DE102019116954A1 - Device and method for setting a gap height - Google Patents

Abstract

The invention relates to a device for setting a gap height (H) between the tips of blades (41) of a turbine stage (17, 19) and / or a compressor stage (14, 15) and a housing (42) surrounding the blades (41) a gas turbine engine (10), characterized by a plurality of structure-borne noise sensors (1, 2, 3, 4, 1a, 2a, 3a, 4a), each of which is located on a sector (A1, A2, A3, A4, A5, A6, A7, A8 ) are arranged on and / or in the housing (42) of a turbine stage (17, 19) and / or a compressor stage (14, 15) of the gas turbine engine (10), the structure-borne noise sensors (1, 2, 3, 4, 1a, 2a , 3a, 4a) are set up and designed in such a way that a rubbing of a rotating blade of the turbine stage (17, 19) and / or the compressor stage (14, 15) can be detected using at least one structure-borne noise signal (K) and a measuring and control unit ( 5), which uses the at least one structure-borne noise signal (K) to determine the geometric location of the rubbing against the spit zen of the blades (41) on the housing (42) can be determined and, depending on this, at least one temperature control device (C1, C2, C3, C4, C5, C6, C7, C8) can be activated which corresponds to the at least one sector (A1, A2, A3 , A4, A5, A6, A7, A8) so that the gap height (H) can be adjusted locally to a target value. The invention also relates to a method for adjusting the gap height (H).

Description

The present disclosure relates to a device for setting a gap height with the features of claim 1 and a method for setting a gap height according to the features of claim 11.

Aircraft gas turbine engines are particularly efficient when the inflowing air flow hits the blades of the compressor stages and / or turbine stages and not between the gaps between the blade tips and the surrounding housing. The so-called tip clearance should therefore be kept as small as possible. For this purpose, devices are known from the prior art, which are referred to as tip clearance control. It is also known to detect the rubbing of blades on a housing surrounding them (e.g. WO 2008/060164 A1 , DE 32 31 587 A1 ), with which, among other things, an adjustment of the gap height is possible.

The object is to provide efficient devices and methods that are suitable for adjusting the gap height.

According to a first aspect, a device having the features of claim 1 is provided.

For this purpose, a plurality of structure-borne noise sensors are used, each of which is arranged in a sector on and / or in the housing of a turbine stage and / or a compressor stage of the gas turbine engine. A sector is understood here to mean a region of the housing, in particular the housing wall, in which an assigned structure-borne sound sensor can preferably receive structure-borne sound signals.

The structure-borne noise sensors are set up and designed in such a way that a rubbing against a rotating blade of the turbine stage and / or the compressor stage can be detected using at least one structure-borne noise signal. The measured frequency range depends in particular on the resonance frequencies that are excited.

Furthermore, with a measuring and control unit based on the at least one detected structure-borne noise signal, the geometric location of the brushing of the tips of the blades on the housing can be determined and depending on this, at least one temperature control device can be activated which is assigned to the at least one sector, so that the gap height is localized a setpoint can be set.

The geometrically determined arrangement of the structure-borne noise sensors enables the rubbing events in the gas turbine engine to be localized without the gap height being measured directly. The measuring and control unit, e.g. a computer that is arranged in the gas turbine engine can then initiate temperature control specifically for this sector. Depending on the desired result, e.g. be more or less cooled or heated in order to adjust the gap height in the sector of the housing.

In one embodiment, the at least one temperature control device can each have a controllable valve system for air as the temperature control medium for cooling or heating. The valve system is addressed by the measuring and control unit. Warm air can e.g. taken from one of the hot areas within the gas turbine engine during operation. Cooling air can be taken from a bypass duct or the compressor during operation; this air is cooled beforehand if necessary.

In one embodiment, the structure-borne noise sensors each have a piezo element, in particular a piezoelectric and / or a piezoresistive element. Strain gauges can also be used. These designs are sufficiently robust for use in an aircraft engine.

For the arrangement of the structure-borne noise sensors and the other units on / or in the housing, different variants are conceivable, which can also be combined with one another. E.g. the sectors, the structure-borne noise sensors and / or at least two temperature control devices can be arranged regularly around the circumference of the housing. It can thus be determined in which angular position of a plane perpendicular to the axis of rotation of the aircraft engine a rubbing event occurs.

