Patent Publication Number: US-10788209-B2

Title: Combustor for gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     The present application is a Continuation of U.S. patent application Ser. No. 13/795,100 filed Mar. 12, 2013, now U.S. Pat. No. 9,127,843, the entire content of which is incorporated herein by reference. 
    
    
     FIELD OF THE INVENTION 
     The present application relates to gas turbine engines and to a combustor thereof. 
     BACKGROUND OF THE ART 
     In conventional fuel nozzle systems such as airblast and in particular air-assist, the nozzle air enters into the large combustor primary zone, losing its axial momentum but gaining radial and tangential momentum which results in diffusing the flow out rapidly. Subsequently, lower air velocity remains to perform secondary droplet break-ups. Furthermore, typical combustion systems deploy a relatively low number of discrete fuel nozzles which individually mix air and fuel as the fuel/air mixture is introduced into the combustion zone. Improvement is desirable. 
     SUMMARY 
     In accordance with an embodiment of the present disclosure, there is provided a combustor comprising: an inner liner; an outer liner spaced apart from the inner liner; an annular combustor chamber formed between the inner and outer liners, the annular combustor chamber having a central axis; fuel nozzles each having an end in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber, the fuel nozzles oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber; a plurality of nozzle air holes defined through the inner liner and the outer liner adjacent to and downstream of the fuel nozzles, the nozzle air holes configured for high pressure air to be injected from an exterior of the liners through the nozzle air holes generally radially into the annular combustor chamber, a central axis of the nozzle air holes having a tangential component relative to the central axis of the annular combustor chamber. 
     In accordance with another embodiment of the present disclosure, there is provided a gas turbine engine comprising a combustor, the combustor comprising: an inner liner; an outer liner spaced apart from the inner liner; an annular combustor chamber formed between the inner and outer liners, the annular combustor chamber having a central axis; fuel nozzles each having an end in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber, the fuel nozzles oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber; a plurality of nozzle air holes defined through the inner liner and the outer liner adjacent to and downstream of the fuel nozzles, the nozzle air holes configured for high pressure air to be injected from an exterior of the liners through the nozzle air holes generally radially into the annular combustor chamber, a central axis of the nozzle air holes having a tangential component relative to the central axis of the annular combustor chamber. 
     In accordance with yet another embodiment of the present disclosure, there is provided a method for mixing fuel and nozzle air in an annular combustor chamber, comprising: injecting fuel in a fuel direction having at least an axial component relative to a central axis of the annular combustor chamber; injecting high pressure nozzle air from an exterior of the annular combustor chamber through holes made in an outer liner of the annular combustor chamber into a fuel flow, the holes being oriented such that nozzle air is generally radially injected and has a tangential component relative to a central axis of the annular combustor chamber; and injecting high pressure nozzle air from an exterior of the annular combustor chamber through holes made in an inner liner of the annular combustor chamber into a fuel flow, the holes being oriented such that nozzle air is generally radially injected and has a tangential component relative to a central axis of the annular combustor chamber, the tangential components of the nozzle air of the inner liner and outer liner being in a same direction. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic cross-sectional view of a turbofan gas turbine engine; 
         FIG. 2  is a longitudinal sectional view of a combustor assembly in accordance with the present disclosure; 
         FIG. 3  is a sectional perspective view of the combustor assembly of  FIG. 2 ; and 
         FIG. 4  is another sectional perspective view of the combustor assembly of  FIG. 2 . 
     
    
    
     DESCRIPTION OF THE EMBODIMENT 
       FIG. 1  illustrates a turbofan gas turbine engine  10  of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan  12  through which ambient air is propelled, a multistage compressor  14  for pressurizing the air within a compressor case  15 , a combustor  16  in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section  18  for extracting energy from the combustion gases. 
     The combustor  16  is illustrated in  FIG. 1  as being of the reverse-flow type, however the skilled reader will appreciate that the description herein may be applied to many combustor types, such as straight-flow combustors, radial combustors, lean combustors, and other suitable annular combustor configurations. The combustor  16  has an annual geometry with an inner liner  20  and an outer liner  30  defining therebetween an annular combustor chamber in which fuel and air mix and combustion occurs. As shown in  FIGS. 2 and 3 , a fuel manifold  40  is positioned inside the combustion chamber and therefore between the inner liner  20  and the outer liner  30 . 
