DE102019125273A1 - Bearing device for an input or output shaft of a planetary gear in a gas turbine engine - Google Patents

Abstract

The invention relates to a mounting device for an input or output shaft of a planetary gear (30) in a gas turbine engine, in particular a safety gear engine (10) of an aircraft, characterized by a roller bearing (50) for the input or output shaft of the planetary gear (30), whose pitch circle diameter (D) is larger than the diameter of the orbit (P), which goes through the center points (M) of the planetary gears (32).

Description

The present disclosure relates to a bearing device for an input and output shaft of a planetary gear in a gas turbine engine with the features of claim 1 and a gas turbine engine with the features of claim 8.

Planetary gears, which are regularly exposed to high mechanical loads, are used in a fan gear engine of an aircraft. In particular, no deformations or displacements of the inlet or outlet shafts of the planetary gear should be transmitted to the planetary gear.

Therefore, there is a need to create bearings for planetary engines that are robust.

In this case, a storage device for an input or output shaft of a planetary gear can be used in a gas turbine engine. The gas turbine engine can in particular be a fan gear engine of an aircraft.

A roller bearing is used for the input or output shaft of the planetary gear, the pitch circle diameter of which is larger than the diameter of the orbit that goes through the center points of the planetary gears. This makes it possible to find an axially compact design for this area of the gas turbine engine

The roller bearing can be designed as a cylindrical roller bearing, a tapered roller bearing, ball bearing or spherical roller bearing.

For a mechanically stable embodiment, the distance between the center line of the roller bearing and the center line of the planetary gear can be less than 1.5 times the pitch circle diameter, in particular less than 1.2 times the pitch circle diameter, in particular equal to the pitch circle diameter ( D. ), be.

The roller bearing can absorb axial and / or radial loads during assembly and operation.

For a secure arrangement of the roller bearing, in one embodiment, the outer ring of the roller bearing can be firmly connected to a static part of the safety gear drive unit.

In the event that the roller bearing is arranged on the output side of the planetary gear, the inner ring of the roller bearing can be coupled to a drive shaft for a fan.

Typically, the outer diameter of the roller bearing can be in the range between 300 and 1000 mm, in particular in the range between 500 and 700 mm.

The task is also addressed by a gas turbine engine with the features of claim 8.

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 comprising 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 (with fan blades) positioned upstream of the core engine.

Arrangements of the present disclosure may be particularly, but not exclusively, advantageous for transmission fans that are driven via a transmission. Correspondingly, the gas turbine engine can comprise a transmission which is driven via the core shaft and the output of which drives the fan in such a way that it has a lower speed than the core shaft. The input for the gearbox can take place directly 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 the compressor so that the turbine and the compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine described and / or claimed herein can be of any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting the turbines and compressors, such as one, two, or three shafts. By way of example only, the turbine connected to the core shaft can be a first turbine, the compressor connected to the core shaft can be a first compressor, and the core shaft can 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 can be positioned axially downstream of the first compressor. The second compressor may be arranged to receive flow from the first compressor (e.g. directly, for example via a generally annular channel).

The transmission can be configured to be driven by the core shaft, which is 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 can be designed in such a way that it is only driven by the core shaft, which is 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 above example ). Alternatively, the transmission can be designed in such a way that it is driven by one or more shafts, for example the first and / or the second shaft in the above example.

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 (for example at its outlet) if a second compressor is provided. As a further example, the flow at the outlet of the compressor can be fed to the inlet of the second turbine if a second turbine is provided. The burner device can be provided upstream of the turbine (s).

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

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

Each fan blade may have a radial span extending from a root (or hub) at a radially inward gas overflow location or 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 can 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 can be in a closed range bounded by two values in the preceding sentence (i.e. the values can be upper or lower limits). These ratios can generally be 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 leading edge) of the blade. The hub-to-tip ratio, of course, relates to the portion of the fan blade overflowing with gas; H. the portion that is radially outside 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, 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 can be in a closed range bounded by two of the values in the preceding sentence (i.e. the values can be 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. As a 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. Merely as a further non-limiting example, the speed of the fan under constant speed conditions for an engine with a fan diameter in the range from 250 cm to 300 cm (for example 250 cm to 280 cm) in the range from 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. Merely as a further non-limiting example, the speed of the fan under constant speed conditions for an engine with a fan diameter in the range from 320 cm to 380 cm in the range from 1200 rpm to 2000 rpm, for example in the range from 1300 rpm. min to 1800 rpm, for example in the range from 1400 rpm to 1600 rpm.

