GB2377973A

GB2377973A - Gas bleed system for a gas turbine

Info

Publication number: GB2377973A
Application number: GB0118077A
Authority: GB
Inventors: Mary Bridget Fletcher
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2001-07-25
Filing date: 2001-07-25
Publication date: 2003-01-29
Also published as: GB0118077D0

Abstract

A gas turbine engine 10 comprises an inlet 12, a compressor section 14 with a low pressure compressor 24 driven via a shaft 32 by a low pressure turbine 30 in a turbine section 18, a high pressure compressor 26 driven via a shaft 34 by a high pressure turbine 28, a combustion section 16 comprising two combustion chambers or zones (42 and 44, fig 2), a power turbine section 20 arranged to drive a load 36 for example an electrical generator via a shaft 38, and an exhaust 22. The compressor section 14 is provided with a bleed valve 60 and a controller 62 arranged to bleed air from the compressor section 14 at predetermined operating conditions through a bleed duct 64 to a position downstream of the turbine section 18 and upstream of the power turbine section 20. The air may be bled from various positions in the compressor section (Figures 3 to 5).

Description

A GAS TURBINE ENGINE The present invention relates to a gas turbine engine, particularly to an industrial or marine gas turbine, but is also applicable to an aero gas turbine engine.

In gas turbine engines emissions of nitrous oxides (NOx) increase when the temperature is too high, due to dissociation of atmospheric nitrogen, and emissions of carbon monoxide (CO) increase when the temperature is too low, due to incomplete combustion.

In recent years the combustion chambers of industrial gas turbine engines have been improved to minimise these emissions, as shown in our International patent application W09207221A, published 30 April 1992. These combustion chambers comprise a primary combustion chamber and a secondary combustion chamber arranged downstream of the primary combustion chamber. A primary fuel and air mixing duct supplies a uniform fuel and air mixture to the primary combustion zone and a secondary fuel and air mixing duct supplies a uniform fuel and air mixture to the secondary fuel and air mixing duct. This arrangement promotes spatial temperature uniformity.

Variations in power and ambient temperature are taken into account by controlling the temperature in the primary and secondary combustion zones by precise control of the fuel flow to each of the primary and secondary combustion zones as described in our International patent application W09309339A, published 13 May 1993.

For a given combustion section, of the gas turbine engine, there is a finite range of power and ambient temperature over which low emissions may be achieved. This is often referred to as the turndown. The most usual situation is high emissions of carbon monoxide (CO) at low power and/or at low ambient temperature operating conditions. A combustion chamber with a single combustion zone and a fuel and air mixing duct has, for example, a 50 K stator outlet temperature turndown. A combustion chamber

with a primary combustion zone, primary fuel and air mixing duct, secondary combustion zone and secondary fuel and air mixing duct has, for example, a 3000K stator outlet temperature turndown. This temperature range may be increased by minimising the variation in turbine stator outlet temperature versus power and ambient temperature.

Other methods may be used to increase the range of ambient temperature and power over which low emissions may be achieved.

One known method is to provide variable inlet guide vanes on the compressor section of the gas turbine engine.

However, variable inlet guide vanes are only helpful on a gas turbine engine comprising a turbine arranged to drive a compressor and a fixed speed load, for example a synchronous alternator. The variable inlet guide vanes are closed to reduce power. This reduces the compressor flow at fixed speed, which means that the stator outlet temperature reduction is minimised. For gas turbine engines with free power turbines the compressor speed is not fixed and therefore variable inlet guide vanes do not reduce emissions.

Another known method is to provide a bleed arrangement on the compressor section of the gas turbine engine and to dump the air out of the gas turbine engine. However, bleeding air from the compressor section makes the gas turbine engine run hotter. The advantage of the bleed arrangement is that it is relatively cheap and easy to provide on a gas turbine engine. The disadvantage is that there is high fuel consumption. Bleeding air from the compressor section achieves low emissions in terms of volume parts per million, which is the most common measurement unit. However, there are trends to measure emissions in tons per year or g/kWh.

Accordingly the present invention seeks to provide a novel gas turbine engine which has an increased range of

ambient temperature and power over which low emissions may be achieved.

Accordingly the present invention provides a gas turbine engine comprising a compressor section, a combustion section, a turbine section and a power turbine section, the compressor section comprising at least one compressor, the turbine section comprising at least one turbine arranged to drive the at least one compressor, the power turbine section being arranged to drive a load, means to selectively bleed gas from the compressor section and means to supply the gas bled from the compressor section to a position between the turbine section and the power turbine section.

