Lean burn combustor

Abstract

A lean-burn combustor (16) comprising: a plurality of lean-burn fuel injectors (50), each lean-burn fuel injector comprising a fuel supply arm (52) and a lean-burn fuel injector head (54) with a lean-burn fuel injector head tip (72), wherein the lean-burn fuel injector head tip has a lean-burn fuel injector head tip diameter (d), the lean-burn fuel injector head comprises a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially with the pilot fuel injector and radially outwardly; and a combustor chamber (60) extending in an axial direction (62) and comprising a radially inner annular wall (64), a radially outer annular wall (66) and a metering plate (68) defining a size and shape of the combustor chamber, wherein the combustor chamber comprises a primary combustion zone (80) having a primary combustion zone depth (D) and a secondary combustion zone (82). The ratio D / d of the primary combustion zone depth to the lean-burn fuel injector head tip diameter is less than 2.4.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to combustion equipment, and in particular to lean burn combustors for gas turbine engines used in aircraft, industrial, and marine applications. BACKGROUND

[0002] Gas turbine engines for aircraft applications typically include a fan, one or more compressors, a combustion system, and one or more turbines arranged in axial flow. The combustion system typically includes a plurality of fuel injectors having fuel spray nozzles that combine fuel and air flow and produce a spray of atomized liquid fuel into a combustion chamber. The mixture of air and atomized liquid fuel is then combusted in the combustion chamber, and the resulting hot combustion products are then expanded through the one or more turbines and drive the one or more turbines thereby.

[0003] There is a continuing need to reduce the environmental impact of gas turbine engines in terms of carbon emissions and nitrogen oxides (NOx), which begin to form at high temperatures and increase exponentially with increasing temperature.

[0004] To address the NOx emissions problem, "lean burn" combustion techniques have been proposed. In lean burn combustion, the air-to-fuel ratio (AFR) is higher than the stoichiometric ratio, which allows the combustion temperature to be kept within known limits to reduce NOx production.

[0005] On the other hand, keeping relatively low combustion temperatures can result in incomplete or weak combustion, which in turn can result in the production of other pollutants, such as carbon monoxide (CO) and unburned hydrocarbons (UHC), and / or flame instability and rumbling, which in turn can cause fatigue failures of components in the engine and / or passenger discomfort, depending on the frequency of the rumbling.

[0006] Gas turbine engines for industrial and marine applications face similar challenges as gas turbine engines for aircraft applications.

[0007] Accordingly, there is a need to provide lean burn combustion systems for aircraft, industrial, and marine engines that allow for the reduction of NOx, as well as CO and UHC, emissions from the engines, and improve engine operability. SUMMARY

[0008] According to a first aspect, there is provided a lean-burn combustor comprising: a plurality of lean-burn fuel injectors, each lean-burn fuel injector comprising a fuel supply arm and a lean-burn fuel injector head with a lean-burn fuel injector head tip, wherein the lean-burn fuel injector head tip has a lean-burn fuel injector head tip diameter (d), the lean-burn fuel injector head comprises a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially with the pilot fuel injector and radially outwardly; and a combustor chamber extending in an axial direction and comprising a radially inner annular wall, a radially outer annular wall and a metering plate provided upstream of the radially inner annular wall and the radially outer annular wall, the metering plate having a plurality of apertures adapted to accommodate the lean-burn fuel injector head tips. The radially inner annular wall, the radially outer annular wall and the metering plate define the size and shape of the combustor chamber, wherein the combustor chamber has a combustor chamber length (L) and comprises a primary combustion zone having a primary combustion zone length (Z) and a primary combustion zone depth (D) and a secondary combustion zone arranged downstream of the primary combustion zone having a secondary combustion zone length (L-Z). According to the first aspect, the ratio of the primary combustion zone depth to the lean-burn fuel injector head tip diameter D / d is less than 2.4.

[0009] In the present disclosure, upstream and downstream are relative to the fuel and air flow through the combustor and front and rear are relative to the lean-burn combustor, i.e. the lean-burn fuel injectors are in the front and the combustor chamber is in the rear.

[0010] The inventors have found a unique combination of dimensionless parameters for the combustor chamber which allows for combustor aerodynamics development to optimize combustion efficiency and minimize NOx and smoke. The lean-burn combustor according to the present disclosure allows for the formation of a so-called S-shaped recirculation zone in the primary combustion zone of the combustor chamber which allows the pilot fuel nozzles to support the main fuel nozzle combustion. In particular, the inventors have found that the combustor chamber according to the present disclosure allows for the pilot fuel and air combustion mixture from the pilot fuel injectors to form an S-shaped flow recirculation. In detail, the pilot fuel and air combustion mixture from the pilot fuel injectors can reach a stagnation point of the primary combustion zone at which the pilot fuel and air mixture local velocity is zero, travel backwards towards the lean-burn fuel injectors and turn (due to the low static pressure in the main flow jet) towards the radially inner annular wall and the radially outer annular wall of the combustor chamber to join the main fuel and air combustion mixture from the main fuel injectors and support its combustion. In other words, the pilot fuel and air combustion mixture from the pilot fuel injectors can flow along an S-shaped trajectory.

[0011] The skilled person will appreciate that when designing a combustor chamber for a lean-burn combustor, aerodynamic studies must be performed on any combustor chamber dimensions in order to optimize the aerodynamics and combustion of the fuel and air mixture. The present inventors have surprisingly found that lean-burn combustors according to the present disclosure can be scaled up and down without impacting combustion efficiency. In other words, due to the fact that the ratio D / d is dimensionless, an S-shaped recirculation zone can be effectively and efficiently formed within the primary combustion zone for a wide range of dimensions of the combustor chamber of a lean-burn combustor according to the present disclosure.

[0012] For example, a lean-burn combustor according to the present disclosure can be dimensioned for fitting into an engine installed on a small, medium and large aircraft.

[0013] In embodiments, the ratio D / d of the primary combustion zone depth to the lean-burn fuel injector head tip diameter can be less than 2.3, for example, less than 2.2, or less than 2.1, or less than 2.0.

[0014] The ratio D / d of the primary combustion zone depth D to the lean-burn fuel injector head tip diameter d can be greater than 1.2. In embodiments, the ratio D / d of the primary combustion zone depth D to the lean-burn fuel injector head tip diameter d can be greater than 1.3, for example, greater than 1.4, or greater than 1.5.

[0015] The lean-burn fuel injector head can extend generally in a longitudinal direction, the longitudinal direction forming an oblique angle a 斜 with the axial direction, the oblique angle a 斜 being comprised between 0° and 10°.

