GB2641387A

GB2641387A - Ventilation mechanism for a gas turbine

Info

Publication number: GB2641387A
Application number: GB2407643.2A
Authority: GB
Inventors: Hübner Gerhard
Original assignee: Rolls Royce Deutschland Ltd and Co KG
Current assignee: Rolls Royce Deutschland Ltd and Co KG
Priority date: 2024-05-30
Filing date: 2024-05-30
Publication date: 2025-12-03
Also published as: GB202407643D0

Abstract

A gas turbine engine, aircraft engine and aircraft are disclosed. In each, a gas turbine comprising a bypass duct 20 has at least one ventilation device 27 that is configured to release core airflow FC out of a core airflow passage FCP. The bleeding device 27 comprises at least one pyrotechnical opening mechanism 100. By explosively opening the core airflow an uncontrolled increase of the rotational speed of the shaft of the gas turbine can be prevented. A method for protecting against turbine overspeed event or shaft failure in a gas turbine is also disclosed, comprising the steps of detecting the event, cutting off fuel and actuating a ventilation device 27 to release core airflow FC.

Description

VENTILATION MECHANISM FOR A GAS TURBINE

FIELD

The present disclosure relates to a ventilation mechanism for gas turbines in general, and more specifically, to gas turbines of engines and/or engines of aircrafts.

BACKGROUND

In subsonic as well as in supersonic flight conditions, the need for additional protection systems against turbine overspeed and/or shaft failure events which may lead to hazardous consequences for aircrafts is becoming increasingly important in civil aviation. Regulations dictate for gas turbines powering civil transportation aircrafts to have a protection system against turbine overspeed and/or a shaft failure event. Such protection systems may be configured to automatically shut off the fuel supply to the engine combustor in case of a detection of an over speeding of the turbine shaft to prevent the turbine against bursting which may lead to the release of hazardous high energy debris impacting the aircraft.

US 2017/284306 Al discloses an aircraft turbine engine, comprising at least one first compressor, an annular combustion chamber and at least one first turbine, which define a first flow duct for a primary flow, which comprises, between said combustion chamber and said first turbine, a device for discharging at least part of said primary flow.

FR 3114608A1 discloses a method for controlling a turbomachine. The method comprises the steps of monitoring whether at least one event occurs indicating that a speed of the member has gone a predetermined threshold; and if the event occurs, dispensing at least a portion of the flow so that this portion of the flow does not supply power to the combustion chamber.

There is a need to further improve gas turbine engines with respect to safety, whilst maintaining efficiency.

It is therefore desirable to address at least one of the above problems.

One objective underlying the present disclosure is to improve gas turbine engines with respect to safety characteristics.

SUMMARY

According to a first aspect of the disclosure, a gas turbine is proposed comprising a bypass duct and a core engine together defining a bypass airflow passage configured to guide a bypass airflow; the core engine defining a core airflow passage, configured to guide a core airflow. The gas turbine has at least one ventilation device, configured to release the core airflow out of the core airflow passage. The ventilation device comprises at least one pyrotechnical opening mechanism.

The present disclosure is based on the idea to explosively ventilate or release the hot gas of the core airflow passage of the core engine by at least one ventilation device. The ventilation device comprises at least one pyrotechnical opening mechanism that is configured to explosively open the core airflow. By explosively opening the core airflow, or the core airflow passage, respectively, an uncontrolled increase of the rotational speed of the shaft of the gas turbine engine can be prevented. The opening may be configured to open explosively upon ignition, creating an opening of a specific area, such as an area at a casing, at a membrane and/or at a duct. The opening may be a one-off use of the ventilation device at least as part of the existing turbine overspeed protection systems of the engine. The disclosure can provide an additional safety mechanism in case of a turbine overspeed and/or a shaft failure, in particular in addition to one or more existing turbine overspeed protection systems of the engine, which may include a fuel cut off. This safety mechanism, i.e. the ventilation device, is configured to ventilate or release the hot gas core airflow in the core airflow passage and can support the restraining of energy from the core of the engine to protect the turbine from hazardous over speeding. Furthermore, it is generally beneficial to have backup options in case of any delay or failure of one of the existing turbine overspeed protection systems. In particular, the actuation of the at least one ventilation device may be linked to the actuation of the existing detection system of the gas turbine engine. In case of an emergency shut down of the engine, the actuation of the ventilation device may be part of a command sequence.

A further advantage particularly applies to supersonic flight conditions where environmental conditions are affecting the inside of the engine to a larger extent. In particular, the supersonic ram air pressure effects on the inside of the core airflow passage of the engine, which results in an additional input of energy to the engine. Therefore, an additional protection mechanism would improve the overall safety. The ventilation or release of the core airflow will limit the supersonic ram air pressure effect across the core, especially on the turbine which could overspeed and burst due to the high stresses caused by the rotational speed.

