FR3117193A1 - Burner with improved primary combustion zone - Google Patents

Abstract

A lean burn burner (16) includes a plurality of lean burn fuel injectors (50), each including a fuel feed arm (52) and a lean burn fuel injector head (54) with a lean burn fuel injector head tip (72), the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radial outward from the pilot fuel injector; and a combustion chamber (60) extending along an axial direction (62) and including a radially inner annular wall (64), a radially outer annular wall (66) and a metering panel (68) defining the size and shape of the combustion chamber, wherein the combustion chamber includes a primary combustion zone (80) and a secondary combustion zone (82). Figure for abstract: [Figure. 5]

Description

Burner with improved primary combustion zone

This disclosure relates to combustion equipment, and in particular to lean burn burners for gas turbine engines for aeronautical, industrial and marine applications.

Technical background

A gas turbine engine for aeronautical applications typically includes, in an axial flow arrangement, a fan, one or more compressors, a combustion system and one or more turbines. The combustion system typically includes a plurality of fuel injectors having fuel spray nozzles that combine fuel and air streams and generate jets of atomized liquid fuel in a combustion chamber. The mixture of air and atomized liquid fuel is then burned in the combustion chamber and the resulting hot combustion products subsequently expand through the turbine(s), and so entrain them.

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

In order to solve the problem of NOx emissions, "lean burn" combustion technology has been proposed. During lean-burn combustion, the air-fuel ratio (AFR) is above a stoichiometric ratio, keeping the combustion temperature within known limits to reduce NOx production.

On the other hand, keeping the combustion temperature relatively low could lead to incomplete or weak combustion, which in turn can lead to 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 failure of components in the engine and/or passenger discomfort, depending on the frequency of the rumble.

Gas turbine engines for industrial and marine applications face similar challenges as gas turbine engines for aeronautical applications.

There is therefore a need to provide a lean burn combustion system for aircraft, industrial and marine engines which makes it possible to reduce engine emissions of NOx, as well as CO and UHC, and to improve operability. of engine.

Disclosure Summary

According to a first aspect, a lean burn burner is provided, which burner comprises: a plurality of lean burn fuel injectors, each comprising a fuel delivery 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 injector head of lean burn fuel comprising a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outward from the pilot fuel injector; and a combustion chamber extending along an axial direction and comprising a radially inner annular wall, a radially outer annular wall and a metering panel provided upstream of the radially inner and radially outer annular walls with a plurality of openings adapted to receive the tip tips of lean burn fuel injectors. The radially inner annular wall, radially outer annular wall, and metering panel define the size and shape of the combustion chamber, where the combustion chamber has a combustion chamber length (L) and includes a primary combustion zone with a primary combustion zone length (Z) and a primary combustion zone depth (D), and a secondary combustion zone with a secondary combustion zone length (L-Z) arranged downstream of the primary combustion zone. According to the first aspect, an L/D ratio of combustion chamber length to primary combustion zone depth is less than 2.0.

In this disclosure, the upstream and downstream positions are defined with respect to the flow of fuel and air through the burner, and the fore and aft positions are defined with respect to the lean burn burner, i.e. that the lean burn fuel injectors are at the front and the combustion chamber is at the rear.

A unique combination of non-dimensional parameters for the combustion chamber has been found, which allows the development of burner aerodynamics to optimize combustion efficiency and minimize NOX and smoke. The lean burn burner according to the disclosure allows a so-called S-shaped recirculation zone to form in the primary combustion zone of the combustion chamber, which allows the pilot fuel nozzle to promote the main fuel nozzle combustion. In particular, a combustion chamber according to the disclosure allows the combustion mixture of pilot fuel and air from the pilot fuel injector to form an S-shaped flow recirculation. In detail, the combustion mixture of pilot fuel and air from the pilot fuel injector can reach a stagnation point in the primary combustion zone, where the local velocity of the pilot fuel and air mixture is zero, flow back to the pilot fuel injectors lean burn fuel, and be diverted (due to low static pressure in the main flow stream) to the radially inner and radially outer annular walls of the combustion chamber to join the combustion mixture of main fuel and air from the main fuel injector and promote its combustion. In other words, the combustion mixture of pilot fuel and air from the pilot fuel injector can flow along an S-shaped path.

Those skilled in the art would understand that when designing a combustor for a lean burn burner, an aerodynamic study should be performed for any combustor size to optimize aerodynamics and combustion of the mixture. fuel and air. It has surprisingly been found that a lean burn burner can be scaled up and down without affecting combustion efficiency. In other words, since the L/D ratio is non-dimensional, for a wide range of combustion chamber sizes of the lean burn burner according to the disclosure, the S-shaped recirculation zone can be formed efficiently. and efficient in the primary combustion zone.

For example, a lean burn burner can be sized for engines suitable for mounting on small, medium and large aircraft.

In embodiments, the L/D ratio of combustion chamber length L to primary combustion zone depth D may be less than 1.9, for example 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 combustion chamber length L to the primary combustion zone depth D may 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.

The lean burn fuel injector head may extend generally along a longitudinal direction, the longitudinal direction forming an angle of inclination α _cant with the axial direction, the angle of inclination α _cant being between 0° and 10°.

The combustion chamber may extend axially between the metering panel (upstream) and an annular outlet (downstream), through which the burnt gas exits the combustion chamber. The annular outlet may be defined by, and between, the radially inner annular wall and the radially outer annular wall of the combustion chamber. In the present disclosure, the combustion chamber length (L) can be defined as an axial distance between the measuring panel and the annular outlet.

The radially outer annular wall may extend substantially axially between the measurement panel and the annular outlet. In embodiments, the radially outer annular wall may form an outer angle α _outer with the axial direction, the outer angle α _outer being 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°.

The radially outer annular wall may include a first part and a second part. The first part of the radially outer annular wall can be arranged upstream of the second part of the radially outer annular wall. The first part and the second part of the radially outer annular wall can be mutually aligned.

The radially inner annular wall may include a first part and a second part. The first part of the radially internal annular wall can be arranged upstream of the second part of the radially internal annular wall. The first part of the radially inner annular wall can be connected to the measurement panel. The second portion of the radially inner annular wall and the second portion of the radially outer annular wall may define the annular outlet of the combustion chamber. The first part of the radially inner annular wall can be arranged at an angle with respect to the second part of the radially inner annular wall. The first portion of the radially inner annular wall may be parallel to the radially outer annular wall. The first part of the radially inner annular wall may be parallel to the axial direction.

The first radially inner annular wall portion, the first radially outer annular wall portion, and the metering panel define the primary combustion zone.

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

In the present disclosure, the primary combustion zone depth (D) can be defined as the 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 may refer to the direction perpendicular to the first portion of the radially inner annular wall and the first portion of the radially outer annular wall.

The second part of the radially inner annular wall can converge towards the second part of the radially outer annular wall in a downstream direction. In embodiments, the second part of the radially inner annular wall can form an internal angle α _inner with the first part of the radially inner annular wall, the internal angle α _inner being between 15° and 50°, for example between 15° and 45°, or between 15° and 40°, or between 20° and 50°, or between 25° and 50°, or between 25° and 45°, or between 25° and 40°.

The second portion of the radially inner annular wall and the second portion of the radially outer annular wall may define the secondary combustion zone. The secondary combustion zone may 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 along the secondary combustion zone length (LZ). The second portion of the radially outer annular wall may extend over a length equal to the length of the secondary combustion zone (LZ). The second part of the radially inner annular wall can extend over a length equal to (LZ)/cos(α _inner ).

The respective inner surfaces of the radially inner annular wall, the radially outer annular wall, and the metering panel can define the size and shape of the combustion chamber where combustion occurs. In some publications, the radially inner annular wall, radially outer annular wall, and metering panel are referred to as combustion liners. In embodiments, the radially inner annular wall, the radially outer annular wall, and the measurement panel may each include respective tiles. The tiles can define the respective inner surfaces of the radially inner annular wall, the radially outer annular wall and the metering panel, and therefore define the size and shape of the combustion 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 measuring panel can face the combustion process in the combustion chamber and can be in contact with the mixture fuel and air, and/or combustion gases.

The inventors of the present disclosure have also found that other non-dimensional parameters can be advantageous when designing a combustor for a lean burn burner with improved combustion efficiency.

In embodiments, a ratio L/d of combustion chamber length L to lean burn fuel injector head tip diameter d may 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 combustion chamber length L to the lean burn fuel injector head tip diameter d may 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.

In embodiments, the ratio D/d of the primary combustion zone depth D to the lean burn fuel injector head tip diameter d may 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 may 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 may be greater than 1.3, for example greater than 1, 4, or greater than 1.5.

In embodiments, a ratio Z/d of primary combustion zone length Z to lean burn fuel injector head tip diameter d may 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 burn fuel injector head tip diameter d may 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.

Those skilled in the art would understand that, as the ratios L/d of the combustion chamber length L to the lean burn fuel injector head tip diameter d, D/d of the primary combustion zone depth D on the lean burn fuel injector head tip diameter d, and Z/d of the primary combustion zone length Z on the lean burn fuel injector head tip diameter d are all not dimensional, they can all be applied to lean burn burners and related combustion chambers, of a wide range of sizes and can help form the S-shaped recirculation zone in the primary combustion zone.

According to a second aspect, a lean burn burner is provided, which burner comprises: a plurality of lean burn fuel injectors, each comprising a fuel delivery 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 injector head of lean burn fuel comprising a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outward from the pilot fuel injector; and a combustion chamber extending along an axial direction and comprising a radially inner annular wall, a radially outer annular wall and a metering panel provided upstream of the radially inner and radially outer annular walls with a plurality of openings adapted to receive the tip tips of lean burn fuel injectors. The radially inner annular wall, radially outer annular wall, and metering panel define the size and shape of the combustion chamber, where the combustion chamber has a combustion chamber length (L) and includes a primary combustion zone with a primary combustion zone length (Z) and a primary combustion zone depth (D), and a secondary combustion zone with a secondary combustion zone length (L-Z) arranged downstream of the primary combustion zone. According to the second aspect, a D/d ratio of the primary combustion zone depth to the lean burn fuel injector head tip diameter is less than 2.4.

In embodiments, the ratio D/d of primary zone depth D to lean burn fuel injector head tip diameter d may be less than 2.3, for example 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 may be greater than 1.2, for example greater than 1.3, or greater than 1 .4, or greater than 1.5.

In embodiments, a ratio L/d of combustion chamber length L to lean burn fuel injector head tip diameter d may 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 combustion chamber length L to the lean burn fuel injector head tip diameter d may be greater than 1.8, for example greater than 1.85, or greater than 1, 9, or greater than 2.0.

In embodiments, a ratio Z/d of primary combustion zone length Z to lean burn fuel injector head tip diameter d may 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 burn fuel injector head tip diameter d may 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.

In embodiments, an L/D ratio of combustion chamber length L to primary combustion zone depth D may 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 combustion chamber length L to the primary combustion zone depth D may 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.

According to a third aspect, a lean burn burner is provided, which burner comprises: a plurality of lean burn fuel injectors, each comprising a fuel feed 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 injector head of lean burn fuel comprising a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outward from the pilot fuel injector; and a combustion chamber extending along an axial direction and comprising a radially inner annular wall, a radially outer annular wall and a metering panel provided upstream of the radially inner and radially outer annular walls with a plurality of openings adapted to receive the tip tips of lean burn fuel injectors. The radially inner annular wall, radially outer annular wall, and metering panel define the size and shape of the combustion chamber, where the combustion chamber has a combustion chamber length (L) and includes a primary combustion zone with a primary combustion zone length (Z) and a primary combustion zone depth (D), and a secondary combustion zone with a secondary combustion zone length (L-Z) arranged downstream of the primary combustion zone. According to the third aspect, an L/d ratio of the combustion chamber length to the lean burn fuel injector head tip diameter is less than 5.

In embodiments, the ratio L/d of the combustion chamber length L to the lean burn fuel injector head tip diameter d may 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 combustion chamber length L to the lean burn fuel injector head tip diameter d may 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.

In embodiments, a ratio D/d of primary zone depth D to lean burn fuel injector head tip diameter d may 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 may be greater than 1.2, for example greater than 1.3, or greater than 1 .4, or greater than 1.5.

In embodiments, an L/D ratio of combustion chamber length L to primary combustion zone depth D may 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 combustion chamber length L to the primary combustion zone depth D may 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.

In embodiments, a ratio Z/d of primary combustion zone length Z to lean burn fuel injector head tip diameter d may 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 burn fuel injector head tip diameter d may 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.

In embodiments, the lean burn burner of the first, second and third aspects described above may comprise a pre-diffuser, arranged upstream of the lean burn fuel injector head and adapted to supply the chamber with combustion of compressed air. In some publications, the pre-diffuser is simply referred to as a diffuser. The pre-diffuser may be generally annular and may include radially inner and radially outer walls defining an outlet for the compressed air. In the present disclosure, a wet space (g) can be defined as the axial distance between a midpoint between the radially inner and radially outer walls of the pre-diffuser at said outlet and a midpoint between the radially inner annular walls and radially external of the combustion chamber at the level of the measuring panel. A ratio g/d of wet space g to lean burn fuel injector head tip diameter d may be less than 1.30, for example less than 1.25, or less than 1.2, or less than 1.15. The ratio g/d of the wet space g to the lean burn fuel injector head tip diameter d may 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.

Those skilled in the art would also understand that the ratio g/d of wet space g to lean burn fuel injector head tip diameter d is non-dimensional and can be applied to lean burn burners , and related combustion chambers, of a wide range of sizes and can help form the S-shaped recirculation zone in the primary combustion zone.

According to a fourth aspect, a lean burn burner is provided, which burner comprises: a plurality of lean burn fuel injectors, each comprising a fuel delivery 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 injector head of lean burn fuel comprising a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outward from the pilot fuel injector; and a combustion chamber extending along an axial direction and comprising a radially inner annular wall, a radially outer annular wall and a metering panel provided upstream of the radially inner and radially outer annular walls with a plurality of openings adapted to receive the tip tips of lean burn fuel injectors. The radially inner annular wall, radially outer annular wall, and metering panel define the size and shape of the combustion chamber. The lean burn burner of the fourth aspect further includes a pre-diffuser, arranged upstream of the lean burn fuel injector head and adapted to supply the combustion chamber with compressed air. The pre-diffuser is generally annular and comprises radially inner and radially outer walls defining an outlet for the compressed air. A wet space (g) is defined as the axial distance between a midpoint between the radially inner and radially outer walls of the pre-diffuser at said outlet and a midpoint between the radially inner and radially outer annular walls of the chamber of combustion at the metering panel, where a g/d ratio of wet space to lean burn fuel injector head tip diameter is less than 1.30.

In embodiments, the ratio g/d of wet space g to lean burn fuel injector head tip diameter d may be less than 1.25, for example less than 1.2, or less than 1.15. The ratio g/d of the wet space g to the lean burn fuel injector head tip diameter d may 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.

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

In embodiments, a ratio D/d of primary zone depth D to lean burn fuel injector head tip diameter d may 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 may be greater than 1.2, for example greater than 1.3, or greater than 1 .4, or greater than 1.5.

In embodiments, an L/D ratio of combustion chamber length L to primary combustion zone depth D may 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 combustion chamber length L to the primary combustion zone depth D may 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.

In embodiments, a ratio L/d of combustion chamber length L to lean burn fuel injector head tip diameter d may 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 combustion chamber length L to the lean burn fuel injector head tip diameter d may 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.

In embodiments, a ratio Z/d of primary combustion zone length Z to lean burn fuel injector head tip diameter d may 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 burn fuel injector head tip diameter d may 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.

According to a fifth aspect, a gas turbine engine is provided, which engine includes a lean burn burner according to any of the aspects described above.

The gas turbine engine of the fifth aspect may be a gas turbine engine for aircraft or for industrial and marine applications.

In embodiments, the gas turbine engine may further include: an engine core including a compressor, a burner, a turbine, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades, wherein the burner is the lean burner according to any of the first, second, third and fourth aspects.

In embodiments, the compressor and the turbine can rotate around an engine main axis of rotation, and the axial direction of the combustion chamber can be parallel to the main engine rotation axis.

As previously indicated, the lean burner according to the disclosure can be dimensioned for engines adapted to be mounted on small, medium and large aircraft. Accordingly, the fan of the gas turbine engine according to the fifth aspect may have a fan diameter greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches) or 420 cm (about 165 inches). The fan diameter can be in an inclusive range between any two values of the values in the previous sentence (i.e. the values can form upper or lower limits), for example in the range from 240 cm to 280 cm or from 330 cm to 380 cm.

Arrangements of the present disclosure may be particularly, but not exclusively, beneficial for blowers which are driven through a gearbox. Accordingly, the gas turbine engine may include a gearbox which receives input from the core shaft and outputs a drive to the fan to drive the fan at a rotational speed lower than that of the core tree. Input to the gearbox can be directly from the core shaft or indirectly from the core shaft, for example via a spur gear and/or shaft. The core shaft can rigidly connect the turbine and the compressor, so that the turbine and the compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine can have any desired number of shafts that connect turbines and compressors, such as one, two or three shafts. For illustrative purposes only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor and the core shaft may be a first core shaft. The engine core may 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.

In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive (for example receive directly, for example via a generally annular duct) a flow from the first compressor.

The gearbox can be arranged to be driven by the core shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. the first core shaft in the example above ). For example, the gearbox may be arranged to be driven only by the core shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. only by the first core shaft , and not by the second kernel tree, in the example above). Alternatively, the gearbox can be arranged to be driven by any one shaft or several shafts, for example the first and/or the second shaft in the example above.

The gearbox can be a speed reducer (as long as the output to the blower has a lower rotational speed than the input from the core shaft). Any type of gearbox can be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the input shaft rotational speed divided by the output shaft rotational speed), for example greater than 2.5, by example in the range of 3 to 4.2 or 3.2 to 3.8, for example of the order of 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. The gear ratio can be, for example, between any two values of the values in the previous sentence. For illustrative purposes only, the gearbox may be a "star" gearbox having a ratio in the range of 3.1 or 3.2 to 3.8.

In one aspect, an aircraft is provided, which aircraft includes a gas turbine engine as described and/or claimed herein. The aircraft in this aspect is the aircraft for which the gas turbine engine was designed to be attached to it.

Those skilled in the art will understand that, unless mutually exclusive, a characteristic or parameter described in relation to any of the above aspects can be applied to any other aspect. Further, unless mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

Embodiments will now be described by way of example only, with reference to the figures, in which:

is a sectional side view of a gas turbine engine;

is a close-up side sectional view of an upstream portion of the gas turbine engine of the ;

is a partially exploded view of a gearbox for a gas turbine engine;

is a partial rear view of a lean burner according to the disclosure;

is a cross-sectional side view of the lean burn burner of the along the arrows AA; and

is a schematic representation of an S-shaped flow recirculation in a primary combustion zone of the lean burn burner of Figures 4 and 5.

Detailed description of the disclosure

With reference to the , a gas turbine engine, generally indicated by 10, has an engine main axis of rotation 9. The engine 10 includes an air intake 12 and a propulsion fan with a plurality of fan blades 23 which generates two flows airflow: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 which receives the core airflow A. The engine core 11 includes , in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment comprising a lean burn burner 16, a high pressure turbine 17, a low pressure turbine 19 and an exhaust nozzle core 20. A nacelle 21 generally surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Bypass airflow B flows through the bypass duct 22 The fan is attached to and driven by the low pressure turbine 19 through the between a shaft 26 and an epicyclic gearbox 30.

In use, core A airflow is accelerated and compressed by low pressure compressor 14 and directed into high pressure compressor 15 where further compression takes place. The compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is burned. The resulting hot combustion products then expand through the high pressure and low pressure turbines 17, 19 and so entrain them before being vented through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 through a suitable interconnecting shaft 27. The fan generally provides most of the propulsive thrust. The epicyclic gearbox 30 is a speed reducer.

It should be noted that the terms "low pressure turbine" and "low pressure compressor", as used herein, can be understood to refer to the turbine stages having the lowest pressure and the compressor stages having the lowest pressure ( i.e. not including the fan) respectively and/or the turbine and compressor stages which are connected together by the interconnecting shaft 26 having the lowest rotational speed in the engine (c' i.e. not including the gearbox output shaft which drives the blower). In some publications, the terms "low pressure turbine" and "low pressure compressor" referred to herein may also be referred to as "intermediate pressure turbine" and "intermediate pressure compressor". When such alternate nomenclature is used, the fan may be referred to as the first compression stage or the compression stage with the lowest pressure.

Other gas turbine engines to which this disclosure may be applied may have alternate configurations. By way of example, such engines may have an alternating number of interconnecting shafts (e.g. two) and/or an alternating number of compressors and/or turbines. Furthermore, the motor may be a gearless motor, i.e. the motor may not include a gear box provided in the kinematic chain from the turbine to the compressor and/or the fan.

The further illustrates the gearbox 30 of the gas turbine engine 10. The low pressure turbine 19 (see ) drives shaft 26, which is coupled to sun gear, or sun gear, 28 of epicyclic gear arrangement 30. A plurality of planet gears 32 which are coupled together by planet carrier 34 are located radially exterior of the sun gear 28 and mesh with it. The planet carrier 34 constrains the planets 32 to precess around the planetary 28 in a synchronized manner while allowing each planet 32 to rotate around its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation around the motor axis 9. A ring gear or a toothed wheel 38 which is coupled, via linkages 40, to a stationary support structure 24 is located radially outward from the satellites 32 and meshes therewith.

The epicyclic gearbox 30 is shown by way of example in more detail in the . Each of sun gear 28, planet wheels 32 and gear wheel 38 includes teeth around its periphery to mesh with the other gears. However, for the sake of clarity, only exemplary parts of the teeth are shown in the . There are four satellites 32 illustrated, although it will be apparent to the skilled reader that more or fewer satellites 32 may be provided within the scope of the claimed invention. Practical applications of a 30 planetary epicyclic gearbox typically include at least three 32 planet gears.

The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the toothed wheel 38 fixed. However, any other suitable type of epicyclic gearbox 30 can be used. As a further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held stationary, with the gear wheel (or ring gear) 38 allowed to rotate. In such an arrangement, fan 23 is driven by gear wheel 38. As a further alternative example, gearbox 30 may be a differential gearbox in which both gear wheel 38 and planet carrier 34 are permitted. to turn.

The arrangement shown in Figures 2 and 3 is given by way of example only and various variations are within the scope of this disclosure. Accordingly, the present disclosure relates to a gas turbine engine having any arrangement of gearbox styles (e.g. star or planetary), support structures, arrangement of input and output shafts and bearing locations.

Figures 4 and 5 illustrate the lean burn burner 16 in more detail.

The lean burn burner 16 includes a plurality of lean burn fuel injectors 50, each including a fuel feed arm 52 and a lean burn fuel injector head 54. The fuel feed arm 52 routes fuel from a delivery system (not shown) to lean burn fuel injector head 54, where fuel and air are mixed.

The lean burn fuel injector head 54 includes 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 includes air swirl devices (not shown for simplicity). According to known arrangements, the lean burn fuel injector head 54 may include three, four or five air swirl devices adapted to provide swirling air streams which atomize fuel from the pilot and main fuel injectors. Air swirl devices may include swirl vanes.

For example, in a three air swirl arrangement, the pilot fuel injector is provided between the inner and outer air swirl devices, the main fuel injector is also provided between the inner and outer air swirl devices. inner and outer air, the pilot fuel injector outer air swirl being the main fuel injector inner air swirl. In a four swirl device arrangement, the pilot fuel injector and the main fuel injector do not share air swirl devices, so each of the pilot fuel injector and the main fuel includes its own set of internal and external air swirl devices. In a five swirl device arrangement, an additional air swirl device is provided between the outer pilot fuel injector air swirl device and the inner fuel injector air swirl device. major.

The lean burn burner 16 further includes a combustion chamber 60 extending along an axial direction 62. In the illustrated embodiment, the axial direction 62 is substantially parallel to the engine main axis of rotation 9 In other embodiments not shown, the axial direction 62 may not be parallel to the main axis of rotation of the engine 9. In other words, the combustion chambers may extend at an angle with respect to to the axial direction 62, for example at an angle between 0° and 20°.

The combustion chamber 60 includes a radially inner annular wall 64, a radially outer annular wall 66 and a metering panel 68 provided upstream of the radially inner and radially outer annular walls 64, 66. Axially opposite the metering panel 68, the chamber combustion chamber 60 has an annular outlet 67, through which the burnt gas exits the combustion chamber 60. The annular outlet is defined between the respective downstream end portions of the radially inner annular wall 64 and the radially outer annular wall 66 of the combustion chamber 60. In other words, the combustion chamber 60 extends axially from the upstream measuring plate 68 and the downstream annular outlet 67 over a length L.

The meter panel 68 is provided with a plurality of apertures 70 to receive the lean burn fuel injectors 50. In detail, the lean burn fuel injectors 50 are connected to the meter panel 68 at a tip 72 of lean burn fuel injector head 54 which is coaxially received in aperture 70.

The lean burn fuel injector head 54 may 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 angle of inclination α _cant defined between the longitudinal direction 55 and the axial direction 62 is 0°. In an embodiment not illustrated, the lean burn fuel injector head 54 may not be coaxial with the opening 70, or in other words the angle of inclination α _cant may be different from 0° , for example between 0° and 10°.

Lean burn fuel injectors 50 are configured to inject fuel and air into combustion chamber 50. A metering panel midpoint 69 is defined at metering panel 68 midway between the wall radially inner annular 64 and the radially outer annular wall 66.

The lean burn fuel injector head tip 72 has a lean burn fuel injector head tip diameter d, which corresponds to the diameter of the aperture 70.

The radially inner annular wall 64 and the radially outer annular wall 66 are connected to the measurement panel 68 at their upstream end portions. Radially inner annular wall 64, radially outer annular wall 66, and metering panel 68 together with respective inner surfaces define the size and shape of combustion chamber 60.

In non-illustrated embodiments, radially inner annular wall 64, radially outer annular wall 66, and measurement panel 68 may each include respective tiles. If present, the tiles define the respective inner surfaces of radially inner annular wall 64, radially outer annular wall 66, and metering panel 68, and therefore the size and shape of combustion 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 measurement panel 68 face the combustion process in the combustion chamber 60 and are in contact with fuel and air mixture and/or combustion gases.

The radially outer annular wall 66 extends substantially axially between the measurement panel 68 and the annular outlet 67. In other words, the radially outer annular wall 66 forms an external angle α _outer with the axial direction 62 substantially equal to 0. °. In embodiments not shown, the radially outer annular wall 66 may extend along a direction which forms with the axial direction 62 an external angle α _outer different from 0°, for example between 0° and 15° .

The radially outer annular wall 66 comprises a first part 74 and a second part 75. The first part 74 of the radially outer annular wall 66 is arranged upstream of the second part 75 of the radially outer annular wall 66. An upstream part of the first part 74 of the radially outer annular wall 66 is connected to the measurement panel 68. A downstream end part of the second part 75 of the radially outer annular wall 66 defines the annular outlet 67 of the combustion chamber 60. In the illustrated embodiment, the first part 74 and the second part 75 of the radially outer annular wall 66 are integral and mutually aligned along the axial direction 62.

The radially internal annular wall 64 comprises a first part 76 and a second part 77. The first part 76 of the radially internal annular wall 64 is arranged upstream of the second part 77 of the radially internal annular wall 64. An upstream part of the first part 76 of the radially internal annular wall 64 is connected to the measurement panel 68. A downstream end part of the second part 77 of the radially internal annular wall 64 as well as the downstream end part of the second part 75 of the radially outer annular wall 66 define the annular outlet 67 of the combustion chamber 60. The first part 76 of the radially inner annular wall 64 is arranged at an angle with respect to the second part 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 Inner element 64 is generally parallel to first portion 74 of radially outer annular wall 66. Second portion 75 of radially inner annular wall 64 converges toward radially outer annular wall 66 in a downstream direction to form annular outlet 67. second part 77 of the radially internal annular wall 64 is arranged at an angle with respect to the first part 76 of the radially internal annular wall 64. In addition, the second part 77 of the radially internal annular wall 64 forms an internal angle α _inner with the first part 76 of the radially internal annular wall 64. The internal angle α _inner is generally between 25° and 40°. As the first part 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 part 77 of the radially inner annular wall 64 is arranged according to the internal angle α _inner relative to the axial direction 62 and to the radially outer annular wall 74.

The combustion chamber 60 includes a primary combustion zone 80 and a secondary combustion zone 82.

Primary combustion zone 80 is defined by first portion 76 of radially inner annular wall 64, first portion 74 of radially outer annular wall 66, and metering panel 68. Primary combustion zone 80 has a cross section annular and extends axially from the measurement panel 68 over 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 over the length Z. In addition, the primary combustion zone 80 extends radially, that is to say in a direction perpendicular to the axial direction 62, over a depth D between the first part 76 of the radially inner annular wall 64 and the first part 74 of the radially outer annular wall 66.

The secondary combustion zone 82, which is arranged downstream of the primary combustion zone 80, is defined by the second part 77 of the radially internal annular wall 64 and the second part 75 of the radially external annular wall 66. In practice, secondary combustion zone 82 extends from a downstream end portion of primary combustion zone 80 to annular outlet 67. Secondary combustion zone 82 extends axially over a length LZ. In the embodiment described, the second part 75 of the radially outer annular wall 66 extends over the same length LZ and the second part 77 of the radially inner annular wall 64 extends over a length equal to (LZ) sin α _inner . The second combustion zone 82 is of annular and frustoconical shape and converges downstream towards the annular outlet 67.

Combustion chamber 60 is sized such that a ratio L/D of combustion chamber length L to primary combustion zone depth D is less than 2.0, for example less than 1.60, and greater than to 1.0, for example greater than 1.25. The L/D ratio being less than 2.0, for example less than 1.60, and greater than 1.0, for example greater than 1.25, makes it possible to optimize the aerodynamics of the mixture of fuel and air from the main and pilot fuel injectors 56, 58 and relative air swirl device and to increase combustion efficiency.

This will be described in more detail with reference to the .

The pilot fuel and air mixture moves along what is called an S-shaped path 86 in the primary combustion zone 80. The pilot fuel and air mixture from the tip of the d the lean burn fuel injector 72 reaches a stagnation point SP where the local pilot fuel and air mixture velocity is zero, and is then deflected rearward toward the radially outer and radially inner annular walls 74, 76 (due to the low static pressure exerted by the main fuel and air mixture 84) where the pilot fuel and air mixture comes into contact and promotes/stabilizes the combustion of the main fuel and air mixture 84 .

The L/D ratio being less than 2.0, for example less than 1.60, and greater than 1.0, for example greater than 1.25, makes it possible to obtain the S-shaped flow recirculation of the mixture of pilot fuel and air in the primary combustion zone 80. In other words, the pilot fuel and air mixture stagnation point SP is in the primary combustion zone 80 and the pilot fuel and air mixture of air mixes with the primary air and fuel mixture 84 in the primary combustion zone 80.

Another non-dimensional parameter can have a positive effect on the formation of the S-shaped path of pilot fuel and air mixture 86 in the primary combustion zone 80.

Combustor 60 may be sized such that a ratio L/d of combustion chamber length L to lean burn fuel injector head tip diameter d is less than 5, or less than 2 .5, and greater than 1.5, or greater than 2.0. In one embodiment, the combustor 60 may have an L/d ratio of 3.5.

Further, the combustion chamber 60 can be sized such that the ratio D/d of the primary combustion zone depth D to the lean burn fuel injector head tip diameter d is between 1.2 and 2.4, preferably between 2.0 and 2.4. In one embodiment, combustor 60 may have a D/d ratio of 2.2.

Additionally, the combustion chamber 60 may be sized such that a ratio Z/d of the primary combustion zone length L to the lean burn fuel injector head tip diameter d is greater than 0, 70 and less than 1.40, preferably between 0.9 and 1.25. In one embodiment, combustor 60 may have a Z/d ratio of 1.05.

The above ratios (L/d, D/d, and Z/d) can help optimize the aerodynamics of the fuel and air mixture from the main and pilot fuel injectors 56, 58 and swirl device. relative air and increase combustion efficiency.

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

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

The pre-diffuser 90 is arranged upstream of the lean burn fuel injector head 54 at a distance g (wet space) from the metering panel 68. The wet space g is defined as the axial distance between the midpoint pre-diffuser 98 and the metering panel midpoint 69. lean burn fuel injector head d may be less than 1.30, e.g., less than 1.15, and greater than 0.65, e.g., greater than 0.85. In one embodiment, the combustor 60 may have a g/d ratio of 1.05.

The arrangement of the pre-diffuser 90 at a distance from the metering panel 68 so that the ratio g/d of the wet space g to the lean burn fuel injector head tip diameter d can be less than 1.30 and greater than 0.65 can further improve the aerodynamics of the main fuel and air and pilot fuel and air mixtures in the combustion chamber 60, and particularly in the combustion zone primary 80.

Although the present disclosure has been described with reference to a turbofan gas turbine engine, it is also possible to use the present disclosure on a turbojet gas turbine engine, a turboshaft engine or a turboprop gas turbine engine. Although the present disclosure has been described with reference to an aeronautical gas turbine engine, it is also possible to use the present disclosure on a marine gas turbine engine, or an industrial gas turbine engine.

Claims (15)

Lean burn burner (16) comprising:
- a plurality of lean burn fuel injectors (50), each comprising a fuel feed arm (52) and a lean burn fuel injector head (54) with a lean burn fuel (72), wherein the lean burn fuel injector head tip (72) has a lean burn fuel injector head tip diameter (d), the injector head of lean burn fuel (54) comprising a pilot fuel injector (56) and a main fuel injector (58), the main fuel injector (58) being arranged coaxially and radially outward from the injector pilot fuel (56); and
- a combustion chamber (60) extending along an axial direction (62) and comprising a radially inner annular wall (64), a radially outer annular wall (66) and a measurement panel (68) provided in upstream of radially inner and radially outer annular walls with a plurality of apertures (70) adapted to receive lean burn fuel injector tip tips (72), the radially inner annular wall (64), the radially inner annular wall external (66) and the measurement panel (68) defining the size and shape of the combustion chamber (60), wherein the combustion chamber (60) has a combustion chamber length (L) and includes an area of primary combustion (80) with a primary combustion zone length (Z) and a primary combustion zone depth (D), and a secondary combustion zone (82) with a secondary combustion zone length (LZ) arranged in downstream of the primary combustion zone (80);
wherein an L/D ratio of combustion chamber length to primary combustion zone depth is less than 2.0. Lean burn burner according to Claim 1, characterized in that the L/D ratio of the length of the combustion chamber to the depth of the primary combustion zone is less than 1.6. Lean burner according to one of the preceding claims, characterized in that the ratio L/D of the length of the combustion chamber to the depth of the primary combustion zone is greater than 1.0, preferably greater than 1, 25. A lean burn burner according to any preceding claim, characterized in that an L/d ratio of combustion chamber length to lean burn fuel injector head tip diameter is less than 5, preferably less than 3; and/or wherein an L/d ratio of combustion chamber length to lean burn fuel injector head tip diameter is greater than 1.5, preferably greater than 2.0. A lean burn burner according to any preceding claim, characterized in that a ratio D/d of primary combustion zone depth to lean burn fuel injector head tip diameter is less than 2 .4, preferably less than 2.0; and/or wherein a D/d ratio of primary combustion zone depth to lean burn fuel injector head tip diameter is greater than 1.2, preferably greater than 1.5. Lean burn burner according to any of the preceding claims, characterized in that a Z/d ratio of primary combustion zone length to lean burn fuel injector head tip diameter is less than 1 .40, preferably less than 1.20; and/or wherein a Z/d ratio of primary combustion zone length to lean burn fuel injector head tip diameter is greater than 0.70, preferably greater than 0.90. Lean burn burner according to any one of the preceding claims, characterized in that in addition a pre-diffuser (90), arranged upstream of the lean burn fuel injector head (54) and adapted to supply the combustion chamber (60) of the compressed air, the pre-diffuser (90) being generally annular and comprising radially internal and radially external walls (92, 94) defining an outlet (96), a wet space (g) being defined as an axial distance between a midpoint (98) between the radially inner and radially outer walls (92, 94) of the pre-diffuser (90) at said outlet (96) and a midpoint (69) between the walls radially inner and radially outer annuli (92, 94) of the combustion chamber (60) at the measurement panel (68), wherein a ratio g/d between the wet space and the tip tip diameter of lean burn fuel injector is less than 1.3, preferably less than 1.15. Lean burn burner according to the preceding claim, characterized in that the ratio g/d between the wet space and the tip diameter of the lean burn fuel injector head is greater than 0.65, preferably greater than 0 .85. Lean burner according to one of the preceding claims, characterized in that the radially outer annular wall (66) of the combustion chamber (60) forms an external angle α _outer with the axial direction (62), the external angle α _outer being between 0° and 15°. Lean burner according to one of the preceding claims, characterized in that the radially inner annular wall (64) of the combustion chamber (60) comprises a first part (76) and a second part (77), the second part (77) forming an internal angle α _inner with the first part (76), the internal angle α _inner being between 15° and 50°, preferably between 25° and 40°. Lean burn burner according to one of the preceding claims, characterized in that the lean burn fuel injector head (54) extends generally along a longitudinal direction (55), the longitudinal direction (55) forming an angle of inclination α _cant with the axial direction (62), the angle of inclination α _cant being between 0° and 10°. Lean burn burner according to one of the preceding claims, characterized in that the radially inner annular wall (64), the radially outer annular wall (66) and the measuring panel (68) are each provided with respective tiles, said tiles defining respective inner surfaces of the radially inner annular wall (64), the radially outer annular wall (66) and the measurement panel (68). A gas turbine engine (10) including a lean burn burner (16) according to any preceding claim. Gas turbine engine according to the preceding claim, further comprising:
- an engine core (11) comprising a compressor (14), a burner, a turbine (19), and a core shaft (26) connecting the turbine (19) to the compressor (14),
- a fan located upstream of the engine core, the fan comprising a plurality of fan blades (23),
wherein the burner is the lean burn burner (16) according to any preceding claim. Gas turbine engine according to Claim 14, characterized in that the compressor (14) and the turbine (19) rotate around an engine main axis of rotation (9), the axial direction (62) of the combustion engine (60) being parallel to the engine main axis of rotation (9).