     In the illustrated embodiment, an upstream end of the combustor  16  has a sequence of zones, namely zones A, B, and C. The manifold  40  is in upstream zone A. A narrowing portion B 1  is defined in mixing zone B. A shoulder B 2  is defined in mixing zone B to support components involved in the mixing of the fuel and air, such as a louver, as described hereinafter. In dilution zone C, the combustor  16  flares to allow wall cooling and dilution air to mix with the fuel and nozzle air mixture coming from the zones B and C of the combustor  16 . A combustion zone is downstream of the dilution zone C. 
     The inner liner  20  and the outer liner  30  respectively have support walls  21  and  31  by which the manifold  40  is supported to be held in position inside the combustor  16 . Hence, the support walls  21  and  31  may have outward radial wall portions  21 ′ and  31 ′, respectively, supporting components of the manifold  40 , and turning into respective axial wall portions  21 ″ and  31 ″ towards zone B. Nozzle air inlets  22  and  32  are circumferentially distributed in the inner liner  20  and outer liner  30 , respectively. According to an embodiment, the nozzle air inlets  22  and nozzle air inlets  32  are equidistantly distributed. The nozzle air inlets  22  and nozzle air inlets  32  are opposite one another across combustor chamber. It is observed that the central axis of one or more of the nozzle air inlets  22  and  32 , generally shown as N, may have an axial component and/or a tangential component, as opposed to being strictly radial. Referring to  FIG. 2 , it is observed that the central axis N is oblique relative to a radial axis R of the combustor  16 , in a plane in which lies a longitudinal axis X of the combustor  16 . Hence, the axial component NX of the central axis N is oriented downstream, i.e., in the same direction as that of the flow of the fuel and air, whereby the central axis N leans towards a direction of flow (for instance generally parallel to the longitudinal axis X). In an embodiment, the central axis N could lean against a direction of the flow. 
     Referring to  FIGS. 3 and 4 , the central axis N of one or more of the nozzle air inlets  22  and  32  may have a tangential component NZ, in addition or in alternative to the axial component NX. For simplicity, in  FIGS. 3 and 4 , only the tangential component NZ of the central axis N is shown, although the nozzle air inlets  22  and  32  may have both an axial and a tangential component. The tangential component NZ is oblique relative to radial axis R in an axial plane, i.e., the axial plane being defined as having the longitudinal axis X of the combustor  16  being normal to the axial plane. In  FIG. 3 , the tangential component NZ is in a counterclockwise direction, while in  FIG. 4 , the tangential component NZ is clockwise. The tangential component NZ may allow an increase residence time of the air and fuel mixture in the downstream mixing zone B of the combustor  16 . 
     Referring to  FIG. 2 , nozzle air inlets  23  and  33  may be located in the narrowing portion B 1  of mixing zone B. Alternatively, as shown in  FIG. 3 , the nozzle air inlets  23  and  33  may be in the upstream zone A. The nozzle air inlets  23  and  33  may form a second circumferential distribution of inlets, if the combustor  16  has two circumferential distributions of inlets (unlike  FIG. 4 , showing a single circumferential distribution). In similar fashion to the set of inlets  22 / 32 , the inlets  23  and  33  are respectively in the inner liner  20  and outer liner  30 . The inlets  23  and  33  may be oriented such that their central axes X may have an axial component and/or a tangential component. 
     Hence, the combustor  16  comprises numerous nozzle air inlets (e.g.,  22 ,  23 ,  32 ,  33 ) impinging onto the fuel sprays produced by the fuel manifold  40 , in close proximity to the fuel nozzles, thereby encouraging rapid mixing of air and fuel. The orientation of the nozzle air inlets relative to the fuel nozzles (not shown) may create the necessary shearing forces between air jets and fuel stream, to encourage secondary fuel droplets breakup, and assist in rapid fuel mixing and vaporization. 
