the leading end LE of the blade track24will have higher pressures than cavities positioned toward the trailing end TE of the blade track24.

In the illustrative embodiment, the turbine shroud20includes the carrier22, the blade track24, and the retainer26as shown inFIG. 4. The carrier22is a metallic component adapted to support the rest of the turbine shroud20and to be coupled to the rest of the engine10. The blade track24is a ceramic-matrix composite component mounted to the carrier22and adapted to withstand the high temperature gasses that act on the turbine blades30. The retainer26engages aft sides of the carrier22and the blade track24to hold the blade track24in place relative to the carrier22.

The carrier22includes a diffuser54, a case hanger52coupled to a radially outer portion of the diffuser54, and a forward bracket56coupled to a radially inner portion of the diffuser54as shown inFIG. 4. The diffuser54receives and distributes air from the compressor14to the first cavity42. The case hanger52is adapted to couple the carrier22to the outer case19. The forward bracket56has an L-shape and engages the blade track24to couple the blade track24to the carrier22.

The blade track24includes a runner62, a forward hanger64coupled to radially outer surface25, and an aft hanger66spaced apart from the forward hanger64and coupled to the radially outer surface25as shown inFIG. 4. The runner62extends circumferentially around at least a portion of the axis11and blocks gasses from passing over the blades30by defining an outer boundary of the gas path moving across the blades30. The forward hanger64has an L-shape and engages the bracket56of the carrier22to couple the blade track24to the carrier22.

The retainer26includes a diffuser mount74, a case hanger72coupled to a radially outer portion of the diffuser mount74, and an aft bracket76coupled to a radially inner portion of diffuser mount74. The diffuser mount74engages with the diffuser54of the carrier22to support an aft end of the diffuser54. The case hanger72is adapted to couple the retainer26to the outer case19. The aft bracket76of the retainer26engages with the aft hanger66of the blade track24.

In the illustrative embodiment, the blade track24is trapped axially between and supported by the forward bracket56and aft bracket76as shown inFIG. 4. A seal member156is positioned between the forward bracket56and the blade track24to seal a portion of the first cavity42. A seal member176is positioned between the aft bracket76and the blade track24to seal a portion of the third cavity46. Alternatively, the seal member156is positioned between the forward bracket56and forward hanger64and the seal member176is positioned between the aft bracket76and aft hanger66. The case hangers52,72are positioned to support the assembled turbine shroud20on the outer case19of the gas turbine engine10.

A pair of bulkheads82,86are coupled to the diffuser54and at least partially define the pressurized cavities42,44,46as shown inFIG. 4. The bulkheads82,86extend radially inward from the diffuser54and engage with the radially outer surface25of the blade track24. A forward bulkhead82is axially positioned between the forward hanger64and the aft hanger66of the blade track24when the blade track24is coupled to the carrier22. An aft bulkhead86is spaced apart from the forward bulkhead82and axially positioned between the forward bulkhead82and the aft hanger66. While two bulkheads82,86are shown, more or less bulkheads are used depending on the number of cavities being created.

The forward bulkhead82includes a divider81coupled to diffuser54and a seal member83positioned between the divider81and blade track24as shown inFIG. 4. Similarly, the aft bulkhead86includes a divider85coupled to diffuser54and a seal member87positioned between the divider85and blade track24. The divider81is formed to include at least one pressure-control hole84. Similarly, the divider85is formed to include at least one pressure-control hole88. The pressure-control holes84,88provide fluid communication between the pressurized cavities42,44,46.

The diffuser54of the carrier22defines a cooling air diffuser plenum58and is formed to include at least one diffuser inlet aperture51. The diffuser inlet aperture51is positioned to pass cooling air received from the compressor14and/or fan assembly12into the cooling air diffuser plenum58as suggested by arrow151. The diffuser54distributes the cooling air such that the cooling air diffuser plenum58has a substantially even pressure throughout.

