FIELD OF INVENTION

The present invention relates to a combustor for a gas turbine and particularly but not exclusively a combustor suitable for burning fuel with a significant hydrogen content or a high hydrocarbon content without flashback.

BACKGROUND OF INVENTION

EP3278029A discloses a dry low emission (DLE) combustion system comprising a pilot burner assembly surrounded by a main burner assembly. The pilot burner comprises a pilot fuel lance and an igniter, both located at the burner surface. The main burner assembly comprises an annular array of vanes known as a radial swirler. Gas fuel is injected from two locations in each swirler slot formed between adjacent vanes of the radial swirler. A first location for main gas injection is in the base of each slot and at the entrance of the slot. The fuel injection hole is in a counter-bore to help mixing of the fuel and air. The second location for main gas injection is via two injection holes in the side of each vane in each slot. This second gas injection is into the crossflow of air passing through the slot. This DLE combustion system design creates two distinct flames formed in the combustion chamber. The flame generated from the main fuel injection which is partially premixed produces the main flame and fuel injected from pilot burner assembly produces a diffusion or pilot flame.

However, this design is very susceptible to flame flashback when there is a high hydrogen content in the fuel. When higher hydrocarbon (HC) fuels such as propane and butane are burnt, this DLE system is also prone to flame flashback. This flashback is mainly due to higher flame speeds of Hydrogen and high HC fuel. Flame flashback is undesirable because it can extinguish the combustion flame, cause increased emissions and damage to components of the combustion system.

SUMMARY OF INVENTION

Thus, an object of the presently disclosed combustor is to provide a combustion system which prevents flame flashbacks yet provides efficient combustion with low emissions. Another object of the presently disclosed combustor is to provide better mixing of fuel and air. Another object of the presently disclosed combustor is to provide a more stable combustion process. The above objects are achieved by a combustor of a gas turbine, the combustor comprising a combustor axis about which is arranged a burner and a combustion chamber, the burner comprises a main fuel injector, a radial swirler and a mixing tube, the mixing tube has an inlet and an outlet, the inlet is connected to the radial swirler and the outlet is connected to the combustion chamber, the radial swirler surrounds the main fuel injector and comprises an annular array of vanes. In use compressed air passes through the radial swirler and forms a main vortex which combines with fuel from the main fuel injector and mixes in the mixing tube and then enters the combustion chamber to burn. The burner comprises a liner that surrounds the main fuel injector,

the liner is connected to the other axial side of the radial swirler to the mixing tube, the liner forms an annular gap with an outer surface of the main fuel injector,

the liner comprises and array of air holes through which compressed air passes in use.

The main fuel injector may comprise an injector tube having a fuel inlet end and a fuel outlet end, the fuel outlet end is located axially closer to the combustion chamber than the fuel inlet end, the fuel outlet end has an array of fuel holes distributed around the tube.

The fuel outlet end may have a cylindrical portion and the array of fuel holes distributed around the tube is within the cylindrical portion.

The fuel outlet end may have a divergent portion and the array of fuel holes distributed around the tube is within the divergent portion.

The fuel outlet end may have a convergent portion and the array of fuel holes distributed around the tube is within the convergent portion.

The fuel holes of the array of fuel holes each have a centre line, and at least one of the fuel holes of the array of fuel holes has its centre line perpendicular to the outer surface in the axial sense; preferably, all the fuel holes of the array of fuel holes have their centre line perpendicular to the outer surface in the axial sense.

The fuel holes of the array of fuel holes each may have a centre line, and at least one of the fuel holes of the array of fuel holes may have its centre line angled with an axial component to the outer surface in the axial sense; preferably, all the fuel holes of the array of fuel holes may have their centre line angled with an axial component to the outer surface in the axial sense. The fuel holes of the array of fuel holes each may have a centre line or the centre line, and at least one of the fuel holes of the array of fuel holes may have its centre line at a tangential angle component relative to the combustor axis; preferably, all the holes of the array of fuel holes may have their centre line at a tangential angle component relative to the combustor axis.

The array of fuel holes may comprise at least two rows of fuel holes, preferably the fuel holes of one row are staggered relative to the fuel holes of another row.

The liner may comprise an array of air holes that, in use, guides air into the annular gap.

The fuel holes of the array of holes each have a centre line or the centre line, and at least one of the air holes of the array of air holes may have its centre line at a tangential angle component relative to the combustor axis; preferably, all the air holes of the array of air holes may have their centre line at a tangential angle component relative to the combustor axis.

The array of air holes may comprise at least two rows of air holes, preferably the air holes of one row are staggered relative to the air holes of another row.

The swirler vanes may be arranged to produce a clockwise or anticlockwise rotating main vortex about the combustor axis and the array of air holes may be arranged to produce an oppositely rotating centre vortex the about the combustor axis relative to the main vortex.

The fuel holes may be arranged to inject fuel in the same rotational direction as the rotating centre vortex.