[0016] The combustor chamber can extend axially between a metering plate (upstream) and an annular outlet (downstream) through which combusted gases exit the combustor chamber. The annular outlet can be defined by, and between, a radially inner annular wall and a radially outer annular wall of the combustor chamber. In the present disclosure, the combustor chamber length (L) can be defined as the axial distance between the metering plate and the annular outlet.

[0017] The radially outer annular wall can extend substantially axially between the metering plate and the annular outlet. In embodiments, the radially outer annular wall can form an outer angle a 外 with the axial direction, the outer angle a 外 being comprised between 0° and 15°, for example, between 0° and 12°, or between 0° and 10°, or between 3° and 15°, or between 5° and 15°.

[0018] The radially outer annular wall can comprise a first portion and a second portion. The first portion of the radially outer annular wall can be arranged upstream of the second portion of the radially outer annular wall. The first and second portions of the radially outer annular wall can be aligned with each other.

[0019] The radially inner annular wall can comprise a first portion and a second portion. The first portion of the radially inner annular wall can be arranged upstream of the second portion of the radially inner annular wall. The first portion of the radially inner annular wall can be connected to the metering plate. The second portion of the radially inner annular wall and the second portion of the radially outer annular wall can define an annular outlet of the combustion chamber. The first portion of the radially inner annular wall can be arranged at an angle to the second portion of the radially inner annular wall. The first portion of the radially inner annular wall can be parallel to the radially outer annular wall. The first portion of the radially inner annular wall can be parallel to the axial direction.

[0020] The first portion of the radially inner annular wall, the first portion of the radially outer annular wall and the metering plate define a primary combustion zone.

[0021] In the present disclosure, a primary combustion zone length (Z) can be defined as an axial length of the primary combustion zone. The first portion of the radially inner annular wall can define the primary combustion zone length (Z). The first portion of the radially outer annular wall can define the primary combustion zone length (Z). The first portion of the radially inner annular wall and the first portion of the radially outer annular wall can have the same length along the axial direction.

[0022] In the present disclosure, a primary combustion zone depth (D) can be defined as a radial distance between the first portion of the radially inner annular wall and the first portion of the radially outer annular wall. The term "radial" as used herein can refer to a direction perpendicular to the first portion of the radially inner annular wall and the first portion of the radially outer annular wall.

[0023] The second portion of the radially inner annular wall can converge in a downstream direction towards the second portion of the radially outer annular wall. In embodiments, the second portion of the radially inner annular wall can form an inner angle a 内 with the first portion of the radially inner annular wall. 内 The inner angle a

[0024] The second portion of the radially inner annular wall and the second portion of the radially outer annular wall can define a secondary combustion zone. The secondary combustion zone can extend between the primary combustion zone and the annular outlet of the combustion chamber. The secondary combustion zone is arranged downstream of the primary combustion zone. The secondary combustion zone extends over a secondary combustion zone length (L-Z). The second portion of the radially outer annular wall can extend over a length equal to the secondary combustion zone length (L-Z). The second portion of the radially inner annular wall can extend over a length equal to (L-Z) / cos(a 内 ).

[0025] The respective inner surfaces of the radially inner annular wall, the radially outer annular wall, and the metering plate can define the size and shape of the combustion chamber, where combustion occurs. In some literature, the radially inner annular wall, the radially outer annular wall, and the metering plate are referred to as a combustion liner. In embodiments, the radially inner annular wall, the radially outer annular wall, and the metering plate can each comprise a respective tile. The tiles can define the respective inner surfaces of the radially inner annular wall, the radially outer annular wall, and the metering plate, and thus the size and shape of the combustor chamber, where combustion occurs. The tiles, or in other words, the inner surfaces of the radially inner annular wall, the radially outer annular wall, and the metering plate, can face the combustion process within the combustion chamber and can be in contact with the fuel and air mixture and / or the combustion gases.

[0026] The inventors of the present disclosure have also discovered that other dimensionless parameters can be advantageous when designing a combustor chamber for a lean-burn combustor with improved combustion efficiency.

[0027] In embodiments, the ratio of the primary combustion zone length Z to the lean-burn fuel injector head tip diameter d, Z / d, can be less than 1.40, for example, less than 1.35, or less than 1.30, or less than 1.25, or less than 1.20. The ratio of the primary combustion zone length Z to the lean-burn fuel injector head tip diameter d, Z / d, can be greater than 0.70, for example, greater than 0.75, or greater than 0.80, or greater than 0.85, or greater than 0.90.

[0028] In embodiments, the ratio of the combustor chamber length L to the primary combustion zone depth D, L / D, can be less than 2.0, for example, less than 1.9, or less than 1.8, or less than 1.75, or less than 1.70, or less than 1.65, or less than 1.60. The ratio of the combustor chamber length L to the primary combustion zone depth D, L / D, can be greater than 1.0, for example, greater than 1.05, or greater than 1.10, or greater than 1.15, or greater than 1.20, or greater than 1.25.

[0029] In embodiments, the ratio of the combustor chamber length L to the lean-burn fuel injector head tip diameter d, L / d, can be less than 5, for example, less than 4.5, or less than 4, or less than 3.5, or less than 3, or less than 2.8, or less than 2.6, or less than 2.5, or less than 2.45, or less than 2.4. The ratio of the combustor chamber length L to the lean-burn fuel injector head tip diameter d, L / d, can be greater than 1.5, for example, greater than 1.7, or greater than 1.8, or greater than 1.85, or greater than 1.9, or greater than 2.0.

[0030] The skilled person will realize that since the ratio Z / d of the primary combustion zone length Z to the lean fuel injector head tip diameter d, the ratio L / D of the combustor chamber length L to the primary combustion zone depth D, and the ratio L / d of the combustor chamber length L to the lean fuel injector head tip diameter d are all dimensionless, they can all be applied to lean burn combustors and associated combustor chambers of a wide range of dimensions, and can contribute to the formation of an S-shaped recirculation zone within the primary combustion zone.

[0031] According to a second aspect, there is provided a lean burn combustor comprising: a plurality of lean fuel injectors, each lean fuel injector comprising a fuel supply arm and a lean fuel injector head with a lean fuel injector head tip, wherein the lean fuel injector head tip has a lean fuel injector head tip diameter (d), the lean fuel injector head comprises a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially with the pilot fuel injector and radially outwardly; and a combustor chamber extending in an axial direction and comprising a radially inner annular wall, a radially outer annular wall, and a metering plate provided upstream of the radially inner annular wall and the radially outer annular wall, the metering plate having a plurality of apertures adapted to accommodate the lean fuel injector head tips. The radially inner annular wall, the radially outer annular wall, and the metering plate define the size and shape of the combustor chamber, wherein the combustor chamber has a combustor chamber length (L) and comprises a primary combustion zone having a primary combustion zone length (Z) and a primary combustion zone depth (D), and a secondary combustion zone arranged downstream of the primary combustion zone having a secondary combustion zone length (L-Z). According to the second aspect, the ratio L / D of the combustor chamber length to the primary combustion zone depth is less than 1.8.

