Patent Application
20140360202
The present invention relates to a fuel injector and a combustion chamber and in particular to a lean burn fuel injector and a gas turbine engine combustion chamber.
Conventionally fuel is supplied into a gas turbine engine combustion chamber via a plurality of fuel injectors. In an annular combustion chamber each fuel injector is located in a respective one of a plurality of apertures in an upstream end of the combustion chamber.
One type of gas turbine engine combustion chamber is known as a rich burn combustion chamber and another type of gas turbine engine combustion chamber is known as a lean burn combustion chamber. In a lean burn type of combustion chamber the fuel and air is mixed such that the fuel to air equivalence ratio is less than one.
Conventionally gas turbine engine combustion chambers use rich burn technology, however rich burn gas turbine engine combustion chambers will not be able to achieve future nitrous oxide (NOx) emission requirements.
Gas turbine engine combustion chambers are being developed which use lean burn technology to reduce the emissions of nitrous oxides (NOx). Lean burn gas turbine engine combustion chambers have lean burn fuel injectors, each of which comprises a pilot fuel injector and a main fuel injector, to enable lean combustion at higher air to fuel ratios than the stoichiometric air to fuel ratio, and to provide high thrusts with low NOx and to supply fuel to the pilot fuel injector at low thrusts to achieve required combustion efficiency, lean blow out margin and altitude relight capability.
One type of fuel injector for a lean burn type of combustion chamber comprises a pilot fuel injector and a main fuel injector. The pilot fuel injector is provided between two sets of air swirlers and the main fuel injector is provided between a further two sets of air swirlers. Generally the pilot fuel injector and the main fuel injector are arranged concentrically and the main fuel injector is arranged around the pilot fuel injector. The first two sets of air swirlers provide swirling flows of air which atomise the fuel from the pilot fuel injector and the second two sets of air swirlers provide swirling flows of air which atomise the fuel from the main fuel injector. Each air swirler comprises a plurality of circumferentially spaced radially extending swirl vanes and the swirl vanes extend between concentric members. The four sets of air swirlers are arranged concentrically. This type of fuel injector is also provided with first and second air splitters between the pilot fuel injector and the main fuel injector. The first air splitter has a converging downstream portion and the second air splitter has a diverging downstream portion.
There have been problems with obtaining satisfactory ignition of the fuel in the combustion chamber or problems with unacceptably high temperatures of the second air splitter and unacceptably low combustion efficiency when the pilot fuel injector only is used.
Therefore the present invention seeks to provide a novel fuel injector for a gas turbine engine combustion chamber which reduces or overcomes the above mentioned problem.
Accordingly the present invention provides a fuel injector comprising a pilot fuel injector and a main fuel injector, the pilot fuel injector comprising at least one pilot air swirler, the main fuel injector comprising a main air blast fuel injector located between an inner main air swirler and an outer main air swirler, a first air splitter located between the at least one pilot air swirler and the inner main air swirler and a second air splitter located between the at least one pilot air swirler and the inner main air swirler, the first air splitter comprising a downstream portion converging to a downstream end, the second air splitter comprising a downstream portion diverging to a downstream end, the downstream end of the second air splitter is downstream of the downstream end of the first air splitter, the downstream end of the second air splitter being downstream of the downstream end of a member defining the outer surface of the outer main air swirler, and the ratio of the distance from the downstream end of the first air splitter to the downstream end of the second air splitter to the diameter of the downstream end of the second air splitter is in the range of 0.22 to 0.30.
The ratio of the distance from the downstream end of the first air splitter to the downstream end of the second air splitter to the diameter of the downstream end of the second air splitter may be in the range of 0.24 to 0.28.
The ratio of the distance from the downstream end of the first air splitter to the downstream end of the second air splitter to the diameter of the downstream end of the second air splitter may be in the range of 0.25 to 0.27.
The pilot fuel injector may comprise a pilot air blast fuel injector located between an inner pilot air swirler and an outer pilot air swirler.
The second air splitter may be located between the first air splitter and the inner main air swirler, an additional air swirler is provided between the first air splitter and the second air splitter to direct air over the second air splitter.
The outer main air swirler may comprise a plurality of swirl vanes arranged in an annular duct, the annular duct is defined by a radially inner surface of an outer wall and a radially outer surface of an inner wall.