In another and separate aspect of the present disclosure and that has the same objectives, there is provided a combustor of a gas turbine, the combustor comprising a combustor axis about which is arranged a burner and a combustion chamber, the burner comprises a main fuel injector, a pilot fuel injector, a radial swirler and a mixing tube, the mixing tube has an inlet and an outlet, the inlet is connected to the radial swirler and the outlet is connected to the combustion chamber, the radial swirler comprises an annular array of vanes extending in the axial direction from a base plate to a top plate, the vanes are arranged to have a swirl angle, the vanes have a trailing edge, in use compressed air passes through the radial swirler and forms a main vortex which combines with fuel from at least the main fuel injector and mixes in the mixing tube and then enters the combustion chamber to burn, wherein the trailing edge is divergent relative to the combustor axis between the base plate and the top plate and in the direction from the base plate towards the top plate.

The trailing edge may have a divergent angle Q from the combustor axis.

The mixing tube may comprise at least a first portion that is divergent between the inlet and the outlet, and the first portion is divergent in the direction from the inlet to the outlet, the first portion has a divergent angle Q from the combustor axis, wherein the trailing edge has the same divergent angle Q as the first portion.

The combustor may comprise a total compressed air flow, wherein a proportion of the compressed air passing through the array of fuel holes in the liner may be in the range 20% to 40%, preferably about 30%, of the total compressed air.

The combustor comprising at least one fuel supply having at least one fuel type, the at least one fuel supply feeds fuel to at least one of the main fuel injectors and / or the pilot fuel injector.

The least one fuel type comprises a fuel composition having at least 5%, preferably at least 20%, by volume hydrogen and / or a high-HC content with a fuel heating value greater than 45MJ/m ³ .

BRIEF DESCRIPTION OF THE DRAWINGS

The above-mentioned attributes and other features and advantages of the present technique and the manner of attaining them will become more apparent and the present technique itself will be better understood by reference to the following description of embodiments of the present technique taken in conjunction with the accompanying drawings, wherein:

FIG. 1 shows part of a turbine engine in a sectional view and in which the present combustor arrangement is incorporated,

FIG. 2 is a schematic cross-section through a known combustor, FIG. 3 is a schematic cross-section through a first embodiment of the presently disclosed combustor, pre-chamber, mixing tube and fuel injection apparatus and which may be incorporated into the turbine engine shown and described with reference to FIG.l,

FIG. 4 is a part schematic cross-section through the combustor and pre-chamber showing a pilot fuel injector in more detail,

FIG. 5 is a view on arrow A shown in FIG. 4 and shows further detail of the pilot fuel injector arrangement,

FIG. 6 is a schematic section through the mixing tube and shows possible details of an array of holes in the mixing tube,

FIG. 7 is a section B-B through the mixing tube shown in FIG. 6 and showing further possible details of the array of holes in the mixing tube,

FIG. 8 is a perspective view of a single swirler vane of the swirler of the known combustor shown in FIG. 2,

FIG. 9 is a perspective view of a single swirler vane of the presently disclosed swirler of the combustor shown in FIG. 3,

FIGS. 10, 11, 12 and 15 are schematic side views, partly in section, of a number of possible arrangements of a main fuel injector of the presently disclosed combustor shown in FIG. 3 inter alia,

FIGS. 13 and 14 are schematic sections C-C shown in FIG. 10 of the main fuel injector and shown details of an array of main fuel injection holes.

DETAILED DESCRIPTION OF INVENTION Figure 1 is a schematic illustration of a general arrangement of a turbine engine 10 having an inlet 12, a compressor 14, a combustor system 16, a turbine system 18, an exhaust duct 20 and a twin-shaft arrangement 22, 24. The turbine engine 10 is generally arranged about an axis 26 which for rotating components is their rotational axis. The shafts of the twin-shaft arrangement 22, 24 may have the same or opposite directions of rotation. The combustor system 16 comprises an annular array of combustor units 36, only one of which is shown. In one example, there are six combustor units evenly spaced about the engine. The turbine system 18 includes a high-pressure turbine 28 drivingly connected to the compressor 14 by a first shaft 22 of the twin-shaft arrangement. The turbine system 18 also includes a low-pressure turbine 30 drivingly connected to a load (not shown) via the second shaft 24 of the twin-shaft arrangement.

The terms radial, circumferential and axial are with respect to the engine's rotational axis 26 or as otherwise stated. The terms upstream and downstream are with respect to the general direction of gas flow through the engine and as seen in FIG.l (and later FIG. 3) is generally from left to right.

The compressor 14 comprises an axial series of stator vanes and rotor blades mounted in a conventional manner. The stator or compressor vanes may be fixed or have variable geometry to improve the airflow onto the downstream rotor or compressor blades. Each turbine 28, 30 comprises an axial series of stator vanes and rotor blades. The stator vanes can be mounted to a radially outer casing or a radially inner drum. The rotor blades are mounted via rotor discs arranged and operating in a conventional manner. A rotor assembly comprises an annular array of rotor blades or blades and the rotor disc.

Each combustor unit 36 is constructed from two walls, an inner wall 37 and an outer wall 39, between which is defined a generally annular space. At the head of the combustor unit 36 is a swirler 40 which comprises a swirl plate, an annular array of swirler vanes 46 and fuel injection points as will be described in more detail later. The swirler 40 is succeeded by a pre-chamber 42 and then a main combustion chamber 38. These combustor unit 36 components are generally arranged about a combustor central axis 44.

