Gas turbine engine having improved core system
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-006/08
F02C-006/00
출원번호
US-0941508
(2004-09-15)
발명자
/ 주소
Orlando,Robert Joseph
Venkataramani,Kattalaicheri Srinivasan
Lee,Ching Pang
출원인 / 주소
General Electric Company
인용정보
피인용 횟수 :
2인용 특허 :
20
초록▼
A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a drive shaft; a booster compressor positioned downstream of the fan section including a plurality of sta
A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a drive shaft; a booster compressor positioned downstream of the fan section including a plurality of stages, where each stage includes a stationary compressor blade row and a rotating compressor blade row connected to the drive shaft and interdigitated with the stationary compressor blade row; and, a combustion system for producing pulses of gas having increased pressure and temperature of a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet. A first source of compressed air from the booster compressor is provided to the combustion system inlet and a second source of compressed air from the booster compressor is provided to cool the combustion system, where the pressure of the compressed air from the second source has a greater pressure than that of the compressed air from the first source.
대표청구항▼
What is claimed is: 1. A gas turbine engine having a longitudinal centerline axis therethrough, comprising: (a) a fan section a forward end of said gas turbine engine including at least a first fan blade row connected to a drive shaft; (b) a booster compressor positioned downstream of said fan sec
What is claimed is: 1. A gas turbine engine having a longitudinal centerline axis therethrough, comprising: (a) a fan section a forward end of said gas turbine engine including at least a first fan blade row connected to a drive shaft; (b) a booster compressor positioned downstream of said fan section including a plurality of stages, each said stage including a stationary compressor blade row and a rotating compressor blade row connected to said drive shaft and interdigitated with said stationary compressor or blade row; and, (c) a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flow provided to an inlet thereof; when a first source of compressed air from said booster compressor is provided to said combustion system inlet and a second source of compressed air from said booster compressor is provided to cool said combustion system. 2. The gas turbine engine of claim 1, wherein compressed air from said second source has a higher pressure than compressed air from said first source. 3. The gas turbine engine of claim 1, wherein pressure of compressed air from said second source is greater than pressure of compressed air from said first source by at least approximately 20%. 4. The gas turbine engine of claim 1, wherein pressure of compressed air from said second source is greater than pressure of compressed air from said first source by at least approximately 50% a. 5. The gas turbine engine of claim 1, wherein pressure of compressed air from said second source is greater than pressure of compressed air from said first source by at least approximately 100%. 6. The gas turbine engine of claim 1, wherein said first source of compressed air originates in said booster compressor upstream of said second source of compressed air. 7. The gas turbine engine of claim 1, wherein said first source of compressed air originates between adjacent stages of said booster compressor. 8. The gas turbine of claim 1, wherein said second source of compressed air originates at an aft end of said booster compressor. 9. The gas turbine engine of claim 1, wherein pressure of compressed air from said second source is greater than pressure of fluid at an exit end of said combustion system. 10. The gas turbine engine of claim 1, wherein said combustion system is a constant volume combustion device. 11. The gas turbine engine of claim 1, wherein sad combustion system is a pulse detonation device. 12. The gas turbine of claim 1, wherein said combustion device includes at least one rotating member for powering said drive shaft. 13. The gas turbine engine of clan 1, wherein said combustion device includes no rotating members. 14. The gas turbine engine of claim 1, further comprising a turbine downstream of and in flow communication with said combustion system which powers said drive shaft. 15. The gas turbine engine of claim 14, wherein compressed air from said second source is provided to cool fluid entering said turbine. 16. The gas turbine engine of claim 1, wherein said engine operates substantially in accordance with an ideal Humphrey cycle. 17. The gas turbine engine of claim 1, wherein said combustion system increases the pressure and temperature of said fluid therein substantially simultaneously. 18. The gas turbine engine of claim 1, wherein a portion of compressed air from said second source is supplied to said combustion system to assist atomization of fuel therein. 19. The gas turbine of claim 1, further comprising a heat exchanger in flow communication with said compressed air of said second source. 20. The gas turbine engine of claim 1, wherein said gas turbine engine is able to generate a maximum of approximately 30,000 pounds of thrust. 21. A method of cooling a combustion system of a gas turbine engine including a booster compressor having a plurality of stages, wherein said combustion system produces pulses of gas having increased pressure and temperature from a fluid flow provided thereto, comprising the following steps: (a) providing a first source of compressed air from said booster compressor to said combustion system; and (b) providing a second source of compressed air from said booster compressor to cool said combustion system; wherein pressure of compressed air from said second source is greater than pressure of compressed air from said first source by a predetermined amount. 22. The method of claim 21, further comprising the step of originating said first source from a first point located between adjacent stages of said booster compressor. 23. The method of claim 22, further comprising the step of originating said second source from a second point located downstream of said first point. 24. The method of claim 21, further comprising the step of providing compressed air from said second source to an initial stage of a turbine in flow communication with said combustion system. 25. The method of claim 21, further comprising the step of cooling compressed air from said second source. 26. A gas turbine engine comprising: (a) a compressor positioned at a forward end of said gas turbine engine having a plurality of stages, each said stage including a stationary compressor blade row and a rotatable blade row connected to a first drive shaft and interdigitated with said first compressor blade row; (b) a combustion system for producing pulses of gas having increased pressure and temperature from a fluid supplied to an inlet thereof; (c) a load connected to a second drive shaft; and, (d) a turbine downstream of and in flow communication with said combustion system for powering said first and second drive shafts; wherein a first source of compressed air from said compressor is provided to said combustion system inlet and a second source of compressed air from said compressor is provided to cool said combustion system. 27. A gas turbine engine having a longitudinal centerline therethrough, comprising: (a) a fan section at a forward end of said gas turbine engine including at least a first fan blade row connected to a drive shaft; (b) a booster compressor positioned downstream of said fan section including a plurality of stages, each said stage including a stationary compressor blade row and a rotating compressor Wade row connected to said drive shaft and interdigitated with said stationary compressor blade row; (c) a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flaw provided to an inlet thereof; and, (d) a low pressure turbine downstream of and in flow communication with said combustion system which powers said drive shaft; wherein a first source of compressed air from said booster compressor is provided to said combustion system inlet and a second source of compressed air from said booster compressor is provided at a forward end of said low pressure turbine so as to mitigate effects of said gas pulses thereon.
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이 특허에 인용된 특허 (20)
Huber David John ; Briesch Michael Scot, Closed loop air cooling system for combustion turbines.
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