Reverse-flow core gas turbine engine with a pulse detonation system
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-003/14
F02K-003/11
F23R-007/00
F02K-003/06
F02C-005/02
F02C-007/06
F02C-007/36
F23G-007/06
F02C-003/107
출원번호
US-0764186
(2013-01-29)
등록번호
US-10094279
(2018-10-09)
국제출원번호
PCT/US2013/023565
(2013-01-29)
국제공개번호
WO2014/120115
(2014-08-07)
발명자
/ 주소
Kupratis, Daniel B.
Moon, Francis R
출원인 / 주소
United Technologies Corporation
대리인 / 주소
Bachman & LaPointe, PC
인용정보
피인용 횟수 :
0인용 특허 :
1
초록▼
The engine (10) includes a low spool (16) disposed aft of an air inlet (12) and a high spool (34) disposed aft of the low spool (16). An intake reverse-duct (44) is disposed radially outward of the high spool (34) and reverses direction of low pressure compressed air from the low spool (16) into a f
The engine (10) includes a low spool (16) disposed aft of an air inlet (12) and a high spool (34) disposed aft of the low spool (16). An intake reverse-duct (44) is disposed radially outward of the high spool (34) and reverses direction of low pressure compressed air from the low spool (16) into a forward-flow high pressure compressor (40) of the high spool (34). A discharge reverse-manifold (48) directs flow of an exhaust gas stream (50} from a forward-flow low pressure turbine (20) into a rearward-flow direction and into at least one pulse detonation firing tube (54). An annular bypass air duct (72) directs cooling air along the engine (10)—The at least, one firing tube is positioned radially outward of the high spool (34), overlies the high spool (34) and is also positioned within the bypass air duct (72).
대표청구항▼
1. A reverse-flow core gas turbine engine having an air inlet and an engine exhaust aft of the air inlet, the engine comprising: a. a low spool disposed aft of the air inlet including a rearward-flow low pressure compressor, and a forward-flow low pressure turbine disposed aft of the rearward-flow l
1. A reverse-flow core gas turbine engine having an air inlet and an engine exhaust aft of the air inlet, the engine comprising: a. a low spool disposed aft of the air inlet including a rearward-flow low pressure compressor, and a forward-flow low pressure turbine disposed aft of the rearward-flow low pressure compressor;b. a high spool disposed aft of the low spool, the high spool including a forward-flow high pressure turbine disposed aft of the forward-flow low pressure turbine, a combustor disposed aft of the forward-flow high pressure turbine, and a forward-flow high pressure compressor disposed aft of the combustor;c. an intake reverse-duct disposed radially outward of the high spool for directing output of the rearward-flow low pressure compressor to the forward-flow high pressure compressor so that the output reverses from rearward-flow to forward-flow to pass through the high spool;d. a discharge reverse-manifold disposed forward of the high spool and radially outward of the intake reverse-duct for receiving an exhaust gas stream from the forward-flow low pressure turbine and for directing the exhaust gas stream from forward-flow to rearward-flow toward the engine exhaust;e. a pulse detonation system including at least one pulse detonation firing tube secured in fluid communication with the discharge reverse-manifold, the at least one pulse detonation firing tube positioned to be radially outward of and to overlie the high spool so that a portion of the pulse detonation firing tube intersects an axis that is perpendicular to an engine center line and which axis passes through the high spool; and,f. the at least one pulse detonation firing tube being configured to mix all of the exhaust gas stream with fuel so that the mixed fuel and exhaust gas stream pulse detonates as the mixed fuel and exhaust gas stream pass through the firing tube toward the engine exhaust, wherein the engine further comprises an annular bypass duct surrounding and extending radially outward of the low spool and the high spool that directs bypass air from the air inlet to the exhaust of the engine, and wherein the one or more pulse detonation firing tubes are positioned within the annular bypass duct and exposed to cooling air passing through the bypass duct, the one or more pulse detonation firing tubes defining two separate flow paths within the annular bypass duct, wherein the first flow path is through the pulse detonation firing tubes and carries the exhaust gas stream from the discharge reverse-manifold, and wherein the second flow path is in the annular bypass duct outside of the pulse detonation firing tubes and carries bypass air. 2. The reverse-flow core gas turbine engine of claim 1, further comprising a plurality of pulse detonation firing tubes, wherein each of the plurality of firing tubes is positioned about an equal distance from adjacent firing tubes, and wherein the plurality of firing tubes are positioned to surround the high spool of the engine. 3. The reverse-flow core gas turbine engine of claim 2, wherein the plurality of pulse detonation firing tubes includes between about eleven and about twenty-two firing tubes. 4. The reverse-flow core gas turbine engine of claim 1, further comprising a plurality of pulse detonation firing tubes and wherein the plurality of firing tubes comprise constant volume combustor tubes. 5. The reverse-flow core gas turbine engine of claim 1, further comprising a plurality of groups of adjacent pulse detonation firing tubes, each of the plurality of groups of adjacent firing tubes being positioned so that each group is about an equal distance from a closest group of firing tubes, and the groups of adjacent pulse detonation firing tubes being positioned to surround the high spool of the engine. 