A gas turbine engine combustion chamber

Abstract: A gas turbine combustion chamber (28) which has primary, secondary and tertiary combustion zones (36,40,44) in flow series has a secondary mixing duct (70) and a tertiary mixing duct (82). The secondary and tertiary mixing ducts (70,82) reduce in cross-sectional area from their intakes (78,88) to their outlet apertures (80,90) to provide an accelerating flow through the mixing ducts (70,82) to prevent the formation of recirculating zones. Fuel injectors (96,112) have fuel discharge apertures (106,120) downstream of any recirculating zones formed at the intakes (78,88). The fuel injectors (96,112) extend across a major portion of the width of the ducts (70,82) to effectively subdivide the ducts (70,82) over at least part of the streamwise length of the ducts (70,82). The portions (107,117) of the fuel injectors (96,112) within the ducts (70,82) are race track shaped in cross-section and the portions (105,115) outside the ducts are aerofoil shaped in cross-section. The fuel injectors (96,112) reduce in dimension perpendicular to the widthwise direction of the duct (70,82).

Description: The present invention relates to a gas turbine engine combustion chamber. In order to meet the emission level requirements, for industrial low emission gas turbine engines, staged combustion is required in order to minimise the quantity of the oxides of nitrogen (NOx) produced. Currently the emission level requirement is for less than 25 volumetric parts per million of NOx for an industrial gas turbine exhaust. The fundamental way to reduce emissions of nitrogen oxides is to reduce the combustion reaction temperature and this requires premixing of the fuel and all the combustion air before combustion takes place. The oxides of nitrogen (NOx) are commonly reduced by a method which uses two stages of fuel injection. Our UK patent no 1489339 discloses two stages of fuel injection to reduce NOx. Our International patent application no WO92/07221 discloses two and three stages of fuel injection. In staged combustion, all the stages of combustion seek to provide lean combustion and hence the low combustion temperatures required to minimise NOx. The term lean combustion means combustion of fuel in air where the fuel to air ratio is low, ie less than the stoichiometric ratio. The present invention is particularly concerned with gas turbine engines which have staged combustion, and more particularly concerned with the secondary fuel and air mixing duct and secondary fuel injection or tertiary fuel and air mixing duct and tertiary fuel injection. In order to inject fuel into the secondary, or tertiary fuel and air mixing ducts, it is known to use cylindrical fuel injectors which extend across the inlet to the mixing duct as described in our copending UK patent application 9310690.4 filed 24 May 1993. This arrangement has suffered from preburning of fuel in the air in the mixing duct whereas the fuel should not burn until it is in the appropriate combustion zone. The fuel burns in the air in the mixing duct because of recirculation of the fuel and air in regions immediately downstream of the fuel injectors and due to hot gases in the combustion zone flowing upstream into the mixing duct. EP0169431A discloses a gas turbine engine combustion chamber comprising a combustion zone and an annular fuel and air mixing duct for conducting a mixture of fuel and air to the combustion zone. A plurality of fuel injectors are provided for injecting fuel into the annular mixing duct. Each fuel injector extends through the upstream end of the annular mixing duct and a portion of each fuel injector is positioned outside the mixing duct. Each fuel injector extends in a downstream direction along the annular mixing duct. Each fuel injector is circular in cross-section and has a plurality of discharge apertures positioned to inject fuel into the annular mixing duct. The discharge apertures inject fuel transversely of the streamwise direction and radially towards the walls of the annular mixing duct. The present invention seeks to provide a combustion chamber which reduces or overcomes these problems. Accordingly the present invention provides a gas turbine combustion chamber comprising at least one combustion zone defined by at least one peripheral wall, means to define at least one fuel and air mixing duct for conducting a mixture of fuel and air to the at least one combustion zone, each mixing duct having an upstream end for receiving air, an intermediate region for receiving fuel and a downstream end for delivering a fuel and air mixture into the at least one combustion zone, each mixing duct reducing in cross-sectional area from its upstream end to its downstream end to produce an accelerating flow therethrough, a plurality of fuel injectors for injecting fuel into the intermediate region of the at least one mixing duct, each fuel injector extends through the upstream end of the mixing duct, a portion of each fuel injector is positioned outside the mixing duct, each fuel injector extending in a downstream direction along the at least one mixing duct to the intermediate region, each fuel injector being effective to subdivide the at least one mixing duct into a plurality of ducts over at least a part of the streamwise length of the