Alternatively or additionally, the sectors, the structure-borne noise sensors and / or the at least two temperature control devices can be arranged axially along the housing. It can thus be seen whether a rubbing event occurs further forward or further back in the gas turbine engine.

If the two possible arrangements are combined with one another, it is possible to obtain a three-dimensional spatial allocation of the rubbing events in order to enable a targeted adjustment of the gap height.

In one embodiment, the measuring and control unit has a means for analyzing the frequency spectrum of the at least one structure-borne noise signal and / or for detecting the speed of the turbine stage and / or the compressor stage. With a Fourier transformation, e.g. the spectrum characteristic of the grazing can be determined.

It is in particular also possible that the measuring and control unit has a model for setting the gap height, which the speed of a turbine stage and / or compressor stage, a temperature, in particular an inlet temperature of a turbine stage, a pressure, in particular an inlet pressure of a turbine stage, a Valve position of the at least one temperature control device, the number of blades of a turbine stage and / or compressor stage, material expansion coefficients of a blade and / or the housing and / or heat exchange surfaces are taken into account. The parameters mentioned each have an influence on the gap height, so that information in the model can be used to carry out particularly efficient temperature control. For example, a temperature rise at one point on the gas turbine engine can indicate that a scraping process is imminent, so that the temperature control device can intervene more quickly when the scraping process then occurs

In one embodiment, the measuring and control unit can have an adaptive means for adjusting the gap height, which activates the at least one temperature control device until no structure-borne sound signal (because there is no rubbing) or a predefined (e.g. just tolerable) structure-borne sound signal can be detected. With the adaptive means, the gap height can be set to a predetermined level in test operation or also in operation. E.g. with the adaptive means, the gap height can be adjusted locally so that a predetermined geometry of the gap, in particular the cross-sectional area of a circular ring around the circumference and / or in the axial extent, is achieved. This can e.g. with the help of a machine learning algorithm on the basis of previous, learned attempts at the given state variables.

The object is also achieved by a method having the features of claim 11.

In a first step, a plurality of structure-borne noise sensors, each assigned to a sector on and / or in the housing of a turbine stage and / or a compressor stage of the gas turbine engine, detect the rubbing of a rotating blade of the turbine stage and / or the compressor stage using at least one structure-borne noise signal.

From this, the geometric location of the contact between the tips of the blades and the housing can then be determined by a measuring and control unit.

In response to this, at least one temperature control device can then be activated which is assigned to at least one sector, so that the gap height is set locally to a target value.

The measuring and control unit can e.g. calculate a frequency spectrum of the at least one structure-borne sound signal and / or determine the speed of the turbine stage and / or the compressor stage.

In one model, the measuring and control unit in one embodiment can use a number of parameters to adjust the gap height, such as e.g. the speed of a turbine stage and / or compressor stage, a temperature, in particular an inlet temperature of a turbine stage, a pressure, in particular an inlet pressure of a turbine stage, a valve position of the at least one temperature control device, the number of blades of a turbine stage and / or compressor stage, material expansion coefficient of a blade and / or the housing and / or heat exchange surfaces.

With an adaptive means for adjusting the gap height, the measuring and control unit can activate the at least one temperature control device until no structure-borne sound signal or a predefined structure-borne sound signal is detected. In particular, with the aid of the adaptive means, the gap height can be set locally so that a predetermined geometry of the gap, in particular a cross-sectional area of a circular ring, can be set around the circumference and / or in the axial extent. For this purpose, e.g. A machine learning algorithm is used that records the relationships between different parameters (e.g. the above) and thus automatically predicts the gap height using a model with the influencing parameters.

As noted elsewhere herein, the present disclosure may apply to a gas turbine engine, e.g. 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 whose output drives the fan so that it has a lower speed than the core shaft. The input for the gearbox can be directly from the core shaft or indirectly via the core shaft, for example via a Spur shaft and / or a spur gear take place. The core shaft may be rigidly connected to the turbine and the compressor so that the turbine and the compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine described and / or claimed herein can be of any suitable general architecture. For example, the gas turbine engine can have any desired number of shafts connecting the turbines and compressors, such as one, two, or three shafts. For example only, the turbine connected to the core shaft can be a first turbine, the compressor connected to the core shaft can be a first compressor and the core shaft can be a first core shaft. The core engine may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.