In operation air 32 is drawn into the engine 10 through the inlet 12 and into the compressor 14 where the successive stages of vanes and blades compress the air before delivering the compressed air 34 into the combustor system 16. The compressed air 34 flows between the inner and outer walls 37, 39 and into the swirler 40. The swirler 40 creates highly turbulent air into which the fuel is injected. The air / fuel mixture is delivered into the pre-chamber 42, where mixing continues, and then into the main combustion chamber 38. In the combustion chamber 38 of the combustion unit 16 the mixture of compressed air and fuel is ignited and burnt. The resultant hot working gas flow is directed into, expands and drives the high-pressure turbine 28 which in turn drives the compressor 14 via the first shaft 22. After passing through the high-pressure turbine 28, the hot working gas flow is directed into the low-pressure turbine 30 which drives the load via the second shaft 24.

The low-pressure turbine 30 can also be referred to as a power turbine and the second shaft 24 can also be referred to as a power shaft. The load is typically an electrical machine for generating electricity or a mechanical machine such as a pump or a process compressor. Other known loads may be driven via the low-pressure turbine. The fuel may be in gaseous and/or liquid form.

The turbine engine 10 shown and described with reference to FIG.l is just one example of a number of engines or turbomachinery in which this invention can be incorporated.

Such engines can be gas turbines or steam turbine and include single, double and triple shaft engines applied in marine, industrial and aerospace sectors.

FIG.2 is a cross-section through part of a known combustor unit 36 of the turbine engine 10 described above. The swirler 40 comprises an annular array of vanes 46 which are angled relative to the combustor axis 44 to impart a swirling flow 55 of mixing air and fuel as is well known. The swirling flow 55 rotates about the combustor axis 44 and flows in a general left to right direction as seen in FIG.2. The vanes 46 form an array of mixing channels 47 between each vane 46. The swirler 44 further comprises main fuel injectors 48A, 48B and pilot fuel injectors 50. The swirler 40 has a pilot surface 52 which faces the pre-chamber 42 and bounds the pre-chamber's upstream axial extent. The pre-chamber 42 is further defined by an annular wall 54 which has parallel sides. The pre-chamber 42 has an inlet 66 and an outlet 68. The outlet 68 forms or is at a lip 70 of the pre-chamber 42 and where the pre-chamber 42 terminates. The pre-chamber 42 walls 54 are then succeeded by the wall(s) 37 of the main combustion chamber 38. From the lip 70 the wall 37 is divergent and opens to the main combustion chamber 38 which has a greater cross-sectional area than that of the pre-chamber 42.

There are two distinct fuel / air mixtures and subsequently combustion flames in the combustion chamber 38; a pilot flame 56 is derived from the pilot fuel supply 50 and the main flame 58 is derived from the main fuel supply 48A, 48B. The pilot and main flames are distinct from one another because of the location of the respective fuel injection points into the air flow in or near to the mixing channel(s) 47. The main fuel injectors 48A, 48B inject fuel into the mixing channel further away from the pilot surface 52 than the pilot fuel injector(s) 50. Thus, the respective fuel / air mixtures form substantially different flame regions with the pilot flame 56 generally radially inward of the main flame 58.

Radial swirlers, as in the case here, have or can be defined as having, a swirl number SN. The swirl number can be calculated as is well known in the art, suffice to say here, that the swirl number can be defined by a relationship between the fluxes of angular and linear momentum of the fuel / air mixture. That is to say, the angular momentum relates to rotational velocity about the combustor axis 44 and the linear momentum relates to the velocity in the axial direction along the combustor axis 44. Thus, the SN is defined herein as the ratio of tangential momentum to axial momentum of the fluid or fuel / air mixture.

The general schematic cross section of FIG.2 shows a Dry Low Emissions (DLE) combustor 36. The known swirler 40 described above has a SN in the region 0.5 to 0.8. This combustor provides a good DLE burner for combusting methane and medium calorific value fuels (MCV fuels) containing hydrocarbons. However, this current design is not suitable for burning fuel with significant (e.g. >5% by weight) hydrogen or

high-hydrocarbon content mainly due to dominance of flame speed on the flow characteristics. The presence of hydrogen increases flame speed and causes flash-back into the pre-chamber 42. This is clearly detrimental and undesirable and can cause extinction of the flame and increased emissions of nitrous oxides, sulphur oxides and unburned hydrocarbons amongst other undesirable combustion by products.

Reference is now made to the present disclosure of a new combustor 70 as shown in FIG.3. The combustor 70 is implemented in the gas turbine 10 as shown in Fig. 1 and the combustor 70 is shown in schematic cross-section. The combustor 70 has a combustor axis 72 about which is arranged a burner 74 and a combustion chamber 76. The burner 74 comprises a main fuel injector 78, a radial swirler 80 and a mixing tube 82. The mixing tube 82 has an inlet 84 and an outlet 86. The inlet 84 of the mixing tube 82 is connected to the radial swirler 80. The outlet 86 or near the outlet-end of the mixing tube 82 is connected to the combustion chamber 76. The radial swirler 80 surrounds or is radially outward, with respect to the combustor axis 72, of the main fuel injector 78 and comprises an annular array of vanes 88. The vanes 88 are arranged to have a swirl angle as known in the art and extend in an axial direction from a base plate 114 to a top plate 116. The base plate 114 and the top plate 116 are generally annular rings. The base plate 114 has an exposed or gas washed surface 158 over which a flow of air 34A passes in use and through the slots defined between adjacent vanes 88. The annular array of vanes 88 form channels between successive vanes. The channels are both radially aligned and tangentially angled. Thus, the flow of air 34A is directed in a radially inward direction as it passes through the annular array of vanes 88 and in a tangential direction or

circumferential direction. Once in the mixing-tube 82 the air (and fuel mixture) rotates or swirls around the combustor axis 72 and generally flows in an axial direction towards the main combustion chamber 76. In this example, the radial swirler 80 comprises channels that are radially aligned, although in other examples the channels may be angled to a radial line up to 15°.

