Structural assembly for a gas turbine engine

Abstract

Structural assembly which includes: - a bearing (5) comprising a statically arranged outer ring (6) and a rotatably arranged inner ring (8), wherein the inner ring (8) is non-rotatably connected to a component (34) rotatable about a longitudinal axis (9) or forms a component of such a component (34), wherein the longitudinal axis (9) defines an axial direction of the bearing, - a housing flange (450) of a support structure (45) to which the statically arranged outer ring (6) is connected, - wherein the outer ring (6) is formed in two parts, each part (61, 62; 63, 64) of the outer ring having a connecting element (612, 622; 632, 642) which is connected to the housing flange (450), wherein the housing flange (450) is arranged axially between the two connecting elements (612, 622; 632, 642), characterized in that that the outer ring (6) is radially divided, wherein the outer ring (6) has a radially inner ring element (61) and a radially outer ring element (62), each forming one of the connecting elements (612, 622), wherein the radially inner ring element (61) forms a running surface of the bearing (5), the radially outer ring element (62) adjoins the radially inner ring element (61) on the outside, and the two ring elements (61, 62) have axial areas (611, 621) in which they overlap in the axial direction, or that the outer ring (6) is axially split, wherein the outer ring (6) has an axially front ring element (63) and an axially rear ring element (64), each of which forms one of the connecting elements (632, 642), wherein one of the ring elements (64) forms a running surface of the bearing (5) and the other ring element (63) is connected to the one ring element (64) at a distance from the running surface.

Description

[0001] The invention relates to a structural assembly for a gas turbine engine according to the preamble of patent claim 1.

[0002] Bearings are known in which a statically arranged outer ring of the bearing forms a flange which is connected to a housing flange of a supporting structure.

[0003] US 4,473,309 A discloses an assembly with a plain bearing comprising a static outer part and a rotatable inner part. The outer part is supported by a housing, the outer part being made of steel and the housing being made of a material with a higher coefficient of thermal expansion than steel.

[0004] US Pat. No. 4,386,812 A discloses a ball bearing arrangement in which an outer ring of the ball bearing is pushed against a bearing seat to position and lock the ball bearing to a shaft shoulder. A T-piece engaging the outer ring is slidable via a thread to press the outer ring against the bearing seat.

[0005] EP 3 267 089 A1 discloses an oil distribution system for components in a housing. It is provided that a ball bearing is supplied via the oil distribution system, with the outer ring of the ball bearing being connected to a housing structure.

[0006] DE 10 2017 106 664 A1 discloses a jet engine having a chamber defined by a housing, in which a rotatable device which can be supplied with hydraulic fluid is arranged, wherein hydraulic fluid introduced into the chamber through the housing and air can be discharged from the chamber.

[0007] A generic structural assembly is known from US 8 052 331 B2.

[0008] The invention is based on the object of providing a structural assembly with a bearing which enables favorable load transfer from the bearing to a support structure.

[0009] This object is achieved by a structural assembly having the features of patent claim 1. Embodiments of the invention are specified in the dependent claims.

[0010] The invention accordingly contemplates a structural assembly that can be a component of a gas turbine engine. The structural assembly comprises a bearing having a statically arranged outer ring and a rotatably arranged inner ring. The inner ring is connected in a rotationally fixed manner to a component rotatable about a longitudinal axis or forms a component of such a component. The longitudinal axis of the rotatable component defines an axial direction of the bearing. The bearing further comprises a housing flange of a support structure, wherein the outer ring is connected to the housing flange.

[0011] It is further provided that the outer ring is split either radially or axially. If it is split radially, it is further provided that the radially inner ring element forms a running surface of the bearing, the radially outer ring element borders radially on the outside of the radially inner ring element and the two ring elements have axial regions in which they overlap in the axial direction. If the outer ring is split axially, the outer ring forms an axially front ring element and an axially rear ring element, each of which each forms one of the connecting elements. In this case, it is further provided that one of the ring elements forms a running surface of the bearing and the other ring element is connected to one of the ring elements at a distance from the running surface.

[0012] In such a structural assembly, the invention provides that the outer ring is formed in two parts, wherein each part of the outer ring has a connecting element which is connected to the housing flange, wherein the housing flange is arranged in the axial direction between the two connecting elements.

