Gas turbine engine variable area exhaust nozzle
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02K-003/075
F02K-001/12
F02C-009/18
F01D-017/14
F02K-003/04
F02K-003/06
F01D-017/08
출원번호
US-0585077
(2009-09-02)
등록번호
US-8904750
(2014-12-09)
우선권정보
GB-0820174.1 (2008-11-05)
발명자
/ 주소
Hillel, Malcolm L.
Brown, Stephen G.
출원인 / 주소
Rolls-Royce PLC
대리인 / 주소
Oliff PLC
인용정보
피인용 횟수 :
0인용 특허 :
10
초록▼
A turbofan gas turbine engine (10) comprises a variable area exhaust nozzle (12) arranged at the downstream end of a casing (17). A control unit (66) analyzes the power produced by the gas turbine engine (10), the flight speed of the gas turbine engine (1) and/or the altitude of the gas turbine engi
A turbofan gas turbine engine (10) comprises a variable area exhaust nozzle (12) arranged at the downstream end of a casing (17). A control unit (66) analyzes the power produced by the gas turbine engine (10), the flight speed of the gas turbine engine (1) and/or the altitude of the gas turbine engine (10). The control unit (66) configures the variable area nozzle (12) at a first cross-sectional area (70A) when the flight speed of the gas turbine engine (10) is less than a first predetermined value. The control unit (66) configures the variable area nozzle (12) at a second, smaller, cross-sectional area (70B) when the flight speed of the gas turbine engine (10) is greater than the first predetermined value and the power produced by the gas turbine engine (10) is greater than a second predetermined value. The control unit (66) configures the variable area nozzle (12) at a third, intermediate, cross-sectional area (70C) when the flight speed of the gas turbine engine (10) is greater than the first predetermined value and the power produced by the gas turbine engine (10) is less than the second predetermined value.
대표청구항▼
1. A gas turbine engine comprising: a casing defining a flow passage through the gas turbine engine;a variable area exhaust nozzle being arranged at the downstream end of the casing, the variable area exhaust nozzle having a downstream end and a cross-sectional area measured at the downstream end of
1. A gas turbine engine comprising: a casing defining a flow passage through the gas turbine engine;a variable area exhaust nozzle being arranged at the downstream end of the casing, the variable area exhaust nozzle having a downstream end and a cross-sectional area measured at the downstream end of the variable area exhaust nozzle;a power sensor configured to measure the power produced by the gas turbine engine;a flight speed sensor configured to measure the flight speed of the gas turbine engine or an altitude sensor configured to measure the altitude of the gas turbine engine;a control device configured to receive measurements of the power produced by the gas turbine engine, the flight speed of the gas turbine engine or the altitude of the gas turbine engine,the control device configured to analyse the power produced by the gas turbine engine, the flight speed of the gas turbine engine or the altitude of the gas turbine engine,the control device configured to configure the cross-sectional area of the variable area exhaust nozzle at a first cross-sectional area in a first mode of operation when the control device determines that the flight speed of the gas turbine engine or the altitude of the gas turbine engine is less than a first predetermined value,the control device configured to configure the cross-sectional area of the variable area exhaust nozzle at a second cross-sectional area that is greater than zero in a second mode of operation when the control device determines that the flight speed of the gas turbine engine or the altitude of the gas turbine engine is equal to or greater than the first predetermined value and the power produced by the gas turbine engine is equal to or greater than a second predetermined value, andthe control device is arranged to configure the cross-sectional area of the variable area exhaust nozzle at a third cross-sectional area in a third mode of operation when the control device determines that the flight speed of the gas turbine engine or the altitude of the gas turbine engine is equal to or greater than the first predetermined value and the power produced by the gas turbine engine is less than the second predetermined value, whereinthe second cross-sectional area is less than both the first cross-sectional area and the third cross-sectional area and the third cross-sectional area is less than the first cross-sectional area. 2. The gas turbine engine as claimed in claim 1, wherein the control device is configured to determine if the power produced by the gas turbine engine has reduced in the third mode of operation of the gas turbine engine, andthe control device is configured to configure the cross-sectional area of the variable area exhaust nozzle at a plurality of cross-sectional areas in the third mode of operation in response to a reduction in the power produced by the gas turbine engine. 3. The gas turbine engine as claimed in claim 1, wherein the first cross-sectional area is a maximum cross-sectional area and the second cross-sectional area is a minimum cross-sectional area. 4. The gas turbine engine as claimed in claim 1, wherein the power sensor includes one of a rotational speed sensor configured to measure a rotational speed of a shaft of the gas turbine engine, a pressure sensor configured to measure a pressure ratio of the gas turbine engine, and turbine entry temperature sensor configured to measure turbine entry temperature of the gas turbine engine. 