BURNER FOR A GAS TURBINE

Abstract

A burner for a gas turbine engine has a combustion chamber and a swirler adapted to guide a swirler air flow to the combustion chamber, wherein the swirler has a first wall confining the swirler air flow as well as a second wall confining the swirler air flow on the same side as and downstream with respect to the swirler air flow from the first wall and being displaced with respect to the first wall in a direction away from the swirler air flow so that a step being able to cause a flow separation of the swirler air flow is formed by the first wall and the second wall, wherein the second wall has a through hole in its surface adapted to inject a liquid fuel into the swirler air flow.

Claims

1. A burner for a gas turbine engine, wherein the burner comprises: a combustion chamber, and a swirler adapted to guide a swirler air flow to the combustion chamber, wherein the swirler comprises a first wall confining the swirler air flow as well as a second wall confining the swirler air flow on the same side as and downstream with respect to the swirler air flow from the first wall and being displaced with respect to the first wall in a direction away from the swirler air flow so that a step being able to cause a flow separation of the swirler air flow is formed by the first wall and the second wall, wherein the second wall has a through hole in its surface adapted to inject a liquid fuel into the swirler air flow.

2. The burner according to claim 1, wherein the swirler comprises at least one further wall confining the swirler air flow on the same side as and downstream with respect to the swirler air flow from the second wall, wherein each of the further walls is displaced with respect to its directly adjacent and with respect to the swirler air flow upstream wall in a direction away from the swirler air flow so that a respective step being able to cause a flow separation of the swirler air flow is formed by two directly adjacent walls, wherein each further wall has a through hole in its surface adapted to inject the liquid fuel into the swirler air flow.

3. The burner according to claim 2, wherein the distance between two neighboured steps is at least 2*L, wherein L is the distance from the step to its with respect to the swirler air flow downstream and closest through hole.

4. The burner according to claim 1, wherein the swirler comprises a multitude of swirler sectors confining the swirler air flow and shaped to cause an angular momentum of the swirler air flow, wherein the swirler sectors are in contact with each of the walls.

5. The burner according to claim 4, wherein the step is located at a radial distance from the burner axis which is from r.sub.1+0.2*(r.sub.2r.sub.1) to r.sub.1+0.8*(r.sub.2r.sub.1), wherein r.sub.2-r.sub.1 is the distance from the radial inner end of the swirler sectors to the radial outer end of the swirler sectors.

6. The burner according to claim 1, wherein the height of each step is from 0.2*L to 0.5*L, wherein L is the distance from the step to its with respect to the swirler air flow downstream and closest through hole.

7. The burner according to claim 6, wherein L is from 4 mm to 20 mm, in particular from 4 mm to 8 mm.

8. Burner The burner according to claim 1, wherein the height of each step is at least 1 mm.

9. The burner according to claim 8, wherein the height of each step is maximum 15% of the swirler channel height (H), wherein the swirler channel height (H) is the distance from the with respect to the swirler air flow upstream wall forming the step to an opposite wall confining the swirler air flow and facing towards the with respect to the swirler air flow upstream wall forming the step.

10. The burner according to claim 1, wherein the diameter of the through hole is from 0.5 mm to 3 mm.

11. The burner according to claim 1, wherein the swirler is adapted to guide the swirler air flow such that the air flow entering the combustion chamber has a flow direction with respect to a main flow direction within the combustion chamber, wherein the flow direction essentially consists of a radial inward component and a component in circumferential direction.

12. The burner according to claim 1, wherein the burner is configured for dry operation only.

13. The burner according to claim 1, wherein the burner is adapted to generate a premixed flame.

14. The burner according to claim 6, wherein L is from 4 mm to 8 mm.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0012] The above mentioned attributes and other features and advantages of this invention and the manner of attaining them will become more apparent and the invention itself will be better understood by reference to the following description of embodiments of the invention taken in conjunction with the accompanying drawings, wherein

[0013] FIG. 1 shows part of a gas turbine in a sectional view and in which the present inventive burner is incorporated,

[0014] FIG. 2 shows a longitudinal section of the burner and a part of the combustion chamber,

[0015] FIG. 3 shows a perspective view of a part of the a swirler of the burner,

[0016] FIG. 4 shows a sectional view of a part of the swirler,

[0017] FIG. 5 shows a top view of the burner,

[0018] FIGS. 6 to 10 show different embodiments for through holes of the swirler.

