阅读说明：本技术 气体涡轮引擎分叉 (Gas turbine engine bifurcation ) 是由 卢西恩·英格利 于 2019-12-23 设计创作，主要内容包括：本发明公开了一种气体涡轮引擎,该气体涡轮引擎包括：引擎核心；环形流动通道,该环形流动通道围绕该引擎核心布置；以及分叉,该分叉在该引擎核心下方延伸并且具有跨越该环形流动通道的长度尺寸。该分叉包括容纳导管的外壳。该导管在该分叉的引擎核心端部处具有入口,并且沿该分叉的长度延伸。该导管包括耐火材料,并且被构造成将流体从该引擎核心排出并通过该分叉至出口。(The invention discloses a gas turbine engine, comprising: an engine core; an annular flow passage disposed around the engine core; and a bifurcation extending below the engine core and having a length dimension spanning the annular flow passage. The bifurcation includes a housing that houses the conduit. The conduit has an inlet at the engine core end of the bifurcation and extends along the length of the bifurcation. The conduit includes a refractory material and is configured to discharge fluid from the engine core and through the bifurcation to an outlet.)

1. A gas turbine engine, the gas turbine engine comprising:

an engine core;

an annular flow passage disposed around the engine core; and

a bifurcation extending below the engine core and having a length dimension spanning the annular flow passage;

wherein the bifurcation includes a housing containing a conduit having an inlet at an engine core end of the bifurcation and extending along a length of the bifurcation, the conduit comprising a refractory material and configured to discharge fluid from the engine core and through the bifurcation to an outlet.

2. The gas turbine engine of claim 1, wherein the housing comprises a contoured outer surface and the cross-sectional profile of the conduit is shaped to follow at least a portion of the profile of the outer surface.

3. The gas turbine engine of claim 1, wherein the conduit comprises a different material than the housing and provides structural reinforcement to the housing.

4. The gas turbine engine of claim 1, the inlet of the duct being located at a lowermost portion of the engine core.

5. The gas turbine engine of claim 1, wherein the bifurcation includes a leading edge and a trailing edge, the conduit being located in the trailing edge of the bifurcation.

6. The gas turbine engine of claim 5, wherein the cross-sectional profile of the conduit is conical or triangular.

7. The gas turbine engine of claim 1, wherein the duct has a substantially constant cross-sectional profile along its length.

8. The gas turbine engine of claim 1, wherein the conduit extends the entire length of the bifurcation so as to isolate the interior of the conduit from the remainder of the interior of the bifurcation.

9. The gas turbine engine of claim 1, wherein the refractory material is titanium.

10. The gas turbine engine of claim 1, comprising one or more exhaust ducts mounted within the duct.

11. The gas turbine engine of claim 1, wherein the engine core includes a fluid trap that collects fluid for treatment through the bifurcated conduit.

12. The gas turbine engine of claim 1, wherein the outlet of the conduit opens to a vent in a nacelle structure surrounding the annular flow passage.

13. The gas turbine engine of claim 1, wherein the outlet of the conduit discharges to the environment.

14. The gas turbine engine of claim 1, wherein the engine core comprises a turbine, a compressor, and a spindle connecting the turbine to the compressor; and the gas turbine engine further comprising:

a fan located upstream of the engine core, the fan comprising a plurality of fan blades;

a fan housing; and

a gearbox that receives an input from the spindle and outputs a drive to the fan to drive the fan at a lower rotational speed than the spindle.

15. The gas turbine engine of claim 14, wherein:

the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle;

the engine core further comprising a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor; and is

The second turbine, the second compressor and the second spindle are arranged to rotate at a higher rotational speed than the first spindle.

Technical Field

The present disclosure relates to a bifurcation of a flow passage for a gas turbine engine, such as for a gas turbine engine.

Background

In gas turbine engines, such as turbofan engines, a fork spans the bypass flow duct. The bifurcation includes a static support/tower structure between the outer fan nacelle and the inner engine core casing. The interior of the bifurcation also houses and protects the fuel lines, hydraulic lines, conduits, electrical wires, communication lines, and other components that support the operation of the engine.

The radially inner end of the bifurcation is disposed within the flame retardant region of the engine core. Thus, the interior of the bifurcation is itself the fire zone. A fire zone represents an area within a gas turbine engine where there is a potential risk of fire. This requires that any component with a fire resistant area must have increased fire resistance and therefore places stringent requirements on the design and construction of the furcation and any piping passing therethrough.

It is an object of the present disclosure to find an alternative bifurcated design configuration.