[0032] In embodiments, the ratio L / D of the combustor chamber length L to the primary combustion zone depth D can be less than 1.75, for example, less than 1.70, or less than 1.65, or less than 1.60. The ratio L / D of the combustor chamber length L to the primary combustion zone depth D can be greater than 1.0, for example, greater than 1.05, or greater than 1.10, or greater than 1.15, or greater than 1.20, or greater than 1.25.

[0033] In embodiments, the ratio L / d of the combustor chamber length L to the lean fuel injector head tip diameter d can be less than 5, for example, less than 4.5, or less than 4, or less than 3.5, or less than 3, or less than 2.8, or less than 2.6, or less than 2.5, or less than 2.45, or less than 2.4. The ratio L / d of the combustor chamber length L to the lean fuel injector head tip diameter d can be greater than 1.8, for example, greater than 1.85, or greater than 1.9, or greater than 2.0.

[0034] In embodiments, the ratio D / d of the primary zone depth D to the lean fuel injector head tip diameter d can be less than 2.4, for example, less than 2.3, or less than 2.2, or less than 2.1, or less than 2.0. The ratio D / d of the primary combustion zone depth D to the lean fuel injector head tip diameter d can be greater than 1.2, for example, greater than 1.3, or greater than 1.4, or greater than 1.5.

[0035] In embodiments, the ratio Z / d of the primary combustion zone length Z to the lean fuel injector head tip diameter d can be less than 1.40, for example, less than 1.35, or less than 1.30, or less than 1.25, or less than 1.20. The ratio Z / d of the primary combustion zone length Z to the lean fuel injector head tip diameter d can be greater than 0.70, for example, greater than 0.75, or greater than 0.80, or greater than 0.85, or greater than 0.90.

[0036] According to a third aspect, there is provided a lean burn combustor comprising: a plurality of lean fuel injectors, each lean fuel injector comprising a fuel supply arm and a lean fuel injector head with a lean fuel injector head tip, wherein the lean fuel injector head tip has a lean fuel injector head tip diameter (d), the lean fuel injector head comprises a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially with the pilot fuel injector and radially outwardly; and a combustor chamber extending in an axial direction and comprising a radially inner annular wall, a radially outer annular wall, and a metering plate provided upstream of the radially inner annular wall and the radially outer annular wall, the metering plate having a plurality of apertures adapted to accommodate the lean fuel injector head tips. The radially inner annular wall, the radially outer annular wall, and the metering plate define the size and shape of the combustor chamber, wherein the combustor chamber has a combustor chamber length (L) and comprises a primary combustion zone having a primary combustion zone length (Z) and a primary combustion zone depth (D), and a secondary combustion zone arranged downstream of the primary combustion zone having a secondary combustion zone length (L-Z). According to the third aspect, the ratio L / d of the combustor chamber length to the lean fuel injector head tip diameter is less than 5.

[0037] In embodiments, the ratio L / d of the combustor chamber length L to the lean fuel injector head tip diameter d can be less than 4.5, for example, less than 4, or less than 3.5, or less than 3, or less than 2.8, or less than 2.6, or less than 2.5, or less than 2.45, or less than 2.4. The ratio L / d of the combustor chamber length L to the lean fuel injector head tip diameter d can be greater than 1.5, for example, greater than 1.7, or greater than 1.8, or greater than 1.85, or greater than 1.9, or greater than 2.0.

[0038] In embodiments, the ratio of the primary zone depth D to the lean fuel injector head tip diameter d, D / d, can be less than 2.4, for example, less than 2.3, or less than 2.2, or less than 2.1, or less than 2.0. The ratio of the primary combustion zone depth D to the lean fuel injector head tip diameter d, D / d, can be greater than 1.2, for example, greater than 1.3, or greater than 1.4, or greater than 1.5.

[0039] In embodiments, the ratio of the combustor chamber length L to the primary combustion zone depth D, L / D, can be less than 2.0, for example, less than 1.9, or less than 1.8, or less than 1.75, or less than 1.70, or less than 1.65, or less than 1.60. The ratio of the combustor chamber length L to the primary combustion zone depth D, L / D, can be greater than 1.0, for example, greater than 1.05, or greater than 1.10, or greater than 1.15, or greater than 1.20, or greater than 1.25.

[0040] In embodiments, the ratio of the primary combustion zone length Z to the lean fuel injector head tip diameter d, Z / d, can be less than 1.40, for example, less than 1.35, or less than 1.30, or less than 1.25, or less than 1.20. The ratio of the primary combustion zone length Z to the lean fuel injector head tip diameter d, Z / d, can be greater than 0.70, for example, greater than 0.75, or greater than 0.80, or greater than 0.85, or greater than 0.90.

[0041] In embodiments, the lean burn combustor of the first, second, and third aspects described above can include a pre-diffuser arranged upstream of the lean fuel injector head and adapted to provide compressed air to the combustor chamber. In some literature, the pre-diffuser is simply referred to as a diffuser. The pre-diffuser can be generally annular and can include a radially inner wall and a radially outer wall defining an outlet for the compressed air. In the present disclosure, a buffer gap (g) can be defined as the axial distance between the midpoint of the radially inner wall and the radially outer wall of the pre-diffuser at the outlet and the midpoint between the radially inner annular wall and the radially outer annular wall of the combustor chamber at the metering plate. The ratio of the buffer gap g to the lean fuel injector head tip diameter d, g / d, can be less than 1.30, for example, less than 1.25, or less than 1.2, or less than 1.15. The ratio of the buffer gap g to the lean fuel injector head tip diameter d, g / d, can be greater than 0.65, for example, greater than 0.7, or greater than 0.75, or greater than 0.8, or greater than 0.85.

[0042] Those skilled in the art will appreciate that the ratio of the buffer gap g to the lean fuel injector head tip diameter d, g / d, is also dimensionless and can be applicable to a wide range of sizes of lean burn combustors and associated combustor chambers and can be helpful in forming an S-shaped recirculation zone within the primary combustion zone.

