US20110303390A1

US20110303390A1 - Combustion Chamber Cooling Method and System

Info

Publication number: US20110303390A1
Application number: US12/815,024
Authority: US
Inventors: James Oakley; David William Artt; Stephen William Thomas Spence
Original assignee: Vykson Ltd
Current assignee: Vykson Ltd
Priority date: 2010-06-14
Filing date: 2010-06-14
Publication date: 2011-12-15

Abstract

A cooling system for use with a gas turbine combustion assembly having a flame tube and a surrounding air delivery sleeve, the system comprising a vessel shaped and dimensioned to surround the air delivery sleeve and adapted for fluid flow there through. A method of cooling a combustion assembly includes the step of directing a flow of fluid over an outer surface of the delivery sleeve.

Description

BACKGROUND

The present disclosure concerns a combustion chamber cooling method and system, and in particular a method and system adapted for use with a gas turbine operable to run on low calorific value gas fuel, for example a “low quality” gas typically containing 20% methane and 80% CO2.
When gaseous fuels are used in a gas turbine engine, the fuel gas must be compressed to a pressure above the engine cycle pressure, so that the gas can be injected into the engine combustion chamber. When low calorific gases are used, the volume of gas required to produce the necessary heat release becomes very large. It is, therefore, beneficial to use a turbo compressor driven by a turbine, the high pressure gas used to power the turbine being extracted from the engine cycle.
Air from the engine's compressor flows into an annulus between a flame tube and an outer sleeve surrounding the flame tube. This high pressure air enters the same tube through rows of holes at various points along the length of the flame tube. The fuel gas is injected into the flame tube where it mixes with the compressor delivery air and burns, producing hot high pressure gas which in turn drives the engines turbine.
If the air entering the combustion chamber from the engines compressor is at a moderate temperature, (for example 120° C.), it will provide an insulating layer between the flame tube and the surface of the outer sleeve, keeping the outer sleeve relatively cool. However, if the air entering the combustion chamber is at a high temperature (for example approximately 500° C.), for example if the engine is operating on a recuperated cycle, the outer sleeve will no longer be cooled or sufficiently cooled, and will become very hot. This causes a large reduction in the strength of the material forming the sleeve, and produces excessive heat loss from the engine, thus reducing the thermal efficiency of the engine.

BRIEF SUMMARY

The present invention has been developed with a view to mitigating the above-mentioned problems of the prior art.
The present disclosure therefore provides a method for cooling a gas turbine combustion assembly having a flame tube and a surrounding air delivery sleeve, the method comprising the step of directing a flow of fluid over an outer surface of the delivery sleeve.
Preferably, the method comprises confining the fluid flow along a substantially annular vessel surrounding the delivery sleeve.
Preferably, the method comprises supplying air for cooling the delivery sleeve by bleeding air from a supply for the air delivery sleeve.
Preferably, the method comprises utilizing the air which has cooled the delivery sleeve to drive a compressor for the fuel gas being fed to the flame tube.
According to a second aspect of the disclosure, there is provided a cooling system for use with a gas turbine combustion assembly having a flame tube and a surrounding air delivery sleeve, the system comprising a vessel shaped and dimensioned to surround the air delivery sleeve and adapted for fluid flow there through.
Preferably, the vessel is substantially annular in form and comprises a fluid inlet and a fluid outlet.
Preferably, the system comprises means for directing air from a supply for the air delivery sleeve to the fluid inlet.
Preferably, the system comprises means for directing air from the fluid outlet to a turbine/compressor assembly for compressing the fuel gas to be supplied to the flame tube.

DESCRIPTION OF THE DRAWINGS

The present disclosure will now be described with reference to the accompanying drawings, in which:

FIG. 1 illustrates a sectioned side elevation of a prior art gas turbine combustion assembly, consisting of an inner flame tube and an outer air delivery sleeve;

FIG. 2 illustrates a sectioned side elevation of a cooling system according to the present disclosure, for use with a gas turbine combustion assembly; and

FIG. 3 illustrates a schematic representation of the cooling system illustrated in FIG. 2, as part of a larger gas turbine assembly.

