Abstract:
A gas turbine engine combustor includes an apparatus for directing air around the axis along which fuel is injected into the combustor to thereby improve the combustor.

Description:
TECHNICAL FIELD  
       [0001]     The present invention relates generally to gas turbine engine combustors and, more particularly, to a low cost combustor configuration having improved performance.  
       BACKGROUND OF THE ART  
       [0002]     Gas turbine combustors are the subject of continual improvement, to provide better cooling, better mixing, better fuel efficiency, better performance, etc. at a lower cost. Also, a new generation of very small gas turbine engines is emerging (i.e. a fan diameter of 20 inches or less, with about 2500 lbs. thrust or less), however larger designs cannot simply be scaled-down, since many physical parameters do not scale linearly, or at all, with size (droplet size, drag coefficients, manufacturing tolerances, etc.). There is, therefore, a continuing need for improvements in gas turbine combustor design.  
       SUMMARY OF THE INVENTION  
       [0003]     In accordance with the present invention there is provided a gas turbine engine combustor comprising a liner enclosing a combustion chamber, the liner including a dome portion at an upstream end thereof, the dome portion having defined therein a plurality of openings each adapted to receive a fuel nozzle and a plurality of holes defined around each opening, each opening having an axis generally aligned with an fuel injection axis of a fuel nozzle received by the opening, the holes adapted to direct air into the combustion chamber in a spiral around the axis of an associated one of said openings.  
         [0004]     In accordance with another aspect there is also provided a gas turbine engine combustor comprising a liner enclosing a combustion chamber, the liner having defined therein a plurality of openings each adapted to receive a fuel nozzle for directing fuel into the combustion chamber generally along an axis of the opening, the liner also having means associated with each opening for directing air into the combustion chamber in a spiral pattern around an axis of the associated opening.  
         [0005]     In accordance with another aspect there is also provided a method of combusting fuel in a gas turbine combustor, the method comprising the steps of injecting a mixture of fuel and air into the combustor along an axis, igniting the mixture to create at least one combustion zone in which the mixture is combusted, and directing air into the combustor around said axis in a spiral pattern;  
         [0006]     Further details of these and other aspects of the present invention will be apparent from the detailed description and Figures included below. 
     
    
     DESCRIPTION OF THE DRAWINGS  
       [0007]     Reference is now made to the accompanying Figures depicting aspects of the present invention, in which:  
         [0008]      FIG. 1  shows a schematic cross-section of a turbofan engine having an annular combustor;  
         [0009]      FIG. 2  shows an enlarged view of the combustor of  FIG. 1 ;  
         [0010]      FIG. 3  shows an enlarged view of an alternate embodiment of a combustor of the present invention, schematically depicting a subset of the holes which may be provided therein;  
         [0011]      FIG. 4  shows an inside end view of the dome of the combustor of  FIG. 2 ;  
         [0012]      FIG. 5  is a view similar to  FIG. 2 , schematically depicting the device in use;  
         [0013]      FIG. 6  is a view similar to  FIG. 3 , schematically depicting an aspect of the device in use; and  
         [0014]      FIG. 7  is similar to  FIG. 6 , but showing one effect of the one aspect of the present invention. 
     
    
     DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS  
       [0015]      FIG. 1  illustrates a gas turbine engine  10  preferably of a type 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, an annular combustor  16  in which compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases which is then redirected by combustor  16  to a turbine section  18  for extracting energy from the combustion gases.  
         [0016]     Referring to  FIG. 2 , the combustor  16  is housed in a plenum  20  defined partially by a gas generator case  22  and supplied with compressed air from compressor  14  by a diffuser  24 . Combustor  16  comprises generally a liner  26  composed of an outer liner  26 A and an inner liner  26 B defining a combustion chamber  32  therein. Combustor  16  preferably has a generally flat dome  34 , as will be described in more detail below. Outer liner  26 A includes a outer dome panel portion  34 A, a relatively small radius transition portion  36 A, a cylindrical body panel portion  38 A, long exit duct portion  40 A, while inner liner  26 B includes an inner dome panel portion  34 B, a relatively small radius transition portion  36 B, a cylindrical body panel portion  38 B, and a small exit duct portion  40 B. The exit ducts  40 A and  40 B together define a combustor exit  42  for communicating with turbine section  18 . The combustor liner  26  is preferably sheet metal. A plurality of holes  44  are provided in liner  26 , a plurality of holes  46  an  46 ′ (see  FIG. 4 ) are provided in dome  34 , and a plurality of holes  48  are provided in transitions  36 , as will be described further below.  
         [0017]     A plurality of air-guided fuel nozzles  50 , having supports  52  and supplied with fuel from internal manifold  54 , communicate with the combustion chamber  32  through nozzle openings  56  to deliver a fuel-air mixture  58  to the chamber  32 . As depicted in  FIG. 2 , the fuel-air mixture is delivered in a cone-shaped spray pattern, and therefore referred to in this application as fuel spray cone  58 .  
