DE102019219474A1 - Motor with freewheel mechanism - Google Patents

Abstract

The proposed solution concerns an engine (10) for an aircraft, with a core engine (11), which has a high pressure turbine (17), a low or medium pressure turbine (19) and a compressor (15) as well as a first the high pressure turbine (17) with the compressor (15) connecting shaft (27), - a multiple fan blades having, rotatable fan (23) which is arranged upstream of the core engine (11), and - a gearbox (30), via which a second, with the Low or medium pressure turbine (19) connected shaft (26) is coupled to the fan (23) in order to transmit a driving force from the low or medium pressure turbine (19) to the fan (23) during operation of the engine (10) and thereby to drive the fan (23) at a lower speed than the second shaft (26). The proposed engine (10) comprises a freewheel device (50) via which the first shaft (27) and the second shaft (26) when transmitting a Driving force from the fan (23) to the low or Mi Oil-pressure turbine (19) can be coupled to one another in a rotationally fixed manner and via which the first shaft (27) and the second shaft (26) are decoupled from one another when a drive force is transmitted from the low-pressure or medium-pressure turbine (19) to the fan (23).

Description

The proposed solution relates to an engine for an aircraft with a gearbox, via which a fan of the engine can be driven at a speed which is lower than the speed of a shaft at the input of the gearbox which transmits the drive force for the fan.

Engines, in particular so-called gas turbine engines or aircraft engines with a gear for reducing the speed in the direction of a fan of the engine, are widely known. In this case, a drive torque transmitted to a (core) shaft from a turbine of the engine is typically transmitted to the fan via the gear unit, which is designed as a reduction gear and is designed, for example, as a planetary gear, in order to drive the fan at a speed that is lower than the speed of the driving shaft . The speed of the shaft is greater than the speed of the fan, for example, by the factor of a stationary gear ratio of a planetary gear.

The fan is typically coupled to a second (core) shaft of the engine via the gearbox, which shaft is connected to a low-pressure or medium-pressure turbine of a core engine of the engine. Another, first (core) shaft, which is then also referred to as a high-pressure shaft, is also provided for connecting a high-pressure turbine of the core engine to an upstream compressor. When the engine is in operation, this first shaft connected to the high-pressure turbine rotates faster than the second shaft intended to drive the fan.

In flight operations, so-called windmilling can occur if the core engine is switched off. In this operating state of the engine, no drive force or drive torque is transmitted from the low-pressure or medium-pressure turbine to the fan. However, the fan is set in rotation by an air flow hitting the fan blades. The air flow hitting the fan during flight operations thus generates a driving force for rotating the fan. This driving force is introduced into the second (low-pressure) shaft via the gearbox. Since the gearbox is designed as a reduction gear for the operation of the engine with an active core engine, the second shaft rotates faster than the fan and thus possibly also faster than the first (high-pressure) shaft. However, just restarting the core engine at a comparatively low high-pressure shaft speed results in a comparatively high fuel-air ratio and, as a result, in a comparatively high temperature of the high-pressure turbine. Due to the thermal expansion of the turbine blades, a large radial gap is required to prevent the turbine blades from running into the housing. Large radial gaps have a negative effect on turbine efficiency, fuel consumption and turbine inlet temperature.

From the DE 10 015 118 669 A1 In another context, an engine with a reduction gear for coupling between a fan and a low-pressure or medium-pressure turbine with a freewheel device is known. By means of this freewheel device, a decoupling is achieved between a fan shaft connected in a rotationally fixed manner to the fan and a shaft of the reduction gear in order to achieve an operative connection between the fan shaft and the shaft of the reduction gear. The aim here is to avoid unwanted braking of the fan shaft in the event of windmilling and thus to prevent the fan from jamming when the core engine is switched off. With a view to the problem addressed above when restarting the engine with a comparatively slowly rotating high-pressure shaft, the DE 10 2015 118 669 A1 nothing to be found.

It is therefore the object of the proposed solution to provide an engine that is improved in this respect.

This object is achieved with an engine according to claim 1.