The diffuser54is also formed to include at least one cavity supply aperture53which provides fluid communication between the cooling air diffuser plenum58and the first pressurized cavity42as shown inFIG. 4. Cooling air within the cooling air diffuser plenum58is passed into the first cavity42through the cavity supply aperture53, as suggested by arrow153. The cooling air enters the first cavity42such that the first cavity42has a substantially even pressure throughout.

Cooling air within the first cavity42is passed into the second cavity44through pressure-control hole84as suggested by arrow184inFIG. 4. The cooling air enters the second cavity44such that the second cavity44has a substantially even pressure throughout. Similarly, cooling air within the second cavity44is passed into the third cavity46through pressure-control hole88as suggested by arrow188. The cooling air enters the third cavity46such that the third cavity44has a substantially even pressure throughout. In some embodiments, seal members83,87are formed and positioned to allow cooling air to pass between the cavities42,44,46instead of, or in conjunction with, the pressure-control holes84,88. In other embodiments, grooves92,94are formed in the outer surface25of the blade track24to allow cooling air to pass between the cavities42,44,46underneath the seal members83,87instead of, or in conjunction with, the pressure-control holes84,88. In yet other embodiments, seal members83,87are formed to include pressure-control holes96,98to allow cooling air to pass between the cavities42,44,46instead of, or in conjunction with, the pressure-control holes84,88.

Pressure-control holes84,88are formed at an axially aft and radially inward slope to direct incoming cooling air toward the blade track24as shown inFIG. 4. For example, the pressure-control hole84includes a first end associated with the first cavity42and a second end associated with the second cavity44that is positioned radially inward of the first end. Similarly, the pressure-control hole88includes a first end associated with the second cavity44and a second end associated with the third cavity46that is positioned radially inward of the first end. In some embodiments, pressure-control holes84,88are positioned substantially parallel with rotational axis11of the gas turbine engine10. The cavity supply aperture53is positioned substantially perpendicular to rotational axis11and directs incoming cooling air toward the blade track24. In some embodiments, cavity supply aperture53may be angled relative to rotational axis11instead of being perpendicular with rotational axis11.

In the illustrative embodiment, the size of arrows151,153,184,188represents the amount of cooling air flowing through the associated holes or apertures. For example, more cooling air is flowing into the first cavity42, as represented by arrow153, than is flowing out of the first cavity42, as represented by arrow184. The relative differential in flow provides the first cavity42with a first pressure. Similarly, more cooling air is flowing into the second cavity44, as represented by arrow184, than is flowing out of the second cavity44, as represented by arrow188. The relative differential in flow provides the second cavity44with a second pressure which is lower than the first pressure within the first cavity42. More cooling air is also flowing into the third cavity46, as represented by arrow188, than is flowing out of the third cavity46moving axially aft through gaps along interfaces with the retainer26. The relative differential in flow provides the third cavity46with a third pressure which is lower than the second pressure within the second cavity44.

In some embodiments, the relative size of the holes and apertures within the turbine shroud20provides the differential amounts of flow through the pressurized cavities42,44,46. In some embodiments, the relative number of holes and apertures within the turbine shroud20provides the differential amounts of flow through the pressurized cavities42,44,46. In some embodiments, the relative size and number of holes and apertures within the turbine shroud20provides the differential amounts of flow through the pressurized cavities42,44,46.

In the illustrative embodiment, the blade track24is formed from ceramic containing materials as shown inFIG. 4. In some embodiments, the blade track24is formed from ceramic-matric composite (CMC) materials. In other embodiments, the blade track24may be formed from metallic or metallic alloy materials. The blade track24is positioned to surround the blades32of turbine wheel assembly30. The blades32may be formed from metallic or metallic alloy materials, or from ceramic containing materials, such as CMC for example. The carrier22and retainer26are formed from metallic or metallic alloy materials. In some embodiments, the carrier22and retainer26are formed from ceramic containing materials, such as CMC for example.