A pilot fuel injector 90, or at least it's fuel outlet, is located downstream of or preferably at the outlet end 86 of the mixing tube 82 and is described in more detail with reference to Fig 4 below. However, briefly, the pilot fuel injector 90 issues pilot fuel directly into the combustion chamber 76 to form a pilot flame or a diffusion flame 118, also denoted as zone 1.

An igniter 170 is located through the wall of the combustion chamber 76 and otherwise operates in known fashion. There may be one or more than one igniter 170 in each combustion chamber 76. The igniter 170 is cooled with a relatively small portion of the compressed air 34.

The mixing tube 82 comprises at least a first portion 108 that is divergent between the inlet 84 and the outlet 86, and the first portion 108 is divergent in the direction from the inlet 84 to the outlet 86. In this embodiment, the mixing tube 82 comprises a second portion 110, which has a generally constant cross-sectional area. The second portion 110 is downstream of the first portion 108 and is connected directly to the first portion 108. The mixing tube 82 is of a unitary construction and may be cast as a monolithic structure or the first and second portions may be made separately and then joined by known techniques. In other embodiments, the first portion 108 may extend from the inlet 84 to the outlet 86 and therefore the mixing tube 82 does not have a constant cross-section second portion 110. Alternatively, the second portion 110 may be divergent, but have a smaller divergent angle than the first portion 108.

The mixing tube 82 comprises an array of holes 112 through which compressed air 34A passes in use. The array of holes 112 is located in the divergent or first portion 108 of the mixing tube 82. In other embodiments, the array of holes 112 may extend over part or all of the second portion 110 in any one of the embodiments described herein.

The straight or constant cross-sectional area second portion 110 of the mixing tube 82 assists in controlling flame propensity; that is the natural tendency of the location of the flame front as indicated by the distance L in Fig.3. Thus, the axial length of the constant cross-sectional area second portion 110 may be selected to control where the front of the main flame is located.

The functionality of the first portion 108 having the array of holes 112 and compressed airflow 34C provides an aerodynamic feature to maintain the main flame in a lifted position away from the combustor's surfaces and prevent the flame from attaching to the surfaces of the combustor. In addition, the second portion 110 with or without an array of holes and additional compressed airflow therethrough, further assists with this aerodynamic feature as well as providing additional axial length of the mixing tube 82 to achieve optimal air and fuel mixing.

The burner 74 further comprises a liner 92 or tubular portion that surrounds the main fuel injector 78. The liner 92 is connected to the other axial end, or opposite end, of the radial swirler 80 to the mixing tube 82. The liner 92 is generally upstream of the radial swirler 80 and the mixing tube 82 is generally downstream of the radial swirler 80. The liner 92 and the main fuel injector 78 are generally coaxial and arranged to have a common centre line that being the combustor's axis 72. The liner 92 forms an annular gap 94 with an outer surface 96 of the fuel injector 78. The liner 92 comprises an array of air holes 98 through which compressed air 34B passes in use and into the gap 94.

The main injector 78 comprises an injector tube 100 having a fuel inlet end 102 and a fuel outlet end 104. The fuel outlet end 104 is located axially closer to the combustion chamber 72 than the fuel inlet end 102. The fuel outlet end 104 has an array of main fuel holes 106 distributed around the injector tube 100. In use, gaseous fuel is supplied through the injector tube 100 and out of the array of main fuel holes 106.

In use, the compressed air 34 being supplied to the combustor 70 enters the burner 74 via three distinct routes for use in the combustion process. For the purposes of describing these distinct routes the compressed air is referred to as compressed air 34A, compressed air 34B and compressed air 34C flows. In addition, some of the compressed air may be used for cooling of combustion components, but this is not shown. For example, a small portion of compressed air 34 is used for cooling the igniter 170 and/or the combustor wall. During operation of the gas turbine 10, the portion of compressed air 34A passes through the radial swirler 78 and forms a main vortex which swirls about the combustor axis 72 as it passes through the mixing tube 82 and into the combustion chamber 76. The portion of compressed air 34B passes through the array of air holes 98 in the liner 92 and into the gap 94 and then passes over the main fuel injection holes 106 and on into the mixing tube 82 where it is joined by the portion of compressed air 34A. The portion of compressed air 34C passes through the array of holes 112 in the mixing tube 82. The compressed air 34C induces forward velocity of the flow of gases (the mixture of compressed air and gaseous fuel) passing through the mixing tube 82 and particularly over and near to an inner surface 120 of the mixing tube 82. This effect is particularly apparent when the main fuel comprises a relatively high proportion of hydrogen and/or hydrocarbon. The portion of compressed air 34C ensures that the combustion flame does not contact the burner components and provides additional velocity for the gas travelling through the mixing tube 82 to help prevent the flame front travelling upstream. The portion of compressed air 34C may help to mix the main fuel 156 and compressed air 34A and 34B flows as a small or local premixer.

