EP3564585A1

EP3564585A1 - Swirler arrangement of a burner

Info

Publication number: EP3564585A1
Application number: EP18170798.5A
Authority: EP
Inventors: Mats Andersson; Nicklas Johansson; Jenny Larfeldt; Daniel Moell; Magnus Persson
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2018-05-04
Filing date: 2018-05-04
Publication date: 2019-11-06

Abstract

The present invention relates to a swirler arrangement (10; 10a) of a burner (52), in particular of a gas turbine engine (14), comprising at least two wings (16, 18; 16a, 18a), wherein a thread-like flow channel (20) is arranged between the at least two wings (16, 18; 16a, 18a), wherein the flow channel (20) is allowing a flow medium (22) to travel along the flow channel (20) in a flow direction (24, 24') and wherein the flow medium (22) is set into a thread-like motion by the flow channel (20) when flowing through the flow channel (20) and wherein each wing (16, 18; 16a, 18a) of the at least two wings (16, 18; 16a, 18a) comprise an upstream part (26) and a downstream part (28; 28a).

To increase the flashback resistance of the burner (52) it is proposed that the downstream part (28; 28a) comprises at least one opening (30; 30a) for injecting fuel (32) into the flow medium (22) while traveling the flow channel (20).

Description

Field of the Invention

The present invention relates to a swirler arrangement of a burner, in particular of a gas turbine engine. The present invention further relates to a burner with such a swirler arrangement as well as to a use of such a swirler arrangement.

Background to the Invention

In a gas turbine engine a working fluid, like air, being compressed by a compressor enters a burner section where combustion of a mixture of the working fluid and a fuel occurs. The resulting combustion gas drives a turbine through the expansion and deflection of the gas through the turbine of the gas turbine engine. The turbine work or a part thereof, is transferred to the compressor through an interconnecting shaft. The compressor discharge temperature in modern gas turbine engines may be well above 420°C. In case the gas turbine engine is operated with highly reactive fuels, like hydrogen, auto-ignition of a hydrogen/air mixture may occur with these temperatures. Since the burner is situated after the compressor it may have surface temperatures similar to the temperature of the compressor discharge. Thus, a risk for flashbacks may be disadvantageously enhanced.
In so-called DLE-combustion systems (DLE: dry low emission) fuel and air are mixed before entering a combustion chamber. It is, for example, known to introduce the fuel by a jet in cross flow arrangement where the fuel is injected perpendicular to the airstream passing through the burner. In the vicinity of these fuel jets there are always low speed zones where the fuel concentration is sufficiently high to be ignited by the hot burner geometry. Consequently, it is currently avoided to use highly reactive fuels in DLE-combustion systems.
It is a first objective of the present invention to provide a swirler arrangement with which the above-mentioned shortcomings can be mitigated, and especially, a secure and reliable operation of the burner can be ensured.
It is a further objective of the present invention to provide a burner, which can be operated safely, especially, with high reactive fuels. A still further objective of the present invention is to provide (a) use(s) of the swirler arrangement and/or the burner that minimises low speed zones and increases a flashback resistance of the gas turbine engine.
These objectives may be solved by a swirler arrangement, a burner and a use according to the subject-matter of the independent claims.

