JP2008298351A - Combustion device for gas turbine engine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a combustion device for a turbine engine capable of uniforming a distribution of a flame temperature in a combustion chamber by providing the air flowing into an annular swirler with the uniform distribution of flow velocity in the entire circumferential direction in a combined combustion method. <P>SOLUTION: This combustion device for a gas turbine engine comprises an annular heat chamber 9 forming the combustion chamber 12, a plurality of fuel injection units 2 having premixed air-fuel supply portions 4 for supplying the premixed air-fuel of fuel F and air CA to form premixed air-fuel combustion region 51 in the combustion chamber 12, a fuel supply unit 20 for supplying the fuel F to the premixed air-fuel supply portions 4, and an annular air inlet 14a for introducing the compressed air CA from a compressor to an upstream side of the fuel injection units 2. The fuel injection units 2 have the swirler for swirling the air, fixed to an air intake, and a circumferential width VWH of a front edge portion in a high dynamic pressure area HA opposed to the air inlet 14a, of a plurality of blades 48 arranged in the circumferential direction of the swirler 43, is determined to be larger than that outside of the area. <P>COPYRIGHT: (C)2009,JPO&INPIT

Description

  The present invention relates to a combustion apparatus for a gas turbine engine having a fuel injection structure of a combined combustion system that combines two combustion systems of a diffusion combustion system and a lean premixed combustion system.

In a gas turbine engine, in consideration of environmental protection, combustion and strict environmental standards are provided with respect to the composition of the exhaust gas discharged by the nitrogen oxides (hereinafter, referred to as N0 _X) to reduce the harmful substances such as It has been demanded. On the other hand, for large gas turbines and aircraft engines, pressure ratios tend to be set higher due to demands for lower fuel consumption and higher output, and accordingly, higher temperatures and higher pressures are being introduced at the combustion equipment inlet. Since the combustion temperature tends to increase due to the increase in the inlet temperature of the combustion apparatus, there is a concern that it may cause N0 _X to increase rather.

Therefore, in recent years, a lean pre-mixture combustion system that can effectively reduce the N0 _X emissions, the composite combustion system combining the combustion method of the two systems of excellent diffusion combustion system ignition performance and the flame holding performance are proposed ing. In the lean premixed combustion method, air and fuel are premixed and burned as a mixture with uniform fuel concentration, so there is no combustion region where the flame temperature is locally high, and the fuel is diluted. by the ability to lower the flame temperature in whole, although there is an advantage that can be effectively reduced N0 _X emissions, admixing uniformly a large amount of air and fuel, very local fuel concentration in the combustion region There is a problem that the fuel becomes thinner and lowers the combustion stability particularly at low load. On the other hand, the diffusion combustion method has an advantage that the flame holding performance is excellent because the fuel and air are burned while being diffused and mixed, so that the blowout hardly occurs even at a low load. Therefore, the combined combustion system can maintain the combustion stability by diffusion combustion at the time of start-up and low load, and can reduce the amount of N0 _X generated by lean premixed gas combustion at high load.

  The combustion apparatus based on the combined combustion system forms a fuel injection portion for spraying fuel so as to form a diffusion combustion area based on a diffusion combustion system in the combustion chamber, and a premix combustion area based on a lean premixed gas combustion system within the combustion chamber. A premixed gas supply unit that supplies a premixed gas of fuel and air is provided. This combustion apparatus supplies fuel only from the fuel injection section at the time of start-up or low load, and supplies fuel from the premixed gas supply section in addition to the fuel injection section at high load (Patent Document 1, Patent Document 1). 2).

JP 2002-115847 A JP 2002-139221 A

  Meanwhile, high-temperature and high-pressure compressed air supplied from the compressor is introduced from the annular diffuser to the upstream side of the annular combustion cylinder at a flow rate of about 100 m / sec. About 15% is only taken into the combustion chamber via the axial flow swirler, but in the above-described combined combustion system, 50% to 70% of the introduced compressed air is put into the combustion chamber via the fuel injection unit. Therefore, an axial flow swirler having a large outer diameter is used in the premixed gas supply unit outside the fuel injection unit. A circular opening having a diameter corresponding to the axial flow swirler having a large outer diameter is formed in the cowl covering the upstream side of the combustion cylinder, and the compressed air is guided to the fuel injection unit through this opening. . In the case of the conventional diffusion combustion system, the circular opening is set to have a diameter of about 1.2 to 2 times the opening flow path height of the diffuser. The diameter is set to be 3 to 4 times the road height.