• - ein Kerntriebwerk, das eine Hochdruckturbine (erste Turbine), eine Nieder- oder Mitteldruckturbine (zweite Turbine) und einen Verdichter sowie eine erste die Hochdruckturbine mit dem Verdichter verbindende (Kern- oder Hochdruck-) Welle umfasst,
• - einen mehrere Fanschaufeln aufweisenden, drehbaren Fan, der - bezogen auf einen Kernluftstrom durch das Kerntriebwerk im Betrieb des Triebwerks - stromaufwärts des Kerntriebwerks angeordnet ist,
• - ein Getriebe, über das eine zweite, mit der Nieder- oder Mitteldruckturbine verbundene (Kern- oder Niederdruck-) Welle mit dem Fan gekoppelt ist, um - über das Getriebe - im Betrieb des Triebwerks eine Antriebskraft von der Nieder- oder Mitteldruckturbine an den Fan zu übertragen und hierbei den Fan mit einer niedrigeren Drehzahl als die zweite Welle anzutreiben,
• - eine Freilaufeinrichtung, über die die erste Welle und die zweite Welle bei der Übertragung einer Antriebskraft von dem Fan an die Nieder- oder Mitteldruckturbine drehfest miteinander koppelbar sind und über die die erste Welle und die zweite Welle bei der Übertragung einer Antriebskraft von der Nieder- oder Mitteldruckturbine an den Fan voneinander entkoppelt sind.

Here, an engine for an aircraft is proposed which comprises at least the following:

• - a core engine comprising a high pressure turbine (first turbine), a low or medium pressure turbine (second turbine) and a compressor as well as a first (core or high pressure) shaft connecting the high pressure turbine to the compressor,
• a rotatable fan which has a plurality of fan blades and which - in relation to a core air flow through the core engine when the engine is in operation - is arranged upstream of the core engine,
• - A gearbox via which a second (core or low pressure) shaft connected to the low or medium pressure turbine is coupled to the fan in order to - via the gearbox - apply a drive force from the low or medium pressure turbine to the during operation of the engine To transmit the fan and to drive the fan at a lower speed than the second shaft,
• - A freewheel device, via which the first shaft and the second shaft in a rotationally fixed manner when a drive force is transmitted from the fan to the low-pressure or medium-pressure turbine can be coupled to one another and via which the first shaft and the second shaft are decoupled from one another when a drive force is transmitted from the low-pressure or medium-pressure turbine to the fan.

The proposed solution is based on the basic idea of providing a freewheel in an engine, for example in a windmilling case and thus the transmission of a drive force from the fan rotating by an air flow to the fan - when the core engine is active, the drive force for the fan transmitting - second (low-pressure) shaft to couple the further first (high-pressure) shaft connected to the high-pressure turbine and vice versa when a drive force is transmitted from the second shaft to the fan - with the engine restarted and the core engine active - the second shaft and the to decouple the first wave from each other.

Accordingly, in one embodiment, the freewheel device is provided for coupling the first shaft and the second shaft in the case of windmilling when the core engine is switched off and for decoupling the first shaft and the second shaft when the core engine is switched on.

This includes in particular an embodiment variant in which the freewheel device is set up to couple the first shaft and the second shaft to one another when the second (low-pressure) shaft rotates faster than the first (high-pressure) shaft. In this case, the freewheel device can furthermore be set up to (again) decouple the first and second shafts, which are coupled to one another in a rotationally fixed manner, when the first shaft, driven by the high-pressure turbine, rotates faster than the second shaft.

With the proposed solution, it can be achieved that in the case of windmilling and the associated rotation of the fan with the core engine switched off, the high-pressure shaft is coupled to the low-pressure shaft in a torque-proof manner when a speed or rotational speed of the low-pressure shaft exceeds a speed of the high-pressure waves. With the fan set in rotation by the flow of air and a resulting fan-side drive force, the high-pressure shaft can be brought to an advantageously higher speed via the gear unit that acts as a reduction gear when the core engine is switched on. A higher speed of the high-pressure shaft then has a positive effect when the core engine is restarted. The first (high-pressure) shaft can thus be driven directly by the fan, which functions as a turbine in windmilling operation, via the free-wheeling device provided. This can significantly improve the restart properties of the engine.

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. By way of example only, the high-pressure turbine connected to the first shaft can be a first turbine, the compressor connected to the core shaft can be a first compressor, and the first shaft can be a first core shaft. The core engine may further comprise 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 can be arranged to receive a flow from the first compressor (for example, to receive it 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 driven only by the core shaft, which is configured to interact (for example in use) with the to rotate the lowest speed (for example only from 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 relates, of course, 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. 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 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 peak load can be defined as dH / U _peak ² , where dH is the enthalpy increase (for example the average 1-D enthalpy increase) across the fan and U _{peak is} the (translational) speed of the fan tip, for example an 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, which is limited 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 preceding sentence (i.e., the values can be 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 preceding sentence (i.e., the values can be 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 material, for example a metal matrix composite and / or an organic matrix composite such as e.g. B. carbon fiber. As a further 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 disk). 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. become.

