Patent Abstract:
The present invention provides in one embodiment a unique gas turbine engine. Another embodiment is a unique gas turbine engine combustion system. Still another embodiment is a unique gas turbine engine combustor. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine combustion systems and combustors.

Full Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     The present application claims the benefit of U.S. Provisional Patent Application 61/290,744, filed Dec. 29, 2009, and is incorporated herein by reference. 
    
    
     GOVERNMENT RIGHTS 
     The present application was made with United States government support under Contract No. F33615-03-D-2357-0002, awarded by the United States Air Force. The United States government may have certain rights in the present application. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to gas turbine engines, and more particularly, to a gas turbine engine combustor. 
     BACKGROUND 
     Combustion systems in gas turbine engines remain an area of interest. Some existing systems have various shortcomings, drawbacks, and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology. 
     SUMMARY 
     One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique gas turbine engine combustion system. Still another embodiment is a unique gas turbine engine combustor. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine combustion systems and combustors. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
         FIG. 1  schematically illustrates a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
         FIG. 2  schematically illustrates a non-limiting example of a gas turbine engine combustion system in accordance with an embodiment of the present invention. 
         FIG. 3  is a perspective view of the gas turbine engine combustion system of the embodiment of  FIG. 2 . 
         FIG. 4  is another perspective view of the gas turbine engine combustion system of the embodiment of  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION 
     For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention. 
     Referring now to the drawings, and in particular,  FIG. 1 , a non-limiting example of a gas turbine engine  10  in accordance with an embodiment of the present invention is schematically depicted. In one form, gas turbine engine  10  is an axial flow machine, e.g., an air-vehicle power plant. In other embodiments, gas turbine engine  10  may be a radial flow machine or a combination axial-radial flow machine. It will be understood that embodiments of the present invention include various gas turbine engine configurations, for example, including turbojet engines, turbofan engines, turboprop engines, and turboshaft engines having axial, centrifugal and/or axi-centrifugal compressors and/or turbines. In addition, embodiments of the present invention include combined cycle engines. 
     In the illustrated embodiment, gas turbine engine  10  includes a compressor  12  having a plurality of blades and vanes  14 ; a diffuser  16 ; a combustion system  18 ; a turbine  20  having a plurality of blades and vanes  22 ; and a shaft  24  coupling compressor  12  with turbine  20 . Combustion system  18  is in fluid communication with compressor  12  and turbine  20 . Turbine  20  is drivingly coupled to compressor  12  via shaft  24 . Compressor  12 , turbine  20  and shaft  24  rotate about an engine centerline  26 . Although only a single spool is depicted, it will be understood that embodiments of the present invention include multi-spool engines having any number of spools. The number of stages of blades and vanes  14  of compressor  12 , and the number of blades and vanes  22  of turbine  20  may vary with the application, e.g., the power output requirements of a particular installation of gas turbine engine  10 . In various embodiments, gas turbine engine  10  may include one or more fans, additional compressors and/or additional turbines in one or more stages. 
     During the operation of gas turbine engine  10 , air is received at the inlet of compressor  12 . Blades and vanes  14  compress the air received at the inlet of compressor  12 , and after having been compressed, the air is discharged via into combustion system  18 . Engine  10  may include a diffuser downstream of compressor  12  to reduce the velocity of the pressurized air discharged from compressor  12 . The pressurized air discharged from compressor  12  is mixed with fuel and combusted in combustion system  18 , and the hot gases exiting combustion system  18  are directed into turbine  20 . Turbine  20  extracts energy from the hot gases to, among other things, generate mechanical shaft power to drive compressor  12  via shaft  24 . In one form, the hot gases exiting turbine  20  are directed into a nozzle (not shown), which provides thrust output for gas turbine engine  10 . In other embodiments, additional compressor and/or turbine stages in one or more additional rotors upstream and/or downstream of compressor  12  and/or turbine  20  may be employed, e.g., in single or multi-spool gas turbine engines. 
     Referring now to  FIGS. 2-4 , a non-limiting example of combustion system  18  in accordance with an embodiment of the present invention is schematically depicted. Combustion system  18  includes a recirculation combustor  28  and a slinger injector  30 . In one form combustor  28  is an annular combustor disposed about centerline  26  of engine  10 . 