Disclosure of Invention

The present disclosure provides a gas turbine engine and bifurcation structure as recited in the appended claims.

According to a first aspect, there is provided a gas turbine engine comprising: an engine core; an annular flow passage disposed around the engine core; and a bifurcation extending below the engine core and having a length dimension spanning the annular flow passage; wherein the bifurcation includes a housing containing a conduit having an inlet at an engine core end of the bifurcation and extending along a length of the bifurcation, the conduit comprising a refractory material and configured to discharge fluid from the engine core and through the bifurcation to an outlet.

The housing may comprise a profiled outer surface and the cross-sectional profile of the conduit may be shaped to follow at least a portion of the profile of the outer surface.

The conduit may comprise a different material than the housing and may provide structural reinforcement of the housing.

The inlet of the duct may be located in the lowermost part of the engine core.

The bifurcation may include a leading edge and a trailing edge, with the conduit being located in the trailing edge of the bifurcation.

The cross-sectional profile of the conduit may be tapered or triangular.

The conduit may have a substantially constant cross-sectional profile along its length.

The catheter may extend the entire length of the bifurcation to isolate the interior of the catheter from the rest of the interior of the bifurcation.

The refractory material may be titanium.

The gas turbine engine may also include one or more exhaust ducts mounted within the duct.

The engine core may include a fluid trap that collects fluid for treatment through the bifurcated conduit.

The outlet of the conduit may lead to a discharge port in the nacelle structure surrounding the annular flow passage.

The outlet of the conduit may be vented to the environment.

The conduit may include an outlet to a drain within the nacelle structure surrounding the annular flow passage.

The engine core may include a turbine, a compressor, and a spindle connecting the turbine to the compressor; and the gas turbine engine further comprises: a fan located upstream of the engine core, the fan may include a plurality of fan blades; a fan housing; and a gearbox receiving an input from the spindle and outputting a drive to the fan to drive the fan at a lower rotational speed than the spindle.

The turbine may be a first turbine, the compressor may be a first compressor, and the spindle may be a first spindle; the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor; and the second turbine, the second compressor and the second spindle may be arranged to rotate at a higher rotational speed than the first spindle.

According to a second aspect, there is provided a bifurcating structure for spanning an annular flow passage around an engine core of a gas turbine engine, the bifurcating structure comprising: a housing having a cross-sectional profile and a first end arranged to be mounted to an engine core, the housing extending along a length of a bifurcated structure, wherein the bifurcated structure includes a conduit having an inlet at the first end and extending through a length of the housing, the conduit having a cross-sectional profile defining a partition within the cross-sectional profile of the housing.

Drawings

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

FIG. 1 is a cross-sectional side view of a gas turbine engine;

FIG. 2 is a close-up cross-sectional side view of an upstream portion of a gas turbine engine;

FIG. 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

FIG. 4 is a partially cut-away isometric view of an aft portion of a gas turbine engine;

FIG. 5 is a close-up side view of the furcation of FIG. 4;

FIG. 6 shows an isolated isometric and top view of a bifurcation;

FIG. 7 shows an isometric view of a bifurcation having a discharge tube and a close-up view of the discharge tube; and is

FIG. 8 shows a cross-sectional side view of a gas turbine engine with indications of different zones.

Detailed Description

Aspects and embodiments of the present disclosure will now be discussed with reference to the drawings. Other aspects and embodiments will be apparent to those skilled in the art.

Fig. 1 shows a gas turbine engine 10 having a main axis of rotation 9. The engine 10 comprises an air intake 12 and a propeller fan 23 which generates two air flows: core stream a and bypass stream B. The gas turbine engine 10 includes an engine core 11 that receives a core gas flow a. The engine core 11 includes, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, a combustion apparatus 16, a high pressure turbine 17, a low pressure turbine 19, and a core exhaust nozzle 20. Nacelle 21 surrounds gas turbine engine 10 and defines bypass duct 22 and bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

The bypass duct 22 comprises one or more outlet guide vanes 25.

In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is led into a combustion device 16, where the compressed air is mixed with fuel and the mixture is combusted. The resulting hot combustion products are then expanded through the high pressure turbine 17 and the low pressure turbine 19 before being discharged through the core exhaust nozzle 20, thereby driving the high pressure turbine 17 and the low pressure turbine 19 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by means of a suitable interconnecting shaft 27. The fan 23 typically provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement of the geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see fig. 1) drives a shaft 26, which shaft 26 is coupled to a sun gear or sun gear 28 of an epicyclic gear arrangement 30. Radially outward of and intermeshed with the sun gear 28 is a plurality of planet gears 32 that are coupled together by a carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously about the sun gear 28, while rotating each planet gear 32 about its own axis. The planet carrier 34 is coupled to the fan 23 via a connecting rod 36 for driving the fan in rotation about the engine axis 9. Radially outward of and intermeshes with the planet gears 32 is a ring gear or ring gear 38 that is coupled to the fixed support structure 24 via a linkage 40.

