This application claims priority to German Patent Application DE102012001777.4 filed Jan. 31, 2012, the entirety of which is incorporated by reference herein.
This invention relates to an annular combustion chamber of a gas turbine. An annular combustion chamber of this type has an upper/radially outer combustion chamber wall and a lower/radially inner combustion chamber wall that together form an annular duct. Air and fuel are supplied to the combustion chamber by the fuel nozzle, and air is also supplied by cooling or air inlet openings on the side walls. Air and fuel are mixed and combusted in the fuel nozzle and in the combustion chamber respectively. The air and the combustion products are passed through the combustion chamber outlet nozzle in the direction of the turbine.
The existing combustion chamber geometries are disadvantageous in that the geometries have weaknesses in terms of the flow guidance of the air. For example, the side wall geometries and the area cross-sections along the combustion chamber axis are not optimally designed from the aerodynamics viewpoint, with the result that the flow is not routed through the combustion chamber in an optimum manner in terms of losses, and flow separations/boundary layer accumulations and wake zones can result close to the combustion chamber walls. This can have a negative effect on the mixing of air and fuel and hence also on flame formation, flame stability and fuel combustion, with the result that emissions of the combustion chamber can be negatively influenced.
The object underlying the present invention is to provide a gas turbine annular combustion chamber of the type specified at the beginning which, while being simply designed and easily and cost-effectively producible, avoids the disadvantages of the state of the art and is characterized by good flow conditions and a high efficiency.
The annular combustion chamber in accordance with the invention is therefore designed such that the respective central axes of the fuel nozzles form an envelope which is rotationally symmetrical to the engine axis and which divides the combustion chamber into an annular and radially outer area and an annular and radially inner area, with the radially outer area and the radially inner area having the same volumes.
The solution in accordance with the invention thus provides an annular combustion chamber in which the air/fuel flows are evenly distributed in the radial direction. Since the central axes of the fuel nozzles for the respective flows leaving the fuel nozzles form a central axis or symmetry axis, these flows are now symmetrically structured in particular in the radial direction. They are not affected by unsuitable combustion chamber wall geometries. It is thus possible to achieve largely undisrupted flow conditions and hence undisrupted combustion conditions, which in turn lead to improved operating conditions. The result in accordance with the invention is a better mixing of fuel and air, air guidance with lower losses inside the combustion chamber, better cooling efficiency, better flame stability, better burn-out and lower emissions.
The present invention thus provides that the design of the side wall geometry is based on the provision of a symmetrical annular combustion chamber which has identical areas relative to the axis of the fuel nozzle.
The following is therefore provided in accordance with the invention:
As mentioned, the combustion chamber centerline M can be arranged at an angle α relative to the engine axis, but it is also possible to align it parallel to the engine axis.
The present invention is described in the following in light of the accompanying drawing, showing an exemplary embodiment. In the drawing,
The gas-turbine engine 10 in accordance with
The intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20, generally referred to as stator vanes and projecting radially inwards from the engine casing 21 in an annular flow duct through the compressors 13, 14. The compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17, respectively.
The turbine sections 16, 17, 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16, 17, 18, and a subsequent arrangement of turbine blades 24 projecting outwards from a rotatable hub 27. The compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.
The length L of the annular combustion chamber results from structural and design requirements, in particular with regard to the necessary flow length and to the flame geometry and ignitability. The respectively necessary areas A are obtained by way of analogy from the design and physical requirements.
Number | Date | Country | Kind |
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10 2012 001 777 | Jan 2012 | DE | national |
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Entry |
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German Search Report dated Sep. 11, 2012 from counterpart application. |
Number | Date | Country | |
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20130192232 A1 | Aug 2013 | US |