DETAILED DESCRIPTION OF INVENTION

[0019] FIG. 1 shows an example of a gas turbine engine 10 in a sectional view. The gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20. The gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10. The shaft 22 drivingly connects the turbine section 18 to the compressor section 14.

[0020] In operation of the gas turbine engine 10, air 24, which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16. The burner section 16 comprises a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28. The combustion chambers 28 and the burners 30 are located inside the burner plenum 26. The compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.

[0021] This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine 18.

[0022] The turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22. In the present example, two discs 36 each carry an annular array of turbine blades 38. However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs. In addition, guiding vanes 40, which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38.

[0023] The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22. The guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.

[0024] The turbine section 18 drives the compressor section 14. The compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48. The rotor blade stages 48 comprise a rotor disc supporting an annular array of blades. The compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48. The guide vane stages include an annular array of radially extending vanes that are mounted to the casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.

[0025] The casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14. A radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48.

[0026] The present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.

[0027] FIG. 2 shows that the burner 30 comprises an inner wall 101 that confines the combustion chamber 28 in a radial direction. Furthermore, the burner 30 comprises a pilot burner 104 and a main burner 105 that are arranged on an axial end of the burner 30. The main burner 105 is arranged radially outside from the pilot burner 104. The burner 30 comprises an outer wall 102 that is arranged radially outside of the inner wall 101. The inner wall 101 and the outer wall 102 are essentially rotationally symmetric around a burner axis 35 of the burner 30. The air 24 is streamed in the space between the inner wall 101 and the outer wall 102 towards the pilot burner 104 and the main burner 105 as indicated by arrows 108, so that the inner wall 101 is cooled and the air 24 is preheated before it enters the combustion chamber 28.

[0028] The burner 30 comprises a swirler 107 located on the main burner 105 for swirling the air before it enters the combustion chamber 28. After passing the space between the inner wall 101 and the outer wall 102 the air 24 passes through the swirler 107 in a direction towards the burner axis 35 and enters the combustion chamber 28. The burner 30 is configured for dry operation only, i.e. it is not configured for the injection of water into the combustion chamber 28.

[0029] The swirler 107 comprises a first axial end 113 that coincides with the main burner 105 and a second axial end 114 being located opposite to the first axial end 113. As it can be seen in FIGS. 3 and 5, the swirler 107 furthermore comprises a multitude of swirler sectors 118 that are in contact with the first axial end 113 and the second axial end 114. The first axial end 113, the second axial end 114 and the swirler sectors 118 confine a swirler air flow 125. The swirler sectors 118 are shaped such that the air flow entering the combustion chamber 28 has a flow direction with respect to the burner axis 35, wherein the flow direction essentially consists of a radial inward component and a component in circumferential direction.

[0030] FIGS. 2 and 3 show that the swirler 107 comprises a first wall 115 that confines the swirler air flow 125 at the first axial end 113 as well as a second wall 116 that confines the swirler air flow 125 on the same side as, i.e. also at the first axial end 113, and downstream with respect to the swirler air flow 125 from the first wall 115. The second wall 116 is displaced with respect to the first wall 115 in an axial direction with respect to the burner axis 35 away from the swirler air flow 125 so that a step 117 being able to cause a flow separation of the swirler air flow 125 is formed by the first wall 115 and the second wall 116. The main burner 105 comprises a through hole 103 that extends through the second wall 116. Via the through hole 103 a liquid fuel can be injected into the swirler air flow 125. The burner 30 does not comprise fuel lances so that the liquid fuel is in contact with the through hole 103 and is immediately in contact with the swirler air flow 125 when leaving the through hole 103.