It is noted that the terms "low pressure turbine" and "low pressure compressor" as used herein may refer to the lowest pressure turbine stage and lowest pressure compressor stage, respectively (i.e., not including the fan 23), and/or the turbine and compressor stages that are connected together by an interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan 23). In some documents, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be referred to as an "intermediate pressure turbine" and an "intermediate pressure compressor". Where such alternative nomenclature is used, the fan 23 may be referred to as the first or lowest pressure compression stage.

The epicyclic gearbox 30 is shown in more detail in figure 3 by way of example. Each of the sun gear 28, planet gears 32, and ring gear 38 includes teeth around its periphery for intermeshing with other gears. However, for clarity, only exemplary portions of the teeth are shown in FIG. 3. Four planet gears 32 are shown, but it will be apparent to those skilled in the art that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of planetary epicyclic gearbox 30 typically include at least three planet gears 32.

The epicyclic gearbox 30 shown by way of example in fig. 2 and 3 is planetary, with a planet carrier 34 coupled to the output shaft via a connecting rod 36, with a ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of another example, the epicyclic gearbox 30 may be a sun arrangement in which the planet carrier 34 is held stationary, allowing the ring gear (or ring gear) 38 to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. By way of another alternative example, the gearbox 30 may be a differential gearbox in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

It should be understood that the arrangements shown in fig. 2 and 3 are exemplary only, and that various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement may be used to position the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of another example, the connections (such as the links 36, 40 in the example of FIG. 2) between the gearbox 30 and other components of the engine 10 (such as the input shaft 26, the output shaft, and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of another example, any suitable arrangement of bearings between rotating and stationary components of the engine (e.g., between input and output shafts from a gearbox and a stationary structure such as a gearbox housing) may be used, and the present disclosure is not limited to the exemplary arrangement of fig. 2. For example, where the gearbox 30 has a sun arrangement (as described above), the skilled person will readily appreciate that the arrangement of the output and support links and the bearing locations will generally differ from that shown by way of example in figure 2.

Accordingly, the present disclosure extends to gas turbine engines having any of a gearbox type (e.g., sun or planetary gear), support structure, input and output shaft arrangements, and bearing locations.

Optionally, the gearbox may drive additional and/or alternative components (e.g., a medium pressure compressor and/or a booster compressor).

Other gas turbine engines to which the present disclosure is applicable may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has split nozzles 18, 20, which means that the flow through the bypass duct 22 has its own nozzle 18 that is separate from and radially outside of the core exhaust nozzle 20. However, this is not limiting and any aspect of the present disclosure may also be applied to engines in which the flow through the bypass duct 22 and the flow through the engine core 11 are mixed or combined before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split) may have a fixed or variable area. Although the described examples relate to turbofan engines, the present disclosure is applicable to any type of gas turbine engine, such as an open rotor (where the fan stages are not surrounded by a nacelle) or, for example, a turboprop engine. In some arrangements, the gas turbine engine 10 may not include the gearbox 30.

The geometry of the gas turbine engine 10 and its components are defined by conventional shafting, including axial (aligned with the axis of rotation 9), radial (in the direction from bottom to top in fig. 1), and circumferential (perpendicular to the page in the view of fig. 1). The axial, radial and circumferential directions are mutually perpendicular.

The following description relates to the bifurcation 42 of the gas turbine engine as shown in FIG. 4. The bifurcation 42 is located below the engine core 11 in fig. 4, while another bifurcation 43, shown in fig. 1 and 2, may be disposed above the engine core 11. Each bifurcation 42, 43 spans the bypass duct 22, for example, aft and/or downstream of the fan outlet guide vanes 25. That is, each bifurcation 42, 43 provides a structural connection between the engine core 11 and the nacelle 21, thereby bifurcating the annular flow path of the bypass duct 22.

Fig. 4 shows a gas turbine engine comprising a lower fork 42. The fork 42 is attached at its upper end to the engine core 11, for example to a housing structure or outer housing wall 50 of the engine core 11. The engine core casing typically has radially inner and outer walls to define an interior volume of the casing in which various engine accessories and lines are mounted. The upper end of the prong 42 abuts the outer housing wall 50 of the housing.