[0013] The invention is based on the idea of ​​providing the most uniform, symmetrical load transfer possible from the bearing to the support structure, largely eliminating bending moments. To this end, the housing flange is arranged axially centrally between two connecting elements formed by two parts of the outer ring. The symmetrical arrangement of the housing flange between the two connecting elements enables uniform load transfer to the support structure and prevents the generation and transmission of bending moments that would occur with a one-sided connection between the outer ring and the housing flange.

[0014] Ideally, the center of mass of the connection between the two connecting elements and the housing flange lies in a plane perpendicular to the longitudinal axis, which also contains the center of mass of the bearing. This prevents any bending moments during load transfer from the bearing to the supporting structure.

[0015] A further advantage associated with the invention is that, due to the fact that a connecting element of the outer ring is attached to both sides of the housing flange, the contact surface available for load transmission by friction to the housing flange is increased.

[0016] The support structure can be any stationary structure capable of supporting loads. For example, it might be a stationary support structure of a gas turbine engine connected to an engine mount. In another example, the support structure might be a stationary gearbox housing.

[0017] The rotatable component to which the inner ring is connected or on which the inner ring is formed is, for example, a shaft or a rotating element of a transmission. One embodiment of this can provide for the rotating element to be formed by a planetary carrier of a planetary transmission or to be connected to such a carrier.

[0018] According to one embodiment of the invention, the connecting elements of the two parts of the outer ring are also each designed as a flange and are flanged to the housing flange. This provides a particularly effective and large-area connection between the housing flange and the connecting elements.

[0019] If the outer ring is radially divided, it can further be provided that the connecting elements extend radially outwards from the respective axial regions of the ring elements on axially opposite sides of the outer ring.

[0020] With a radial division of the outer ring, the radially inner ring element and the radially outer ring element overlap each other in the axial direction. One embodiment provides for the radially outer ring element and the radially inner ring element to be connected to each other in a rotationally fixed manner via a press fit.

[0021] In the case of an axial division of the outer ring, it can be provided that the connecting elements extend radially outwards on axially opposite sides of the outer ring from axial regions of the axially front ring element and the axially rear ring element.

[0022] The axially front ring element and the axially rear ring element are connected to each other in a rotationally fixed manner, for example via bolts or the like.

[0023] The bearing is, for example, a rolling bearing, with rolling elements arranged between the inner ring and the outer ring. However, the invention is not limited to this, and the bearing can alternatively be designed, for example, as a plain bearing, with an inner surface of the outer ring and an outer surface of the inner ring forming the plain bearing.

[0024] A further embodiment provides that the structural assembly comprises a planetary gear, with the inner ring being shrunk onto a planetary carrier of the planetary gear or an element coupled to the planetary carrier. The planetary carrier can be part of a geared turbofan gas turbine engine.

[0025] In a further aspect of the invention, the invention relates to a gas turbine engine with a structural assembly according to the invention. The gas turbine engine may comprise: - an engine core comprising a turbine, a compressor and a turbine shaft in the form of a hollow shaft connecting the turbine to the compressor; - a fan positioned upstream of the engine core, the fan comprising a plurality of fan blades; and - a planetary gear mechanism, the input of which is connected to the turbine shaft and the output of which is connected to the fan shaft, the planetary gear mechanism comprising a planet carrier which is supported on a support structure of the gas turbine engine via a bearing according to claim 1.

[0026] A design for this can provide that - the turbine is a first turbine, the compressor is a first compressor and the turbine shaft is a first turbine shaft; - the engine core further comprises a second turbine, a second compressor, and a second turbine shaft connecting the second turbine to the second compressor; and - the second turbine, the second compressor and the second turbine shaft are arranged to rotate at a higher speed than the first turbine shaft.

[0027] It should be noted that the present invention is described with reference to a cylindrical coordinate system having the coordinates x, r, and φ. Where x indicates the axial direction, r the radial direction, and φ the angle in the circumferential direction. The axial direction is identical to the machine axis of a gas turbine engine in which the structural assembly is arranged. Starting from the x-axis, the radial direction points radially outward. Terms such as "in front of," "behind," "front," and "rear" refer to the axial direction and the flow direction in the engine, respectively. Terms such as "outer" or "inner" refer to the radial direction.

[0028] As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may include an engine core including a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (with fan blades) positioned upstream of the engine core.

