Power plants – Combustion products used as motive fluid
Reexamination Certificate
2000-12-11
2003-03-25
Freay, Charles G. (Department: 3746)
Power plants
Combustion products used as motive fluid
C060S755000, C060S757000
Reexamination Certificate
active
06536201
ABSTRACT:
THE FIELD OF THE INVENTION
The present invention relates generally to a turbine cooling structure in a gas turbine engine and more specifically to an improved configuration of a combustor/turbine successive dual cooling arrangement.
BACKGROUND OF THE INVENTION
In a conventional gas turbine engine comprising a compressor, combustor and turbine, both the combustor and the turbine require cooling due to beating thereof by hot combustion gases.
Within the combustor, fuel fed through the fuel nozzle is mixed with compressed air provided by the high pressure compressor and ignited to drive turbines with the hot gases emitted through the combustor. Within the metal combustor, the gases burn at approximately 3,500° to 4,000° Fahrenheit. The combustion chamber is fabricated of a metal which can resist extremely high temperatures. However, even highly resistant metal will melt at approximately 2,100° to 2,200° Fahrenheit. Therefore, it is important to adequately cool the hot combustor wall of a gas turbine engine for safe engine operation.
As is well known in the art, the combustion gases are prevented from directly contacting the material of the combustor through use of a cool air film which is directed along the internal surfaces of the combustor. The combustor has a number of louver openings through which compressed air is fed parallel to the hot combustor walls. Eventually the cool air curtain degrades and is mixed with the combustion gases. However, in such air film cooling arrangements, the cooling air mixed with the combustion gases increases CO emissions. Thus, while cooling techniques used on the combustor liner may be advantageous in increasing maximum engine temperature, they deleteriously increase CO formation and emission.
The use of air film cooling is limited by the amount of air available exclusively for cooling the combustor wall. Generally, as the amount of cooling air is increased to cool the engine components, the amount of air available for the combustor is decreased, which results in increasing NO
x
formation and emission.
Efforts have been made to cool the combustor wall of a gas turbine engine while avoiding the increase of emissions. For example, U.S. Pat. No. 5,687,572, issued to Schrantz et al. on Nov. 18, 1997, discloses a combustor for a gas turbine engine having a porous outer metallic shell and a thin-walled, nonporous ceramic liner the backside of which is impingement cooled. All air flow used for impingement cooling is re-injected into the combustion process itself, preferably, primarily in the dilution zone of the combustion process so that there is no loss of pressurized air flow from a thermodynamic standpoint, which is advantageous to reduce NO
x
formation, and also no film cooling on the interior surface of the ceramic liner is introduced to induce CO formation.
In another example, U.S. Pat. No. 5,758,504 issued to Abreu et al. on Jun. 2, 1998 discloses a combustor construction including an interior liner having a plurality of angled holes extending therethrough, arranged in a pre-established pattern defining a centroid, and an exterior liner having a plurality of holes extending therethrough at about 90 degrees. At least a portion of the holes in the exterior liner are radially aligned with the centroid of the holes in the interior to reduce the use of cooling air flow per unit length of the combustor wall, thereby resulting in reduction of CO emissions.
In addition to the combustor cooling, in a turbine section of a gas turbine engine, both rotating turbine components such as blades, disks and retainers, and stationary turbine components such as vanes, shrouds and frames are directly exposed to high temperature combustion gases discharged from the combustor and routinely require cooling. Cooling of the turbine, especially the rotating components, is critical to the proper function and safe operation of the engine. Failure to adequately cool a turbine disk and its blades, for example, by providing cooling air deficient in supply pressure, volumetric flow rate or temperature margin, may be detrimental to the life and mechanical integrity of the turbine.
Balanced with the need to adequately cool the turbine is the desire for high levels of engine operating efficiency, which translate into lower fuel consumption and lower operating costs. Since turbine cooling air is typically drawn from one or more stages of the compressor and channeled by various means such as pipes, ducts and internal passageways to the desired components, such air is not available to be mixed with the fuel, ignited in the combustor and undergo work extraction in the primary gas flow path of the turbine, total cooling flow bled from the compressor is therefore treated as a parasitic loss in the engine operating cycle, it being desirable to keep such loses to a minimum.
Efforts have been made to minimize compressor bleed and concomitant cycle losses, for example, by attempting to control bleed source or cooling circuit parameters, such as source pressure, pressure drop, flow rate or temperature. One example is disclosed in U.S. Pat. No. 5,555,721 issued to Bourneuf et al. on Sep. 17, 1996. Burneuf et al. describe a turbine cooling supply circuit for a gas turbine engine in which the flow of coolant through the engine is directed to minimize temperature rise prior to discharge into the turbine. In addition to being used for combustion, compressor discharge pressurized air, which is disposed within a combustor casing, is utilized to cool components of the turbine section subject to the hottest combustion gases, namely the stage one nozzle, a stage one shroud and the stage one disk. Additional bleed sources for turbine cooling air include an impeller tip forward bleed flow and impeller tip aft bleed flow which are provided to additionally cool the stage two nozzle and stage two shroud respectively, as well as other turbine components. Bourneuf et al. do not address the cooling of the combustor wall and it would be understood from the drawings attached thereto that a film cooling arrangement is intended to be used.
It has been realized that directing air for cooling, rather than combustion control, limits the degree of combustion emission optimization, and the minimization of the amount of combustor cooling is critical to the design of a state of the art low emission gas turbine combustion system. Therefore, there have been continuous efforts in the industry to develop combustor/turbine cooling apparatus and methods for low emission gas turbine engines.
SUMMARY OF THE INVENTION
It is one object of the present invention to provide a low emission gas turbine combustion system using an improved cooling method.
It is another object of the present invention to provide a cooling system for a gas turbine engine to significantly reduce the coolant volume in combustor liner cooling.
It is a further object of the present invention to provide a combustor/turbine successive dual cooling to permit all the air typically used to cool the hot end of the engine downstream of the combustor to be used as combustor cooling as well.
In general terms, a method for cooling a gas turbine engine combustor and turbine section comprises, providing a structure; enabling pressurized cooling air to form air flow impingement on an outer surface of a combustor wall for backside cooling of the combustor wall; directing the air flow immediately upon the impingement thereof along the outer surface of the combustor wall, downstream towards a turbine section for further cooling the combustor wall; and providing an access to exhaust combustor backside cooling air flow for cooling the turbine section.
In accordance with one aspect of the present invention, a cooling apparatus for a gas turbine engine having a combustor comprises a wall adapted to be attached to a combustor wall of a gas turbine engine. The wall is in a spaced apart and substantially parallel relationship with respect to an outer surface of the combustor wall to form an air passage between the wall and the combustor wall for conducting cooling air to cool
Acton Jeffrey Scott
Cowburn David Edwin
Fischer Bernhard
Fryer Michael Andrew
Kojovic Aleksandar
Bachman & LaPointe P.C.
Pratt & Whitney Canada Corp.
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