The mixture of main fuel and compressed air is then burned and forms a premixed zone 2 and having a main flame front 122. As can be seen by the dashed lines 122', 122”, 122'”, an increasing proportion of hydrogen in the fuel causes the flame front to move axially towards the main fuel injector 78 or generally upstream. Thus, where the flame has a flame front 122'” the fuel has a higher hydrogen content than the flame that has a flame front 122. The main fuel/air premixing zone 2 is generally radially inward of the pilot fuel mixing zone 1 albeit for the axial extent of the pilot fuel mixing zone 1.

The presently described combustor 70 is intended for mainly gas fuel burning, but liquid fuel may be used. Pilot fuel is injected from the downstream edge 130 of the mixing tube 82 as shown in Fig.3. The pilot fuel is injected through pilot fuel injection holes 132 on the circumferential edge or surface 130 of the mixing tube 82. The circumferential edge or surface 130 is annular and its centre line is coaxial with the combustor axis 72. The number of pilot fuel injection holes 132 may be between and including 6 to 18 and may be generally evenly distributed around the circumference of the surface 130. The pilot fuel injector 90 arrangement is intended to stabilise the main flame and avoid any blow off at low frequency combustion dynamics. The main gas fuel is injected from the main fuel injector 78 located along the centre line 72 of the combustor 70 and burner 74. The radial swirler 80 is truncated to accommodate the compressed air 34A flow and to diffuse the compressed air 34A, 34B flows through the conical section or first portion 108 of the burner 74. The truncated swirler 80 stabilises the flame through the radially inner and low velocity region of the swirling vortex of compressed air and main fuel. The expansion of the flow area and swirl generated through truncated radial swirler 80 and angled array of holes 112 in the first portion 108 further assists mixing of air and fuel.

The axial length D of the second portion 110 helps to keep the flame position away from the main gas injector 78 and other surfaces of the burner 74. The axial length D is 0.4L in this example, but typically may be between 0.2L and 0.8L depending on the constituents of the fuel. During a method of operating the gas turbine 10 and having the presently described new combustor 70: compressed air 34C passing through the array of holes 112 in the mixing tube 82 is approximately 20% of the total compressed air 34; compressed air 34A passing through the swirler 80 is approximately 60% of the total compressed air 34; and compressed air 34B passing through the liner 92 of the main fuel injector 78 is approximately 20% of the total compressed air 34. In other examples, compressed air 34C passing through the array of holes 112 in the mixing tube 82 may be in the range 10% to 30% of the total compressed air 34; compressed air 34A passing through the swirler 80 may be in the range 30% to 50% of the total compressed air 34; and compressed air 34B passing through the liner 92 of the main fuel injector 78 may be in the range 20% to 40%, preferably about 30%, of the total compressed air 34. As mentioned previously some of the compressed air 34 may be used for cooling of the combustion components and this portion of compressed air is not accounted for; thus, the percentages shown are based on the total amount of compressed air 34 minus any compressed air not passing through the components as combustion mixing air i.e. the total of compressed air 34A + 34B + 34C. This method of operating the gas turbine 10 includes supplying main fuel comprising at least a 5% by weight hydrogen or high-hydrocarbon content. More specifically, the present combustor is designed such that it is capable of burning main fuel having a hydrogen content greater than 20% and/or a high-HC content with a fuel heating value greater than 45MJ/m ³ . Typical high-HC fuels suitable for burning in this combustor comprise a high percentage of propane (5-50% of the total amount of main fuel) and/or butane (50-100% of the total amount of main fuel) and/or the remaining percentage of fuel may be methane. The composition of the pilot fuel may be the same as the main fuel but does not need to be the same.

Reference is now made to Fig. 4 and Fig. 5 which illustrate the pilot fuel injector 90. Fig. 4 is a schematic and enlarged view of part of the pilot fuel injector 90 as can be seen in Fig.3. Fig. 5 is a schematic view looking axially along arrow A in Fig.4. At the outlet end 126 of the mixing tube 82 is a rim 128 that extends radially outwardly from the mixing tube 82. The rim 128 has a surface 130 that faces the combustion chamber; as shown the surface 130 is on a plane that is perpendicular to the combustor axis 72. A pilot fuel gallery 124 extends around the outlet end 126 of the mixing tube 82 and on an opposite side of the rim 128 to the surface 130. A pilot feed pipe 134 extends through a wall of the combustion chamber 76 to the pilot fuel gallery 124. An array of pilot fuel injector ports 132 extend from the pilot fuel gallery 124 to the surface 130. There are 8 pilot fuel injector ports 132 equally spaced around the rim 128 although in other embodiments 4 or more pilot fuel injector ports 132 can be used and particularly 8 to 12 pilot fuel injector ports 132 are most likely. As shown, the pilot fuel injector ports 132 are aligned with the combustor axis 72 and normal to the surface 130 such that the pilot fuel is initially directed in the axial direction. However, the pilot fuel injector ports 132 may be inclined towards or away from the combustor axis 72 in other embodiments. Similarly, the surface 130 may be inclined towards or away from the combustor axis 72 in other embodiments. Furthermore, the pilot fuel injector ports 132 may be inclined in a tangential direction either in the same direction as the main swirling gas flow or in the opposite direction as the main swirling gas flow passing through and out of the mixing tube 82. The pilot fuel injectors ports 132 are located radially outwardly of the mixing tube 82 and inject fuel directly into the main combustion chamber 76 and downstream of the mixing tube 82. In other words, the pilot fuel injectors ports 132 do not inject pilot fuel into the mixing tube 82.