When the gas turbine engine is in use, 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 _tip ² , where dH is the enthalpy increase (e.g. the average 1-D enthalpy increase) across the fan and U _{tip is} the (translational) speed of the fan tip, e.g. at the front edge of the tip , (which can be defined as the fan tip radius at the front edge multiplied by the angular velocity). The fan peak load at constant speed conditions can be more 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, which is limited by two of the values in the previous sentence (ie the values can form upper or lower limits).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, the bypass ratio being defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at constant velocity 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 be (lie). The bypass ratio can be in a closed range bounded by two of the values in the preceding sentence (i.e., the values can be upper or lower limits). The bypass channel can be essentially ring-shaped. The bypass duct can be located radially outside the core engine. The radially outer surface of the bypass duct can be defined by an engine nacelle and / or a fan housing.

The total pressure ratio of a gas turbine engine described and / or claimed herein can be defined as the ratio of the back pressure upstream of the fan to the back pressure at the outlet of the super high pressure compressor (before the inlet to the burner device). As a non-limiting example, the total 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 (lie). The total pressure ratio can be in a closed range bounded by two of the values in the preceding sentence (i.e. the values can be 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 of the specific thrust of a jet engine, which is described and / or claimed may be less than (or in the order of) 110 N kg ^-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 (lying). 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 can have any maximum thrust desired. As a non-limiting example, a gas turbine described and / or claimed herein can be used to generate 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 550kN. The maximum thrust can be in a closed range bounded by two of the values in the preceding sentence (i.e. the values can be upper or lower limits). The thrust referred to above may be the maximum net thrust under standard atmospheric conditions at sea level plus 15 ° C (ambient pressure 101.3 kPa, temperature 30 ° C) with a static engine.

In use, the temperature of the flow at the inlet of the high pressure turbine can be particularly high. This temperature, which can be referred to as TET, can be measured at the exit to the combustion device, for example immediately upstream of the first turbine blade, which in turn can be referred to as a nozzle guide vane. At constant speed, the TET can be at least (or in the order of magnitude of): 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K. The TET at constant speed 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 maximum TET when the engine is in use can be at least (or in the order of magnitude): 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K, for example. The maximum TET 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 maximum TET can occur, for example, in a condition of high thrust, for example in an MTO condition (MTO - maximum take-off thrust - maximum take-off thrust).

A fan blade and / or aerofoil of a fan blade described and / or claimed herein can be made from any suitable material or combination of materials. For example, at least a part of the fan blade and / or the blade can be at least partly made of a composite, for example a metal matrix composite and / or a composite with an organic matrix, such as e.g. B. carbon fiber. As another example, at least a portion of the fan blade and / or the blade can be at least in part made of a metal, such as metal. A titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade can include at least two sections made using different materials. For example, the fan blade may have a protective leading 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 using titanium or a titanium-based alloy, for example. Thus, by way of example only, the fan blade may have a carbon fiber or aluminum based body (such as an aluminum-lithium alloy) with a leading edge made of titanium.

A fan described and / or claimed herein may include a central portion from which the fan blades may extend, for example in a radial direction. The fan blades can be attached to the central section in any desired manner. For example, each fan blade can include a fixation device that can engage a corresponding slot in the hub (or disc). Only as an example, such a fixing device can be in the form of a dovetail, which can be inserted into a corresponding slot in the hub / disc and / or brought into engagement therewith in order to fix the fan blade to the hub / disc. As another example, the fan blades can be formed integrally with a central portion. Such an arrangement can be referred to as a blisk or a bling. Any suitable method can be used to manufacture such a blisk or bling. For example, at least a part of the fan blades can be machined from a block and / or at least a part of the fan blades can be welded, e.g. B. linear friction welding, can be attached to the hub / disc.

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

The fan of a gas turbine described and / or claimed herein can have any desired number of fan blades, for example 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 can conventionally be defined as the conditions during the middle part of the flight, for example the conditions to which the aircraft and / or the engine are exposed between (in terms of time and / or distance) the end of the climb and the start of the descent. will.