In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor 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 gearbox 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 rotate (for example in use) at the lowest speed (for example only by the first core shaft and not the second core shaft in the above example ). Alternatively, the transmission can be designed in such a way that it is driven by one or more shafts, for example the first and / or the second shaft in the above example.

In a gas turbine engine as described and / or claimed herein, a combustor may be provided axially downstream of the fan and compressor (or compressors). For example, the burner device can be located directly downstream of the second compressor (for example at its outlet) if a second compressor is provided. As a further example, the flow at the outlet of the compressor can be fed to the inlet of the second turbine if a second turbine is provided. The burner device can be provided upstream of the turbine (s).

The or each compressor (for example the first compressor and the second compressor as described above) can comprise any number of stages, for example several 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 (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 can have a radial span that extends 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 previous sentence (i.e. the values can be upper or lower limits). These ratios can generally be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip can both be measured at the leading edge (or the axially leading edge) of the blade. The hub-to-tip ratio, of course, relates to the portion of the fan blade overflowing with gas; H. the section 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 previous sentence (that is, 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 of 320 cm to 380 cm in the range of 1200 rpm to 2000 rpm, for example in the range of 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 (e.g. the average 1-D enthalpy increase) across the fan and U _{peak is} the (translational) speed of the fan tip, e.g. at the front edge of the tip , (which can be defined as the fan tip radius at the front edge multiplied by the angular velocity). The fan peak load at constant speed conditions can be more than (or on the order of): 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38 , 0.39, or 0.4 (all units in this section are Jkg ^-1 K ^-1 / (ms ^-1 ) ² ). The fan peak load can be in a closed range bounded by two of the values in the previous sentence (ie the values can form upper or lower 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 (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 overall 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 from two of the values in the previous sentence are limited (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 as described and / or claimed herein can have any maximum thrust desired. As a non-limiting example only, a gas turbine described and / or claimed herein can be used to generate a maximum thrust of at least (or on the order of): 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN , 300 kN, 350 kN, 400 kN, 450 kN, 500 kN or 550kN. The maximum thrust can be in a closed range bounded by two of the values in the preceding sentence (i.e., the values can be upper or lower limits). The thrust referred to above may be the maximum net thrust under standard atmospheric conditions at sea level plus 15 ° C (ambient pressure 101.3 kPa, temperature 30 ° C) with a static engine.

In use, the temperature of the flow at the inlet of the high pressure turbine can be particularly high. This temperature, which can be referred to as TET, can be measured at the exit to the combustion device, for example immediately upstream of the first turbine blade, which in turn can be referred to as a nozzle guide vane. At constant speed, the TET can be at least (or in the order of magnitude of): 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K. The TET at constant speed can be in a closed range bounded by two of the values in the previous set (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, 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 may 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, 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 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 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 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 here 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 that the aircraft and / or the engine between (in terms of time and / or Distance) at the end of the climb and the start of the descent.

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 speed condition. For some aircraft, the constant speed condition outside these ranges may be below Mach 0.7 or above Mach 0.9, for example.

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 to 11,100 m, for example in the order of 11,000 m. The constant velocity conditions can correspond to standard atmospheric conditions at any given altitude in these areas.

By way of example only, the constant speed conditions may correspond to: a forward Mach number of 0.8; a pressure of 23,000 Pa and a temperature of -55 ° C.

As they are used throughout here, “constant speed” or “constant speed conditions” can mean the aerodynamic design point. Such an aerodynamic design point (or ADP - Aerodynamic Design Point) can correspond to the conditions (including, for example, the Mach number, environmental conditions and thrust requirement) for which the fan company is designed. This can mean, for example, the conditions under which the fan (or the gas turbine engine) has the optimum 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., conditions during the mid-flight portion) 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 herein can be applied to any aspect and / or combined with any other feature or parameter described herein, if they can not mutually exclusive.