Preferably the compressor section comprises a first compressor and a second compressor, the turbine section comprises a first turbine and a second turbine, the first turbine being arranged to drive the first compressor and the second turbine being arranged to drive the second compressor.

The means to selectively bleed gas from the compressor section may comprise means to bleed gas from a position between the first compressor and the second compressor.

The means to selectively bleed gas from the compressor may comprise means to bleed gas from a position at the downstream end of the compressor section.

The means to selectively bleed gas from the compressor may comprise means to bleed gas from a position between the upstream end and the downstream end of the first compressor.

The means to selectively bleed gas from the compressor may comprise means to bleed gas from a position between the upstream end and the downstream end of the second compressor.

Preferably the load comprises an electrical generator, a pump, a propeller or a drive shaft.

The power turbine may drive the load via a gearbox.

Preferably the combustion section comprises at least one combustion chamber, the combustion chamber comprises a primary combustion zone and a secondary combustion zone arranged downstream of the primary combustion zone, at least one primary fuel and air mixing duct to supply a mixture of fuel and air to the primary combustion zone and at least one secondary fuel and air mixing duct to supply a mixture of fuel and air to the secondary combustion zone.

The present invention will be more fully described by way of example with reference to the accompanying drawings in which :- Figure 1 shows a gas turbine engine according to the present invention.

Figure 2 shows an enlarged cross-sectional view through the combustion section of the gas turbine engine shown in figure 1.

Figure 3 shows an alternative gas turbine engine according to the present invention.

Figure 4 shows a further gas turbine engine according to the present invention.

Figure 5 shows another gas turbine engine according to the present invention.

An industrial gas turbine engine 10, as shown in figure 1, comprises in flow series an inlet 12, a compressor section 14, a combustion section 16, a turbine section 18, a power turbine section 20 and an exhaust 22.

The compressor section 14 comprises in flow series a low pressure compressor 24 and a high pressure compressor 26.

The turbine section 18 comprises in flow series a high pressure turbine 28 and a low pressure turbine 30. The low pressure turbine 30 is arranged to drive the low-pressure compressor 24 via a shaft 32 and the high pressure turbine 28 is arranged to drive the high pressure compressor 26 via a shaft 34. The power turbine 20 is arranged to drive a load, for example an electrical generator 36 via a shaft 38.

The combustion section 16 is shown more clearly in figure 2. The combustion section 16 comprises a plurality of tubular combustion chambers 40. Each combustion chamber 40 comprises a primary combustion zone 42 and a secondary combustion zone 44 arranged downstream of the primary combustion zone 42. Each combustion chamber 40 comprises a primary fuel and air mixing duct 46 and a secondary fuel and air mixing duct 48. Each primary fuel and air mixing duct 46 is arranged to supply a mixture of fuel and air into the primary combustion zone 42 of the respective combustion chamber 40 and each secondary fuel and air mixing duct 48 is arranged to supply a mixture of fuel and air into the secondary combustion zone 44 of the respective combustion chamber 40. Each primary fuel and air mixing duct 46 is provided with one or more primary fuel injectors 50 and each secondary fuel and air mixing duct 48 is provided with one or more secondary fuel injectors 52.

The compressor section 14 is provided with a bleed valve 60 arranged to bleed air from the compressor section 14. The bleed valve 60 is controlled by a controller 62 which opens the valve 60 at predetermined operating conditions of the gas turbine engine 10. The bleed valve 60 is arranged to bleed air from a position upstream of the high pressure compressor 26 and downstream of the low pressure compressor 24. The bleed valve supplies air bled from the compressor section 14 to a bleed duct 64 and the bleed duct supplies the air bled from the compressor section 14 to a position downstream of the turbine section 18 and upstream of the power turbine section 20.

In operation of the gas turbine engine 10 below a predetermined power and/or below a predetermined temperature the controller 62 opens the bleed valve 60 and air is bled off the compressor section 14. The air bled from the compressor section 14 flows through the bleed duct 64 and is supplied into the hot gases flowing from the turbine section 18 to the power turbine section 20. The

bleeding of the air from the compressor section 14 makes the gas turbine engine 10 run hotter. The bled air supplied to the power turbine section 20 reduces the expansion ratio of the high pressure turbine 28 and the low pressure turbine 30 in the turbine section 18 and hence reduces the speed of rotation of, and the airflow through, the low pressure compressor 24 and the high pressure compressor 26 in the compressor section 14, for a given fuel flow. The bled air supplied to the power turbine section 20 also does work in driving the power turbine section 20.