Summary of the Invention

Accordingly, the present invention provides a swirler arrangement of a burner, in particular of a gas turbine engine, comprising at least two wings, wherein a thread-like flow channel is arranged between the at least two wings, wherein the flow channel is allowing a flow medium to travel along the flow channel in a flow direction and wherein the flow medium is set into a thread-like motion by the flow channel when flowing through the flow channel and wherein each wing of the at least two wings comprise an upstream part and a downstream part.
It is proposed that the downstream part comprises at least one opening for injecting fuel into the flow medium while traveling the flow channel.
Due to the inventive matter the risk for the exhausted or injected fuel to come into contact with the hot burner geometry can be minimized. Hence, the flashback resistance of the burner can be increased. Consequently, a secure operation of the gas turbine engine even when operated with highly reactive fuels can be ensured.
Even if a chosen term is used in the singular or in a specific numeral form in the claims and the specification the scope of the patent (application) should not be restricted to the singular or the specific numeral form. It should also lie in the scope of the invention to have more than one or a plurality of the specific structure(s).
A turbine engine may be any engine feasibly for a person skilled in the art and is preferably a gas turbine engine. A gas turbine engine is preferably a whole plant, with at least a compressor section, a combustion section and a turbine section, further sections or devices may be provided. The combustion section may comprise the one or several burners, for example, one or several main burner(s) and/or one or several pilot burner(s). Further, a swirler arrangement is intended to mean an arrangement of parts being able or arranged in such a way that the flow medium is set into a swirling motion. Further, it comprises at least one mixing device, like a swirler, premixer or mixer. The swirler arrangement may be also called swirl generator.
In this context a "thread-like flow channel" is intended to mean a channel having or providing, respectively, a path or shape of a thread or a convolution. Furthermore, the phrase "a thread-like flow channel is arranged between the at least two wings" should be understood in that at least parts or sections of the wings form or built the shape of at least one part of the flow channel. Hence, the wings or parts thereof restrict the flow path of the flow medium. Moreover, at least the section of the flow channel between the two wings has a thread-like shape. Thus, the flow medium is influencable by the shape and/or orientation of the at least two wings or parts thereof into a thread-like motion. There may be other sections of the flow channel for the flow medium that may have a non-thread-like path, like a linear or circular path.
Moreover, a flow direction is intended to mean the direction from a section of the turbine engine being located upstream of the burner of a combustion section, like the compressor section, to a section of the turbine engine being located downstream of the burner of the combustion section, like the turbine section. Hence, the flow direction is a general direction "along" the turbine engine. In selected parts of the flow direction the actual direction may vary from the general direction. A flow medium may be any medium feasible for a person skilled in the art, like a fluid, a gas or a gas mixture, and is preferably air.
In this context the "upstream part" should be understood as the first region of the wing coming into contact with a particular fraction of the flow medium. Consequently, the "downstream part" is located downstream of the upstream part and is coming later in contact with said particular fraction of the flow medium. Moreover, the upstream part is the part of the wing connecting the wing with its support structure. The downstream part is extending from the upstream part into free space. Furthermore, the opening may also be called exhaust port or fuel injection point.
An aerodynamically suitable swirler arrangement can be provided when each wing of the at least two wings has an aerofoil-like shape with a trailing edge. Moreover, it may comprise a leading edge. Preferably, the upstream part comprises the leading edge and the downstream part comprises the trailing edge. Beneficially, the at least one opening is arranged at the trailing edge. Hence, an injection of the fuel can be advantageously realised in direction of the flow of the flow medium traveling along the flow channel. Due to this, the risk of the injected fuel contacting hot surfaces of the burner geometry, like areas of (a) adjacent positioned wing(s), can be avoided. In this case the hot geometry are the parts of the swirl generator exposed to the surrounding flow medium (air) flow, which may have a temperature close to 450°C. The fuel can be injected beneficially without coming in contact with parts of the swirl generator exposed to the surrounding hot airflow.
Moreover, it is provided that the swirler arrangement comprises a plurality of openings. Thus, fuel injection can be done efficiently. Furthermore, it may be advantageous when the openings of the plurality of openings are arranged over a flow width of the flow channel. Hence, customised injection may be performed depending on the position of the respective opening in relation to the flow channel. The openings of the plurality of openings can be distributed in any fashion feasible for a person skilled in the art, like randomly, in doublets, triplets or groups. Preferably, the openings of the plurality of openings are evenly distributed along the flow width of the flow channel or an outlet edge of the wing. As a result, a homogeneous injection over the width of the flow channel can be performed.
The opening may have any shape, like circular, oval, egg-shaped, rectangular, triangular etc., feasible for a person skilled in the art. Preferably, the opening has a circular shape. Hence, injection can be performed homogeneously. Further, such circular openings are easy to manufacture.
The at least one opening is surrounded by a rim surface. It is furthermore provided that the rim surface is embodied in such that the exhalable fuel adopts a flow component oriented in flow direction of the flow medium along the flow channel. Hence, properties of fuel injection can be selected or influenced according to specific needs of the combustion section or in dependence of characteristics of the combustion section and its components. For example, the rim surface may form/have a cylindrical shape and may be inclined in reference to an outer surface (side face) of the wing in which the opening is arranged. Preferably, the inclination built/has an angle with the flow direction of the flow medium passing the opening of about 120° to 180°, preferably of about 135° to 170° and most preferably of about 140° to 160°.