The radially inner surface of the outer wall of the annular duct may converge to a minimum diameter downstream of the swirl vanes, the radial width of the annular duct at the trailing edges of the swirl vanes is in the range of 1.1 to 1.3 times the radial width of the annular duct at the minimum diameter of the radially inner surface of the outer wall of the annular duct.
The radially outer surface of the inner wall of the annular duct may converge to a minimum diameter downstream of the swirl vanes, the radial distance of convergence of the radially outer surface of the inner wall is in the range of 0.5 to 1.0 times the radial width of the annular duct at the minimum diameter of the radially inner surface of the outer wall of the annular duct.
The axial length of the annular duct from the trailing edges of the swirl vanes to the minimum diameter of the radially inner surface of the outer wall of the annular duct may be in the range of 1.7 to 2.5 times the radial width of the annular duct at the minimum diameter of the radially inner surface of the outer wall of the annular duct.
The downstream end of the second air splitter may be downstream of the downstream end of the inner wall of the annular duct.
The fuel injector may be provided in a combustion chamber. The combustion chamber may comprise an igniter and the igniter is positioned downstream of the at least one fuel injector. The combustion chamber may be a gas turbine engine combustion chamber.
The gas turbine engine may be a turbofan gas turbine engine, a turbo-jet gas turbine engine, a turbo-shaft gas turbine engine or a turbo-prop gas turbine engine. The gas turbine engine may be an aero gas turbine engine, a marine gas turbine engine, an industrial gas turbine engine or an automotive gas turbine engine.
The present invention also provides a method of operating a combustion chamber, the combustion chamber comprising an igniter and at least one fuel injector, the igniter being positioned downstream of the at least one fuel injector, the fuel injector comprising a pilot fuel injector and a main fuel injector, the pilot fuel injector comprising at least one pilot air swirler, the main fuel injector comprising a main air blast fuel injector located between an inner main air swirler and an outer main air swirler, a first air splitter located between the at least one pilot air swirler and the inner main air swirler and a second air splitter located between the at least one pilot air swirler and the inner main air swirler, the first air splitter comprising a downstream portion converging to a downstream end, the second air splitter comprising a downstream portion diverging to a downstream end, the downstream end of the second air splitter is downstream of the downstream end of the first air splitter, the downstream end of the second air splitter is downstream of the downstream end of a member defining the outer surface of the outer main air swirler and the ratio of the distance from the downstream end of the first air splitter to the downstream end of the second air splitter to the diameter of the downstream end of the second air splitter is in the range of 0.22 to 0.30, the method comprising supplying pilot fuel to the pilot fuel injector and supplying main fuel to the main fuel injector, atomising the pilot fuel using a swirling flow of air from the at least one pilot air swirler, atomising the main fuel using swirling flows of air from the inner main air swirler and the outer main air swirler, producing an S shaped flow path for the pilot fuel supplied from the pilot fuel injector to the main fuel supplied by the main fuel injector, and mixing the pilot fuel with the main fuel and air flow upstream of the igniter.
The present invention will be more fully described by way of example with reference to the accompanying drawings, in which:—
A turbofan gas turbine engine 10, as shown in
The combustion chamber 15, as shown more clearly in
Each fuel injector 56 comprises a fuel feed arm 58 and a fuel injector head 60, as shown in
A fuel injector 56 according to the present invention is shown more clearly in
The fuel injector head 60 also comprises a fourth generally annular member 88 spaced coaxially around the third member 70, a fifth member 90 spaced coaxially around the fourth member 88 and a sixth member 92 spaced coaxially around the fifth member 90. A plurality of circumferentially spaced swirl vanes 94 extend radially between the fourth member 88 and the fifth member 90 to form a third air swirler 93. The fifth member 90 has a greater axial length than the fourth member 88 and the fourth member 88 is positioned at the downstream end 90B of the fifth member 90 and a generally annular duct 96 is defined between the fourth member 88 and the fifth member 90 and the swirl vanes 94 extend across the annular duct 96. The fifth member 90 has one or more internal fuel passages 98 which are arranged to receive fuel from the second internal fuel passage 64 in the fuel feed arm 58. The one or more fuel passages 98 are arranged to supply fuel to a fuel swirler 100 which supplies a film of fuel onto the radially inner surface 102 at the downstream end 90B of the fifth member 90. A plurality of circumferentially spaced swirl vanes 104 extend radially between the fifth member 90 and the sixth member 92 to form a fourth air swirler 103. A generally annular duct 106 is defined between the downstream end 90B of the fifth member 90 and the sixth member 92 and the swirl vanes 104 extend across the annular duct 106. The downstream end 88B of the fourth member 88 is conical and is divergent in a downstream direction. In operation the main fuel supplied by the internal fuel passages 98 and fuel swirler 100 onto the radially inner surface 102 of the fifth member 90 is atomised by swirling flows of air C and D from the swirl vanes 94 and 104 of the third and fourth air swirlers 93 and 103 respectively. The fourth member 88, the fifth member 90, the sixth member 92, the third swirler 93 and the fourth swirler 103 form the main fuel injector 61 and in this case the main fuel injector 61 is a main air blast fuel injector 61. The third and fourth air swirlers 93 and 103 are the inner and outer main air swirlers respectively for the main air blast fuel injector 61.