Furthermore, it is generally beneficial to cover a wide range of operating conditions under certification levels of safety and the reduction of critical conditions as turbine overspeed and/or a shaft failure event.

Further developments of the invention can be found in the dependent claims and show particularly advantageous possibilities to implement the above-described concept in light of the object of the invention and regarding further advantages.

The pyrotechnical opening mechanism may comprise a carrier, in particular ring carrier. The pyrotechnical opening mechanism may further comprise an arm device, wherein the arm device comprises at least one electro-pyrotechnical detonator and at least one pyrotechnical charge.

In a development, the disclosure further proposes a pyrotechnical opening mechanism that may comprise a carrier, in particular a ring carrier, that may be coupled and/or attached to an area or location preferred. The carrier may have a circular or non-circular shape or dimension, as the purpose may be to carry the arm device and to place and fix the arm device at a preferred area. The arm device may comprise at least one low-or high-energy electro-pyrotechnical detonator and at least one pyrotechnical charge, wherein detonators are used for various purposes, both military and civilian ones, but will here be described mainly in relation to applications configured for detonation or blasting in gas turbine engines. Electro-pyrotechnical detonators may be configured to be activated by high voltages for increased safety.

The pyrotechnical opening mechanism may comprise a pyrotechnically actuated device, such as exploding bolts, pin pushers/ retractors, pyrotechnical valves, separation devices, shaped charges and cable/ rope cutters, or similar pyrotechnical disconnection / release systems.

The arm device may be a ring. The pyrotechnical opening mechanism may comprise an igniter. The arm device may comprise an interface to the igniter, wherein the igniter is connectable or connected to an Electronic Engine Control.

The arm device may have another shape and dimension as the main purpose is to maintain, carry and protect the pyrotechnical charge from uncontrolled and unwanted detonation. The interface to the igniter may support a controlled and commanded detonation of the pyrotechnical charge of the arm device. In particular, the automatic control of the detonation through the Electronic Engine Control (EEC) is preferred. The EEC may be configured to control the igniter.

The pyrotechnical opening mechanism may be configured to be activated via a shaft failure signal from the EEC to the igniter.

The Electronic Engine Controller EEC may be configured to detect a turbine overspeed and/or a shaft failure and to identify it as such. The automatic order of relevant safety steps may be adapted according to the requirements and/or the actual configuration of the turbine overspeed protection system. In particular, the EEC may be configured to automatically send a shaft failure signal to the igniter, which is interfacing the pyrotechnical opening mechanism. The automatic detection and command via signals may be part and another relevant safety aspect of the existing turbine overspeed protection system.

The igniter may be configured to generate a detonation signal upon receiving the shaft failure signal, wherein the detonation signal is configured to generate a spark and/or a short circuit or the like signal suitable to initiate a detonation.

In general, such pyrotechnical devices are initiated by a remotely controlled electrical signal that causes an electric igniter to trigger detonation. The remote control may be manual, or preferably an automatic control via the Electronic Engine Controller (EEC). In particular, the igniter initiates detonation, which includes sending a detonation signal or a detonation pulse which may be converted into a spark and/or short circuit to cause a detonation of the pyrotechnical charge. Another advantage of an automatic initiation of the detonation is the quick reaction time from detection to detonation, and the controlled and quick response of the pyrotechnical charge.

The spark and/or the short circuit may be configured to cause an explosive response of the pyrotechnical charge.

Sparks and/or short circuits represent a cost-effective and highly reliable way to cause an explosive response of the pyrotechnical charge. Furthermore, only a limited amount of energy is required to initiate the explosive response of the pyrotechnical charge. Although the resultant pressure wave does not result in high forces being introduced into the ventilation device, the pyrotechnical opening mechanisms are mounted such that they cannot move in the remaining components of the ventilation device. The detonation induced by means of one and/or multiple sparks is applicable to the automatically command of the EEC.

The ventilation device may be arranged on the inner casing of the turbine section, in particular upstream of the low-pressure turbine.

By arranging the ventilation device on or at the inner casing of the turbine section, in particular upstream of the low-pressure turbine, for example adjacent to and/or coupled to the inner casing, the core airflow passage can be explosively opened when the ventilation device is initiated, and the core airflow may be ventilated or released out of the core airflow passage. This is beneficial, as it utilizes existing construction and it is easy to assemble. In supersonic conditions this would have a significant advantage for the ram air ventilation or release to the turbine in case of a turbine overspeed and/or a shaft failure. The electronic controller may be configured to control a command sequence between the components of the turbine overspeed protection system of the turbine. The ventilation device can be considered as an additional component to the turbine overspeed protection system.