Furthermore, it is advantageous when the at least one opening is constant along its extension. Hence, injection can be performed with constant pressure. Expressed differently, the at least one opening is neither narrowed nor enlarged along its extension. Or in other words, two opposed arranged portions of the rim surface are arranged basically in parallel towards each other or are arranged equidistantly along the extension or depth of the opening. Generally, it may be also possible to orient opposed portions of the rim surface differently. For example, the rim surface may have a cone-like shape either in direction of injection or against the injection direction. Due to this, the opening may be function like a nozzle.
In a further embodiment of the invention, the downstream part or the trailing edge, respectively, ends in a tip, wherein the opening is arranged in the tip. Hence, fuel is directly injected from the tip. As a result, the fuel is injected directly in flow direction of the flow medium. Hence, the injected fuel does not have to change direction, minimising the risk of low speed zones in fuel rich regions. Moreover, the mixing of the injected fuel and the passing flow medium travelling the flow channel mixes downstream of the training edge or tip of the wing. Hence, an attachment of fuel to the downstream part of the wing or its hot surfaces can be avoided. This in turn, minimises the risk for flashbacks.
According to a preferred refinement of the invention, the at least one opening is arranged at a suction side of each wing. Hence, the fuel can be injected into the flow channel of the flow medium easily. In other words, the opening is arranged in a side face of the downstream part or trailing edge, respectively, and is opening out into the flow channel.
The flow of the flow medium is accelerating from an outer part of the swirler arrangement towards a "throat" of the swirler arrangement. There is a low speed zone at the upstream part of the wing where the flow is starting to accelerate. Hence, in this embodiment the fuel is injected slightly upstream of the wing tip in a low speed zone of the burner. Consequently, the tip is located downstream of the opening.
A thick trailing edge of the wing has large re-circulation zones behind it as a result of the surrounding flow. Keeping the wing trailing edge to a minimum reduces or minimizes these re-circulation zones, and thereby the low speed zones. Hence, it is especially advantageous, when the downstream part or the trailing edge, respectively, ends in a sharp tip. The phrase "sharp tip" should be understood as a tapering/tapered or pointed tip. Thus, the trailing edge thickness should be very small to minimise the risk of low speed zones in regions where fuel is present.
The highest velocity of the flow medium is found close to the flow channel between two wings (throat). Low speed zones are not desired here since this throat will partially define the pressure drop across the burner of the combustion section. The conditions in the region behind the tip are a result of this high speed zone in the throat. When the high speed flow of flow medium passes the sharp wing tip, vortices will form with local low speed zones close to the tip as a result. However, the sharp tip can minimise the risk for low speed zones after or downstream, respectively, of the wing. The construction of the swirl generator is a fine balance between the amount of vortices behind the tip and an acceptable pressure drop across the burner.
In addition, it is advantageous when each wing of the at least two wings comprises an inside fuel manifold suppling the at least one opening with fuel. Thus, the fuel feed can be realised constructively easily. Preferably, each wing may comprise a "storage" cavity for the fuel in the upstream part of the wing. A section of the fuel manifold may connect the storage cavity with the opening.
According to a further aspect of the invention the swirler arrangement may comprise three or a plurality of wings, wherein at least two of which form/built/shape the/said tread-like flow channel. Hence, the needed swirling motion of the flow medium can be effectively created.
A material of each wing may be any material feasible for a skilled person, and beneficially, the material of the wing is a material selected out of the group consisting of: a nickel based super alloy, Inconel (In) 1792, In 939, In 738, Alloy 247 or Hastelloy X. Hence, a wide range of suitable materials are available. Favourably, the material is Hastelloy X. According to a preferred refinement, each wing is manufactured by additive manufacturing. In other words, each wing is built or formed by a method of manufacturing, wherein each wing is manufactured by additive manufacturing or 3D printing, respectively.
The invention further refers to a burner comprising at least one swirler arrangement as described above and a mixing zone. It is proposed that the swirler arrangement is arranged upstream of the mixing zone.
Moreover, the invention further refers to a turbine engine, especially a gas turbine engine, with a burner comprising at least one swirler arrangement as described above and a mixing zone.
Due to these inventive matters a burner or a whole combustion section and a turbine engine that are less prone to failures can be provided. Especially, the flashback resistance of the burner of the combustion section can be increased since the risk for the exhausted or injected fuel to come into contact with the hot burner geometry can be minimized. Moreover, a secure operation of the turbine engine even when operated with highly reactive fuels can be ensured.
The invention further relates to (a) use(s) of the above described swirler arrangement and/or the above described burner. Favourably, it is provided that the flow medium being swirled is primary air and the exhausted/injected fuel is natural gas with a high content of hydrogen. In this context, the phrase "with high hydrogen content" should be understood as a hydrogen content above 10% and preferably above 15%. Beneficially, the exhausted/injected fuel is natural gas with hydrogen content above 15%.
The previously given description of advantageous embodiments of the invention contains numerous features which are partially combined with one another in the dependent claims. Expediently, these features can also be considered individually and be combined with one another into further suitable combinations. Furthermore, features of the method, formulated as apparatus features, may be considered as features of the assembly and, accordingly, features of the assembly, formulated as process features, may be considered as features of the method.
The above-described characteristics, features and advantages of the invention and the manner in which they are achieved can be understood more clearly in connection with the following description of exemplary embodiments which will be explained with reference to the drawings. The exemplary embodiments are intended to illustrate the invention, but are not supposed to restrict the scope of the invention to combinations of features given therein, neither with regard to functional features. Furthermore, suitable features of each of the exemplary embodiments can also be explicitly considered in isolation, be removed from one of the exemplary embodiments, be introduced into another of the exemplary embodiments and/or be combined with any of the appended claims.