     Purged air inlets  24  and  34  may be respectively defined in the inner liner  20  and the outer liner  30 , and be positioned in the upstream zone A of the combustor  16 . In similar fashion to the sets of nozzle air inlets  22 / 32 , a central axis of the purged air inlets  24  and  34  may lean toward a direction of flow with an axial component similar to axial component NX, as shown in  FIG. 2 . Purged air inlets  24  and  34  produce a flow of air on the downstream surface of the manifold  40 . As shown in  FIGS. 2, 3 and 4 , sets of cooling air inlets  25  and  35 , and cooling air inlets  25 ′ and  35 ′, respectively in the inner liner  20  and the outer liner  30 , may be circumferentially distributed in the mixing zone B downstream of the sets of nozzle air inlets  23  and  33 . The cooling air inlets  25 ,  25 ′,  35 ,  35 ′ may be in channels defined by the liners  20  and  30  and mixing walls  50  and  60  (described hereinafter). Cooling air inlets  25 ,  25 ′,  35  and  35 ′ may produce a flow of air on flaring wall portions of the inner liner  20  and outer liner  30 . 
     Referring to  FIG. 4 , dilution air inlets  26  and  36  are circumferentially distributed in the dilution zone C of the combustor  16 , respectively in the inner liner  20  and outer liner  30 . According to an embodiment, the dilution air inlets  26  and  36  are equidistantly distributed, and opposite one another across combustor chamber. It is observed that the central axis of one or more of the dilution air inlets  26  and  36 , generally shown as D, may have an axial component and/or a tangential component, as opposed to being strictly radial. Referring to  FIG. 4 , the central axis D is oblique relative to a radial axis R of the combustor  16 , in a plane in which lies a longitudinal axis X of the combustor  16 . Hence, the axial component DX of the central axis D is oriented downstream, i.e., in the same direction as that of the flow of the fuel and air, whereby the central axis D leans towards a direction of flow (for instance generally parallel to the longitudinal axis X). In an embodiment, the central axis D could lean against a direction of the flow. 
     Still referring to  FIG. 4 , the central axis D of one or more of the dilution air inlets  26  and  36  may have a tangential component DZ, in addition or in alternative to the axial component DX. For simplicity, in  FIG. 4 , one inlet is shown with only the axial component DX, while another is shown with only the tangential component DZ. It should however be understood that the inlets  26  and  36  may have both the axial component DX and the tangential component DZ. The tangential component DZ is oblique relative to radial axis R in an axial plane, i.e., the axial plane being defined as having the longitudinal axis X of the combustor  16  being normal to the axial plane. In  FIG. 4 , the tangential component DZ is in a counterclockwise direction. It is thus observed that the tangential component DZ of the central axes D may be in an opposite direction than that of the tangential component NZ of the central axes N of the nozzle air inlets  22 ,  23 ,  32 , and/or  33 , shown as being clockwise. The opposite direction of tangential components DZ and NZ may enhance fluid mixing to render the fuel and air mixture more uniform, which may lead to keeping the flame temperature relatively low (and related effects, such as lower NOx and smoke emissions, low pattern factor, and enhanced hot-section durability). The opposite tangential direction of dilution air holes relative to the nozzle air holes cause the creation of a recirculation volume immediately upstream of the penetrating dilution jets, further enhancing fuel-air mixing before burning, in a relatively small combustor volume. It is nonetheless possible to have the tangential components of nozzle air inlets and dilution air inlets being in the same direction, or without tangential components. 
     Referring to  FIG. 4 , a plurality of cooling air inlets  27  may be defined in the inner liner  20  and outer liner  30  (although not shown). The outer liner  30  has a set of dilution air inlets  37  in an alternating sequence with the set of dilution air inlets  36 . The dilution air inlets  37  have a smaller diameter than that of the dilution air inlets  36 . This alternating sequence is a configuration considered to maximize the volume of dilution in a single circumferential band, while providing suitable structural integrity to the outer liner  30 . 
     Referring to  FIGS. 2 to 4 , the manifold  40  is schematically shown as having fuel injector sites  41  facing downstream on an annular support  42 . The annular support  42  may be in the form of a full ring, or a segmented ring. The fuel injector sites  41  are circumferentially distributed in the annular support  42 , and each accommodate a fuel nozzle (not shown). It is considered to use flat spray nozzles to reduce the number of fuel injector sites  41  yet have a similar spray coverage angle. As shown in  FIGS. 3 and 4 , the number of nozzle air inlets (e.g.,  22 ,  23 ,  32 , and  33 ) is substantially greater than the number of fuel injector sites  41 , and thus of fuel nozzles of the manifold  40 . Moreover, the continuous circumferential distribution of the nozzle air inlets relative to the discrete fuel nozzles creates a relative uniform air flow throughout the upstream zone A in which the fuel stream is injected. 