Thus the present invention reduces the emissions at low power and/or low temperature and in addition provides a decrease in fuel consumption, an increase in thermal efficiency. It is believed that there is a 10% reduction in fuel consumption by supplying the bled air to the power turbine section 20 compared to dumping the bled air overboard at low power conditions. Also there is a reduction in noise. At 75% power the amount of air required to be bled from the compressor section 14 and supplied to the power turbine section 20 is half of that required to be bled from the compressor section 14 and dumped overboard. This provides a 5% improvement in efficiency. At power levels less than 75% the improvement in efficiency of the present invention would be greater.

A gas turbine engine 110, as shown in figure 3, is substantially the same as that shown in figure 1 and like parts are denoted by like numerals. The gas turbine engine 110 in figure 3 differs in that the bleed valve 60 is arranged to bleed air from the downstream end of the high pressure compressor 26.

A gas turbine engine 210, as shown in figure 4, is substantially the same as that shown in figure 1 and like parts are denoted by like numerals. The gas turbine engine 210 in figure 4 differs in that the bleed valve 60 is arranged to bleed air from a position between the upstream

end and the downstream end of the high pressure compressor 26.

A gas turbine engine 310, as shown in figure 5, is substantially the same as that shown in figure 1 and like parts are denoted by like numerals. The gas turbine engine 310 in figure 5 differs in that the bleed valve 60 is arranged to bleed air from a position between the upstream end and the downstream end of the high pressure compressor 24.

The power turbine may be arranged to drive other loads for example a pump, a propeller or a drive shaft.

Any suitable type of valve may be used for the bleed valve.

Claims

Claims :- 1. A gas turbine engine comprising a compressor section, a combustion section, a turbine section and a power turbine section, the compressor section comprising at least one compressor, the turbine section comprising at least one turbine arranged to drive the at least one compressor, the power turbine section being arranged to drive a load, means to selectively bleed gas from the compressor section and means to supply the gas bled from the compressor section to a position between the turbine section and the power turbine section.
2. A gas turbine engine as claimed in claim 1 wherein the compressor section comprises a first compressor and a second compressor, the turbine section comprises a first turbine and a second turbine, the first turbine is arranged to drive the first compressor and the second turbine is arranged to drive the second compressor.
3. A gas turbine engine as claimed in claim 2 wherein the means to selectively bleed gas from the compressor section comprises means to bleed gas from a position between the first compressor and the second compressor.
4. A gas turbine engine as claimed in claim 1 or claim 2 wherein the means to selectively bleed gas from the compressor comprises means to bleed gas from a position at the downstream end of the compressor section.
5. A gas turbine engine as claimed in claim 1 or claim 2 wherein the means to selectively bleed gas from the compressor comprises means to bleed gas from a position between the upstream end and the downstream end of the first compressor.
6. A gas turbine engine as claimed in claim 1 or claim 2 wherein the means to selectively bleed gas from the compressor comprises means to bleed gas from a position between the upstream end and the downstream end of the second compressor.

<Desc/Clms Page number 9>
7. A gas turbine engine as claimed in any of claims 1 to 6 wherein the load comprises an electrical generator, a pump, a propeller or a drive shaft.
8. A gas turbine engine as claimed in any of claims 1 to 7 wherein the power turbine drives the load via a gearbox.
9. A gas turbine engine as claimed in any of claims 1 to 8 wherein the combustion section comprises at least one combustion chamber, the combustion chamber comprises a primary combustion zone and a secondary combustion zone arranged downstream of the primary combustion zone, at least one primary fuel and air mixing duct to supply a mixture of fuel and air to the primary combustion zone and at least one secondary fuel and air mixing duct to supply a mixture of fuel and air to the secondary combustion zone.
10. A gas turbine engine substantially as hereinbefore described with reference to figure 1 of the accompanying drawings.
11. A gas turbine engine substantially as hereinbefore described with reference to figure 3 of the accompanying drawings.
12. A gas turbine engine substantially as hereinbefore described with reference to figure 4 of the accompanying drawings.
13. A gas turbine engine substantially as hereinbefore described with reference to figure 5 of the accompanying drawings.