Patent Abstract:
A hybrid combustor for a gas turbine engine includes a plurality of circularly arrayed ceramic can combustors whose outlets communicate with the inlet of an annular, metal combustor. The combustion process is continuous through the plurality of can combustors and into the single annular combustor. Preferably only fuel-rich combustion occurs within each of the can combustors, and fuel-lean combustion continues within the single annular combustor.

Full Description:
TECHNICAL FIELD 
     This invention pertains to combustors for gas turbine engines, and pertains more particularly to an improved hybrid combustor incorporating the ceramic can combustors and a metallic annular combustor. 
     1. Background of the Invention 
     Gas turbine engine efficiency increases with increased temperature. To this end, it has been proposed to utilize ceramic components within gas turbine engines, particularly at the highest temperature locations therein, to increase gas turbine engine maximum temperatures. Utilization of ceramics, such as ceramic matrix composites, in the combustor of the gas turbine engine is therefore highly desirable. 
     However, ceramic material such as ceramic matrix composites are sensitive to the temperature difference through the thickness of the material. The temperature difference between the hot interior and the cooler exterior generate thermal stresses resulting in cracking of the ceramic matrix. This limits the allowable wall thickness of the design making it difficult to produce a conventional annular ceramic combustor configuration of a reasonably large diameter which needs larger wall thickness to withstand the buckling pressures associated with the larger diameters. Ceramic designs are thus limited by small diameter, low pressure drop, low heat loading, or a reduced combination of such factors, which ultimately limit the combustor performance. 
     2. Summary of the Invention 
     Accordingly, it is an important object of the present invention to provide an improved combustor for a gas turbine engine which utilizes ceramic materials in a geometric configuration which avoids the problems normally associated with such use of ceramics. More particularly, it is an important object of the present invention to provide a hybrid combustor having a plurality of can-type ceramic combustors disposed in a circular array, along with a conventional metallic annular combustor construction. summary, the present invention contemplates a plurality of ceramic can combustors each having a cylindrical ceramic wall, wherein primary, fuel-rich combustion occurs, along with a single annular, metallic combustor which receives the exhaust of the fuel-rich burn from all of the can combustors, along with pressurized air flow from the combustor inlet. Fuel-lean combustion continues to occur in the annular metallic combustor as a continuation of the fuel-rich combustion process in each of the can combustors. In this manner the ceramic cylindrical walls of the can combustors can be made of relatively small diameter to minimize thermal stresses and buckling forces thereon. 
     These and other objects and advantages of the present invention are specifically set forth in or will become apparent from the following detailed description of a preferred embodiment of the invention when read in conjunction with the accompanying drawings. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a schematic, perspective representation of a hybrid combustion constructed in accordance with the principles of the present invention; 
     FIG. 2 is a cross-sectional plan view of the hybrid combustor of the present invention; and 
     FIG. 3 is a front elevational view of a portion of the combustor of the present invention. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
     Referring now more particularly to the drawings, a gas turbine engine combustor  10  generally includes a plurality of can combustors  12  disposed in a circular array about the central axis  14  of an associated annular combustor  16 . As best depicted in FIG. 2, the gas turbine engine combustor  10  includes an annular outer casing  18  having a pressurized air inlet  20 , an exhaust  22 , and a fuel supply duct  24  leading to a fuel nozzle  26  associated with each of the can combustors  12 . Each fuel nozzle  26  in conventional fashion receives air for primary combustion from the pressurized air inlet as illustrated by arrows  28 , and may include a primary swirler  30  (FIG. 1) so as to deliver a finely mixed mixture of fuel and air into the primary combustion zone within each of the can combustors  12 . 
