Gas turbine engine and method for operating a gas turbine engine
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F23R-003/28
F23N-005/02
F02C-009/32
F02C-009/28
F02C-007/228
F02C-009/26
출원번호
US-0136067
(2013-12-20)
등록번호
US-10228139
(2019-03-12)
발명자
/ 주소
Cummings, III, William G.
출원인 / 주소
Rolls-Royce Corporation
대리인 / 주소
Fishman Stewart PLLC
인용정보
피인용 횟수 :
0인용 특허 :
20
초록▼
A method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft may include measuring a gas temperature downstream of the combustor using a plurality of temperature measurement devices circumferentially disposed about an engine centerline and varying the
A method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft may include measuring a gas temperature downstream of the combustor using a plurality of temperature measurement devices circumferentially disposed about an engine centerline and varying the output of at least one fuel injector in a first direction and the output of at least another fuel injector in a second direction, while maintaining a thrust output of the gas turbine engine during the flight operation. A gas turbine engine may include temperature measurement devices circumferentially disposed about an engine centerline and a system configured for varying the output of at least one fuel injector in a first direction and the output of at least another fuel injector in a second direction, while maintaining a power output of the gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines.
대표청구항▼
1. A method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft, comprising: providing a fuel flow into a combustor via a plurality of fuel injectors circumferentially disposed about an engine centerline, and combusting fuel in the combustor;measuring
1. A method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft, comprising: providing a fuel flow into a combustor via a plurality of fuel injectors circumferentially disposed about an engine centerline, and combusting fuel in the combustor;measuring a gas temperature downstream of the combustor using a plurality of temperature measurement devices that are equal in quantity to the plurality of fuel injectors and circumferentially disposed about the engine centerline such that each of the plurality of temperature measurement devices is disposed circumferentially between respective circumferentially adjacent pairs of the plurality of fuel injectors;varying, by way of a full authority digital engine controller (FADEC), a first output of an individual one of the plurality of fuel injectors in a first direction by a first amount of fuel flow for a predetermined time period and a second output of other ones of the plurality of fuel injectors in a second direction that is opposite the first direction by a second amount of fuel flow for the predetermined time period, such that a thrust output of the gas turbine engine during the flight operation remains the same as prior to the varying; anddetermining which of the plurality of temperature measurement devices is impacted by the varying the first output of the individual one of the plurality of fuel injectors in the first direction by determining which of the plurality of temperature measurement devices experiences a change in a gas temperature measurement following the varying, andtrimming the second output of at least one of the other ones of the plurality of fuel injectors based on the gas temperature measurement from the plurality of temperature measurement devices that are impacted by the varying of the first output of the individual one of the plurality of fuel injectors in the first direction to yield a uniform measured gas temperature between the plurality of temperature measurement devices. 2. The method of claim 1, further comprising at least one of: providing the plurality of temperature measurement devices arranged in a circumferentially staggered relationship to the plurality of fuel injectors;providing the plurality of temperature measurement devices on a leading edge of at least one of the turbine vane; andproviding second stage vanes and third stage vanes, and the plurality of temperature measurement devices on at least one of the second stage vanes and the third stage vanes. 3. The method of claim 1, wherein the providing of the fuel flow into the combustor is providing the fuel flow into an annular combustor. 4. The method of claim 1, wherein the varying of the first output of the individual one of the plurality of fuel injectors in the first direction is performed at the same time as the varying of the second output of the other ones of the plurality of fuel injectors in the second direction is performed such that a total fuel output to the plurality of fuel injectors remains constant during the varying. 5. The method of claim 4, wherein the varying of the second output of the other ones of the plurality of fuel injectors is performed sequentially in a circumferential direction. 6. The method of claim 1, further comprising trimming the second output of the at least one of the other ones of the plurality of fuel injectors by operating a controller associated with the individual one of the plurality of fuel injectors. 7. A gas turbine engine, comprising: a compressor;a combustor in fluid communication with the compressor;a turbine in fluid communication with the combustor; anda system for providing fuel flow to the combustor, the system including a first controller; a plurality of fuel injectors circumferentially disposed about an engine centerline and operative to discharge fuel into the combustor;a plurality of second controllers, each second controller corresponding to at least one fuel injector of the plurality of fuel injectors, wherein each second controller is communicatively coupled to the first controller and is configured to control a flow of fuel discharged by the corresponding at least one fuel injector under the direction of the first controller; anda plurality of temperature measurement devices that are equal in quantity to the plurality of fuel injectors and circumferentially disposed about the engine centerline downstream of the combustor, and each of the plurality of temperature measurement devices is disposed circumferentially between respective circumferentially adjacent pairs of the plurality of fuel