Brief Description of the Drawings

The present invention will be described with reference to drawings in which:

FIG 1:: shows schematically a sectional view of a gas turbine engine comprising a burner with an inventive swirler arrangement,
FIG 2:: shows a sectional view of the burner of FIG 1 with the swirler arrangement,
FIG 3:: shows a sectional view of the swirler arrangement of FIG 2,
FIG 4:: shows in a perspective view one wing of the swirler arrangement of FIG 2 and
FIG 5:: shows a sectional view of an alternatively embodied swirler arrangement.

Detailed Description of the Illustrated Embodiments

The present invention is described with reference to an exemplary turbine engine 14 having a single shaft 62 or spool connecting a single, multi-stage compressor section 54 and a single, one or more stage turbine section 58. 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.
The terms upstream and downstream refer to the flow direction 24' of the flow medium 22 and/or airflow and/or working gas flow through the engine 14 unless otherwise stated. If used, the terms axial, radial and circumferential are made with reference to a rotational axis 60 of the engine 14.
FIG 1 shows an example of a gas turbine engine 14 in a sectional view. The gas turbine engine 14 comprises, in flow series, an inlet 54, a compressor section 56, a combustion section 12 and a turbine section 58, which are generally arranged in flow series and generally in the direction of a longitudinal or rotational axis 60. The gas turbine engine 14 further comprises a shaft 62 which is rotatable about the rotational axis 60 and which extends longitudinally through the gas turbine engine 14. The shaft 62 drivingly connects the turbine section 58 to the compressor section 56.
In operation of the gas turbine engine 14, flow medium 22 or air 64 which is taken in through the air inlet 54 is compressed by the compressor section 56 and delivered to the combustion section or burner section 12. The burner section 12 comprises a burner plenum 66, one or more combustion chambers 68 defined by a double wall can 70 and at least one burner 72 fixed to each combustion chamber 68. The at least one combustion chambers 68 and the burners 72 are located inside the burner plenum 66. The compressed air passing through the compressor section 56 enters a diffuser 74 and is discharged from the diffuser 74 into the burner plenum 66 from where a portion of the air 64 enters the burner 72 and is mixed with a gaseous or liquid fuel (details see below). The air/fuel mixture is then burned and the combustion gas 76 or working gas from the combustion is channelled via a transition duct 78 to the turbine section 58.
This exemplary gas turbine engine 14 has a cannular combustor section arrangement 80, which is constituted by an annular array of combustor cans 70 each having the burner 72 and the combustion chamber 68, the transition duct 78 has a generally circular inlet that interfaces with the combustion chamber 68 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 section 58.
The turbine section 58 comprises a number of turbine assemblies embodied as blade carrying discs 82 or turbine wheels 84 attached to the shaft 62. In the present example, the turbine section 58 comprises two discs 82 each carry an annular array of turbine aerofoils, each embodied as a turbine blade. However, the number of blade carrying discs 82 could be different, i.e. only one disc 82 or more than two discs 82. In addition, turbine assemblies embodied as turbine cascades 86 are disposed between the turbine blades. Each turbine cascade 86 carries an annular array of aerofoils, in the form of guiding vanes, which are fixed to a stator 88 of the gas turbine engine 14, and an inner and outer platform. Between the exit of the combustion chamber 68 and the leading turbine blades inlet guiding vanes or nozzle guide vanes 90 are provided. They turn and accelerate the flow of working gas 76 onto the turbine blades.
The combustion gas 76 from the combustion chamber 68 enters the turbine section 58 and drives the turbine blades which in turn rotate the shaft 62. The guiding vanes 90 serve to optimise the angle of the combustion or working gas 76 on to the turbine blades. The turbine section 58 drives the compressor section 56. The compressor section 56 comprises an axial series of turbine assemblies embodied as guide vane stages 92 and rotor blade stages 94. The rotor blade stages 94 comprise a rotor disc 82 supporting an annular array of turbine blades.