The crotch 42 is attached to the lowest point of the engine core 11, i.e., the lowest point of the outer housing wall 50 of the engine core 11. In this regard, the outer housing wall 50 is curved in profile, reaching its lowest point partway between the front and rear ends of the housing. That is, the housing bulges or widens part way along its length.

The outer housing wall 50 of the housing defines an inner annular wall of the bypass conduit 22. The bifurcation 42 extends between the outer housing wall 50 and the outer wall 52 of the bypass duct 22 for attachment to the nacelle 21. The height of the bifurcation 42 thus completely spans the bypass conduit 22.

The bifurcation may be described as being located/aligned at a bottom dead center of the gas turbine engine 10 or the engine core 11 of the gas turbine engine.

Additional bifurcations 43 may be provided but are not shown in fig. 4. The additional/upper branch 43 may be positioned diametrically opposite the lower branch 42. The further bifurcation 43 may be positioned between the engine core 11 and the tower (not shown) mounting the engine to the wing of the aircraft.

The profiled outer wall 44 defines the outer surface of the bifurcation 42, i.e. the surface which will be gas scrubbed by the bypass duct 22 in use. The outer wall 44 defines a fairing/enclosure extending between the engine core 11 and the inner and outer walls 50, 52 of the bypass duct 22.

The outer wall 44 of the bifurcation 42 defines a closed interior that houses and protects the components extending between the engine core 11 and the nacelle 21. These components may include, among others: a fuel line; a hydraulic line; a conduit; an electric wire; a communication line; and other components that support engine operation.

The bifurcation 42 provides additional structural rigidity, i.e., support struts, between the engine core 11 and the nacelle 21.

The bifurcated outer wall 44 may comprise a sheet metal material, a composite material (e.g., a composite laminate), or a plastic material (e.g., a molded piece). The sheet metal material may be a Ti CP sheet or a Ti/64 sheet material.

In the example of fig. 4-8, the bifurcation 42 includes a conduit 46 that extends the length of the bifurcation (i.e., between the engine core 11 and the nacelle 21). The conduit may be located within the outer wall 44 of the bifurcation, for example as an internal section of the bifurcated shell. Alternatively, the conduit 46 may be disposed on an exterior surface of the bifurcated outer wall 44. That is, the conduit may be formed as a bifurcated contiguous portion/section attached to the outer wall 44.

The conduit 46 extends along/through the bifurcation to provide a discharge path from the engine core 11 to the nacelle 21 (e.g., to a discharge point in the nacelle) to communicate with the ambient environment (i.e., outside of the engine power plant).

The conduit 46 includes an inlet 54 that extends/opens through the wall 50 into the interior of the engine core 11, for example into the engine core housing, to provide a fluid pathway from the engine core 11 into the conduit 46.

As shown in fig. 5, at least a portion of the inlet 54 is disposed on a lowermost portion 56 of the outer housing wall 50 of the engine core 11. The engine core 11 and/or the wall 50 includes a fluid trap 57 positioned adjacent a lowermost portion 56 of the engine core 11. A fluid trap 57 is fluidly connected to the inlet 54.

The fluid trap 57 may include a reservoir or basin-like structure in the outer housing wall 50, such as around the ends of the bifurcation and/or the inlet 54.

As shown in fig. 6, the diverging profile may widen toward the inlet end of the conduit 46 (i.e., adjacent the outer housing wall 50) to define a mouth or edge 63 on the interior of the outer housing wall 50. In this regard, the bifurcation may curve/taper toward the outer housing wall 50, i.e., to provide an outwardly curved edge that abuts the outer housing wall 50.

The conduit 46 includes an outlet 55 that is open at an end of the conduit 46 opposite the inlet. The outlet 55 is fluidly connected to the ambient environment via an exhaust port adjacent the outlet 55 (i.e., within the nacelle 21).

The conduit 46 may include an outwardly extending edge 66 at the inlet end, e.g., a mouth structure. The edge 66 may be shaped to correspond to the shape of the wall 50 and/or the edge 66 of the conduit 46, such as at the interface between the conduit 46 and the wall 50.

The conduit 46 comprises a refractory material. The conduit may comprise titanium. The conduit may comprise grade 1 titanium (commercially pure titanium). The conduit may comprise a material different from the material of the bifurcated outer wall 44. The conduit may be made of the same material as the housing wall 50.

