DE102020101324A1 - Assembly in a gas turbine engine and method for detecting failure of a thrust bearing - Google Patents

Abstract

The invention relates to an assembly in a gas turbine engine and a method for detecting a failure of a thrust bearing. The assembly comprises a stator (5), a rotor (6), an interstage seal (7) and an interstage cavity (8). The interstage seal (7) comprises a radially outwardly directed groove (71) which receives a rail (54) of the stator (5). It is provided that the groove (71) of the interstage seal (7) has an axial width (713) which is greater than the axial width of the rail (54). This has the effect that when the interstage seal (7) is moved axially forwards, the axially front groove wall (711) comes at an axial distance from the rail (54) and the groove (71) forms an axial gap (714) through which a gas stream (106) can flow from the intermediate stage cavity (8) into the main flow path (4), whereby a pressure drop occurs in the intermediate stage cavity (8), the pressure drop or a temperature rise triggered by it being measured.

Description

The invention relates to an assembly in a gas turbine engine according to the preamble of claim 1 and a method for detecting a failure of a thrust bearing.

The failure of one or more of the thrust bearings that support the high pressure turbine shaft of the gas turbine engine can fail in gas turbine engines. In such a case, the rotors of the high pressure turbine move upstream, which can result in contact between rotating and static components of the high pressure turbine. The aim is to recognize such a situation quickly in order to be able to initiate suitable countermeasures.

There are also non-contact seals known that seal a gap between a rotor and a stator of a gas turbine engine. It is from the DE 10 2015 226 732 A1 known to integrate a temperature sensor in such a seal which detects the temperature of a gas flowing through the seal.

The present invention is based on the object of providing an assembly in a gas turbine engine which makes it possible to detect the failure of a thrust bearing in a simple manner. Furthermore, a method for the detection of a failure of a drawer bearing is to be provided.

This object is achieved by an assembly with the features of claim 1 and a method with the features of claim 12. Refinements of the invention are given in the dependent claims.

According to a first aspect of the invention, the invention then considers an assembly in a gas turbine engine which comprises a stator and a rotor of a turbine, the rotor being arranged downstream of the stator, the stator having guide vanes and a radially inner platform, the rotor comprising rotating blades and a turbine disk , and the guide vanes and the blades extend in a main flow path of the gas turbine engine. The assembly further includes an interstage seal that forms an axial gap with the turbine disk.

The interstage seal forms a radially outwardly directed groove which receives a radially inwardly extending rail of the radially inner platform, the groove having an axially forward, radially extending groove wall and an axially rearward, radially extending groove wall. The axially front groove wall is provided and designed to lie tightly against the radially extending rail during normal operation of the gas turbine engine, so that no gas can flow through the groove. There is also an intermediate stage cavity formed between the rotor disk of a further rotor arranged upstream of the stator and the intermediate stage seal, the intermediate stage cavity being acted upon by secondary air with a static pressure during normal operation of the gas turbine engine which is greater than the static pressure prevailing in the main flow path of the gas turbine engine Print.

It is provided that the groove of the interstage seal has an axial width which is greater than the axial width of the rail extending radially inward. This means that when the rotor moves axially forwards (as is the case, for example, in the event of the failure of a thrust bearing) and the interstage seal is also moved axially forwards, the axially front groove wall is at an axial distance from the radial one extending rail device and the groove forms an axial gap through which a gas stream can flow from the interstage cavity into the main flow path downstream of the stator. This causes a pressure drop in the interstage cavity. The assembly further comprises means which measure the pressure drop in the interstage cavity or a temperature rise triggered by this, the measurement being able to take place directly or indirectly.

The present invention is based on the idea of designing the interstage cavity in such a way that the pressure conditions in the interstage cavity change in the event of an axial upstream displacement of the rotor, this change in the pressure conditions or a temperature rise triggered by this change in the interstage cavity being measured and evaluated as a parameter indicating the failure of a drawer bearing.

A change in the pressure conditions in the intermediate stage cavity is based on the fact that, when the rotor and thus the intermediate stage cavity are displaced axially upstream, an axial gap is opened in the area of the groove, through which a gas stream flows from the intermediate stage cavity into the main flow path downstream of the stator, which leads to leads to a pressure drop in the interstage cavity. The pressure drop in the interstage cavity in turn causes hot gas to flow from the main flow path upstream of the stator into the interstage cavity, since after the pressure drop in the interstage cavity the pressure in the main flow path upstream of the stator is greater than the pressure in the Interstage cavity. The pressure drop in the interstage cavity can be measured by a pressure sensor or, alternatively, a temperature rise caused by the hot gas flowing into the interstage cavity from the main flow path can be measured.

Accordingly, one embodiment of the invention provides that the means comprise a temperature sensor, with hot gas flowing from the main flow channel into the interstage cavity in the event of a pressure drop in the interstage cavity and the temperature sensor being arranged in such a way that it measures a temperature increase associated therewith.

For example, it can be provided that the temperature sensor is designed to directly measure the temperature of the gas that flows through a contact-free seal that is arranged in a gap between the radially inner platform and the further rotor arranged upstream of the stator and which separates the interstage cavity from the main flow path. The gas flows through the non-contact seal in one or the other direction, depending on the pressure ratio. In the event of a pressure drop in the interstage cavity, hot gas flows from the main flow channel into the interstage cavity, the temperature sensor measuring the temperature of the hot gas.