at least one mixing duct, each fuel injector having a plurality of discharge apertures positioned to inject fuel into the intermediate region of the at least one mixing duct, said discharge apertures injecting fuel transversely of the streamwise direction, said fuel injector being directed towards adjacent fuel injectors. Each fuel injector may extend the full length of the at least one mixing duct, to subdivide the at least one mixing duct into a plurality of ducts over the full streamwise length of the at least one mixing duct. A plurality of walls may extend in a downstream direction along the at least one mixing duct, each wall being effective to subdivide the at least one mixing duct into a plurality of ducts over at least a part of the streamwise length of the at least one mixing duct. Each fuel injector may extend over an upstream portion of the mixing duct, each wall extends over a downstream portion of the mixing duct, the downstream end of each fuel injector being positioned substantially immediately upstream of the upstream end of the corresponding wall such that the fuel injector and the wall cooperate to subdivide the at least one mixing duct into a plurality of ducts over the full streamwise length of the at least one mixing duct. The at least one fuel injector may extend over an upstream portion of the mixing duct, the fuel injector reducing in cross-sectional area from its upstream end to its downstream end. The downstream end of the fuel injector preferably has a relatively sharp edge. Preferably the portion of the fuel injector positioned within the mixing duct has a race track cross-section. Preferably the portion of the fuel injector outside the mixing duct has an aerofoil cross-section. Preferably the fuel injector extends in a first direction transversely relative to the streamwise direction across a major portion of the at least one mixing duct. Preferably the fuel injector has at least a portion of substantially constant dimension in the first direction, the portion is arranged between the upstream end and the intermediate region of the mixing duct. Preferably the portion of the fuel injector positioned outside the mixing duct reduces in cross-sectional area towards the portion of the fuel injector positioned within the mixing duct. Preferably the fuel injector reduces in dimension in a second direction transversely relative to the streamwise direction, between the upstream end and the intermediate region of the mixing duct, the second direction is perpendicular to the first direction. Preferably there is a uniform reduction in dimension in the second direction. The combustion chamber may have a primary combustion zone and a secondary combustion zone downstream of the primary combustion zone, the at least one fuel and air mixing duct delivers the fuel and air mixture into the secondary combustion zone. The peripheral wall may be annular, the at least one fuel and air mixing duct is arranged around the primary combustion zone. The combustion chamber may have a primary combustion zone, a secondary combustion zone downstream of the primary combustion zone and a tertiary combustion zone downstream of the secondary combustion zone, the at least one fuel and air mixing duct delivers the fuel and air mixture into the tertiary combustion zone. The peripheral wall may be annular, the at least one fuel and air mixing duct is arranged around the secondary combustion zone. The at least one fuel and air mixing duct may be defined at its radially inner extremity and radially outer extremity by a pair of annular walls. Preferably a plurality of equi-circumferentially spaced fuel injectors are provided. Preferably the combustion chamber is surrounded by a combustion chamber casing, a fuel manifold to supply fuel to the at least one fuel injector. The present invention also provides a gas turbine engine fuel injector comprising a member reducing in cross-sectional area in the longitudinal direction from a first end to a second end, the member reducing in dimension in a first direction perpendicular to the longitudinal direction from the first end to the second end, the member having a passage extending longitudinally therethrough for the supply of fuel from the first end towards the second end, the member having a plurality of discharge apertures at a predetermined distance from the second end, the discharge apertures being spaced apart in a second direction which is substantially perpendicular to both the first direction and the longitudinal direction, the apertures being arranged to direct fuel substantially perpendicularly to the second direction. There may be a uniform reduction in dimension in the first direction. Preferably at least a portion of the member has a substantially constant dimension in the second direction. Preferably the at least a portion of the member is adjacent the second end of the member. Preferably a portion of the fuel injector reduces in dimension in the second direction between the first end of the member and the portion of the member having a constant dimension in the second direction. Preferably the portion of the member which has a substantially constant dimension in the first direction has a race track cross-section. Preferably the portion of the member which reduces in dimension in the second