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| 국가/구분 | United States(US) Patent 등록 |
|---|---|
| 국제특허분류(IPC7판) |
|
| 출원번호 | US-0180275 (2011-07-11) |
| 등록번호 | US-8984895 (2015-03-24) |
| 발명자 / 주소 |
|
| 출원인 / 주소 |
|
| 대리인 / 주소 |
|
| 인용정보 | 피인용 횟수 : 0 인용 특허 : 326 |
A method and apparatus are disclosed for a gas turbine spool design combining metallic and ceramic components in a way that controls clearances between critical components over a range of engine operating temperatures and pressures. In a first embodiment, a ceramic turbine rotor rotates just inside
A method and apparatus are disclosed for a gas turbine spool design combining metallic and ceramic components in a way that controls clearances between critical components over a range of engine operating temperatures and pressures. In a first embodiment, a ceramic turbine rotor rotates just inside a ceramic shroud and separated by a small clearance gap. The ceramic rotor is connected to a metallic volute. In order to accommodate the differential rates of thermal expansion between the ceramic rotor and metallic volute, an active clearance control system is used to maintain the desired axial clearance between ceramic rotor and the ceramic shroud over the range of engine operating temperatures. In a second embodiment, a ceramic turbine rotor rotates just inside a ceramic shroud which is part of a single piece ceramic volute/shroud assembly. As temperature increases, the ceramic volute expands at approximately the same rate as ceramic shroud and tends to increase the axial clearance gap between the ceramic rotor and ceramic shroud, but only by a small amount compared to a metallic volute attached to the shroud in the same way.
1. A gas turbine engine, comprising: at least one turbo-compressor spool assembly, wherein the at least one turbo-compressor spool assembly comprises a compressor in mechanical communication with a turbine, a volute directing an inlet gas towards an inlet of a rotor of the turbine and a shroud adjac
1. A gas turbine engine, comprising: at least one turbo-compressor spool assembly, wherein the at least one turbo-compressor spool assembly comprises a compressor in mechanical communication with a turbine, a volute directing an inlet gas towards an inlet of a rotor of the turbine and a shroud adjacent to the rotor of the turbine, the shroud directing an outlet gas towards an outlet of the at least one turbo-compressor spool assembly; anda clearance control device to substantially maintain, during the at least one turbo-compressor spool assembly operation, an operational clearance between the rotor and shroud at a level no greater than about 110% of a non-operational clearance between the rotor and shroud when the at least one turbo-compressor spool assembly is non-operational; andwherein the clearance control device comprises: (a) a metallic shroud carrier connected to an engine housing and case and to the shroud, the shroud being ceramic, (b) a labyrinth metallic seal sleeve, and (c) the volute comprising a labyrinth seal engaging the labyrinth metallic seal sleeve, the labyrinth seal and seal sleeve sealing substantially against gas flow. 2. The engine of claim 1, wherein an inlet gas to the turbine is heated by a fuel combustor, wherein the inlet gas has a temperature of from about 1,000 K to about 1,400 K, and the outlet gas has a temperature less than the inlet gas, the outlet gas temperature ranging from about 900 K to about 1,200 K, whereby the shroud is subjected to a temperature differential ranging from about 200 K to about 400 K. 3. The engine of claim 2, wherein the rotor and shroud comprise a ceramic material of substantially identical thermal expansion characteristics and wherein the volute interfaces with the ceramic shroud. 4. The engine of claim 2, wherein the shroud and the volute interfacing with the shroud each comprise a substantially identical ceramic composition. 5. The engine of claim 3, wherein the volute comprises circumferential rings and grooves to form the labyrinth seal. 6. The engine of claim 5, wherein the shroud carrier is positioned between the volute and ceramic shroud and wherein a coefficient of thermal expansion of the shroud carrier is larger than a coefficient of thermal expansion of the ceramic shroud. 7. The engine of claim 1, wherein the clearance control device comprises an armature attached to an engine component and to the shroud carrier, the armature being cooled, during at least one turbo-compressor spool assembly operation, by a cooling fluid having a temperature less than the outlet gas temperature. 8. The engine of claim 7, wherein the cooling fluid is a gas removed from an input gas to at least one of the compressor, a combustor, and a recuperator. 9. The engine of claim 7, wherein the cooling fluid has a temperature of from about 400 to about 800 K and wherein the armature is metallic. 