Patent Publication Number: US-9429324-B2

Title: Fuel injector with radial and axial air inflow

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to a fuel injector used in, for example, a gas turbine engine and including a combined fuel injector configured by combining a plurality of fuel nozzles, and particularly to a main injector of the fuel injector. 
     2. Description of the Related Art 
     In recent years, in consideration of the environment, there is a need for a reduction of NOx (nitrogen oxide) emitted from gas turbine engines. The NOx to be emitted from the gas turbine engine is generated mainly by oxidization of nitrogen in inflow air when fuel is supplied to the inflow air and combusted at high temperature. Meanwhile, the amount of CO2 emission of the gas turbine engine, that is, fuel consumption decreases as an exhaust gas at an exit of a combustor increases in temperature. Therefore, to reduce the CO2, the fuel needs to be combusted at high temperature by increasing a fuel-air ratio. According to a fuel nozzle of a combustor of a conventional gas turbine engine, the fuel is directly sprayed to a combustion chamber without premixing the fuel with the air. Therefore, before the fuel is adequately mixed with the air, the fuel combusts, and regions where a flame temperature is significantly higher than an average value are generated locally. The amount of NOx generation increases exponentially with the flame temperature. Therefore, a large amount of NOx is generated from the local regions where the flame temperature is high. On this account, according to the conventional combustion method, when the temperature of the exhaust gas at the exit of the combustor is increased, the amount of NOx emission increases sharply. 
     To reduce the local regions where the flame temperature is high, a lean premix combustion method is effective. According to this method, the fuel and the air are premixed, and a fuel-air mixture in which the fuel in the form of a mist is dispersed in the air is supplied to the combustion chamber and combusted therein. Meanwhile, according to the lean premix combustion method, in a case where the output of the gas turbine engine is low and the fuel-air ratio is low, the flame is unstable and incomplete combustion tends to occur as compared to a case where the fuel is directly sprayed to the combustion chamber. Here, a concentric fuel injector has been devised. This fuel injector is configured such that a pilot injector and a main injector provided outside the pilot injector are provided coaxially. When the output of the gas turbine engine is low, the fuel is directly sprayed from only the pilot injector to the combustion chamber to maintain stable combustion. When the output of the gas turbine engine is intermediate or high, that is, when the amount of NOx emission is large, the amount of fuel injected directly from the pilot injector is reduced, and a pre-mixture generated by the main injector is also injected to the combustion chamber. With this, the amount of NOx emission is reduced. Regarding a gas turbine engine for aircrafts, the output of the gas turbine engine is substantially low (lower than about 40% of the rated output) in a state of each of ground idle, flight idle, and approach, the output of the gas turbine engine is substantially intermediate (about 40 to 80% of the rated output) in a cruising state, and the output of the gas turbine engine is substantially high (about 80 to 100% of the rated output) in a state of each of climb and takeoff. 
     According to the concentric fuel injector, the temperature of the gas at an entrance of the combustor and the fuel-air ratio when the output of the gas turbine engine is intermediate are respectively lower than those when the output of the gas turbine engine is high, and the flame temperature when a main pre-mixture is combusted is low. Therefore, although the amount of NOx emission is generally not so large, the main pre-mixture tends to cause the incomplete combustion, and the combustion efficiency tends to be low. On this account, when the output of the gas turbine engine is intermediate, flame holding of the main pre-mixture by combustion flame of the pilot injector becomes important. Meanwhile, when the output of the gas turbine engine is high, the flame temperature is adequately high, so that the flame stabilizes only by the main pre-mixture, and the combustion efficiency has no problem. However, the amount of NOx emission tends to be large. Therefore, uniformization of the main pre-mixture needs to be further considered. To maintain satisfactory performance when the output of the gas turbine engine is intermediate or high, a fuel injector has been proposed, in which: two air channels through which air inflows in a radial direction are formed on the main injector; and fuel injection holes corresponding to one fuel supply system are formed on each air channel, that is, the fuel injection holes corresponding to each of two fuel supply systems are formed (see Japanese Laid-Open Patent Application Publication No. 2003-262337). According to this fuel injector, when the output of the gas turbine engine is intermediate, the fuel is injected only to one air channel through which the pre-mixture is supplied to a position closer to the pilot injector. With this, the flame holding of the pre-mixture by the combustion flame of the pilot injector is promoted, and this improves the combustion efficiency. Meanwhile, when the output of the gas turbine engine is high, the fuel is injected to both air channels, and the pre-mixture is generated more uniformly. Thus, the NOx emission is suppressed. 
     According to the fuel injector of Japanese Laid-Open Patent Application Publication No. 2003-262337, since the fuel is injected to the two air channels, the fuel injection holes corresponding to each of the two fuel supply systems are formed. Forming the fuel injection holes corresponding to each of a plurality of fuel supply systems as above increases structural complexity of the fuel injector. Generally, according to the concentric fuel injector, when the output of the gas turbine engine is low, the fuel flows only to the pilot injector, and the fuel in a fuel channel in the main injector stays therein. The staying fuel causes carbonization (coking) at a certain temperature or higher and accumulates on a wall surface of the fuel channel to close the fuel channel. As a countermeasure against this problem, in the main injector, a cooling structure for preventing the coking needs to be provided for each of the fuel supply systems of the fuel injection holes. However, providing the cooling structures respectively for the plurality of fuel supply systems of the fuel injection holes within a limited space in the main injector extremely increases the structural complexity of the main injector. Moreover, in a case where the fuel injection holes to be used are switched in accordance with the output, a control mechanism is required, and it is extremely difficult to secure reliability for ensuring appropriate selection of the fuel injection holes in any scene. 
     SUMMARY OF THE INVENTION 
     The present invention addresses the above described conditions, and an object of the present invention is to provide a fuel injector capable of realizing the NOx reduction by adequately mixing the air and the fuel when the output of the gas turbine engine is intermediate or high, that is, when the main injector is operating, without increasing the complexity of devices and control mechanisms. 
     To achieve the above object, a fuel injector according to the present invention includes: a pilot injector configured to spray fuel so as to form a first combustion region in a combustion chamber; and a main injector provided coaxially with the pilot injector so as to surround the pilot injector and configured to supply a fuel-air mixture that is a mixture of the fuel and air to form a second combustion region in the combustion chamber, wherein the main injector includes: a first inflow channel configured to take the air therethrough and give the air a major flow component in an axial direction; a second inflow channel configured to take the air therethrough, give the air a major flow component in a radial direction, and cause the air therein to meet the air from the first inflow channel; and a main fuel injecting portion configured to inject the fuel only to the second inflow channel. Here, the air having the major flow component in the axial direction may include a small flow component in a radial direction or a circumferential direction. The air having the major component in the radial direction may include a small flow component in the axial direction or the circumferential direction. 
