DE102020122721A1 - gear part - Google Patents

Abstract

The invention relates to a gear part (50) gear part (50) of a planetary gear (30) which is connected to a planet carrier (34) for planet gears (32), the gear part (50) at least partially extending axially along the planet gears (32). and at least partially surrounds the planet wheels (32), characterized in that at least one oil guide means (61a, 61 b) is arranged on a surface of the transmission part (50), in particular on a surface (60) which faces the planet wheels (32). is, so that at least one oil flow (62) is led away from a lubricating point (32, 38, 38a).

Description

The present disclosure relates to a transmission part with the features of claim 1.

Planetary gears are used, for example, in motor vehicles or gas turbine engines. Due to the complex pairing of gears and planets in a planetary drive and the sometimes high mechanical loads, the lubrication of these planetary gears is a challenge. In the so-called fan gearboxes in particular, the planetary gearboxes have to work maintenance-free for a very long time.

The task is to develop lubricants for planetary gears that are particularly gentle on the oil.

According to a first aspect, the subject matter of claim 1 has a planetary gear with a gear part which is connected to a planet carrier for planetary gears and which at least partially has an axial extension along the planetary gears and at least partially surrounds the planetary gears. At least one oil guiding means is arranged on a surface of the gear part, in particular on a surface facing the planet gears (i.e. an inner side), so that at least one oil flow is guided away from a lubrication point. At the lubrication point, the oil would be subjected to further mechanical loads and would also heat up. The at least one oil guide means prevents this by guiding the oil in a targeted manner.

In one embodiment, at least one oil flow with an axial flow component can be guided to the edge of the planet carrier and/or can be guided with an axial flow component into a gap between two ring gears. In these cases, the oil flow is avoided in the lubricating point, e.g. in the toothing between ring gears and planetary gears.

The at least one oil flow (i.e. oil that moves e.g. along the transmission part) gets an axial directional component towards the edge of the planet carrier. In this way, the flow of oil is directed away from the engagements between the planet wheel and ring gear in particular and, for example, is diverted axially outwards through the at least one oil guide means, so that the oil flow can leave the transmission part. This prevents, for example, the oil from being stressed again by the ring gear on the outside of the gearbox.

In one embodiment, the at least one oil guiding means has an indentation or groove that guides the oil flow in the desired direction. Additionally or alternatively, in one embodiment, a wall can also be used whose elevation does not exceed the distance between the transmission part and the planet. In a further embodiment, which can be combined with one or both of the aforementioned embodiments, the at least one oil guide means has a three-dimensional contouring of the surface. In particular, the contouring can have a concave and/or convex area.

Furthermore, the at least one oil guide means can be arranged in the surface in the projection essentially between 30 and 50°, in particular 45°, perpendicular to the axis of rotation of the planet wheel. Efficient guidance of the oil flows is possible at these angles.

The at least one oil guide means can guide the at least one oil flow radially outwards to the edge of the planet carrier. A radial component of the removal is caused by the centrifugal force, with the oil guiding means providing an additional axial directional component or it influences the already existing (intensify or reverse) directional component.

The transmission part can be set up and designed to transmit torque from one side of the planetary carrier to the other.

In a further embodiment, the at least one flow of oil comes partially or entirely from a slide bearing of a planet wheel.

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine, such as an aircraft engine. Such a gas turbine engine may include a core engine that includes a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having fan blades) positioned upstream of the core engine.

Arrangements of the present disclosure may be particularly, but not exclusively, beneficial for geared fans driven via a gearbox (a planetary gearbox with a rotating planet carrier). Accordingly, the gas turbine engine may include a gearbox driven by the core shaft and the output of which drives the fan to rotate at a slower speed than the core shaft. The input for the gearbox can be direct from the core shaft or indirectly via the core shaft, for example via a spur shaft and/or a spur gear. The core shaft can be connected to the turbine and the compressor be rigidly connected so that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

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 spools connecting the turbine and compressor, such as one, two, or three spools. For example only, the turbine coupled to the core shaft may be a first turbine, the compressor coupled to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The core engine 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.