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 condition may be outside of 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 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 up 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; one 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 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.

The attached figures illustrate possible variants of the proposed solution by way of example.

• 1 eine vergrößerte Schnittdarstellung einer Ausführungsvariante eines vorgeschlagenen (Gasturbinen-) Triebwerks mit einer als Überholkupplung ausgebildeten Freilaufeinrichtung für eine vom Betriebszustand des Triebwerks abhängige Ein- und Entkopplung zwischen einer Niederdruckwelle und einer Hochdruckwelle des Triebwerks;
• 2 eine Seitenschnittansicht eines Gasturbinentriebwerks, auf dem die Ausführungsvariante der 1 basiert;
• 3 eine Schnittansicht eines stromaufwärtigen Abschnitts des Gasturbinentriebwerks der 2;
• 4 eine zum Teil geschnittene Ansicht eines Getriebes für ein Gasturbinentriebwerk der 1 bis 4.

Here show:

• 1 an enlarged sectional view of a variant of a proposed (gas turbine) engine with a freewheeling device designed as an overrunning clutch for coupling and decoupling between a low-pressure shaft and a high-pressure shaft of the engine, depending on the operating state of the engine;
• 2 a side sectional view of a gas turbine engine on which the variant embodiment of 1 based;
• 3 FIG. 10 is a sectional view of an upstream portion of the gas turbine engine of FIG 2 ;
• 4th a partially sectioned view of a transmission for a gas turbine engine of 1 to 4th .

2 represents a gas turbine engine 10 (e.g. an aircraft) 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 flows: a core air flow 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 via a second (low pressure) shaft 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 from and thereby drive 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 a suitable first (high pressure) wave 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 3 shown. The low pressure turbine 19th (please refer 1 ) drives the second 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 linkage 36 with the fan 23 coupled to the effect of its rotation around the engine axis 9 to drive. A External gear or ring gear 38 with an internal toothing, the linkage 40 with a stationary support structure 24 is coupled, is located by the planet gears 32 radially outside and combs with 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 gearbox 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 4th 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 3 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 3 and 4th epicyclic planetary gear shown as an example 30th is a planetary gear in which the planet carrier 34 via linkage 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. As another example, the planetary gear 30th be a star arrangement in which the planet carrier 34 is held fixed, allowing the ring gear (or outer gear) 38 to rotate. With such an arrangement the fan becomes 23 from the ring gear 38 driven. As another alternative example, the transmission 30th be a differential gear that allows both the ring gear 38 as well as the planet carrier 34 rotate.

Each of the planet gears 32 is on a bearing pin 35 of the planet carrier 34 rotatably mounted, in the present case via a plain bearing which, in the specific example, is lubricated with a lubricant, for example an engine oil. Alternatively, the planet gears are 32 stored by means of a ball or roller bearing.

It goes without saying that the in 3 and 4th 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 linkages 36 , 40 in the example of 3 ) between the gearbox 30th and other parts of the gas turbine engine 10 (such as the second shaft provided 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 3 limited. 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 are different in the case of a star arrangement (described above) of the transmission 30th usually by those who exemplified in 3 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 arrangements, and bearing positions.

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 shafts. As another example, the 2 The gas turbine engine shown has a split flow nozzle 20th , 22nd on what that means the current through the bypass duct 22nd has its own nozzle, that of the engine core nozzle 20th is separate and radially outward therefrom. 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 turbofan 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 or are defined by a conventional axis system that has an axial direction (that of the axis of rotation 9 aligned), a radial direction (in the direction from bottom to top in 2 ) and a circumferential direction (perpendicular to the view in 2 ) includes. The axial, radial and circumferential directions are perpendicular to one another.

The 1 shows a further development of the (gas turbine) engine in an enlarged sectional view 10 the 2 and 3 according to the proposed solution.