         [0018]     In use, high-speed compressed air enters plenum  20  from diffuser  24 . The air circulates around combustor  16 , as will be discussed in more detail below, and eventually enters combustion chamber  32  through a plurality of holes  44  in liner  26 , holes  46  and  46 ′ in dome  34 , and holes  48  in transition  36 . Once inside the combustor  16 , the air is mixed with fuel and ignited for combustion. Combustion gases are then exhausted through exit  42  to turbine section  18 .  
         [0019]     Referring to  FIG. 3 , as mentioned combustor  16  has holes  44 ,  46  and  48  therein (represented schematically in this Figure by the indication of their centrelines only) provided for cooling of the liner  26 . (For clarity of explanation, holes  46 ′ will be temporarily ignored.) It will be understood that effusion cooling is often achieved by directing air though angled holes in a combustor liner. Therefore, holes  46  in dome panel  34  are angled outwardly away from nozzle  50 , while holes  44  are angled downstream in the combustor. Holes  48  in transition portions  36 A,B are provided generally parallelly to body panel portion  38 A,B to direct cooling air in a louver-like fashion along the interior of body panel portions  38 A,B to cool them. It will be noted in this embodiment that transition portions  36 A,B are frustoconical with relatively small radii connections to their respective dome and body panels.  
         [0020]     Referring now to  FIG. 4 , holes  46  in dome panels  34 A,B, include holes  46 ′, which provided preferably in a concentric circular configuration around nozzle opening  56  and angled generally tangentially relative to an associated opening  56  to deliver air in a circular or helical pattern around opening  56 . The entry/exit angle of holes  46 ′ is indicated by the arrows in  FIG. 4 , and is noted to be generally tangential to opening  56  when viewed in this plane. The patterns of holes  46 ′ around openings  56  may interlace, for example as in region  62  indicated in  FIG. 4 . Holes  46  may also interlace with holes  46 ′ in a region, such as region  62  for example.  
         [0021]     Referring to  FIG. 5 , in use, air entering combustor  16  through holes  46 ′ will tend to spiral around nozzle opening  56  in a helical fashion, and thus create a vortex around fuel spray cone  58 , as will be discussed in further detail below. Holes  46 ′ are preferably provided in the flat end portion of dome panels  34 , to provide better control over the vortex created, as will also be discussed further below.  
         [0022]     The combustor  16  is preferably provided in sheet metal, and may be made by any suitable method. Holes  44 ,  46 , and  48  are preferably drilled in the sheet metal, such as by laser drilling. It will be appreciated in light of the description, however, that holes  48  in transition  36  are provided quite close to body panels  38 A,B, and necessarily are so to provide good film cooling of body panels  38 A,B. This configuration, however, makes manufacturing difficult since the drilling of holes  48  may inadvertently compromise the body panel behind this hole, and thereby result in a scrapped part. While drilling can be controlled with great precision, such precision adds to the cost of the part. According to the present invention, however, providing combustor  16  with small radius transition portions  36 A,B and a flat dome permits drilling to completed less precisely and with minimal risk of damaging the adjacent body panels. This is because manufacturing tolerances for drilled holes provided on curved or conical surfaces are much larger than the comparable tolerances for drilling on a flat, planar surface. Thereby, maximizing the flat area of the combustor dome, the present invention provides an increase area over which cooling holes may be more accurately provided. This is especially critical in heat shield-less combustor designs (i.e. in which the liner has no inner heat shield, but rather the dome is directly exposed to the combustion chamber), since the cooling of the dome therefore become critical, and the cooling pattern must be precisely provided therein. By improving the manufacturing tolerances of the combustor dome, the chance of holes not completely drilled-through, or drilling damage occurring on a liner surface downstream of the drilled hole (i.e. caused by the laser or drilling mechanism hitting the liner after completing the hole) are advantageously reduced. Thus, by making the dome end flat, holes may be drilled much closed to the “corners” (i.e. the intersection between the dome and the side walls), with reduced risk of accidentally damaging the liner side walls downstream of the hole (i.e. by over-drilling). Although a flat dome, depending on its configuration, may present dynamic or buckling issues in larger-sized configurations, the very small size of a combustor for a very small gas tribune engine will in part reduce this tendency. This aspect of the invention is thus particularly suited for use in very small gas turbine engines. In contrast, conventional annular reverse-flow combustors have curved domes to provide stability against dynamic forces and buckling. However, as mentioned, this typical combustor shape presents interference and tolerance issues, particularly when providing an heat shield-less combustor dome.  