DETAILED DESCRIPTION

Referring now to the accompanying drawings, there is illustrated a cooling system, generally indicated as 10, for use with a conventional gas turbine combustion assembly as illustrated in FIG. 1, which comprises an inner flame tube F which is fed with gaseous fuel via a feed pipe P, the flame tube F being surrounded by an outer sleeve S into which pressurized air is supplied, and injected into the flame tube F via an array of holes in the outer wall of the flame tube F.
Referring now in particular to FIGS. 2 and 3, the system 10 is illustrated in combination with the conventional flame tube F and outer sleeve S, the system 10 comprising a vessel 12 which is shaped and dimensioned to be concentrically located about the outer sleeve S, in order to substantially enclose the outer sleeve S. The system 10 further comprises a fluid inlet 14 communicating with the vessel 12 at one end thereof, and a fluid outlet 16, preferably communicating with the vessel 12 at an opposed end thereof, although it will be appreciated that the fluid inlet 14 and outlet 16 need not be disposed in the positions illustrated. This arrangement provides an annular enclosure about the outer sleeve S through which fluid can be pumped by means of the fluid inlet 14 and fluid outlet 16. In this way, a cooling fluid, preferably air, can be pumped through the vessel 12, in order to withdraw heat from the outer sleeve S, thereby maintaining same at an acceptable working temperature.
In order to further improve the performance of the system 10, and therefore the overall efficiency of any gas turbine (not shown) to which the system 10 is fitted, the air to be pumped through the vessel 12 is preferably withdrawn directly from the compressed air which feeds the outer sleeve S, via a bleed line 18 which taps into the relatively cool (approximately 200° C.) air supply, and feeds same directly to the fluid inlet 14. In addition, the air exiting the vessel 12, via the fluid outlet 16, will have been heated to a moderate temperature, which is then fed, via an exhaust line 20, to a secondary turbine 22, through which the heated air is expanded. This secondary turbine 22 drives a compressor 24, to which is fed the fuel gas for supply to the flame tube F, which is therefore compressed by the compressor 24. The compressed fuel gas is then fed via the feed pipe P into the flame tube F. Thus the heat withdrawn from the outer sleeve S via the system 10 is returned to the cycle by using the heated exhausted air from the vessel 12 to compress the fuel gas being fed to the flame tube F.
The system 10 of the present disclosure thus enables the temperature of a combustion assembly used in, for example, a recuperated cycle, to be maintained within a working temperature range, and to use the heat removed from the combustion assembly during cooling to be recycled in compressing the fuel gas supplied to the combustion assembly.
The present invention is not limited to the embodiment described herein, and which may be amended or modified without departing from the scope of the present invention.

Claims

1. A method for cooling a gas turbine combustion assembly having a flame tube and a surrounding air delivery sleeve, the method comprising the step of directing a flow of fluid over an outer surface of the delivery sleeve.

2. A method of claim 1 further comprising confining the fluid flow along a substantially annular vessel surrounding the delivery sleeve.

3. A method of claim 1 further comprising supplying air for cooling the delivery sleeve by bleeding air from a supply for the air delivery sleeve.

4. A method of claim 3 comprising the further step of utilizing the air that has cooled the delivery sleeve to drive a compressor for the fuel gas being fed to the flame tube.

5. A cooling system for use with a gas turbine combustion assembly having a flame tube and a surrounding air delivery sleeve, the system comprising a vessel shaped and dimensioned to surround the air delivery sleeve and adapted for fluid flow there through.

6. A cooling system of claim 5, wherein the vessel is substantially annular in form and comprises a fluid inlet and a fluid outlet.

7. A cooling system of claim 5, wherein the system comprises means for directing air from a supply for the air delivery sleeve to the fluid inlet.

8. A cooling system of claim 5, wherein the system comprises means for directing air from the fluid outlet to a turbine/compressor assembly for compressing the fuel gas to be supplied to the flame tube.

9. A cooling system for use with a gas turbine combustion assembly having a flame tube and a surrounding air delivery sleeve, the system comprising a substantially annular vessel concentrically surrounding the air delivery sleeve and adapted for fluid flow there through, and a means for directing air, wherein the vessel includes a fluid inlet and a fluid outlet.