[0029] Arrangements of the present disclosure may be particularly, but not exclusively, advantageous for fans driven by a gearbox. Accordingly, the gas turbine engine may include a gearbox that receives input from the core shaft and outputs drive to the fan for driving the fan at a lower speed than the core shaft. The input to the gearbox may be directly from the core shaft or indirectly from the core shaft, for example, via a spur shaft and / or spur gear. The core shaft may be rigidly connected to the turbine and compressor so that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

[0030] The gas turbine engine 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 turbines and compressors, such as one, two, or three shafts. For example 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 may be arranged to rotate at a higher speed than the first core shaft.

[0031] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive flow from the first compressor (e.g., directly, e.g., via a generally annular channel).

[0032] The gearbox may be arranged to be driven by the core shaft configured to rotate (e.g., in use) at the lowest 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 configured to rotate (e.g., in use) at the lowest speed (e.g., only by the first core shaft and not the second core shaft in the example above). Alternatively, the gearbox may be arranged to be driven by one or more shafts, e.g., the first and / or second shaft in the example above.

[0033] In a gas turbine engine described and / or claimed herein, a combustion chamber may be provided axially downstream of the fan and the compressor(s). For example, the combustion chamber may be located directly downstream of the second compressor (e.g., at its outlet) if a second compressor is provided. As a further example, the flow at the outlet of the compressor may be fed to the inlet of the second turbine if a second turbine is provided. The combustion chamber may be provided upstream of the turbine(s).

[0034] The or each compressor (e.g., the first compressor and the second compressor as described above) may comprise any number of stages, for example, multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (in that their angle of attack may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other.

[0035] The or each turbine (e.g., the first turbine and the second turbine as described above) may comprise any number of stages, for example, multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades. The row of rotor blades and the row of stator blades may be axially offset from each other.

[0036] Each fan blade may be defined with a radial span extending from a root (or hub) at a radially inward gas-overflow location or at a 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 on the order 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 an inclusive range bounded by two of the values ​​in the previous sentence (i.e., the values ​​may form upper or lower limits). These ratios may be collectively referred to as the hub-to-tip ratio.The hub radius and the tip radius can both be measured at the leading edge portion (or the axially forwardmost edge) of the blade. The hub-to-tip ratio, of course, refers to the gas-overflow portion of the fan blade, i.e., the portion located radially outside any platform.

[0037] The radius of the fan can be measured between the centerline of the engine and the tip of the fan blade at its leading edge. The diameter of the fan (which may simply be twice the radius of the fan) may be greater than (or on the order of): 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) or 390 cm (about 155 inches). The fan diameter may be within an inclusive range bounded by two of the values ​​in the previous sentence (i.e., the values ​​may be upper or lower limits).

[0038] The speed of the fan may vary during use. Generally, the speed is lower for fans with a larger diameter. By way of non-limiting example only, the speed of the fan under constant speed conditions may be less than 2500 rpm, for example, less than 2300 rpm. By way of another non-limiting example only, the speed of the fan under constant speed conditions for an engine with a fan diameter in the range of 250 cm to 300 cm (for example, 250 cm to 280 cm) may also be in the range of 1700 rpm to 2500 rpm, for example, in the range of 1800 rpm to 2300 rpm, for example, in the range of 1900 rpm to 2100 rpm.Merely as a further non-limiting example, the speed of the fan at constant speed conditions for an engine having a fan diameter in the range of 320 cm to 380 cm may be in the range of 1200 rpm to 2000 rpm, for example in the range of 1300 rpm to 1800 rpm, for example in the range of 1400 rpm to 1600 rpm.

[0039] During operation of the gas turbine engine, the fan (with associated fan blades) rotates around a rotational axis. This rotation causes the tip of the fan blade to rotate at a speed U Spitze The work performed by the fan blades on the flow results in an increase in the enthalpy dH of the flow. A fan peak load can be expressed as dH / U Spitze 2 where dH is the enthalpy rise (e.g. the average 1-D enthalpy rise) across the fan and U Spitzeis the (translational) speed of the fan tip, for example at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip loading at constant speed conditions may be more than (or of the order of): 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (where all units in this section are Jkg -1 K -1 / (ms -1 ) 1 The peak fan load may be within an inclusive range bounded by two of the values ​​in the previous sentence (ie, the values ​​may be upper or lower limits).