By way of example only, the forward speed under the constant speed condition may be at any point in the range of 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 in the order of Mach 0.8, in the order of Mach 0.85 or in the range from 0.8 to 0, 85, lie. Any speed within these ranges can 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.

By way of example only, the constant velocity conditions may correspond to standard atmospheric conditions at an altitude that 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 feet), for example in the range from 10,500 m to 11,500 m, for example in the range from 10,600 m to 11,400 m, for example in the range from 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 from 10,900 m to 11,100 m, for example in the order of 11,000 m. The constant velocity conditions can correspond to standard atmospheric conditions at any given altitude in these areas.

By way of 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 they are used throughout here, “constant speed” or “constant speed conditions” can mean the aerodynamic design point. Such an aerodynamic design point (or ADP - Aerodynamic Design Point) can correspond to the conditions (including, for example, the Mach number, environmental conditions and thrust requirement) for which the fan company is designed. This can mean, for example, the conditions under which the fan (or the gas turbine engine) has the optimum degree of efficiency according to its design.

In operation, a gas turbine engine described and / or claimed herein can be operated at the constant speed conditions defined elsewhere herein. Such constant speed conditions may be determined by the constant speed conditions (e.g., the conditions during the middle part of flight) of an aircraft to which at least one (e.g., two or four) gas turbine engine (s) may be attached to provide thrust.

It will be understood by those skilled in the art that a feature or parameter described in relation to any of the above aspects can be applied to any other aspect, provided that they are not mutually exclusive. Furthermore, any feature or parameter described here can be applied to any aspect and / or combined with any other feature or parameter described here, insofar as 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 Gastu rb i nentriebwerk;
• 4 eine schematische Schnittansicht einer Ausführungsform einer Lagerungsvorrichtung auf der Abtriebsseite eines Planetengetriebes;
• 5 eine Detailschnittansicht einer weiteren Ausführungsform einer Lagerungsvorichtung auf der Abtriebsseite eines Planetengetriebes;
• 6 eine weitere Ausführungsform mit einem Wälzlager auf der Antriebsseite.

Embodiments will now be described by way of example with reference to the figures; in the figures show:

• 1 a side sectional view of a gas turbine engine;
• 2 Fig. 3 is a side sectional close-up view of an upstream portion of a gas turbine engine;
• 3rd a partially cut-away view of a transmission for a gas turbine engine;
• 4th a schematic sectional view of an embodiment of a storage device on the output side of a planetary gear;
• 5 a detailed sectional view of a further embodiment of a storage device on the output side of a planetary gear;
• 6th another embodiment with a roller bearing on the drive side.

1 represents a gas turbine engine 10 with a main axis of rotation 9 The gas turbine engine 10 includes an air inlet 12th and a fan 23 that creates two air streams: a core air stream A. and a bypass air flow B. . The gas turbine engine 10 includes a core 11 that is the core airflow A. records. The core engine 11 includes, in axial flow order, a low pressure compressor 14th , a high pressure compressor 15th , an incinerator 16 , a high pressure turbine 17th , a low pressure turbine 19th and a core thrust nozzle 20th . An engine nacelle 21 surrounds the gas turbine engine 10 and defines a bypass channel 22nd and a bypass thrust nozzle 18th . The bypass airflow B. flows through the bypass channel 22nd . The fan 23 is about a wave 26th and an epicyclic planetary gear 30th on the low pressure turbine 19th attached and is driven by this.

In operation, the core air flow A. by the low pressure compressor 14th accelerated and compressed and in the high pressure compressor 15th where further compression takes place. The one from the high pressure compressor 15th compressed air is discharged into the incinerator 16 where it is mixed with fuel and the mixture is burned. The resulting hot products of combustion then propagate through the high pressure and low pressure turbines 17th , 19th and thereby propel them before they are used to provide a certain thrust through the nozzle 20th be expelled. The high pressure turbine 17th drives the high pressure compressor 15th by means of a suitable connecting shaft 27 at. The fan 23 generally provides the majority of the thrust. The epicyclic planetary gear 30th is a reduction gear.

