WO2020239417A1

WO2020239417A1 - Planetary transmission of a gas turbine engine, and gas turbine engine

Info

Publication number: WO2020239417A1
Application number: PCT/EP2020/063029
Authority: WO
Inventors: David Krüger
Original assignee: Rolls-Royce Deutschland Ltd & Co Kg
Priority date: 2019-05-31
Filing date: 2020-05-11
Publication date: 2020-12-03
Also published as: DE102019114685A1

Abstract

A planetary transmission (30) of a gas turbine engine having at least one planet carrier (34) is described, which planet carrier is configured with at least one planet gear journal (42). At least one planet gear (32) is arranged rotatably on the planet gear journal (42). The at least one planet gear journal (45) is produced from a ceramic which is distinguished by a ratio between the modulus of elasticity thereof and the density thereof of greater than 25 GPa*cm³/g. A gas turbine engine for an aircraft is also described. A method for producing a planet gear journal from silicon nitride is additionally also proposed.

Description

Planetary gear of a gas turbine engine and gas turbine engine

The present disclosure relates to a planetary gear of a gas turbine engine with at least one planet carrier, which is designed with at least one planet gear pin. The present disclosure also relates to a gas turbine engine for an aircraft. Additionally, the present disclosure relates to a method of making a planetary gear pin.

From US 2016/0377167 A1 a planetary gear of a gas turbine engine is known. The planetary gear comprises a sun gear, a ring gear fixed to the housing and a rotatable planet carrier via which a fan can be driven. A plurality of planet gears are in mesh with the sun gear and with the ring gear.

Furthermore, from DE 10 2014 105 024 A1, a planetary gear transmission with a planetary gear carrier with a plurality of planet bolts for receiving planetary gears is known. The planet pins can be made of metal, a metal alloy or ceramic. In this context, it is stated that metals, metal alloys or ceramics offer the advantage of a high modulus of elasticity or E modulus and high strength. Furthermore, it should be noted that planetary pins only make up a small proportion of a planetary gear and thus the total weight of a planetary gear is only slightly increased by making the planetary pins, for example, in metal, whereas the rigidity can be greatly increased.

A planetary gear and a gas turbine engine with a planetary gear are to be made available, each of which is characterized by a low component weight. Furthermore, a method for producing a planet gear bolt is to be specified. This object is achieved with a planetary gear and with a gas turbine engine and with a method of patent claims 1, 8 and 10, respectively

According to a first aspect, a planetary gear of a gas turbine engine is provided with at least one planet carrier, which is designed with at least one planet wheel pin on which at least one planet wheel is rotatably arranged Density greater than 25 GPa * cm ³ / g is marked

The present disclosure is based on the knowledge that in known planetary gears whose components are made of steel, whose planet carriers are designed to be rotatable and which are operated in particular at high rotational speeds and under high loads, the tendency of planet gear pins to bend is undesirably high. The bending of the steel-made planetary gear bolts that occurs during operation influences the center distances of mutually meshing gears and the distribution of the gear meshing forces.

If, for example, a planetary gear is mounted on a steel planetary gear pin via a plain bearing, bending of the planetary gear pin affects the bearing gap between the planetary gear and the planetary gear pin. If the bend of the planet gear bolt increases too much, this has a significant influence on the bearing properties of such a plain bearing. Planetary gear bolts are therefore mostly designed as a function of a permissible bending and not taking into account material strength. This often leads to planetary gear bolts being designed with large external dimensions and high component weight. Furthermore, contrary to prevailing opinion, the stiffness of steel for use in conventional planetary gear systems is too low to be able to effectively limit the bending of planetary gear bolts. So-called high-performance ceramics are usually characterized by a lower density than steel, which is why the use of such high-performance ceramics enables a significant weight reduction compared to steel smaller component dimensions and, in turn, can be implemented with a lower component weight. This offers the possibility of creating a planetary gear with a lower component weight without increasing the tendency of a planetary gear pin to bend. In addition, the tendency of planetary gear bolts to bend can also be significantly reduced if a planetary gear bolt made of ceramic is provided with the same dimensions as a planetary gear bolt made of steel.