Reference is now made to Fig. 6 and Fig. 7 which illustrate the first portion or divergent portion 108 of the mixing tube 82 and the array of holes 112 in more detail. Fig. 7 is a section B-B shown in Fig. 6. In this embodiment, the array of holes 112 is located in the divergent portion 108. In other embodiments, the array of holes 112 may be located at least partly in the second portion 110. Any specific arrangement of the mixing tube 82 and the array of holes 112 are at least partly governed by the flame speed properties of the fuel and its content of hydrogen or high-HC fuel. Undesirable flashback from the combustion chamber into the mixing tube 82 is prevented by providing an airflow into the mixing tube 82 in a significant quantity and at a significant speed. This is in conjunction with the divergent portion 108, such that in the direction of increasing cross-sectional area of the mixing tube 82 (i.e. axially and towards the combustion chamber 76), each successive row of holes in the array of holes 112 adds to the total amount of airflow and maintains its velocity. The desired velocity and quantity of airflow prevents potential flashback of the flame front. It should be noted that the object of the airflow through the array of holes 112 is to create a barrier to the flame front and prevent flashback, it is not the object to mix air with the fuel for combustion purposes within the mixing tube 82. However, the air 34C passing through the array of holes 112 will inevitably mix with the fuel. It should be noted that the array of holes 112 are not for the main purpose of 'mixing holes' or 'mixer holes' that are usually substantially larger holes and substantially fewer holes such that the mixing air jet injects the air into, rather than surrounds, the combustion flame and/or air-fuel mixture. Details of the array of holes 112 and airflow are now described as too is the mixing tube 82.

The divergent portion 108 is generally symmetrical about the combustor axis 72. The wall of the divergent portion 108 has a constant divergent angle Q relative to the combustor axis 72. The divergent angle Q is approximately 20° but may be between 10° and 25°. Each hole of the array of holes 112 has a centreline 136 and the centreline 136 has an angle a to the combustor axis 72, where a is approximately 70° and may be between 60° and 80°. The centreline 136 of the holes 112 also has a swirl or tangential angle b relative to a radial line 138 from the combustor axis 72, the tangential angle b is approximately 10° and may be between 0° and 20°. The swirl angle creates a swirling flow as shown by arrow 140.

The array of holes (112) in the mixing tube (82) comprises a series of rows of holes (112A, 112B, 112C, 112D), each row of holes (112A, 112B, 112C, 112D) extends around the first portion (108) of the mixing tube (82).

The array of holes 112 is a series of rows of holes 112A, 112B, 112C, 112D and there may be 10 to 15 rows of holes 112A, 112B, 112C, 112D and each row of holes 112A, 112B, 112C, 112D comprises between and including 10 to 25 holes, preferably 15 to 20 holes. In other embodiments, and dependent of the size of the divergent portion 108 there may be more or less rows of holes and more or less number of holes in each row 112A, 112B, 112C, 112D. Furthermore, the series of rows of holes comprises the holes of one row, e.g. 112A, circumferentially offset from the holes of the next row, e.g. 112B, in the axial direction.

Any one or more of the holes of the array of holes 112 has a cross-sectional area between and including 0.7mm ² and 20.0mm ² . The diameter of the hole(s) is between and includes lmm and 5mm and preferably between 3mm and 5mm. In this exemplary embodiment all the holes have the same cross-sectional area; however, the cross-sectional area of the holes may increase in downstream direction to increase the flow of compressed air 34C as the surface area of the mixing tube increases. Thus at least one first hole 133 of the array of holes has a smaller cross-sectional area than at least one second hole 135, the at least one first hole 133 (e.g. in row 112A) is nearer the inlet 84 than the at least one second hole 135 (e.g. in row 112B). Instead of increasing cross-sectional area or as well as, the number of holes per row may increase in the downstream direction, again this is to increase the amount of compressed air 34C entering the mixing tube 82 as its diameter increases in the downstream direction. Thus, there may be more holes in row 112A than is row 112B and so on along the axial direction of the mixing tube 82.

The diameter of holes on the mixing tube 82 is governed by the flame speed properties of the fuel and specifically the content of Hydrogen gas and/or the high-HC fuel. It has been found that the total number of holes in the array of holes 112 is between 200 to 250 over an axial distance of 40 to 50mm. The density of holes in the array of holes 112 is between 1 hole/lOOmm ² and 6 holes/lOOmm ² . In other examples, where the mixing tube 82 is longer the axial distance may be greater than 50mm. It has been found that the array of holes 112 can be arranged in 10 rows in the axial direction and possibly up to 25 rows along the axial direction. This being based on a diameter of the mixing tube 82 being 50mm at its inlet and 70mm at its outlet. It has been found that the design of the array of holes and mixing tube 82 is intended to provide a velocity of air from each hole in the region of 70-100m/s.