[0040] Gas turbine engines 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 through the bypass duct to the mass flow rate through the core at constant speed conditions. In some arrangements, the bypass ratio may be more than (on the order 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 an inclusive range bounded by two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits). The bypass duct may be substantially annular. The bypass duct may be located radially outward of the engine core. The radially outer surface of the bypass duct can be defined by an engine nacelle and / or a fan casing.

[0041] The total pressure ratio of a gas turbine engine described and / or claimed herein may be defined as the ratio of the back pressure upstream of the fan to the back pressure at the exit of the ultrahigh-pressure compressor (before entering the combustion chamber). As a non-limiting example, the total pressure ratio of a gas turbine engine described and / or claimed herein at constant speed may be more than (or on the order of): 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio may be within an inclusive range bounded by two of the values ​​in the preceding sentence (i.e., the values ​​may be upper or lower limits).

[0042] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow through the engine. At constant speed conditions, the specific thrust of an engine described and / or claimed here may be less than (or of the order of): 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg -1 s or 80 Nkg -1 s. The specific thrust may be within a range limited by two of the values ​​in the previous sentence (i.e., the values ​​may form upper or lower limits). Such engines can be particularly efficient compared to conventional gas turbine engines.

[0043] A gas turbine engine described and / or claimed herein may have any desired maximum thrust. By way of non-limiting example only, a gas turbine described and / or claimed herein may be capable of producing a maximum thrust of at least (or on the order of): 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range bounded by two of the values ​​in the preceding sentence (i.e., the values ​​may be upper or lower limits). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C) with the engine running statically.

[0044] In use, the flow temperature at the inlet of the high-pressure turbine can be particularly high. This temperature, which may be referred to as the TET, can be measured at the exit to the combustion chamber, for example immediately upstream of the first turbine blade, which in turn may be referred to as a nozzle guide vane. At constant speed, the TET may be at least (or of the order of): 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at constant speed may lie within an inclusive range bounded by two of the values ​​in the previous sentence (i.e., the values ​​may form upper or lower limits). The maximum TET in use of the engine may, for example, be at least (or of the order of): 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may lie within an inclusive range bounded by two of the values ​​in the previous sentence (ieThe values ​​can represent upper or lower limits. Maximum TET can occur, for example, under a high-thrust condition, such as an MTO (Maximum Take-Off Thrust) condition.

[0045] A fan blade and / or a blade portion of a fan blade described and / or claimed herein may be made from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or blade may be made at least in part from a composite material, such as a metal matrix composite and / or an organic matrix composite, such as carbon fiber. As another example, at least a portion of the fan blade and / or blade may be made at least in part 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 made using different materials.For example, the fan blade may include a protective leading edge constructed using a material that can withstand impacts (e.g., from birds, ice, or other material) better than the rest of the blade. Such a leading edge may, for example, be constructed using titanium or a titanium-based alloy. Thus, by way of example only, the fan blade may include a carbon fiber or aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0046] A fan described and / or claimed herein may include a central section from which the fan blades may extend, for example, in a radial direction. The fan blades may be attached to the central section in any desired manner. For example, each fan blade may include a locating device engageable with a corresponding slot in the hub (or disc). As one example only, such a locating device may be in the form of a dovetail that can be inserted into and / or engaged with a corresponding slot in the hub / disc to fix the fan blade to the hub / disc. As another example, the fan blades may be integrally formed with a central section. Such an arrangement may be referred to as a blisk or bling.Any suitable method may be used to manufacture such a blisk or bling. For example, at least a portion of the fan blades may be machined from a block and / or at least a portion of the fan blades may be attached to the hub / disk by welding, such as linear friction welding.

[0047] The gas turbine engines described and / or claimed herein may or may not be equipped with a VAN (Variable Area Nozzle). Such a variable area nozzle may allow variation of the bypass duct exit area during use. The general principles of the present disclosure may apply to engines with or without a VAN.

[0048] The fan of a gas turbine described and / or claimed herein may have any desired number of fan blades, for example 16, 18, 20 or 22 fan blades.

[0049] As used herein, constant speed conditions of an aircraft to which the gas turbine engine is attached may mean. Such constant speed conditions may conventionally be defined as the conditions during the middle part of the flight, for example, the conditions to which the aircraft and / or engine are subjected between (in time and / or distance) the end of the climb and the beginning of the descent.