     According to this configuration, the air of the first inflow channel and the fuel-air mixture of the second inflow channel meets at a certain angle. Therefore, after the air and the fuel-air mixture meet, the air and the fuel are adequately mixed with each other in a comparatively short distance. Thus, the NOx reduction can be realized when the output of the gas turbine engine is intermediate or high, that is, when the main injector is operating. In addition, since the fuel is injected only to the second inflow channel, a fuel channel and its cooling structure can be simplified. 
     In the present invention, it is preferable that the main injection vale inject the fuel to the second inflow channel from the main fuel injecting portion provided at a portion which defines a boundary between the first inflow channel and the second inflow channel. According to this configuration, when the output of the gas turbine engine is intermediate, that is, when the momentum of the injection of the main fuel is small, the injected fuel just reaches a region close to the inject holes, as compared to when the output of the gas turbine engine is high, that is, when the momentum thereof is large. As a result, the fuel is injected mainly to a position close to the main fuel injecting portion in the air flow of the second inflow channel. Therefore, when the air flow of the second inflow channel meets the air flow of the first inflow channel to be changed to the air flow in the axial direction and is then injected to the combustion chamber, the fuel in the form of a mist flows on a radially inward side as compared to when the output of the gas turbine engine is high. To be specific, when the output of the gas turbine engine is intermediate, the main fuel in the form of a mist gets close to the first combustion region where the combustion state is stable, as compared to when the output of the gas turbine engine is high. As a result, the flame holding effect by the flame in the first combustion region can be easily obtained at the time of combustion. Thus, the combustion efficiency improves. Moreover, the portion which defines a boundary between the first inflow channel and the second inflow channel can generally secure a space widely in many cases. Therefore, a structure, such as a cooling structure for preventing coking, in the main fuel injecting portion can be easily, spatially arranged. 
     In the present invention, it is preferable that: a first swirling unit and a second swirling unit be respectively attached to an entrance of the first inflow channel and an entrance of the second inflow channel; and the second swirling unit include a plurality of swirling portions, the swirling portion located closest to the main fuel injecting portion causes inflow air to flow straight in a substantially radially inward direction, and the remaining swirling portions give a swirl velocity component to inflow air. 
     According to this configuration, in the vicinity of the main fuel injecting portion of the second inflow channel, the air flow simply flowing straight in a substantially radially inward direction is generated by the swirling portion located closest to the main fuel injecting portion. Meanwhile, the swirl air flow is generated by the remaining swirling portions at a position away from the main fuel injecting portion in the second inflow channel. When the output of the gas turbine engine is intermediate, that is, when the flow quantity of the fuel is small and the injection velocity of the fuel is low, the momentum of the fuel is small. Therefore, most of the fuel injected from the main fuel injecting portion cannot reach the swirl air flow generated by the remaining swirling portions. On this account, the fuel in the form of a mist is not diffused in the radial direction by the swirl air flow and flows in the radially inward direction together with the air flow flowing in the radially inward direction. Thus, the fuel-air mixture is generated such that the fuel distribution is positioned on an inner peripheral side of the air channel of the main injector. As a result, the fuel-air mixture having a high fuel concentration is supplied to a region which is located on a radially inward side and close to the first combustion region. Thus, the combustion efficiency when the output of the gas turbine engine is intermediate further improves. 
     When the output of the gas turbine engine is high, that is, when the flow quantity of the fuel injected from the main fuel injecting portion is large and the injection velocity of the fuel is high, the momentum of the fuel is large. Therefore, the fuel is inject to a wide range of the second inflow channel, a part of the injected fuel flows in the radially inward direction as with when the output of the gas turbine engine is intermediate, and the remaining fuel reaches the swirl air flow generated by the other swirling portions to flow in a radially outward direction. As a result, when the output of the gas turbine engine is high, the fuel-air mixture is uniformly generated in the entire air channel of the main injector. Thus, the NOx reduction is realized. As above, by such a simple configuration, the fuel distribution suitable for output conditions is realized, and a desired performance can be obtained. 
     In a preferred mode of the present invention, a position of an exit end of the pilot injector coincides with or is upstream of a position of an exit end of the main injector in the axial direction. In this case, it is preferable that a ratio W/Dm that is a ratio of an axial distance W between the exit ends to an inner diameter Din of the exit end of the main injector be 0.25 or less. According to this configuration, the fuel-air mixture from the main injector promptly contacts the first combustion region in the vicinity of the exit of the pilot injector. As a result, when the output of the gas turbine engine is intermediate, the fuel-air mixture of the main injector starts combusting from a further upstream side, so that the combustion efficiency improves. 
     In the present invention, it is preferable that the fuel injector further include an annular dividing wall configured to define a boundary between the pilot injector and the main injector, and a ratio T/Dp that is a ratio of a radial width T of an exit end of the dividing wall to an inner diameter Dp of the exit end of the pilot injector be 0.02 to 0.15. According to this configuration, since the dividing wall is adequately small (thin), the fuel-air mixture from the main injector easily contacts the first combustion region when the output of the gas turbine engine is intermediate. As a result, the flame holding is easily realized by the flame of the first combustion region. Thus, the combustion efficiency can be improved. 
     In the present invention, it is preferable that the fuel injector further include an annular dividing wall configured to define a boundary between the pilot injector and the main injector, and a virtual extended inner peripheral surface extending from an exit end of an inner peripheral surface of the dividing wall in a downstream direction and a virtual extended outer peripheral surface extending from an exit end of an outer peripheral surface of the dividing wall in the downstream direction extend in parallel with each other in the downstream direction or gradually separate from each other as they extend in the downstream direction. According to this configuration, when the output of the gas turbine engine is low, that is, when the main injector is not operating, the interference of the air from the main injector and the first combustion region of the pilot injector is prevented. Thus, high ignitability of the pilot injector and high combustion efficiency can be maintained. 
     In the present invention, it is preferable that a radially inner surface of the first inflow channel include: an inside flare portion formed in a vicinity of an exit end of the radially inner surface and configured to increase in diameter toward a downstream side; and an inside reduced-diameter portion provided upstream of the inside flare portion and configured to reduce in diameter toward the downstream side. According to this configuration, the first inflow channel is shaped so as to get close to the pilot injector once at the inside reduced-diameter portion and then widen at the inside flare portion located in the vicinity of the exit end. As a result, in the vicinity of the immediately downstream side of the exit end of the pilot injector, the fuel-air mixture of the main injector easily contacts the first combustion region, so that high combustion efficiency when the output of the gas turbine engine is intermediate can be maintained. 