direction has an aerofoil cross-section. Preferably the second end of the member has a sharp edge. The present invention will be more fully described by way of of example with reference to the accompanying drawings, in which:- Figure 1 is a view of a gas turbine engine having a combustion chamber assembly according to the present invention. Figure 2 is an enlarged longitudinal cross-sectional view through the combustion chamber shown in figure 1. Figure 3 is a cross-sectional view in the direction of arrows A-A in figure 2. Figure 4 is a cross-sectional view in the direction of arrows B-B in figure 2. Figure 5 is an enlarged partial view in the direction of arrow C in figure 2 showing a single fuel injector. Figure 6 is a cross-sectional view in the direction of arrows D-D in figure 5. Figure 7 is a cross-sectional view in the direction of arrows E-E in figure 5. Figure 8 is a cross-sectional view in the direction of arrows F-F in figure 5. Figure 9 is a cross-sectional view in the direction of arrows G-G in figure 5. Figure 10 is a cross-sectional view in the direction of arrows H-H in figure 5. An industrial gas turbine engine 10, shown in figure 1, comprises in axial flow series an inlet 12, a compressor section 14, a combustion chamber assembly 16, a turbine section 18, a power turbine section 20 and an exhaust 22. The turbine section 18 is arranged to drive the compressor section 14 via one or more shafts (not shown). The power turbine section 20 is arranged to drive an electrical generator 26 via a shaft 24. However, the power turbine section 20 may be arranged to provide drive for other purposes. The operation of the gas turbine engine 10 is quite conventional, and will not be discussed further. The combustion chamber assembly 16 is shown more clearly in figures 2 to 5. The combustion chamber assembly 16 comprises a plurality of, for example nine, equally circumferentially spaced tubular combustion chambers 28. The axes of the tubular combustion chamber 28 are arranged to extend in generally radial directions. The inlets of the tubular combustion chambers 28 are at their radially outermost ends and their outlets are at their radially innermost ends. Each of the tubular combustion chambers 28 comprises an upstream wall 30 secured to the upstream end of an annular wall 32. A first, upstream, portion 34 of the annular wall 32 defines a primary combustion zone 36, a second, intermediate portion 38 of the annular wall 32 defines a secondary combustion zone 40 and a third downstream portion 42 of the annular wall 32 defines a tertiary combustion zone 44. The downstream end of the first portion 34 has a frustoconical portion 46 which reduces in diameter to a throat 48. The second portion 38 of the annular wall 32 has a greater diameter than the first portion 34. A frustoconical portion 50 interconnects the throat 48 and the upstream end of the second portion 38. The downstream end of the second portion 38 has a frustoconical portion which reduces in diameter to a throat 54. The third portion 42 of the annular wall 32 has a greater diameter than the second portion 38. A frustoconical portion 56 interconnects the throat 54 and the upstream end of the third portion 42. The upstream wall 30 of each of the tubular combustion chambers 28 has an aperture 58 to allow the supply of air and fuel into the primary combustion zone 36. A first radial flow swirler 60 is arranged coaxially with the aperture 58 in the upstream wall 30 and a second radial flow swirler 62 is arranged coaxially with the aperture 58 in the upstream wall 30. The first radial flow swirler 60 is positioned axially downstream, with respect to the axis of the tubular combustion chamber, of the second radial flow swirler 62. The first radial flow swirler 60 has a plurality of fuel injectors 64, each of which is positioned in a passage formed between two vanes of the swirler. The second radial flow swirler 62 has a plurality of fuel injectors 72, each of which is positioned in a passage formed between two vanes of the swirler. The first and second radial flow swirlers 60 and 62 are arranged such that they swirl the air in opposite directions. For a more detailed description of the use of the two radial flow swirlers and the fuel injectors positioned in the passages formed between the vanes see our international patent application no W092/07221. The primary fuel and air is mixed together in the passages between the vanes of the first and second radial flow swirlers 60 and 62. An annular secondary fuel and air mixing duct 70 is provided for each of the tubular combustion chambers 28. Each secondary fuel and air mixing duct 70 is arranged coaxially around the primary combustion zone 36. Each of the secondary fuel and air mixing ducts 70 is defined between a second annular wall 72 and a third annular wall 74. The second annular wall 72 defines the radially inner extremity of the secondary fuel and air mixing duct 70 and third annular wall 74 defines the radially outer extremity of the secondary fuel and air mixing duct 70. The axially upstream end 76 of the second annular wall 72 is secured to a side plate of the first radial flow swirler 60. The axially upstream ends of the second and third annular walls 72 and 74 are substantially in the same plane perpendicular to the axis of the