This application claims priority to German Patent Application DE102018106572.8 filed Mar. 20, 2018 and German Patent Application DE102019106633.6 filed Mar. 15, 2019. The entirety of both applications are incorporated by reference herein.

The present disclosure relates to a gearbox with a rotatably mounted structural component that is formed with a rotationally symmetrical duct. Further, the present disclosure relates to a gas turbine engine for an aircraft.

A gearbox of a jet engine is known from practice. The gearbox comprises a sun gear, a ring gear fixedly attached at the housing, and a rotatable planetary carrier via which a fan can be driven. Multiple planetary gears are in engagement with the sun gear and the ring gear. Via a collecting duct that is connected to the planetary carrier of the gearbox, oil is conducted in the direction of the tooth meshings between the planetary gears and the ring gear. The collecting duct extends in the circumferential direction of the planetary carrier and is embodied to be open radially inside. Via the radially inner opening, oil is introduced via a supply feed.

However, the oil that is guided into the collecting duct is disadvantageously not accelerated to the degree which is necessary for further conduction in the direction of the planetary gear bearings inside the collecting duct. As a result, there is the danger of an undersupply to the planetary gear bearings as well as to the area of the tooth meshings between the structural components of the gearbox. Among other things, this is due to the fact that the wall friction between the collecting duct and the oil which is introduced into the collecting duct is not sufficient to accelerate the oil inside the collecting duct in the radial direction towards the direction of the planetary gear bearings.

What is to be provided is a gearbox and a gas turbine engine with a gearbox in which the oil supply of a rotatable structural component of the gearbox is respectively ensured.

This objective is achieved through a gearbox and a gas turbine engine with features as disclosed herein.

Provided according to a first aspect is a gearbox with a rotatably mounted structural component that is formed with an at least approximately rotationally symmetrical duct. Starting from a radially inner area, oil can be guided into the duct from an oil supply that is fixedly attached at the housing. Starting from the oil supply, an introduction direction of the oil into the duct encloses an angle in a range between 45° and 90° with the radial extension direction of the duct. Alternatively, there is also the possibility that, starting from the oil supply, the introduction of oil into the duct encloses an angle of between 60° and 90° with the axial extension direction of the duct. Further, the introduction direction of the oil in the circumferential direction of the duct encloses an angle with the radial extension direction that is larger than or equal to 0° and smaller than 90°.

In this manner, an impulse is applied to the oil when it is introduced into the duct from the oil supply, such that the oil in the duct is accelerated to an extend as it is for example necessary for it to be conducted further in the direction of a bearing of the rotatable structural component, and an undersupply of the bearing or of tooth meshings between structural components of the gearbox is prevented.

What is understood by the term duct in the present case is an outer wall-like boundary that delimits an inner area that is duct-shaped at least in certain areas. Here, the boundary itself is embodied to be at least approximately rotationally symmetrical independently of the installations that are arranged in the inner area of the boundary.

Further, starting from the oil supply, the introduction of the oil into the duct encloses an angle of between 75° and 90°, preferably of between 80° and 90°, with the axial extension direction of the duct. If the extension direction of the oil encloses an angle with the axial extension direction of the duct that is within the last-mentioned angular range, an impulse is applied to the oil already as it is being introduced into the duct. This impulse ensures that the oil is accelerated inside the duct to an extent as it is necessary for further conduction, for example in the direction of a bearing of the rotatable structural component, and an undersupply of the bearing or also of the tooth meshings between structural components of the engine is avoided.

If the duct has at least one outlet opening for the oil in a radially outer area, the oil is accelerated in the direction of the outlet opening by the centrifugal force that is applied to the oil during rotation, and from there is conducted further in the necessary manner in the direction of the areas of the gearbox which are to be impinged with oil, such as a bearing and/or a toothing.

According a further aspect of the present disclosure, the structural component is a rotating shaft, preferably a sun gear, a planetary carrier, a planetary gear and/or a ring gear. In that case, a bearing or a tooth meshing of a planetary gearbox can be supplied with oil to the necessary extent, and a reliable and safe operation of the entire gearbox can be ensured in a constructionally simple manner.

