IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
US-0283182
(2011-10-27)
|
등록번호 |
US-9140196
(2015-09-22)
|
우선권정보 |
CH-1786/10 (2010-10-27) |
발명자
/ 주소 |
|
출원인 / 주소 |
|
대리인 / 주소 |
Buchanan Ingersoll & Rooney PC
|
인용정보 |
피인용 횟수 :
0 인용 특허 :
4 |
초록
▼
A method for controlling a gas turbine, including during transient operating states, and such a gas turbine are provided. The gas turbine includes a compressor for compressing inducted combustion air, a combustion chamber for generating hot gas by combusting a fuel with the aid of the compressed com
A method for controlling a gas turbine, including during transient operating states, and such a gas turbine are provided. The gas turbine includes a compressor for compressing inducted combustion air, a combustion chamber for generating hot gas by combusting a fuel with the aid of the compressed combustion air, and a multistage turbine for expanding the generated hot gas and performing work. The controlling of the gas turbine is carried out in accordance with the hot gas temperature which is derived from a plurality of other measured operating variables of the gas turbine. A reliable controlling of the gas turbine is achieved, even during rapid changes, by pressure measurements being gathered exclusively at different points of the gas turbine for derivation of the hot gas temperature.
대표청구항
▼
1. A method for controlling a gas turbine, wherein the gas turbine comprises a compressor for compressing inducted combustion air, a combustion chamber for generating hot gas by combusting a fuel with the aid of the compressed combustion air, and a multistage turbine for expanding the generated hot
1. A method for controlling a gas turbine, wherein the gas turbine comprises a compressor for compressing inducted combustion air, a combustion chamber for generating hot gas by combusting a fuel with the aid of the compressed combustion air, and a multistage turbine for expanding the generated hot gas and performing work, wherein the method comprises: controlling the gas turbine in accordance with a hot gas temperature which is derived from a plurality of other measured operating variables of the gas turbine,wherein a derivation of the hot gas temperature is formed exclusively through pressure measurements at different spatial points in a flow direction in a region of hot gas flow of the gas turbine,wherein the pressure measurements are carried out in the hot gas flow which is generated in the combustion chamber, andwherein the pressure measurements include a pressure loss measurement in a main flow of the gas turbine. 2. The method as claimed in claim 1, wherein the measuring positions, between which the pressure loss is determined, lie downstream of flames which are present in the combustion chamber of the gas turbine. 3. The method as claimed in claim 2, wherein the measuring positions, between which the pressure loss is determined, lie exclusively in the combustion chamber. 4. The method as claimed in claim 3, wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured. 5. The method as claimed in claim 4, wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula T=const**(n)·(pΔp)2·(p·pamb0p0·pamb)1,31-11,31=F(p,Δp) wherein n is the rotational speed, pamb is the ambient pressure and p0 is the pressure upstream of the turbine and pamb0 is the associated ambient pressure, both at a nominal point. 6. The method as claimed in claim 2, wherein a turbine flange is located in the gas turbine at the transition between the combustion chamber and the turbine, and wherein the measuring positions, between which the pressure loss is determined, lie exclusively in the turbine flange. 7. The method as claimed in claim 6, wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured. 8. The method as claimed in claim 7, wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula T=const**(n)·(pΔp)2·(p·pamb0p0·pamb)1,31-11,31=F(p,Δp) wherein n is the rotational speed, pamb is the ambient pressure and p0 is the pressure upstream of the turbine and pamb0 is the associated ambient pressure, both at a nominal point. 9. The method as claimed in claim 2, wherein a turbine flange is located in the gas turbine at the transition between the combustion chamber and the turbine, and wherein the measuring positions, between which the pressure loss is determined, lie both in the combustion chamber and in the turbine flange. 10. The method as claimed in claim 9, wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured. 11. The method as claimed in claim 10, wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula T=const**(n)·(pΔp)2·(p·pamb0p0·pamb)1,31-11,31=F(p,Δp) wherein n is the rotational speed, pamb is the ambient pressure and p0 is the pressure upstream of the turbine and pamb0 is the associated ambient pressure, both at a nominal point. 12. The method as claimed in claim 2, wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured. 13. The method as claimed in claim 12, wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula T=const**(n)·(pΔp)2·(p·pamb0p0·pamb)1,31-11,31=F(p,Δp) wherein n is the rotational speed, pamb is the ambient pressure and p0 is the pressure upstream of the turbine and pamb0 is the associated ambient pressure, both at a nominal point. 14. The method as claimed in claim 1, wherein in addition to the pressure loss measurement, a pressure directly upstream of a first blade row of the turbine is measured. 15. The method as claimed in claim 14, wherein the hot gas temperature is derived from the pressure loss measurement and the pressure measurement directly upstream of the first blade row of the turbine in accordance with the formula T=const**(n)·(pΔp)2·(p·pamb0p0·pamb)1,31-11,31=F(p,Δp) wherein n is the rotational speed, pamb is the ambient pressure and p0 is the pressure upstream of the turbine and pamb0 is the associated ambient pressure, both at a nominal point. 16. The method as claimed in claim 1, wherein in the pressure measurements, possible pulsation signals, which emanate from combustion chamber pulsations, are suppressed by means of one of time-based averaging and filtering. 17. The method as claimed in claim 1, wherein the controlling of the gas turbine comprises controlling the gas turbine during transient operating states of the gas turbine. 18. The method as claimed in claim 1, wherein the different spatial points are arranged in series in the flow direction. 19. The method as claimed in claim 1, wherein the pressure measurements are made through at least two different spatial points in a flow direction in a region of hot gas flow. 20. A gas turbine comprising: a compressor for compressing inducted combustion air;a combustion chamber for generating hot gas by combusting a fuel with the aid of the compressed combustion air;a multistage turbine for expanding the generated hot gas and performing work;a machine control unit;at least two pressure sensors for measuring a pressure loss in the hot gas flow, the pressure sensors being arranged at different points in series in a flow direction in the region of the hot gas flow which is generated in the combustion chamber, and the pressure sensors being connected to the machine control unit; anda turbine flange arranged in the gas turbine at a transition between the combustion chamber and the turbine,wherein the machine control unit is configured to control the gas turbine in accordance with the hot gas temperature which is derived from a plurality of other measured variables of the gas turbine,wherein a derivation of the hot gas temperature is formed exclusively by means of pressure measurements at the different points of the gas turbine, andwherein at least one of the pressure sensors is located in a region of the turbine flange. 21. The gas turbine as claimed in claim 20, wherein the pressure sensors are located in the region of the combustion chamber. 22. The gas turbine as claimed in claim 20, wherein the pressure sensors are located in the region of the turbine flange. 23. The gas turbine as claimed in claim 20, comprising: a third pressure sensor located directly upstream of a first blade row of the turbine and being connected to the machine control unit of the gas turbine. 24. The gas turbine as claimed in claim 23, wherein the pressure sensors are connected to the machine control unit of the gas turbine via a device for suppressing pulsation signals. 25. The gas turbine as claimed in claim 20, wherein the pressure sensors are connected to the machine control unit of the gas turbine via a device for suppressing pulsation signals.
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