tubular combustion chamber 28. The secondary fuel and air mixing duct 70 has a secondary air intake 78 defined radially between the upstream end of the second annular wall 72 and the upstream end of the third annular wall 74. At the downstream end of the secondary fuel and air mixing duct 70, the second and third annular walls 72 and 74 respectively are secured to the frustoconical portion 50 and the frustoconical portion 50 is provided with a plurality of equi-circumferentially spaced apertures 80. The apertures 80 are arranged to direct the fuel and air mixture into the secondary combustion zone 40 in the tubular combustion chamber 28, in a downstream direction towards the axis of the tubular combustion chamber 28. The apertures 80 may be circular or slots and are of equal flow area. The secondary fuel and air mixing duct 70 reduces gradually in cross-sectional area from the intake 78 at its upstream end to the apertures 80 at its downstream end. The second and third annular walls 72 and 74 of the secondary fuel and air mixing duct 70 are shaped to produce an aerodynamically smooth duct 70. The shape of the secondary fuel and air mixing duct 70 therefore produces an accelerating flow through the duct 70 without any regions where recirculating flows may occur. An annular tertiary fuel and air mixing duct 82 is provided for each of the tubular combustion chambers 28. Each tertiary fuel and air mixing duct 82 is arranged coaxially around the secondary combustion zone 40. Each of the tertiary fuel and air mixing ducts 82 is defined between a fourth annular wall 84 and a fifth annular wall 86. The fourth annular wall 84 defines the radially inner extremity of the tertiary fuel and air mixing duct 82 and the fifth annular wall 86 defines the radially outer extremity of the tertiary fuel and air mixing duct 82. The axially upstream ends of the fourth and fifth annular walls 84 and 86 are substantially in the same plane perpendicular to the axis of the tubular combustion chamber 28. The tertiary fuel and air mixing duct 82 has a tertiary air intake 88 defined radially between the upstream end of the fourth annular wall 84 and the upstream end of the fifth annular wall 86. At the downstream end of the tertiary fuel and air mixing duct 82, the fourth and fifth annular walls 84 and 86 respectively are secured to the frustoconical portion 56, and the frustoconical portion 56 is provided with a plurality of equi-circumferentially spaced apertures 90. The apertures 90 are arranged to direct the fuel and air mixture into the tertiary combustion zone 44 in the tubular combustion chamber 28, in a downstream direction towards the axis of the tubular combustion chamber 28. The apertures 90 may be circular or slots and are of equal flow area. The tertiary fuel and air mixing duct 82 reduces gradually in cross-sectional area from the intake 88 at its upstream end to the apertures 90 at its downstream end. The fourth and fifth annular walls 84 and 86 of the tertiary fuel and air mixing duct 82 are shaped to produce an aerodynamically smooth duct 82. The shape of the tertiary fuel and air mixing duct 82 therefore produces an accelerating flow through the duct 82 without any regions where recirculating flows may occur. A plurality of secondary fuel systems 92 are provided, to supply fuel to the secondary fuel and air mixing ducts 70 of each of the tubular combustion chambers 28. The secondary fuel system 92 for each tubular combustion chamber 28 comprises an annular secondary fuel manifold 94 arranged coaxially with the tubular combustion chamber 28 at the upstream end of the tubular combustion chamber 28. The secondary fuel manifold is defined by the casing 124, but it may be positioned outside or inside the casing 124. Each secondary fuel manifold 94 has a plurality, for example thirty two, of equi-circumferentially spaced secondary fuel injectors 96. Each of the secondary fuel injectors 90 comprises a hollow member 98 which extends axially with respect to the tubular combustion chamber 28, from the secondary fuel manifold 94 in a downstream direction through the intake 78 of the secondary fuel and air mixing duct 70 and into the secondary fuel and air mixing duct 70. Each hollow member 98 extends in a downstream direction along the secondary fuel and air mixing duct 70 to a position, sufficiently far from the intake 78, where there are no recirculating flows in the secondary fuel and air mixing duct 70 due to the flow of air into the duct 70. Each hollow member 98 extends in a first direction, ie radially across the secondary fuel and air mixing duct 70, transversely relative to the streamwise direction, across a major portion of the secondary fuel and air mixing duct 70. Each hollow member 98 has the same dimension in the first direction at one portion 107 along its length, and radially with respect to the tubular combustion chamber 28. Each hollow member 98 has a gradual reduction in dimension in a second direction, perpendicular to the first direction and transversely relative to the streamwise direction, between a first end 100 secured to the secondary fuel manifold 94 and a second end 102 in the secondary fuel and air mixing duct 70. The hollow member 98 reduces in dimension in the first direction between the first end 100 and the portion 107. Thus each