The temperature sensor can be connected to evaluation means which evaluate the temperature of the temperature sensor to detect the failure of a thrust bearing. It can be provided, for example, that the failure of a thrust bearing is recognized when the measured temperature exceeds a predefined threshold value.

The temperature sensor is arranged, for example, in the radially inner platform of the stator.

It is pointed out that, in the context of the present disclosure, the term “turbine disk” is used both to designate the actual turbine disk and to designate a blade root with which the respective rotor blade is connected to the turbine disk. In this respect, there is no differentiation.

Another embodiment of the invention provides that the interstage seal has a downstream, radially extending rear side which is arranged such that it is axially spaced from the turbine disk of the rotor during normal operation of the gas turbine engine. The gap that extends between the rear side and the turbine disk during normal operation of the gas turbine engine forms part of the gap between the stator and rotor. It is provided that the rear side of the interstage seal is provided with wedge-shaped circumferential grooves. The depth of the circumferential grooves thus decreases in the circumferential direction until it runs out into the rear of the interstage seal, whereupon a new circumferential groove begins.

The wedge-shaped circumferential grooves are provided and designed to build up a dynamic pressure when the turbine disk approaches the rear side - which occurs when the rotor moves overall upstream due to the failure of a thrust bearing - which creates friction between the turbine disk and the rear side of the interstage seal reduced. In a sense, an air bearing is provided between the turbine disk and the rear.

One embodiment provides that the air required to build up a dynamic pressure is provided via existing pre-swirl nozzles, which are designed as bores in the rear, with air entering the circumferential grooves through the pre-swirl nozzles. It can be provided that the pre-swirl nozzles are arranged in such a way that air enters the circumferential grooves through the pre-swirl nozzles at the deepest points of the circumferential grooves.

A further embodiment provides that, in order to reduce friction between the turbine disk and the rear side of the interstage seal, the rear side and / or the turbine disk are provided with a friction-reducing coating in the potential contact area.

According to a further aspect of the invention, the present invention relates to a gas turbine engine with an assembly according to claim 1. An embodiment provides that the gas turbine engine comprises a high-pressure turbine which has a first stage and a second stage, each stage having a stator and a rotor, and wherein the interstage seal of the assembly is disposed between the stator and the rotor of the second stage of the high pressure turbine.

• - Beaufschlagen der Zwischenstufenkavität mit Sekundärluft derart, dass im Normalbetrieb des Gasturbinentriebwerks der in der Zwischenstufenkavität anliegende statische Druck größer ist als der im Hauptströmungspfad des Gasturbinentriebwerks stromaufwärts des Stators herrschende statische Druck,
• - fortlaufendes Messen des in der Zwischenstufenkavität herrschenden Druckes und/oder der dort herrschenden Temperatur,
• - durch das Messen Erkennen eines Druckabfalls und/oder eines Temperaturanstiegs in der Zwischenstufenkavität, wenn ein solcher vorliegt,
• - Erkennen des Versagens eines Schublagers, wenn der Druckabfall und/oder der Temperaturanstieg einen jeweils zugeordneten Grenzwert unterschreitet bzw. überschreitet.

In a further aspect of the invention, the present invention relates to a method for detecting a failure of a thrust bearing that supports a turbine shaft of a gas turbine engine, the method being carried out using an assembly according to claim 1. The procedure consists of the following steps:

• - Applying secondary air to the intermediate stage cavity in such a way that, during normal operation of the gas turbine engine, the static pressure present in the intermediate stage cavity is greater than the static pressure prevailing in the main flow path of the gas turbine engine upstream of the stator,
• - Continuous measurement of the pressure in the interstage cavity and / or the temperature there,
• - by measuring the detection of a pressure drop and / or a temperature increase in the interstage cavity, if such is present,
• - Detection of the failure of a thrust bearing when the pressure drop and / or the temperature rise falls below or exceeds a respectively assigned limit value.

One embodiment of the method provides that, in the event of a pressure drop in the intermediate stage cavity, hot gas flows from the main flow channel into the intermediate stage cavity and a temperature sensor measures a temperature rise triggered by this.

One embodiment for this provides that the temperature sensor measures the temperature of the gas flowing through a seal which is arranged in a gap between the radially inner platform and the further rotor of the assembly arranged upstream of the stator.

Another embodiment of the method provides that when the turbine disk approaches the rear of the interstage seal, friction between the turbine disk and the rear is reduced by blowing secondary air into the wedge-shaped circumferential grooves formed in the rear of the interstage seal. It can be provided that the secondary air is blown in through pre-swirl nozzles which are formed in the rear of the interstage seal.

It should be noted that the present invention is described with reference to a cylindrical coordinate system which has the coordinates x, r and φ. Here x indicates the axial direction, r the radial direction and φ the angle in the circumferential direction. The axial direction is identical to the machine axis of the gas turbine engine, the axial direction pointing from the engine inlet in the direction of the engine outlet. Starting from the x-axis, the radial direction points radially outwards. Terms such as “in front”, “behind”, “front” and “rear” relate to the axial direction or the direction of flow in the engine. Terms like “outer” or “inner” refer to the radial direction.

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

Arrangements of the present disclosure may be particularly, but not exclusively, advantageous for fans that are driven via a transmission. Accordingly, the gas turbine engine may include a transmission that receives an input from the core shaft and provides drive for the fan to drive the fan at a lower speed than the core shaft. The input for the transmission can take place directly from the core shaft or indirectly from the core shaft, for example via a spur shaft and / or a spur gear. The core shaft may be rigidly connected to the turbine and the compressor so that the turbine and the compressor rotate at the same speed (with the fan rotating at a lower speed).