     In the present invention, it is preferable that a ratio Q 1 /Q 2  that is a ratio of a flow quantity Q 1  of the air flowing through the first inflow channel to a flow quantity Q 2  of the air flowing through the second inflow channel be in a range from 3/7 to 7/3. According to this configuration, the fuel concentration does not become high locally in the air channel of the main injector. Therefore, the flame temperature at the time of the combustion can be suppressed to a low level, and the generation of the NOx can be suppressed. In addition, the damages on the wall surface by the flashback or auto ignition under high temperature and pressure can be avoided. 
     According to the fuel injector of the present invention, after the air having the major component in the axial direction in the first inflow channel and the fuel-air mixture having the major component in the radial direction in the second inflow channel meet, the air and the fuel are adequately mixed with each other in a comparatively short distance. Thus, when the output of the gas turbine engine is intermediate or high, that is, when the main injector is operating, the amount of NOx emission becomes small. In addition, since the fuel is injected only to the second inflow channel, a fuel channel and its cooling structure can be simplified. In a case where the fuel is injected to the second inflow channel from the main fuel injecting portion provided at the portion which defines a boundary between the first inflow channel and the second inflow channel, and the output of the gas turbine engine is intermediate, that is, the injection velocity of the fuel is lower than that when the output of the gas turbine engine is high, and a fuel spray penetration distance is short, the fuel having been injected only to the second inflow channel flows on a radially inward side, that is, on the pilot injector side. Therefore, the fuel gets close to the first combustion region where the combustion state is stable. As a result, the flame holding effect by the first combustion region is easily obtained. Thus, the combustion efficiency improves. Meanwhile, when the output of the gas turbine engine is high, that is, when the injection velocity of the fuel is high and the fuel spray penetration distance is long, the fuel-air mixture is generated uniformly in the entire air channel of the main injector, so that the NOx reduction is further realized. As above, by such a simple configuration, the fuel distribution suitable for the output conditions is realized, and a desired performance can be obtained. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a cross-sectional view showing a combustor of a gas turbine engine including a fuel injector according to one embodiment of the present invention. 
         FIG. 2  is a longitudinal sectional view showing the fuel injector in detail. 
         FIG. 3  is a longitudinal sectional view showing the fuel injector when viewed from an axially upstream side. 
         FIG. 4A  is a cross sectional view taken along line IV-IV of  FIG. 2 . 
         FIG. 4B  is a longitudinal sectional view showing a modification example of an outside swirler. 
         FIG. 5  is an enlarged longitudinal sectional view showing a main air channel of the fuel injector. 
         FIG. 6  is a longitudinal sectional view showing a state of the fuel injector when the output of the gas turbine engine is high or intermediate. 
         FIG. 7  is a longitudinal sectional view showing a state of the fuel injector when the output of the gas turbine engine is low. 
         FIG. 8  is an enlarged longitudinal sectional view showing the vicinity of a tip end portion of a nozzle of the fuel injector. 
         FIG. 9A  is an enlarged longitudinal sectional view showing the main air channel of the fuel injector when the output of the gas turbine engine is intermediate. 
         FIG. 9B  is a diagram showing a fuel injection state of  FIG. 9A  when viewed from a downstream side of the channel. 
         FIG. 10A  is an enlarged longitudinal sectional view showing the main air channel of the fuel injector when the output of the gas turbine engine is high. 
         FIG. 10B  is a diagram showing the fuel injection state of  FIG. 10A  when viewed from the downstream side of the channel. 
         FIG. 11  is a longitudinal sectional view showing the fuel injector according to another embodiment of the present invention in detail. 
     
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     Hereinafter, preferred embodiments of the present invention will be explained in reference to the drawings. 
       FIG. 1  shows a combustor  1  of a gas turbine engine including a fuel injector  2  according to one embodiment of the present invention. The combustor  1  mixes fuel with compressed air supplied from a compressor (not shown) of the gas turbine engine, combusts the obtained mixture, and supplies a high temperature and pressure combustion gas, generated by this combustion, to drive the turbine. 
     The combustor  1  is an annular type, and an annular outer casing  5  and an annular inner casing  7  provided inside the annular outer casing  5  constitute a combustor housing  3  including an annular internal space. The annular outer casing  5  and the annular inner casing  7  are provided coaxially with an engine rotation central axis C. In the annular internal space of the combustor housing  3 , an annular combustor liner  9  is provided coaxially with the combustor housing  3 . The combustor liner  9  is configured such that: an annular outer liner  11  and an annular inner liner  13  provided inside the annular outer liner  11  are provided coaxially with each other; and an annular combustion chamber  4  is formed in the combustor liner  9 . A plurality of fuel injectors  2  configured to inject the fuel to the combustion chamber  4  are arranged on an upstream wall of the combustor liner  9  coaxially with the engine rotation central axis C, that is, in a circumferential direction of the combustor liner  9  at regular intervals. Each of the fuel injectors  2  includes a pilot injector  6  and a main injector  8 . The main injector  8  is provided coaxially with a central axis C 1  of the pilot injector  6  so as to surround an outer periphery of the pilot injector  6  and generates a fuel-air mixture. Each fuel injector  2  is supported on the combustor housing  3  by a stem portion  27  attached to the combustor housing  3  by fastening members  19 . An ignition plug  1 G configured to perform ignition is provided so as to extend in a radial direction of the combustor liner  9  and penetrate the outer casing  5  and the outer liner  11 , and a tip end of the ignition plug IG is located close to the fuel injector  2 . 
     Compressed air CA is supplied from the compressor through an annular air induction passage  21  to the annular internal space of the combustor housing  3 . This compressed air CA is supplied to the fuel injector  2  and is also supplied to the combustion chamber  4  through a plurality of air introducing holes  23  formed on the outer liner  11  and inner liner  13  of the combustor liner  9 . The stem portion  27  forms a fuel pipe unit U. The fuel pipe unit U includes a first fuel supply system F 1  configured to supply the fuel to the pilot injector  6  and a second fuel supply system F 2  configured to supply the fuel to the main injector  8 . 
     A downstream portion of the fuel injector  2  is supported by an outer support  29  via a flange  25 A and a supporting body  25 B. The flange  25 A and the supporting body  25 B are provided on an outer peripheral portion of the downstream portion of the fuel injector  2 , and the outer support  29  is formed integrally with the outer liner  11 . The outer liner  11  is supported by the outer casing  5  using a liner fixing pin P. The outer support  29  projects in a radially inward direction of the fuel injector  2  and is protected from high temperature of the combustion chamber  4  by a heat shield  17  internally fitted in the outer support  29 . A first-stage nozzle TN of the gas turbine engine is connected to a downstream end portion of the combustor liner  9 . 