• 1 eine Seitenschnittansicht eines Gasturbinentriebwerks;
• 2 eine Seitenschnittgroßansicht eines stromaufwärtigen Abschnitts eines Gasturbinentriebwerks;
• 3 eine zum Teil weggeschnittene Ansicht eines Getriebes für ein Gasturb i nentriebwerk;
• 4 eine schematische Ansicht einer Ausführungsform einer Vorrichtung zur Einstellung einer Spalthöhe um einen Umfang herum;
• 5 eine schematische Ansicht einer weiteren Ausführungsform mit einer Temperierluftzuführung;
• 6 eine schematische Darstellung einer Ausführungsform zu Einstellung der Spalthöhe entlang der axialen Erstreckung des Gasturb i nentriebwerks.

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

• 1 a side sectional view of a gas turbine engine;
• 2 Figure 3 is a side sectional close-up view of an upstream portion of a gas turbine engine;
• 3 a partially cut-away view of a transmission for a gas turbine engine;
• 4th a schematic view of an embodiment of a device for adjusting a gap height around a circumference;
• 5 a schematic view of a further embodiment with a temperature control air supply;
• 6th a schematic representation of an embodiment for setting the gap height along the axial extent of the gas turbine engine.

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

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

In 1 and 2 are shovels 41 of the compressor stages 14th , 15th and the turbine stages 17th , 19th shown schematically. The blades become radially outward 41 from a housing 42 enclosed. Between the tip of the blades 41 and the case 42 a gap forms with the height H off (see 2 ). This gap should be kept as small as possible. Embodiments which accomplish this are described below.

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

It should be noted that the terms “low-pressure turbine” and “low-pressure compressor”, as used here, can be understood to mean the turbine stage with the lowest pressure and the compressor stage with the lowest pressure, respectively (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 writings, 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 3 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 within the scope of the claimed invention, more or fewer planetary gears 32 can be provided. Practical applications of an epicyclic planetary gear 30th generally include at least three planetary gears 32 .

This in 2 and 3 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 external 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.

It goes without saying that the in 2 and 3 The arrangement shown is exemplary only and various alternatives are within the scope of the present disclosure. Any suitable arrangement for positioning the 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 2 ) between the gearbox 30th and other parts of the gas turbine engine 10 (such as the input shaft 26th , the output shaft and the established structure 24 ) have some degree of rigidity or flexibility. As another example, any suitable arrangement of the Bearing between rotating and stationary parts of the gas turbine engine 10 (for example, between the input and output shafts of the transmission and the specified structures, such as the transmission housing), and the disclosure is not limited to the exemplary arrangement of FIG 2 limited. 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 star arrangement (described above) of the transmission 30th usually by those who exemplify in 2 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 connecting shafts. As another example, the 1 gas turbine engine shown has a split flow nozzle 20th , 22nd on, which means that the flow 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 turbo engine, the disclosure may be applied to any type of gas turbine engine such as a gas turbine engine. B. in an open rotor (in which the fan stage is not surrounded by an engine nacelle) or a turboprop engine. In some arrangements, the gas turbine engine comprises 10 possibly no transmission 30th .

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 1 ) and a circumferential direction (perpendicular to the view in 1 ) includes. The axial, radial and circumferential directions are perpendicular to one another.

As mentioned above, the gap height should H between the tips of the blades 41 and the case 42 be kept as small as possible so that the gas turbine engine 10 has the highest possible efficiency. For the sake of simplicity, the case is 42 of the compressor stages 14th , 15th and the turbine stages 17th , 19th considered here as a unit, although a much more complex structure may be required.

That the turbine stages 17th , 19th and / or the compressor stages 14th , 15th surrounding housing 42 should be in the gap between the blade tips and the housing 42 adjusted, in particular by the gap distances H to minimize.

The relevant part of the housing should 42 be cooled, regulating the gap height H particularly efficient.

Some embodiments are described below for this purpose.

When there is no gap between the tip of the shovel 41 and the case 42 exists, what is known as rubbing occurs, ie there is direct contact between the tip of the blade 41 and the case 42 .

In the event of such a contact, a rubbing noise is generated which has characteristic properties and e.g. can be recognized on the basis of a measured frequency spectrum. An evaluation can then take place together with the repetition frequency of the rubbing (speed multiplied by the number of blades of the relevant turbine section and / or compressor section).

Owing to thermal inhomogeneities, the rubbing is often not evenly distributed over the entire circumference, but always at specific geometric locations.

The location of the grazing events can be within the gas turbine engine 10 be localized by evaluating the acoustic events.

