Patent Application
20120304654
This invention relates to internal cooling within a gas turbine; and more particularly, to an apparatus for providing better and more uniform cooling in a combustion liner of the turbine.
Traditional gas turbine combustors use diffusion (i.e., non-premixed) combustion in which fuel and air enter the combustion chamber separately. The process of mixing and burning produces flame temperatures exceeding 3900° F. Since conventional combustors and/or transition pieces having liners are generally capable of withstanding a maximum temperature on the order of only about 1500° F. for about ten thousand hours (10,000 hrs.), steps to protect the combustor and/or transition piece must be taken. This has typically been done by film-cooling which involves introducing relatively cool compressor air into a plenum formed by the combustor liner surrounding the outside of the combustor. In this prior arrangement, the air from the plenum passes through louvers in the combustor liner and then passes as a film over the inner surface of the liner, thereby maintaining combustor liner integrity.
Because diatomic nitrogen rapidly disassociates at temperatures exceeding about 3000° F. (about 1650° C.), the high temperatures of diffusion combustion result in relatively large NOx emissions. One approach to reducing NOx emissions has been to premix the maximum possible amount of compressor air with fuel. The resulting lean premixed combustion produces cooler flame temperatures and thus lower NOx emissions. Although lean premixed combustion is cooler than diffusion combustion, the flame temperature is still too hot for prior conventional combustor components to withstand.
Furthermore, because the advanced combustors premix the maximum possible amount of air with the fuel for NOx reduction, little or no cooling air is available, making film-cooling of the combustor liner and transition piece premature at best. Nevertheless, combustor liners require active cooling to maintain material temperatures below limits. In dry low NOx (DLN) emission systems, this cooling can only be supplied as cold side convection. Such cooling must be performed within the requirements of thermal gradients and pressure loss. Thus, means such as thermal barrier coatings in conjunction with “backside” cooling have been considered to protect the combustor liner and transition piece from destruction by such high heat. Backside cooling involved passing the compressor discharge air over the outer surface of the transition piece and combustor liner prior to premixing the air with the fuel.
With respect to the combustor liner, one current practice is to convectively cool the liner, or to provide continuous linear turbulators on the exterior surface of the liner. The continuous liner turbulators are evenly spaced and non-interrupted. The various known techniques enhance heat transfer but with undesirable effects on thermal gradients and pressure losses. Turbulators work by providing a blunt body in the flow which disrupts the flow creating shear layers and high turbulence to enhance heat transfer on the surface, but they also increase pressure drop which is undesirable.
A low heat transfer rate from the liner can lead to high liner surface temperatures and ultimately loss of strength. Several potential failure modes due to the high temperature of the liner include, but are not limited to, spallation of the thermal barrier coating, cracking of the aft sleeve weld line, bulging and triangulation. These mechanisms shorten the life of the liner, requiring replacement of the part prematurely.
Accordingly, there remains a need for enhanced levels of active cooling with minimal pressure losses at higher firing temperatures than previously available while extending a combustion inspection interval to decrease the cost to produce electricity.
According to one aspect of the present invention, a combustor for a turbine is provided. The combustor includes a plurality of fuel nozzles and a combustion zone is aligned with a combustion process associated with each of the fuel nozzles. A combustion liner includes a plurality of turbulator groups, and each of the turbulator groups has or more individual turbulators. Each of the turbulator groups is aligned with a hot streak caused by the combustion zone associated with the fuel nozzle. Each of the turbulator groups are circumferentially spaced from a neighboring turbulator group.
According to another aspect of the present invention, a combustor for a turbine is provided. The combustor has a plurality of fuel nozzles and a combustion zone is aligned with a combustion process associated with each of the fuel nozzles. A combustion liner includes a plurality of turbulator groups, and each of the turbulator groups has one or more individual turbulators. The turbulator groups are substantially aligned with a hot streak in the combustion liner caused by the combustion zone associated with the fuel nozzles. Each of the turbulator groups are circumferentially spaced from a neighboring turbulator group.
These and other features will become apparent from the following detailed description, which, when taken in conjunction with the annexed drawings, where like parts are designated by like reference characters throughout the drawings, and disclose embodiments.
With reference to
Still referring to
Hot gases from the combustion section in combustion liner 12 flow therefrom into section 16. There is a transition region indicated generally at 46 in
The hot streaks 420 generally contain hotter temperatures than the surrounding regions not included in the hot streak regions (e.g., the regions between hot streaks 420). Further, each individual hot streak region will contain sub-regions or areas of various temperatures. Accordingly, an improved turbulator configuration is proposed to cool these hot streak regions more effectively while reducing pressure drop over the combustion liner 400.