  However, the high-temperature and high-pressure compressed air introduced from the annular diffuser to the upstream side of the cowl is annular with a width corresponding to the height of the opening of the diffuser, that is, the radial dimension of the air inlet to the combustion device. A high dynamic pressure region distributed in a band shape is formed, and this high dynamic pressure region faces only the central portion of each opening having a large diameter of the cowl. Therefore, the axial swirler of the premixed gas supply unit includes a portion into which high dynamic pressure compressed air that has passed through the central portion in the gas turbine radial direction at the opening of the cowl flows and a portion into which compressed air that does not have high dynamic pressure flows. And the distribution of the air flow rate passing through the axial swirler becomes non-uniform.

  In this way, when air passes through the axial swirler with a non-uniform flow distribution, uneven mixing of fuel and air occurs in the premixing passage of the premixed gas supply section, so the flame temperature in the combustion chamber is nonuniform. This causes an adverse effect such as an increase in the amount of NOx generated in the high-temperature flame part, a decrease in the life of the liner due to the generation of hot spots on the liner, and a decrease in the life of the turbine due to uneven distribution of the outlet temperature of the combustor. In addition, if the air velocity distribution derived from the axial swirler is uneven, a good backflow region cannot be formed in the combustion chamber, which hinders good ignitability and stable combustion, and reduces combustion efficiency. Invite.

  In order to solve the above-mentioned problem, it is conceivable to install the cowl far away from the diffuser or to increase the opening channel height of the diffuser, but when using any of these means, Since the combustor's axial length increases and the combustor becomes larger, it is difficult to adopt.

  The present invention is a combined combustion system that combines two combustion systems, ie, a diffusion combustion system and a lean premixed combustion system, in which the air flowing into the annular swirler has a uniform velocity distribution in the entire circumferential direction, An object of the present invention is to provide a combustion apparatus for a turbine engine that can make the flame temperature distribution in the combustion chamber uniform by mixing.

  In order to achieve the above object, a combustion apparatus for a gas turbine engine according to the present invention includes an annular combustion cylinder that forms a combustion chamber, and a premixed fuel and air that forms a premixed gas combustion region in the combustion chamber. A plurality of fuel injection units arranged in the circumferential direction of the combustion cylinder having a pre-mixture supply section for supplying a mixture, a fuel supply unit for supplying fuel to the pre-mixture supply section, and concentric with the combustion cylinder And an annular air inlet for introducing compressed air from a compressor upstream of the fuel injection unit, the fuel injection unit having a swirler for air swirling fixed to the air intake. In the annular high dynamic pressure region where the plurality of blades arranged in the circumferential direction of the swirler are opposed to the air introduction port, the circumferential width of the front edge portion is set larger than the outside of this region.

  According to this configuration, since the circumferential width of the front edge portion of the blade is set large in the region facing the high dynamic pressure region in the swirler, the air flow path formed between the blades As the directional width increases, the opening area of the inlet is reduced, and the amount of air taken into the air flow path is reduced. Therefore, an increase in the amount of air passing through the swirler is suppressed in a portion included in the high dynamic pressure region in the swirler. On the other hand, in the air flow path into which air with low dynamic pressure flows in the swirler, the opening width of the inlet is increased because the circumferential width of the front edge of the blade is set small, so that the air passage passes through the inside. The amount of air increases. As a result, the same amount of air is taken into each air flow path between the blades of the annular swirler. As a result, the air flows out from the outlet at a substantially uniform flow velocity over the entire circumferential direction of the swirler. To do.

  For this reason, in the premixed gas supply section, the mixture concentration of fuel and air is made uniform over the entire circumferential direction. As a result, the flame temperature in the combustion chamber becomes uniform, so the amount of NOx generated due to the high temperature flame is reduced. It is possible to suppress the occurrence of hot spots on the wall surface of the liner constituting the combustion chamber, and to ensure a sufficient lifetime of the liner. In addition, a sufficient life of the turbine can be ensured as the outlet temperature distribution of the combustor becomes uniform. Furthermore, since the velocity distribution of the air-fuel mixture supplied from the pre-air mixture supply unit is made uniform, a good backflow region can be stably formed in the combustion chamber, so that stable combustibility, good ignition performance and high combustion efficiency are achieved. Can be secured.