[0043] According to a fourth aspect, there is provided a lean-burn combustor comprising: a plurality of lean-burn fuel injectors, each lean-burn fuel injector comprising a fuel supply arm and a lean-burn fuel injector head with a lean-burn fuel injector head tip, wherein the lean-burn fuel injector head tip has a lean-burn fuel injector head tip diameter (d), the lean-burn fuel injector head comprises a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially with the pilot fuel injector and radially outwardly; and a combustor chamber extending in an axial direction and comprising a radially inner annular wall, a radially outer annular wall and a metering plate provided upstream of the radially inner annular wall and the radially outer annular wall, the metering plate having a plurality of apertures adapted to accommodate the lean-burn fuel injector head tips. The radially inner annular wall, the radially outer annular wall and the metering plate define the size and shape of the combustor chamber. The lean-burn combustor of the fourth aspect further comprises a pre-diffuser arranged upstream of the lean-burn fuel injector heads and adapted to provide compressed air to the combustor chamber. The pre-diffuser is generally annular and comprises a radially inner wall and a radially outer wall defining an outlet for the compressed air. A buffer gap (g) is defined as the axial distance between a midpoint of the radially inner wall and the radially outer wall of the pre-diffuser at the outlet and a midpoint between the radially inner annular wall and the radially outer annular wall of the combustor chamber at the metering plate, wherein the ratio g / d of the buffer gap to the lean-burn fuel injector head tip diameter is less than 1.30.

[0044] In embodiments, the ratio g / d of the buffer gap g to the lean-burn fuel injector head tip diameter d can be less than 1.25, for example, less than 1.2, or less than 1.15. The ratio g / d of the buffer gap g to the lean-burn fuel injector head tip diameter d can be greater than 0.65, for example, greater than 0.7, or greater than 0.75, or greater than 0.8, or greater than 0.85.

[0045] The combustor chamber of the lean-burn combustor of the fourth aspect has a combustor chamber length (L) and can define a primary combustion zone having a primary combustion zone length (Z) and a primary combustion zone depth (D), and a secondary combustion zone arranged downstream of the primary combustion zone having a secondary combustion zone length (L-Z).

[0046] In embodiments, the ratio D / d of the primary zone depth D to the lean-burn fuel injector head tip diameter d can be less than 2.4, for example, less than 2.3, or less than 2.2, or less than 2.1, or less than 2.0. The ratio D / d of the primary combustion zone depth D to the lean-burn fuel injector head tip diameter d can be greater than 1.2, for example, greater than 1.3, or greater than 1.4, or greater than 1.5.

[0047] In embodiments, the ratio L / D of the combustor chamber length L to the primary combustion zone depth D can be less than 2.0, for example, less than 1.9, or less than 1.8, or less than 1.75, or less than 1.70, or less than 1.65, or less than 1.60. The ratio L / D of the combustor chamber length L to the primary combustion zone depth D can be greater than 1.0, for example, greater than 1.05, or greater than 1.10, or greater than 1.15, or greater than 1.20, or greater than 1.25.

[0048] In embodiments, the ratio L / d of the combustor chamber length L to the lean fuel injector head tip diameter d can be less than 5, for example, less than 4.5, or less than 4, or less than 3.5, or less than 3, or less than 2.8, or less than 2.6, or less than 2.5, or less than 2.45, or less than 2.4. The ratio L / d of the combustor chamber length L to the lean fuel injector head tip diameter d can be greater than 1.5, for example, greater than 1.7, or greater than 1.8, or greater than 1.85, or greater than 1.9, or greater than 2.0.

[0049] In embodiments, the ratio Z / d of the primary combustion zone length Z to the lean fuel injector head tip diameter d can be less than 1.40, for example, less than 1.35, or less than 1.30, or less than 1.25, or less than 1.20. The ratio Z / d of the primary combustion zone length Z to the lean fuel injector head tip diameter d can be greater than 0.70, for example, greater than 0.75, or greater than 0.80, or greater than 0.85, or greater than 0.90.

[0050] According to a fifth aspect, there is provided a gas turbine engine comprising a lean burn combustor according to any of the aspects described above.

[0051] The gas turbine engine of the fifth aspect can be a gas turbine engine for an aircraft or for industrial and marine applications.

[0052] In embodiments, the gas turbine engine can further comprise: an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor; and a fan upstream of the engine core, the fan comprising a plurality of fan blades, wherein the combustor is a lean burn combustor according to any of the first, second, third, and fourth aspects.

[0053] In embodiments, the compressor and turbine can rotate about a main engine rotation axis, and the axial direction of the combustor chamber can be parallel to the main engine rotation axis.

[0054] As noted previously, the lean burn combustor according to the present disclosure can be sized for adaptation to an engine mounted on small, medium, and large aircraft. Thus, the fan of the gas turbine engine according to the fifth aspect can have a fan diameter greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (on the order of 100 inches), 260 cm, 270 cm (on the order of 105 inches), 280 cm (on the order of 110 inches), 290 cm (on the order of 115 inches), 300 cm (on the order of 120 inches), 310 cm, 320 cm (on the order of 125 inches), 330 cm (on the order of 130 inches), 340 cm (on the order of 135 inches), 350 cm, 360 cm (on the order of 140 inches), 370 cm (on the order of 145 inches), 380 cm (on the order of 150 inches), 390 cm (on the order of 155 inches), 400 cm, 410 cm (on the order of 160 inches), or 420 cm (on the order of 165 inches). The fan diameter can be within an inclusive range bounded by any two of the foregoing values (i.e., these values can form an upper or lower bound), e.g., in a range from 220 cm to 420 cm, or 240 cm to 380 cm, 240 cm to 280 cm, or 330 cm to 380 cm.

[0055] The arrangements of the present disclosure can be particularly (but not exclusively) beneficial for fans driven via a gearbox. Thus, the gas turbine engine can include a gearbox that receives input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox can be directly from the core shaft, or indirectly from the core shaft, e.g., via spur shafts and / or gears. The core shaft can rigidly connect the turbine and the compressor such that the turbine and the compressor rotate at the same speed (with the fan rotating at a lower speed).

[0056] The gas turbine engine as described and / or claimed herein can have any suitable overall structure. For example, the gas turbine engine can have any desired number of shafts connecting the turbine and the compressor, e.g., one, two, or three shafts. Purely by way of example, the turbine connected to the core shaft can be a first turbine, the compressor connected to the core shaft can be a first compressor, and the core shaft can be a first core shaft. The engine core can further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft can be arranged to rotate at a higher rotational speed than the first core shaft.