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

With such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor may be arranged to receive flow from the first compressor (e.g., receive directly, e.g., via a generally annular channel).

The gearbox can be arranged to be driven by the core shaft configured to rotate (e.g. in use) at the lowest speed (e.g. the first core shaft in the above example). For example, the gearbox can be arranged to only rotate from the core shaft configured to rotate (e.g. in use) at the lowest speed (e.g. only from the first core shaft and not the second Core shaft in the above example) to be driven. Alternatively, the transmission can be arranged to be driven by one or more shafts, for example the first and / or the second shaft in the above example.

In a gas turbine engine described and / or claimed here, a combustion chamber may be provided axially downstream of the fan and the compressor (s). For example, the combustion chamber 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 combustion chamber 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 (in that their angle of attack can be variable). The row of rotor blades and the row of stator blades can be axially offset from one another.

The or each turbine (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 be defined with a radial span extending from a root (or hub) at a radially inward gas overflow location or at a 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of): 0.4, 0.39, 0.38, 0.37, 0.36, 0, 35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26 or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip can be in an inclusive range bounded by two of the values in the preceding sentence (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 portion (or axially most forward edge) of the blade. The hub-to-tip ratio, of course, relates to the portion of the fan blade overflowing with gas; H. the portion that is radially outside of any platform.

The radius of the fan can be measured between the centerline of the engine and the tip of the fan blade at its leading edge. The diameter of the fan (which can be simply twice the radius of the fan) can be greater than (or on the order of): 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), or 390 cm (about 155 inches). The fan diameter can be in an inclusive range 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 with use. In general, the speed is lower for fans with a larger diameter. By way of non-limiting example only, the speed of the fan under constant speed conditions may be less than 2500 rpm, for example less than 2300 rpm. Merely as a further non-limiting example, the speed of the fan under constant speed conditions for an engine with a fan diameter in the range from 250 cm to 300 cm (for example 250 cm to 280 cm) in the range from 1700 rpm to 2500 rpm, for example in the range from 1800 rpm to 2300 rpm, for example in the range from 1900 rpm to 2100 rpm. Merely as a further non-limiting example, the speed of the fan under constant speed conditions for an engine with a fan diameter in the range from 320 cm to 380 cm in the range from 1200 rpm to 2000 rpm, for example in the range from 1300 rpm min to 1800 rpm, for example in the range from 1400 rpm to 1600 rpm.

When the gas turbine engine is in use, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tip of the fan blade to move _{at a speed 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, for example at the front edge of the tip (which can be defined as the fan tip radius at the front edge multiplied by the angular speed). The fan peak load at constant speed conditions can be more than (or on the order of): 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38 , 0.39, or 0.4 (where all units in this section are Jkg ^-1 K ^-1 / (ms ^-1 ) ² ). The fan peak load can be in an inclusive range bounded by two of the values in the preceding sentence (ie, the values can be upper or lower limits).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, the bypass ratio being defined as the ratio of the mass flow rate of 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 (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 can be in an inclusive 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 engine core. The radially outer surface of the bypass duct can be defined by an engine nacelle and / or a fan housing.

The total pressure ratio of a gas turbine engine described and / or claimed herein can be 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 combustion chamber). As a non-limiting example, the total pressure ratio of a gas turbine engine described and / or claimed herein at constant speed may be greater than (or on the order of): 35, 40, 45, 50, 55, 60, 65, 70, 75 (lie). The total pressure ratio can be in an inclusive range bounded by two of the values in the preceding sentence (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. Under constant speed conditions, the specific thrust of an engine described and / or claimed here can be less than (or in the order of magnitude of): 110 Nkg ^-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. The specific thrust can lie in an inclusive range which is limited by two of the values in the preceding 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): 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN , 450kN, 500kN or 550kN. The maximum thrust can be in an inclusive range bounded by two of the values in the preceding sentence (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 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C) with a static engine.

In use, the temperature of the flow at the inlet of the high pressure turbine can be particularly high. This temperature, which can be referred to as TET, can be measured at the exit to the combustion chamber, for example immediately upstream of the first turbine blade, which in turn can be referred to as a nozzle guide vane. At constant speed, the TET can be at least (or on the order of): 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at constant speed can be in an inclusive range bounded by two of the values in the preceding sentence (i.e., the values can be upper or lower limits). The maximum TET when the engine is in use can, for example, be at least (or on the order of): 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET can be in an inclusive 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 a blade portion of a fan blade described and / or claimed herein can be made from any suitable material or combination Materials are made. For example, at least a portion of the fan blade and / or the blade can be made at least in part of a composite, for example a metal matrix composite and / or an organic matrix composite, such as e.g. B. carbon fiber. As another example, at least a portion of the fan blade and / or the blade can be at least in part made of a metal, such as metal. A titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade can include at least two sections made using different materials. For example, the fan blade may have a leading edge that is made using a material that can withstand impact (such as birds, ice, or other material) better than the rest of the blade. Such a leading edge can be made using titanium or a titanium-based alloy, for example. Thus, by way of example only, the fan blade may have a carbon fiber or aluminum based body (such as an aluminum-lithium alloy) with a leading edge made of titanium.

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

The gas turbine engines 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 exit cross-section of the bypass channel to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine described and / or claimed herein can have any desired number of fan blades, for example 16, 18, 20 or 22 fan blades.

