[특허]Secondary flow arrangement for slotted rotor

United Technologies Corporation

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특허 상세정보

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Secondary flow arrangement for slotted rotor

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F01D-005/06 F01D-011/00
미국특허분류(USC)	416/096R; 416/198A
출원번호	US-0459474 (2012-04-30)
등록번호	US-8961132 (2015-02-24)
발명자 / 주소	Suciu, Gabriel L. Dye, Christopher M. Ackermann, William K. Muron, Stephen P. Alvanos, Ioannis Merry, Brian D. Salve, Arthur M. Norris, James W.
출원인 / 주소	United Technologies Corporation
대리인 / 주소	Carlson, Gaskey & Olds, PC
인용정보	피인용 횟수 : 4 인용 특허 : 14

초록

A rotor for a gas turbine engine includes a plurality of blades which extend from a rotor disk and at least one spacer adjacent to the plurality of blades. A flow passage is defined between the rotor disk and the blades and spacer. A plurality of inlets are formed within the spacer to pump air into the flow passage.

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1. A rotor for a gas turbine engine comprising: a rotor disk defined along an axis of rotation, the rotor disc including a rotor outer peripheral surface;a plurality of blades which extend from the rotor disk, wherein the blades are supported on platforms that have a blade inner surface that faces the rotor outer peripheral surface;at least one spacer positioned adjacent the plurality of blades to define a flow passage between the rotor disk and the blades and spacer, wherein the spacers include a spacer outer peripheral surface and a spacer inner peripheral surface that faces the rotor outer peripheral surface, and wherein the flow passage is defined between the rotor outer peripheral surface and the blade and spacer inner surfaces; anda plurality of inlets formed within the at least one spacer to pump air into the flow passage, wherein the inlets extend through the at least one spacer from at least one of the spacer outer and inner peripheral surfaces to an end face of the at least one spacer such that air flows in a generally axial direction in the flow passage from the at least one spacer toward the rotor disk. 2. The rotor as recited in claim 1, wherein the plurality of blades includes at least a first set of blades and a second set of blades spaced axially aft of the first set of blades, and wherein the at least one spacer comprises at least a first spacer positioned upstream of the first set of blades and a second spacer positioned between the first and second sets of blades, and wherein the plurality of inlets is formed within the first spacer. 3. The rotor as recited in claim 1, wherein the flow passage includes an outlet configured to direct cooling airflow into a turbine section. 4. The rotor as recited in claim 3, wherein the turbine section comprises a high pressure turbine. 5. The rotor as recited in claim 4, wherein the plurality of blades comprise compressor blades. 6. The rotor as recited in claim 1, wherein the plurality of blades are integrally formed as one piece with the rotor disk. 7. The rotor as recited in claim 1, wherein the plurality of blades are high pressure compressor blades. 8. The rotor as recited in claim 1, wherein the at least one spacer is integrally formed as one piece with the rotor disk. 9. The rotor as recited in claim 1, wherein the plurality of blades comprise compressor blades, and wherein the at least one spacer comprises an inlet spacer positioned upstream of all stages of an associated compressor. 10. A rotor for a gas turbine engine comprising: a rotor disk defined along an axis of rotation;a plurality of blades which extend from the rotor disk, wherein the plurality of blades are formed from a first material and the rotor disk is formed from a second material that is different from the first material, and wherein the plurality of blades are bonded to the rotor disk at an interface;at least one spacer positioned adjacent the plurality of blades to define a flow passage between the rotor disk and the blades and spacer; anda plurality of inlets formed within the at least one spacer to pump air into the flow passage. 11. A rotor for a gas turbine engine comprising: a rotor disk defined along an axis of rotation;a plurality of blades which extend from the rotor disk;at least one spacer positioned adjacent the plurality of blades to define a flow passage between the rotor disk and the blades and spacer, wherein the at least one spacer is formed from a first material and an associated rotor ring is formed from a second material that is different from the first material, and wherein the at least one spacer is bonded to the rotor ring at an interface; anda plurality of inlets formed within the at least one spacer to pump air into the flow passage. 12. A rotor for a gas turbine engine comprising: a rotor disk defined along an axis of rotation;a plurality of blades which extend from the rotor disk;at least one spacer positioned adjacent the plurality of blades to define a flow passage between the rotor disk and the blades and spacer, wherein the flow passage is sealed by axial seals extending axially along the blades and tangential seals extending circumferentially about the axis of rotation between the at least one spacer and the plurality of blades; anda plurality of inlets formed within the at least one spacer to pump air into the flow passage. 