Generally, the conduit will be made of a material that has been proven to be fire resistant by material selection. Examples are Ti ═ 0.45mm thick, aly (aluminum) ═ 6mm thick, and steel ═ 0.40mm thick.

The conduit 46 has a wall thickness. The thickness can be varied in conjunction with material selection to ensure that its fire specification is achieved. An example is a Ti CP conduit, which may have a wall thickness between 0.9mm and 2.0mm, for example with a practical value of 1.2mm or between 1.1mm and 1.6 mm.

The bifurcation 42 and/or the conduit 46 extends in a generally radial direction from the engine core 11. The bifurcation 42 and/or the conduit 46 may be angled relative to the engine core 11, i.e., obliquely angled relative to the engine axis 9. The bifurcation 42 and/or the conduit 46 may be angled (e.g., from its radially inner end to its outer end) toward the front of the gas turbine engine (i.e., toward the fan 23 or the engine inlet 12).

The conduit 46 may be parallel to the longitudinal axis of the bifurcation 42, or may be at a different angle to the longitudinal axis of the bifurcation 42.

As shown in fig. 6, the outer wall 44 of the bifurcation 42 defines an aerodynamic shape, i.e., a fairing. The cross-section of the housing may take the form of an airfoil or a tear drop. The bifurcation 42 includes a leading edge 58 and a trailing edge 60. Leading edge 58 is directed toward the front of the gas turbine engine, while trailing edge 60 is directed toward the rear of the gas turbine engine.

In the examples described herein, the conduit 46 is configured to provide a structural portion of the bifurcation 42 and/or its outer wall 44. The bifurcation may have a strength, stiffness, or hardness equivalent to that of the outer wall 44.

Additionally or alternatively, the conduit may be shaped to define a bifurcated partition, thereby isolating the bifurcated first interior volume from further interior volumes of the partition within the conduit. The conduit may comprise a wall or wall portion extending between (e.g. in cross-section) the diverging opposing side walls. In the example of fig. 6, the opposing sidewalls of the bifurcation 42 each extend from the leading edge 58 to the trailing edge 50, i.e., such that the bifurcation has a longitudinal axis in cross-section. The walls of the conduit may extend completely between the side walls so as to provide a reinforcing partition across the bifurcation. The walls of the conduit may be oriented transverse/perpendicular to the longitudinal axis of the bifurcation and/or the engine.

The conduit 46 is configured to conform to and/or form a portion of the cross-sectional profile of the bifurcation 42. In the example shown in fig. 6, the conduit 46 is located in the bifurcated trailing edge 60. Thus, the conduit 46 adopts a tapered or substantially triangular cross-sectional shape so as to conform to the cross-sectional shape of the inner surface of the outer wall 44. It will be appreciated that the catheter 46 may have any cross-sectional shape that conforms to the contours of the bifurcation, depending on where the catheter 46 is positioned within the bifurcation 42. For example, if the conduit 46 is located at the leading edge 58, the conduit 46 will assume a more spherical cross-sectional shape (i.e., generally bullet-shaped) of the inner surface of the housing at the leading edge 58.

In the example of fig. 6, the wall 47 of the conduit defines a dividing wall, separating the interior of the conduit from the remainder of the bifurcated interior.

The cross-sectional area of the conduit 46 is adapted to ensure that the outlet/exhaust port (i.e., the flame-protected area) of the engine housing can be vented of possible leakage fluid for a period of time sufficient to meet conventional regulations. This can be calculated by calculating the maximum possible fluid leakage rate in the zone and in the worst case from the minimum pressure difference between the interior of the core zone/shell and the ambient pressure. In this example, the cross-sectional area of the conduit may be greater than or equal to 15, 20, or 25 square inches (96.8, 129.0, or 161.3 square centimeters, respectively).

The cross-sectional area of the conduit 46 may be greater than or equal to, for example, 0.06, 0.08, 0.10, 0.12, 0.14, 0.16, 0.18, or 0.2 times the cross-sectional area of the bifurcation. However, since the bifurcation size is affected by the structural requirements and service wiring through the bifurcation, the flow requirements of the cross-sectional area of the conduit are ultimately generally independent of the area of the bifurcation.

As shown in fig. 7, the conduit 46 includes at least one drain 62. The exhaust duct 62 may be connected to specific portions of the gas turbine engine (e.g., within the engine core 11) that have specific exhaust requirements. The drain 62 may drain the fluid to the surrounding environment at the nacelle and/or the drain 65. The discharge tube 62 is mounted within the conduit 46 using one or more mounts 64, such as a plurality of mounts 64 spaced along the length of the conduit. Each of the mounts 64 may support a plurality of discharge tubes 62 that extend through the conduit. Thus, the conduit may provide a common path for multiple exhaust pipes and fluid collection/pooling in the engine core casing itself.