As used herein, constant speed conditions may mean constant speed conditions of an aircraft on which the gas turbine engine is mounted. Such constant speed conditions can conventionally be defined as the conditions during the middle part of the flight, for example the conditions to which the aircraft and / or the engine are exposed between (in terms of time and / or distance) the end of the climb and the start of the descent. 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 travel condition. For some aircraft, the cruise control conditions 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 standard atmospheric conditions at an altitude that is in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (about 38,000 feet), for example in Range from 10,500 m to 11,500 m, for example in the range from 10,600 m to 11,400 m, for example in the range from 10,700 m (about 35,000 feet) to 11,300 m, for example in the range from 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example in the order of 11,000 m, correspond. The constant velocity conditions can correspond to standard atmospheric conditions at any given altitude in these areas.

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

As they are used throughout here, “constant speed” or “constant speed conditions” can mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (including, for example, Mach number, environmental conditions, and thrust requirement) for which the blower is designed to operate. 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 use, a gas turbine engine described and / or claimed herein can be operated at the constant speed conditions defined elsewhere herein. Such constant speed conditions may be determined by the constant speed conditions (e.g., the conditions during the middle part of flight) of an aircraft to which at least one (e.g. 2 or 4) gas turbine engine may be attached to provide thrust.

It will be understood by those skilled in the art that a feature or parameter described in relation to any of the above aspects can be applied to any other aspect, provided that they are not mutually exclusive. Furthermore, any feature or parameter described here can be applied to any aspect and / or combined with any other feature or parameter described here, insofar as they are compatible not mutually exclusive.

• 1 eine Seitenschnittansicht eines Gasturbinentriebwerks;
• 2 eine Seitenschnittgroßansicht eines stromaufwärtigen Abschnitts eines Gastu rbi nentriebwerks;
• 3 eine zum Teil weggeschnittene Ansicht eines Getriebes für ein Gastu rbi nentriebwerk;
• 4 eine Baugruppe eines Gasturbinentriebwerks, die einen Stator, einen stromabwärts des Stators angeordneten Rotor und eine Zwischenstufendichtung umfasst, wobei die Anordnung der Elemente der Baugruppe der Anordnung im Normalbetrieb des Gasturbinentriebwerks entspricht;
• 5 die Baugruppe der 4, wobei der Rotor nach einem Versagen eines Schublagers stromaufwärts gewandert ist, so dass eine Turbinenscheibe des Rotors in Kontakt mit einer Rückseite der Zwischenstufendichtung getreten ist;
• 6 ein Ausführungsbeispiel einer erfindungsgemäßen Baugruppe eines Gasturbinentriebwerks, wobei die Baugruppe einen Stator, einen stromabwärts des Stators angeordneten Rotor und eine Zwischenstufendichtung umfasst, wobei die Zwischenstufendichtung eine nach außen gerichtete radiale Nut aufweist, in die eine Schiene des Stators eingreift, wobei die Nut der Zwischenstufendichtung eine axiale Breite aufweist, die größer ist als die axiale Breite der Schiene, und wobei die Anordnung der Elemente der Baugruppe der Anordnung im Normalbetrieb des Gasturbinentriebwerks entspricht;
• 7 die Baugruppe der 6, wobei der Rotor nach einem Versagen eines Schublagers stromaufwärts gewandert ist, wodurch ein axialer Spalt im Bereich der Nut geöffnet ist;
• 8 einen Schnitt durch die Baugruppe der 7 entlang der Linie A-A der 7;
• 9 die Baugruppe der 6 unter zusätzlicher Darstellung der Drücke, die in der Baugruppe vorliegen, und daraus resultierender Strömungen;
• 10 die Baugruppe der 7 unter zusätzlicher Darstellung der Drücke, die in der Baugruppe nach einem Wandern des Rotors stromaufwärts vorliegen, und daraus resultierender Strömungen; und
• 11 den Ablauf eines Verfahrens zur Erkennung eines Versagens eines Schublagers.

The invention is explained in more detail below with reference to the figures of the drawing on the basis of several exemplary embodiments. Show it:

• 1 a side sectional view of a gas turbine engine;
• 2 Fig. 3 is a side sectional close-up view of an upstream portion of a gas turbine engine;
• 3 a partially cut-away view of a transmission for a Gasturbi nentriebwerk;
• 4th an assembly of a gas turbine engine comprising a stator, a rotor disposed downstream of the stator and an interstage seal, the arrangement of the elements of the assembly corresponding to the arrangement during normal operation of the gas turbine engine;
• 5 the assembly of the 4th wherein the rotor has migrated upstream after a failure of a thrust bearing such that a turbine disk of the rotor has made contact with a rear side of the interstage seal;
• 6th an embodiment of an assembly according to the invention of a gas turbine engine, the assembly comprising a stator, a downstream of the stator arranged rotor and an interstage seal, wherein the interstage seal has an outwardly directed radial groove in which a rail of the stator engages, the groove of the interstage seal a has an axial width which is greater than the axial width of the rail, and wherein the arrangement of the elements of the assembly corresponds to the arrangement in normal operation of the gas turbine engine;
• 7th the assembly of the 6th wherein the rotor has moved upstream after a failure of a thrust bearing, whereby an axial gap is opened in the region of the groove;
• 8th a section through the assembly of the 7th along the line AA of the 7th ;
• 9 the assembly of the 6th with additional representation of the pressures that are present in the assembly and the resulting flows;
• 10 the assembly of the 7th with additional representation of the pressures which are present in the assembly after the rotor has moved upstream, and the flows resulting therefrom; and
• 11 the sequence of a method for the detection of a failure of a drawer bearing.