[0050] By way of example only, the forward speed at the constant-speed condition may be at any point in the range of Mach 0.7 to 0.9, for example, 0.75 to 0.85, for example, 0.76 to 0.84, for example, 0.77 to 0.83, for example, 0.78 to 0.82, for example, 0.79 to 0.81, for example, on the order of Mach 0.8, on the order of Mach 0.85, or in the range of 0.8 to 0.85. Any speed within these ranges may be the constant-speed condition. For some aircraft, the constant-speed conditions may be outside these ranges, for example, below Mach 0.7 or above Mach 0.9.

[0051] By way of example only, the constant speed conditions may correspond to standard atmospheric conditions at an altitude which is in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (about 38,000 ft), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m (about 35,000 ft) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range of 10,900 m to 11,100 m, for example in the order of 11,000 m. The constant speed conditions may correspond to standard atmospheric conditions at any given altitude in these regions.

[0052] Just as an example, constant speed conditions might correspond to a forward Mach number of 0.8; a pressure of 23,000 Pa; and a temperature of -55 degrees C.

[0053] As used throughout this article, "constant speed" or "constant speed conditions" may refer to the aerodynamic design point (ADP). Such an aerodynamic design point (ADP) may correspond to the conditions (including, for example, Mach number, ambient conditions, and thrust requirement) for which the fan is designed to operate. This may, for example, refer to the conditions at which the fan (or gas turbine engine) is designed to achieve optimal efficiency.

[0054] In use, a gas turbine engine described and / or claimed herein may be operated at the constant speed conditions defined elsewhere herein. Such constant speed conditions may be determined by the constant speed conditions (e.g., conditions during the mid-flight) of an aircraft to which at least one (e.g., two or four) gas turbine engines may be attached to provide thrust.

[0055] It will be understood by those skilled in the art that a feature or parameter described in relation to one of the above aspects may be applied to any other aspect, provided they are not mutually exclusive. Furthermore, any feature or parameter described herein may be applied to any aspect and / or combined with any other feature or parameter described herein, provided they are not mutually exclusive.

[0056] The invention is explained in more detail below with reference to the figures of the drawing using several exemplary embodiments. They show: Fig. 1 is a side sectional view of a gas turbine engine; Fig. Figure 2 is a side sectional close-up view of an upstream section of a gas turbine engine; Fig. 3 is a partially cutaway view of a gearbox for a gas turbine engine; Fig. 4 shows an embodiment of a structural assembly comprising a bearing with a two-part outer ring connected to a housing flange of a support structure, wherein the outer ring is radially split; and Fig. 5 shows a further embodiment of a structural assembly comprising a bearing with a two-part outer ring connected to a housing flange of a support structure, the outer ring being axially split.

[0057] Fig. 1 illustrates a gas turbine engine 10 having a main rotational axis 9. The engine 10 includes an air inlet 12 and a thrust fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow order, a low-pressure compressor 14, a high-pressure compressor 15, a combustor 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core exhaust nozzle 20. An engine nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass air flow 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 epicycloidal gear 30.

[0058] In use, the core airstream A is accelerated and compressed by the low-pressure compressor 14 and passed into the high-pressure compressor 15 where further compression occurs. The compressed air exhausted from the high-pressure compressor 15 is passed into the combustor 16 where it is mixed with fuel and the mixture is combusted. The resulting hot combustion products then propagate through, and drive, the high-pressure and low-pressure turbines 17, 19 before being expelled through the nozzle 20 to provide some thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 through a suitable connecting shaft 27. The fan 23 generally provides the majority of the thrust. The epicycloidal gear 30 is a reduction gear.

[0059] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Fig. 2. The low-pressure turbine 19 (see Fig. 1) drives the shaft 26, which is coupled to a sun gear 28 of the epicycloidal gear assembly 30. A plurality of planet gears 32, coupled together by a planet carrier 34, are located radially outward from the sun gear 28 and mesh therewith. The planet carrier 34 restricts the planet gears 32 to synchronously orbiting the sun gear 28 while allowing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the fan 23 via linkages 36 to drive its rotation about the engine axis 9. An outer gear or ring gear 38, coupled to a stationary support structure 24 via linkages 40, is located radially outward from the planet gears 32 and meshes therewith.