As shown in FIG. 8, a gas turbine engine of the type described herein is shown with the contours of its various/separate regions. The engine comprises a first zone 67 and a second zone 68. The engine may also include a third zone 70.

The first zone 67 comprises a forward region of the nacelle 21, referred to as the fan zone, which is not a fire protected zone. Components located within the first zone 67 may place less stringent requirements on the refractory design.

The second zone 68 includes the main/rearward portion of the engine core 11, i.e., the engine core housing interior. The second region 68 is a flame-protected region. Thus, components located within the second zone 68 have more stringent requirements for refractory design.

It can be seen that the conduit 46 is provided with a second region 68, i.e. in communication with the second region 68 to allow fluid to be discharged therefrom. Thus, the interior of the conduit 46 places more stringent requirements on the refractory design.

However, the remainder of the bifurcation 42 (i.e., the remainder of the interior of the wall 44 outside of the conduit 46) is isolated from the flame-resistant region 68 by the conduit wall and is provided with a first region 67. Thus, the housing may have less stringent requirements for a fire resistant design.

During operation, fluid may leak into the engine core 11 (e.g., due to a fuel tube burst). Under the force of gravity, the fluid flows toward the lowermost portion 56 of the engine core 11 and into a fluid trap 57. The fluid then enters the conduit 46 via the inlet 54 and exits through the conduit towards the outlet 55. Finally, the fluid is discharged to ambient pressure (outside of the engine power plant) via an exhaust port.

Any fluid leakage in the engine core 11 may be drained through the conduit under the force of gravity. However, an actuator (e.g., a pump) may be provided in the engine core 11 or the conduit 46 to expedite the removal of fluid.

Thus, the conduit provides a venting feature to the fire protected area so that the area can be quickly emptied of leaking fluid (e.g., fuel) while isolating other areas from the fire risk.

As described elsewhere herein, the present disclosure may relate to a gas turbine engine. Such gas turbine engines may include an engine core including a turbine, a combustor, a compressor, and a spindle connecting the turbine to the compressor. Such gas turbine engines may include a fan (with fan blades) located upstream of the engine core.

The arrangement of the present disclosure may be particularly, but not exclusively, beneficial for fans driven via a gearbox. Accordingly, the gas turbine engine may include a gearbox that receives an input from the spindle and outputs a drive to the fan to drive the fan at a lower rotational speed than the spindle. The input to the gearbox may come directly from the spindle or indirectly from the spindle, for example via spur gear shafts and/or gears. The spindle may rigidly connect the turbine and compressor such that the turbine and 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 may have any desired number of shafts connecting the turbine and the compressor, such as one shaft, two shafts, or three shafts. By way of example only, the turbine connected to the spindle may be a first turbine, the compressor connected to the spindle may be a first compressor, and the spindle may be a first spindle. The engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor. The second turbine, the second compressor and the second spindle may be arranged to rotate at a higher rotational speed than the first spindle.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive the flow from the first compressor (e.g. directly, e.g. via a substantially annular duct).

The gearbox may be arranged to be driven by a spindle (e.g. the first spindle in the above example) which is configured (e.g. in use) to rotate at the lowest rotational speed. For example, the gearbox may be arranged to be driven only by the spindles (e.g. only the first spindle, not the second spindle in the above example) that are configured to rotate (e.g. in use) at the lowest rotational speed. Alternatively, the gearbox may be arranged to be driven by any one or more shafts, such as the first shaft and/or the second shaft in the above examples.

The gearbox is a reduction gearbox (since the fan output is at a lower rate than the input from the spindle). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "sun" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range 3 to 4.2, for example around 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 may be between any two values in the preceding sentence. Higher gear ratios may be more suitable for "planetary" type gearboxes. In some arrangements, the gear ratio may be outside of these ranges.

In any gas turbine engine as described and/or claimed herein, the combustor may be disposed axially downstream of the fan and the one or more compressors. For example, where a second compressor is provided, the combustor may be located directly downstream of (e.g., at the outlet of) the second compressor. By way of another example, where a second turbine is provided, the flow at the combustor outlet may be provided to the inlet of the second turbine. The combustor may be disposed upstream of one or more turbines.