     Each can combustor  12  includes a cylindrical outer metal liner  32  and a continuous cylindrical inner ceramic wall  34 . For fuel-rich can combustors, the ceramic wall  34  is preferably non-perforated. Preferably the ceramic wall  34  is made of a ceramic matrix composite material. If desired, metal supports  36  may extend radially inwardly from the outer metal wall liner  32  to position the ceramic wall  34  centrally therewithin without inducing thermal stresses on the ceramic wall  34 . Defined between outer metal liner  32  and inner ceramic wall  34  is a ring-shaped, annular air space  40  extending axially along the can  12 . At the inlet end, the outer metal liner  32  extends radially inwardly to the fuel nozzle  26 . A floating metal grommet  42  effectively seals between and intersecures the outer metal liner  12  with the fuel nozzle  26 . As best depicted in FIG. 3, the inlet end of the outer liner  32  includes a plurality of inlet air passages  44  disposed in a full circular array for allowing pressurized air from the inlet  20  to enter the annular air space  40  for axial flow therealong on the exterior side of the ceramic wall  34 . 
     Annular metal combustor  16  conventionally includes inner and outer metal walls  44 ,  46  disposed in an annular configuration normally surrounding the turbine section of the gas turbine engine. As desired, the metal walls  44 ,  46  may have small openings  48  therein for film or effusion cooling of the metal walls  44 ,  46 . 
     The inlet end of annular combustor  16  includes a plurality of relatively large openings  49  each of which receives the corresponding exhaust end of the associated can combustor  12 . Outer metal liner  32  of each can combustor is rigidly secured to the annular combustor walls  44 ,  46  such as by a plurality of welded brackets  50 . Accordingly, each of the can combustors  12  is rigidly secured to the annular combustor  16  through associated metal liner  32 . The annular air passage  40  of each can combustor  12  opens into the inlet of the annular combustor  16 , as depicted by arrows  52 , to inject pressurized air received from inlet  20  directly in to the annular combustor  16  to support secondary combustion therein as described in greater detail below. In conventional fashion, the outlet end of the annular combustor  16  is appropriately secured to the combustor casing  18  for delivery of hot combustion products through the exhaust  22 . 
     In operation, pressurized air inlet flow from the compressor section of the gas turbine engine is delivered through air inlet  20  inside the annular outer combustor casing  18  in a generally axial direction. Fuel is delivered through each fuel nozzle  26  to mix with air for primary combustion to be delivered in to the interior of each can combustor  12 . Primary combustion occurs inside the ceramic wall  34  of each can combustor  12 . Preferably this is a fuel-rich burn combustion process inside each ceramic can combustor  12 . If transition to fuel-lean combustion is desired in the can combustors  12 , openings along the length of wall  34  may be included instead of the nonperforated configuration shown. 
     To minimize thermal stress across the ceramic wall  34 , its thickness is minimized. Minimization of the thickness of ceramic wall  34  reduces the temperature differential thereacross and therefore minimizes the thermal stresses imposed thereon. Additionally, the annular air passage  40  through which pressurized air flow is delivered provides cooling to the ceramic can  34  and the outer liner  32  to maintain material temperatures of both components within acceptable ranges. It is because of the necessity to minimize the thickness of the ceramic wall  34  that makes it unacceptable for use as a relatively large annular combustor, since the necessary thinness of the wall would subject it to buckling. 
     The combustion process inside each can combustor  12  continues throughout the axial length thereof and through the openings  49  into the annular combustor  16 . That is, the flame front created in the primary combustion zone within each can combustor  12  extends through the associated opening  49  and into the interior of the annular combustor  16 . 
     Significant pressurized air flow is injected into the annular combustor  16  through the annular air passage  40  as depicted by arrows  52  in FIG.  2 . The combustion process initiated in each of the can combustors continues within the annular combustor  16  with secondary, fuel-lean combustion occurring therewithin. Because the annular combustor is a continuous, circular configuration, the combustion process therewithin expands circumferentially into a continuous, ring-like combustion front. In this manner, the present invention provides all of the attendant advantages associated with conventional annular combustors, and in particular the elimination of thermal patterning therein. As noted, fuel-lean secondary combustion continues within the annular combustor  16  until the combustion process is completed therewithin. The exhaust products from the combustor  10  are delivered through exhaust  22  to drive the turbine section of the gas turbine engine. 
     Various alterations and modifications to the foregoing detailed description of a preferred embodiment of the invention will be apparent to those skilled in the art. Accordingly, the foregoing should be considered exemplary in nature and not as limiting to the scope and spirit of the invention as set forth in the appended claims.

Technology Classification (CPC): 5