       FIG. 2  is a longitudinal sectional view showing the fuel injector  2  of  FIG. 1  in detail. The pilot injector  6  provided at a center portion of the fuel injector  2  includes a central body  10 , an inside tubular body  12 , an outside cylindrical body  14 , and an inner shroud  15 . The central body  10  is provided on the central axis C 1 . The inside tubular body  12  is provided coaxially with the central body  10 , is formed integrally with the stem portion  27 , and forms a main body of the pilot injector  6 . The outside cylindrical body  14  is provided outside the inside tubular body  12  and coaxially with the inside tubular body  12 . The inner shroud  15  is an annular dividing wall provided outside the outside cylindrical body  14  and coaxially with the outside cylindrical body  14 . The inner shroud  15  defines a boundary between the pilot injector  6  and the main injector  8 . A venturi nozzle-shaped pilot outer peripheral nozzle  18  is formed at a downstream portion of an inner peripheral surface of the inner shroud  15 . As shown in  FIG. 3 , except for a portion where the pilot outer peripheral nozzle  18  is formed, the stem portion  27  is formed in a long and thin shape having a width smaller than an inner diameter of a below-described inside swirler  30 . 
     The inside tubular body  12  of the pilot injector  6  shown in  FIG. 2  is supported by a base portion  19  ( FIG. 1 ) connected to the fuel pipe unit U ( FIG. 1 ) of the first fuel supply system F 1 . A strut  28  configured to support the central body  10  on the inside tubular body  12  is fixed inside the inside tubular body  12 . An annular center nozzle  20  is formed between the central body  10  and the inside tubular body  12  and forms an inside air channel concentrically with the central axis C 1 . The diameter of the central body  10  gradually increases on a downstream side of the strut  28  such that the air flow in the center nozzle  20  accelerates toward the downstream side. An annular pilot fuel channel  22  configured to communicate with the first fuel supply system F 1  is formed in a downstream portion of the inside tubular body  12 . An outside air channel  24  is formed between the inside tubular body  12  and the outside cylindrical body  14 , and a supplemental air channel  26  is formed between the outside cylindrical body  14  and the inner shroud  15 . 
     The inside swirler  30  is provided upstream of the outside air channel  24 , and an outside swirler  32  is provided upstream of the supplemental air channel  26 . The inside swirler  30  swirls the air around the central axis C 1  of the pilot injector  6 . The outside swirler  32  is a diffuser type which more strongly swirls the air than the inside swirler  30 . To be specific, swirling directions of the swirlers  30  and  32  are the same as each other, and a swirling angle of the outside swirler  32  is larger than that of the inside swirler  30 . The swirling angle is an exit attachment angle of a blade with respect to a flat surface including the central axis C 1 . As above, the pilot injector  6  includes the outside air channel  24 , the supplemental air channel  26 , the central body  10 , the strut  28 , and the swirlers  30  and  32 . It is preferable that the swirling angle of air jet that is air flow ejected from the center nozzle  20  be less than 10° at an exit of the center nozzle. For example, in a case where air flow field on an upstream side of the fuel injector  2  is stable or in a case where there are limitations regarding manufacture, the central body  10  and the strut  28  may be simplified by devising an inside shape of the inside tubular body  12 . The exit swirling angle of the inside swirler  30  is, for example, 30° and preferably 20 to 50°. The exit swirling angle of the outside swirler  32  is, for example, 50° and preferably 40 to 60°. 
     As shown in  FIG. 4A , regarding the outside swirler  32 , an entrance angle (angle of a front edge with respect to the axial direction) θ 1  of each vane (blade) is set to be larger than an exit angle (angle of a rear edge with respect to the axial direction) θe, and each air channel widens toward the downstream side. To be specific, the outside swirler  32  includes a plurality of diffuser vanes  32   a , which are smoothly curved in the circumferential direction such that an effective cross-sectional area of the air channel in a direction perpendicular to the air flow becomes large. As shown in  FIG. 4B , the outside swirler  32  may include a plurality of diffuser vanes  32   b , each of whose vane height (radial height of the channel) increases toward the downstream side so that the air channel widens. The outside swirler  32  may be a normal swirler including a plurality of vanes configured such that the cross-sectional area of the air channel in the direction perpendicular to the air flow is constant or decreases from the entrance toward the exit. 
     The pilot fuel channel  22  of  FIG. 2  is formed on the inside tubular body  12  and is located between the center nozzle  20  and the outside air channel  24 . The fuel from the first fuel supply system F 1  is injected from a pilot fuel injecting portion  22   a , formed at a downstream end of the pilot fuel channel  22 , toward the center nozzle. The pilot fuel injecting portion  22   a  is a pre-filmer type including an annular opening through which the fuel is injected in an annular film shape. Each of a downstream portion  16   b  of an outer peripheral portion  16  of the inside tubular body  12  and a downstream portion  14   b  of the outside cylindrical body  14  is shaped to taper toward the downstream side. The outer peripheral portion  16  is formed at an outer peripheral side of the pilot fuel channel  22 . With this, the pilot fuel channel  22  and the outside air channel  24  incline by the downstream portions  16   b  and  14   b  toward the inside air channel  20  in the radially inward direction. A downstream end  16   a  of the outer peripheral portion  16  of the inside tubular body  12  and a downstream end  14   a  of the outside cylindrical body  14  are located on a downstream side of the vicinity of the exit of the center nozzle  20 . To be specific, the pilot fuel injecting portion  22   a  that is the downstream end of the pilot fuel channel  22  and an exit end  24   a  of the outside air channel  24  face the vicinity of an exit  20   a  of the center nozzle  20 . 
     The pilot outer peripheral nozzle  18  is formed by an inner peripheral surface of a downstream portion of the inner shroud (dividing wall)  15 , the downstream portion being located downstream of the outside swirler  32 . The pilot outer peripheral nozzle  18  includes a pilot flare portion  18   b  and a pilot reduced-diameter portion  18   c . The pilot flare portion  18   b  is provided in the vicinity of an exit end  18   a  of the pilot outer peripheral nozzle  18  and increases in diameter toward the downstream side. The pilot reduced-diameter portion  18   c  is provided upstream of the pilot flare portion  18   b  and reduces in diameter toward the downstream side. To be specific, the inner diameter of the pilot outer peripheral nozzle  18  becomes minimum at a narrow portion  18   d  that is a boundary between the pilot flare portion  18   b  and the pilot reduced-diameter portion  18   c . As above, the pilot outer peripheral nozzle  18  is shaped to narrow once and then widens toward the downstream side. The pilot flare portion  18   b  inclines at a tilt angle θ 1  with respect to the direction of the central axis C 1 . In the present embodiment, the tilt angle θ 1  is 20° and preferably 15 to 30°. As long as the tilt angle θ 1  is in this range, a pilot combustion region A 1  that is a below-described first combustion region can appropriately spread in a radially outward direction. Thus, high combustion efficiency can be maintained. 