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 receiving, e.g. via a generally annular duct).

The gearbox may be configured 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 transmission may be configured to be driven only by the core shaft that is configured to rotate (e.g., in use) at the lowest speed (e.g., only 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, for example the first and/or the second shaft in the example above.

In a gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor (or compressors). For example, the burner device can be located directly downstream of the second compressor (e.g. at its outlet) if a second compressor is provided. As another example, if a second turbine is provided, the flow at the exit of the compressor may be directed to the inlet of the second turbine. The burner device may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and the second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may include a row of rotor blades and a row of stator blades, which may be variable (i.e., variable pitch) stator blades. The row of rotor blades and the row of stator blades may be axially offset from one another.

The or each turbine (such as the first turbine and the second turbine as described above) may include any number of stages, such as multiple stages. Each stage may include 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 one another.

Each fan blade may have a radial span extending from a root (or hub) at a radially inner gas flow location or extending from a 0% span position to a 100% span tip. 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 in a closed range bounded by two values in the previous sentence (i.e. the values may form an upper or lower limit). These ratios can be generically referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip can both be measured at the leading edge (or the axially most forward edge) of the blade. The hub-to-tip ratio is of course related to the portion of the fan blade over which the gas flows, i. H. the portion radially outward of any platform.

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 can generally be twice the radius of the fan) can be greater than (or on the order of): 250 cm (about 100 inches), 260 cm (about 103 inches), 270 cm (about 105 inches) , 280 cm (approx. 110 inches), 290 cm (approx. 115 inches), 300 cm (approx. 120 inches), 310 cm (approx 123 inches), 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm (about 139 inches), 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), or 390 cm (about 155 inches). The fan diameter can be in a closed range bounded by two of the values in the previous sentence (ie the values can form upper or lower limits).

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

In use of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tip of the fan blade to move at a speed U _tip . The work done by the fan blades on the flow results in an increase in the enthalpy dH of the flow. A fan tip load can be defined as dH/U _peak ² , where dH is the enthalpy rise (e.g. the average 1-D enthalpy rise) across the fan and U _peak is the (translational) velocity of the fan tip, e.g. at the leading edge of the tip , is (which can be defined as the leading edge fan tip radius times the angular velocity). Fan peak loading at constant speed conditions can be greater than (or on 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 ) ² ). The fan peak load can be in a closed range bounded by two of the values in the previous sentence (ie the values can form upper or lower bounds).

Gas turbine engines according to the present disclosure may have any desired bypass ratio, where 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 constant speed conditions. In some arrangements the bypass ratio can be more than (or 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 (lie). The bypass ratio can be in a closed range bounded by two of the values in the previous sentence (i.e. the values can form upper or lower limits). The bypass channel can be essentially ring-shaped. The bypass duct can be located radially outside of the core engine. The radially outer surface of the bypass duct may be defined by an engine nacelle and/or a fan casing.

The overall pressure ratio of a gas turbine engine, as described and/or claimed herein, may be defined as the ratio of the ram pressure upstream of the fan to the ram pressure at the exit of the super high pressure compressor (prior to the entrance to the combustor). As a non-limiting example, the overall pressure ratio of a gas turbine engine described and/or claimed herein at constant speed may be greater than (or on the order of): 35, 40, 45, 50, 55, 60, 65, 70, 75 (to lie). The overall pressure ratio can be in a closed range bounded by two of the values in the previous sentence (i.e. the values can form upper or lower limits).

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 herein 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 (lie). The specific thrust can be in a closed range bounded by two of the values in the previous sentence (ie the values can form upper or lower limits). Such engines can be particularly efficient compared to conventional gas turbine engines.

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 engine described herein and/or bean capable of generating a maximum thrust of at least (or of 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 550kN to be able to. The maximum thrust can be in a closed range bounded by two of the values in the previous sentence (ie the values can form upper or lower limits). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15°C (ambient pressure 101.3 kPa, temperature 30°C) with the engine static.

In use, the temperature of the flow at the entrance to the high pressure turbine can be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine blade, which in turn may be referred to as a nozzle vane. At constant speed the TET can be (are) at least (or in the order of): 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K. The TET at constant velocity can be in a closed range bounded by two of the values in the previous sentence (i.e. the values can form upper or lower bounds). For example, the maximum TET in use of the engine may be at least (or on the order of): 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, or 2000K. The maximum TET can be in a closed range bounded by two of the values in the previous sentence (i.e. the values can form upper or lower bounds). For example, the maximum TET may occur at a high thrust condition, such as an MTO (Maximum Take-Off Thrust) condition.

A fan blade and/or a fan blade aerofoil 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 formed at least in part from a composite such as a metal matrix composite and/or an organic matrix composite such as e.g. B. carbon fiber, are produced. As another example, at least a portion of the fan blade and/or blade may be formed at least in part from a metal such as aluminum. 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 sections made using different materials. For example, the fan blade may have a leading protective edge made using a material that can withstand impact (e.g., from birds, ice, or other material) better than the rest of the blade. Such a leading edge can be made, for example, using titanium or a titanium-based alloy. Thus, as just one example, the fan blade may have a carbon fiber or aluminum based body (such as an aluminum-lithium alloy) with a titanium leading edge.

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 can be attached to the center 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). By way of example only, such a fixation device may be in the form of a dovetail which may 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 center section. Such an arrangement may be referred to as a blisk or a 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 an ingot and/or at least a portion of the fan blades may be welded, such as by welding. B. linear friction welding, to be attached to the hub / disc.

The gas turbine engines described and/or claimed herein may or may not be provided with a VAN (Variable Area Nozzle). Such a nozzle with a variable cross-section can, during operation, allow the exit cross-section of the bypass channel to be varied. The general principles of the present disclosure may apply to engines with or without a VAN.

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

As used herein, constant speed conditions may mean the constant speed conditions of an aircraft on which the gas turbine engine is mounted mean. Such constant speed conditions may conventionally be defined as the conditions during the middle part of flight, for example the conditions experienced by the aircraft and/or engine between (in terms of time and/or distance) the end of the climb and the beginning of the descent. will.

By way of example only, the forward speed at the constant speed condition may be at any point in the range Mach 0.7 to 0.9, e.g. 0.75 to 0.85, e.g. 0.76 to 0.84, e.g. 0.77 to 0 .83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example of the order of Mach 0.8, of the order of Mach 0.85 or in the range of 0.8 to 0, 85 lie. Any speed within these ranges can be the cruise condition. For some aircraft, cruise conditions may be outside of these ranges, for example below Mach 0.7 or above Mach 0.9.

As an example only, the constant velocity conditions may correspond to standard atmospheric conditions at an altitude that is in the range 10,000 m to 15,000 m, for example in the range 10,000 m to 12,000 m, for example in the range 10,400 m to 11,600 m (about 38,000 feet), 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 feet) 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 of the order of 11,000 m. The constant velocity conditions can correspond to standard atmospheric conditions at any given altitude in these ranges.

As an example only, the constant speed conditions may correspond to: a forward Mach number of 0.8; a pressure of 23,000 Pa and a temperature of -55 °C.

As used throughout, "constant speed" or "constant speed conditions" can mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (including, for example, Mach number, environmental conditions, and thrust requirement) for which the fan operation is designed. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to be at its optimum efficiency.

In operation, 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 dictated by the constant speed conditions (e.g., mid-flight conditions) of an aircraft on which at least one (e.g., two or four) gas turbine engine(s) may be mounted to provide thrust.

It will be understood by those skilled in the art that a feature or parameter described in relation to one aspect above 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 compatible not mutually exclusive.

• 1 eine Seitenschnittansicht eines Gasturbinentriebwerks;
• 2 eine Seitenschnittgroßansicht eines stromaufwärtigen Abschnitts eines Gasturbinentriebwerks;
• 3 eine zum Teil weggeschnittene Ansicht eines Getriebes für ein Gasturbinentriebwerk;
• 4 eine Detailansicht einer Ausführungsform eines Getriebeteils;
• 5 eine Ansicht auf eine Innenfläche einer weiteren Ausführungsform eines Getriebeteils;
• 6 eine Ansicht auf eine Innenfläche einer weiteren Ausführungsform eines Getriebeteils;
• 7 eine Ansicht auf eine Innenfläche einer weiteren Ausführungsform eines Getriebeteils mit einer Ölführung in eine Lücke zwischen zwei Verzahnungen in den Hohlrädern;
• 8 eine Ansicht auf eine Innenfläche einer weiteren Ausführungsform eines Getriebeteils mit einem konkaven Ölführungsmittel;
• 9 eine Ansicht auf eine Innenfläche einer weiteren Ausführungsform eines Getriebeteils mit einem konvexe Ölführungsmittel.

Exemplary embodiments will now be described with reference to the figures; show in the figures:

• 1 Figure 12 is a side sectional view of a gas turbine engine;
• 2 Figure 12 is a close-up side sectional view of an upstream portion of a gas turbine engine;
• 3 Figure 12 is a partially cut-away view of a transmission for a gas turbine engine;
• 4 a detailed view of an embodiment of a transmission part;
• 5 a view of an inner surface of a further embodiment of a transmission part;
• 6 a view of an inner surface of a further embodiment of a transmission part;
• 7 a view of an inner surface of a further embodiment of a transmission part with an oil guide in a gap between two teeth in the ring gears;
• 8th a view of an inner surface of a further embodiment of a transmission part with a concave oil guide means;
• 9 a view of an inner surface of a further embodiment of a transmission part with a convex oil guide means.

Before embodiments of a transmission part 50 of a planetary gear 30 in connection with the 4 until 9 are shown, the technical context of a planetary gear 30 in a gas turbine engine 10 is described below.

1 1 shows a gas turbine engine 10 having a main axis of rotation 9. The engine 10 includes an air intake 12 and a fan 23 which produces two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 which receives the core airflow A. The core engine 11 includes, 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 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 planetary gear 30.

During operation, the core air flow A is accelerated and compressed by the low-pressure compressor 14 and conducted into the high-pressure compressor 15, where further compression takes place. The compressed air discharged from the high pressure compressor 15 is directed into the combustor 16 where it is mixed with fuel and the mixture is burned. The resultant hot products of combustion then propagate through and thereby drive the high 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 bulk of the thrust. The epicyclic planetary gear 30 is a reduction gear.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in 2 shown. The low-pressure turbine 19 (see 1 ) drives the shaft 26 which is coupled to a sun gear 28 of the epicyclic planetary gear set 30 . A plurality of planetary gears 32, which are coupled to one another by a planetary carrier 34, are located radially outward of the sun gear 28 and mesh therewith. The planetary carrier 34 guides the planetary gears 32 to orbit synchronously about the sun gear 28 while allowing each planetary gear 32 to rotate about its own axis. Planet carrier 34 is coupled to fan 23 via linkages 36 to drive its rotation about engine axis 9 . An outer gear or ring gear 38, which is coupled to a stationary support structure 24 via linkage 40, is radially outward of the planetary gears 32 and meshes therewith.

It is 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 interconnected by the lowest speed connecting shaft 26 in the engine (ie not comprising the gearbox output shaft driving 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." Using such alternative nomenclature, the fan 23 may be referred to as a first compression stage or lowest pressure compression stage.

The epicyclic planetary gear 30 is used in 3 shown in more detail as an example. The sun gear 28, planet gears 32 and ring gear 38 each include teeth on their periphery to allow meshing with the other gears. However, for the sake of clarity, only exemplary sections of the teeth are shown in 3 shown. Although four planetary gears 32 are illustrated, those skilled in the art will recognize that more or fewer planetary gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicyclic planetary gear 30 generally include at least three planetary gears 32.