The compressor section 56 also comprises a stationary casing 96 that surrounds the rotor stages 94 in circumferential direction 98 and supports the vane stages 92. The guide vane stages 92 include an annular array of radially extending turbine aerofoils embodied as vanes that are mounted to the casing 96. 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 92 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.
The casing 96 defines a radially outer surface 100 of a passage 102 of the compressor section 56. A radially inner surface 104 of the passage 102 is at least partly defined by a rotor drum 106 of the rotor which is partly defined by the annular array of blades.
FIG 2 shows a sectional view of a burner 72 in the form of a so-called DLE-burner 52 (DLE: dry low emission) of the combustion section 12. The burner 52 comprises in flow direction 24' of the flow medium 20 an entry zone 108, a swirler arrangement 10 or swirl generator, a mixing zone 50 or reacting zone and a combustion zone 110. Compressed or primary air 64 is entering the swirler arrangement 10 from the compressor section 56 via the entry zone 108. Fuel 32, like natural gas with high hydrogen content, exemplarily, above 15%, is injected in the swirl generator and mixed in the mixing zone 50 with the primary air 64 (details see below). Hence, the swirler arrangement 10 is arranged upstream of the mixing zone 50. Downstream of the mixing zone 50 the fuel/air-mixture is burnt in the combustion zone 110.
The swirler arrangement 10 comprises a plurality and in this exemplary embodiment four wings 16, 18 (see FIG 3, only two wings are marked with reference numerals). Each wing 16, 18 has an aerofoil-like shape with a leading edge 34 and a trailing edge 36. For better understanding only one wing 16 will be described in the following text and when needed the wing 16 will be described in reference to a second wing 18. In respect to the flow direction 24 of the flow medium 22 the wing 16 comprises an upstream part 26 and a downstream part 28. The upstream part 26 comprises the leading edge 34 and the downstream part 28 the trailing edge 36. Moreover, the wing 16 is connected with a support structure 112 of the swirler arrangement 10 via its leading edge 34 or the upstream part 26 (not shown in detail). The downstream part 28 or a tail 114 of the wing 16 that ends in the trailing edge 36 extends from the upstream part 26 into free space and basically in radial direction 116. Further, it points in direction of a pressure side 118 of the adjacent wing 18.
The swirler arrangement further comprises thread-like flow channels 20 wherein one of the flow channels 20 each is arranged between two wings 16, 18. Thus, two wings 16, 18 each form a flow channel 20. Each flow channel 20 is allowing the flow medium 22 to travel along the flow channel 20 in the flow direction 24. This flow direction 24 may vary from the flow direction 24' of the flow medium 22 traveling through the turbine engine 14. When flowing through the flow channel 20 the flow medium 22 is set into a thread-like motion by the flow channel 20 or the shape and orientation of the wings 16, 18 or their tails 114.
Furthermore, as could be especially seen in FIG 4, which shows one wing 16 of the swirler arrangement 10 in a perspective view, the downstream part 28 of the wing 16 comprises a plurality of openings 30 for injecting fuel 32 into the flow medium 22 while the flow medium 22 travels the flow channel 20. Hence, the openings 30 are arranged at the trailing edge 36 and are arranged homogeneously over a flow width w of the flow channel 20 (see FIG 2). The openings 30 have a circular shape.
The downstream part 28 of the wing 16 of the embodiment shown in FIG 3 ends in a blunt or flat tip 42. The openings 30 are arranged in the tip 42. Further, each opening 30 is surrounded by a rim surface 38. This rim surface 38 is embodied in such that the exhaustable fuel 32 adopts a flow component oriented in flow direction 24 of the flow medium 22 along the flow channel 20. In other words, the fuel 32 in injected with a 0° angle in respect to the flow direction 24 of the flow medium 22. Therefore the rim surface 38 is oriented in parallel to a side face 120 of the downstream part 28 or a suction side 44 of the wing 16. Hence, the opening 30 is constant or, in other words, neither narrowed nor enlarged along its extension 40.