hollow member 98 reduces in cross-sectional area from its first end 100 to its second end 102. Each hollow member 98 has a passage 104 which extends longitudinally from the first end 100 of the hollow member 98 at the secondary fuel manifold 94 towards but to a position spaced from the second end 102 of the hollow member 98. The second end 102 of each hollow member 98 has a plurality of discharge apertures 106. The apertures 106 are spaced apart in the first direction and are arranged to direct fuel perpendicularly to the first direction, ie in the second direction. There are apertures 106 provided to discharge fuel from both sides of the hollow member 98 in the second direction, but in opposite directions. The passage 104 interconnects with the discharge apertures 106 to supply fuel from the secondary fuel manifold 94 to the discharge apertures 106. It can be seen that the discharge apertures 106 on each hollow member 98 are thus spaced apart radially with respect to the secondary fuel and air mixing duct 70 and that they discharge fuel generally in circumferential directions. Thus each fuel injector 96 discharges fuel towards the adjacent fuel injectors 96. The hollow members 98 of the fuel injectors 96 extend across a major portion of the secondary fuel and air mixing ducts 70 such that they effectively aerodynamically divide the duct 70 into a number of separate mixing ducts. The fuel injectors 96 thus divide the secondary fuel and air mixing duct 70 into separate mixing ducts as well as serving to supply fuel into the separate mixing ducts. There is negligible mass flow between the radially inner and outer ends of the hollow member 98 and the annular walls 72 and 74 defining the secondary fuel and air mixing duct 70. The fuel injectors 96 extend only part of the length of the secondary fuel and air mixing duct 70. The hollow members 98 are aerofoil shaped in cross-section over the region 105, as shown in figures 6 and 7, but the hollow members 98 blend, as shown in figure 8, to a race track shape cross-section in region 107, as shown in figures 9 and 10. The hollow members 98 are aerofoil shaped at region 105 to allow a smooth aerodynamic flow of air transversely of the hollow members 98, within the casing 124, without disturbance to the first and second radial flow swirlers 60 and 62. The hollow members 98 are race track shaped at region 107 to provide a smooth aerodynamic flow of air lengthwise of the hollow members 98 into the secondary fuel and air mixing duct 70. The second end 102 of the hollow members 98 is a very thin edge so that substantially no, or very little, turbulence is generated by the air flow passing through the secondary fuel and air mixing duct 70 along the hollow members 98 as it leaves the second end 102. A plurality of tertiary fuel systems 108 are provided, to supply fuel to the tertiary fuel and air mixing ducts 82 of each of the tubular combustion chambers 28. The tertiary fuel system 108 for each tubular combustion chamber 28 comprises an annular tertiary fuel manifold 110 arranged coaxially with the tubular combustion chamber 28. The tertiary fuel manifold 110 is positioned outside the casing 124, but may be positioned in the casing 124. Each tertiary fuel manifold 110 has a plurality, for example thirty two, of equi-circumferentially spaced tertiary fuel injectors 112. Each of the tertiary fuel injectors 112 comprises a hollow member 114 which extends initially radially inwardly and then axially with respect to the tubular combustion chamber 28 from the tertiary fuel manifold 110 in a downstream direction through the intake 88 of the tertiary fuel and air mixing duct 82 and into the tertiary fuel and air mixing duct 82. Each hollow member 114 extends in a downstream direction along the tertiary fuel and air mixing duct 82 to a position, sufficiently far from the intake 88, where there are no recirculating flows in the tertiary fuel and air mixing duct 82 due to the flow of air into the duct 82. Each hollow member 114 extends in a first direction, ie radially across the tertiary fuel and air mixing duct 82, transversely relative to the streamwise direction, across a major portion of the tertiary fuel and air mixing duct 82. Each hollow member 114 has the same dimension in the first direction at all positions along its length which are within the tertiary fuel and air mixing duct 82. Each hollow member 114 has a gradual reduction in dimension in a second direction, perpendicular to the first direction and transversely relative to the streamwise direction, between a first end 116 and secured to the tertiary fuel manifold 110 and a second end 118 in the tertiary fuel and air mixing duct 82. Thus each hollow member 114 reduces in cross-sectional area from its first end 116 to its second end 118. Each hollow member 114 has a passage 120 which extends longitudinally from the first end 116 of the hollow member 114 at the tertiary fuel manifold 110 towards but to a position spaced from the second end 118 of the hollow member 114. The second end 118 of each hollow member 114 has a plurality of discharge apertures 122. The apertures 122 are spaced apart in the first direction and are arranged to direct fuel perpendicularly to the first direction, ie in the second direction. There are apertures 122 provided to