1 represents a gas turbine engine 10 with a main axis of rotation 9 represent. The engine 10 includes an air inlet 12th and a thrust fan 23 , which creates two air flows: a core air flow A and a bypass air flow B. The gas turbine engine 10 includes a core 11 , which takes in the core air flow A. The engine core 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 air flow B flows through the bypass duct 22nd . The blower 23 is about a wave 26th and a Epicycloidal gears 30th on the low pressure turbine 19th attached and is driven by this.

In use, the core air flow A is through 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 means of a suitable connecting shaft 27 on. The blower 23 generally provides the majority of the thrust. The epicycloidal gear 30th is a reduction gear.

An exemplary arrangement for a geared blower 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 the epicycloidal gear assembly 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 limits the planet gears 32 on it, in sync with the sun gear 28 to orbit while allowing each planetary gear to move 32 can rotate around its own axis. The planet carrier 34 is about linkage 36 with the blower 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 with it.

It is noted that the terms “low pressure turbine” and “low pressure compressor”, as used here, can be understood to mean the lowest pressure turbine stage and the lowest pressure compressor stage (i.e. not the fan 23 include) and / or the turbine and compressor stage through the connecting shaft 26th with the lowest speed in the engine (meaning that it is not the gearbox output shaft that controls 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”. Using such alternative nomenclature, the fan 23 may be referred to as a first compression stage or the lowest pressure compression stage.

The epicycloidal 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 periphery for 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 epicycloidal gear 30th generally include at least three planetary gears 32 .

This in 2 and 3 epicycloidal gears shown by way of 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 epicycloidal gearing can be used 30th be used. As another example, the epicycloidal 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 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 engine 10 and / or to connect the transmission 30th with the 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 engine 10 (such as the input shaft 26th , the output shaft and the established structure 24 ) have some degree of rigidity or flexibility. As another example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (e.g., between the input and output shafts of the transmission and the fixed structures such as the transmission housing) can be used and the disclosure is not to the exemplary arrangement of 2 limited. For example, it is readily apparent to a person skilled in the art that the arrangement of the output and the support rods and the bearing positions in a star arrangement (described above) of the transmission 30th usually by those who exemplified in 2 would differ.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gear types (e.g., star or planetary), support structures, input and output shaft arrangements, and bearing 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 The 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 fan 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 noted, the high pressure turbine drives 17th the high pressure compressor 15th through a connecting shaft 27 which is also referred to as a high pressure turbine shaft. The high pressure turbine shaft 27 is supported by at least two bearings. One of the camps 270 is schematically in the 1 shown.

In the context of the present invention is the design of the high pressure turbine 17th is of importance, the invention providing an embodiment of an assembly of the high pressure turbine that allows the failure of one or more of the bearings in a simple manner 270 the high pressure turbine shaft 27 to recognize. However, the principles of the present invention apply to any arrangement of a rotor or a stator in a flow path.

the 4th For a better understanding of the background of the present invention, Figure 12 shows an assembly which does not implement the present invention. The illustration is in a sectional plane that also includes the engine axis. The assembly includes elements of a high pressure turbine 17th a gas turbine engine, which is basically any gas turbine engine, such as the gas turbine engine 10 the 1 can act.

The assembly has a stator 5 and a rotor 6th on showing a second stage of the high pressure turbine 17th form. A first stage of the high pressure turbine 17th has a stator (not shown) immediately downstream of a combustion chamber and a rotor 60 on, the upstream of the stator 5 is arranged. The stator 5 includes guide vanes 51 , a radially inner platform 52 and a radially outer platform 53 that in one housing 91 is stored. The rotor 6th includes blades 61 and a turbine disk 62 . The blades 61 are each are about a blade root 63 in the turbine disk 62 attached. In the context of the present description, the blade root 63 but as part of the turbine disk 62 viewed, that is, there is no further terminology between the blade root 63 and turbine disk 62 differentiated. The further rotor 60 also includes blades 601 and a turbine disk 602 .

The blades 601 , 61 and the guide vanes 51 each extend in the main flow path 4th of the gas turbine engine. The section of the main flow path 4th between the blades 601 of the further rotor 60 and the guide vanes 51 of the stator 5 is here with the reference number 41 and the portion of the main flow path 4th between the guide vanes 51 of the stator 5 and the blades 61 of the rotor 6th with the reference number 42 marked.

The assembly also includes an interstage seal 7th that have a gap between the stator 5 and the rotor 6th seals, namely a gap 81 between the interstage seal 7th of the stator 5 and the turbine disk 62 (which in this consideration, as explained, the blade root 63 with includes). The interstage seal 7th includes a section 72 which is used to form a non-contact labyrinth seal 94 with form elements 64 the turbine disk 62 meshes with the labyrinth seal 94 is designed so that secondary air or cooling air through the labyrinth seal 94 through to cool the blade 61 can be provided.

The interstage seal 7th further comprises a radially outwardly directed groove 71 , the one axially forward, radially extending groove wall 711 and an axially rearward, radially extending groove wall 712 having. The groove 71 takes a rail or wall with essentially no play 54 on that part of the radially inner platform 52 and which extends radially inward. In the normal state of the gas turbine engine shown, it is provided that the axially front groove wall 711 axially on the facing side of the rail 54 so that there is no flow connection between an interstage cavity 8th , which will be discussed later, and the section 42 of the main flow path 4th consists.