13. A gas turbine engine comprising: a compressor section including a rotor disk rotatable about an axis, a plurality of blades comprising at least a first set of blades and a second set of blades spaced axially aft of the first set of blades, and a plurality of spacers comprising at least a first spacer positioned upstream of the first set of blades and a second spacer positioned between the first and second sets of blades;a flow passage defined between an outer peripheral surface of the rotor disk and inner surfaces of the blades and the spacers;a plurality of inlets formed within the first spacer to pump air into the flow passage, wherein the inlets extend through the first spacer from at least one of outer and inner peripheral surfaces of the first spacer to an end face of the first spacer such that air flows in a generally axial direction in the flow passage from the first spacer toward the first set of blades; anda turbine section configured to receive air pumped out of the flow passage. 14. The gas turbine engine as recited in claim 13, wherein the compressor section comprises a high pressure compressor and the turbine section comprises a high pressure turbine. 15. The gas turbine engine as recited in claim 13, wherein the plurality of inlets comprise discrete openings that are circumferentially spaced apart from each other about the axis. 16. The gas turbine engine as recited in claim 13, wherein the plurality of blades includes a third set of blades positioned axially aft of the second set of blades and wherein the plurality of spacers includes a third spacer positioned between the second and third sets of blades, and wherein the flow passage extends in a generally axial direction from a location starting at the inlets at the first spacer and terminating at an outlet into the turbine section positioned aft of the third set of blades. 17. The gas turbine engine as recited in claim 16, including a turbine casing section positioned aft of the third set of blades to define a turbine cavity that receives air exiting the flow passage. 18. The gas turbine engine as recited in claim 13, wherein the blades are formed from a first material and the rotor disk is formed from a second material that is different from the first material, and wherein the blades are bonded to the rotor disk at an interface. 19. The gas turbine engine as recited in claim 13, wherein at least one of the first and second spacers comprise a plurality of seals extending outwardly from a rotor ring, and wherein the seals are formed from a first material and the rotor ring is formed from a second material that is different from the first material, and wherein the seals are bonded to the rotor ring at an interface. 20. The gas turbine engine as recited in claim 13, wherein the first spacer comprises an inlet spacer that upstream of all compressor blades. 21. A gas turbine engine comprising: a compressor section including a rotor disk rotatable about an axis, a plurality of blades comprising at least a first set of blades and a second set of Hades spaced axially aft of the first set of blades, and a plurality of spacers comprising at least a first spacer positioned upstream of the first set of blades and a second spacer positioned between the first and second sets of blades;a flow passage defined between an outer peripheral surface of the rotor disk and inner surfaces of the blades and the spacers;a plurality of inlets formed within the first spacer to pump air into the flow passage; and a turbine section configured to receive air pumped out of the flow passage;a turbine section configured to receive air pumped out of the flow passage anda plurality of axial seals and tangential seals that cooperate to seal the flow passage. 22. The gas turbine engine as recited in claim 21, wherein the axial seals extend along a length of platform edges for adjacent blades. 23. The gas turbine engine as recited in claim 21, wherein the tangential seals extend circumferentially about the axis between fore and aft edges of the spacers and an associated fore and aft edge of platforms for the first and second sets of blades.

이 특허에 인용된 특허 (14)

이 특허를 인용한 특허 피인용횟수: 4

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특허 상세정보

Secondary flow arrangement for slotted rotor

초록

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이 특허에 인용된 특허 (14)

이 특허를 인용한 특허 피인용횟수: 4

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특허 상세정보

Secondary flow arrangement for slotted rotor

초록

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이 특허에 인용된 특허 (14)

이 특허를 인용한 특허 피인용횟수: 4

IPC 상위 출원인

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상세정보조회

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IPC 상세정보(version : 201201)

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