To supply the openings 30 with fuel 32 the wing 16 comprises an inside fuel manifold 48 with a storage cavity 122 (the fuel manifold 48 is only shown in broken lines for the wing 16). Moreover, a material of the wing 16 is a material selected out of the group consisting of: a nickel based super alloy, Inconel (In) 1792, In 939, In 738, Alloy 247 or as preferred Hastelloy X. Furthermore, each wing 16, 18 is manufactured by additive manufacturing.
Due to this construction a risk for low speed zones in fuel rich regions can be minimised. Moreover, this fuel injection angle hinders the possibility of fuel contacting hot surfaces of the burner geometry and thus minimises the risk for flashbacks. Further, it allows the advantageous use of the swirler arrangement 10 in a DLE-burner 52 as well as a beneficial use of the burner 52 itself.
The fuel pressure is dependent on the actual load of the engine 14, wherein the maximum fuel pressure is 25 bar. Moreover, the pressure and temperature are similar to the compressor outlet conditions, which are engine dependent. For example, for the so-called SGT-800 gas turbine engine 14 the pressure is about 20 bar and the temperature is about 450°C. The equivalence ratio (ratio of actual fuel to air ratio to the stoichiometric fuel to air ratio) is around 0.5, i.e. the fuel air mixture is fuel lean.
In FIG 5 an alternative embodiment of the swirler arrangement 10 is shown. Components, features and functions that remain identical are in principle substantially denoted by the same reference characters. To distinguish between the embodiments, however, the letter "a" has been added to the different reference characters of the embodiment in FIG 5. The following description is confined substantially to the differences from the embodiment in FIG 1 to 4, wherein with regard to components, features and functions that remain identical reference may be made to the description of the embodiment in FIG 1 to 4.
FIG 5 shows an alternative swirler arrangement 10a of a combustion section 12 of a gas turbine engine 14. The embodiment from FIG 5 differs in regard to the embodiment according to FIG 1 to 4 in that the openings 30a are positioned in the downstream part 28a slightly upstream of the actual trailing edge 36.
In detail, the openings 30a are arranged at a suction side 44 of the wing 16a to inject fuel 32 into a flow medium 22 that is traveling a thread-like flow channel 20 that is arranged between two wings 16a, 18a. The flow channel 20 is allowing the flow medium 22 to travel along the flow channel 20 in a flow direction 24, wherein the flow medium 22 is set into a thread-like motion by the flow channel 20 when flowing through the flow channel 20. In other words, the openings 30a are arranged in a side face 120 of the downstream part 28a or trailing edge 36, respectively, and are opening out into the flow channel 22.
The openings 30a are surrounded by a rim surface 38 that is embodied in such that the exhaustable fuel 32 adopts a flow component oriented in flow direction 24 of the flow medium 22 along the flow channel 20. In this exemplary embodiment the rim surface 38 has a cylindrical shape and is inclined in reference to the side face 120 of the wing 16 in which the openings 30a are arranged. The inclination has an angle α with the flow direction 24 of the flow medium 22 passing the openings 30a of about 135°. Further, each opening 30a is constant or neither narrowed nor enlarged along its extension 40.
Moreover, the downstream part 28a ends in a sharp tip 46. Due to this, low speed zones downstream of the tip 46 can be advantageously avoided. Furthermore, the orientation of the rim surface 38 or its injection angle α allows the fuel 32 to be injected without the risk for the fuel 32 contacting surfaces of the hot burner geometry as well as avoids low speed zones in the vicinity of the fuel injection points (openings 30a).
It should be noted that the term "comprising" does not exclude other elements or steps and "a" or "an" does not exclude a plurality. Also elements described in association with different embodiments may be combined. It should also be noted that reference signs in the claims should not be construed as limiting the scope of the claims.
Although the invention is illustrated and described in detail by the preferred embodiments, the invention is not limited by the examples disclosed, and other variations can be derived therefrom by a person skilled in the art without departing from the scope of the invention.