6. A reverse-flow core gas turbine engine having an air inlet and an engine exhaust aft of the air inlet, the engine comprising: a. a low spool disposed aft of the air inlet including a rearward-flow low pressure compressor, a forward-flow low pressure turbine disposed aft of the rearward-flow low pressure compressor and a low pressure shaft secured between the low pressure turbine and the low pressure compressor;b. a high spool disposed aft of the low spool, the high spool including a forward-flow high pressure turbine disposed aft of the forward-flow low pressure turbine, a combustor disposed aft of the forward-flow high pressure turbine, a forward-flow high pressure compressor disposed aft of the combustor, and a high pressure shaft secured between the high pressure turbine and the high pressure compressor;c. an intake reverse-duct disposed radially outward of the high spool for directing output of the rearward-flow low pressure compressor to the forward-flow high pressure compressor so that the output reverses from rearward-flow to forward-flow to pass through the high spool;d. a discharge reverse-manifold disposed forward of the high spool and radially outward of the intake reverse-duct for receiving an exhaust gas stream from the forward-flow low pressure turbine and for directing the exhaust gas stream from forward-flow to rearward-flow toward the engine exhaust;e. a pulse detonation system including at least one pulse detonation firing tube secured in fluid communication with the discharge reverse-manifold, the at least one pulse detonation firing tube positioned to be radially outward of and to overlie the high spool so that a portion of the pulse detonation firing tube intersects an axis that is perpendicular to an engine center line and which axis passes through the high spool;f. the at least one pulse detonation firing tube being configured to mix all of the exhaust gas stream with fuel so that the mixed fuel and exhaust gas stream pulse detonates as the mixed fuel and exhaust gas stream pass through the firing tube toward the engine exhaust, wherein the engine further comprises an annular bypass duct surrounding and extending radially outward of the low spool and the high spool that directs bypass air from the air inlet to the exhaust of the engine, and wherein the one or more pulse detonation firing tubes are positioned within the annular bypass duct and exposed to cooling air passing through the bypass duct, the one or more pulse detonation firing tubes defining two separate flow paths within the annular bypass duct, wherein the first flow path is through the pulse detonation firing tubes and carries the exhaust gas stream from the discharge reverse-manifold, and wherein the second flow path is in the annular bypass duct outside of the pulse detonation firing tubes and carries bypass air; andg. wherein flow through the engine extends sequentially through the rearward-flow low pressure compressor in a downstream direction, through the intake reverse-duct to the forward-flow high pressure compressor, the combustor, the forward-flow high pressure turbine and the forward-flow low pressure turbine in a forward direction, and through the discharge reverse-manifold to a rearward direction through the at least one pulse detonation firing tube. 7. The reverse-flow core gas turbine engine of claim 6, further comprising a plurality of pulse detonation firing tubes, wherein each of the plurality of firing tubes is positioned about an equal distance from adjacent firing tubes, and wherein the plurality of firing tubes are positioned to surround the high spool of the engine. 8. The reverse-flow core gas turbine engine of claim 6, wherein the one or more pulse detonation firing tubes comprise pulse detonation shock tubes. 9. A method of operating a reverse-flow gas turbine engine, the method comprising: a. directing flow of inlet air through an air inlet of the engine;b. then, compressing the air in a downstream direction through a rearward-flow low pressure compressor of a low spool of the engine;c. then, directing flow of the compressed air through an intake reverse-duct, and then into a forward-flow high pressure compressor, then combusting the compressed air by directing the compressed air with fuel through a combustor, then directing a combusted gas stream from the combustor and through a forward-flow high pressure turbine of a high spool, and then through a forward-flow low pressure turbine in a forward direction;d. then, directing all of the exhaust gas stream from the forward-flow low pressure turbine through a discharge reverse-manifold and then in a rearward direction through at least one pulse detonation firing tube overlying the high spool; ande. mixing all of the exhaust gas stream with fuel within the at least one pulse detonation firing tube and pulse detonating the mixed fuel and exhaust gas stream within the at least one pulse detonation firing tube, wherein the engine further comprises an annular bypass duct surrounding and extending radially outward of the low spool and the high spool that directs bypass air from the air inlet to the exhaust of the engine, and wherein the one or more pulse detonation firing tubes are positioned within the annular bypass duct and exposed to cooling air passing through the bypass duct, the one or more pulse detonation firing tubes defining two separate flow paths within the annular bypass duct, wherein the first flow path is through the pulse detonation firing tubes and carries the exhaust gas stream from the discharge reverse-manifold, and wherein the second flow path is in the annular bypass duct outside of the pulse detonation firing tubes and carries bypass air. 10. A method of operating the reverse-flow gas turbine engine of claim 9, further comprising directing the exhaust gas stream from the forward-flow low pressure turbine through a plurality of pulse detonation firing tubes positioned about an equal distance from adjacent firing tubes and positioned to surround the high spool of the engine. 11. A method of operating the reverse-flow gas turbine engine of claim 9, further comprising directing the exhaust gas stream from the forward-flow low pressure turbine through a plurality of groups of adjacent pulse detonation firing tubes positioned so that each group is about an equal distance from a closest group of firing tubes, and positioned to surround the high spool of the engine. 12. A method of operating the reverse-flow gas turbine engine of claim 9, wherein the step of directing the exhaust gas stream through at least one pulse detonation firing tube overlying the high spool further comprises generating shock waves within the at least one pulse detonation firing tube.
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이 특허에 인용된 특허 (1)
Louis G. Hunter ; Billy D. Couch ; Paul E. Hagseth, Combined cycle pulse combustion/gas turbine engine.
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