[0031] As it can be seen in FIGS. 3 and 4, an atomisation region 119 is formed within the swirler air flow 125 directly beginning from where the liquid fuel enters the swirler air flow 125. A large part of the atomisation region 125 overlaps with the vortex caused by the flow separation of the swirler air flow 125 on the step 117 which results in a particular efficient atomisation of the liquid fuel and mixing of the liquid fuel with air.

[0032] The step 117 is located at a radial distance from the burner axis 35 which is from r.sub.1+0.2*(r.sub.2r.sub.1) to r.sub.1+0.8*(r.sub.2r.sub.1), wherein r.sub.1 is the radial distance from the burner axis to the radial inner end of the swirler sectors 118 and r.sub.2 is the radial distance from the burner axis to the radial outer end of the swirler sectors 118. The height h of each step 117 is from 0.2*L to 0.5*L, wherein L is the distance from the step 117 to its with respect to the swirler air flow 125 downstream and closest through hole 103. The height h of each step 117 is maximum 15% of the swirler channel height H. The swirler channel height H is the distance from the with respect to the swirler air flow 125 upstream wall 115 forming the step 117 to an opposite wall confining the swirler air flow 125 and facing towards the with respect to the swirler air flow 125 upstream wall 115 forming the step 117.

[0033] After the premixing of the liquid fuel with the air, the mixture enters the combustion chamber 28, where the combustion of the mixture occurs. The flame in the combustion chamber 28 has an inner recirculation zone 110 that stabilises the flame by transporting hot combustion products to the unburned air/fuel mixture, and an outer recirculation zone 111.

[0034] As can be seen in the FIGS. 3 and 4 fuel is injected perpendicularly with respect to the direction of the air flow 125. This fuel injection angle can be also described as parallel to the burner axis 35. The through hole 103 and/or at least its outlet end or nozzle is also arranged generally parallel to the burner axis 35 and therefore perpendicular to the direction of air flow 35. In this arrangement the fuel injected is mixed particularly well by the vortices created by the airflow over the step 117. The stated dimensions H, h and L are particularly well suited to this arrangement of a perpendicular fuel injection relative to the airflow direction. The terms perpendicular and parallel are intended to be approximate and directions angled up to 30 away from nominally parallel or perpendicular are intended to be within the scope of these terms as used here. The direction of the fuel injected is a nominal angle of the fuel spray centre-line rather than a cone angle of the fuel spray.

[0035] The vortices created by the step 117 are particularly suited to providing good mixing of the fuel and air under low or part power conditions where there would otherwise be less mixing than desirable to minimise emissions.

[0036] As it can be seen in FIG. 5 the swirler 107 and the step 117 have the form of an ellipse, wherein other forms, e.g. a circle, are also conceivable. At least one through hole 103 is located between two adjacent swirler sectors.

[0037] It is conceivable that the swirler 107 comprises at least one further wall confining the swirler air flow 125 on the first axial end 113 and downstream with respect to the swirler air flow 125 from the second wall 116, wherein each of the further walls is displaced in an axial direction with respect to its directly adjacent and with respect to the swirler air flow 125 upstream wall in a direction away from the swirler air flow 125 so that a respective step being able to cause a flow separation of the swirler air flow 125 is formed by two directly adjacent walls, wherein each further wall has a through hole 103 in its surface adapted to inject the liquid fuel into the swirler air flow 125. The distance between two neighboured steps is at least 2*L. It is conceivable that the steps are arranged parallel to each other.

[0038] FIGS. 6 to 10 show possible geometries for the through holes 103. The first through hole 121 according to FIG. 6 has the shape of a circle with a missing sector having an angle of 90. The second through hole 122 according to FIG. 7 has the shape of a ring. The through hole 123 according to FIG. 8 consists of a plurality of elongate through holes that are arranged tilted with respect to each other. The through hole 124 according to FIG. 9 has the form of a circle. FIG. 10 shows a perspective view of a plate 126 containing the through hole 124 according to FIG. 9. The through holes 103 can be formed as an assembly of several joint layers of metal.

[0039] Although the invention is described in detail by the preferred embodiment, the invention is not constrained by the disclosed examples and other variations can be derived by the person skilled in the art, without leaving the extent of the protection of the invention.