discharge fuel from both sides of the hollow member 114 in the second direction, but in opposite directions. The passage 120 interconnects with the discharge apertures 122 to supply fuel from the tertiary fuel manifold 110 to the discharge apertures 122. It can be seen that the discharge apertures 122 on each hollow member 120 are thus spaced apart radially with respect to the tertiary fuel and air mixing duct 82 and that they discharge fuel generally in circumferential directions. Similarly the hollow members 114 of the fuel injectors 112 extend across a major portion of the tertiary fuel and air mixing ducts 82 such that they effectively aerodynamically divide the duct 82 into a number of separate mixing ducts. The fuel injectors 112 thus divide the tertiary fuel and air mixing duct 82 into separate mixing ducts as well as serving to supply fuel into the separate mixing ducts. There is negligible mass flow between the radially inner and outer ends of the hollow member 114 and the annular walls 84 and 86 defining the tertiary fuel and air mixing duct 82. The fuel injectors 112 extend only part of the length of the tertiary fuel and air mixing duct 82. The hollow members 114 are aerofoil shaped in cross-section over the region 115, as shown in figure 2, but the hollow members 114 are race track shape in cross-section in region 117 as shown in figure 2. The hollow members 114 are aerofoil shaped at region 115 to allow a smooth aerodynamic flow of air transversely of the hollow members 114, within the casing 124, without disturbance to the first and second radial flow swirlers 60 and 62 and to the secondary fuel and air mixing duct 70. The hollow members 114 are race track shaped at region 117 to provide a smooth aerodynamic flow of air lengthwise of the hollow members 117 into the tertiary fuel and air mixing duct 82. The second end 118 of the hollow members 114 is a very thin edge so that substantially no, or very little, turbulence is generated by the air flow passing through the tertiary fuel and air mixing duct 82 along the hollow members 114 as it leaves the second end 118. The secondary and tertiary fuel manifolds 94 and 110 are positioned outside the combustion casing 124 which encloses the tubular combustion chamber 28. In operation there is an accelerating flow of air through the secondary and tertiary fuel and air mixing ducts 70 and 82 respectively due to the aerodynamically smooth shape of the ducts and due to the fact that the secondary and tertiary fuel and air mixing ducts 70, 82 reduce in cross-sectional area between their intakes 78, 88 at their upstream ends and the apertures 80, 90 at their downstream ends. The accelerating flow of air through the mixing ducts 70 and 82 reduces or prevents the formation of recirculating zones in the mixing ducts 70 and 82, and this in turn reduces or eliminates the possibility of burning of the fuel injected into the mixing ducts 70 and 82. The fuel injectors 96 and 112 extend from respective fuel manifolds 94 and 110 positioned outside the combustion chamber casing 124. The locating of fuel manifolds outside the combustion chamber casing 124 has the advantage that there is no possibility of fuel leaking from the fuel manifolds into the mixing ducts 70 and 82 and hence the possibility of fires in the mixing duct 70 and 82 is reduced. It is not necessary to have seals internally of the combustion chamber casing for this design, nor is it necessary to have supply pipes with expansion/contraction capability. The distances from the discharge apertures 106, 122 to the respective apertures 80, 90 is maintained as large as is possible for optimum mixing of the fuel and air while ensuring that the discharge apertures 106, 122 are sufficiently far away from the intakes 78, 88 of the mixing ducts 70, 82 such that any fuel injected from the injectors 96, 112 does not migrate into any recirculating zones at the intakes 78, 88 of the mixing ducts 70, 82. It is possible that fuel injectors at all positions around the annular mixing ducts have the same degree of tapering. However, it may be possible to vary the degree of tapering of the fuel injectors at various positions around the annular mixing ducts. The invention has described fuel injectors which extend only part of the length of the mixing duct. However, if the mixing duct is substantially straight, the fuel injectors may extend the full length of the mixing duct to fully divide the mixing duct into separate mixing ducts. In this case the fuel injectors may have constant cross-sectional area throughout the length of the mixing duct. It may be possible to subdivide the mixing duct at its downstream end with radially extending walls. For example the tertiary fuel and air mixing duct 82 has radial walls 126 indicated by the broken lines in figure 2. The downstream ends 118 of the fuel injectors 112 are positioned immediately adjacent to, or close to, the upstream ends of the walls 126 such that the fuel injectors 112 and walls 126 cooperate to completely divide the tertiary fuel and air mixing duct 82 from the intake 88 to the apertures 90. The fuel injectors may have constant cross-sectional area throughout the length of the tertiary mixing duct. The walls may be secured to both annular walls 84 and 86 or secured to only one of the walls 84,86.