     Combustor  28  includes an outer annular combustion liner  32 , an annular end wall  34 , a continuous annular fuel injection zone  36 , a plurality of compressor discharge air injectors  38 , and a plurality of air scoops  40 . Combustor  28  is disposed inside an engine case  42  about centerline  26 . In one form, combustor  28  is a single-sided annular combustor. In other embodiments, combustor  28  may not be a single-sided combustor. As a single-sided combustor, combustor  28  does not include a fixed inner annular combustion liner to help form the primary combustion zone or primary zone PZ of combustor  28 . In other embodiments, combustor  28  may include an inner annular combustion liner that extends partially or fully along the length of outer annular combustion liner  32  to contain the PZ. 
     Compressor discharge air injectors  38  and air scoops  40  are parts of a ducting system that enforces recirculation in combustor  28  to stabilize a combustion process in the form of a flame in the PZ of combustor  28 . The PZ of combustor  28  is structured to support a recirculation vortex  44  sufficient to prevent fuel from hitting the walls of combustor  28 , e.g., outer annular combustion liner  32  and annular end wall  34 , by trapping the fuel and burning it in the primary zone. In one form, annular end wall  34  has a geometric shape that facilitates the formation of recirculation vortex  44  inside combustor  28 . In one form, end wall  34  has a cross-sectional shape of a dome, and may be referred to as a dome panel. In other embodiments, end wall  34  may have other geometric configurations. In one form, end wall  34  and combustion liner  32  are integrally formed. In other embodiments, end wall  34  and combustion liner  32  may be separately formed and subsequently assembled and/or joined together. 
     Compressor discharge air injectors  38  are spaced apart circumferentially around combustor  28 . Compressor discharge air injectors  38  are operative to initiate recirculation vortex  44  in the PZ inside combustor  28 . In one form, air injectors  38  are formed integrally with combustor  28 . In other embodiments, air injectors  38  may be separately formed and subsequently affixed to combustor  28 . In one form, compressor discharge air injectors  38  extend inside of combustor  28 . In one form, compressor discharge air injectors  38  are curved tubes that extend from end wall  34  to the PZ inside combustor  28 . In one form, air injectors  38  have an end  46  that is disposed adjacent to and abuts recirculation vortex  44 . In another form, end  46  is disposed partially or completely inside recirculation vortex  44 . 
     Compressor discharge air injectors  38  include an opening  48  and an opening  50 . In one form, opening  48  is positioned to receive the total pressure of the air discharged by compressor  12  and diffused by diffuser  16 . In other embodiments, injectors  38  may be positioned otherwise. Opening  50  is positioned to discharge pressurized air into recirculation vortex  44 . The illustrated embodiment employs  20  compressor discharge air injectors  38 . Greater or lesser numbers of air injectors  38  may be employed in other embodiments. 
     Air scoops  40  are spaced apart circumferentially around combustor  28 . In one form, air scoops  40  extend inside combustor  28 . Air scoops  40  are operative to confine the recirculation vortex  44  initiated by air injectors  38 . In other embodiments, air scoops  40  may not extend inside combustor  28 . In one form, air scoops  40  are formed integrally with combustor  28 . In other embodiments, air scoops  40  may be separately formed and subsequently affixed to combustor  28 . In one form, air scoops  40  are curved tubes that extend from combustion liner  32  to the PZ inside combustor  28 . In one form, air scoops  40  have an end  52  that is disposed adjacent to and abuts recirculation vortex  44 . In other embodiments, end  52  may be disposed partially or completely inside recirculation vortex  44 . In still other embodiments, end  52  may not be adjacent to recirculation vortex  44 , and air scoops  40  may be otherwise configured to confine recirculation vortex  44 . 
     Air scoops  40  include an opening  54  and an opening  56 . In one form, opening  54  is positioned to receive the static pressure of the air discharged by compressor  12  and diffused by diffuser  16 . In other embodiments, air scoops  40  may be positioned otherwise. Opening  56  is positioned to discharge pressurized air to limit the extent of recirculation vortex  44 . The illustrated embodiment employs  20  air scoops  40 . Greater or lesser numbers of air scoops  40  may be employed in other embodiments. 
     Slinger injector  30  is operative to inject fuel into combustor  28 . In one form, slinger injector  30  is a body of revolution. Slinger injector  30  includes an opening  58  for injecting fuel F into combustor  28  in a fuel injection plane  60  that intersects with continuous annular fuel injection zone  36 . In one form, opening  58  is a circumferentially continuous opening at the forward end  62  of slinger injector  30  that provides a circumferentially continuous discharge of fuel into continuous annular fuel injection zone  36 . In other embodiments, opening  58  may employ a plurality of discrete openings in the body of revolution for discharging fuel. In one form, slinger injector  30  rotates with engine shaft  24 , i.e. at the same rotational speed as engine shaft  24  and about the same axis as engine shaft  24 , i.e., centerline  26 . In other embodiments, slinger injector  30  may rotate at a different speed than shaft  24  and/or may rotate about a different axis of rotation than engine shaft  24 . In one form, slinger injector  30  is affixed to shaft  24 . In other embodiments, slinger injector  30  may be integral with shaft  24  or may be otherwise coupled to shaft  24 . 