A first group of turbulators 430 is aligned with a hot streak or combustion zone of a fuel nozzle, while a second group of turbulators 440 is aligned with another combustion zone (or hot streak) associated with a different fuel nozzle. Each individual turbulator may comprise a raised rib or raised portion having any desired shape for the specific application. The regions between the hot streaks do not have the turbulators 430, 440, and this feature reduces pressure drop in areas where turbulators are not required, and provides a more uniform circumferential temperature profile that reduces the global/overall liner stress. The first group of turbulators 430 may contain turbulators having variable axial spacing. For example, a turbulator sub-group 431 contains multiple turbulators having an axial spacing of L1, a turbulator sub-group 432 contains multiple turbulators having an axial spacing of L2, and a turbulator sub-group 433 contains multiple turbulators having an axial spacing of L3. As shown, L3 is greater than L1, and L1 is greater than L2.
In this example, the hottest portion of the hot streak 420 is covered by the turbulator sub-group 432, a medium temperature portion of the hot streak is covered by the turbulator sub-group 431 and the coolest part of the hot streak is covered by turbulator sub-group 433. It can be seen that the turbulators may be configured to have the closest axial spacing in hotter regions, while cooler hot streak regions may have turbulators with a greater axial spacing. In addition, each group and/or sub-group of turbulators may be circumferentially spaced from a neighboring group of turbulators. For example, the first sub-group of turbulators 431 may be circumferentially spaced by a distance C1 from the second sub-group of turbulators 441. Each sub-group may also have substantially the same or a different circumferential spacing between a neighboring turbulator sub-group. Turbulator sub-group 441 may be spaced substantially the same or a different circumferential distance away from the sub-group turbulators 431, and sub-group turbulators 442 may be spaced the same or a different circumferential distance away from the sub-group turbulators 432. Further, each individual turbulator in a single subgroup may have variable axial spacing from adjacent individual turbulators in the same sub-group.
An advantage of this configuration is that the hottest regions of the hot streaks have greater cooling by the use of closely spaced turbulators, while cooler regions require less cooling and can employ turbulators having a greater axial spacing. Another advantage is that pressure drop is increased the most only in regions with the greatest cooling needs (e.g., the area covered by turbulators 432), and other areas have reduced pressure drop due to fewer turbulators or the presence of no turbulators (e.g., the regions between hot streaks 420).
The turbulators 532 may be located in the hottest or highest temperature portion of the hot streak, while the turbulators 533 may be located in a cooler or lower temperature portion of the hot streak. The turbulators 531 may be located in a portion of the hot streak having a temperature between the areas covered by turbulators 532 and 533. This configuration limits the maximum pressure drop to only those areas having the highest temperatures, and reduces the pressure drop for other areas of the hot streak and reduces the pressure drop even further for portions of the combustion liner outside the hot streaks.
The increased height H2 of the turbulators 632 can help to further cool the hotter portions of the combustion liner in the hotter portions of the hot streak, by increasing turbulence to thereby increase heat transfer. In some applications or in some regions of the hot streak, it may be desirable to increase the height of at least some of the individual turbulators as well as the inter-turbulator axial spacing distance. In medium temperature regions, a medium height H1 may be used, while in cooler regions of the hot streak a lower height H3 may be used for inducing turbulence.
It can be seen that an increase in turbulation (and hence heat transfer) and a reduction in overall pressure drop can be obtained by circumferentially spacing groups of turbulators on a combustion liner in a gas turbine. A group of turbulators is substantially aligned with a hot streak associated with the combustion products of a fuel nozzle, and individual sub-groups of turbulators may have various heights and/or axial spacing between neighboring turbulators.
It is noted that the terms “first,” “second,” and the like, as well as “primary,” “secondary,” and the like, herein do not denote any amount, order, or importance, but rather are used to distinguish one element from another, and the terms “a” and “an” herein do not denote a limitation of quantity, but rather denote the presence of at least one of the referenced item. As used herein the term “about”, when used in conjunction with a number in a numerical range, is defined being as within one standard deviation of the number “about” modifies. The suffix “(s)” as used herein is intended to include both the singular and the plural of the term that it modifies, thereby including one or more of that term (e.g., the turbulator includes one or more turbulators).
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