The third member 70 and the fourth member 88 are first and second splitters respectively and are used to separate, or split, the flows of fuel and air from the pilot fuel injector 59 from the fuel and air from the main fuel injector 61.
The fuel injector head 60 also comprises a plurality of circumferentially spaced swirl vanes 108 which extend radially between the third member 70 and the fourth member 88 to form a fifth air swirler 107. In operation the swirl vanes 108 of the fifth air swirler 107 provide a swirling flow of air E over the radially inner surface of the fourth member 88.
The sixth member 92 has a radially inner surface 110, the radially inner surface 110 of the sixth member 92 is generally circular in cross-section in a plane perpendicular to the axis Y of the fuel injector head 60 of the fuel injector 56. The radially inner surface 110 of the downstream end 92B of the sixth member 92 converges to a minimum diameter 114 at a plane arranged perpendicular to the axis Y of the fuel injector head 60 containing the downstream end 90B of the fifth member 90 and then the radially inner surface 110 of the downstream end 92B of the sixth member 92 diverges downstream of the downstream end 90B of the fifth member 90. In particular the radially inner surface 110 of the downstream end 92B of the sixth member 92 downstream of the swirl vanes 104 converges to the minimum diameter 114.
The fifth member 90 has a radially outer surface 112, the radially outer surface 112 of the fifth member 90 is generally circular in cross-section in a plane perpendicular to the axis Y of the fuel injector head 60 of the fuel injector 56. The radially inner surface 102 of the fifth member 90 is generally circular in cross-section in a plane perpendicular to the axis Y of the fuel injector head 60 of the fuel injector 56. The radially outer surface 112 of the downstream end 90B of the fifth member 90 converges to a minimum diameter 116 at a plane arranged perpendicular to the axis Y of the fuel injector head 60. The minimum diameter 116 of the radially outer surface 112 of the downstream end 90B of the fifth member 90 is positioned upstream of the minimum diameter 114 of the radially inner surface 110 of the downstream end 92B of the sixth member 92.
The advantage of the present invention is that the required aerodynamics are achieved for the fuel injector resulting in minimisation of NOx, maximising combustion efficiency, achieving the required ignition characteristics and lean blow out level and the required temperature for the second splitter.
The radius R1 of the radially inner surface 110 of the sixth member 92 at the minimum diameter 114 is in the range of 0.9 to 1.1 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92. The radius R2 of the radially inner surface 110 of the sixth member 92 at the trailing edges of the swirl vanes 104 is in the range 0.9 to 1.1 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92. The radius R3 of the radially outer surface 112 of the fifth member 90 at the trailing edges of the swirl vanes 104 is in the range 0.9 to 1.1 times the radial width T of the annular duct 106 at the minimum diameter 114 of the radially inner surface 110 of the sixth member 92.
The radially inner surface 110 of the sixth member 92 is defined by a tangent connecting the radius R2 and the radius R1 and the length L4 is the length of a tangent from the radius R2 at the angle α1. The angle α1 of the divergent portion of the radially inner surface 110 of the sixth member 92 downstream of the position of minimum diameter 114 is in the range of 35° to 45°. The radially outer surface 112 of the fifth member 90 is defined by a tangent connecting the radius R3 and the minimum diameter point 116.