Here is downstream of the transmission 30th with the core engine switched on 11 as a reduction gear for a drive torque to be transmitted from the low-pressure turbine in the direction of the fan 2 19th acts, a freewheel device 50 intended. About this freewheel device 50 are the first (high pressure) wave 27 and the second (low pressure) wave 26th of the engine 10 rotatably coupled to each other during windmilling. So can the fan 23 with the core engine switched off 11 and thus not from the low-pressure turbine 19th to the fan 23 The drive torque to be transmitted to the fan during flight operations 23 hitting air flow are set in rotation. Through the transmission 30th leads the rotation of the fan 23 to a rotation of the low pressure turbine 19th coupled second wave 26th with a speed that is higher than the speed of the fan 23 . The second wave 26th can turn faster during a windmilling than the first one with the high-pressure turbine 17th connected shaft 27 . In the embodiment variant shown, in this case the freewheeling device, which is designed as an overrunning clutch, is used 50 the second from the fan 23 about the gearbox 30th driven shaft 26th with the first wave 27 non-rotatably coupled. If the speed of the second (low-pressure) wave is therefore during windmilling 26th higher than the speed of the first (high pressure) wave 27 become the two waves 26th and 27 via the freewheel device 50 rotatably coupled with each other, so that one by the rotation of the fan 23 applied driving force also to that with the high pressure turbine 27 connected first wave 27 is transmitted. In this way the first shaft rotates 27 during windmilling at a higher speed than without a free-wheeling device 50 . When the engine is restarted 10 then has this higher speed compared to conventional engines 10 improved restart behavior in flight operations with one engine 10 equipped aircraft result.

Is the core engine 11 active again and a driving force generated on the turbine side in the direction of the high-pressure compressor becomes active 15th and the fan 23 are transmitted via the freewheel device 50 the two waves 26th and 27 again decoupled from each other as soon as the first (high pressure) shaft has a speed that is the speed of the second (low pressure) shaft 26th increases. The normal operation of the engine 10 thus remains of the freewheel device 50 and influences.

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

21

Engine nacelle

22nd

Bypass duct

23

fan

24

stationary support structure

26th

second (core) wave / low pressure wave

27

first (core) wave / high pressure wave

28

Sun gear

30th

transmission

32

Planetary gear

34

Planet carrier

35

Bearing pin

36

Linkage

38

Ring gear

40

Linkage

50

Freewheel device

A.

Core airflow

B.

Bypass airflow

QUOTES INCLUDED IN THE DESCRIPTION

This list of the documents listed by the applicant was generated automatically and is included solely for the better information of the reader. The list is not part of the German patent or utility model application. The DPMA assumes no liability for any errors or omissions.

Patent literature cited

• DE 10015118669 A1 [0005]
• DE 102015118669 A1 [0005]

Claims (5)

Engine for an aircraft, with - a core engine (11) which has a high pressure turbine (17), a low or medium pressure turbine (19) and a compressor (15) and a first shaft connecting the high pressure turbine (17) to the compressor (15) (27) comprises - a multiple fan blades, rotatable fan (23) which is arranged upstream of the core engine (11), and - a transmission (30) via which a second, connected to the low or medium pressure turbine (19) Shaft (26) is coupled to the fan (23) in order to transmit a drive force from the low or medium pressure turbine (19) to the fan (23) during operation of the engine (10) and here the fan (23) with a low To drive speed as the second shaft (26), characterized in that the engine (10) comprises a freewheel device (50), via which the first shaft (27) and the second shaft (26) when transmitting a driving force from the fan ( 23) to the low or medium pressure turbine (19) dr are permanently coupled to one another and via which the first shaft (27) and the second shaft (26) are decoupled from one another when a drive force is transmitted from the low or medium pressure turbine (19) to the fan (23). Engine after Claim 1 , characterized in that the freewheel device (50) for coupling the first shaft (27) and the second shaft (26) in the case of windmilling with the core engine (11) switched off and for decoupling the first shaft (27) and the second shaft (26 ) is provided when the core engine (11) is switched on. Engine after Claim 1 or 2 , characterized in that the freewheel device (50) is set up to couple the first shaft (27) and the second shaft (26) to one another when the second shaft (26) rotates faster than the first shaft (27). Engine after Claim 3 , characterized in that the freewheel device (50) is set up to decouple the first and second shafts (27, 26), which are coupled to one another in a rotationally fixed manner, when the first shaft (27), driven by the high-pressure turbine (17), rotates faster than the second shaft (26). Engine according to one of the preceding claims, characterized in that the gear (30) is designed as a planetary gear.

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