[0057] In this arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive (e.g. directly receive, e.g. via a generally annular conduit) flow from the first compressor.

[0058] The gearbox can be arranged to be driven by the spindle (e.g. the first spindle in the above example) which is configured to rotate (e.g. in use) at the lowest rotational speed. For example, the gearbox can be arranged to be driven by only the spindle (e.g. only the first spindle in the above example, and not the second spindle) which is configured to rotate (e.g. in use) at the lowest rotational speed. Alternatively, the gearbox can be arranged to be driven by any one or more of the shafts, e.g. the first and / or second shafts in the above example.

[0059] The gearbox can be a reduction gearbox (as the output to the fan is a lower rotational speed than the input from the spindle). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), e.g. greater than 2.5, e.g. in the range from 3 to 4.2 or 3.2 to 3.8, e.g. approximately or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. For example, the gear ratio can be between any two of the values in the preceding sentence. Purely by way of example, the gearbox can be a "star" gearbox having a ratio in the range from 3.1 or 3.2 to 3.8.

[0060] According to one aspect, there is provided an aircraft comprising a gas turbine engine as described and / or claimed herein. The aircraft according to this aspect is an aircraft to which the gas turbine engine has been designed to be attached.

[0061] The skilled person will appreciate that features or parameters described in relation to any one of the above aspects can be applied to any other aspect, except where mutually exclusive. Furthermore, any feature or parameter described herein can be applied to any aspect and / or in combination with any other feature or parameter described herein, except where mutually exclusive. BRIEF DESCRIPTION OF DRAWINGS

[0062] Embodiments will now be described, by way of example only, with reference to the accompanying drawings in which:

[0063] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0064] Figure 2 is Figure 1 is a close-up cross-sectional side view of an upstream portion of the gas turbine engine of

[0065] Figure 3 is a partial cutaway view of a gear case for a gas turbine engine;

[0066] Figure 4 is a partial rear view of a lean burn combustor according to the present disclosure;

[0067] Figure 5 is Figure 4 is a cross-sectional side view of the lean burn combustor of

[0068] Figure 6 is Figure 4 and Figure 5 is a schematic representation of S-shaped flow recirculation in the primary combustion zone of the lean burn combustor of DETAILED DESCRIPTION

[0069] Referring to Figure 1 , a gas turbine engine (generally designated 10) has an engine main axis of rotation 9. The engine 10 includes an intake 12 and a propulsive fan having a plurality of fan blades 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 includes, in serial axial flow, a low pressure compressor 14, a high pressure compressor 15, a combustion arrangement 16 including a lean burn combustor, a high pressure turbine 17, a low pressure turbine 19, and a core exhaust nozzle 20. A nacelle 21 generally surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan is attached to and driven by the low pressure turbine 19 via a shaft 26 and epicyclic gear case 30.

[0070] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and is directed into the high pressure compressor 15 where further compression occurs. The compressed air discharged from the high pressure compressor 15 is directed into the combustion arrangement 16 where it is mixed with fuel and the mixture is combusted. The resulting hot combustion products are then expanded through the high and low pressure turbines 17, 19 before being discharged through the nozzle 20 and thereby drive the high and low pressure turbines to provide some of the propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 through a suitable interconnecting shaft 27. The fan generally provides the majority of the propulsive thrust. The epicyclic gear case 30 is a reduction gear case.

[0071] Note that the terms "low pressure turbine" and "low pressure compressor" as used herein can be considered to mean the lowest pressure turbine stage and the lowest pressure compressor stage (i.e., excluding the fan) and / or the turbine and compressor stages that are connected together by the interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft that drives the fan). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein can alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor". Where such alternative terminology is used, the fan can be referred to as the first or lowest pressure, compression stage.

[0072] Other gas turbine engines to which the present disclosure can be applicable can have alternative configurations. By way of example, such engines can have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Additionally, the engine can be a gearless engine, i.e., the engine can not include a gearbox that provides in the drive train from the turbine to the compressor and / or fan.

[0073] Figure 2 The gearbox 30 of the gas turbine engine 10 is illustrated in more detail. The low pressure turbine 19 (see Figure 1 ) drives a shaft 26 that is coupled to a sun or sun gear 28 of the epicyclic gear arrangement 30. Radially outward of, and meshing with, the sun gear 28 is a plurality of planet gears 32 that are coupled together by a carrier 34. The carrier 34 constrains the planet gears 32 to revolve in synchronism about the sun gear 28, whilst enabling each planet gear 32 to rotate about its own axis. The carrier 34 is coupled via a link 36 to the fan 23 so as to drive its rotation about the engine axis 9. Radially outward of, and meshing with, the planet gears 32 is an annulus or ring gear 38 that is coupled via a link 40 to the static support structure 24.

[0074] Figure 3 The epicyclic gearbox 30 is shown in more detail by way of example. Each of the sun gear 28, planet gears 32 and ring gear 38 comprises teeth about its periphery to mesh with the other gears. However, for the sake of clarity, Figure 3 only exemplary portions of the teeth are illustrated. Four planet gears 32 are illustrated, but it will be apparent to those skilled in the art that more or fewer planet gears 32 can be provided within the scope of the claimed application. Practical applications of the planetary epicyclic gearbox 30 typically include at least three planet gears 32.

[0075] Figure 2 and Figure 3The epicyclic gearbox 30 illustrated by way of example in Figures is of the planetary type, in which the planet carrier 34 is coupled to the output shaft via a connecting rod 36, while the ring gear 38 is fixed. However, any other suitable type of epicyclic gearbox 30 can be used. By way of further example, the epicyclic gearbox 30 can be of the star arrangement, in which the planet carrier 34 remains fixed, the ring gear 38 being allowed to rotate. In this arrangement, the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 can be a differential gearbox, in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0076] It will be appreciated that, Figure 2 and Figure 3 The arrangements shown in Figures are by way of example only, and various alternatives are within the scope of the present disclosure. Thus, the present disclosure extends to gas turbine engines having any arrangement of gearbox style (e.g. star or planetary), support structure, input and output shaft arrangement, and bearing location.

[0077] Figure 4 and 5 The lean burn combustor 16 is illustrated in more detail.

[0078] The lean burn combustor 16 comprises a plurality of lean burn fuel injectors 50, each of which comprises a fuel supply arm 52 and a lean burn fuel injector head 54. The fuel supply arm 52 delivers fuel from a distribution system (not illustrated) to the lean burn fuel injector head 54, where the fuel and air are mixed.

[0079] The lean burn fuel injector head 54 comprises a pilot fuel injector 56 and a radially outer main fuel injector 58. The main fuel injector 58 is arranged coaxially around the pilot fuel injector 56. The lean burn fuel injector head 54 further comprises air swirler(s) (not illustrated for simplicity). According to known arrangements, the lean burn fuel injector head 54 can comprise three, four or five air swirler(s) adapted to provide a swirling air flow that atomises the fuel from the pilot and main fuel injectors. The air swirler(s) can comprise swirler vanes.

[0080] For example, in a three air swirler arrangement, the pilot fuel injector is provided between the inner and outer air swirler, the main fuel injector is also provided between the inner and outer air swirler, the pilot fuel injector outer air swirler is the main fuel injector inner air swirler. In a four swirler arrangement, the pilot and main fuel injectors do not share air swirler(s), such that each of the pilot and main fuel injectors comprises its own set of inner and outer air swirler(s). In a five swirler arrangement, an additional air swirler is provided between the outer air swirler of the pilot fuel injector and the inner air swirler of the main fuel injector.

[0081] The lean-burn combustor 16 further comprises a combustor chamber 60 extending along an axial direction 62. In the illustrated embodiment, the axial direction 62 is substantially parallel to the engine main rotation axis 9. In other non-illustrated embodiments, the axial direction 62 can not be parallel to the engine main rotation axis 9. In other words, the combustor chamber can extend at an angle with the axial direction 62, for example, at an angle comprised between 0° and 20°.

[0082] The combustor chamber 60 comprises a radially inner annular wall 64, a radially outer annular wall 66 and a metering plate 68 provided upstream of the radially inner and outer annular walls 64, 66. Axially opposite the metering plate 68, the combustor chamber 60 is characterized by an annular outlet 67 through which combusted gases exit the combustor chamber 60. The annular outlet is defined between respective downstream end portions of the radially inner and outer annular walls 64, 66 of the combustor chamber 60. In other words, the combustor chamber 60 extends axially from the upstream metering plate 68 and the downstream annular outlet 67 by a length L.

[0083] The metering plate 68 is provided with a plurality of orifices 70 for accommodating the lean-burn fuel injectors 50. In detail, the lean-burn fuel injectors 50 are connected to the metering plate 68 at a tip 72 of the lean-burn fuel injector head 54 (accommodated coaxially in the orifice 70).

[0084] The lean-burn fuel injector head 54 can extend generally along a longitudinal direction 55. In the illustrated embodiment, the longitudinal direction 55 is parallel to the axial direction 62. In other words, an oblique angle a 斜 defined between the longitudinal direction 55 and the axial direction 62 is 0°. In non-illustrated embodiments, the lean-burn fuel injector head 54 can be de-axial with the orifice 70, or in other words, the oblique angle a 斜 may be different from 0°, for example, comprised between 0° and 10°.

[0085] The lean-burn fuel injectors 50 are configured to inject fuel and air into the combustor chamber 50. A metering plate midpoint 69 is defined at an intermediate between the radially inner and outer annular walls 64, 66 at the metering plate 68.

[0086] The lean-burn fuel injector head tip 72 is characterized by a lean-burn fuel injector head tip diameter d, which corresponds to a diameter of the orifice 70.

[0087] The radially inner and outer annular walls 64, 66 are connected to the metering plate 68 at their upstream end portions. The radially inner and outer annular walls 64, 66 and the metering plate 68 define the size and shape of the combustor chamber 60 with respective inner surfaces.

[0088] In an embodiment not shown, the radially inner annular wall 64, the radially outer annular wall 66, and the metering plate 68 may each include corresponding tiles. If present, the tiles define the respective inner surfaces of the radially inner annular wall 64, the radially outer annular wall 66, and the metering plate 68, and thus define the size and shape of the combustor chamber 60 (where combustion occurs). The tiles, or in other words, the inner surfaces of the radially inner annular wall 64, the radially outer annular wall 66, and the metering plate 68 face the combustion process within the combustion chamber 60 and are in contact with the fuel-air mixture and / or combustion gases.

[0089] The radially outer annular wall 66 extends substantially axially between the metering plate 68 and the annular outlet 67. In other words, the radially outer annular wall 66 forms an exterior angle α with the axial direction 62 that is substantially equal to 0°. 外 In an embodiment not shown, the radially outer annular wall 66 may form an exterior angle α different from 0° along the axial direction 62. 外 The direction extends, for example, the outer angle includes between 0° and 15°.

[0090] The radially outer annular wall 66 includes a first portion 74 and a second portion 75. The first portion 74 of the radially outer annular wall 66 is disposed upstream of the second portion 75 of the radially outer annular wall 66. The upstream portion of the first portion 74 of the radially outer annular wall 66 is connected to the metering plate 68. The downstream end portion of the second portion 75 of the radially outer annular wall 66 defines an annular outlet 67 of the combustion chamber 60. In the illustrated embodiment, the first portion 74 and the second portion 75 of the radially outer annular wall 66 are integral and aligned with each other in the axial direction 62.

[0091] The radially inner annular wall 64 includes a first portion 76 and a second portion 77. The first portion 76 of the radially inner annular wall 64 is disposed upstream of the second portion 77 of the radially inner annular wall 64. The upstream portion of the first portion 76 of the radially inner annular wall 64 is connected to the metering plate 68. The downstream end portion of the second portion 77 of the radially inner annular wall 64 and the downstream end portion of the second portion 75 of the radially outer annular wall 66 define an annular outlet 67 of the combustion chamber 60. The first portion 76 of the radially inner annular wall 64 is arranged at an angle to the second portion 77 of the radially inner annular wall 64. The first portion 76 of the radially inner annular wall 64 is generally parallel to the axial direction 62. The first portion 76 of the radially inner annular wall 64 is generally parallel to the first portion 74 of the radially outer annular wall 66. The second portion 75 of the radially inner annular wall 64 converges toward the radially outer annular wall 66 in the downstream direction to form the annular outlet 67. The second portion 77 of the radially inner annular wall 64 is arranged at an angle to the first portion 76 of the radially inner annular wall 64. Furthermore, the second portion 77 of the radial inner annular wall 64 forms an interior angle α with the first portion 76 of the radial inner annular wall 64. 内 . Internal angle α 内Typically, this angle is between 25° and 40°. Since the first portion 76 of the radially inner annular wall 76 and the radially outer annular wall 74 are generally parallel to the axial direction 62, the second portion 77 of the radially inner annular wall 64 is at an interior angle α with respect to the axial direction 62 and the radially outer annular wall 74. 内 Arrangement.

[0092] The burner chamber 60 includes a primary combustion zone 80 and a secondary combustion zone 82.

[0093] The primary combustion zone 80 is defined by a first portion 76 of a radially inner annular wall 64, a first portion 74 of a radially outer annular wall 66, and a metering plate 68. The primary combustion zone 80 is annular in cross-section and extends axially from the metering plate 68 to a length Z. In the illustrated embodiment, both the first portion 74 of the radially outer annular wall 66 and the first portion 76 of the radially inner annular wall 64 extend axially to a length Z. Furthermore, the primary combustion zone 80 extends radially (i.e., in a direction perpendicular to the axial direction 62) to a depth D between the first portion 76 of the radially inner annular wall 64 and the first portion 74 of the radially outer annular wall 66.

[0094] The secondary combustion zone 82, located downstream of the primary combustion zone 80, is defined by a second portion 77 of a radially inner annular wall 64 and a second portion 75 of a radially outer annular wall 66. In practice, the secondary combustion zone 82 extends from the downstream end portion of the primary combustion zone 80 to the annular outlet 67. The secondary combustion zone 82 extends axially to a length LZ. In the described embodiment, the second portion 75 of the radially outer annular wall 66 extends to the same length LZ, and the second portion 77 of the radially inner annular wall 64 extends to a length equal to (LZ)·sinα. 内 The length of the second combustion zone 82 is annular and truncated conical, converging downstream toward the annular outlet 67.

[0095] The dimensions of the combustion chamber 60 are determined such that the ratio D / d of the primary combustion zone depth D to the diameter d of the lean fuel injector tip is between 1.2 and 2.4, preferably between 2.0 and 2.4. In an embodiment, the combustion chamber 60 has a ratio D / d of 2.2. A ratio D / d between 1.2 and 2.4, preferably between 2.0 and 2.4, allows for optimization of the aerodynamics of the fuel and air mixture from the main fuel injector and the pilot fuel injectors 56, 58 and associated air swirlers, and improves combustion efficiency.

[0096] This will refer to Figure 6 To describe in more detail.

[0097] The pilot fuel-air mixture travels along a so-called S-shaped trajectory 86 within the primary combustion zone 80. The pilot fuel-air mixture from the tip 72 of the lean fuel injector head reaches a stagnation point SP, where the pilot fuel-air mixture has a local velocity of zero, and then turns backward toward the radially outer and radially inner annular walls 74, 76 (due to the low static pressure exerted by the main fuel-air mixture 84), where the pilot fuel-air mixture contacts and supports / stabilizes the combustion of the main fuel-air mixture 84.

[0098] A D / d ratio between 1.2 and 2.4, preferably between 2.0 and 2.4, allows for the S-shaped flow recirculation of the pilot fuel-air mixture within the primary combustion zone 80. In other words, the pilot fuel-air mixture stagnation point SP is located within the primary combustion zone 80, and the pilot fuel-air mixture mixes with the main fuel-air mixture 84 within the primary combustion zone 80.

[0099] Other dimensionless parameters can have a positive impact on the formation of the S-shaped trajectory 86 of the pilot fuel-air mixture within the primary combustion zone 80.

[0100] The dimensions of the combustion chamber 60 can be determined such that the ratio Z / d of the primary combustion zone length L to the diameter d of the lean fuel injector tip is greater than 0.7 and less than 1.40, preferably between 0.9 and 1.25. In an embodiment, the combustion chamber 60 may have a ratio Z / d of 1.05.

[0101] Furthermore, the dimensions of the combustion chamber 60 can be determined such that the ratio L / D of the burner chamber length L to the primary combustion zone depth D is less than 2.0, for example, less than 1.60, and greater than 1.0, for example, greater than 1.25. In an embodiment, the burner chamber 60 may have a ratio L / D of 1.5.

[0102] Furthermore, the dimensions of the burner chamber 60 can be determined such that the ratio L / d of the burner chamber length L to the diameter d of the lean fuel injector tip is less than 5, or less than 2.5, and greater than 1.5, or greater than 2.0. In an embodiment, the burner chamber 60 may have a ratio L / d of 3.5.

[0103] The ratios (Z / d, L / D, and L / d) above can help optimize the aerodynamics of the fuel and air mixture from the main fuel injector and the pilot fuel injectors 56, 58 and the associated air vortex, and improve combustion efficiency.

[0104] It should be noted that all the above ratios (D / d, Z / d, L / D, and L / d) are dimensionless and therefore applicable to a wide range of lean-burn burners. For example, D can include between 90 mm and 150 mm, for example, between 110 mm and 140 mm; d can include between 60 mm and 100 mm, for example, between 70 mm and 85 mm; Z can include between 50 mm and 130 mm, for example, between 60 mm and 110 mm; and L can include between 100 mm and 200 mm.

[0105] The lean burner 16 further includes a pre-diffuser 90 for supplying compressed air from the high-pressure compressor 15 to the lean fuel injector head 54. The pre-diffuser is annular and includes a radially inner wall 92 and a radially outer wall 94 defining an outlet 96 for compressed air. The outlet pre-diffuser midpoint 98 is defined midway between the radially inner wall 92 and the radially outer wall 94 at the outlet 96.

[0106] A pre-diffuser 90 is disposed upstream of the lean fuel injector head 54 at a distance g (buffer gap) from the metering plate 68. The buffer gap g is defined as the axial distance between the midpoint 98 of the outlet pre-diffuser and the midpoint 69 of the metering plate. The pre-diffuser 90 is spaced from the combustor chamber 60 such that the ratio g / d of the buffer gap g to the diameter d of the lean fuel injector head tip can be less than 1.30, for example, less than 1.15, and greater than 0.65, for example, greater than 0.85. In an embodiment, the combustor chamber 60 can have a ratio g / d of 1.05.

[0107] Arranging the pre-diffuser 90 at a distance from the metering plate 68 such that the ratio of the buffer gap g to the diameter d of the lean fuel injector tip g / d can be less than 1.30 and greater than 0.65 can further improve the aerodynamics of the pilot fuel and main fuel-air mixture within the combustor chamber 60 (and especially within the primary combustion zone 80).

[0108] Although this disclosure has been described with reference to turbofan gas turbine engines, its use in turbojet gas turbine engines, turboshaft gas turbine engines, or turboprop gas turbine engines is equally possible. Similarly, although this disclosure has been described with reference to aviation gas turbine engines, its use in marine gas turbine engines or industrial gas turbine engines is equally possible.

Claims

1. A lean-burn combustor (16) comprising: - a plurality of lean-burn fuel injectors (50), each lean-burn fuel injector comprising a fuel supply arm (52) and a lean-burn fuel injector head (54) with a lean-burn fuel injector head tip (72), wherein the lean-burn fuel injector head tip has a lean-burn fuel injector head tip diameter d, the lean-burn fuel injector head comprises a pilot fuel injector (56) and a main fuel injector (58) arranged coaxially with the pilot fuel injector and radially outwardly; and - a combustor chamber (60) extending in an axial direction (62) and comprising a radially inner annular wall (64), a radially outer annular wall (66) and a metering plate (68) provided upstream of the radially inner annular wall (64) and the radially outer annular wall (66), the metering plate having a plurality of orifices (70) adapted to accommodate the lean-burn fuel injector head tips (72), the radially inner annular wall (64), the radially outer annular wall (66) and the metering plate (68) defining the size and shape of the combustor chamber (60), wherein the combustor chamber (60) has a combustor chamber length L and comprises a primary combustion zone (80) having a primary combustion zone length Z and a primary combustion zone depth D, and a secondary combustion zone (82) arranged downstream of the primary combustion zone (80) having a secondary combustion zone length; wherein the ratio of the primary combustion zone depth to the lean-burn fuel injector head tip diameter D / d is less than 2.

4.

2. The lean burner of claim 1, wherein, The ratio of the primary combustion zone depth to the lean-burn fuel injector head tip diameter D / d is less than 2.

0.

3. The lean burner of claim 1, wherein, The ratio of the primary combustion zone depth to the lean-burn fuel injector head tip diameter D / d is greater than 1.

2.

4. The lean burner of claim 1, wherein, The ratio of the primary combustion zone depth to the lean-burn fuel injector head tip diameter D / d is greater than 1.

5.

5. The lean-burn combustor of any one of claims 1-4, wherein, The ratio of the primary combustion zone length to the lean-burn fuel injector head tip diameter Z / d is less than 1.

40.

6. The lean burner of any one of claims 1-4, wherein, The ratio of the primary combustion zone length to the lean-burn fuel injector head tip diameter Z / d is less than 1.

20.

7. The lean-burn combustor of any one of claims 1-4, wherein, The ratio of the primary combustion zone length to the lean-burn fuel injector head tip diameter Z / d is greater than 0.

70.

8. The lean burner of any one of claims 1-4, wherein, The ratio of the primary combustion zone length to the lean-burn fuel injector head tip diameter Z / d is greater than 0.

90.

9. The lean burner of any one of claims 1-4, wherein, The ratio of the combustor chamber length to the primary combustion zone depth L / D is less than 2.

0.

10. The lean burner of any one of claims 1-4, wherein, The ratio of the combustor chamber length to the primary combustion zone depth L / D is less than 1.

60.

11. The lean-burn combustor of any one of claims 1-4, wherein, The ratio of the combustor chamber length to the primary combustion zone depth L / D is greater than 1.

0.

12. The lean-burn combustor of any one of claims 1-4, wherein, The ratio of the combustor chamber length to the primary combustion zone depth L / D is greater than 1.

25.

13. The lean-burn combustor of any one of claims 1-4, wherein, The ratio of the combustor chamber length to the lean-burn fuel injector head tip diameter L / d is less than 5.

14. The lean-burn combustor of any one of claims 1-4, wherein, The ratio of the combustor chamber length to the lean-burn fuel injector head tip diameter L / d is less than 3.

15. The lean-burn combustor of any one of claims 1-4, wherein, The ratio of the combustor chamber length to the lean-burn fuel injector head tip diameter L / d is greater than 1.

5.

16. The lean-burn combustor of any one of claims 1-4, wherein, The ratio L / d of the combustor chamber length to the lean fuel injector head tip diameter is greater than 2.

0.

17. The lean-burn combustor of any one of claims 1-4, further comprising a pre-diffuser (90) disposed upstream of the lean fuel injector head (54) and adapted to provide compressed air to the combustor chamber (60), the pre-diffuser (90) being generally annular and comprising a radially inner wall (92) and a radially outer wall (94) defining an exit (96), a buffer gap g being defined as an axial distance between a midpoint (98) of the radially inner wall (92) and the radially outer wall (94) of the pre-diffuser (90) at the exit (96) and a midpoint (69) between the radially inner annular wall (64) and the radially outer annular wall (66) of the combustor chamber (60) at the metering plate (68), wherein, The ratio g / d between the buffer gap and the lean fuel injector head tip diameter is less than 1.

3.

18. The lean burner of claim 17, wherein, The ratio g / d between the buffer gap and the lean fuel injector head tip diameter is less than 1.

15.

19. The lean burner of claim 17, wherein, The ratio g / d between the buffer gap and the lean fuel injector head tip diameter is greater than 0.

65.

20. The lean burner of claim 17, wherein, The ratio g / d between the buffer gap and the lean fuel injector head tip diameter is greater than 0.

85.

21. The dilution combustor of any one of claims 1-4, wherein, The radially outer annular wall (66) of the combustor chamber (60) forms an outer angle a with the axial direction (62) 外 The outer angle a 外 is comprised between 0° and 15°.

22. The lean burner of any one of claims 1-4, wherein, Said radially inner annular wall (64) of said combustor chamber (60) comprises a first portion (76) and a second portion (77), said second portion (77) forming an internal angle a with said first portion (76) 内 , said internal angle a 内 is comprised between 15° and 50°.

23. The lean burner of claim 22, wherein, the inner angle a 内 comprises between 25° and 40°.

24. The dilution combustor of any one of claims 1-4, wherein, The lean fuel injector head (54) extends generally in a longitudinal direction (55) forming an oblique angle a 斜 with the axial direction (62). The oblique angle a 斜 is comprised between 0° and 10°.

25. The dilution combustor of any one of claims 1-4, wherein, The radially inner annular wall (64), the radially outer annular wall (66) and the metering plate (68) are each provided with a respective tile defining a respective inner surface of the radially inner annular wall (64), the radially outer annular wall (66) and the metering plate (68).

26. A gas turbine engine (10) comprising a lean burn combustor (16) according to any one of claims 1-25.

27. The gas turbine engine according to claim 26, further comprising: - an engine core (11) comprising a compressor (14), the lean burn combustor (16), a turbine (19) and a core shaft (26) connecting the turbine (19) to the compressor (14), - a fan upstream of the engine core, the fan comprising a plurality of fan blades (23).

28. A gas turbine engine as claimed in claim 27 wherein, The compressor (14) and turbine (19) rotate around an engine main rotation axis (9), the axial direction (62) of the combustor chamber (60) being parallel to the engine main rotation axis (9).

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