The already mentioned interstage cavity 8th is between the rotor disk 602 of the further rotor 60 and the interstage seal 7th educated. During normal operation of the gas turbine engine, secondary air is applied to the intermediate stage cavity, which creates a pressure in the intermediate stage cavity 8th which is larger than that in the main flow path 4th prevailing pressure, in particular greater than the pressure in the section 41 of the main flow path 4th is present. This prevents hot gases from getting out of the main flow path 4th into the interstage cavity 8th can flow in.

In the connecting section between the section 41 of the main flow path 4th and the interstage cavity 8th is another non-contact labyrinth seal 93 arranged. In the area of the labyrinth seal 93 there is a temperature sensor 92 that is in the radially inner platform 52 of the stator 5 is stored. The temperature sensor 92 monitors the temperature of the flow through the labyrinth seal 93 . For example, if, due to a malfunction, hot gases from the main flow path 4th into the interstage cavity 8th would flow, the temperature sensor would 92 detect this.

The temperature sensor 92 is about a bracket 920 that also have electrical wiring for the sensor 92 includes, in the stator 5 arranged with the outer end of the bracket 920 in the area of the housing 91 is arranged so that an electrical connection with other components such as an evaluation unit can be made.

The interstage seal 7th further comprises a downstream, radially extending back 73 , which are axially spaced from the turbine disk in the normal operation of the gas turbine engine shown 62 of the rotor 6th is arranged, the turbine disk 62 an outer surface that also extends radially 620 in this area, which is parallel to the back 73 runs. Between the back 73 and the turbine disk 62 runs part of the gap already mentioned 81 between the stator 5 and the rotor 6th .

The backside 73 includes a fin 74 which tapers towards its end and obliquely into the gap 81 protrudes.

In the section of the interstage seal 7th who have favourited the back 73 are a plurality of circumferentially spaced pre-swirl nozzles 76 designed, which serve to set secondary air in rotation and in the direction of the turbine disk 62 deflect and accelerate, so that more effective cooling of the turbine disk 62 and the blade 61 he follows. The secondary air or cooling air is branched off from a compression stage of the compressor of the gas turbine engine.

In the event of failure of one or more of the thrust bearings for supporting the high pressure turbine shaft of the gas turbine engine, which is connected to the rotors 6th , 60 is coupled, the rotors wander 6th , 60 axially forward. This causes the rotating turbine disk 62 with components of the static interstage seal 7th gets in contact. This illustrates the 5 .

According to the 5 are the rotors 6th , 60 axially forward (ie in the illustration of the 4th and 5 to the left) hiked. The labyrinth seal is accordingly 93 essentially repealed. The gap 81 between the back 73 the interstage seal 7th and the outer surface 620 the turbine disk 62 is closed with the back 73 in contact with the outer surface 620 stepped. The transition between the in the 4th normal situation shown and the situation of 5 takes place successively, with the fin first 74 in contact with the outer surface 620 the turbine disk 62 occurs and is gradually removed. However, this is still not critical. However, it is to be regarded as critical that the rear side 73 and the turbine disk 62 come into full contact with each other and accordingly there is high friction between these two parts, which leads to material abrasion on the one hand and to an increase in temperature on the other.

The one in the 5 The situation shown can be detected quickly and reliably on the one hand. On the other hand it is beneficial to reduce the friction between the back 73 and the turbine disk 62 reduce in the event of such a situation.

To provide such functionalities, the 6th an assembly according to the invention, which compared to the assembly of 4th with regard to the design of the radially outwardly directed groove 71 the interstage seal 7th and with regard to the design of the back 73 the interstage seal 7th is modified. The other elements of the assembly are identical, so the relevant description of the 4th Is referred to. the 6th shows additionally schematically represented evaluation means 95 that came with the temperature sensor 92 are connected and the one through the temperature sensor 92 can evaluate recorded temperature change.

The radially outwardly directed groove 71 who have favourited the rail 54 the radially inner platform 52 takes up, points to the assembly of the 6th has an axial width that is greater than the axial width of the rail 54 . Accordingly, in the representation of the 6th the axially rear groove wall 712 axially spaced from the rail 54 arranged. The axially anterior groove wall 711 however, is in axial contact with the rail 54 so that no air from the interstage cavity 8th in the section 42 of the main flow path 4th can flow.

the 7th shows - according to the 5 - the state of the assembly when the rotors 6th , 60 have moved axially forward due to failure of one or more of the thrust bearings. In the 7th it is easy to see that the axial width 713 the groove 71 is greater than the axial width of the radially inwardly extending rail 54 the radially inner platform 52 of the stator 5 . Because the rotor 62 after contacting the back 73 the interstage seal 7th moved axially forward overall, and due to the increased axial width of the groove 713 , the axially front groove wall gets 711 at an axial distance from the rail 54 , leaving a gap 714 in the groove 71 forms through which a gas flow from the interstage cavity 8th in the section 42 the flow path can flow. This can also be made possible by the fact that the groove wall 712 not over the entire circumference of the interstage seal 7th is present, but forms local recesses for a gas passage. In either case, the gap is 714 released when the axially front groove wall 711 off the rail 54 lifts off and up to the axially rear groove wall 712 to the rail 54 applies.

The one with the opening of the gap 714 changing pressure conditions and flows in the considered assembly are related to the 9 and 10 explained.

the 9 corresponds to the 6th , whereby only the static pressures prevailing in the individual areas are also shown. In normal operation of the gas turbine engine according to the 6th and 9 is the greatest pressure P1 in the interstage cavity 8th before, as this is acted upon with secondary air. The pressure P1 is greater than the pressure P2 that is in the 41 of the main flow path 4th prevails so that a gas flow 101 through the labyrinth seal 93 from the interstage cavity 8th in the main flow path 4th he follows.