     The downstream end  16   a  of the outer peripheral portion  16  of the inside tubular body  12  and the downstream end  14   a  of the outside cylindrical body  14  are located slightly upstream of the narrow portion  18   d  of the pilot outer peripheral nozzle  18 . As described above, the downstream portion  14   b  of the outside cylindrical body  14  tapers toward the downstream side. To correspond to this tapered shape, the pilot outer peripheral nozzle  18  includes the pilot reduced-diameter portion  18   c  which narrows once toward the downstream side. With this, the channel area of the supplemental air channel  26  does not drastically increase on a radially outer side of the downstream portion  14   b  of the outside cylindrical body  14 . Therefore, the separation of the air flow along an outer peripheral surface of the outside cylindrical body  14  can be suppressed, and the outer peripheral surface of the outside cylindrical body  14  can be prevented from burning out by the combustion gas in the combustion chamber  4 . 
     The air having flowed through the pilot injector  6  except for the air jet flowing through the center nozzle  20  diffuses toward an outer peripheral side by the swirling. Regarding the air flow immediately after the exit of the fuel injector  2 , negative pressure is generated in the vicinity of the central axis C 1  by strong swirling of the air mainly from the main injector  8 , and a radially inward pressure gradient and a radially outward centrifugal force are balanced. However, the strong swirling air flow from the main injector  8  spreads, decays, and weakens as it flows toward the downstream side. Therefore, the pressure in the vicinity of the central axis C 1  gradually recovers toward the downstream side. On this account, on the central axis C 1  located downstream of the fuel injector  2 , an adverse pressure gradient is generated, that is, the pressure is higher on the downstream side than on the upstream side. As a result, a recirculation region X ( FIG. 1 ) in which reverse flow from the downstream side toward the upstream side occurs is formed. 
     Meanwhile, the pilot fuel injecting portion  22   a  injects fuel F to the air flowing through the center nozzle  20 . The air jet from the center nozzle  20  flows substantially straight in an axially downstream direction, is mixed with ambient air in the recirculation region X, and disappears. Then, the fuel in the form of a mist reaches a center portion of the recirculation region X and vaporizes and combusts in the recirculation region X to form the pilot combustion region A 1 . If the momentum of the air jet having been emitted from the center nozzle  20  is large, a concave portion Xa may be formed on the recirculation region X in a process in which the air jet gets into the recirculation region X and disappears. 
     The air having flowed through the pilot injector  6  spreads in the radially outward direction while swirling along the pilot flare portion  18   b . With this, the recirculation region X ( FIG. 1 ) formed by the air from the pilot injector  6  can moderately spread in the radially outward direction. The pilot combustion region A 1  ( FIG. 6 ) is formed by injecting the fuel from the pilot injector  6  to the moderately spread recirculation region X. Therefore, high combustion efficiency can be maintained even when the output of the gas turbine engine is low. 
     Referring back to  FIG. 2 , the main injector  8  fitted on the outer periphery of the pilot injector  6  will be explained. The main injector  8  includes a ring portion  34  and an outer shroud  36 . The ring portion  34  is provided on a radially outer side of the inner shroud  15  and coaxially with the inner shroud  15  and is formed integrally with the stem portion  27 . The outer shroud  36  is provided on an axially downstream side of the ring portion  34 . An annular first air channel  38  is formed between the inner shroud  15  and the ring portion  34 . The annular first air channel  38  is an inflow channel through which the air having a major flow component in the axial direction of the fuel injector  2  is taken, that is, the air is taken in a state where an axial flow component of the air in the vertical cross section including the central axis C 1  in  FIG. 2  is larger than a radial flow component thereof. An annular second air channel  42  is formed between the ring portion  34  and the outer shroud  36 . The second air channel  42  is an inflow channel through which the air having a major flow component in the radial direction of the fuel injector  2  is taken, that is, the air is taken in a state where the radial flow component of the air in the vertical cross section including the central axis C 1  in  FIG. 2  is larger than the axial flow component thereof. To be specific, a downstream end surface of the ring portion  34  forms one side wall of the second air channel  42 , and an upstream portion of an inner peripheral surface  37  of the outer shroud  36  forms another side wall of the second air channel  42 . The ring portion  34  defines a boundary between the first air channel  38  and the second air channel  42 . 
     The first air channel  38  extends from an entrance of a below-described main inside swirler  46  up to an inner peripheral rear end edge  34   a  of the ring portion  34 . The second air channel  42  extends from an entrance of a below-described main outside swirler  48  up to the inner peripheral rear end edge  34   a  of the ring portion  34 . A premixing chamber  58  where the air flow from the first air channel  38  and the air flow from the second air channel  42  meet is located downstream of these two channels  38  and  42  and is formed between the outer shroud  36  and the inner shroud  15 . A main channel  56  is constituted by the first air channel  38 , the second air channel  42 , and the premixing chamber  58 . 
     An annular main fuel injecting portion  40  connected to the second fuel supply system F 2  is formed in the ring portion  34  which defines a boundary between the first air channel  38  and the second air channel  42 . When the output of the gas turbine engine is low, the fuel is not supplied to the main injector  8 . Only when the output of the gas turbine engine is intermediate or high, the fuel is supplied from the second fuel supply system F 2  to the main injector  8 . The main fuel injecting portion  40  injects the fuel only to the second air channel  42 . The injected fuel is mixed with the air flow from the main outside swirler  48  and the air flow from the main inside swirler  46  in the premixing chamber  58 . Thus, a pre-mixture is produced. The pre-mixture is supplied to and combusted in the combustion chamber  4 . With this, a premix combustion region A 2  shown in  FIG. 6  is formed. 
     As shown in  FIG. 7 , when the output of the gas turbine engine is low, that is, when the fuel is not supplied to the main injector  8 , a main air flow E having flowed through the swirlers  46  and  48  is supplied to the combustion chamber  4  through the premixing chamber  58 . 
     A downstream portion of the inner peripheral surface  37  of the outer shroud  36  shown in  FIG. 2  forms a main exit flare  43  of the main injector  8 . The main exit flare  43  widens from a base end portion  43   a  that is an upstream end thereof toward an exit end  43   b  that is a downstream end thereof. The base end portion  43   a  is a portion which projects most in the radially inward direction. To be specific, an outer peripheral surface of the main channel  56  that is the air channel of the main injector  8  widens toward an exit end thereof. The vicinity of the exit end  43   b  of the main exit flare  43  inclines at a tilt angle θ 2  with respect to the central axis C 1 . With this, as shown in FIG.  7 , the main air flow E spreads in the radially outward direction and can be prevented from significantly interfering with the pilot combustion region A 1  when the output of the gas turbine engine is low. The tilt angle θ 2  of the main exit flare  43  shown in  FIG. 2  is about 35° and preferably 20 to 50°. As long as the tilt angle θ 2  is in this range, the recirculation region X can adequately spread in the radially outward direction and the flame holding performance can be improved while preventing the interference with the pilot combustion region A 1 . 