This in 2 and 3 The epicyclic planetary gear set 30 exemplified is a planetary gear set in which the planet carrier 34 is coupled to an output shaft via linkages 36 with the ring gear 38 fixed. As another alternative example, the transmission 30 may be a differential where both the ring gear 38 and the planetary carrier 34 are allowed to rotate.

It is understood that the in 2 and 3 The arrangement shown is exemplary only and various alternatives are within the scope of the present disclosure. Any suitable arrangement for positioning the gearbox 30 within the engine 10 and/or connecting the gearbox 30 to the engine 10 may be used, for example only. As another example, the connections (e.g., linkages 36, 40 in the example of FIG 2 ) between the gearbox 30 and other parts of the Engine 10 (such as the input shaft 26, output shaft and fixed structure 24) may have some degree of rigidity or flexibility. As another example, any suitable arrangement of bearings between rotating and stationary parts of engine 10 (e.g., between the input and output shafts of the transmission and fixed structures such as the transmission case) may be used and the disclosure is disclosed not to the exemplary arrangement of 2 limited.

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

Optionally, the transmission can drive auxiliary and/or alternative components (e.g. the medium-pressure compressor and/or a booster).

Other gas turbine engines to which the present disclosure may have application may have alternative configurations. For example, such engines can have an alternative number of compressors and/or turbines and/or an alternative number of connecting shafts. As a further example, this points to 1 The gas turbine engine shown has a split flow nozzle 20, 22, which means that the flow through the bypass duct 22 has its own nozzle which is separate from the engine core nozzle 20 and radially outward therefrom. However, this is not limiting and any aspect of the present disclosure may also apply to engines where flow through bypass duct 22 and flow through core 11 are upstream of (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. be mixed or combined. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. For example, although the example described relates to a turbofan engine, the disclosure may apply to any type of gas turbine engine, such as a turbofan engine. 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.

The geometry of the gas turbine engine 10 and components thereof is or are defined by a conventional axis system having an axial direction (which is aligned with the axis of rotation 9), a radial direction (in the bottom-up direction in 1 ) and a circumferential direction (perpendicular to the view in 1 ) includes. The axial, radial and circumferential directions are perpendicular to one another.

The planet gears 32 of the planetary gear 30 are usually supported by plain bearings, not shown here, which are lubricated with oil. Due to the rotating movements in the planetary gear 30, this oil tends to be guided radially outwards.

In the 4 shows the mounting of a planet wheel 32 (not visible here itself) in a transmission part 50, the inner surface 60 of the transmission part 50 being visible in the view. The transmission part 50 here connects two parts of the planet carrier 34 (one on the input side, one on the output side), ie it transmits a torque

When the planetary wheel 32 rotates about its axis of rotation R, oil flows 62 are established which are directed from the axial outside inwards. With a superimposed radial movement, the oil streams 62 flow outwards in the direction of the ring gear 38, with the toothing (ring gear) 38a lying on the inside here. The teeth 38a are divided here, i.e. a gap 39 is arranged between the teeth 38a.

It is fundamentally undesirable for the oil to flow into the ring gear teeth 38a, since the oil heats up there and causes the oil to age. As shown below, oil guide means 61 are used to guide oil flows 61 axially and radially outwards or to guide the oil flows 61 in a targeted manner into the gap between toothings.

In the 3 an intersection curve AA is given, the intersection of which is in 5 is shown. In 5 1 shows an embodiment in which two grooves or recesses 61a are arranged in the surface 60, which are arranged in such a way that the oil flows 62 are respectively directed axially and outwardly. There, the oil flows 62 can then leave the transmission part 50 without further mechanical stress.

The oil guide means 61a in the (curved) surface 60 in the projection are essentially between 30 and 50°, in particular 45°, perpendicular to the axis of rotation R of the planet wheel 32.

the 6 until 9 show similar views of area 60 as in FIG 5 shown.