If the planetary gear pin is made of silicon nitride, the planetary gear according to the present disclosure is characterized by a low component weight and, in addition, there is little bending of the planetary gear pin during operation of the planetary gear.

In addition to high strength and low density, silicon nitride is also characterized by a desired high rigidity. In the present preferred application of a planetary gear unit of a gas turbine engine, the planet gear pin and the planet carrier can be made as stiff as possible while the planet gear unit is lightweight. It can thus be achieved that the toothing between the planetary gears and a sun gear and / or a ring gear exhibit the smallest possible skew division with respect to one another under load. Such a misalignment is promoted, among other things, by the torque transmission between the mutually meshing gears and possibly also by certain flight maneuvers

Since silicon nitride, in addition to its high strength and low density, has a high modulus of elasticity of up to 400 GPa, which is roughly double Equivalent to the modulus of elasticity of steel, planetary gear pins made of silicon nitride can be made with the lowest possible component weight and high rigidity at the same time. Furthermore, silicon nitride has strengths of up to 1000 MPa. The high rigidity of silicon nitride thus offers the possibility of designing planetary gear bolts with smaller dimensions than planetary gear bolts made of steel, without worsening their tendency to bend. In combination with the lower density of silicon nitride compared to steel, the weight saving effect on the entire planetary gear as an assembly is multiplied. This in turn results from the fact that lighter planet gear bolts allow the planet carrier to be designed with lower strength and that the planet carrier as a whole can also be designed with a low component weight. The weight savings in turn enable the use of bearings with lower load capacity and other structures.

Furthermore, there is the possibility of producing planetary gear bolts from silicon nitride with a very smooth surface, which in turn promotes the plain bearing properties of the plain bearing between the planetary gear bolt and the planetary gears

In further embodiments of the planetary gear according to the present disclosure, the planetary gear meshes with at least one sun gear and / or with at least one ring gear.

Furthermore, there is the possibility that the planet carrier is carried out in a rotation test, with what is known as two-shaft operation is possible by means of the planetary gear according to the present disclosure

If the planet carrier is designed to be rotatable, a so-called three-shaft operation can be represented via the planetary gear according to the present disclosure. If the planet carrier comprises two cheeks spaced apart from one another in the axial direction and if the planet gear pin is connected to the planet carrier in a rotationally test manner at the end in the area of the cheeks, the planetary gear according to the present disclosure is characterized by a simple structural design

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

Arrangements of the present disclosure may be particularly, but not exclusively, advantageous for fans that are driven via a transmission. Correspondingly, the gas turbine engine can comprise a gearbox that receives an input from the core shaft and delivers drive for the fan for driving the fan at a lower speed than the core shaft. The input for the gearbox can be directly from the core shaft or indirectly from the core shaft, for example via a stimulus and / or a stim gear. The core shaft can be rigidly connected to the turbine and the twisted 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 can have any desired number of shafts connecting turbines and compressors, such as one, two, or three shafts. Merely by way of example, 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 instinct Werkkem 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.

With 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., receive directly, e.g., via a generally annular channel).

The gearbox can be arranged 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 above example). For example, the transmission can be arranged to be driven only 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 arranged to be driven by one or more shafts, for example the first and / or the second shaft in the above example.

In a gas turbine engine described and / or claimed here, a combustion chamber can be provided axially downstream of the fan and the compressor (s). For example, the combustion chamber can be 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 may 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 several stages. Each stage can include a series of rotor blades and a series of stator blades, which can be variable stator blades (in that their angle of attack 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 (e.g. the first turbine and the second turbine as described above) can comprise any number of stages, e.g. 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 be defined with a radial span extending from a root (or hub) at a radially inward gas overflow location or at a 0% span position to a tip at a 100% span position The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of): 0.4, 0.39, 0.38, 0.37, 0.36, 0.35 , 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26 or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip can be in an inclusive range bounded by two of the values in the preceding sentence (ie, 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 portion (or the axially furthest edge) of the blade. The hub-to-tip ratio, of course, relates to the portion of the fan blade overflowing with gas, ie the portion which 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 be simply 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 an inclusive range limited by two of the values in the preceding sentence (ie the values can form upper or lower limits).