The or each compressor (e.g. the first and second compressors as described above) may comprise any number of stages, for example a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes, which may be variable stator vanes (as the angle of incidence of the row of stator vanes may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (e.g. the first and second turbines as described above) may comprise any number of stages, for example a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas wash position or 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or about) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26 or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be within a range of inclusion defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit). These ratios may be generally referred to as hub-to-tip ratios. Both the radius at the hub and the radius at the tip may be measured at the leading (or axially forwardmost) portion of the blade. Of course, the hub-to-tip ratio refers to the gas scrubbing portion of the fan blade, i.e., the portion radially outside of any platform.

The radius of the fan may be measured between the engine centerline and the tip at the leading edge of the fan blade. The fan diameter (which may be only twice the fan radius) may be greater than (or about) any one of: 250cm (about 100 inches), 260cm, 270cm (about 105 inches), 280cm (about 110 inches), 290cm (about 115 inches), 300cm (about 120 inches), 310cm, 320cm (about 125 inches), 330cm (about 130 inches), 340cm (about 135 inches), 350cm, 360cm (about 140 inches), 370cm (about 145 inches), 380cm (about 150 inches), or 390cm (about 155 inches). The fan diameter may be within the inclusive range defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit).

The rotational speed of the fan may vary during use. Generally, for fans having larger diameters, the rotational speed is lower. By way of non-limiting example only, the rotational speed of the fan at cruise conditions may be less than 2500rpm, such as less than 2300 rpm. By way of further non-limiting example only, for an engine with a fan diameter in the range of 250cm to 300cm (e.g. 250cm to 280cm), the rotational speed of the fan at cruise conditions may be in the range of 1700rpm to 2500rpm, such as in the range of 1800rpm to 2300rpm, such as in the range of 1900rpm to 2100 rpm. By way of further non-limiting example only, for an engine with a fan diameter in the range of 320cm to 380cm, the rotational speed of the fan at cruise conditions may be in the range of 1200rpm to 2000rpm, such as in the range of 1300rpm to 1800rpm, such as in the range of 1400rpm to 1600 rpm.

In use of the gas turbine engine, a fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tips of the fan blades to rotate at a speed U_{Tip end}And (4) moving. The work done by the fan blades 13 on the flow results in an enthalpy rise dH for the flow. Fan tip load may be defined as dH/U_{Tip end} ²Where dH is the enthalpy rise across the fan (e.g., 1-D average enthalpy rise), and U_{Tip end}Is the (translational) speed of the fan tip, e.g. at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular speed). The fan tip load at cruise conditions may be greater than (or about) any one of: 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph are Jkg)^-1K^-1/(ms^-1)²). The fan tip load may be within the inclusive range defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit)。

A gas turbine engine according to the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of flow through the bypass duct to the mass flow rate of flow through the core at cruise conditions. In some arrangements, the bypass ratio may be greater than (or about) any one of: 10. 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5 or 17. The bypass ratio may be within the inclusive range defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit). The bypass conduit may be substantially annular. The bypass duct may be located radially outward of the engine core. The radially outer surface of the bypass duct may be defined by the nacelle and/or the fan casing.

The overall pressure ratio of the gas turbine engine described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the highest pressure compressor outlet (before entering the combustor). By way of non-limiting example, the overall pressure ratio at cruise of a gas turbine engine as described and/or claimed herein may be greater than (or about) any one of: 35. 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within the inclusive range defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit).

The specific thrust of the engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of the engines described and/or claimed herein may be less than (or about) any of the following: 110Nkg^-1s、105Nkg^-1s、100Nkg^-1s、95Nkg^-1s、90Nkg^-1s、85Nkg^-1s or 80Nkg^-1And s. The specific thrust force may be within the inclusive range defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit). Such engines may be particularly efficient compared to conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described and/or claimed herein may produce a maximum thrust of at least (or about) any of: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN or 550 kN. The maximum thrust may be within the inclusive range defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions, at sea level, plus 15 ℃ (ambient pressure 101.3kPa, temperature 30 ℃), with the engine at rest.

In use, the temperature of the flow at the inlet of the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the outlet of the combustor, for example immediately upstream of a first turbine vane, which may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or about) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be within an inclusive range defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit). The maximum TET of the engine in use may be, for example, at least (or about) any of: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be within an inclusive range bounded by any two values in the preceding sentence (i.e., the values may form an upper or lower limit). The maximum TET may occur, for example, under high thrust conditions, such as Maximum Takeoff (MTO) conditions.