Claims: A gas turbine combustion chamber (28) comprising at least one combustion zone (40) defined by at least one peripheral wall (38), means (72,74) to define at least one fuel and air mixing duct (70) for conducting a mixture of fuel and air to the at least one combustion zone (40), each mixing duct (70) having an upstream end (78) for receiving air, a region for receiving fuel and a downstream end (80) for delivering a fuel and air mixture into the at least one combustion zone (40), a plurality of fuel injectors (96) for injecting fuel into the at least one mixing duct (70), each fuel injector (96) having a plurality of discharge apertures (106) positioned to inject fuel into the mixing duct (70), said discharge apertures (106) injecting fuel transversely of the streamwise direction, each fuel injector (96) extends through the upstream end (78) of the mixing duct (70), a portion (105) of each fuel injector is positioned outside the mixing duct (70), each fuel injector (96) extends in a downstream direction along the at least one mixing duct (70) to at least an intermediate region, each fuel injector (96) injects fuel into the intermediate region of the at least one mixing duct (70), characterised in that each mixing duct (70) reduces in cross-sectional area from its upstream end (78) to its downstream end (80) to produce an accelerating flow therethrough, each fuel injector (96) is effective to subdivide the at least one mixing duct (70) into a plurality of ducts over at least a part of the streamwise length of the at least one mixing duct (70), said fuel injection being directed towards adjacent fuel injectors (96). A combustion chamber (28) as claimed in claim 1 wherein each fuel injector (96) extends the full length of the at least one mixing duct (70), to subdivide the at least one mixing duct (70) into a plurality of ducts over the full streamwise length of the at least one mixing duct (70). A combustion chamber as claimed in claim 1 wherein a plurality of walls (126) extend in a downstream direction along the at least one mixing duct (82), each wall (126) being effective to subdivide the at least one mixing duct (82) into a plurality of ducts over at least a part of the streamwise length of the at least one mixing duct (82). A combustion chamber as claimed in claim 3 wherein each fuel injector (112) extends over an upstream portion of the mixing duct (82), each wall (126) extends over a downstream portion of the mixing duct (70), the downstream end (118) of each fuel injector (112) being positioned substantially immediately upstream of the upstream end of the corresponding wall (126) such that the fuel injector (112) and the wall (126) cooperate to subdivide the at least one mixing duct (82) into a plurality of ducts over the full streamwise length of the at least one mixing duct (82). A combustion chamber as claimed in claim 1 wherein the at least one fuel injector (96) extends over an upstream portion of the mixing duct (70), the fuel injector (96) reducing in cross-sectional area from its upstream end (100) to its downstream end (102). A combustion chamber as claimed in claim 5 wherein the downstream end (102) of the fuel injector (96) has a relatively sharp edge. A combustion chamber as claimed in any of claims 1 to 6 wherein the portion (107) of the fuel injector (96) positioned within the mixing duct (70) has a race track cross-section. A combustion chamber as claimed in any of claims 1 to 7 wherein the portion (105) of the fuel injector (96) outside the mixing duct (70) has an aerofoil cross-section. A combustion chamber as claimed in claim 5 or claim 6 wherein the fuel injector (96) extends in a first direction transversely relative to the streamwise direction across a major portion of the at least one mixing duct (70). A combustion chamber as claimed in claim 9 wherein the fuel injector (96) has at least a portion (107) of substantially constant dimension in the first direction, the portion (107) is arranged between the upstream end (78) and the intermediate region of the mixing duct (70). A combustion chamber as claimed in any of claims 1 to 10 wherein the portion (105) of the fuel injector (96) positioned outside the mixing duct (70) reduces in cross-sectional area towards the portion (107) of the