In particular in the case of circumferential planetary gears, a transfer from the stationary oil system to the planetary gears that are rotating about the sun gear of the gearbox is ensured.

The oil supply can comprise an oil nozzle, with its outlet opening being arranged radially and/or axially at a distance to an inlet opening for the oil that is provided in the radially inner area of the duct.

If at least one oil nozzle is present, which in the installation position of the gearbox is arranged centrally inside the radially inner area of the duct, an impulse as it is necessary for a sufficient oil supply to loads of the gearbox can be applied to the oil with small effort as it is introduced into the duct.

If at least two oil nozzles are provided, which in the installation position of the gearbox are respectively arranged between a rotational axis of the duct and the duct as well as centrally inside the radially inner area of the duct, an impulse as it is necessary for a sufficient oil supply to loads of the gearbox can be applied to the oil.

According to a further aspect, at least eight oil nozzles, preferably 10 or more oil nozzles, are provided. Here, it can be provided that, in the installation position of the gearbox, respectively one oil nozzle is arranged between a rotational axis of the duct and the inner radial area of the duct as well as centrally inside the duct, while the further oil nozzles are positioned at a distance to the two oil nozzles in the circumferential direction with respect to the two centrally arranged oil nozzles radially inside the inner area of the duct.

The present disclosure provides a non-closed oil transfer unit that is characterized by a self-adjusting system of nozzles that are arranged across the circumference. Via the oil nozzles, lubricating and cooling agents are sprayed into a rotating duct or groove. A supply pressure is built up inside the rotating collecting groove and where applicable the following distribution lines as a result of the centrifugal force. The non-closed system that comprises the oil nozzles and the collecting duct allows for the possibility of different liquid levels in the rotating system. In turn, the supply pressure in the rotating system depends on the present liquid level. What results is a self-adjusting and robust supply system in which the loads are supplied in a manner that is considerably less dependent on the back pressure. If the back pressure, i.e. the pressure present in the load and the pressure loss to the load is in an acceptable pressure range and accordingly is neither too low nor too high, the adjustment can occur through the liquid level in the rotating supply lines. If the oil volume flow increases, so does the supply pressure, and more oil is pressed in the direction of the loads, such as for example a bearing. In the reverse case, the supply pressure drops if the volume flow decreases. Through a suitably set back pressure, running dry of the oil system is avoided, with a reduced but continuous supply taking place instead.

The gearbox that is introduced here comprises an oil transfer unit without wearing-out contact seals. Further, the gearbox has a self-adjusting system of oil nozzles which are arranged about the circumference of the duct and via which the lubricating and cooling agent or the oil is injected into a rotating collecting duct, in most instances a groove.

As is explained elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine can comprise an engine core which comprises a turbine, a combustion space, a compressor, and core shaft that connects the turbine to the compressor. Such a gas turbine engine can comprise a fan (with fan blades) that is positioned upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven by means of a gearbox. Correspondingly, the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive for the fan for driving the fan with a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example by means of a spur shaft and/or gear. The core shaft may be rigidly connected to the turbine and the compressor, such that the turbine and the compressor rotate at the same rotational speed (with the fan rotating at a lower rotational speed). Here, the gearbox can be embodied as a planetary gearbox as it is described in more detail above.

The gas turbine engine as described and claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may 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, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) a flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only by the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by one or multiple shafts, for example the first and/or second shaft in the above example.

In a gas turbine engine as described and claimed herein, a combustion space may be provided axially downstream of the fan and the compressor (or the compressors). For example, the combustion space may be located directly downstream of the second compressor (for example at the exit thereof), if a second compressor is provided. By way of further example, the flow at the exit of the combustor may be provided to the inlet of the second turbine, if a second turbine is provided. The combustion space may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and the second compressor according to the above description) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (i.e. in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset with respect to each other.

The or each turbine (for example the first turbine and second turbine according to the above description) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset with respect to each other.

Each fan blade may be defined with a radial span width extending from a root (or hub) at a radially inner gas-washed location, or from a 0% span position to a tip with a 100% span width. Here, 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) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in a closed range bounded by any two values in the previous sentence (i.e., the values may represent upper or lower bounds). These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading (or the axially forwardmost) edge of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion that is located radially outside any platform.

The radius of the fan may be measured between the engine centerline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any 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 (about 150 inches) cm or 390 cm (about 155 inches). The fan diameter may be in a closed range bounded by any two of the values in the previous sentence (i.e. the values may represent upper or lower bounds).

The rotational speed of the fan may vary during use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be in the range from 1700 rpm to 2500 rpm, for example in the range of between 1800 rpm to 2300 rpm, for example in the range of between 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of between 320 cm to 380 cm may be in the range of between 1200 rpm to 2000 rpm, for example in the range of between 1300 rpm to 1800 rpm, for example in the range of between 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with the associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utipis the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as the fan tip radius at the leading edge multiplied by the angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (with all units in this paragraph being Jkg−1K−1/(ms−1)2). The fan tip loading may be in a closed range bounded by any two of the values in the previous sentence (i.e. the values may represent upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is 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 cruise conditions. In some arrangements, the bypass ratio may be greater 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. The bypass ratio may be in a closed range bounded by any two of the values in the previous sentence (i.e. the values may represent upper or lower bounds). The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan housing.

The overall pressure ratio of a gas turbine engine as described and claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest-pressure compressor (before entry into the combustion space). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruising speed may be greater than (or on the order of): 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in a closed range bounded by any two of the values in the previous sentence (i.e. the values may represent upper or lower bounds).

Specific thrust of a gas turbine engine may be defined as the net thrust of the gas turbine engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine as described and/or claimed herein 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 may be in a closed range bounded by any two of the values in the previous sentence (i.e. the values may represent upper or lower bounds). Such gas turbine engines may be particularly efficient as compared to conventional gas turbine engines.

A gas turbine engine as described and claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing 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 550 kN. The maximum thrust may be in a closed range bounded by any two of the values in the previous sentence (i.e. the values may represent upper or lower bounds). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 deg C.), with the engine being static.

In use, the temperature of the flow at the entry to the high-pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustion space, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruising speed, the TET may be at least (or on the order of): 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K. The TET at cruising speed may be in a closed range bounded by any two of the values in the previous sentence (i.e. the values may represent upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of): 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. The maximum TET may be in a closed range bounded by any two of the values in the previous sentence (i.e. the values may represent upper or lower bounds). The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade as described herein may be manufactured from any suitable material or combination of materials. For example, at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fiber. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminum based material (such as an aluminum-lithium alloy) or a steel based material. The fan blade may comprise at least two regions that are manufactured by using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fiber or aluminum based body (such as an aluminum lithium alloy) with a titanium leading edge.

A fan as described herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture device which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture device may be present in the form of a dovetail that may be inserted into a corresponding slot in the hub/disc and/or may engage with the same in order to fix the fan blade to the hub/disc. By way of further example, the fan blades may be formed integrally with a central portion. Such an arrangement may be referred to as a blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow for the exit area of the bypass duct 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 engine as described and claimed herein may have any desired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may refer to the cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of descent.

Purely by way of example, the forward speed at cruise condition may be any point in the range from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85, or in the range from 0.8 to 0.85. Any single speed within these ranges may be the cruise condition. For some aircrafts, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions at an altitude that is in the range from 10000 m to 15000 m, for example in the range from 10000 m to 12000 m, for example in the range from 10400 m to 11600 m (around 38000 ft), for example in the range from 10500 m to 11500 m, for example in the range from 10600 m to 11400 m, for example in the range from 10700 m (around 35000 ft) to 11300 m, for example in the range from 10800 m to 11200 m, for example in the range from 10900 m to 11100 m, for example on the order of 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to the following: a forward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of −55 deg C.

As used anywhere herein, “cruise” or “cruise conditions” may refer to the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach Number, environmental conditions and thrust requirement) in which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or the gas turbine engine) is designed to have optimum efficiency.

During operation, a gas turbine engine as described and claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example two or four) gas turbine engine may be mounted in order to provide propulsive thrust.

A person skilled in the art will appreciate that, except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or can be combined with any other feature or parameter described herein.