The ventilation device may be arranged between the inner casing of the turbine section and bypass duct at an axial position upstream of the low-pressure turbine.

By arranging the ventilation device between the inner casing of the turbine section and the bypass duct in particular at an axial position upstream of the low-pressure turbine, for example adjacent to and/or coupled to the inner casing, the core airflow can be ventilated or released out of the core airflow passage. This is beneficial, as existing construction can be used and it is easy to assemble. In supersonic conditions, this would have a significant advantage for the ram air ventilation or release to the turbine in case of a turbine overspeed and/or a shaft failure. The electronic controller may be configured to control a command sequence between the components of the turbine overspeed protection system of the gas turbine. The ventilation device can be considered as an additional component of the turbine overspeed protection system.

The ventilation device may comprise an ejector. The ejector may be configured to ventilate into the bypass airflow passage.

An ejector may be configured to ventilate or guide, and to selectively drive the core airflow through the inner casing of the turbine to the bypass airflow passage. The ejector is a device that allows a portion of the core airflow to be conducted into the bypass airflow passage. The ejector may comprise a sectioned ejector body. The sectioned ejector body forms an ejector airflow passage configured to catch the hot and highly pressurised core airflow and to ventilate it into the bypass airflow passage. The use of multiple ejectors, e.g. a plurality of 2, 3, 4, 5, 6 ejectors, in particular arranged equally around the circumference of the gas turbine, may support the quick extraction of the hot core airflow.

The ejector may be a Venturi-type valve.

According to this development, an ejector may be profiled as a Venturi-type valve configuration. Venturi-type valves are most often used in spaces that require precise and accurate airflow control. Venturi-type valves may be configured such that the core airflow can be actively drawn into the bypass airflow passage upon activation of the ventilation device. A Venturi-type arrangement of the ejector body results in precise airflow control over a wide range of pressures without the need for actuated airflow control. The venturi effect makes the ejector immune to dust and lint contamination.

The ejector may be at least partially coated with a heat-resistant coating.

The ejector may -specifically in its function as a component of a ventilation device benefit from a heat-resistant coating. A heat-resistant coating may let the ejector and/or the sectioned ejector body withstand the hot core airflow as long as required and provide a shielding to the bypass duct against hot core airflow. This may significantly reduce the burning intensity and fire spread, also when coating is applied in fire safety, particularly for individual elements of structures such as castings, ducts, columns, and frames. To increase the fire resistance of surfaces, the heat-resistant coating may be used as a passive fire protection method.

The ventilation device may be integrated in an outlet guide vane strut. The ventilation device may be configured to release the core airflow through the at least one outlet guide vane strut, in particular along a longitudinal axis of the at least one outlet guide vane strut, to an outer environment of a nacelle of the gas turbine, i.e. outside of the gas turbine. Integrating the ventilation device into at least one, preferably multiple, outlet guide vane struts is beneficial in terms of cost saving, easy assembly and using existing gas turbine features. In this constellation it is preferred to install at least one pyrotechnical opening mechanism. Further, the ventilation device may comprise at least one discardable plug to support the effect of the ventilation to release the hot core airflow through the bypass duct casing to outside of the nacelle, in particular at a longitudinal end of the outlet guide vane strut, in particular opposite of the pyrotechnical opening mechanism. The ventilation device may comprise one pyrotechnical opening mechanism, in particular arranged at one end of the outlet guide vane strut in or at the inner casing of the gas turbine, and one discardable plug, in particular arranged at the other end of the outlet guide vane strut in or at the nacelle. The discardable plug may be configured to be released upon contact with the core airflow, which in particular is released into the outlet guide vane strut by the pyrotechnical opening mechanism. The discardable plug may comprise a material that is configured to react upon a change, in particular increase, in pressure and/or temperature and/or chemical composition of the contacted fluid. The discardable plug may be a releasable cover and/or comprise a membrane, or a plug, or the like predetermined breaking zone. The discardable plug ensures protection and environmental sealing of the outlet guide vane strut during normal operation of the gas turbine engine, as well as a smooth nacelle surface.