[0060] It should be noted that the terms "low-pressure turbine" and "low-pressure compressor," as used herein, may be construed to mean the lowest-pressure turbine stage and the lowest-pressure compressor stage, respectively (i.e., not including the fan 23) and / or the turbine and compressor stages connected by the lowest-speed connecting shaft 26 in the engine (i.e., not including the transmission output shaft that drives the fan 23). In some writings, the "low-pressure turbine" and "low-pressure compressor" referred to herein may alternatively be known as the "intermediate-pressure turbine" and "intermediate-pressure compressor." When using such alternative nomenclature, the fan 23 may be referred to as a first compression stage or lowest-pressure compression stage.

[0061] The epicycloidal gear 30 is Fig. 3 is shown in more detail. The sun gear 28, the planet gears 32 and the ring gear 38 each include teeth around their periphery for meshing with the other gears. However, for the sake of clarity, only exemplary portions of the teeth are shown in Fig. 3. Although four planetary gears 32 are illustrated, it will be apparent to those skilled in the art that more or fewer planetary gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicycloidal gear 30 generally include at least three planetary gears 32.

[0062] The Fig. 2 and Fig. The epicycloidal gear train 30 illustrated by way of example in Figure 3 is a planetary gear train in which the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 being fixed. However, any other suitable type of epicycloidal gear train 30 may be used. As another example, the epicycloidal gear train 30 may be a star arrangement in which the planet carrier 34 is held fixed, while the ring gear (or external gear) 38 is allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the gear train 30 may be a differential gear train in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0063] It is understood that the Fig. 2 and Fig. 3 is merely exemplary and various alternatives are within the scope of the present disclosure. For example only, any suitable arrangement for positioning the gearbox 30 in the engine 10 and / or for connecting the gearbox 30 to the engine 10 may be used. As a further example, the connections (e.g., the linkages 36, 40 in the example of Fig. 2) between the transmission 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft, and the fixed structure 24) may have a certain degree of rigidity or flexibility. As another example, any suitable arrangement of bearings between rotating and stationary parts of the engine (for example, between the input and output shafts of the transmission and the fixed structures, such as the transmission housing) may be used, and the disclosure is not limited to the exemplary arrangement of Fig. 2. For example, it will be readily apparent to those skilled in the art that the arrangement of output and support rods and bearing positioning in a star arrangement (described above) of the gearbox 30 will generally differ from those shown, for example, in Fig. 2 would differ.

[0064] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gear types (e.g., star or planetary), support structures, input and output shaft arrangement, and bearing positioning.

[0065] Optionally, the gearbox can drive secondary and / or alternative components (e.g. the medium-pressure compressor and / or a booster compressor).

[0066] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative number of connecting shafts. As another example, the Fig. 1 has a split-flow nozzle 20, 22, meaning that the flow through the bypass duct 22 has its own nozzle separate from and radially outward from the engine core nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the 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-flow) may have a fixed or variable area. Although the described example relates to a turbofan engine, the disclosure may be applied to any type of gas turbine engine, such as an open-rotor (where the fan stage is not surrounded by an engine nacelle) or a turboprop engine.In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0067] The geometry of the gas turbine engine 10 and components thereof is or are defined by a conventional axis system having an axial direction (aligned with the rotational axis 9), a radial direction (in the direction from bottom to top in Fig. 1) and a circumferential direction (perpendicular to the view in Fig. 1). The axial, radial, and circumferential directions are perpendicular to each other.

[0068] In the Fig. 2 shows that the planet carrier 34 is supported on a bearing 5. The bearing 5 has an outer ring that is supported on a strut or wall structure 45, which is part of a supporting structure of the gas turbine engine. Within the scope of the present invention, the design of the bearing 5 is important. It should be noted that the following description of the bearing 5 is not limited to its use in a gas turbine engine in connection with the support of a planetary gear. The bearing 5 can be used in essentially any context.

[0069] According to the embodiment of the Fig. 4, the bearing 5 comprises a rotating inner ring 8 connected to a rotatable component, a static outer ring 6, and rolling elements 7 arranged therebetween, which are designed, for example, as balls or rollers. However, this design of the bearing 5 is to be understood merely as an example.

