SENSOR ARRANGEMENT AND MEASUREMENT METHOD FOR A TURBOMACHINE

A sensor arrangement with a sensor element for measuring at least one physical and/or chemical fluid characteristic in a turbomachine is provided. The sensor element detects the at least one fluid characteristic inside a non-contact seal, in particular a labyrinth seal, between a rotor stage and a stator stage, wherein during operation the sensor element is in contact with the fluid flow along the flow path inside the labyrinth seal.

REFERENCE TO RELATED APPLICATION

This application claims priority to German Patent Application No. 10 2015 226 732.6 filed on Dec. 24, 2015, the entirety of which is incorporated by reference herein.

BACKGROUND

The invention relates to a sensor device for a turbomachine and a measurement method.

In turbomachines, in particular in aircraft engines, fluid flows play a role in many areas. Thus, it is for example necessary to use cooling air and to monitor its temperature in different areas of the turbomachine. A sudden temperature rise of the cooling air can be an indication of a problem, or it may itself cause a problem. It may also be necessary to monitor working fluids, such as for example oil flows.

SUMMARY

Therefore, there is the objective to create sensor arrangements and measurement methods which facilitate an efficient and reliable monitoring of fluids inside the turbomachine.

At that, the sensor element detects at least one fluid characteristic inside a non-contact seal between a rotor stage and a stator stage, wherein during operation the sensor element is in (direct or indirect) contact with the fluid flow along the flow path inside the non-contact seal. Thus, the detection takes place directly inside the seal and not in an upstream or downstream position.

In one embodiment, a pressure gradient is present across the flow path, so that a fluid is transported in the non-contact seal, in particular the labyrinth seal.

In a further embodiment, the rotor stage and the stator stage are arranged inside a turbine, in particular a high-pressure turbine, wherein the seal with the sensor element is located near the gas path, in order to provide a seal against the entry of hot gases and if necessary to detect the entry of a hot gas.

Here, the sensor element can be arranged at the beginning, in the middle or at the end of the flow path through the non-contact seal, in particular the labyrinth seal. In addition or as an alternative, the sensor element can also be arranged inside a measuring chamber, which is connected through a channel with the actual measuring point in the seal, and thus the fluid to be detected first has to be guided to the sensor element through a channel.

When the sensor arrangement is applied, the sensor element can measure a temperature of the fluid flow, in particular of an air flow, a mass flow of the fluid flow, in particular an oil flow, and/or a composition of the fluid flow, in particular of a combustion gas. For this purpose, for example optical sensors can be used that can detect the composition, the conductivity or the optical permeability of the fluid flow.

Here, it is also possible that the sensor element is coupled to a control device52by means of which a control signal can be emitted based on the measurement of the sensor element51, which may for example inform the pilot in the cockpit about the measurement, or may initiate an automatic switch-off of the aircraft engine.

The objective is also achieved through a measurement method with features as described herein.

DETAILED DESCRIPTION

The aircraft engine10according toFIG. 1shows a general example of a turbomachine.

In most cases, the aircraft engine10is embodied in a per se known manner as a multi-shaft engine and comprises, arranged in succession in flow direction, an air inlet11, a fan12rotating inside a housing, where applicable a medium-pressure compressor13, a high-pressure compressor14, a combustion chamber15, a high-pressure turbine16, where applicable a medium-pressure turbine17and a low-pressure turbine18, as well as an exhaust nozzle19, which all are arranged around a central engine axis1.

The medium-pressure compressor13and the high-pressure compressor14respectively comprise multiple stages, of which each has an arrangement of fixedly attached stationary guide vanes20extending in the circumferential direction, which are generally referred to as stator vanes and which protrude radially inwards from the engine shroud21through the compressors13,14into a ring-shaped flow channel. Further, the compressors have an arrangement of compressor rotor blades22, which protrude radially outwards from a rotatable drum or disc26and which are coupled to turbine rotor hubs27of the high-pressure turbine16or of the medium-pressure turbine17.

The turbines16,17,18have similar stages, comprising an arrangement of stationary guide vanes23, which protrude radially inwards from the housing21through the turbines16,17,18into the ring-shaped flow channel, and a subsequent arrangement of turbine blades24which protrude outwards from the rotatable turbine rotor hub27. During operation, the compressor drum or compressor disc26and the blades22arranged thereon as well as the turbine rotor hub27and the turbine blades24arranged thereon rotate around the engine central axis1.

Often there is a gap between the stators and the rotors of the compressors13,14and the turbines16,17,18directly at the flow channel or also further inside the engine, through which fluids, such as for example cooling air or oil, may leak. These gaps are usually provided with non-contact seals40, such as for example a labyrinth seal. Here, it is important for the functionality of the aircraft engine10that the fluid flow through the seal40is monitored.

FIGS. 2 and 3show embodiments of a labyrinth seal40with a sensor arrangement50for fluid monitoring. In a labyrinth seal40, the sealing effect is based on an elongation of the flow path inside the gap to be sealed, whereby the flow resistance is increased. Typically, a labyrinth seal40has meshing parts (combing).

The sectional planes ofFIGS. 2 and 3comprise the engine axis1and thus extend in the radial direction.

FIGS. 2 and 3respectively show an exemplary embodiment in which the flow direction extends from left to right in the aircraft engine10. Thus, the turbine flow first flows through a stator stage30with guide vane leafs before impinging on a rotor stage29with rotor blade leafs.

InFIG. 2, the labyrinth seal40is arranged between the stator disc30and the rotor disc29. The fluid flow41—in this case air, for example—flows radially form inside out along the flow path S. At that, the flow path is delimited radially inwards and outwards by the dashed line.

Here, a sensor element51is arranged directly at the flow path S as a part of a sensor device50, by means of which a fluid characteristic, namely the temperature inside the labyrinth seal40, is measured. In the process, the sensor element51is in (direct or indirect) contact with the fluid flow41along the flow path S inside the labyrinth seal40. In the present case, the sensor element41is arranged approximately in the middle of the flow path S, and is oriented in such a manner that the fluid flow41flows towards the sensor element41.

InFIG. 3, an alternative sensor arrangement50inside a labyrinth seal40is shown, wherein the description ofFIG. 2may be referred to. Here, too, the sensor element51is arranged approximately in the middle of the flow path S. But in this case, the flow impinges on the sensor element51tangentially.

In alternative embodiments, the sensor element41is arranged at the beginning or at the end of the flow path S. It is also possible for the sensor element51to be arranged not at a right angle to the flow path S, but at an acute or obtuse angle. In any case, it has direct contact with the fluid flowing inside the labyrinth seal40.

In the embodiments ofFIGS. 2 and 3, the sensor element51of the sensor arrangement50is connected to a control device52that can emit a control signal53if for example a certain measurement value—here a temperature value—is exceeded or undershot.

If, for example due to a malfunction, hot gas would flow from the outside into the labyrinth seal40, the temperature rise would be detected. This could then be translated into a control signal53by the control device52.

InFIGS. 4 to 6, some such cases are shown in a schematic manner. In all three cases, the sensor element51is arranged approximately in the middle of the fluid flow41. The sensor arrangement is respectively arranged in the stator stage30.

FIG. 4shows a labyrinth seal40that is arranged between a rotor stage29to the left and a stator stage30to the right. What is shown here is a case in which hot gas flows from the outside radially inward through the labyrinth seal40in the event of a damage to the aircraft engine10. The sensor element51detects the temperature rise that is an indicator for the damage and passes the information on to the control device52. If a defined threshold value is exceeded, the pilot in the cockpit could be informed by a signal. Another possibility would be an automatic switch-off of the aircraft engine.

FIG. 5shows a further application of the sensor arrangement50which is designed for the detection of oil leaks. Here, a non-contact seal40is arranged between a stator stage30above and a rotor stage29that is located below the same. A sensor element51, for example of an optical oil sensor, for example detects a leak of a bearing chamber in the event that the oil flow exceeds a certain amount. The measurement result is forwarded to the control device52.

InFIG. 6, a labyrinth seal40is arranged between a stator stage30to the left and a rotor stage29to the right. Here, an optical sensor detects whether combustion products and/or solid bodies flow through the labyrinth seal40together with the fluid. In this case, the optical characteristics (for example the extinction) of the fluid would change, so that a case of error due to the entry of combustion gases via the seal element could be detected. Also in this case, the measurement result is passed on to the control device52.

FIG. 7shows a perspective view of a non-contact seal40between a stator stage and a rotor stage of a turbine.FIG. 7shows the fluid flow41that occurs due to an unplanned suctioning in of a hot gas for example. In this case, the sensor element51does not operate directly at the site of the fluid flow41. However, in order to also be able to immediately detect the hot gas before it advances further into the aircraft engine10, the fluid flow41is suctioned through a pressure gradient that is present above the stator stage, and is guided past the sensor element via the air channel54. By designing a short channel, a delayed response to the entry of the hot gas is minimized, but also a simplified design, here in particular of a straight sensor element51, is facilitated. This simple constructional design of the sensor facilitates an exchange of the element with engines that are still mounted at the aircraft.

In the embodiment shown herein, the sensor arrangement50has an approximately cylindrical sensor element51.

Parts list