     As clearly shown in  FIG. 5 , the second air channel  42  is smoothly curved toward the combustion chamber  4  as it extends toward the downstream side. Air CA 2  from the exit of the second air channel  42  and air CA 1  from the exit of the first air channel  38  meet at an intersection angle α at an intersection point J of the premixing chamber  58 . The intersection angle α is preferably in a range from 40 to 80° in order to generate strong turbulence of the air flow when the air CA 1  from the exit of the first air channel  38  and the air CA 2  from the exit of the second air channel  42  meet. 
     A plurality of main fuel injection holes  44  are formed on the main fuel injecting portion  40  so as to be located at a portion of the second air channel  42  and arranged in the circumferential direction at regular intervals, the portion of the second air channel  42  being located upstream of the intersection point J. The plurality of main fuel injection holes  44  inject the fuel to the second air channel  42  from the upstream side (left side in  FIG. 5 ) to the downstream side (right side in  FIG. 5 ) in the axial direction. The main fuel injection holes  44  may be arranged at irregular intervals. The main fuel injection holes  44  are open on an axially upstream wall surface of the second air channel  42  and inject the fuel by a plane jet method. Preferably, five or more main fuel injection holes  44  are arranged in the circumferential direction. An angle β between the flow of the air of the second air channel  42  and the flow of the fuel injected from the main fuel injection holes  44  is substantially 90° in the vicinity of the main fuel injection holes  44 . The angle β is preferably 70 to 90° in order to promote the atomization of the fuel by the air flow. 
     The fuel-air mixture generated by injecting the fuel from the main fuel injection holes  44  toward the air flow CA 2  in the second air channel  42  meets the air CA 1  flowing in the axial direction in the first air channel  38 . Since the fuel-air mixture meets the air CA 1  at a certain angle, the air turbulence further promotes the mixing of the air and the fuel. After the fuel-air mixture and the air CA 1  meet, the fuel-air mixture is further mixed in the premixing chamber  58  and then sprayed to the combustion chamber  4 . 
     Here, a ratio Q 1 /Q 2  is preferably 3/7 to 7/3, the ratio Q 1 /Q 2  being a ratio of a flow quantity Q 1  of the air CA 1  flowing through the first air channel  38  to a flow quantity Q 2  of the air CA 2  flowing through the second air channel  42 . If the flow quantity ratio is out of this range, the fuel and the air are unlikely to be mixed with each other, and the generation of the NOx may not be adequately suppressed. In addition, the possibility of the damages on the wall surface by flashback or auto ignition under high temperature and pressure may increase. 
     The main inside swirler  46  that is a first swirling unit is attached to an entrance of the first air channel  38 . The main outside swirler  48  that is a second swirling unit is attached to an entrance of the second air channel  42 . The main outside swirler  48  includes a first swirler  50  and a second swirler  52 , which are swirling portions arranged in the axial direction of the main injector  8 . Swirl blades of the first swirler  50  provided close to the main fuel injection holes  44  is set such that the air having passed through the first swirler  50  simply flows straight in the substantially radially inward direction. Swirl blades of the second swirler  52  provided away from the main fuel injection holes  44  is set such that the air having passed through the second swirler  52  is swirled around the central axis C 1 . 
     When the output of the gas turbine engine is intermediate, that is, when the flow quantity of the fuel from the main fuel injection holes  44  is small and the momentum of the fuel of the main fuel inject holes  44  is small, most of the injected fuel just reaches the air flow having flowed through the first swirler  50  in the radially inward direction. Therefore, the fuel is not diffused in the radial direction by the swirling of the second swirler  52  and flows in the radially inward direction. Thus, the fuel-air mixture is generated on a radially inward side of the main channel  56 . 
     Meanwhile, when the output of the gas turbine engine is high, that is, when the flow quantity of the fuel from the main fuel injection holes  44  is large and the momentum of the fuel of the main fuel injection holes  44  is large, a part of the injected fuel flows in the radially inward direction together with the air flow in the radially inward direction as with when the output of the gas turbine engine is intermediate, but the remaining fuel reaches the swirl flow having flowed through the second swirler  52  and generates the fuel-air mixture, which flows in the radially outward direction together with the swirl flow. As a result, when the output of the gas turbine engine is high, the fuel-air mixture is generated uniformly in the entire main channel  56 . 
     The main outside swirler  48  may be a single swirler. In this case, the main outside swirler  48  includes swirl blades, each of which is formed in such a twisted shape that: the air flowing through a portion, closest to the main fuel injection holes  44 , of the swirl blade flows straight in the substantially radially inward direction; and the swirling component increases as the portion where the air flows is away from the main fuel injection holes  44 . It should be noted that each of the first swirler  50  and the second swirler  52  may be constituted by a plurality of swirlers arranged in the axial direction. 
     A main inside flare portion  54   b  which increases in diameter toward the downstream side is formed in the vicinity of an exit end  54   a  of a radially inner surface  54  of the first air channel  38  shown in  FIG. 2 , and a main inside reduced-diameter portion  54   c  which reduces in diameter toward the downstream side is formed upstream of the main inside flare portion  54   b . The exit end  54   a  of the radially inner surface  54  of the first air channel  38  is located slightly downstream of the base end portion  43   a  of the main exit flare  43 . 
     As shown in  FIG. 7 , a virtual extended inner peripheral surface VP 1  and a virtual extended outer peripheral surface VP 2  gradually separate from each other as they extend in the downstream direction. The virtual extended inner peripheral surface VP 1  is a surface extending from the exit end  18   a  of the inner peripheral surface of the inner shroud  15  in the downstream direction, and the virtual extended outer peripheral surface VP 2  is a surface extending from the exit end  54   a  of the outer peripheral surface of the inner shroud  15  in the downstream direction. The virtual extended inner peripheral surface VP 1  and the virtual extended outer peripheral surface VP 2  may be arranged in parallel with each other. In other words, these surfaces VP 1  and VP 2  may be arranged in any manner as long as these surfaces VP 1  and VP 2  do not intersect with each other on a downstream side of the pilot outer peripheral nozzle  18 . 
     A radial thickness of an exit end surface  15   a  of the inner shroud  15  is set to be thin. As shown in  FIG. 8 , a ratio T/Dp is preferably in a range from 0.02 to 0.15, the ratio T/Dp being a ratio of a distance T between the exit end  18   a  of the inner peripheral surface of the inner shroud  15  and the exit end  54   a  of the outer peripheral surface of the inner shroud  15 , that is, a radial width T of the exit end surface  15   a  of the inner shroud  15  to an inner diameter Dp of the exit end  18   a  of the pilot outer peripheral nozzle  18 . If the ratio T/Dp is less than 0.02, the main air flow E and the pilot combustion region A 1  in  FIG. 7  are too close to each other and strongly interfere with each other. This deteriorates the combustion efficiency, ignitability, and flame holding performance of the pilot injector  6  when the output of the gas turbine engine is low. In contrast, if the ratio T/Dp exceeds 0.15, the pilot combustion region A 1  and the premix combustion region A 2  that is a second combustion region in  FIG. 6  are largely spaced apart from each other in the radial direction. This deteriorates the flame holding effect obtained by the pilot flame of the main injector  8  when the output of the gas turbine engine is intermediate, so that the combustion efficiency decreases. 