The advantage of the arrangement of the arrangement of the fourth swirler, the main outer air swirler, 103 of the fuel injector 56 as described is that the aerodynamic flow of air from the fourth swirler 103 attaches to the radially inner surface 110 of the sixth member 92 both upstream and downstream of the minimum diameter 114 of the radially inner surface 110 and the air from the fourth swirler 103 attaches to the radially outer surface 112 of the fifth member 90. This maximises the mixing of the air from the third and fourth swirlers 93 and 103 and hence maximises the mixing of the fuel from the downstream end of the radially inner surface 102 of the fifth member 90 into the air from the third and fourth swirlers 93 and 103 respectively. This minimises emission and maximises combustion efficiency. The angle α1 of the diverging portion of the radially inner surface 110 of the downstream end 92B of the sixth member 92 is optimised to direct the flow of fuel and air towards the igniter, to enable ignition of the fuel and air. The presence of the diverging portion of the radially inner surface 110 of the downstream end 92 of the sixth member 92 downstream of the minimum diameter 114 of the radially inner surface 110 of the sixth member 92 stabilises the flow cone angle temporarily, which is beneficial in suppressing combustion instabilities and producing consistent ignition. Thus, this arrangement prevents flow separation of the air flow from, or significant boundary layer thickness build up on, the radially inner surface 110 of the sixth member 92 and the radially outer surface 112 of the fifth member 90 which impair mixing of the air from the third and fourth swirlers 93 and 103 respectively and the mixing of the fuel into the air from the third and fourth air swirlers 93 and 103.
As mentioned previously the pilot fuel is supplied by the internal fuel passages 78 and the fuel swirler 80 onto the radially inner surface 82 of the second member 68 and the pilot fuel is atomised by swirling flows of air A and B from the swirl vanes 72 and 84 of the first and second air swirlers 71 and 83 respectively. The first member 66, the second member 68, the third member 70, the first swirler 71 and the second swirler 83 form the pilot fuel injector 59 and in this case the pilot fuel injector 59 is a pilot air blast fuel injector. In addition the main fuel supplied by the internal fuel passages 98 and fuel swirler 100 onto the radially inner surface 102 of the fifth member 90 is atomised by swirling flows of air C and D from the swirl vanes 94 and 104 of the third and fourth air swirlers 93 and 103 respectively. The fourth member 88, the fifth member 90, the sixth member 92, the third swirler 93 and the fourth swirler 103 form the main fuel injector 61 and in this case the main fuel injector 61 is a main air blast fuel injector 61. Thus, the pilot fuel and the main fuel are both atomised by high air velocity accelerating the fuel from the respective pre-filming surface, and thus the fuel pressure does not affect the fuel atomisation. An advantage of this type of fuel injector is that pilot fuel only is supplied to the pilot fuel injector 59 for and during a “cold day” take-off, in order to obtain satisfactory combustion efficiency. The pilot fuel injector is provided with an increased number of fuel passages 78, fuel passages 78 with a greater diameter etc. in order to provide a greater flow of fuel to the pilot fuel injector 59 during a “cold day” take off. On the contrary if pilot fuel is supplied to the pilot fuel injector 59 and main fuel is supplied to the main fuel injector 61 for and during a “cold day” take-off, unsatisfactory combustion efficiency, lower than a predetermined level of efficiency, is achieved. Unacceptable combustion efficiency is obtained by supplying pilot fuel and main fuel to the fuel injector because the temperature in the combustion chamber is not hot enough for the main fuel to burn at lean conditions. The modified pilot fuel injector 59 also enables relight ignition to be achieved by supplying pilot fuel only to the pilot fuel injector 59.
Although the present invention has been described with reference to the use of a separate first air splitter, a separate second air splitter and an air swirler positioned between the first and second air splitters it may be possible to provide an arrangement in which the first air splitter and the second air splitter diverge from a downstream end of a single annular member.
Although the present invention has been described with reference to a gas turbine engine combustion chamber it may be possible to use the present invention in other types of combustion chambers.
Although the present invention has been described with reference to a turbofan gas turbine engine it is equally possible to use the present invention on a combustion chamber of a turbo-jet gas turbine engine, a turbo-shaft gas turbine engine or a turbo-prop gas turbine engine. Although the present invention has been described with reference to an aero gas turbine engine it is equally possible to use the present invention on a combustion chamber of a marine gas turbine engine, an industrial gas turbine engine or an automotive gas turbine engine.
| Number | Date | Country | Kind |
|---|---|---|---|
| 1310261.1 | Jun 2013 | GB | national |