10. A method, comprising: providing an engine comprising at least one turbo-compressor spool assembly, wherein the at least one turbo-compressor spool assembly comprises a compressor in mechanical communication with a turbine, a volute adjacent to a rotor of the turbine directing an inlet gas towards an inlet of the turbine rotor, and a shroud adjacent to the turbine rotor, the shroud directing an outlet gas towards an outlet of the at least one turbo-compressor spool assembly;substantially maintaining, during the at least one turbo-compressor spool assembly operation, an operational clearance between the rotor and shroud at a level no greater than about 110% of a non-operational clearance between the rotor and shroud when the at least one turbo-compressor spool assembly is non-operational; andwherein the engine further comprises (a) a metallic shroud carrier connected to an engine housing and case and to the shroud, the shroud being ceramic, (b) a labyrinth metallic seal sleeve, and (c) the volute comprising a labyrinth seal engaging the labyrinth metallic seal sleeve, the labyrinth seal and seal sleeve sealing substantially against gas flow. 11. The method of claim 10, wherein the inlet gas to the turbine is heated by a fuel combustor, the inlet gas has a temperature of from about 1,000 K to about 1,400 K, and the outlet gas has a temperature less than the inlet gas, the outlet gas temperature ranging from about 900 K to about 1,200 K, whereby the shroud is subjected to a temperature differential ranging from about 200 K to about 400 K. 12. The method of claim 11, wherein the rotor and shroud each comprise a ceramic material of substantially identical thermal expansion characteristics and wherein the volute is in mechanical communication with the ceramic shroud. 13. The method of claim 11, wherein the shroud is in mechanical communication with the volute, and the shroud and volute each comprise a substantially identical ceramic composition. 14. The method of claim 13, wherein the volute comprises circumferential rings and grooves to form the labyrinth seal. 15. The method of claim 12, wherein the shroud carrier is positioned between the volute and ceramic shroud and wherein a coefficient of thermal expansion of the shroud carrier is larger than a coefficient of thermal expansion of the ceramic shroud. 16. The method of claim 10, wherein the engine further comprises an armature attached to an engine component and to the shroud carrier and further comprising: contacting at least one of the shroud carrier and armature, during the at least one turbo-compressor spool assembly operation, with a cooling fluid having a temperature less than the outlet gas temperature to cool the at least one of the shroud carrier and armature. 17. The method of claim 16, wherein the cooling fluid is a gas removed from an input gas to at least one of the compressor, a combustor, and a recuperator. 18. The method of claim 16, wherein the cooling fluid has a temperature of from about 400 to about 800 K and wherein the armature is nonceramic. 19. A gas turbine engine, comprising: at least one turbo-compressor spool assembly, wherein the at least one turbo-compressor spool assembly comprises a compressor in mechanical communication with a turbine, a volute directing an input gas to a rotor of the turbine, and a shroud adjacent to the turbine rotor, the shroud directing an outlet gas towards an outlet of the at least one turbo-compressor spool assembly, wherein the volute and shroud each comprise a ceramic material to maintain, during the at least one turbo-compressor spool assembly operation, at least an operational clearance between the rotor and shroud of no more than about 110% of a non-operational clearance between the rotor and shroud when the at least one turbo-compressor spool assembly is non-operational; andwherein the gas turbine engine further comprises (a) a metallic shroud carrier connected to an engine housing and case and to the shroud (b) a labyrinth metallic seal sleeve, and (c) the volute comprising a labyrinth seal engaging the labyrinth metallic seal sleeve, the labyrinth seal and seal sleeve sealing substantially against gas flow. 20. The engine of claim 19, wherein the rotor comprises a ceramic material and further comprising: a clearance control device to substantially maintain, during the at least one turbo-compressor spool assembly operation, the operational clearance between the rotor and shroud at a level no greater than the non-operational clearance between the rotor and shroud when the at least one turbo-compressor spool assembly is non-operational. 21. The engine of claim 19, wherein the ceramic composition is one or more of alumina, cordierite, silicon carbide, silicon nitride, and mullite. 22. The engine of claim 19, wherein the rotor comprises a ceramic material and wherein the rotor, volute, and shroud have substantially the same coefficient of thermal expansion and thermal contraction.
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