Fig. 8 shows a conventional swirler vane 46, one of the annular array of vanes 46 which are tangentially angled relative to the combustor axis 44 to impart a swirling flow 55 of mixing air and fuel as is well known and shown in Fig. 2. This swirler vane has a trailing edge 142 that is parallel to the combustor axis 72. Referring to Fig. 9, a presently disclosed swirler vane 88 is shown. A trailing edge 144 of the swirler vane 88 is truncated such that the trailing edge 144 has a divergent angle f and here, the divergent angle f maintains the same divergent angle f as the divergent portion 108 of the mixing tube 82. The trailing edge 144 is effectively a generally triangular shaped surface which ends in an apex 172 at its axially downstream end. The generally triangular shaped surface is also arcuate about the combustor axis 72 to form effectively a divergent surface or edge 144 that has the divergent angle f and here, the divergent angle f is such that the surface 144 is continuous with the divergent portion 108 and preferably has the same divergent angle Q. The annular array of vanes 88 defines an outlet plane 174 on which the surface 144 is part of. The outlet plane 174 comprises at least a portion that is divergent in the direction from the main fuel injector 78 towards the combustion chamber 76. The outlet plane 174 is generally annular about the combustor axis 72 and has the divergent angle f from the combustor axis, preferably the divergent angle f is between and includes 10° and 25°. Preferably, the divergent angle f equals the divergent angle Q of the first portion of mixing tube. This ensures a smooth and continuous surface over which the compressed air and compressed air / fuel mixture can flow.

As defined above, the present radial swirler 880 is arranged to have a geometric SN between and including 0.5 to 0.8 and in this example 0.6. The truncated swirler vanes 88 provide stabilisation of the main flame by virtue of the centre region of the fuel / air vortex having a lower angular and axial velocity than the fuel / air vortex further radially outward. The expansion of the flow area of the swirler 80 and the swirl generated through truncated radial swirler 80 along with the tangentially angled array of holes 112 on the conical or divergent portion 108 enhances mixing of the fuel and air. The number of swirler vanes 88 and slots defined therebetween is 12, but the number of swirler vanes 88 and slots may be any number in the range 8 to 16.

Reference is now made to Figures 10-16 which show details of the main fuel injector 78 and various alternative embodiments thereof. In Fig. 10 the fuel outlet end 104 has a cylindrical portion 146 and the array of fuel holes 106 is distributed around the injector tube 100 within the cylindrical portion 146. In Fig. 11 the fuel outlet end 104 has a divergent portion 148 and the array of fuel holes 106 is distributed around the injector tube 100 within the divergent portion 148. In Fig. 12 the fuel outlet end 104 has a convergent portion 150 and the array of fuel holes 106 is distributed around the injector tube 100 within the convergent portion 150.

The divergent portion 148 has a divergent angle m which is from the combustor axis 72 to the outer surface 96 and in this example is 15°, but in other embodiments the divergent angle may be equal to the divergent angle Q of the swirler vanes 88 and/or the first portion 108 of the mixing tube 82. Thus, the divergent angle m may be up to and including 25°. In this embodiment, the divergent angle m is numerically less than the divergent angle Q of the swirler vanes 88 and/or the first portion 108 of the mixing tube 82.

The convergent portion 150 has a convergent angle d which is from the combustor axis 72 to the outer surface 96 and in this example is 10°, but in other embodiments the convergent angle d may be up to and including 25°.

For all of the embodiments of the main fuel injector 78, the fuel holes 107 of the array of fuel holes 106 each have a centre line 152, and the fuel holes have a centre line 152 at an angle l to the respective outer surface 96. The angle is in the axial sense, i.e. angled along the combustor axis 72. In these embodiments, all the fuel holes 107 of the array of fuel holes 106 have their centre line 152 perpendicular, i.e. l = 90°, to the outer surface 96 in the axial sense. However, in each of the embodiments herein, at least one of the fuel holes 107 of the array of fuel holes 106 has its centre line 152 angled l in the range 90° to 60° relative to the outer surface 96 in the axial sense. Furthermore, the array of holes 106 may comprise holes that are perpendicular to the outer surface 96 and holes that are angled l up to 60° in the axial direction and relative to the outer surface 96.

Holes 107 that are angled l from perpendicular to the outer surface 96 and with an angular component in the axial direction may be angled in the downstream direction or upstream direction. Some holes 107 in the array of hole 106 may be angled l in the downstream direction and some hole 107 may be angled in the upstream direction.

Referring to Fig. 13, which is a schematic cross-section through the injector tube 100 of Fig.10 looking downstream, the holes 107 of the array of fuel holes 106 have their centre lines 152 radially aligned, that is their centre line 152 passes through the axis 72. It should be appreciated that as stated above, these holes may be angled with an axial component either upstream or downstream, in other words into or out of the plane of the paper.