  In the present invention, the fuel injection unit further includes a fuel injection unit that sprays fuel so as to form a diffusion combustion region in the combustion chamber, and the fuel injection unit is a central part of the annular premixed gas supply unit The main part of the fuel supply unit is arranged upstream of the air intake, and is located at the center of the annular main fuel supply that is concentric with the air intake. The fuel injection unit may have a pilot fuel supply unit that supplies pilot fuel. As a result, the fuel injection unit and the premixed gas supply unit can be appropriately arranged to reduce the size of the fuel injection unit.

  In the present invention, the circumferential width of the leading edge of the blade is the largest in the radial center of the annular high dynamic pressure region, and the smallest in the remote region farthest from the high dynamic pressure region. It is preferable to change gradually in the intermediate region between. Thereby, the front edge part of each blade | wing arrange | positioned along the circumferential direction to the annular swirler becomes the circumferential width corresponding to the velocity distribution of the air flowing into the swirler, and the swirler flows out over the entire circumferential direction of the swirler. The flow rate of the air to be made can be made more uniform.

  In the present invention, it is preferable that the maximum value of the circumferential width of the leading edge of the blade is set to 1.2 to 8 times the minimum value. More preferably, it is in the range of 2 to 7 times. The high dynamic pressure region is generated in an annular belt-like arrangement along the circumferential direction of the gas turbine, and the air flowing into the swirler generally has a large flow velocity in the portion included in the high dynamic pressure region, and is the highest from the high dynamic pressure region. It becomes about 1.3 to 3 times the flow rate of the distant part. In order to make this flow velocity uniform at the outlet of the swirler, it is preferable to set the maximum value and the minimum value of the circumferential width of the front edge of the blade as described above. For example, the minimum width of the blade in the circumferential direction is about 1 mm due to processing restrictions, and the maximum value is about 6 mm due to weight reduction requirements.

  According to the combustion apparatus for a gas turbine engine of the present invention, since the circumferential width of the front edge portion of the blade in the high dynamic pressure region in the swirler is set large, the air that has passed through the annular swirler is As a result, the premixed gas in which the mixture concentration of fuel and air is made uniform over the entire circumferential direction of the annular premixed gas supply section is generated. The flame temperature is also made uniform, the amount of NOx generated can be suppressed, and a sufficient lifetime of the liner constituting the combustion chamber can be ensured. In addition, as the temperature distribution at the outlet of the combustor is made uniform, sufficient life of the turbine can be secured, and a good backflow region can be stably formed in the combustion chamber to achieve stable combustion, good ignition performance and high combustion efficiency. Can be secured.

  Hereinafter, preferred embodiments of the present invention will be described with reference to the drawings. FIG. 1 shows a head of a combustor 1 constituting a combustion apparatus for a gas turbine engine according to an embodiment of the present invention. The combustor 1 combusts an air-fuel mixture generated by mixing fuel with compressed air supplied from a compressor (not shown) of a gas turbine engine, and sends high-temperature and high-pressure combustion gas generated by the combustion to the turbine. To drive the turbine.

  The combustor 1 is an annular type, an annular inner casing 8 is concentrically disposed inside an annular outer casing 7, and has a cyclic internal space that is concentric with a gas turbine axis (rotation axis). A housing 6 is configured. In the annular inner space of the combustion device housing 6, a combustion cylinder 9 in which an annular inner liner 11 is concentrically disposed inside the annular outer liner 10 is disposed concentrically with the combustion device housing 6. The combustion cylinder 9 has an annular combustion chamber 12 formed therein, and a plurality (14 in this embodiment) of fuel injection units 2 that inject fuel into the combustion chamber 12 on the top wall 9a of the combustion cylinder 9. Are arranged at equal intervals on a single circle concentric with the combustion cylinder 9. Each fuel injection unit 2 includes a fuel injection unit (pilot fuel injection nozzle) 3 and a premixed gas supply unit (main fuel injection unit) concentrically provided with the fuel injection unit 3 so as to surround the outer periphery of the fuel injection unit 3. Nozzle) 4. Details of the fuel injection unit 3 and the premixed gas supply unit 4 will be described later.

  Two spark plugs 13 for penetrating through the outer casing 7 and the outer liner 10 are provided so as to face the radial direction of the combustion cylinder 9 and have their tips opposed to the fuel injection unit 2. Therefore, in the combustor 1, the combustible air-fuel mixture from the two fuel injection units 2 facing the two spark plugs 13 is first ignited, and the flame caused by this combustion is combusted from the adjacent fuel injection units 2. Then, it propagates while moving one after another, and the combustible air-fuel mixture from all the fuel injection units 2 is ignited.

  FIG. 2 is an enlarged longitudinal sectional view taken along line II-II in FIG. Compressed air CA fed from the compressor is introduced into the annular internal space of the combustion device housing 6 through the internal passage of the annular diffuser 14 from the air inlet 14a at the downstream end thereof. The compressed air CA is supplied to the fuel injection unit 2 through an opening 27 a provided in the cowl 27 that covers the upstream end of the combustion cylinder 9, and is guided to the outside of the combustion cylinder 9 by the cowl 27. Nine outer liners 10 and inner liners 11 are supplied into the combustion chamber 12 through a plurality of dilution air inlets 17.

  The fuel piping unit 18 includes a first fuel supply system F1 that supplies fuel for diffusion combustion to the fuel injection unit 3 and a second fuel that supplies fuel for lean premixed combustion to the premixed gas supply unit 4. A supply system F <b> 2 is provided and is supported by the outer casing 7. A fuel supply unit 20 connected to the fuel pipe unit 18 supplies the fuel F from the fuel pipe unit 18 to the fuel injection unit 2. The fuel supply unit 20 is disposed so as to penetrate the opening 27a of the cowl 27, and details thereof will be described later. The fuel injection unit 2 is supported by the outer liner 10 via a flange 5A provided on the outer periphery thereof and a support 5B provided on the outer liner 10, and the outer liner 10 is supported on the outer casing 7 by a liner fixing pin P. Yes. A turbine first stage nozzle TN is connected to the downstream end of the combustion cylinder 9.

  The fuel injection unit 3 is provided at the center of the fuel injection unit 2. The fuel injection unit 3 includes an inner and outer double annular air swirler 33 and a diffusion nozzle 32. The fuel injection unit 3 is fed from the first fuel supply system F 1 of the fuel pipe unit 18 and is supplied to the fuel supply unit 20. The fuel F injected from the fuel injection hole of the pilot fuel supply unit 31 is atomized by the compressed air CA that has passed through the swirler 33, and then sprayed into the combustion chamber 12 through the diffusion nozzle 32, whereby the diffusion combustion region 50 Form. The annular premixed gas supply unit 4 is provided so as to surround the outer periphery of the fuel injection unit 3. The premixed gas supply unit 4 has a premixing passage 42 and an inner and outer double annular air swirler 43, and is supplied from the second fuel supply system F <b> 2 of the fuel pipe unit 18 to be supplied to the fuel supply unit 20. The fuel F injected into the premixing passage 42 from the fuel injection holes arranged at equal intervals in the circumferential direction in the annular main fuel supply unit 41 is mixed with the compressed air CA that has passed through the swirler 43 to prepare the premixed gas. This is generated and injected into the combustion chamber 12 to form a premixed combustion region 51.

  The fuel injection unit 3 and the premixed gas supply unit 4 take in the compressed air CA through the inner and outer double swirlers 33 and 43 fixed to the respective annular air intakes. The fuel supply unit 20 has a configuration in which a pilot fuel supply unit 31 and a main fuel supply unit 41 are connected to a unit base body 19. The fuel supply unit 20 has a predetermined number of fuel injection units 3 with the pilot fuel supply unit 31 fitted into the annular swirler 33 and the main fuel supply unit 41 fitted into the annular swirler 43. They are fastened to each other by fitting pins (not shown).

  The fuel F is supplied to the fuel injection unit 3 from the first fuel supply system F1 in the entire load region. The premixed gas supply unit 4 includes a second fuel at a high load of, for example, 70% or more with respect to the total load, and at a high load and a low load, for example, at a medium load of 40 to 70% of the full load. Fuel F is supplied from the supply system F2. The premixed gas supply unit 4 supplies only the compressed air CA to the combustion chamber 12 through the swirler 43 because the fuel F is not supplied at a low load of 40% or less with respect to the total load.

  3 is a perspective view showing a relative relationship among the cowl 27, the fuel supply unit 20, and the fuel injection unit 2 constituting the combustor 1. As shown in FIG. In the fuel supply unit 20, an annular main fuel supply unit 41 concentric with the inner and outer double swirler 43 of the premixed gas supply unit 4 is fitted into the swirler 43 from the upstream side (front side in the figure), And the pilot fuel supply part 31 located in the center part of the main fuel supply part 41 is inserted into the inside and outside double swirler 33 of the fuel injection part 3 in FIG. It is connected to. The inlet 43 a of the swirler 43 forms an air intake of the premixed gas supply unit 4, and the inlet 33 a of the swirler 33 forms an air intake of the fuel injection unit 3. The main part of the fuel supply unit 20 excluding the main fuel supply part 41 is located upstream of these air intakes.

  Further, the cowl 27 is disposed on the upstream side with respect to the swirlers 33 and 43 of the fuel injection unit 3 and the premixed gas supply unit 4 in the fuel supply unit 2. An opening 27 a for introducing the compressed air CA into the fuel injection unit 2 is formed in the cowl 27 in a circular shape when viewed from the axial direction of the unit 2.

  4 is a view in the direction of arrow A in FIG. 3 (the axial direction of the fuel injection unit 2 in FIG. 2), and the opening 27a of the cowl 27 has an outer diameter of the inner and outer double swirler 43 of the premixed gas supply unit 4. It is a larger diameter circle. The air introduction hole 14a that is the outlet of the diffuser 14 in FIG. 2 has an annular shape that is concentric with the gas turbine shaft center, and the compressed air CA having a high flow velocity flows out from the air introduction hole 14a. Therefore, as shown in FIG. A high dynamic pressure area HA having a width W corresponding to the opening flow path height (diameter dimension) of the introduction port 14a and annularly distributed in the gas turbine circumferential direction Q is generated in a belt shape. It passes through the central portion of the gas turbine radial direction R in 27a, that is, the central portion of the gas turbine radial direction R in the inner / outer double swirler 43 of the premixed gas supply unit 4. For example, a high dynamic pressure region HA having a width W of about 0.4D passes through a central portion of the annular swirler 43 having an outer diameter D.

  Therefore, in the swirler 43 of the premixed gas supply unit 4, the compressed air CA having a high dynamic pressure flows into a portion entering the high dynamic pressure region HA, and the compression of the low dynamic pressure is performed in a portion other than the high dynamic pressure region HA. Air CA flows in. That is, the compressed air CA flows into the swirler 43 with a non-uniform flow velocity distribution. In this combustion apparatus, the compressed air CA is designed to prevent the occurrence of problems due to the local flow of the flow velocity with a nonuniform distribution. This point will be described below.

  5A is an enlarged view showing only the inner and outer double swirlers 43 of the premixed gas supply unit 4 extracted from FIG. 4, and FIG. 5B is a gas turbine diameter of the compressed air CA flowing into the swirler 43. It is a distribution map of the flow velocity in the direction R, and is shown in association with the position of the swirler 43 in (a). As shown in (b), when the flow velocity of the compressed air CA flowing in the high dynamic pressure region HA in the swirler 43 is V, both ends of the swirler 43 in the gas turbine radial direction R, that is, farthest from the high dynamic pressure region HA. The flow rate of the compressed air CA flowing into the portion is about 0.5V.

  In this manner, the compressed air CA flows into the swirler 43 with a distribution in which the flow velocity is uneven in location. On the other hand, as shown in FIG. 5A, the swirler 43 has a circumferential width of the front edge portions (edge portions on the near side in the drawing) of the plurality of blades 48 arranged at equal intervals in the circumferential direction thereof. Is set to a different value corresponding to the flow velocity distribution of the compressed air CA flowing in. That is, the blade 48A located in the central region of the annular high dynamic pressure region HA, that is, the central region in the gas turbine radial direction R of the swirler 43 is set to the maximum circumferential width VWH, and the high dynamic pressure region HA The blade 48B located in the farthest remote area LA, that is, the area on both ends in the gas turbine radial direction R of the swirler 43 is set to the minimum circumferential width VWL, and the intermediate area MA between these two areas HA, LA The circumferential width of the blade 48C positioned at is set to a value VWM that gradually decreases from the maximum value VWH to the minimum value VWL.

  In the flow velocity distribution of the compressed air CA flowing into the swirler 43, the ratio of the maximum flow velocity V to the minimum flow velocity is usually 1.3 to 3 times. In the flow velocity distribution shown in FIG. 5B, this ratio is twice, and the circumferential width VWH of the blade 48A at the center of the high dynamic pressure region HA corresponds to the blade 48B in the remote region LA. Is set to be three times the circumferential width VWL. The circumferential width VWM of the blade 48C in the intermediate area MA is set to an intermediate value between VWL and VWH. Thereby, each circumferential width VW of each blade 48 of the swirler 43 becomes a value corresponding to the flow velocity distribution of the inflowing compressed air CA.

  The dimension of the opening width IW of the air flow path 49 is determined by the relationship between the circumferential length of the swirler 43 and the number of blades, but when the minimum circumferential width VWL is around 1 mm due to processing restrictions, the opening width IW is 8 to 8 mm. Generally, it is 12 mm. That is, IW = 8-12VWL. Considering this, in order to make the flow velocity distribution at the outlet of the swirler 43 uniform, the circumferential width VWH of the blade 48A at the center of the high dynamic pressure region HA is set to be equal to the circumferential width VWL of the blade 48B in the remote region LA. It is preferable to set to 1.2 to 8 times. More preferably, it is 2 to 7 times. For example, as described above, the minimum width of the blade in the circumferential direction is about 1 mm due to processing restrictions, and the maximum value is about 6 mm due to weight reduction requirements. In addition, the front edge part of each blade | wing 48 is shaped so that it may have roundness in order to suppress air resistance.

  6A is an enlarged cross-sectional view showing the main fuel supply unit 41 and the inner swirler 43 of the inner and outer doubles in the premixed gas supply unit 4, and FIG. 6B is a diagram showing a high movement in the swirler 43 of FIG. It is an expanded sectional view of a portion which enters pressure field HA. As shown in FIG. 6A, the fuel F supplied through the passage of the main fuel supply section 41 is provided on the fuel injection hole 44 provided on the passage wall and the inner peripheral wall of the premixing passage 42. The fuel is introduced into the downstream of the swirler 43 through the fuel introduction hole 47 and atomized by the swirling airflow of the compressed air CA from the swirler 43, and the fuel F and the compressed air CA are sufficiently mixed in the premixing passage 42 of FIG. The premixed gas is generated, and this premixed gas is supplied into the combustion chamber 12. In this embodiment, a plurality of fuel supply holes 52 including the fuel injection holes 44 and the fuel introduction holes 47 that are opened in the radial direction are provided at equal intervals in the circumferential direction. The amount of fuel F that passes through is set to be the same. The number of the fuel supply holes 52 is arbitrary, but in this embodiment, it is set to ½ of the number of blades 48.

  The circumferential width of the front edge portion of each blade 48 in FIG. 5A is set to a larger value as the flow velocity of the compressed air CA flowing in increases, but as shown in FIG. 6B. In addition, the circumferential width of the trailing edge (right edge in the drawing) of each blade 48 is set to a small constant value. Therefore, the opening width IW of the inflow port 49a of the air flow path 49 formed between the blades 48 is the smallest in the portion entering the high dynamic pressure region HA, and gradually increases toward the region away from the high dynamic pressure region HA. Become bigger. On the other hand, the opening widths OW of the outlets 49b of the air channels 49 are all the same value.

Therefore, in the air flow path 49 into which the compressed air CA having the highest flow velocity, that is, the compressed air CA having the highest dynamic pressure flows, the inlet 49a having the smallest opening area is formed by the blade 48A having the largest circumferential width VWH. Therefore, the inflow air amount is suppressed. Here, when the opening width of the outlet 49b of the air flow path 49 is OW, the blade width VW is considered in order to make the flow in the air flow path 49, that is, between the blades, an accelerated flow with less risk of separation. The inlet width (IW-VW) of the substantial air flow path 49 is
(IW-VW) = OW
Is desirable. Here, the opening width OW of the outlet 49 b is OW = cos α × IW, where α is the angle of the chord line 55 of the blade 48 with respect to 43 C with respect to the axis of the swirler 43.

  Therefore, the inlet width (IW−VW) of the air flow path 49 is such that the portion where the inlet 49a having the smallest opening area is formed (high dynamic pressure region) has a smaller flow acceleration than the other portions. 6 and the opening width OW of the outflow port 49b, the compressed air CA flowing out from the outflow port 49b of each air flow path 49 is schematically shown in FIG. As shown in FIG. 5, all of them have a substantially uniform flow rate. Further, since the opening widths OW of the outlets 49b of the air channels 49 are all set to be the same as described above, the flow rates of the compressed air CA flowing out from the outlets 49b of the air channels 49 are almost the same. It becomes. On the other hand, as described above, the supply amounts of the fuel F passing through the fuel injection holes 44 and the fuel introduction holes 47 are all set to be substantially the same.

  Therefore, in the premixing passage 42, the compressed air CA and the fuel F are mixed at a uniform mixing ratio over the entire circumferential direction of the premixed gas supply unit 4. Therefore, the premixed gas generated in the premixing passage 42 is supplied as a premixed gas supply, as schematically shown in FIG. The mixture concentration is constant throughout the circumferential direction of the section 4, and this premixed gas is supplied into the annular combustion chamber 12 of FIG. As a result, the flame temperature in the combustion chamber 12 becomes uniform, so that the amount of NOx generated due to the high-temperature flame can be suppressed, and the walls of the liners 10 and 11 constituting the combustion chamber 12 have a temperature distribution. Since the non-uniform hot spots do not occur, the liners 10 and 11 can have a sufficient life without being damaged. In addition, since the outlet temperature distribution of the combustor 2 is also made uniform, the first nozzle TN of the turbine can ensure a sufficient life by suppressing the occurrence of burning and high thermal stress. Furthermore, since the velocity distribution of the air-fuel mixture supplied from the pre-air mixture supply unit 4 is made uniform, a good backflow region can be stably formed in the combustion chamber 12, so that stable combustibility, good ignitability, and high Combustion efficiency can be ensured.

  In this embodiment, as shown in FIG. 6B, a straight blade 48 having a wide leading edge is illustrated and described. However, the blade 48 may have an airfoil cross-sectional shape as described above. The same effect as can be obtained.

It is a schematic front view which shows the combustion apparatus of the gas turbine engine which concerns on one Embodiment of this invention. It is an expanded sectional view along the II-II line of FIG. It is a perspective view which shows the principal part structure of a combustion apparatus same as the above. FIG. 4 is a view in the direction of arrow A in FIG. 3. (A) is the enlarged view which showed only the swirler of FIG. 4, (b) The distribution map of the flow velocity of the compressed air which flows into a swirler. (A) is an expanded sectional view which shows the main fuel supply part and inner swirler of a premixed gas supply part, (b) is an expanded sectional view of the part which enters into the high dynamic pressure area | region in the swirler of (a).

Explanation of symbols

2 Fuel injection unit 3 Fuel injection unit 4 Premixed gas supply unit 9 Combustion cylinder 12 Combustion chamber 14a Air inlet 20 Fuel supply unit 31 Pilot fuel supply unit 41 Main fuel supply unit 43 Swirler 48 (48A, 48B, 48C) Blade 50 Diffusion combustion region 51 Premixed gas combustion region F Fuel CA Compressed air HA High dynamic pressure region VWH, VWL, VWM Circumferential width LA Remote region MA Intermediate region

Claims (4)

An annular combustion cylinder forming a combustion chamber;
A plurality of fuel injection units arranged in a circumferential direction of the combustion cylinder, having a premixed gas supply section for supplying a premixed gas of fuel and air so as to form a premixed gas combustion region in the combustion chamber;
A fuel supply unit for supplying fuel to the premixed gas supply unit;
Concentric with the combustion cylinder, and provided with an annular air inlet for introducing compressed air from a compressor upstream of the fuel injection unit,
The fuel injection unit has an air swirling swirler fixed to the air intake, and a plurality of blades arranged in the circumferential direction of the swirler are in an annular high dynamic pressure region facing the air inlet. A combustion apparatus for a gas turbine engine in which the circumferential width of the front edge portion is set to be larger than outside this region.   2. The fuel injection unit according to claim 1, further comprising a fuel injection unit that sprays fuel so as to form a diffusion combustion region in the combustion chamber, the fuel injection unit being the center of the annular premixed gas supply unit The main part of the fuel supply unit is located upstream of the air intake, and is located at the center of the annular main fuel supply part concentric with the air intake. A combustion apparatus for a gas turbine engine, comprising: a pilot fuel supply unit that supplies pilot fuel to the fuel injection unit.   3. The circumferential width of the leading edge of the blade according to claim 1, wherein the circumferential width is maximum at a radial center of the annular high dynamic pressure region and minimum at a remote region farthest from the high dynamic pressure region. A combustion apparatus for a gas turbine engine that is gradually changing in an intermediate region between these regions.   4. The combustion apparatus for a gas turbine engine according to claim 1, wherein the maximum value of the circumferential width of the leading edge of the blade is 1.2 to 8 times the minimum value.

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