     A liner interface comprising a ring  43  and locating pins  44  or the like support means may be used as an interface between the support walls  21  and  31  of the inner liner  20  and outer liner  30 , respectively, and the annular support  42  of the manifold  40 . Hence, as the manifold  40  is connected to the combustor  16  and is inside the combustor  16 , there is no relative axial displacement between the combustor  16  and the manifold  40 . 
     As opposed to manifolds located outside of the gas generator case, and outside of the combustor, the arrangement shown in  FIGS. 2-4  of the manifold  40  located inside the combustor  16  does not require a gas shielding envelope, as the liners  20  and  30  act as heat shields. The manifold  40  is substantially concealed from the hot air circulating outside the combustor  16 , as the connection of the manifold  40  with an exterior of the combustor  16  may be limited to a fuel supply connector projecting out of the combustor  16 . Moreover, in case of manifold leakage, the fuel/flame is contained inside the combustor  16 , as opposed to being in the gas generator case. Also, the positioning of the manifold  40  inside the combustor  16  may result in the absence of a combustor dome, and hence of cooling schemes or heat shields. 
     Referring to  FIGS. 2 and 4 , mixing walls  50  and  60  are respectively located in the inner liner  20  and outer liner  30 , against the shoulders B 2  upstream of the narrowing portion B 1  of the mixing zone B, to define a straight mixing channel. The mixing walls  50  and  60  form a louver. Hence, the mixing walls  50  and  60  concurrently define a mixing channel of annular geometry in which the fuel and nozzle air will mix. The mixing walls  50  and  60  are straight wall sections  51  and  61  respectively, which straight wall sections  51  and  61  are parallel to one another in a longitudinal plane of the combustor  16  (i.e., a plane of the page showing  FIG. 2 ). The straight wall sections  51  and  61  may also be parallel to the longitudinal axis X of the combustor  16 . Other geometries are considered, such as quasi-straight walls, a diverging or converging relation between wall sections  51  and  61 , among other possibilities. For instance, a diverging relation between wall sections  51  and  61  may increase the tangential velocity of the fluid flow. It is observed that the length of the straight wall sections  51  and  61  (along longitudinal axis X in the illustrated embodiment) is several times greater than the height of the channel formed thereby, i.e., spacing between the straight wall sections  51  and  61  in a radial direction in the illustrated embodiment. Moreover, the height of the channel is substantially smaller than a height of the combustion zone downstream of the dilution zone C. According to an embodiment, the ratio of length to height is between 2:1 and 4:1, inclusively, although the ratio may be outside of this range in some configurations. The presence of narrowing portion B 1  upstream of the mixing channel may cause a relatively high flow velocity inside the mixing channel. This may for instance reduce the flashback in case of auto-ignition during starting and transient flow conditions. The configuration of the mixing zone B is suited for high air flow pressure drop, high air mass flow rate and introduction of high tangential momentum, which may contribute to reaching a high air flow velocity. 
     The mixing walls  50  and  60  respectively have lips  52  and  62  by which the mixing annular chamber flares into dilution zone C of the combustor  16 . Moreover, the lips  52  and  62  may direct a flow of cooling air from the cooling air inlets  25 ,  25 ′,  35 ,  35 ′ along the flaring wall portions of the inner liner  20  and outer liner  30  in dilution zone C. 
     Hence, the method of mixing fuel and nozzle air is performed by injecting fuel in a fuel direction having axial and/or tangential components, relative to the central axis X of the combustor  16 . Simultaneously, nozzle air is injected from an exterior of the combustor  16  through the holes  32 ,  33  made in the outer liner  30  into a fuel flow. The holes  32 ,  33  are oriented such that nozzle air has at least a tangential component NZ relative to the central axis X of the combustor  16 . Nozzle air is injected from an exterior of the combustor  16  through holes  22 ,  23  made in the inner liner  20  into the fuel flow. The holes  22 ,  23  are oriented such that nozzle air has at least the tangential component NZ relative to the central axis X, with the tangential components NZ of the nozzle air of the inner liner  20  and outer liner  30  being in a same direction. Dilution air may be injected with a tangential component DZ in an opposite direction. 
     The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. Other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.