5. The gas turbine engine as claimed in claim 4, wherein the power sensor includes a sensor configured to measure the rotational speed of a fan shaft of the gas turbine engine and a sensor configured to measure the temperature at the intake of the gas turbine engine, andthe control device determines the rotational speed of the fan shaft divided by the root of the temperature at the intake of the gas turbine engine. 6. The gas turbine engine as claimed in claim 4, wherein the pressure sensor includes a bypass pressure sensor configured to measure the pressure in a bypass duct of the gas turbine engine and an intake pressure sensor configured to measure the pressure at the intake of the gas turbine engine, andthe control device determines the pressure in the bypass duct of the gas turbine engine divided by the pressure at the intake of the gas turbine engine. 7. The gas turbine engine as claimed in claim 1, wherein the casing is a fan casing and the variable area exhaust nozzle is a variable area fan nozzle. 8. The gas turbine engine as claimed in claim 7, further comprising: an intake being arranged at the upstream end of the fan casing,wherein the fan casing extends from the intake to the variable area fan exhaust nozzle without a further air intake. 9. The gas turbine engine as claimed in claim 1, wherein the first mode of operation is take-off conditions, the second mode of operation is at high power conditions at high altitude and the third mode of operation is at cruise conditions. 10. The gas turbine engine as claimed in claim 1, wherein the gas turbine engine is a turbofan gas turbine engine having a fan and a core engine, the core engine having a core exhaust nozzle and the fan having a fan exhaust nozzle, the fan exhaust nozzle being arranged around the core exhaust nozzle, the casing is a fan casing and the variable area exhaust nozzle is the fan exhaust nozzle. 11. The gas turbine engine as claimed in claim 10, wherein the core engine comprising a series of compressors, a combustor and a series of turbines, all of the turbines being arranged in flow series, an intake being arranged at the upstream end of the fan casing, the fan casing extending from the intake to the variable area fan exhaust nozzle without a further air intake, the variable area fan exhaust nozzle having a downstream end and a cross-sectional area, the cross-sectional area being at the downstream end of the variable area fan exhaust nozzle, and the second cross-sectional area of the variable area fan exhaust nozzle is greater than zero. 12. A method of operating a gas turbine engine having a casing defining a flow passage through the gas turbine engine, a variable area exhaust nozzle being arranged at the downstream end of the casing, the variable area exhaust nozzle having a downstream end and a cross-sectional area measured at the downstream end of the variable area exhaust nozzle, the method comprising: measuring the power produced by the gas turbine engine,measuring the flight speed of the gas turbine engine or measuring the altitude of the gas turbine engine,analysing the power produced by the gas turbine engine, the flight speed of the gas turbine engine or the altitude of the gas turbine engine,configuring the cross-sectional area of the variable area exhaust nozzle at a first cross-sectional area in a first mode of operation when the flight speed of the gas turbine engine or the altitude of the gas turbine engine is less than a first predetermined value,configuring the cross-sectional area of the variable area exhaust nozzle at a second cross-sectional area that is greater than zero in a second mode of operation when the flight speed of the gas turbine engine or the altitude of the gas turbine engine is equal to or greater than the first predetermined value and the power produced by the gas turbine engine is equal to or greater than a second predetermined value, andconfiguring the cross-sectional area of the variable area exhaust nozzle at a third cross-sectional area in a third mode of operation when the flight speed of the gas turbine engine or the altitude of the gas turbine engine is equal to or greater than the first predetermined value and the power produced by the gas turbine engine is less than the second predetermined value, whereinthe second cross-sectional area is less than the first cross-sectional area and the third cross-sectional area is greater than the second cross-sectional area and less than the first cross-sectional area. 13. The method as claimed in claim 12, further comprising: determining if the power produced by the gas turbine engine has reduced in the third mode of operation of the gas turbine engine and configuring the cross-sectional area of the variable area exhaust nozzle at a plurality of cross-sectional areas in the third mode of operation in response to a reduction in the power produced by the gas turbine engine. 14. The method as claimed in claim 12, wherein the first cross-sectional area is a maximum cross-sectional area and the second cross-sectional area is a minimum cross-sectional area. 15. The method as claimed in claim 9, wherein measuring the power produced by the gas turbine engine includes at least one of measuring a rotational speed of a shaft of the gas turbine engine, measuring a pressure ratio of the gas turbine engine, and measuring turbine entry temperature of the gas turbine engine. 16. The method as claimed in claim 15, wherein measuring the power produced by the gas turbine engine includes measuring the rotational speed of a fan shaft of the gas turbine engine, measuring the temperature at the intake of the gas turbine engine, and determining the rotational speed of the fan shaft divided by the root of the temperature at the intake of the gas turbine engine. 