[0070] The outer ring 6 of the bearing 5 is connected to a support structure 45, which is part of a load-bearing support structure, for example, of a gas turbine engine. The support structure 45 can be, for example, a fastening arm, a strut, a wall, or a housing element. To connect the bearing 5 to the support structure 45, the support structure forms a flange 450 whose side surfaces extend in the radial direction. The flange 450 is referred to below as the housing flange, although, as explained, it is not necessarily part of a housing, but rather of a support structure in general.

[0071] The outer ring 6 is formed in two parts and comprises a radially inner ring element 61 and a radially outer ring element 62. The radially inner ring element 61 has an axial region 611 and a connecting element 612 that protrudes perpendicularly from the axial region 611. The radially outer ring element 62 also has an axial region 621 and a connecting element 622 that protrudes perpendicularly from the axial region 621.

[0072] The axial region 611 of the radially inner ring element 61 forms, with its inner surface, a running surface of the bearing 5. The axial region 621 of the radially outer ring element 62 borders radially on the outside of the axial region 611. Provision can be made for the axial region 611 and the axial region 621 of the two ring elements 61, 62 to be connected to one another in a rotationally fixed manner by a press fit. Alternatively, a rotationally fixed connection can be provided via form-locking elements (not shown), or a connection can be made only indirectly via the connection of the connecting elements 621, 622 to the housing flange.

[0073] The two connecting elements 612, 622 of the two ring elements 61, 62 extend radially outward on axially opposite sides of the outer ring 6. They each form a flange that is connected to the housing flange 450 via bolts 91. Thus, the connecting elements 612, 622 are arranged on both sides of the housing flange 450, and the housing flange 450 is thus arranged in the axial direction between the two connecting elements 612, 622. This ensures a uniform load transfer from the bearing 5 to the housing flange 450 and the support structure 45, without bending moments being introduced into the support structure 45.

[0074] It is intended that both the center of mass of the connection of the two connecting elements 612, 622 to the housing flange 450 and the center of mass of the bearing 5 lie in the same plane E. The plane E is perpendicular to the axial direction of the bearing 5, which is defined by the axis of rotation of a rotating element to which the inner ring 8 is connected. By arranging both centers of mass in the same plane, the introduction of bending moments or tilting moments into the support structure 45 is avoided. Such introduction of bending moments occurs, however, if only one connecting element is provided, which is flanged to the housing flange 450 on one side.

[0075] The Fig. 5 shows an embodiment of a bearing which differs from the embodiment of the Fig. 4 in that the outer ring 6 is axially divided and not radially divided as in the embodiment of the Fig. 4.

[0076] According to the embodiment of the Fig. 5, the bearing 5 comprises a rotating inner ring 8 connected to a rotatable component, a static outer ring 6, and rolling elements 7 arranged therebetween. The bearing 5 is connected to a support structure 45, which is part of a load-bearing support structure, for example, of a gas turbine engine. To connect the outer ring 6 of the bearing 5 to the support structure 45, the support structure forms a housing flange 450 whose side surfaces extend in the radial direction.

[0077] The outer ring 6 is formed in two parts and comprises an axially front ring element 63 and an axially rear ring element 64. The axially front ring element 63 comprises an axial region 631 and a connecting element 632, which protrudes perpendicularly from the axial region 631. The axially rear ring element 64 also has an axial region 641 and a connecting element 642, which protrudes perpendicularly from the axial region 641.

[0078] The axial region 641 of the axially rear ring element 64 forms a running surface of the bearing 5 on its inner side. The axially front ring element 63 is located axially in front of the ring element 64. The two ring elements 63, 64 are connected to one another via bolts 92.

[0079] The two connecting elements 632, 642 of the two ring elements 63, 64 extend on axially opposite sides of the outer ring 6. They form, as in the embodiment of the Fig. 4 each comprise a flange which is connected to the housing flange 450 via bolts 91, which are only shown schematically. The housing flange 450 is thus arranged in the axial direction between the two connecting elements 632, 642, so that a uniform load transfer from the bearing 5 to the housing flange 450 and the support structure 45 takes place, which is essentially free of bending moments. As in the exemplary embodiment of the Fig. 4, it should be provided that the center of gravity of the connection of the two connecting elements 632, 642 with the housing flange 450 and the center of gravity of the bearing 5 lie in the same plane.