Claims

Swirler arrangement (10; 10a) of a burner (52), in particular of a gas turbine engine (14), comprising at least two wings (16, 18; 16a, 18a), wherein a thread-like flow channel (20) is arranged between the at least two wings (16, 18; 16a, 18a), wherein the flow channel (20) is allowing a flow medium (22) to travel along the flow channel (20) in a flow direction (24, 24') and wherein the flow medium (22) is set into a thread-like motion by the flow channel (20) when flowing through the flow channel (20) and wherein each wing (16, 18; 16a, 18a) of the at least two wings (16, 18; 16a, 18a) comprise an upstream part (26) and a downstream part (28; 28a),
characterised in that the downstream part (28; 28a) comprises at least one opening (30; 30a) for injecting fuel (32) into the flow medium (22) while traveling the flow channel (20).
Swirler arrangement according to claim 1, wherein
each wing (16, 18; 16a, 18a) of the at least two wings (16, 18; 16a, 18a) has an aerofoil-like shape with a trailing edge (36) and wherein the at least one opening (30; 30a) is arranged at the trailing edge (36).
Swirler arrangement according to claim 1 or 2,
characterised by
a plurality of openings (30; 30a) arranged over a flow width (w) of the flow channel (20).
Swirler arrangement according to any one of the preceding claims, wherein
the at least one opening (30; 30a) has a circular shape.
Swirler arrangement according to any one of the preceding claims, wherein
the at least one opening (30; 30a) is surrounded by a rim surface (38) embodied in such that the exhaustable fuel (32) adopts a flow component oriented in flow direction (24) of the flow medium (22) along the flow channel (20).
Swirler arrangement according to any one of the preceding claims, wherein
the at least one opening (30; 30a) is constant along its extension (40).
Swirler arrangement according to any one of the preceding claims, wherein
the downstream part (28) ends in a tip (42), wherein the at least one opening (30) is arranged in the tip (42).
Swirler arrangement according to any one of claim 1 to 6, wherein
the at least one opening (30a) is arranged at a suction side (44) of each wing (16a, 18a).
Swirler arrangement according to claim 8, wherein
the downstream part (28a) ends in a sharp tip (46).
Swirler arrangement according to any one of the preceding claims, wherein
each wing (16, 18; 16a, 18a) of the at least two wings (16, 18; 16a, 18a) comprises an inside fuel manifold (48) suppling the at least one opening (30; 30a) with fuel (32).
Swirler arrangement according to any one of the preceding claims, characterised by
three or a plurality of wings (16, 18; 16a, 18a), wherein at least two (16, 18; 16a, 18a) of which form the tread-like flow channel (20).
Swirler arrangement according to any one of the preceding claims, wherein
a material of the wing (16, 18; 16a, 18a) is a material selected out of the group consisting of: a nickel based super alloy, Inconel (In) 1792, In 939, In 738, Alloy 247 or Hastelloy X.
Swirler arrangement according to any one of the preceding claims, wherein
each wing (16, 18; 16a, 18a) is manufactured by additive manufacturing.
Burner (52) comprising at least one swirler arrangement (10; 10a) according to any one of claims 1 to 13 and a mixing zone (50),
wherein the swirler arrangement (10; 10a) is arranged upstream of the mixing zone (50).
Use of the swirler arrangement (10; 10a) according to any one of claims 1 to 13 and/or a burner (52) according to claim 14 characterised in that
a flow medium (22) being swirled is primary air (64) and an exhausted fuel (32) is natural gas with a hydrogen content above 15%.