Referring to Fig. 14, which is a schematic cross-section through the injector tube 100 of Fig.10 looking downstream and showing an alternative embodiment to Fig.13, the holes 107 of the array of fuel holes 106 have their centre lines 152 having a tangential angle p relative to a radial line from the combustor axis 72; thus, their centre line 152 does not pass through the axis 72. The tangential angle p is 30°, and may be up to 45 °. As shown in Fig.14, the holes 107 are arranged to produce jets of fuel 156 in a generally clockwise direction, but it is possible for the holes 107 to be oppositely inclined to the radial line to produce jets of fuel 156 in a generally anticlockwise direction about the combustor axis 72. It should be appreciated that as stated above, these holes may be angled with an axial component either upstream or downstream, in other words into or out of the plane of the paper.

Reference is now made to Fig.15 which shows an alternative arrangement of the main fuel injector 78. To further enhance mixing of the main fuel 156 and compressed air 34B, the array of fuel holes 106 comprises three rows of fuel holes 106A, 106B and 106C where a first row of fuel holes 106A is staggered relative to the next and second row of fuel holes 106B. Further, the third row of holes 106C may also be staggered relative to the second row of fuel holes 106B. Other embodiments may have two, four or more rows of holes in a similarly staggered arrangement. The staggering of immediately adjacent rows of holes may be a pure geometric staggering as seen in Fig.15 or an 'aerodynamic' staggering which is related to any swirl angle of the compressed air 34B flow direction. In other words, the stagger or off-set of the holes 107 in one row 106A to the holes 107 in the next row 106B is viewed in the direction of the compressed air 34B flow.

Furthermore, Fig. 13 shows an alternative arrangement of the outlet end 148 of the main fuel injector 78 and specifically the axial location of (three rows of in this example) fuel injection holes 106 relative to the exposed surface 158 of the base plate 114 of the radial swirler 80. Arrow 154 indicates the general downstream and axial direction in the combustor. The exposed surface 158 has an extended plane 160 shown in dashed lines.

In this embodiment, some of the holes 107 of the array of fuel injection holes 106 are located axially upstream of the exposed surface 158 of the base plate 114 of the radial swirler 80. Some of the fuel injection holes 106 are located axially downstream of the plane 160 of the exposed surface of the base plate 114. In the other embodiments shown and described herein, some or all the fuel injection holes 107 of the array of fuel injection holes 106 may be located either upstream or downstream or a combination of upstream and downstream of the exposed surface of the base plate 114. In all the embodiments herein, the main fuel injection holes 106 are located axially upstream of the top plate 116 of the radial swirler 80. The axial location of the array of holes 106 is dependent on the relative amount of compressed air 34B passing through the liner 92. Where a higher percentage of compressed air 34B is used the array of holes 106 are further upstream (with respect to arrow 156 in Fig.15) and the injector tube 100 extends further

downstream. Where a lower percentage of compressed air 34B is used the array of holes 106 are further downstream (with respect to arrow 156 in Fig.15).

One object of the portion of compressed air 34B that passes through the array of air holes 98 in the liner 92 is to straighten the compressed air 34B in the gap 94 such that the air 34B is travelling substantially or completely in the axial direction. In other words, the compressed air 34B flowing in the gap 94 does not have any swirl or a minimal swirl. The holes in the array of air holes 98 are orientated such that their centre lines (162) are in a radial direction only with no swirl or tangential component. In the embodiments shown these holes in the array of air holes 98 do not have an axial angle, but it is possible in other embodiment for (at least one of) the air holes to have their centreline 162 at an axial angle w, i.e. the angle to an axial line that is parallel to the axis 72 as shown in Fig.

15. As mentioned, the axial angle w is preferably 90, but may be between 90° and 30°. Another object of array of air holes 98 in the liner 92 is to provide a turbulent air flow to aid mixing with the main fuel flow 156. In another embodiment, where the compressed air 34B in the gaps 94 has a swirl component, at least one and preferably all of the air holes of the array of air holes (98) has its centre line at a tangential angle component relative to the combustor axis (72) and in a counter-rotating direction to the compressed air 34B such that the total resultant airflow is axial or near axial. The tangential angle (e.g. as shown similarly by angle p in Fig. 14) may be up to and including 60°.

This new combustor 70 improves the aerodynamics for mixing of fuel and air in and issuing from the from the burner 74 and holds the main flame away from the

components' surfaces whilst allowing the use of fuels with higher flame speeds. To optimise any design of this new combustor 70, any one or more of the following parameters may be adjusted in particular:

• the number, size and spacing of pilot fuel injection holes 132 in each pilot burner 90,

• the number, size and spacing of the main gas injection holes 106,

• the angles of the (centre line 152) of the main gas injection holes 106,

• the flame propensity distance L for different fuels and therefore the axial length of at least the second portion 110 of the mixing tube 82, • the diameter Dp of pilot injector 90 (see Fig. 5) may be selected from a value between 0.4 to 0.6 x Dc, where Dc is the diameter of the combustion chamber 76 (see Fig.3),

• the any one or more of the divergent angle Q, angle f, angle a to the

combustor axis 72 of the holes 112 and swirl or tangential angle b of the centreline 136 of the holes 112 shown and described with reference to Fig.6 may be adjusted.

All the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least some of such features and/or steps are mutually exclusive.

Each feature disclosed in this specification (including any accompanying claims, abstract and drawings) may be replaced by alternative features serving the same, equivalent or similar purpose, unless expressly stated otherwise. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features.

The invention is not restricted to the details of the foregoing embodiment(s). The invention extends to any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed.