DE102021209401A1 - Planetary gear with a sealing system and gas turbine engine with a planetary gear - Google Patents

Abstract

The invention relates to a planetary gear (30) with at least one sealing system (100), the annular sealing surface (101) of which is arranged on a rotatable component (120), the sealing surface (101) sealing axially against a stationary component (130), characterized in that that the sealing surface (101) in the axial direction (AX) delimits a sealing gap (102) and the sealing surface (101) at least partially has a non-linear contour (103). The invention also relates to a gas turbine engine (10).

Description

The present disclosure relates to a planetary gear set having the features of claim 1 and a gas turbine engine having the features of claim 13.

In mechanical systems, here in particular planetary gears, it is often necessary to provide seals. Planetary gears, which are used, for example, in a gas turbine engine of an aircraft, have to work maintenance-free for a very long time, so that the task in particular is to create efficient, robust and cost-effective planetary gears.

The planetary gear has at least one sealing system, the annular sealing surface of which is arranged on a rotatable component, the sealing surface sealing axially against a stationary component. The sealing surface delimits a sealing gap in the axial direction and the sealing surface has at least partially a non-linear contour. Due to the non-linear contour of the sealing surface, the sealing gap can have a particularly steeply rising shape.

In one embodiment, at least one hydrostatic reservoir for oil is arranged on the rotatable component. This reservoir can be arranged at the wide end of the sealing gap, i.e. the sealing gap narrows and rises steeply away from the reservoir.

The sealing system has at least one minimum sealing gap between the rotatable component and the stationary component, in which the distance between the sealing system and the stationary component is minimal compared to other areas of the sealing gap. This minimum sealing gap can be arranged at the end of the steeply converging sealing gap.

In one embodiment, the non-linear contour at least partially has a convex section, in particular a strictly convex section, with the radius of curvature of the convex section pointing in the direction of the sealing gap and/or the stationary component. The sealing gap can thus converge steeply at least on one side. A concave curvature of the sealing gap would rather result in an asymptotic convergence.

The non-linear contour can assume different shapes in the cross section perpendicular to the axial direction. The contour can thus comprise at least part of an arc of a circle or part of an ellipse. It is also possible for the contour to be curved at least partially monotonously, in particular strictly monotonously. The contour can also have a polynomial or an exponential shape in the cross section perpendicular to the axial direction.

The non-linear contour can have a constant shape in the radial direction of the sealing system, which makes it easier to produce the contour.

The sealing system can have a large number of segments S, which each have the same non-linear contour among one another. A segment can each have a non-linear contour or the sealing surface in a segment has two non-linear contours, with the two non-linear contours of the segment being arranged symmetrically along the circumferential direction, e.g. opposite the hydrostatic reservoir. For example, the sealing system can have between 60 and 12, in particular 24, segments.

The object is also achieved by a gas turbine engine having the features of claim 13.

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 that are gear driven. 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 may be rigidly connected to 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 described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts sen that connect turbines and compressors, for example, one, two or three shafts. 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 upper or lower limits). 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 (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm (about 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) be (lie) . The fan diameter 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 speed of the fan can vary during operation. In general, the speed is lower for fans with a larger diameter. 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 having 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 of 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 ultra-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 (lay). 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 described and/or claimed herein is capable of producing 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. 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 static engine.

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. 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. become.

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 ranging from 10,000 m to 15,000 m, for example in the range from 10,000 m to 12,000 m, for example in the range from 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 mit einer Ausführungsform eines Planetengetriebes;
• 2 eine Seitenschnittgroßansicht eines stromaufwärtigen Abschnitts eines Gasturbinentriebwerks;
• 3 eine zum Teil weggeschnittene Ansicht eines Getriebes für ein Gasturbinentriebwerk;
• 4 eine Schnittansicht durch ein erstes Dichtungssystem, das in einer Ausführungsform eines Planetengetriebes einsetzbar ist;
• 5 eine Schnittansicht durch ein zweites Dichtungssystem, das in einer Ausführungsform eines Planetengetriebes einsetzbar ist;
• 6 eine Aufsicht auf eine Dichtfläche mit einer Vielzahl von Segmenten von Dichtflächen.

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

• 1 Figure 12 is a sectional side view of a gas turbine engine having an embodiment of a planetary gear set;
• 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 sectional view through a first sealing system that can be used in an embodiment of a planetary gear;
• 5 a sectional view through a second sealing system that can be used in an embodiment of a planetary gear;
• 6 a top view of a sealing surface with a plurality of segments of sealing surfaces.

Before embodiments and details of a planetary gear 30 with a sealing system 100 are described (see 6 ), is associated with the 1 until 3 a field of application described, namely a gas turbine engine 10 of an aircraft. In the 4 and 5 details of a profiling for oil distribution in the sealing system 100 are then described, which can be used in embodiments of the planetary gear 30.

1 Figure 13 illustrates a gas turbine engine 10 having a primary axis of rotation 9. The engine 10 includes an air intake 12 and a fan 23 which creates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 which receives the core air flow 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, coupled via linkage 40 to a stationary support structure 24, is radially outward of the planetary gears 32 and in mesh 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 (i.e. not including 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. However, any other suitable type of planetary gear 30 may be used. As another example, the planetary gear set 30 may be a wye arrangement in which the planetary carrier 34 is held fixed while allowing the ring gear (or ring gear) 38 to rotate. With such an arrangement, the fan 23 is driven by the ring gear 38 . 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 transmission 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) some degree of rigidity or flexibility exhibit. 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 not to the exemplary arrangement of 2 limited. For example, those skilled in the art will readily appreciate that the output and support linkage arrangements and bearing locations for a star configuration (described above) of the transmission 30 will typically differ from those exemplified in FIG 2 are shown would differ.

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.

It may be necessary to provide sealing systems 100 at various points of the planetary gear 30 .

In 4 Illustrated is a partial view of a sealing system 100 that may be employed in a first embodiment of a planetary gear set 30 such as that used in a gas turbine engine 10 .

Such a sealing system 100 can be arranged on the shaft which rotates with the planet carrier 34, for example. The sealing system 100 must create an adequate seal with the oil transfer coupling, the housing of which is a stationary component 130 here.

In 4 a top view of a sealing gap 102 is shown, which is arranged between a rotatable component 120 (direction of rotation n) and a stationary component 130 . The sealing system 100 is an axial seal; the axial direction AX is in the 4 shown.

A single segment S of the sealing system 100 is shown, which as a whole extends in the form of a ring around an axis of rotation (see FIG 6 ). Between 12 and 60 segments S, for example, are arranged along the ring, which are basically all structurally identical to one another. If the sealing system 100 z. B. has 24 segments S, then each segment S covers an angular range of 15 °.

The segment S of the sealing system in the embodiment according to 4 has a depression on the left edge as a hydrodynamic reservoir for oil. A minimum sealing gap 104 is arranged on the right-hand side, ie the sealing gap 102 is narrowest here.

The sealing gap 102 has a sealing surface 101 on the rotating component 120, which delimits the sealing gap 102 in the axial direction AX. This sealing surface 101 has a non-linear contour 103 at least in part. Which has an exponential course in the cross section shown here. Starting from the hydrostatic reservoir 105 to the minimum sealing gap 104, the sealing gap 102 becomes strictly exponentially smaller; namely steeply increasing to the area of the minimum sealing gap 104. Am There is an edge at the transition from the convex non-linear contour 103 to the minimum sealing gap 104 .

A non-strictly exponential progression, on the other hand, would also include sections that are not exponential.

In alternative embodiments, the non-linear contour 103 can also be part of an arc of a circle, an ellipse or a polynomial.

In any case, the non-linear contours would at least partially have a convex section, in particular a strictly convex section. The non-linear contour 103 opens towards the sealing gap 102 so that the radius of curvature points in the direction of the sealing gap 102 and/or in the direction of the stationary component 103 .

Generally, this non-linear transition ensures a fluidically smooth transition from the hydrostatic reservoir 105, which also improves oil distribution for a bearing. Such a design can also do without special cooling of the sealing system 100 .