injectors, wherein the temperature measurement devices are communicatively coupled to the first controller and configured to supply measured gas temperature data to the first controller,wherein the the first controller is configured to control a fuel flow into the combustor via the plurality of fuel injectors for combusting the fuel in the combustor by measuring a gas temperature downstream of the combustor using the plurality of temperature measurement devices, and varying an first output of an individual one of the plurality of fuel injectors in a first direction by a first amount of fuel flow for a predetermined time period and a second output of other ones of the plurality of fuel injectors in a second direction that is opposite the first direction by a second amount of fuel flow for the predetermined time period such that a power output of the gas turbine engine is maintained the same as prior to the varying, anddetermine which of the plurality of temperature measurement devices are impacted by the varying the first output of the individual one of the plurality of fuel injectors in the first direction by determining which of the plurality of temperature measurement devices experiences a change in a gas temperature measurement following the varying, andtrim the second output of at least one of the other ones of the plurality of fuel injectors based on the gas temperature measurement from the plurality of temperature measurement devices that are impacted by the varying of the first output of the individual one of the plurality of fuel injectors in the first direction to yield a uniform measured gas temperature as between the plurality of temperature measurement devices. 8. The gas turbine engine of claim 7, wherein the first controller includes a full authority digital engine controller (FADEC). 9. The gas turbine engine of claim 7, wherein the combustor is an annular combustor. 10. The gas turbine engine of claim 7, wherein the first controller is configured to perform the varying of the first output of the individual one of the plurality of fuel injectors in the first direction at the same time as the varying of the second output of the other ones of the plurality of fuel injectors in the second direction such that a total fuel output to the plurality of fuel injectors remains constant during the varying. 11. The gas turbine engine of claim 10, wherein the first controller is configured to perform the varying of the second output of the other ones of the plurality of fuel injectors sequentially in a circumferential direction. 12. The gas turbine engine of claim 7, wherein the first controller is configured for trimming the second output of the at least one of the other ones of the plurality of fuel injectors by operating at least one of the plurality of second controllers under the direction of the first controller. 13. The gas turbine engine of claim 7, wherein the first controller is configured to vary the first output of the individual one of the plurality of fuel injectors in the first direction and the second output of the at least one of the other ones of the plurality of fuel injectors in a second direction, while maintaining a constant thrust output of the gas turbine engine during flight operations in an aircraft. 14. A gas turbine engine, comprising: a compressor;a combustor in fluid communication with the compressor;a plurality of fuel injectors operative to provide fuel for combustion in the combustor;a turbine in fluid communication with the combustor; anda controller including a full authority digital engine controller (FADEC) configured for balancing the fuel flow provided from the plurality of fuel injectors, the controller being communicatively coupled to each of a plurality of temperature measurement devices with at least a first communication link and to each of the plurality of fuel injectors with at least a second communication link, the plurality of temperature measurement devices being equal in quantity to the plurality of fuel injectors, each of the plurality of temperature measurement devices being disposed circumferentially between respective circumferentially adjacent pairs of the plurality of fuel injectors,wherein the controller is configured to: vary a first output of an individual one of the plurality of fuel injectors in a first direction by a first amount of fuel flow for a predetermined time period and a second output of other ones of the plurality of fuel injectors in a second direction that is opposite the first direction by a second amount of fuel flow for the predetermined time period while maintaining a thrust output of the gas turbine engine;determine which of the plurality of temperature measurement devices are impacted by the varying the first output of the individual one of the plurality of fuel injectors in the first direction by determining which of the plurality of temperature measurement devices experiences a change in a gas temperature measurement following the varying, andtrim the second output of at least one of the plurality of fuel injectors based on the gas temperature measurement from the plurality of temperature measurement devices that are impacted by the varying of the first output of the individual one of the plurality of fuel injectors in the first direction to yield a uniform measured gas temperature as between the plurality of temperature measurement devices.
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이 특허에 인용된 특허 (20)
Cornwell, Michael D.; Myhre, Douglas C.; Eriksen, O. Harald; Goeke, Jerry L., Active combustion control system for gas turbine engines.
Brocard Jean-Marie (Rubelles FRX) Hebert Pierre G. J. (Crosne FRX) de Jongh Thierry N. (Bonneuil-Matours FRX), Fuel supply system for gas turbine engines.
Beebe Kenneth W. (Schenectady NY) Davis L. Berkly (Schenectady NY) Iasillo Robert J. (Schenectady NY), Fuel trim system for a multiple chamber gas turbine combustion system.
Kothnur,Vasanth Srinivasa; Ali,Mohamed Ahmed; Vandale,Daniel Doyle; Roberts,Matthew Eugene, Method and apparatus for actuating fuel trim valves in a gas turbine.
Kothnur,Vasanth Srinivasa; Ali,Mohamed Ahmed; Vandale,Daniel Doyle; Roberts,Matthew Eugene, Method and apparatus for actuating fuel trim valves in a gas turbine.
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