         [0023]     Referring to  FIG. 6 , in some combustor installations, flow restrictions may exist upstream of dome  34 , which may be caused, for example, by a small clearance h between case  22  and combustor  16  (in this case) and/or by the presence of airflow obstructions outside the combustor outside the combustor dome, such as (referring again to  FIG. 2 ) the supports  52 , the fuel manifold  54  and/or igniters (not shown) or other obstructions. These flow restrictions typically result in higher flow velocity between case  22  and liner  26  than is present in engines without such geometries, and these velocities are especially high around the outer liner/dome intersection, and may result in a “wake area” being generated (designated schematically by the shaded region  60 ), in which the air pressure will be lower than the surrounding flow. Consequently, air entering combustor  16  through effusion holes  46  adjacent wake area  60  will have relatively lower momentum (represented schematically by the relative thickness of flow arrows in  FIG. 6 ), which negatively impacts cooling performance. This problem is particularly acute in the next generation of very small gas turbofan engines, having a fan diameter of 20 inches or less, 2500 lbs. thrust or less. Larger prior art gas turbines have the ‘luxury’ of a relatively larger cavity around the liner and thus may avoid such restrictions altogether. However, in very small turbofans, space is at an absolute a premium, and such flow restrictions are all but unavoidable.  
         [0024]     Referring again to  FIGS. 3 and 6 , exacerbating the problem created by the wake area, in a combustor configuration where the effusion cooling holes  46  in the upper half of dome  34 A are directed away from the combustor centre, air entering these holes must thus essentially reverse direction relative to the air flow outside the combustor adjacent the wake area. This further reduces the momentum of air entering in the combustion chamber in this area. Consequently, very low cooling effectiveness results adjacent this area inside the liner, and thus can undesirably permit the flame to stabilize close to the combustor outer wall. This results in the upper half of the dome and combustor outer liner being very hot compared to bottom half/inner liner, since the dome cooling holes in this portion of the combustor have the same general direction as the air flow in plenum  22 .  
         [0025]     To address this problem, the cooling hole pattern of the present invention improves the flow in the wake area by reducing the overall drag coefficient (C d ) in the wake area by providing holes  46 ′ in addition to holes  46 , and thus permitting more direct entry of air into the combustor (since holes  46 ′ are not angled as harshly relative to the primary flow in plenum  20 , and thus air may enter combustor  16  at a higher momentum though holes  46 ′ than through holes  46 . This higher momentum air exiting from holes  46 ′ assists holes  46  in pushing away fuel from the liner walls to impede flame stabilization near the wall liner wall.  
         [0026]     Perhaps more importantly, however, the spiral or helical flow also helps to constrain the lateral extent of fuel spray cone  58 . Referring again to  FIG. 5 , as mentioned above the pattern of holes  46 ′ causes air inside the liner to spiral or spin in a vortex around the fuel nozzle and away from dome  34  and into combustion chamber  32 . This helps keep the fuel spray away from dome panel  34  as well as the upstream portions of the outer and inner liner panels adjacent to the dome by narrowing the width of the fuel spray cone. Although the skilled reader will appreciate that the size of fuel spray cone  58  can also be controlled by the nozzle characteristics (e.g. the spray cone can be narrowed by using more air in the nozzle swirler, or providing a nozzle having a narrower nozzle cone), such nozzle-based modes of control are less preferable than the present solution, since the present invention makes use of cooling air already in use to cool the combustor wall (which permits improved efficiency over using increase guide air), and permits a shorter combustor length since a narrower spray generated from the nozzle swirler will require a longer combustor liner or otherwise cause burning of the LED  40 A by fuel impingement of fuel thereon. Thus, the present invention facilitates both efficiency and size reduction improvements.  
         [0027]     The spiral flow inside the liner also provides better fuel/air mixing and thus also improves the re-light characteristic of the engine, because the spiral flow ‘attacks’ the outer shell of the fuel spray cone, which is consists of the lower density of fuel particles, and thus improves fuel-air mixing in the combustion chamber.  
         [0028]     As a result of the hole pattern of the present invention, a novel combustor air flow pattern results. Conventionally, combustor internal aerodynamics provide either single torroidal or double torroidal flows inside the liner, however the present invention results in new aerodynamic pattern due to spiral flow introduced inside the liner.  
         [0029]     The present invention is believed to be best implemented with a combustor having a flat dome panel. Although the invention may also be applied to conical, curved or other shaped dome panels, it is believed that the spiral flow which is introduced inside the liner will be inferior to that provided by the present hole pattern in a flat dome panel.  
         [0030]     The above description is meant to be exemplary only, and one skilled in the art will recognize that further changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, the invention may be provided in any suitable annular combustor configuration, and is not limited to application in turbofan engines. It will also be understood that holes  46 ′ need not be provided in a concentric circular configuration, but in any suitable pattern. Holes  46  and  46 ′ need not be provided in distinct regions of the dome  34 , and may instead be interlaced in overlapping regions. Holes  46 ′ around adjacent nozzle openings  56  may likewise be interlaced with one another. The direction of vortex flow around each nozzle is preferably in the same direction, though not necessarily so. Each nozzle does not require a vortex, though it is preferred. Although the use of holes for directing air is preferred, other means such as slits, louvers, etc. may be used in place of or in addition to holes. Still other modifications 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.