If now the number of segments S (see 6 ) is kept relatively small, e.g. at 24, then the length of the convergent part (ie here the non-linear contour 103) increases in the circumferential direction, which can increase the lubricating film thickness,

In 4 it is indicated that a further segment S, which is structurally identical to the one shown, connects to the left-hand edge. Such a segment S would also follow on the right, with all segments S together then forming the ring-shaped sealing system 100 (see Fig 6 ).

This means that the embodiment according to 4 has asymmetrical segments S, which means that such a sealing system 100 can have only one direction of rotation n (here clockwise).

If the sealing system 100 is also to be used in a planetary gear 30 that rotates in two opposite directions (see double arrow n in 5 ) should work, then an embodiment according to 5 be used.

Two non-linear contours 103 are used here in a segment S, which are arranged symmetrically to the center line of the hydrostatic reservoir 105 . The basic mode of operation corresponds to that of the embodiment 4 , so that reference can be made to the above description.

In the 6 a front view of the sealing system 100 is shown, in which the ring-shaped sealing system 100 is divided into 24 segments S, a segment S here encompassing an angular range of φ=15°.

Sealing systems of the type described here can also be used for other sealing tasks, such as in internal combustion engines or wind turbines. The sealing systems 100 have also been described here in connection with oil that is used as a lubricant. In principle, sealing systems 100 of this type can also be used for sealing against other fluids.

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

40

linkage

100

sealing system

101

sealing surface

102

sealing gap

103

non-linear contour

104

minimum sealing gap

105

hydrostatic reservoir

120

rotatable component

130

stationary component

A

core airflow

AX

axial direction

B

bypass airflow

N

direction of rotation

S

segment

Claims (13)

Planetary gear (30) with at least one sealing system (100), the annular sealing surface (101) of which is arranged on a rotatable component (120), the sealing surface (101) sealing axially against a stationary component (130), characterized in that the sealing surface (101) delimits a sealing gap (102) in the axial direction (AX) and the sealing surface (101) at least partially has a non-linear contour (103). Planetary gear (30) after claim 1 , characterized in that the sealing system (101) on the rotatable component (120) has at least one hydrostatic reservoir (105) for oil. Planetary gear (30) after claim 1 or 2 , characterized in that the sealing system (101) between the rotatable component (120) and the stationary component (130) has at least one minimum sealing gap (104) in which the distance between the sealing system (100) and the stationary component (130) is opposite other areas of the sealing gap (102) is minimal. Planetary gear (30) according to at least one of the preceding claims, characterized in that the non-linear contour (103) at least partially has a convex section, in particular a strictly convex section, with the radius of curvature in the direction of the sealing gap (102) and / or of the stationary component (103). Planetary gear (30) according to at least one of the preceding claims, characterized in that the non-linear contour (103) has at least part of an arc of a circle or part of an ellipse in cross section perpendicular to the axial direction (AX). Planetary gear (30) according to at least one of the preceding claims, characterized in that the non-linear contour (103) in cross section perpendicular to the axial direction (AX) is at least partially monotonically, in particular strictly monotonically, curved. Planetary gear (30) according to at least one of the preceding claims, characterized in that the non-linear contour (103) has a polynomial or an exponential shape in cross section perpendicular to the axial direction (AX). Planetary gear (30) according to at least one of the preceding claims, characterized in that the non-linear contour (103) has a constant shape in the radial direction of the sealing system (100). Planetary gear (30) according to at least one of the preceding claims, characterized in that the sealing system (100) has a large number of segments (S) which each have the same non-linear contour (103) as one another. Planetary gear (30) after claim 9 , characterized in that in a segment (S), the sealing surface (101) has a non-linear contour (103). Planetary gear (30) after claim 9 , characterized in that in a segment (S), the sealing surface (101) has two non-linear contours (103), the two non-linear contours (103) of the segment (S) being arranged symmetrically along the circumferential direction. Planetary gear (30) according to at least one of claims 9 until 11 , characterized in that the sealing system (100) has between 60 and 12, in particular 24, segments (S). 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 planetary gear (30) according to at least one of the preceding claims, which can be driven by the core shaft (26), the fan (23) being drivable by means of the planetary gear (30) at a lower speed than the core shaft (26).

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