[0080] It should be understood that the invention is not limited to the embodiments described above, and various modifications and improvements may be made without deviating from the concepts described herein. For example, the bearing may be designed as a plain bearing instead of a rolling bearing.

[0081] It is understood that any of the described features may be used separately or in combination with any other features, provided they are not mutually exclusive, and the disclosure extends to and encompasses all combinations and subcombinations of one or more features described herein. Where ranges are defined, these include all values ​​within these ranges, as well as all subranges that fall within a range.

Claims

[1] Structural assembly comprising: - a bearing (5) having a statically arranged outer ring (6) and a rotatably arranged inner ring (8), wherein the inner ring (8) is connected in a rotationally fixed manner to a component (34) rotatable about a longitudinal axis (9) or forms a component of such a component (34), wherein the longitudinal axis (9) defines an axial direction of the bearing, - a housing flange (450) of a support structure (45) to which the statically arranged outer ring (6) is connected, - wherein the outer ring (6) is formed in two parts, wherein each part (61, 62; 63, 64) of the outer ring has a connecting element (612, 622; 632, 642) which is connected to the housing flange (450), wherein the housing flange (450) is arranged in the axial direction between the two connecting elements (612, 622; 632, 642), characterized by , that the outer ring (6) is radially divided, wherein the outer ring (6) has a radially inner ring element (61) and a radially outer ring element (62), each forming one of the connecting elements (612, 622), wherein the radially inner ring element (61) forms a running surface of the bearing (5), the radially outer ring element (62) adjoins the radially inner ring element (61) radially on the outside, and the two ring elements (61, 62) have axial regions (611, 621) in which they overlap in the axial direction, or that the outer ring (6) is axially divided, wherein the outer ring (6) has an axially front ring element (63) and an axially rear ring element (64), each of which forms one of the connecting elements (632, 642), wherein one of the ring elements (64) forms a running surface of the bearing (5) and the other ring element (63) is connected to the one ring element (64) at a distance from the running surface. [2] Structural assembly according to claim 1, characterized by that the connecting elements (612, 622; 632, 642) of the two parts (61, 62; 63, 64) of the outer ring (6) are also each designed as a flange and are flanged to the housing flange (450). [3] Structural assembly according to claim 1, characterized by that, if the outer ring (6) is radially divided, the connecting elements (621, 622) extend radially outwards on axially opposite sides of the outer ring (6) from the respective axial regions (611, 621) of the ring elements (61, 62). [4] Structural assembly according to one of the preceding claims, characterized by that, if the outer ring (6) is radially divided, the radially outer ring element (62) and the radially inner ring element (61) are connected to one another in a rotationally fixed manner via a press fit. [5] Structural assembly according to claim 1, characterized bythat, if the outer ring (6) is axially divided, the connecting elements (632, 642) extend radially outwards on axially opposite sides of the outer ring (6) from axial regions (631, 632) of the axially front ring element (63) and the axially rear ring element (64). [6] Structural assembly according to claim 1 or claim 5, characterized by that, if the outer ring (6) is axially divided, the axially front ring element (63) and the axially rear ring element (64) are connected to one another in a rotationally fixed manner via bolts (92) or the like. [7] Structural assembly according to one of the preceding claims, characterized by that the center of mass of the connection between the two connecting elements (612, 622; 632, 642) and the housing flange (450) lies in a plane (E) in which the center of mass of the bearing (5) also lies. [8] Structural assembly according to one of the preceding claims, characterized bythat the bearing (5) is designed as a rolling bearing, with rolling elements (7) being arranged between the inner ring (8) and the outer ring (6). [9] Structural assembly according to one of the preceding claims, characterized by that the inner ring (8) is shrunk onto the planet carrier (34) of a planetary gear (30) or an element connected to such a planet carrier (34). [10] Gas turbine engine (10) for an aircraft, comprising: - an engine core (11) comprising a turbine (19), a compressor (14) and a turbine shaft (26) designed as a hollow shaft connecting the turbine to the compressor; - a fan (23) positioned upstream of the engine core (11), the fan (23) comprising a plurality of fan blades and being driven by a fan shaft; and - a planetary gear (30), the input of which is connected to the turbine shaft (26) and the output of which is connected to the fan shaft, the planetary gear comprising a planet carrier which is supported on a support structure (45) of the gas turbine engine (10) via a structural assembly according to claim 1.

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