17. The method as claimed in claim 15, wherein measuring the pressure ratio of the gas turbine engine includes measuring the pressure in a bypass duct of the gas turbine engine, measuring the pressure at the intake of the gas turbine engine, and determining the pressure in the bypass duct of the gas turbine engine divided by the pressure at the intake of the gas turbine engine. 18. The method as claimed in claim 12, wherein the casing is a fan casing and the variable area exhaust nozzle is a variable area fan nozzle. 19. The method as claimed in claim 18, wherein the gas turbine engine includes: an intake being arranged at an upstream end of the fan casing, andthe fan casing extends from the intake to the variable area fan exhaust nozzle without a further air intake. 20. The method as claimed in claim 12, wherein the first mode of operation is take-off conditions, the second mode of operation is at high power conditions at high altitude and the third mode of operation is cruise conditions. 21. The method of operating a gas turbine engine as claimed in claim 12, wherein the gas turbine engine is a turbofan gas turbine engine having a fan and a core engine, the core engine having a core exhaust nozzle and the fan having a fan exhaust nozzle, the fan exhaust nozzle being arranged around the core exhaust nozzle, the casing is a fan casing and the variable area nozzle is the fan exhaust nozzle. 22. The method as claimed in claim 21, wherein the core engine comprising a series of compressors, a combustor and a series of turbines, all of the turbines being arranged in flow series, an intake being arranged at the upstream end of the fan casing, the fan casing extending from the intake to the variable area fan exhaust nozzle without a further air intake, the variable area fan exhaust nozzle having a downstream end and a cross-sectional area, the cross-sectional area being at the downstream end of the variable area fan exhaust nozzle, and the second cross-sectional area of the variable area fan exhaust nozzle is greater than zero. 23. A method of operating a turbofan gas turbine engine having a fan and a core engine, the core engine having a core exhaust nozzle and the fan having a fan exhaust nozzle, the fan exhaust nozzle being arranged around the core exhaust nozzle, a fan casing defining a flow passage through the turbofan gas turbine engine, and a variable area fan exhaust nozzle being arranged at the downstream end of the fan casing, the variable area fan exhaust nozzle having downstream end and a cross-sectional area measured at the downstream end of the variable area fan exhaust nozzle, the method comprising: in a first mode of operation, arranging the cross-sectional area of the variable area fan exhaust nozzle at a maximum cross-sectional area,in a second mode of operation, arranging the cross-sectional area of the variable area fan exhaust nozzle at a minimum cross-sectional area that is greater than zero, andin a third mode of operation, arranging the cross-sectional area of the variable area fan exhaust nozzle at an intermediate cross-sectional area, whereinthe first mode of operation is take-off conditions, the second mode of operation is at high power conditions at high altitude, the third mode of operation is cruise conditions, and the second mode of operation is at a higher power than the third mode of operation. 24. The method as claimed in claim 23, wherein the gas turbine engine includes: an intake being arranged at an upstream end of the fan casing, andthe fan casing extends from the intake to the variable area fan exhaust nozzle without a further air intake. 25. A method of operating a turbofan gas turbine engine comprising a fan and a core engine, the core engine having a core exhaust nozzle and the fan having a fan exhaust nozzle,the core engine comprising a series of compressors, a combustor and a series of turbines, all of the turbines being arranged in flow series,the fan exhaust nozzle being arranged around the core exhaust nozzle,a fan casing having an upstream end and a downstream end, the fan casing defining a flow passage through the turbofan gas turbine engine,a variable area fan exhaust nozzle arranged at the downstream end of the fan casing,the variable area fan exhaust nozzle having a downstream end and a cross-sectional area measured at the downstream end of the variable area fan exhaust nozzle,the method comprising: in a first mode of operation arranging the cross-sectional area of the variable area fan exhaust nozzle at a maximum cross-sectional area;in a second mode of operation arranging the cross-sectional area of the variable area fan exhaust nozzle at a minimum cross-sectional area, the second cross-sectional area of the variable area fan exhaust nozzle is greater than zero; andin a third mode of operation arranging the cross-sectional area of the variable area fan exhaust nozzle at an intermediate cross-sectional area,wherein the first mode of operation is take-off conditions, the second mode of operation is at high power conditions at high altitude, the third mode of operation is cruise conditions, and the second mode of operation is at a higher power than the third mode of operation. 26. The method as claimed in claim 25, wherein the turbofan gas turbine engine includes: an intake being arranged at the upstream end of the fan casing, andthe fan casing extends from the intake to the variable area fan exhaust nozzle without a further air intake.
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이 특허에 인용된 특허 (10)
Pollak Robert R. (North Palm Beach FL) Khalid Syed J. (Palm Beach Gardens FL) Marcos Juan A. (Jupiter FL), Active geometry control system for gas turbine engines.
Duesler Paul W. ; Loffredo Constantino V. ; Prosser ; Jr. Harold T. ; Jones Christopher W., Variable area fan exhaust nozzle having mechanically separate sleeve and thrust reverser actuation systems.
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