     Slinger injector  30  is supplied with fuel by a fuel metering and delivery system (not shown). Slinger injector  30  is operative to impart centrifugal pressurization of the fuel by rotation of slinger injector  30 , and to thereby inject the fuel into combustor  28 . That is, as slinger injector  30  rotates, it centrifugally forces fuel from the interior of the body of revolution to exit opening  58  and enter continuous annular fuel injection zone  36  of combustor  28  about fuel injection plane  60 . In one form, slinger injector  30  is designed so that the radially discharged fuel exiting slinger injector  30  helps to form recirculation zone  44  in combustor  28 . 
     The size and shape of combustor  28 ; the size, shape, orientation and location of air injectors  38  and air scoops  40 ; the penetration of air injectors  38  and air scoops  40  into combustor  28 ; and the location of fuel injection plane  60  may be selected based on analysis via computational fluid dynamics. 
     Embodiments of the present invention include a gas turbine engine, comprising: a compressor; a turbine; and a combustion system in fluid communication with the compressor and the turbine, wherein the combustion system includes a slinger injector and a single-sided recirculation combustor. 
     In a refinement, the combustor is an annular combustor having a continuous annular fuel injection zone. 
     In another refinement, the engine further comprises an engine shaft coupling the compressor and the turbine, wherein the slinger injector rotates with the engine shaft. 
     In yet another refinement, the slinger injector is affixed to or integral with the engine shaft. 
     In still another refinement, the combustor includes a compressor discharge air injector that extends inside of the combustor. 
     In yet still another refinement, the compressor discharge air injector is operative to initiate a recirculation vortex in a primary zone of the combustor. 
     In a further refinement, the compressor discharge air injector includes an inlet exposed to a total pressure of air exiting the compressor. 
     In a yet further refinement, the compressor discharge air injector is a plurality of discrete air injectors spaced apart circumferentially around the combustor. 
     In a still further refinement, an air scoop extends inside the combustor and is operative to confine a recirculation vortex initiated by the compressor discharge air injector. 
     Embodiments of the present invention include a gas turbine engine, comprising: a compressor; a turbine; means for containing a combustion process; and means for introducing fuel into the means for containing. 
     In a refinement, the means for containing the combustion process is a single-sided recirculation combustor. 
     In another refinement, the means for containing the combustion process includes means for initiating a recirculation vortex in a primary zone of the combustor. 
     In yet another refinement, the means for containing the combustion process includes means for confining the recirculation vortex. 
     In still another refinement, the means for introducing fuel is rotating. 
     In yet still another refinement, an engine shaft couples the compressor and the turbine, and the means for introducing fuel rotates with the engine shaft. 
     Embodiments include a gas turbine engine combustion system, comprising: a recirculation combustor operative to receive pressurized air from a gas turbine engine compressor and discharge combustion products to a gas turbine engine turbine, including: an outer annular combustion liner; an annular end wall; a continuous annular fuel injection zone; and a compressor discharge air injector that extends inside of the combustor from one or both of the outer annular combustion liner and the annular end wall, wherein the compressor discharge air injector is operative to initiate a recirculation vortex in a primary zone of the combustor. 
     In a refinement, wherein the compressor discharge air injector is a plurality of discrete air injectors spaced apart circumferentially around the combustor. 
     In another refinement, an air scoop extends inside the combustor, and the air scoop is operative to confine the recirculation vortex initiated by the compressor discharge air injector. 
     In yet another refinement, the air scoop is a plurality of discreet air scoops spaced apart circumferentially around the combustor. 
     In still another refinement, the air scoop includes an end disposed within the recirculation vortex initiated by the compressor discharge air injector. 
     In yet still another refinement, the compressor discharge air injector is positioned on the combustor to receive a total pressure of air discharged by the compressor; and the air scoop is positioned to receive a static pressure of air discharged by the compressor. 
     In a further refinement, the combustor is a single-sided combustor. 
     In a yet further refinement, the system includes a slinger injector operative to sling fuel into the continuous annular fuel injection zone. 
     In a still further refinement, the slinger injector is operative to provide a circumferentially continuous discharge of fuel into the continuous annular fuel injection zone. 
     While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.

Technology Classification (CPC): 5