     The exit end  18   a  of the pilot outer peripheral nozzle  18  of  FIG. 8  is located upstream of the exit end  43   b  of the main exit flare  43 . Specifically, a ratio W/Dm is preferably 0.25 or lower, and more preferably in a range from 0.1 to 0.25, the ratio W/Dm being a ratio of an axial distance W between the exit ends  18   a  and  43   b  to an inner diameter Dm of the exit end  43   b  of the main exit flare  43 . If the ratio W/Dm is less than 0.1, the flame holding effect obtained by the pilot flame deteriorates. Thus, the improvement effect of the combustion efficiency slightly decreases. However, if the combustion efficiency is adequately high, the exit end  18   a  of the pilot outer peripheral nozzle  18  and the exit end  43   b  of the main exit flare  43  may coincide with each other in the axial direction. Even if the ratio W/Dm is set to more than 0.25, the improvement of the flame holding effect is limited. 
     According to the above configuration, when the output of the gas turbine engine is low, the fuel is supplied from the first fuel supply system F 1  only to the pilot injector  6  in the fuel injector  2  in  FIG. 2 . The air having flowed through the pilot injector  6  except for the air having flowed through the center nozzle  20  diffuses toward the outer peripheral side by the swirling. The pilot fuel injecting portion  22   a  injects the fuel F to the air in the center nozzle  20 . The air jet having been emitted from the center nozzle  20  flows substantially straight in the axially downstream direction, is mixed with the ambient air in the recirculation region X, and disappears. Then, most of the fuel in the form of a mist reaches the center portion of the recirculation region X and vaporizes and combusts in the recirculation region X. Thus, the interfere of the fuel F with the main air flow by the diffusing of the fuel F toward the outer peripheral side is suppressed. As a result, the combustion efficiency, ignitability, and flame holding performance of the pilot injector  6  when the output of the gas turbine engine is low can be improved. 
     Moreover, the virtual extended inner peripheral surface VP 1  extending from the exit end  18   a  of the inner peripheral surface of the inner shroud  15  in the downstream direction and the virtual extended outer peripheral surface VP 2  extending from the exit end  54   a  of the outer peripheral surface of the inner shroud  15  in the downstream direction gradually separate from each other as they extend in the downstream direction. Therefore, the interference of the main air flow E with the pilot combustion region A 1  can be suppressed, and the ignitability, flame holding performance, and combustion efficiency of the pilot injector  6  when the output of the gas turbine engine is low can be further improved. 
     The outside swirler  32  provided on a radially outer side of the inside swirler  30  includes the diffuser vanes  32   a  ( FIGS. 4A and 4B ) formed such that the air channel widens toward the downstream side. As above, in a case where the center nozzle  20  is provided in the vicinity of the central axis C 1  of the pilot injector  6 , and the momentum of the air jet having been emitted from the center nozzle  20  is large, as shown in  FIG. 8 , the recirculation region X is shaped to be concave in the vicinity of the central axis C 1  toward the downstream side. This may deteriorate the combustion efficiency, ignitability, and flame holding performance of the pilot injector  6 . Even in this case, if the diffuser-type outside swirler  32  is provided on the radially outer side of the inside swirler  30 , the air velocity at the exit of the outside swirler  32  becomes lower than that of a normal swirler. Therefore, as shown by a broken line X 1  in  FIG. 8 , the recirculation region X spreads toward the upstream side in the vicinity of the exit of the outside swirler  32 . As a result, the flame of the pilot injector  6  stabilizes, so that the combustion efficiency, ignitability, and flame holding performance of the pilot injector  6  can be prevented from being deteriorated. 
     Further, the reverse flow region can be moderately spread in the radially outward direction by swirl flow S generated by the outside swirler  32  configured to generate a swirl velocity component stronger than that of the inside swirler  30  of the pilot injector  6  in  FIG. 7 . 
     Since the pilot fuel injecting portion  22   a  is a pre-filmer type configured to inject the fuel in an annular film shape, a shear surface area of the air with respect to the fuel increases, and the atomization of the fuel is promoted. As a result, the NOx reduction when the output of the gas turbine engine is low can be realized. 
     When the output of the gas turbine engine is intermediate or high, the fuel is supplied to both the pilot injector  6  and the main injector  8 . As shown in  FIG. 5 , in the main injector  8 , the fuel F is injected to the second air channel  42 , and the air CA 2  having the major component in the radial direction and the fuel F are mixed with each other. Next, fuel-air mixture M 1  and the air CA 1  flowing through the first air channel  38  and having the major component in the axial direction meet in the premixing chamber  58  at a certain angle. With this, the mixing of the fuel and the air is further promoted, so that the air and the fuel are adequately mixed with each other in a comparatively short distance, and the NOx reduction can be realized. In addition, since the fuel is injected only to the second air channel  42 , a fuel channel and its cooling structure can be simplified. 
     The main fuel injecting portion  40  of  FIG. 2  injects the fuel F toward the second air channel  42  from a portion K which defines a boundary between the first air channel  38 . and the second air channel  42 . Therefore, when the output of the gas turbine engine is intermediate, that is, when the momentum of the injection of the main fuel is small, the injected fuel just reaches a region close to the injection holes  44 , as compared to when the output of the gas turbine engine is high, that is, when the momentum thereof is large. As a result, the fuel is injected mainly to a position close to the main fuel injecting portion  40  in the air flow of the second air channel  42 . Therefore, when the air flow of the second air channel  42  meets the air flow of the first air channel  38  to be changed to the air flow in the axial direction and is then injected to the combustion chamber  4 , the fuel in the form of a mist flows on a radially inward side as compared to when the output of the gas turbine engine is high. To be specific, when the output of the gas turbine engine is intermediate, the main fuel in the form of a mist gets close to the pilot combustion region A 1  where the flame is stable in  FIG. 6 , as compared to when the output of the gas turbine engine is high. As a result, the flame holding effect by the flame in the pilot combustion region A 1  can be easily obtained. Thus, the combustion efficiency improves. Moreover, the portion K which defines a boundary between the first air channel  38  and the second air channel  42  can generally secure a space widely in many cases. Therefore, a structure, such as a cooling structure for preventing coking, in the main fuel injecting portion  40  can be easily, spatially arranged. 
     The main inside swirler  46  is attached to the entrance of the first air channel  38 , and the main outside swirler  48  is attached to the entrance of the second air channel  42 . By the first swirler  50 , located close to the main fuel injection holes  44 , of the main outside swirler  48 , as shown in  FIG. 9A , a region M where the air flows straight in the substantially radially inward direction is formed in the vicinity of the main fuel injection holes  44  in the second air channel  42 . Meanwhile, a swirling region where the air flows in the radially outward direction by the second swirler  52  is formed at a position away from the main fuel injection holes  44 . When the output of the gas turbine engine is intermediate, that is, when the flow quantity of the fuel is small and the injection velocity of the fuel is low, most of the fuel F injected from the main fuel injection holes  44  do not reach the strong swirl flow generated by the second swirler  52 , stays in the flow moving straight in the radially inward direction by the first swirler  50 , and flows in the radially inward direction. Therefore, fuel-air mixture Y 1  is generated on the inner side of the main channel  56 . As a result, the fuel-air mixture Y 1  which is comparatively thick is ejected to a position close to the pilot combustion region A 1  ( FIG. 6 ). Thus, the combustion efficiency when the output of the gas turbine engine is intermediate further improves by the flame holding effect obtained by the pilot combustion region A 1 . 
     When the output of the gas turbine engine is high, that is, when the flow quantity of the fuel is large and the injection velocity of the fuel is high, as shown in  FIGS. 10A and 10B , a part of the fuel F having been injected from the main fuel injection holes  44  stays in the flow moving straight in the radially inward direction by the first swirler  50  and forms the fuel-air mixture Y 1  flowing in the radially inward direction. Meanwhile, the remaining fuel flows with the swirl flow generated by the second swirler  52  and forms fuel-air mixture Y 2  flowing in the radially outward direction. As a result, when the output of the gas turbine engine is high, the uniform fuel-air mixture Y 2  is generated in the entire main channel  56 . Thus, the NOx reduction can be realized. As above, by such a simple configuration, fuel distribution suitable for output conditions is realized, and a desired performance can be obtained. 
     As shown in  FIG. 6 , the exit end  18   a  of the pilot outer peripheral nozzle  18  is located upstream of the exit end  43   b  of the main exit flare  43 . Therefore, a pre-mixture M 2  of the main channel  56  promptly contacts the pilot combustion region A 1  in the vicinity of the exit of the pilot outer peripheral nozzle  18 , so that the combustion efficiency when the output of the gas turbine engine is intermediate further improves. 
     As shown in  FIG. 8 , in a case where the ratio W/Dm is set to 0.25 or less, the ratio W/Dm being a ratio of the axial distance W between the exit end  18   a  of the pilot outer peripheral nozzle  18  and the exit end  43   b  of the main exit flare  43  to the inner diameter Dm of the exit end  43   b  of the main exit flare  43 , the main pre-mixture promptly contacts the pilot combustion region A 1  ( FIG. 6 ) in the vicinity of the exit end  18   a  of the pilot outer peripheral nozzle  18 . Therefore, the flame holding effect of the main injector  8  by the pilot flame when the output of the gas turbine engine is intermediate becomes large. Thus, the combustion efficiency further improves. 
     Since the ratio T/Dp is 0.02 to 0.15, the ratio T/Dp being a ratio of the radial width T of the exit end surface  15   a  of the annular inner shroud  15  which defines a boundary between the pilot injector  6  and the main injector  8  to the inner diameter Dp of the exit end  18   a  of the pilot outer peripheral nozzle  18 , the main pre-mixture promptly contacts the pilot combustion region A 1  in the vicinity of a region located downstream of the exit end  18   a  of the pilot outer peripheral nozzle  18 . Therefore, the combustion efficiency when the output of the gas turbine engine is intermediate can be further improved. 
     As shown in  FIG. 6 , the radially inner surface  54  of the first air channel  38  of the main injector  8  is shaped so as to get close to the pilot injector  6  once at the inside reduced-diameter portion  54   c  and then widen at the inside flare portion  54   b  located in the vicinity of the exit end  54   a . With this, in the vicinity of the region located downstream of the exit end  18   a  of the pilot outer peripheral nozzle  18 , the pre-mixture of the main injector  8  tends to contact the pilot combustion region A 1 , so that high combustion efficiency when the output of the gas turbine engine is intermediate can be maintained. Meanwhile, when the output of the gas turbine engine is low, on the downstream side of the exit end  54   a  of the radially inner surface  54  of the first air channel  38  of the main injector  8 , the air having flowed through the main injector  8  is adequately diffused in the radially outward direction by the inside flare portion  54   b . Thus, the interference of the air having flowed through the main injector  8  with the pilot combustion region A 1  of the pilot injector  6  can be suppressed, so that high combustion efficiency when the output of the gas turbine engine is low can be maintained. 
     Further, since the main exit flare  43  of the main injector  8  is shaped to widen toward its exit end, the air from the main injector  8  spreads in the radially outward direction. Therefore, the recirculation region X can moderately spread in the radially outward direction while avoiding the interference of the air from the main injector  8  with the air from the pilot injector  6 . Thus, high combustion efficiency can be obtained even when the output of the gas turbine engine is low. 
     In addition, since the ratio Q 1 /Q 2  is in a range from 3/7 to 7/3, the ratio Q 1 /Q 2  being a ratio of the flow quantity Q 1  of the air flowing through the first air channel  38  to the flow quantity Q 2  of the air flowing through the second air channel  42 , the flow quantity ratio does not become unbalanced. As a result, the fuel concentration does not become high locally. On this account, the flame temperature at the time of the combustion can be suppressed to a low level, and the generation of the NOx can be suppressed. In addition, the damages on the wall surface by the flashback or auto ignition under high temperature and pressure can be avoided. 
     In the above embodiment, the pilot fuel injecting portion  22   a  shown in  FIG. 2  is a pre-filmer type configured to inject the fuel in an annular film shape. However, the present embodiment is not limited to this. For example, as shown in  FIG. 11 , a plane jet type pilot fuel injecting portion  22   b  may be used. The pilot fuel injecting portion  22   b  is provided with a plurality of small holes through which the fuel F is injected in the radially inward direction, the plurality of small holes being arranged at regular intervals in the circumferential direction. With this, the fuel F is supplied in the radial direction to the center nozzle  20  from the plurality of small holes arranged in the circumferential direction. 
     The foregoing has explained a preferred embodiment of the present invention in reference to the drawings. However, various additions, modifications, and deletions may be made within the spirit of the present invention. Therefore, such modified embodiments are included within the range of the present invention. 
     As this invention may be embodied in several forms without departing from the spirit of essential characteristics thereof, the present embodiments are therefore illustrative and not restrictive, since the scope of the invention is defined by the appended claims rather than by the description preceding them, and all changes that fall within metes and bounds of the claims, or equivalence of such metes and bounds thereof are therefore intended to be embraced by the claims.