The speed of the fan can vary during use. In general, the speed is lower for fans with a larger diameter. By way of non-limiting example only, the speed of the fan under constant speed conditions may be less than 2500 rpm, for example less than 2300 rpm. 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 USpeak. The Work done on the flow by the fan blades 13 results in an increase in enthalpy dH of the flow. A fan tip load can be defined as dH / USpitze2, where dH is the enthalpy increase (e.g. the average 1 -D enthalpy increase) across the fan and USpHze is the (translation) 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 an inclusive range bounded by two of the values in the preceding sentence (ie the values can be 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 flow through the bypass duct to the mass flow rate of flow through the core at constant velocity conditions. In some arrangements, the bypass ratio can be more than (on the order of): 10, 10.5, 11, 11, 5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5 , 16, 16.5 or 17 are (lie). The bypass ratio can lie in an inclusive range bounded by two of the values in the preceding seed (i.e. the values can form upper or lower limits). The bypass channel can be essentially ring-shaped. The bypass channel can be located radially outside the engine core. 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 expressed as the ratio of the back pressure upstream of the fan to the back pressure at the outlet of the supercharger (in front of the entrance to the combustion chamber). As a non-limiting example, the total pressure ratio of a gas turbine engine described and / or claimed herein at constant speed may be greater than (or on the order of): 35, 40, 45, 50, 55, 60, 65, 70, 75 (lie). The total pressure ratio can be in an inclusive range bounded by two of the values in the preceding sentence (ie 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. Under constant speed conditions, the specific thrust of an engine described and / or claimed here may be less than (or on the order of): 110 Nkg-1s, 105 Nkg-1s, 100 Nkg-1s, 95 Nkg-1s, 90 Nkg -1s, 85 Nkg-1s or 80 Nkg-1s. The specific thrust can be in an inclusive range bounded by two of the values in the preceding sentence (i.e. the values can form upper or lower limits). Such engines can be particularly efficient compared to conventional gas turbine engines.

A gas turbine engine as 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): 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300 kN, 350kN, 400kN, 450kN, 500kN or 550kN. The maximum thrust can be in an inclusive range bounded by two of the values in the preceding sentence (ie the values can be upper or lower limits). The thrust referred to above may be the net maximum thrust under standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees 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 chamber, 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 on the order of): 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at constant speed can be in an inclusive range bounded by two of the values in the preceding sentence (ie the values can form upper or lower limits). The maximum TET when the engine is in use can, for example, be at least (or on the order of): 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET can lie in an inclusive range bounded by two of the values in the preceding 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 a blade portion 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 portion of the fan blade and / or the blade can be made at least in part 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 can have a front protective edge, which is Turn of a material that can withstand impact (such as birds, ice or other material) better than the rest of the shovel. Such a leading edge can be manufactured using titanium or a titanium-based alloy, for example. Thus, as an 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 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 portion of the fan blades can be machined from a block and / or at least a portion of the fan blades can be welded, e.g. B. linear friction welding, can be attached to the hub / disc.

The gas turbine engines that are described and / or claimed here may or may not be provided with a VAN (Variable Area Nozzle - nozzle with variable cross section). Such a nozzle with a variable cross-section can allow a variation in the exit cross-section of the bypass channel in use allow. 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 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 0.7-0.9, e.g. 0.75-0.85, e.g. 0.76-0.84, e.g. 0.77-0 .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 travel condition. In 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 mean 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 Range from 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, correspond. 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 degrees C.

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

In use, 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. 2 or 4) gas turbine engine may be attached to provide thrust.

According to a further aspect of the present disclosure, a method for producing a planetary gear pin from silicon nitride is proposed. First, a container is filled with powdery silicon nitride and closed closed. To sinter the silicon nitride, the filled container is subjected to pressure and thermal energy on all sides from the outside

In this case, the container is preferably a canister made of steel. Such a canister is usually welded

In a variant of the method according to the present disclosure, the filled container for sintering the silicon nitride is subjected to temperatures of up to 2000 ° C. and pressures between 180 MPa and 250 MPa, preferably of approximately 200 MPa

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 herein can be applied to any aspect and / or combined with any other feature or parameter described herein, if they can not mutually exclusive.

Embodiments will now be described by way of example with reference to the figures.

Further advantages and advantageous developments of the invention emerge from the patent claims and the exemplary embodiment described in principle with reference to the drawing.

It shows

1 is a longitudinal sectional view of a gas turbine engine; FIG. 2 is an enlarged partial longitudinal sectional view of a flow mouth section of a gas turbine engine; FIG.

3 shows a single illustration of a transmission for a gas turbine engine; and

FIG. 4 shows a sectional view of the transmission along a section line IV-IV identified in more detail in FIG. 3.

Figure 1 shows a gas turbine engine 10 with a main axis of rotation 9. The engine 10 comprises an air inlet 12 and a thrust fan 23, which generates two air flows, a core air flow A and a bypass air flow B. The gas turbine engine 10 includes a core 11 which receives the core air flow A The Engine core 11 comprises, in axial flow sequence, a low-pressure compressor 14, a high-pressure compressor 15, a combustion device 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core thrust nozzle 20. An engine nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass air flow 18 through the bypass duct 22. The fan 23 is attached to the low-pressure turbine 19 via a shaft 26 and an epicycloid gear 30 and is driven by the latter.

In use, the core air flow A is accelerated and compressed by the low-pressure compressor 14 and passed into the high-pressure compressor 15, where further compression takes place.The compressed air expelled from the high-pressure compressor 15 is passed into the combustion device 16 where it is mixed with fuel and the mixture is burned. The resulting hot combustion products then propagate through the high pressure and low pressure turbines 17, 19 and thereby drive them before they are ejected through the nozzle 20 to provide some thrust. The high pressure turbine 17 drives the high pressure compressor 15 through a suitable connecting shaft le 27 at. The fan 23 generally provides the majority of the thrust. The epicycloidal gear 30 is a reduction gear.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2 The low pressure turbine 19 (see Figure 1) drives the shaft 26, which is coupled to a sun gear 28 of the epicycloidal gear arrangement 30 , are located radially outward from and mesh with the sun gear 28. The planet carrier 34 restricts the planet gears 32 to orbit synchronously around the sun gear 28 while allowing each planet gear 32 to rotate about its own axis. Planet carrier 34 is coupled to fan 23 via linkage 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 located radially on the outside of planetary gears 32 and meshes with them

It is 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 (ie that they do not include the fan 23 include) and / or the turbine and compressor stages, which are connected to one another by the connecting shaft 26 with the lowest speed in the engine (ie that it does not include the transmission output shaft which drives the fan 23). In some documents, the " Low Pressure Turbine "and the" Low Pressure Compressor "referred to herein may alternatively be known as the" Medium Pressure Turbine "and" Medium Pressure Compressor ". Using such alternative nomenclature, the fan 23 may be referred to as a first compression stage or compression stage with the lowest pressure.

The epicycloidal gear 30 is shown in more detail by way of example in FIG. 3. The sun gear 28, the planet gears 32 and the ring gear 38 each include Teeth around their periphery to mesh with the other gears. However, for the sake of clarity, only exemplary portions of the teeth are shown in Figure 3. Although four planet gears 32 are shown, it will be obvious to those skilled in the art that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicycloidal gear 30 generally include at least three planet gears 32.

The epicycloidal gear 30 shown by way of example in FIGS. 2 and 3 is a planetary gear in which the planet carrier 34 is coupled to an output shaft via linkage 36, the ring gear 38 being fixed. However, any other suitable type of epicycloidal gear 30 can be used. As another example, the epicyclic gear 30 may be a star arrangement in which the planet carrier 34 is held in place, allowing the ring gear (or outer gear) 38 to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the transmission 30 may be a differential gear that allows both the ring gear 38 and the planetary carrier 34 to rotate.

It goes without saying that the arrangement shown in FIGS. 2 and 3 is merely exemplary and various alternatives are within the scope of protection of the present disclosure. Any suitable arrangement for positioning the transmission 30 in the engine 10 and / or for connecting the transmission 30 to the engine 10 can be used merely as an example. As another example, the connections (e.g., the linkages 36, 40 in the example of Figure 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) have some degree of rigidity or flexibility. As another example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example, between the input and output shafts of the gearbox and the fixed structures such as the gearbox housing) are used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, it is readily apparent to a person skilled in the art that the arrangement of the output and support rods and bearing positions in a stem arrangement (described above) of the transmission 30 is usually would differ from those shown by way of example in FIG.

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

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 a further example, the gas turbine engine shown in Figure 1 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 nozzle 20 and radially outside thereof, however, this is not restrictive and any Aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed or combined 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 fan engine, the disclosure may apply to any type of gas turbine engine such as z. B. in an open rotor (in which the fan stage is not surrounded by an engine nacelle) or a T urboprop-T engine.

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

The planet gears 32 are each rotatably arranged on planet gear bolts 42. The planet carrier 34 has, in the manner shown in more detail in FIG. 4, two cheeks 43, 44 which are axially spaced from one another and which are firmly connected to one another. The ends of the planet gear bolts 42 are connected to the planet carrier 34 in a rotationally fixed manner in the region of the cheeks 43, 44. Furthermore, the planetary gear bolts 42 are each made of silicon nitride. In addition, outer sides 45 of the planetary gear bolts each form an inner running surface of a sliding bearing for the planetary gears 32.

It goes without saying that the invention is not restricted to the embodiments described above and various modifications and improvements can be made without departing from the concepts described here. 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.

Claims

Patent claims

1. Planetary gear (30) of a gas turbine engine (10) with at least one planet carrier (34) which is executed with at least one planet gear pin (42) on which at least one planet gear (32) is rotatably arranged, the at least one planet gear pin (42 ) is made from a ceramic which is characterized by a ratio between its modulus of elasticity and its density greater than 25 GPa * cm ³ / g

2. Planetary gear according to claim 1, characterized in that the planet wheel bolt (42) is made of silicon nitride

3. Planetary gear according to claim 1 or 2, characterized in that the planet gear (32) is in engagement with at least one sun gear (28)

4. Planetary gear according to one of claims 1 to 3, characterized in that at least one ring gear (38) is provided which meshes with the at least one planet gear (32)

5. Planetary gear according to one of claims 1 to 4, characterized in that the planet carrier (34) is designed to be rotationally test

6. Planetary gear according to one of claims 1 to 4, characterized in that the planet carrier (32) is rotatable

7. Planetary gear according to one of claims 1 to 6, characterized in that the planet carrier (32) comprises two cheeks (43, 44) spaced apart from one another in the axial direction, the planet gear pin (42) at the end in the region of the cheeks (43, 44) ) rotated with the planet carrier (34) is connected

8. A gas turbine engine (10) for an aircraft, comprising an engine core (11) which comprises a turbine (19), a compressor (14) and a core shaft (26) connecting the turbine (19) to the compressor (14) ;

a fan (23) positioned upstream of the engine core (11), the fan comprising a plurality of fan blades and

a gearbox (30) which receives an input from the core shaft (26) and provides drive for the fan (23) for driving the fan (23) at a lower speed than the core shaft (26), the gearbox (30) as a Planetary gear according to one of claims 1 to 7 is carried out

9. Gas turbine engine according to claim 8, characterized in that 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 engine core (11) further comprises a second turbine (17), a second compressor (15) and a second core shaft (27) which connects the second turbine (17) to the second compressor (15); and

the second turbine (17), the second compressor (15) and the second core shaft (27) are arranged to rotate at a higher speed than the first core shaft (26).

10. A method for producing a planet gear pin (42) from silicon nitride, characterized by the following process steps:

- Filling a container with powdered silicon nitride

- Closing the filled container

- Sintering the silicon nitride by applying pressure and thermal energy to the filled container from the outside.

11. The method according to claim 10, characterized in that the filled

Container for sintering the silicon nitride at temperatures of up to 2000 ° C and pressures between 180 MPa and 250 MPa, preferably about 200 MPa.