An exemplary arrangement for a geared fan gas turbine engine 10 is in 2 shown. The low pressure turbine 19th (please refer 1 ) drives the wave 26th at that with a sun gear 28 of the epicyclic planetary gear 30th is coupled. Several planet gears 32 by a planet carrier 34 are coupled to each other, are located by the sun gear 28 radially outside and mesh with it. The planet carrier 34 guides the planetary gears 32 so that they are in sync around the sun gear 28 orbit while it allows each planetary gear to move 32 can rotate around its own axis. The planet carrier 34 is about a fan wave 36 with the fan 23 coupled to the effect of its rotation around the engine axis 9 to drive. An outer gear or ring gear 38 that is about linkage 40 with a stationary support structure 24 is coupled, is located by the planet gears 32 radially outside and combs it.

It should be noted that the terms “low-pressure turbine” and “low-pressure compressor”, as used here, can be understood to mean the turbine stage with the lowest pressure and the compressor stage with the lowest pressure (i.e. that it is not the fan 23 include) and / or the turbine and compressor stages that are driven by the connecting shaft 26th with the lowest speed in the engine (meaning that it is not the transmission output shaft that drives the fan 23 drives, includes) are interconnected, mean. In some scriptures, the “low pressure turbine” and “low pressure compressor” referred to here may alternatively be known as the “medium pressure turbine” and “medium pressure compressor”. When using such alternative nomenclature, the fan 23 may be referred to as a first compression stage or the lowest pressure compression stage.

The epicyclic planetary gear 30th is in 3rd shown in more detail as an example. The sun gear 28 who have favourited planet gears 32 and the ring gear 38 each include teeth around their circumference to enable meshing with the other gears. However, for the sake of clarity, only exemplary sections of the teeth are shown in FIG 3rd shown. Although four planet gears 32 are shown, it is obvious to those skilled in the art that more or fewer planetary gears are within the scope of the claimed invention 32 can be provided. Practical applications of an epicyclic planetary gear 30th generally include at least three planetary gears 32 .

This in 2 and 3rd epicyclic planetary gear shown as an example 30th is a planetary gear in which the planet carrier 34 about the fan wave 36 is coupled to an output shaft, the ring gear 38 is fixed. However, any other suitable type of planetary gear can be used 30th be used.

It goes without saying that the in 2 and 3rd The arrangement shown is exemplary only and various alternatives are within the scope of the present disclosure. Any suitable arrangement for positioning the transmission can be used merely as an example 30th in the gas turbine engine 10 and / or to connect the transmission 30th with the gas turbine engine 10 be used. As another example, the connections (e.g. the fan shaft, linkage 36 , 40 in the example of 2 ) between the gearbox 30th and other parts of the gas turbine engine 10 (such as the input shaft 26th , the output shaft and the established structure 24 ) have some degree of rigidity or flexibility. As another example, any suitable arrangement of the bearings between rotating and stationary parts of the gas turbine engine can be used 10 (For example, between the input and output shafts of the transmission and the specified structures, such as the transmission housing), and the disclosure is not limited to the exemplary arrangement of FIG 2 limited.

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

Other gas turbine engines to which the present disclosure may find 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 another example, the 1 The gas turbine engine shown has a split flow nozzle 20th , 22nd on, which means that the flow is through the bypass duct 22nd has its own nozzle, that of the engine core nozzle 20th is separate and radially outward from it. However, this is not limiting and any aspect of the present disclosure may also apply to engines in which the flow is through the bypass duct 22nd and the current through the core 11 mixed or combined in front of (or upstream) a single nozzle which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) can have a fixed or variable range. For example, while the example described relates to a turbo engine, the disclosure may be applied to any type of gas turbine engine such as a gas turbine engine. B. in an open rotor (in which the fan stage is not surrounded by an engine nacelle) or a turboprop engine.

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

In 4th Figure 3 illustrates one embodiment of a storage device used in a gas turbine engine 10 (here a fan gear engine) is arranged, which is essentially the 2 corresponds to.

The planetary gear 30th is on the output side via the planet carrier 34 via a spline connection with the fan shaft 36 connected that ultimately the fan 23 drives. The sun gear is used on the drive side 28 the drive of the planetary gears 32 that are in the ring gear 38 circulate.

There is a roller bearing on the output side 50 arranged, whose pitch circle diameter D. is larger than the diameter of the orbit P passing through the midpoints M. the planetary gears 32 goes (see also 3rd , in which the geometric relationships for the planetary gear 30th are shown).

The pitch circle diameter D. is to be understood here as the diameter of the circle that goes through the centers of the rolling elements.

With such a design, ie a roller bearing 50 , which has a comparatively large diameter, the planet carrier 34 axially close to the planetary gear 30th be moved. The pitch circle diameter D. is therefore also larger than the diameter of the planet carrier 38 on the output side.

It should be noted that the bearing device can also be used as a unit of rolling bearings 50 and the planetary gear 30th can be understood.

In the 5 Another embodiment of the storage device is shown, which is also on the output side of the planetary gear 30th is arranged. The planetary gear 30th is shown here only halfway, with the line MP the center plane (perpendicular to the main axis of rotation 9 ) of the planetary gear 30th shows.

The roller bearing is axially further forward 50 arranged, the inner ring 51 with the fan wave 36 connected is. The outer ring 52 of the rolling bearing 50 is associated with a static part of the gas turbine engine 10 connected.

The pitch circle diameter D. is larger than the diameter of the orbit P passing through the midpoints M. the planetary wheels 32 goes.

The distance between the center line MW of the rolling bearing 50 and the center line of the planetary gear MP is smaller than 1.4 times the pitch circle diameter D. .

In the 6th a further embodiment is shown. This has a roller bearing 50 for the input shaft of the planetary gear 30th whose pitch diameter is larger than the diameter of the orbit that goes through the centers of the planetary gears.

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 can be used separately or in combination with any other features, provided that they are not mutually exclusive, and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

List of reference symbols

9

Main axis of rotation

10

Gas turbine engine

11

Core engine

12th

Air inlet

14th

Low pressure compressor

15th

High pressure compressor

16

Incinerator

17th

High pressure turbine

18th

Bypass thrust nozzle

19th

Low pressure turbine

20th

Core thruster

21st

Engine nacelle

22nd

Bypass duct

23

fan

24

stationary support structure

26th

wave

27

Connecting shaft

28

Sun gear

30th

transmission

32

Planetary gears

34

Planet carrier

36

Fan wave

38

Ring gear

40

Linkage

50

roller bearing

51

Inner ring of the rolling bearing

52

Outer ring of the rolling bearing

A.

Core airflow

B.

Bypass airflow

D.

Pitch diameter rolling bearing

M.

Center planetary gear

MP

Center line planetary gear

MW

Center line rolling bearing

P

Orbit planetary gears

Claims (9)

Bearing device for an input or output shaft of a planetary gear (30) in a gas turbine engine, in particular a safety gear engine (10) of an aircraft, characterized by a roller bearing (50) for the input or output shaft of the planetary gear (30), the pitch circle diameter (D ) is greater than the diameter of the orbit (P), which goes through the center points (M) of the planetary gears (32). Storage device according to Claim 1 , characterized in that the roller bearing (50) is designed as a cylindrical roller bearing, as a tapered roller bearing, ball bearing or spherical roller bearing. Storage device according to Claim 1 or 2 , characterized in that the distance between the center line (MW) of the roller bearing (50) and the center line of the planetary gear (MP) is less than 1.5 times the pitch circle diameter (D), in particular less than 1.2 times the Pitch circle diameter (D), in particular, is equal to the pitch circle diameter (D). Storage device according to at least one of the preceding claims, characterized in that axial and / or radial loads can be absorbed with the roller bearing (50) during assembly and operation. Bearing device according to at least one of the preceding claims, characterized in that the outer ring (51) of the roller bearing (50) is firmly connected to a static part of the safety gear drive unit (10). Bearing device according to at least one of the preceding claims, characterized in that the inner ring (52) of the roller bearing (50) is coupled to a drive shaft (36) for a fan (23). Bearing device according to at least one of the preceding claims, characterized in that the outer diameter of the roller bearing (50) is in the range between 300 and 1000 mm, in particular in the range between 500 and 700 mm. A gas turbine engine (10) for an aircraft comprising: a core engine (11) comprising 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) comprising a plurality of fan blades; and a gear (30) which can be driven by the core shaft (26), wherein the fan (23) can be driven by means of the gear (30) at a lower speed than the core shaft (26), the gear (30) having a Storage device according to at least one of the Claims 1 to 7th is coupled. Gas turbine engine after Claim 8 wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the core shaft is a first core shaft (26); the core engine (11) further comprises a second turbine (17), a second compressor (15) and a second core shaft (27) connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher speed than the first core shaft.

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