Counter swirl annular combustor

Details Counter swirl annular combustor Counter swirl annular combustor

2002-03-21

2004-01-13

Kim, Ted (Department: 3746)

Power plants

Combustion products used as motive fluid

Coaxial combustion products generator and turbine

C060S748000, C060S752000

Reexamination Certificate

active

06675587

ABSTRACT:

TECHNICAL FIELD
This invention relates to combustors for a gas turbine and in particular to the efficient mixing of fuel and air within the combustor.
BACKGROUND OF THE INVENTION
In a typical gas turbine engine, the working medium gases are flowed into the combustor where they are mixed with fuel. The combustor provides a combustion chamber where the fuel and air mixture is burned as thoroughly as possible. In an annular combustor, the fuel is metered and injected into the combustor by multiple nozzles along with combustion air having a designated amount of swirl.
To facilitate mixing of the air and fuel mixture as the combustion gases move downstream in the combustion chamber, a plurality of cross-flow apertures are used within the outer and inner liners of the combustor. These apertures introduce additional air (air jets) into the combustion chamber downstream of the fuel nozzles. Generally, the manner in which typical annular three and two zone combustors mix the fuel and air will be mentioned below. Annular combustors employ annular rows of holes within the outer and inner liners, respectively. Referring to
FIG. 1
, in a three zone combustor (two stage air admission) having primary, intermediate and dilution zones, there is generally a first row of holes in the outer liner and a corresponding first row of holes in the inner liner. There is also a second row of holes in the outer liner and a corresponding second row of holes in the inner liner. The second rows of holes in the inner and outer liners are downstream of the first rows of holes in the outer and inner liners. The first rows of holes reduce the formation of the hot streaks while the second rows of holes facilitate an exit temperature profile acceptable to gas turbine engine rotor design. The overall length of the combustor can be reduced, which has been heretofore recognized. Such a reduction in overall length is accomplished by eliminating one of the air admission stages. Thus, a two zone combustor design (single stage air admission), having primary and dilution zones, does not employ the second rows of holes. Therefore, the air jets from the first rows of holes in the outer and inner liners cool the center of the combustor. Unfortunately, as a consequence of having only the first rows of holes, hot streaks can form along the walls of the inner and outer liners or in the gaps between the jets.
If thorough mixing is not achieved during combustion of the mixture, the result will be a non-uniform temperature variation of the combustion products as they exit the combustor. Consequently, the downstream gas turbine parts, such as the first stage turbine vanes, are subjected to localized overheating. This overheating has the effect of degrading the durability of the downstream gas turbine parts. Further, this overheating of the downstream gas turbine parts requires increased cooling air to compensate for the overheating. Consequently, this increase of cooling air supplied to downstream gas turbine parts decreases overall gas turbine efficiency.
Therefore, what is needed is a combustor apparatus that more thoroughly mixes the fuel and air mixture with the results being an enhanced uniform exit temperature distribution thus eliminating hot streaks in the turbine.
SUMMARY OF THE INVENTION
The above discussed and other drawbacks and deficiencies are overcome or alleviated by the present invention.
Accordingly, the present invention provides a combustion apparatus for a gas turbine engine with enhanced mixing of the combustion gases (fuel and air mixture) within the combustion chamber and a reduction of peak temperatures at the exit plane of the combustion chamber. Thus, the apparatus described herein provides a more uniform temperature distribution that reduces the formation of hot streaks and advantageously cools the inner and outer liners of the combustion chamber.
In accordance with the present invention, the apparatus includes an annular combustion chamber having an inner liner and an outer liner coaxially disposed relative to each other to form a combustion zone therebetween. A plurality of fuel injectors is configured to swirl the fuel and air mixture injected therefrom into the combustion zone. The inner liner includes a plurality of apertures circumferentially arranged. Similarly, the outer liner includes a plurality of apertures circumferentially arranged. The apertures are spaced apart circumferentially along the respective inner and outer liners. Further, the apertures in the outer liner are circumferentially positioned such that there is only one such aperture between each of the fuel injectors. Similarly, the apertures in the inner liner are circumferentially positioned such that there is only one such aperture between each of the fuel injectors. In this way, the air jets exhausted into the combustion chamber from the apertures located in both the inner and outer liners advantageously oppose the direction of swirl of the fuel and air mixture that is injected from each of the fuel injectors. Thus, the fuel and air mixture is thoroughly mixed with the results being an enhanced uniform exit temperature distribution thus eliminating hot streaks in the turbine.
Also, the air jets exhausted from the apertures traverses the combustion zone to impinge the respective opposing liners. In the preferred embodiment, the apertures are elliptical in shape offering enhanced air penetration into the combustion zone and enhanced mixing of the air jets with the fuel and air mixture to provide an improved exit temperature profile.
The above discussed and other features and advantages of the present invention will be appreciated and understood by those skilled in the art from the following drawings and detailed description.


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