fuel injector (96) positioned within the mixing duct (70). A combustion chamber as claimed in claim 9 or claim 10 wherein the fuel injector (96) reduces in dimension in a second direction transversely relative to the streamwise direction, between the upstream end (78) and the intermediate region of the mixing duct (70), the second direction is perpendicular to the first direction. A combustion chamber as claimed in claim 12 wherein there is a uniform reduction in dimension in the second direction. A combustion chamber as claimed in any of claims 1 to 13 wherein the combustion chamber (28) has a primary combustion zone (36) and a secondary combustion zone (40) downstream of the primary combustion zone (36), the at least one fuel and air mixing duct (70) delivers the fuel and air mixture into the secondary combustion zone (40). A combustion chamber as claimed in claim 14 wherein the peripheral wall (38) is annular, the at least one fuel and air mixing duct (70) is arranged around the primary combustion zone (36). A combustion chamber as claimed in any of claims 1 to 15 wherein the combustion chamber (28) has a primary combustion zone (36), a secondary combustion zone (40) downstream of the primary combustion zone (36) and a tertiary combustion zone (44) downstream of the secondary combustion zone (40), the at least one fuel and air mixing duct (82) delivers the fuel and air mixture into the tertiary combustion zone (44). A combustion chamber as claimed in claim 16 wherein the peripheral wall (42) is annular, the at least one fuel and air mixing duct (82) is arranged around the secondary combustion zone (40). A combustion chamber as claimed in claim 15 or claim 17, wherein the at least one fuel and air mixing duct (70) is defined at its radially inner extremity and radially outer extremity by a pair of annular walls (72,74). A combustion chamber as claimed in claim 18 comprising a plurality of equi-circumferentially spaced fuel injectors (96). A gas turbine engine fuel injector (96) comprising a member (98), the member (98) having a passage (104) extending longitudinally therethrough for the supply of fuel from a first end (100) towards a second end (102), the member (98) having a plurality of discharge apertures (106), the member (98) reduces in cross-sectional area in the longitudinal direction from the first end (100) to the second end (102), the member (98) reducing in dimension in a first direction perpendicular to the longitudinal direction from the first end (100) to the second end (102), the discharge apertures (106) are located a predetermined distance from the second end (102), characterized in that the discharge apertures (106) are spaced apart in a second direction which is substantially perpendicular to both the first direction and the longitudinal direction, the discharge apertures (106) are arranged to direct the fuel substantially in the first direction. A fuel injector (96) as claimed in claim 20 in which there is a uniform reduction in dimension in the first direction. A fuel injector (96) as claimed in claim 20 or claim 21 in which at least a portion (107) of the member (98) has a substantially constant dimension in the second direction. A fuel injector as claimed in claim 22 in which the at least a portion (107) of the member (98) is adjacent the second end (102) of the member (98). A fuel injector as claimed in claim 23 in which a portion (105) of the member (98) reduces in dimension in the second direction between the first end (100) of the member (98) and the portion (107) of the member (98) having a constant dimension in the second direction. A fuel injector (96) as claimed in any of claims 22 to 24 in which the portion (107) of the member (98) which has a substantially constant dimension in the first direction has a race track cross-section. A fuel injector (96) as claimed in claim 24 in which the portion (105) of the member (98) which reduces in dimension in the second direction has an aerofoil cross-section. A fuel injector (96) as claimed in any of claims 20 to 26 wherein the second end (102) of the member (98) has a sharp edge.

Inventors: RICHARDSON BRIAN, RICHARDSON, BRIAN
Applicants: ROLLS ROYCE PLC, ROLLS-ROYCE PLC
Application year: 1995
Main classification: F23R3
IPC codes: F23R3, F02C7