The pressure P1 in the interstage cavity 8th is also greater than the pressure P4 in the section 42 of the main flow path 4th and the pressure P3 in the gap 81 between the stator 5 and the rotor 6th . At the same time, P4 is smaller than P2. A further gas stream flows accordingly 103 through the pre-swirl nozzles 76 , which extends into the rotor blades via cavities (not shown) 61 can extend. Via a gas stream 102 is cooling air over the labyrinth seal 94 in the gap 81 blown in. Because the pressure P4 in the section 42 of the main flow path 4th is also smaller than the pressure P3 in the gap 81 , there is another gas flow 104 from the gap 81 in the section 42 .

The pressure conditions and flows change when the rotors fail after one or more thrust bearings 6th , 60 migrate axially forward, as in the 10 shown in the 8th and also shows the prevailing pressures and currents.

Through the opened gap 714 the groove 71 a gas stream now flows 106 from the interstage cavity 8th in the section 42 of the main flow path 4th in which the smallest pressure P4 prevails. This gas flow 106 causes a pressure drop in the interstage cavity 8th . This pressure drop in turn causes the pressure P1 in the interstage cavity 8th is now lower than the pressure P2 in the area 41 of the main flow path (upstream of the stator 5 ). Accordingly, hot gas now flows in a gas stream 105 due to the enlarged gap between the rotor 60 and the stator 5 into the interstage cavity 8th . This is done by the temperature sensor 92 at which the gas flow 105 flows past, detected. The temperature increase is caused by the temperature sensor 92 associated evaluation means 95 (see. 6th ) and evaluated with regard to the detection of the failure of a drawer bearing. In the event of a detection, for example, the pilot of an aircraft to which the gas turbine engine is attached can be informed accordingly and / or further suitable measures can be initiated.

Coming back to that 6th and 7th a feature is explained below that in the case of direct contact between the turbine disk 62 and the back wall 73 the resulting friction is reduced. This is done using the 8th explained, which is a sectional view along the line AA of 7th represents, thus reproducing the state when the turbine disk 62 and the back wall 73 stay in contact.

According to the 8th is the back 73 the interstage seal 7th with wedge-shaped circumferential grooves 75 provided, which extend in the circumferential direction and a dynamic pressure in the event of an approach of the turbine disk 62 to the back 73 build, similar to the principle of an air bearing. The air required for this passes through the pre-swirl nozzles 76 at the deepest points of the circumferential grooves 75 in this one. This measure reduces the axial pressure difference across the interstage seal 7th reduced and ideally even vice versa, which leads to or contributes to the axially front groove wall 721 the groove 71 off the rail 54 lifts off so that the axial gap 714 is opened and which in terms of the 9 and 10 explained pressure changes and temperature changes occur.

Alternatively or in addition, it can be provided that the rear 73 and / or the outer surface or contact surface 620 the turbine disk is provided with a friction-reducing coating. Such a coating 97 is due to an increased line thickness on the back 73 the 8th indicated.

It should be noted that sensing a pressure drop in the interstage cavity 8th via one of the temperature sensor 92 measured temperature rise is only to be understood as an example. This variant has the advantage that no additional sensor has to be used because the temperature sensor 92 typically anyway to monitor the flow through the labyrinth seal 93 is available. In principle, however, a pressure drop in the intermediate stage cavity can also be detected directly or indirectly in another way, for example via an additional sensor, which can be a pressure sensor.

the 11 shows schematically and by way of example a method for detecting a failure of a thrust bearing in relation to the 6-10 described assembly.

In a first step 111 becomes the interstage cavity 8th acted upon with secondary air in such a way that, during normal operation of the gas turbine engine, the one in the intermediate stage cavity 8th applied static pressure is greater than that in the main flow path 4th in front of the stator 5 prevailing static pressure. According to step 112 a continuous direct or indirect measurement of the pressure prevailing in the intermediate stage cavity and / or the temperature prevailing there takes place, whereby a continuous measurement does not exclude that measured values are only recorded or evaluated at certain time intervals.

In a further step 113 becomes a pressure drop and / or a temperature rise in the interstage cavity 8th recognized when such a function is present, for example by measuring a temperature rise by the temperature sensor 92 . From a pressure drop and / or a temperature rise according to step 114 a failure of a thrust bearing is recognized when the pressure drop and / or the temperature rise falls below or exceeds a respectively assigned limit value.

It should be understood that the invention is not limited to the embodiments described above and various modifications and improvements can be made without departing from the concepts described herein. It is also pointed out that any of the features described can be used separately or in combination with any other features, provided they are not mutually exclusive. The disclosure extends to and includes all combinations and subcombinations of one or more features described herein. If areas are defined, these include all values within these areas as well as all sub-areas that fall into one area.

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 102015226732 A1 [0003]

Claims (15)

An assembly in a gas turbine engine comprising: - A stator (5) and a rotor (6) of a turbine (17), wherein o the rotor (6) is arranged downstream of the stator (5), o the stator (5) has guide vanes (51) and a radially inner platform (52), o the rotor (6) comprises rotor blades (61) and a turbine disk (62), and o the guide vanes (51) and the rotor blades (61) extend in a main flow path (4) of the gas turbine engine, - an interstage seal (7) which forms an axial gap (81) with the turbine disk (62), - wherein the interstage seal (7) forms a radially outwardly directed groove (71) which receives a radially inwardly extending rail (54) of the radially inner platform (52), the groove (71) having an axially forward, radially extending groove wall (711) and an axially rear, radially extending groove wall (712), and wherein the axially front groove wall (711) is provided and designed to lie tightly against the radially extending rail (54) during normal operation of the gas turbine engine, so that no gas can flow through the groove (71), - An intermediate stage cavity (8) which is formed between the rotor disk (602) of a further rotor (60) arranged upstream of the stator (5) and the intermediate stage seal (7), the intermediate stage cavity (8) being supplied with secondary air during normal operation of the gas turbine engine static pressure is applied which is greater than the static pressure prevailing in the main flow path (4) of the gas turbine engine, - wherein the groove (71) of the interstage seal (7) has an axial width (713) which is greater than the axial width of the radially inwardly extending rail (54), such that when the rotor (6) axially to moved forward and the interstage seal (7) is moved axially forward, the axially front groove wall (711) is axially spaced from the radially extending rail (54) and the groove (71) forms an axial gap (714) which a gas stream (106) can flow from the intermediate stage cavity (8) into the main flow path (4) downstream of the stator (5), and a pressure drop in the intermediate stage cavity (8), triggered thereby, and - Means (92) are provided which measure the pressure drop in the intermediate stage cavity (8) or a temperature rise triggered by this. Assembly according to Claim 1 , characterized in that the means comprise a temperature sensor (92), wherein in the event of a pressure drop in the intermediate stage cavity (8) hot gas flows from the main flow channel (4) into the intermediate stage cavity (8) and the temperature sensor (92) is arranged such that it measures an associated temperature rise. Assembly according to Claim 2 , characterized in that the temperature sensor (92) is designed to measure the temperature of the gas flowing through a non-contact seal (93) in a gap between the radially inner platform (52) and the other, upstream of the stator (5) arranged rotor (60) is arranged and separates the intermediate stage cavity (8) from the main flow path (4). Assembly according to Claim 2 or 3 , characterized in that evaluation means (95) connected to the temperature sensor (92) are provided which evaluate the temperature of the temperature sensor (92) in order to detect a failure of a thrust bearing. Assembly according to one of the Claims 2 until 4th , characterized in that the temperature sensor (92) is arranged in the radially inner platform (52) of the stator (5). Assembly according to one of the preceding claims, characterized in that the intermediate stage seal (7) has a downstream, radially extending rear side (73) which is arranged in such a way that, during normal operation of the gas turbine engine, it is axially spaced from the turbine disk (62) of the rotor ( 6), the rear side (73) of the interstage seal (7) being provided with wedge-shaped circumferential grooves (75). Assembly according to Claim 6 , characterized in that wedge-shaped circumferential grooves (75) are provided and designed to build up a dynamic pressure when the turbine disk (62) approaches the rear side (73), which creates friction between the turbine disk (62) and the rear side (73). the intermediate stage seal (7) is reduced. Assembly according to Claim 6 or 7th , characterized in that the interstage seal (7) also forms pre-swirl nozzles (76) in the rear side (73) which are arranged such that air enters the circumferential grooves (75) through the pre-swirl nozzles (76). Assembly according to Claim 8 , characterized in that the pre-swirl nozzles (76) are arranged in such a way that air enters the circumferential grooves (75) through the pre-swirl nozzles (76) at the lowest points of the circumferential grooves (75). Assembly according to one of the Claims 6 until 9 , characterized in that the rear side (73) of the interstage seal (7) and / or the turbine disk (62) are provided with a friction-reducing coating (97) in the potential contact area (620). Gas turbine engine (10) with an assembly according to Claim 1 wherein the gas turbine engine (10) comprises a high pressure turbine (17) comprising a first stage and a second stage, each stage having a stator and a rotor, and wherein the interstage seal (7) of the assembly between the stator (5) and the rotor (6) of the second stage of the high pressure turbine (17) is arranged. Method for detecting a failure of a thrust bearing supporting a turbine shaft of a gas turbine engine using an assembly according to FIG Claim 1 which comprises the steps: - loading (111) the intermediate stage cavity (8) with secondary air in such a way that, during normal operation of the gas turbine engine, the static pressure present in the intermediate stage cavity (8) is greater than that in the main flow path (4) of the gas turbine engine upstream of the stator ( 5) prevailing static pressure, - continuous measurement (112) of the pressure and / or the temperature prevailing in the intermediate stage cavity (8), - by measuring detection (113) of a pressure drop and / or temperature rise in the intermediate stage cavity (8) if such is present, detection (114) of the failure of a thrust bearing when the pressure drop and / or the temperature rise falls below or exceeds a respectively assigned limit value. Procedure according to Claim 12 , characterized in that in the event of a pressure drop in the intermediate stage cavity (8), hot gas flows from the main flow channel (4) into the intermediate stage cavity (8) and a temperature sensor (92) measures a temperature rise triggered thereby. Procedure according to Claim 13 , characterized in that the temperature sensor (92) measures the temperature of the gas flowing through a seal (93) which is arranged in a gap between the radially inner platform (52) and the further rotor (5) arranged upstream of the stator (5). 60) of the assembly is arranged. Method according to one of the Claims 12 until 14th , characterized in that when the turbine disk (62) approaches the rear side (73) of the interstage seal (7), friction between the turbine disk (62) and the rear side (73) is reduced by secondary air in the rear side (73) the interstage seal (7) formed wedge-shaped circumferential grooves (75) is blown.

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