A plurality of structure-borne noise sensors are used for this purpose 1 , 2 , 3 , 4th each one sector A1 , A2 , A3 , A4 on the circumference of a turbine stage 17th , 19th and / or a compressor stage 14th , 15th of the gas turbine engine 10 are arranged. This is in 4th shown schematically in a front view.

The circle shows 43 the envelope of the blades 41 in the ideal shape, with the blades 41 are not shown here for the sake of simplicity. With the gap height H The housing is then radially spaced from this 42 .

In the embodiment shown, the circumference of the cross section is uniform in the four sectors A1 , A2 , A3 , A4 divided up. Each of these is one of the structure-borne noise sensors 1 , 2 , 3 , 4th to arranged, these being set up and designed so that a rotating blade can rub against it 41 the turbine stage 17th , 19th and / or the compressor stage 14th , 15th based on at least one structure-borne sound signal K is detectable. In the 4th is a structure-borne noise event K due to a grazing in the sector A1 shown. In principle, it is possible that there is more than one structure-borne sound signal around the circumference K present.

The structure-borne sound sensors 1 , 2 , 3 , 4th can be designed as piezoelectric elements, for example, which react very sensitively to structure-borne noise, ie vibrations of certain frequencies.

With a measuring and control unit 5 , e.g. a computer inside the gas turbine engine 10 , is based on the structure-borne noise signal K the locus of the brushing of the tips of the blades 41 on the housing 42 determinable. The structure-borne sound sensor 1 of the first sector A1 For example, a stronger acoustic signal is received than the opposite structure-borne sound sensor 3 in the sector A3 .

Due to the different distances of the structure-borne sound event K to the four structure-borne noise sensors 1 , 2 , 3 , 4th can be the location of the structure-borne noise due to the known geometric dimensions K on the perimeter of the gas turbine engine 10 to be determined. For example, the one at the structure-borne noise sensors 1 , 2 , 3 , 4th measured structure-borne noise each have a different amplitude level. This depends on the distance between the contact position and the structure-borne sound sensors 1 , 2 , 3 , 4th . The position of the rubbing on the circumference of the housing can be determined from the ratio of the signal amplitudes of the structure-borne noise sensors to one another.

As the compressor stages 14th , 15th and the turbine stages 17th , 19th can also extend in the axial direction, via an axial arrangement of structure-borne sound sensors 1 , 2 , 3 , 4th not only a determination of the rubbing event in the circumferential direction, but also alternatively or additionally in the axial direction. This is based on an embodiment according to 6th are explained in more detail.

In any case, depending on the results of the structure-borne noise determination and analysis, at least one temperature control device is used C1 , C2 , C3 , C4 activated that den at least one sector A1 , A2 , A3 , A4 is assigned so that the gap height H can be set locally to a setpoint.

By evaluating structure-borne noise events K is not a direct measurement of the gap height H required.

In 5 will affect the function of the temperature control devices C1 , C2 , C3 , C4 , C5 , C6 , C7 , C8 more precisely, with eight instead of four sectors A1 , A2 , A3 , A4 , A5 , A6 , A7 , A8 around the perimeter of the case 42 are arranged. The measuring and control unit 5 and the connections to the sensor devices 1 , 2 , 3 , 4th , 1a , 1b , 1c , 1c and the temperature control devices C1 , C2 , C3 , C4 , C5 , C6 , C7 , C8 are not shown here for reasons of clarity.

Every sector has A1 , A2 , A3 , A4 , A5 , A6 , A7 , A8 one temperature control device each C1 , C2 , C3 , C4 , C5 , C6 , C7 , C8 with an associated valve system 7th that the respective air supply according to the required temperature in this sector C1 , C2 , C3 , C4 , C5 , C6 , C7 , C8 and thus the expansion or contraction of the structures in the housing 42 can cause. Cold or warm air can be used, depending on the required expansion or contraction.

The case 42 is through the over the valve systems 7th Regulated supply of colder air changed in diameter. With the housing diameter, the gap height can be determined H of the turbine blades 41 to the housing 42 adjusted.

In 5 is for the sector A1 for example the incoming air I. via the valve system not shown in detail here 7th is controlled, and the exiting air O designated. In an alternative embodiment, the exiting air can O via a valve system 7th to be controlled.

In the case 42 structure-borne noise sensors are installed at certain intervals on the circumference 1 , 2 , 3 , 4th , 1a , 2a, 3a, 4a (preferably piezo elements, piezoelectric, piezoresistive elements or strain gauges).

The one at the structure-borne noise sensors 1 , 2 , 3 , 4th , 1a , 2a, 3a, 4a measured structure-borne noise each has a different amplitude level. This depends on the distance between the contact position and the structure-borne sound sensors. The position of the rubbing on the circumference of the housing can be determined from the ratio of the signal amplitudes of the structure-borne noise sensors to one another.

In one embodiment, an adaptive algorithm can be used. This is the diameter of the case 42 as long as changed by increasing a cooling air flow until in one or more sectors A1 , A2 , A3 , A4 , A5 , A6 , A7 , A8 a rubbing occurs. If a rubbing process is perceived at one point, ie there is a structure-borne noise event K , then the supply of colder air to the sector of the housing part is reduced, so that this housing section is expanded. The reduction of the cooling air is continued until the rubbing stops, ie no structure-borne sound events K more to be registered. Usually, this value of the rubbing is stored with the associated variables and the values are used for the future setting of the cooling air flows. A value is then set that maintains a predefined distance from the rubbing.

The case diameter is enlarged on the basis of the learned data (in the sense of machine learning) from the system states of the gas turbine engine 10 when brushing the tips of the blades 41 . The settings should be optimized for constant conditions, especially cruising.

By adjusting the different sectors A1 , A2 , A3 , A4 , A5 , A6 , A7 , A8 it is possible the surrounding case 42 "To pull round" and thus to achieve an additional increase in efficiency, which would not be achievable with a total adjustment (one valve) for the entire circumference of the housing.

The embodiments of the devices and methods can also be used for the hot zones of the compressor stages 14th , 15th be applied.

When brushing the tips of the blades 41 , e.g. a turbine blade on one edge of the surrounding housing 42 are the effective blade lengths and the effective diameter over the circumference of the housing due to the not precisely known temperatures 42 not precisely determinable.

The surrounding case 42 (ie the sector concerned) is changed in diameter with cooling air so that the distance to the blade tips is minimal. A brush against the blades 41 should be excluded in nominal operation. Therefore, the rubbing situations are evaluated with the measured system states of the engine and the conditions of the required diameter of the housing are influenced in such a way that rubbing does not occur in nominal operation.

Learning about the rubbing situations with the corresponding system states (state variables) can take place via targeted rubbing on the test stands or during flight tests. These parameters can be determined either typically for an engine series or once for an individual gas turbine engine 10 or, alternatively, several times in the life of an individual engine, in order to compensate for the drop in performance over the aging process.

If the turbine blade tips are rubbing against the surrounding housing 42 recognized, the contact events and the contact intensity are recognized together with the state variables of the gas turbine engine 10 stored in tables. The state variables of the gas turbine engine are partly measured directly via sensors, and partly calculated from models.

Table entries are used to set the necessary cooling air flow (valve position) to set certain gap heights H the turbine blade tips to the housing 42 to reach.

The table entries are determined on the basis of the recognized stripes, taking into account the system states of the gas turbine engine that exist at the time of the strikes.

As already mentioned above, other embodiments can also use structure-borne sound signals K capture in the axial direction. In 6th is an exemplary and schematic perspective view of part of a housing 42 shown, for example, rotating blades 41 a turbine stage 17th , 19th surrounds. The case 42 is shown here for simplicity as a circular cylinder, whereby in real gas turbine engines 10 the shape of the case 42 is regularly more complex

On the case 42 are sectors A1 , A2 , A3 shown, which are arranged axially one behind the other. In every sector A1 , A2 , A3 is a structure-borne sound sensor 1 , 2 , 3 arranged, in an analogous manner to a circumferential arrangement, the occurrence of structure-borne noise signals K can localize in the axial direction.

A combination of an arrangement of structure-borne sound sensors 1 , 2 , 3 , 4th , 1a , 1b , 1c , 1d in the circumferential direction and in the axial direction allows a spatial detection of structure-borne sound signals K in the gas turbine engine 10 . If at the same time also the temperature control devices C1 , C2 , C3 , C4 , C5 , C6 , C7 , C8 are arranged in the circumferential direction and in the axial direction (in 6th not shown), a complex spatial detection and adjustment of the gap height can also be carried out H implement,

It goes without saying that the invention does not apply to the embodiments described above is limited and various modifications and improvements can be made without departing from the concepts described herein. Any of the features can be used separately or in combination with any other features, provided that they are not mutually exclusive, and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

List of reference symbols

1-4, 1a-1d

Structure-borne sound sensor

5

Measuring and control unit

6th

adaptive agent

7th

Valve system

9

Main axis of rotation

10

Gas turbine engine

11

Core engine

12

Air inlet

14th

Low pressure compressor

15th

High pressure compressor

16

Incinerator

17th

High pressure turbine

18th

Bypass thrust nozzle

19th

Low pressure turbine

20th

Core thruster

21st

Engine nacelle

22nd

Bypass duct

23

fan

24

stationary support structure

26th

wave

27

Connecting shaft

28

Sun gear

30th

transmission

32

Planetary gears

34

Planet carrier

36

Linkage

38

Ring gear

40

Linkage

41

Shovels

42

casing

43

Envelope of the blades (idealized)

A.

Core airflow

B.

Bypass airflow

A1-A8

Sector on the perimeter of the housing

C1-C8

Temperature control device of a sector

I.

Incoming air as temperature control medium

H

Gap height

K

Structure-borne sound signal

L.

Cooling air

O

outflowing air as temperature control medium

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

• WO 2008/060164 A1 [0002]
• DE 3231587 A1 [0002]

Claims (17)

Device for setting a gap height (H) between the tips of blades (41) of a turbine stage (17, 19) and / or a compressor stage (14, 15) and a housing (42) surrounding the blades (41) in a gas turbine engine (10 ), characterized by a plurality of structure-borne noise sensors (1, 2, 3, 4, 1a, 2a, 3a, 4a), each of which is assigned to a sector (A1, A2, A3, A4, A5, A6, A7, A8) on and / or a turbine stage (17, 19) and / or a compressor stage (14, 15) of the gas turbine engine (10) are arranged in the housing (42), the structure-borne sound sensors (1, 2, 3, 4, 1a, 2a, 3a, 4a ) are set up and designed so that a rubbing against a rotating blade (41) of the turbine stage (17, 19) and / or the compressor stage (14, 15) can be detected on the basis of at least one structure-borne noise signal (K) and a measuring and control unit ( 5), which uses the at least one structure-borne noise signal (K) to determine the geometric location of the rubbing against the tips of the blades (41) can be determined on the housing (42) and, depending on this, at least one temperature control device (C1, C2, C3, C4, C5, C6, C7, C8) can be activated, which corresponds to at least one sector (A1, A2, A3, A4, A5, A6, A7, A8) is assigned so that the gap height (H) can be adjusted locally to a target value. Device according to Claim 1 , characterized in that the at least one temperature control device (C1, C2, C3, C4, C5, C6, C7, C8) each have a valve system (7) for air (L) as the temperature control medium, which can be regulated by the measuring and control unit (5) Have cooling or heating. Device according to Claim 1 or 2 , characterized in that the structure-borne sound sensors (1, 2, 3, 4, 1a, 2a, 3a, 4a) each have a piezo element, in particular a piezoelectric element, a piezoresistive element and / or a strain gauge. Device according to at least one of the preceding claims, characterized in that the sectors (A1, A2, A3, A4, A5, A6, A7, A8), the structure-borne noise sensors (1, 2, 3, 4, 1a, 2a, 3a, 4a ) and / or at least two temperature control devices (C1, C2, C3, C4, C5, C6, C7, C8) are arranged regularly around the circumference of the housing (42). Device according to at least one of the preceding claims, characterized in that the sectors (A1, A2, A3, A4, A5, A6, A7, A8), the structure-borne sound sensors (1, 2, 3, 4, 1a, 2a, 3a, 4a ) and / or the at least two temperature control devices (C1, C2, C3, C4, C5, C6, C7, C8) are arranged axially along the housing (42). Device according to at least one of the preceding claims, characterized in that the measuring and control unit (5) has a means for analyzing the frequency spectrum of the at least one structure-borne noise signal (K) and / or for detecting the speed of the turbine stage (17, 19) and / or the compressor stage (14, 15). Device according to at least one of the preceding claims, characterized in that the measuring and control unit (5) has a model for setting the gap height (H), which the speed of a turbine stage (17, 19) and / or compressor stage (14, 15) , a temperature, in particular an inlet temperature of a turbine stage (17, 19), a pressure, in particular an inlet pressure of a turbine stage (17, 19), a valve position of the at least one temperature control device (C1, C2, C3, C4, C5, C6, C7, C8), the number of blades of a turbine stage (17, 19) and / or compressor stage (14, 15), material expansion coefficients of a blade (41) and / or of the housing (42) and / or heat exchange surfaces are taken into account. Device according to at least one of the preceding claims, characterized in that the measuring and control unit (5) has an adaptive means (6) for adjusting the gap height (H), which the at least one temperature control device (C1, C2, C3, C4, C5 , C6, C7, C8) activated until no structure-borne sound signal (K) or a predefined structure-borne sound signal (K) can be detected. Device according to Claim 8 , characterized in that the gap height (H) is locally adjustable with the adaptive means (6) so that a predetermined geometry of the gap, in particular the cross-sectional area of a circular ring around the circumference and / or in the axial extent, can be set. Device according to Claim 8 or 9 , characterized in that the adaptive means (6) has a machine learning algorithm. Method for setting a gap height (H) between the tips of blades (41) of a turbine stage (17, 19) and / or a compressor stage (14, 15) and a housing (41) surrounding the blades (41) in a gas turbine engine (10 ), characterized in that a) a plurality of structure-borne sound sensors (1, 2, 3, 4, 1a, 2a, 3a, 4a), each of which corresponds to a sector (A1, A2, A3, A4, A5, A6, A7, A8 ) on and / or in the housing (42) of a turbine stage (17, 19) and / or a compressor stage (14, 15) of the gas turbine engine (10) are assigned, the rubbing of a rotating blade of the turbine stage (17, 19) and / or the compressor stage (14, 15) using at least one structure-borne noise signal (K), b) the geometric location of the contact between the tips of the blades (41) and the housing (42) is determined by a measuring and control unit (5), and c) at least one temperature control device (C1, C2, C3, C4, C5) as a function thereof , C6, C7, C8) is activated, which is assigned to at least one sector (A1, A2, A3, A4, A5, A6, A7, A8) so that the gap height (H) is set locally to a target value. Procedure according to Claim 11 , characterized in that the control unit (5) calculates a frequency spectrum of the at least one structure-borne noise signal (K) and / or detects the speed of the turbine stage and / or the compressor stage. Procedure according to Claim 11 or 12th , characterized in that the measuring and control unit (5) has a model for setting the gap height (H), the speed of a turbine stage (17, 19) and / or compressor stage (14, 15), a temperature, in particular an inlet temperature a turbine stage (17, 19), a pressure, in particular an inlet pressure of a turbine stage (17, 19), a valve position of the at least one temperature control device (C1, C2, C3, C4, C5, C6, C7, C8), the number of blades a turbine stage (17, 19) and / or compressor stage (14, 15), material expansion coefficients of a blade (41) and / or of the housing (42) and / or heat exchange surfaces are taken into account. Method according to at least one of the Claims 11 to 13 , characterized in that the measuring and control unit (5) has an adaptive means (6) for adjusting the gap height (H), which the at least one temperature control device (C1, C2, C3, C4, C5, C6, C7, C8) activated until no structure-borne sound signal (K) or a predefined structure-borne sound signal (K) can be detected. Procedure according to Claim 14 , characterized in that the gap height (H) is locally adjustable with the adaptive means (6) so that a predetermined geometry of the gap, in particular a cross-sectional area of a circular ring around the circumference and / or in the axial extent, can be set. Procedure according to Claim 14 or 15th , characterized in that the adaptive means (6) has a machine learning algorithm. A gas turbine engine (10) for an aircraft comprising: a core engine (11) comprising a turbine (19), a compressor (14) and a core shaft (26) connecting the turbine to the compressor; a fan (23) positioned upstream of the core engine (11), the fan (23) comprising a plurality of fan blades; and a gear (30) that can be driven by the core shaft (26), wherein the fan (23) can be driven by means of the gear (30) at a lower speed than the core shaft (26), with a device for setting a gap height ( H) with the characteristics of Claims 1 to 10 .

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