The inner shape of the outlet guide vane strut may act like a bleed valve. To withstand the hot core airflow, the outlet guide vane strut structure may receive a heat-resistant coating to withstand the passage of hot gas. An outlet guide vane strut, configured to guide the core airflow through the bypass duct casing and nacelle to the outside of the nacelle of the gas turbine, may prevent the bypass duct and nacelle being damaged, as core airflow may have a high temperature and/or pressure, which can be guided through to the outside, in particular without a direct contact to the bypass duct and nacelle. The angular positioning of the outlet guide vane strut with respect to the circumference is preferably chosen outside of an angular keep out zone, such that the aircraft is not impacted by the ejected core gas airflow. Preferably the outlet guide vane strut position may be arranged outside of the angular keep out zone corresponding to a ventilation or releasing of the core airflow with a sufficient distance from the aircraft. By arranging the outlet guide vane strut outside of this angular keep out zone, the core airflow may be directed away from the aircraft in order to avoid critical damage which may lead to a hazardous condition. The angular keep out zone may differ depending on the respective aircraft configuration and gas turbine design, e.g. embedded engine arrangement, wing-mounted engines, fuselage mounted engines or placement above the wing.

According to a second aspect, an aircraft engine is proposed, comprising: a gas turbine according to the first aspect, a turbine overspeed protection system, which is configured to control the ventilation device, wherein the at least one ventilation device is part of the turbine overspeed protection system.

In use, a gas turbine engine described and/or claimed herein may operate at different flight conditions defined elsewhere herein. Such flight conditions may be determined by the flight conditions (for example subsonic or supersonic conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

According to a third aspect, an aircraft is proposed, comprising the gas turbine according to the first aspect of the disclosure, and/or an aircraft engine according to the second aspect of the disclosure.

The aircraft may be, among other types, a civil or military aircraft, or a space aircraft such as a space shuttle. In particular, the aircraft may be a supersonic aircraft, as a ventilation device according to the present disclosure is particularly suitable to alleviate the effects of a shaft failure and/or turbine overspeed event at high flight velocities.

According to a fourth aspect, a method for protecting against a turbine overspeed event and/or a shaft failure event in a gas turbine which comprises an overspeed protection system, is proposed, the method comprising: detecting a turbine overspeed event and/or a shaft failure event, cutting off fuel, and actuating a ventilation device to release a core airflow out of a core airflow passage of the gas turbine.

The proposed method includes the finding, that an overspeed protection system may be used further to actuate the ventilation device to switch from a closed to an open position to ventilate or release energy from the core of the engine. The opening actions may be commanded by the electronic controller, for example by providing a failure signal and/or a detonation signal.

The method may further comprise: moving variable stator vanes into a closed position.

Such moving of variable stator vanes can restrict airflow through a compressor. The method may further comprise: setting handling bleed valves to open position blocking airflow through a compressor to a bypass airflow passage of the gas turbine engine.

By moving the variable stator vanes into a closed position, further energy from the core of the engine may be restrained. In particular, it is further pointed out that the bleed valves may have two actuating positions only, namely, the opened position and the closed position, or may alternatively be continuously controllable between the opened position and the closed position.

The open position includes that the air in the gas path in the core airflow can flow to a secondary channel such as a bypass airflow passage.

According to a fifth aspect, an electronic controller, configured to perform the method according to a fourth aspect.

It is generally beneficial to use an electronic controller for commanding (such as the EEC). In particular, the electronic controller may be configured to set the variable stator vanes, the bleed valves and/or at the ventilation device to a predetermined position. When controlled by the EEC, the previously mentioned components may be continuously controllable between a fully opened position and a fully closed position.

It shall be understood that the gas turbine according to the first aspect of the disclosure, the aircraft engine according to the second aspect of the disclosure, the aircraft according to the third aspect of the disclosure, the method according to the fourth aspect of the disclosure, and the electronic controller according to the fifth aspect of the disclosure comprise identical or similar developments, in particular as described in the dependent claims. Therefore, a development of one aspect of the disclosure is also applicable to another aspect of the

disclosure.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure will be explained in more detail with exemplary embodiments described with reference to the figures, wherein: FIG. 1A shows a schematic cross section in which a first embodiment of a gas turbine engine where a ventilation device is illustrated; FIG. 1B shows an enlarged plan view and a sectional view of the ventilation device wherein the pyrotechnical opening mechanism is closed and integrated into an ejector; FIG. 1C shows in addition to FIG. 1B the ventilation device wherein the pyrotechnical opening mechanism is open; FIG. 2A shows in addition to FIG. 1A another embodiment of the ventilation device; FIG. 2B shows in addition to FIG. 1B the pyrotechnical opening mechanism and discardable plug closed and integrated into an outer guide vane strut; FIG. 2C shows in additional to FIG. 2B the pyrotechnical opening mechanism and discardable plug open; FIG. 3 shows an enlarged plan view and a sectional view along the line A-A of the pyrotechnical opening mechanism comprising an interface to an igniter wherein the igniter is connected to an EEC; FIG. 4 shows a flowchart of a method for activating the ventilation device in case of turbine overspeed and/or a shaft failure; and FIG. 5 shows a block diagram of the command chain in case of a turbine overspeed and/or a shaft failure event.

DETAILED DESCRIPTION

Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

FIG. 1A schematically illustrates a sectional view of a gas turbine G of an aircraft engine E. Components of the aircraft engine E are arranged along a longitudinal engine axis A. A radial axis R of the turbofan engine extends perpendicularly to the engine axis A. Air is drawn in at the inlet by means of a fan (not shown here). The fan is driven by a turbine 23 via a low-pressure shaft 22. The turbine 23 is fluidically connected to a compressor, which has an intermediate-pressure compressor 15 and a high-pressure compressor 16 and may optionally also comprise further compressor stages such as a medium pressure compressor (not shown here). To generate thrust, the fan supplies the intermediate-pressure compressor 15 and the high-pressure compressor 16 as well as the bypass airflow passage FBP with air. This creates a core airflow FC that runs through the core airflow passage FCP of the core engine EC and a bypass airflow FB that runs through the bypass airflow passage FBP. The air compressed in the compressor 15 is mixed with fuel in the combustor 18 and burned. The hot gas produced is used to drive the turbine 23, which can comprise a high-pressure turbine 19 and a low-pressure turbine 21. Optionally, the turbine 23 may comprise further turbine stages such as a medium-pressure turbine (not shown here). The energy released during combustion is used by the turbine 23 to drive a low-pressure shaft 22 and thus to drive the fan, in order to then generate the required thrust via the air conveyed into the bypass airflow passage FBP. Both the bypass airflow FB from the bypass airflow passage FBP and the core airflow FC from the core airflow passage FCP flow out at an annular exit cross section 26 into an outer environment EN of the gas turbine G. The annular exit cross section 26 usually has a thrust nozzle with a centrally arranged exhaust cone 25. To reduce noise, a mixer may be located in the area of the outlet as part of an exhaust section 24 (not shown here). The special contour of the mixer deflects and mixes the main flow from the core airflow FC and the bypass airflow FB from the bypass airflow passage FBP of the gas turbine G in such a way that the resulting turbulence reduces the audible noise level.

At the rear inner casing 60 of the core engine EC, at least one, preferably two (as shown here), ventilation devices 27 are arranged. The ventilation devices 27 are coupled to the inner casing 60 of the turbine section 23 and arranged upstream of the low-pressure turbine 21. The ventilation devices 27 each comprise at least one pyrotechnical opening mechanism 100, shown in closed position CP.

The proposed solution can also be applied to any other turbine and/or turbine stage in any type of gas turbine engine of the kind shown in FIG. 1A or of any other kind with a different design, for example such as an open-rotor or turboprop engine or a geared turbofan.

In the context of the present disclosure, one ventilation device 27 may be already beneficial to protect the low-pressure turbine 21 in case the low-pressure shaft 22 fails and/or the turbine 23 overspeeds. The ventilation device 27 may be part of the existing protection system of the gas turbine G and may also be configured to be commanded by an electronic controller 220. A command sequence in case of a turbine overspeed and/or a shaft failure 50 is shown in more detail, as a block diagram, in FIG. 4. In FIG. 1A, the gas turbine engine E comprises two ventilation devices 27. In case of a turbine overspeed and/or a shaft failure 50, the electronic controller 220 may be configured to command a sequenced opening of the ventilation devices 27, which means that the pyrotechnical opening mechanisms 100 are set explosively to open, to ventilate, i.e. release, the core airflow FC of the core engine EC. In particular the engine electronic controller EEC as part of the turbine overspeed protection system 230 may send a signal to open the ventilation devices 27. Open includes in particular the igniting, and explosively opening or bursting, of the pyrotechnical opening mechanisms 100. Preferably the ventilation devices 27 may be configured to open at the same time. In FIG. 1B and FIG. 2B the ventilation devices 27 are shown in more detail and at least as two different embodiments. In particular at supersonic conditions, the ventilation of the core airflow FC may have a significant effect on the avoidance of a hazardous condition for the gas turbine engine E, subsequently for the aircraft.

Furthermore, the existing turbine overspeed protection system 230 of the gas turbine engine E may comprise a fuel supply shut-off mechanism to the engine, a variable stator vane mechanism and handling bleed valves. The compressor, in which the variable stator vane mechanism and the handling bleed valves are implemented, may be an intermediate-pressure compressor 15, a medium-pressure compressor or a high-pressure compressor 16 of the aircraft engine E. The ventilation device 27 provides an additional safety component in case of a turbine overspeed and/or a shaft failure 50 to the existing turbine overspeed protection system 230 of the gas turbine engine. In operating conditions such as supersonic flight conditions this may have a significant benefit and effect on releasing the energy of the core airflow FC of the core engine EC. Also, the ventilation device 27 may be implemented in several, or all, of the compressors of a gas turbine G. FIG. 1B is an exemplary cross-sectional view of a core engine EC with a ventilation device 27 as shown in 1A. Illustrated in an enlarged and detailed view is the ventilation device 27 comprising an ejector 40. The illustrated ejector 40 is comprising a frame 42, a sectioned ejector body 46, an ejector airflow passage FEP and the pyrotechnical opening mechanism 100 in closed position CP, wherein the pyrotechnical opening mechanism 100 is preferably integrated into the ejector frame 42 of the lower ejector body 46.2 and locally arranged and coupled to the inner casing 60 of the core engine EC. In this embodiment the ejector airflow passage FEP is located between the sectioned ejector body 46 so that the bypass airflow FB from the bypass airflow passage FBP may pass through and towards the exhaust section 24. The upper ejector body 46.1 as well as the lower ejector body 46.2 are at least partially coated with a heat-resistant coating 29.1 in order to withstand the hot core airflow FC and protect the bypass duct 20 and the nacelle 30 from damage. The ejector 40 may be of annular configuration, meaning that the upper ejector body 46.1 and the lower ejector body 46.2 are joined together such as to form a duct, that is arranged within the bypass airflow passage FBP. The ejector 40 may be substantially rotational symmetric with respect to an ejector axis AE, or have another suitable, e.g. aerodynamically optimized, shape.

The lower ejector body 46.2 comprises in this embodiment a Venturi-type valve 48, configured to release the hot core airflow FC into the bypass airflow FB in case of a turbine overspeed and/or a shaft failure 50.

FIG. 1C is an exemplary illustration of another embodiment of FIG. 1B, wherein the difference is that the pyrotechnical opening mechanism 100 is in open position OP wherein the ventilation device 27 with the Venturi-type valve 48 ventilates, or releases, through the hole in the inner casing 60 of the core engine EC the hot core airflow FC into the bypass airflow FB, via a venturi-effect.

FIG. 2A is an exemplary illustration of another embodiment, wherein the difference to FIG. 1A is that the ventilation devices 27 are arranged between the inner casing 60 of the turbine section 23 and a nacelle 30. The discardable plugs 32 are arranged on an outer surface of the nacelle 30. The ventilation devices 27 are arranged at an axial position upstream of the low-pressure turbine 21. The illustrated ventilation devices 27 are arranged at the turbine section 23 and coupled to the inner casing 60 of the core engine CE and the nacelle 30 casing. Each ventilation device 27 comprises in this embodiment one pyrotechnical opening mechanism 100 and one discardable plug 32 shown both in closed position CP. In other embodiments, the ventilation device 27 may comprise two pyrotechnical opening mechanisms 100, one on or in the inner casing 60, and one on or in the nacelle 30.

FIG. 2B is an enlarged and detailed view of the embodiment shown in FIG. 2A. It is visible that the ventilation device 27 and the discardable plug 32 are integrated into at least one outer guide vane strut 28. The at least one outer guide vane strut 28 is located in the bypass airflow passage FBP and arranged between the nacelle 30 casing and the rear inner casing 60 of the core engine EC, comprising at least one pyrotechnical opening mechanism 100 and at least one discardable plug 32, here shown in closed position CP. In particular at least one pyrotechnical opening mechanism 100 is arranged and coupled to the rear inner casing 60 of the core engine EC and at least one discardable plug 32 is arranged and coupled to the nacelle 30 casing. The inside of the outer guide vane strut 28 is at least partially coated with a heat-resistant coating 29.2 in order to withstand the hot core airflow FC and protect from damage.

The bypass airflow FB flows around the outer guide vane strut 28.

FIG. 2C is an another exemplary illustration of an embodiment of FIG. 2B, wherein the difference is that the at least one pyrotechnical opening mechanism 100 and the discardable plug 32 are in open position OP such that the ventilation device 27 ventilates, or releases, through a hole in the inner casing 60 of the core engine EC and a hole in the bypass duct 20 and nacelle 30 the hot core airflow FC out of the gas turbine G to an outer environment EN. FIG. 3 is an exemplary illustration of another embodiment in an enlarged plan view and in a sectional view along the line A-A. Such implementation may be applied to the previously shown, or other, embodiments. The pyrotechnical opening mechanism 100 comprises an arm device 112 exemplary illustrated as a ring, which is arranged around a ring carrier 114 with assembly holes 108. The arm device 112 and the ring carrier 114 are rotationally symmetric around a ventilation axis AV, which may be substantially oriented along a ventilation flow path of the released core airflow FC. The arm device 112 comprises a pyrotechnical charge 124 which is integrated into the arm device 112, an interface 120, and an electro-pyro detonator 122, wherein at least the interface 120 is connected to an igniter 116, wherein the igniter 116 is connected to an EEC. The EEC may transmit a shaft failure signal 118 to the igniter 116, the igniter 116, upon receiving the shaft failure signal 118, transmits a detonation signal 110, causing a spark/ short circuit 130 at the interface 120 with the electro-pyro detonator 122. The at least one electro-pyro detonator 122 transmits the spark/ short circuit 130 to the pyrotechnical charge 124 which may blast.

FIG. 4 illustrates exemplary steps of a method for protecting against a turbine overspeed and/or a shaft failure 50 in a gas turbine G. The method comprises the detection of a turbine overspeed event and/or a shaft failure event S1. The method may further comprise cutting off the fuel (i.e. interrupting the supply of fuel), S2. Optionally, as shown here, the method may further comprise setting the bleed valves to the fully open position, S3. Optionally, as shown here, the method may further comprise moving the variable stator vanes into the closed position, S4. The method further comprises setting the ventilation device 27 to explosively open the core airflow passage FCP, i.e. release the core airflow FC of the core airflow passage FCP, S5. This particularly includes activating the pyrotechnical opening mechanism 100 According to step S1, the electronic engine controller EEC detects a turbine overspeed and/or a shaft failure 50 and identifies it as such. The order of steps S1, S2, S3, S4, S5 may be adapted, or one or more of the steps S1, S2, S3, S4 may be omitted, according to the requirements and/or the actual configuration of the turbine overspeed protection system. In particular, two or more of the steps S2, S3, S4, S5 may be executed in a parallel manner.

Cutting off the fuel prevents that the combustor would continue to create hot gas that needs to be directed outwards through the ventilation device 27. Preferably, the fuel cut-off, S2 precedes, or is simultaneously to, the activation of the pyrotechnical opening mechanism 100, S5. Optionally, a confirmation signal may be provided after the fuel cut off is established. Such confirmation signal may be a condition to be fulfilled prior to the explosive opening of the core airflow passage FCP, S5.

FIG. 5 illustrates explanatory and simplified a block diagram schematically representing components of a turbine overspeed and/or a shaft failure 50 detection in accordance with the present disclosure. The turbine overspeed protection system 230 comprises a failure detection means 240, configured to send an overspeed and/or shaft failure signal 200 to the electronic controller 220. The electronic controller 220 is configured, upon receiving the shaft failure signal 200, to trigger the fuel metering unit 140 via a fuel shut-off signal 260, to cut off the fuel supply to the gas turbine engine combustor.

The electronic controller 220 is further configured, upon receiving the shaft failure signal 200, to open the handling bleed valves 160 of the high-pressure compressor 16. Opening and closing of the handling bleed valves 140 is controlled by the electronic controller 220 via a handling bleed valve open signal 280.

By means of a variable stator vanes close signal 300, the electronic controller 220 is configured to set the variable stator vanes 180 into a closed position, in which airflow through the high-pressure compressor 16 may be partially blocked. It may be provided that the electronic controller 220 determines both the amount of the change in position of the handling bleed valves 140 and the change of the position of the variable stator vanes 180 before setting the ventilation device 27 to the open position. Further the electronic controller 220 is configured to set the at least one ventilation device 27 to its open position by means of an actuation signal 320, so that the core airflow passage FCP is ventilated.

The electronic controller 220 is further able to determine if the handling bleed valves 160, the variable stator vanes 180 and/or the ventilation device 27 have been set to closed and/or open position by determining the rate of change of the position of them as indicated by the signal provided by the turbine overspeed protection system 230.

The above description is intended for illustrative purposes only and is not intended to limit the scope of the present disclosure in any way. The method described herein can be performed in any suitable order unless contradicted by context. Various features of the various embodiments disclosed herein can be combined in different combinations to create new embodiments within the scope of the present disclosure.

List of reference signs (part of the description)

Low-pressure compressor 16 High-pressure compressor 18 Combustor 19 High-pressure turbine Bypass duct 21 Low-pressure turbine 22 Low-pressure shaft 23 Turbine section 24 Exhaust section Cone 26 Annular exit cross section 27 Ventilation device 28 Outer guide vane strut 29.1 Heat-resistant coating ejector 29.2 Heat-resistant coating outer guide vane strut Nacelle 32 Discardable plug 40 Ejector 42 Frame 46 Ejector body 46.1 Upper ejector body 46.2 Lower ejector body 48 Venturi-type valve Turbine overspeed and/or a shaft failure Inner casing Pyrotechnical opening mechanism 108 Assembly hole 110 Detonation signal 112 Arm device 114 Ring carrier 116 Igniter 118 Shaft failure signal 120 Interface 122 Electro-pyro detonator 124 Pyrotechnical charge Spark/ Short circuit Fuel metering unit Handling bleed valve 180 Variable stator vanes Overspeed/ shaft failure signal 220 Electronic controller 230 Turbine overspeed protection system 240 Failure detection means 260 Fuel shut-off signal 280 Handling bleed valve close signal 300 Variable stator vanes close signal 320 Actuation signal A Longitudinal or rotational axis AE Ejector axis AV Ventilation axis E Aircraft engine EC Core engine EEC Engine Electronic Controller EN Outer environment of the nacelle / gas turbine FB Bypass airflow FBP Bypass airflow passage FC Core airflow FCP Core airflow passage FEP Ejector airflow passage FSP Strut airflow passage G Gas Turbine OP Open position CP Closed position R Radial axis 51 Detection of failure S2 Fuel cut off S3 Set handling bleed valve to open position S4 Set variable stator vanes to closed position S5 Ventilation device opening

Claims

CLAIMS1. A gas turbine comprising a bypass duct (20) and a core engine (EC) together defining a bypass airflow passage (FBP) configured to guide a bypass airflow (FB); the core engine (EC) defining a core airflow passage (FCP), configured to guide a core airflow (FC); wherein the gas turbine has at least one ventilation device (27), configured to release the core airflow (FC) out of the core airflow passage (FCP); and the ventilation device (27) comprises at least one pyrotechnical opening mechanism (100).
2. The gas turbine of claim 1, wherein the pyrotechnical opening mechanism (100) comprises a ring carrier (114), and an arm device (112), wherein the arm device (112) comprises an at least one electro-pyrotechnical detonator (122) and at least one pyrotechnical charge (124).
3. The gas turbine of claim 1 or 2, wherein the pyrotechnical opening mechanism (100) comprises an igniter (116), and the arm device (112) comprises an interface (120) to the igniter (116), wherein the igniter (116) is connected to an Electronic Engine Control (EEC).
4. The gas turbine of any preceding claim, wherein the pyrotechnical opening mechanism (100) is configured to be activated via a shaft failure signal (118) from the Electronic Engine Control (EEC)to the igniter (116).
5. The gas turbine of claim 4, wherein the igniter (116), upon receiving the shaft failure signal (118), is configured to generate a detonation signal (110), wherein the detonation signal (110) is configured to generate preferably a spark and/or a short circuit (130).
6. The gas turbine of claim 5, wherein the spark and/or the short circuit (130) is configured to cause an explosive response of the pyrotechnical charge (124).
7. The gas turbine of any preceding claim, wherein the ventilation device (27) is arranged on the inner casing (60) of the turbine section (23) in particular arranged upstream of the low-pressure turbine (21).
8. The gas turbine of any preceding claim, wherein the ventilation device (27) is arranged between the inner casing (60) of the turbine section (23) and bypass duct (20) at an axial position upstream of the low-pressure turbine (21).
9. The gas turbine of any preceding claim, wherein the ventilation device (27) comprises an ejector (40), configured to ventilate into the bypass airflow passage (FBP).
10. The gas turbine of claim 9, wherein the ejector (40) is a Venturi-type valve.
11. The gas turbine of claim 9 or 10, wherein the ejector is at least partially coated with a heat-resistant coating (29.1, 29.2).
12. The gas turbine of any one of claims 1 to 9, wherein the ventilation device (27) is integrated in an outlet guide vane strut (28), wherein the ventilation device (27) is configured to release the core airflow (FC) through the at least one outlet guide vane strut (28), to an outer environment (EN) of a nacelle (30) of the gas turbine (G).
13. An aircraft engine (E) including: a gas turbine (G) according to preceding claim; and a turbine overspeed protection system (230), which is configured to control the ventilation device (27) wherein the at least one ventilation device (27) is part of the turbine overspeed protection system (230).
14. An aircraft including the gas turbine of any one of claims 1 to 12, and/or an aircraft engine (E) of claim 13.
15. A method for protecting against a turbine overspeed event and/or a shaft failure event in a gas turbine which comprises an overspeed protection system (230), the method comprising the steps of: detecting (S1) a turbine overspeed event and/or a shaft failure event; cutting off (S2) fuel; and actuating (S5) a ventilation device (27) to release a core airflow (FC) out of a core airflow passage (FCP) of the gas turbine (G).
16. The method of claim 15, further comprising: moving (S4) variable stator vanes (180) into a closed position which restrict airflow through a compressor; and/or setting (S3) handling bleed valves (160) to open position releasing airflow through a compressor to a bypass airflow passage (FBP) of the gas turbine engine.
17. An electronic controller (220) that is configured to perform the method of claim 15 or 16. 10