In 6 an embodiment is shown which produces a similar effect by the surface 60 is contoured, eg by a bevel, which directs the oil axially and outwardly. In the 8th a concave area (ie as an example of a contouring) is designed as an oil guide means 61b, with which oil streams 62 are directed into a gap 39 between the teeth 38a (see 4 ) are managed. This means that the oil is not subjected to mechanical stress again. In the 9 a convex area is shown as a further example of contouring, with which the oil flow 62 is guided in a targeted manner to the edges of the sliding bearing device.

In the 7 it is shown that the oil flows 62 can also be guided in the direction of the gap 39 between the teeth 38a by a suitable contouring or correspondingly arranged grooves as oil guide means 61a.

In principle, recesses and contours can also be combined with one another in order to achieve targeted guidance of the oil flows 62

It should be understood that the invention is not limited to the embodiments described above, and various modifications and improvements can be made without departing from the concepts described herein. Any of the features may be employed separately or in combination with any other feature, provided they are not mutually exclusive, and the disclosure extends to and encompasses all combinations and sub-combinations of one or more features described herein.

Reference List

9

main axis of rotation

10

gas turbine engine

11

core engine

12

air intake

14

low-pressure compressor

15

high-pressure compressor

16

incinerator

17

high pressure turbine

18

bypass exhaust nozzle

19

low pressure turbine

20

core thruster

21

engine nacelle

22

bypass channel

23

fan

24

stationary support structure

26

wave

27

connecting shaft

28

sun gear

30

Gears, planetary gears

32

planet gears

34

planet carrier

36

linkage

38

ring gear

38a

Ring Gear

39

gap between teeth

40

linkage

50

gear part

60

surface in the gear part

61a

Oil guide means (recess, groove)

61 b

Oil guide means as contouring of the surface on the transmission part

62

oil flow

A

core airflow

B

bypass airflow

R

Axis of rotation of the planet wheel

Claims (11)

Transmission part (50) of a planetary gear (30) which is connected to a planet carrier (34) for planetary gears (32), the transmission part (50) at least partially having an axial extent along the planetary gears (32) and the planetary gears (32) at least partially surrounded, characterized in that at least one oil guide means (61a, 61b) is arranged on a surface of the transmission part (50), in particular on a surface (60) which faces the planet gears (32), so that at least one oil flow (62 ) is led away from a lubrication point (32, 38, 38a). transmission part claim 1 , characterized in that the at least one oil flow (62) can be guided with an axial flow component to the edge of the planetary carrier (34) and/or can be guided with an axial flow component into a gap (39) between two toothings (38a). Transmission part (50) after claim 1 or 2 , characterized in that the at least one oil guide means (61a) has a depression or groove. Transmission part (50) according to at least one of the preceding claims, characterized characterized in that the at least one oil guide means (61a) has a wall. Transmission part (50) according to at least one of the preceding claims, characterized in that the at least one oil guide means (61b) has a three-dimensional contouring of the surface (60), in particular a concave and/or convex area. Transmission part (50) according to at least one of claims 2 until 5 , characterized in that the at least one oil guide means (61a) is arranged in the projection essentially between 30 and 50°, in particular 45°, perpendicularly to the axis of rotation (R) of the planet wheel (32). Transmission part (50) according to at least one of the preceding claims, characterized in that the at least one oil guide means (61a, 61b) guides the oil flow (60) radially outwards to the edge of the planet carrier (34). Transmission part (50) according to at least one of the preceding claims, characterized in that it is set up and designed to transmit a torque from one side of the planet carrier to the other. Transmission part (50) according to at least one of the preceding claims, characterized in that at least one oil flow (60) originates entirely or partially from a slide bearing of a planet wheel (32). Transmission device with at least one transmission part (50) according to one of the preceding claims. A gas turbine engine (10) for an aircraft, comprising: a core engine (11) including a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) positioned upstream of the core engine (11), the fan (23) including a plurality of fan blades; and a gear (30) drivable by the core shaft (26), the fan (23) being drivable by the gear (30) at a lower speed than the core shaft (26), the gear (30) being a gear device after claim 10 having.

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