The fan blades and/or airfoil portions of fan blades described and/or claimed herein may be fabricated from any suitable material or combination of materials. For example, at least a portion of the fan blade and/or airfoil may be at least partially fabricated from a composite material, such as a metal matrix composite material and/or an organic matrix composite material, such as carbon fiber. By way of further example, at least a portion of the fan blade and/or airfoil may be fabricated at least partially from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum lithium alloy) or a steel-based material. The fan blade may include at least two regions fabricated using different materials. For example, a fan blade may have a protective leading edge that may be manufactured using a material that is better resistant to impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge may be manufactured, for example, using titanium or a titanium-based alloy. Thus, by way of example only, the fan blade may have a carbon fiber or have an aluminum-based body with a titanium leading edge (such as an aluminum lithium alloy).

A fan as described and/or claimed herein may include a central portion from which fan blades may extend, for example, in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fastener that may engage a corresponding slot in the hub (or disk). By way of example only, such fasteners may be of dovetail form that may be inserted into and/or engage corresponding slots in the hub/disk to secure the fan blade to the hub/disk. By way of further example, the fan blade may be integrally formed with the central portion. Such an arrangement may be referred to as a blisk or a blisk ring. Any suitable method may be used to manufacture such a blisk or blisk. For example, at least a portion of the fan blade may be machined from a block and/or at least a portion of the fan blade may be attached to the hub/disk by welding (such as linear friction welding).

The gas turbine engines described and/or claimed herein may or may not be provided with Variable Area Nozzles (VANs). Such variable area nozzles may allow the outlet area of the bypass duct to vary in use. The general principles of the present disclosure may be applied to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24, or 26 fan blades.

As used herein, cruise conditions may refer to cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may generally be defined as conditions of intermediate cruise, such as conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between the climb apex and the descent start point.

By way of example only, the forward speed at cruise conditions may be at any point in the range from mach 0.7 to mach 0.9, such as 0.75 to 0.85, such as 0.76 to 0.84, such as 0.77 to 0.83, such as 0.78 to 0.82, such as 0.79 to 0.81, such as about mach 0.8, about mach 0.85, or 0.8 to 0.85. Any single speed within these ranges may be a cruise condition. For some aircraft, cruise conditions may be outside of these ranges, such as below mach 0.7 or above mach 0.9.

By way of example only, the cruise conditions may correspond to standard atmospheric conditions at an altitude within the following ranges: 10000m to 15000m, for example in the range 10000m to 12000m, for example in the range 10400m to 11600m (about 38000 feet), for example in the range 10500m to 11500m, for example in the range 10600m to 11400m, for example in the range 10700m (about 35000 feet) to 11300m, for example in the range 10800m to 11200m, for example in the range 10900m to 11100m, for example about 11000 m. Cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

By way of example only, the cruise conditions may correspond to: forward mach number 0.8; a pressure of 23000 Pa; and a temperature of-55 ℃.

As used anywhere herein, "cruise" or "cruise conditions" may refer to aerodynamic design points. Such aerodynamic design points (or ADPs) may correspond to conditions under which the fan is designed to operate (including, for example, one or more of mach number, ambient conditions, and thrust requirements). For example, this may refer to the condition where the fan (or gas turbine engine) is designed to have optimal efficiency.

In use, the gas turbine engine described and/or claimed herein may be operated at cruise conditions as defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., intermediate cruise conditions) of an aircraft on which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

The gas turbine engine and/or bifurcated structure of the present disclosure may provide various advantages.

The present disclosure allows fluids, such as fuel, oil, or water, to be discharged from the core.

The present disclosure allows fluid to be expelled under the influence of gravity and/or a pressure differential between the interior of the core region and the environment.

The present disclosure provides a refractory pathway in which flammable fluids may be exhausted from an engine core.

The present disclosure allows fluid to be expelled within a predetermined period of time.

The present disclosure allows for efficient packaging of bifurcations.

The present disclosure may provide support for an exhaust tube within a conduit.

The present disclosure provides increased geometric stiffness and structural rigidity for the bifurcation.

The present disclosure reduces the need for an entire bifurcation or other joining area requiring fire protection, resulting in weight and cost benefits.

The skilled person will appreciate that features or parameters described in relation to any of the above aspects or embodiments may be applied to or combined with any other aspect or embodiment, unless mutually exclusive.

It is to be understood that the present invention is not limited to the above-described embodiments, and various modifications and improvements may be made without departing from the concept described herein. Any feature may be used alone or in combination with any other feature or features unless mutually exclusive, and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein.