DE102019113530A1 - Device for sealing a gap between two components of a turbine of a gas turbine engine - Google Patents

Abstract

The invention relates to a device for sealing a gap (57) between two components (4, 6) of a turbine of a gas turbine engine. The device comprises: a turbine guide vane ring (400) with a plurality of turbine guide vane segments (4), each turbine guide vane segment (4) having an outer platform (41) with a radially extending first wall (5) which is provided with a straight sealing edge (51); and a component (6) adjoining the outer platform (41) downstream in each case and having a radially extending second wall (7) which forms a flat surface (70). The gap (57) to be sealed runs between the first wall (5) and the second wall (7). A longitudinally extending sealing element (8) with an L-shaped cross-section is provided, which has two first and second legs (81, 82) each running in a straight line and arranged perpendicular to one another, the sealing element (8) with its first leg (81 ) is connected with axial play and radial play with the outer platform (41) or with its second leg (82) with axial play and radial play with the adjacent component (6), with the first leg (81) on the to seal the gap (57) in the circumferential direction straight sealing edge (51) of the first wall (5) rests and the second leg (82) rests on the flat surface (70) of the second wall (7).

Description

The invention relates to a device for sealing a gap between two components of a turbine of a gas turbine engine according to the preamble of claim 1.

In the high-pressure turbine of a gas turbine engine, it is necessary to seal a gap in the first turbine stage behind the combustion chamber, which gap extends between the radially outer platform of the turbine guide vane ring and the liner segment adjoining it in the flow direction, the liner segment being in the area of the turbine blade ring the first turbine stage limits the main flow path through the gas turbine engine radially on the outside. Above such a gap there is a considerable pressure drop in the radial direction of, for example, up to 10 bar, which results from the pressure difference between the air under a lower pressure in the main flow path and the bleed air under a higher pressure, the radial outside of the main flow path flows and is used in particular to cool the high pressure turbine. An efficient sealing of such a gap is therefore of great importance.

In order to provide such a seal, it is known that the turbine guide vane segments of a turbine guide vane ring have sealing edges (so-called "chordal seals") running in a straight line in the circumferential direction, which ideally each lie along a straight line on a flat surface of the adjoining liner segment. However, there is the problem that deformations of the sealing edge and / or deformations of the flat surface of the adjacent liner segment on which the sealing edge rests and also relative movements between the adjacent components can result in small gaps or openings that impair the sealing function. Because of the high pressure drop across the gap, even small gaps can lead to considerable leakage losses, which must be avoided.

Similar problems can also arise with other turbine stages.

The present invention is based on the object of providing a device for sealing a gap between two components of a turbine of a gas turbine engine that provides an improved seal between the two components.

This object is achieved by a device with the features of claim 1. Refinements of the invention are given in the dependent claims.

Accordingly, the invention relates to a device for sealing a gap between components of a turbine of a gas turbine engine, which comprises a turbine guide vane ring which has a plurality of turbine guide vane segments, each turbine guide vane segment comprising an outer platform and at least one guide vane, and wherein the outer platform has a radially extending first wall which is provided at its radially outer end with a sealing edge running in a straight line in the circumferential direction. The device further comprises a component adjoining the outer platform of the turbine guide vane segment downstream, which component has a radially extending second wall which in the axial direction directly adjoins the first wall of the outer platform and forms a flat surface towards the first wall . The gap to be sealed runs between the radially extending first wall of the outer platform and the radially extending second wall of the adjoining component.

According to the invention, a longitudinally extending, in cross-section L-shaped sealing element is provided, which has two rectilinear first and second legs arranged perpendicular to each other, the sealing element with its first leg with axial play and radial play with the outer platform or with its second leg is connected to the adjacent component with axial play and radial play. To seal the gap, the first leg rests on the sealing edge of the first wall that runs in a straight line in the circumferential direction, and the second leg rests on the flat surface of the second wall. During operation of the gas turbine engine, when there is a pressure drop across the gap to be sealed, the L-shaped sealing element is pressed firmly against the sealing edge of the first wall and the flat surface of the second wall.

The present invention is based on the knowledge that by providing an L-shaped sealing element which is loosely attached, ie with an axial play and a radial play, and which rests with its one leg on the sealing edge and with its other leg on the flat surface of the adjacent wall rests, the gap extending between the two components can be safely and reliably sealed. The axial play and radial play of the attachment of the L-shaped sealing element provides the necessary flexibility so that the sealing element can adapt to relative movements between the outer platform and the adjoining component or associated relative movements between the two radially extending walls, so that the The occurrence of gaps or openings along the sealing surface is prevented and the sealing function despite such Relative movements are retained. It should be noted that the simultaneous implementation of an axial play and a radial play also allows the L-shaped sealing element to tilt or pivot to a certain extent about its longitudinal axis, so that tilting of the two adjacent components can also be compensated for.

In addition, it can be provided that the L-shaped sealing element is elastic in such a way that elastic deformations of the first wall and / or the second wall, which can occur during the operation of the gas turbine engine, can be compensated for at least to a certain degree, whereby the Sealing function between the two components is further improved. An elastic deformability of the L-shaped sealing element can result from the material of the sealing element without any special measures. For example, it can be provided that the L-shaped sealing element is designed as a metal sheet, the thickness and ductile properties of which are dimensioned in such a way that such deformability automatically exists.

It should be noted that the term L-shaped sealing element according to the present invention is to be understood in such a way that the two legs of the sealing element that form the “L” can of course have a different width, but also the same width. It is further pointed out that the L-shaped sealing element has a first straight leg and a second straight leg, which are arranged perpendicular to one another, but this does not preclude the L-shaped sealing element from forming further surfaces. For example, it can be provided that a further surface protrudes vertically and radially downward from the first leg. The L-shape is thus contained in the sealing element, but can be supplemented by other shapes in design variants.

According to one embodiment of the invention it is provided that the L-shaped sealing element is connected along its length at several fastening points with axial play and radial play with the outer platform or with the adjacent component. The L-shaped sealing element is thus fastened at certain points by means of corresponding fastening elements. A consequence of the loose fastening of the L-shaped sealing element, provided with axial play and radial play, by means of spaced fastening points is an additional tiltability of the sealing element about an axis which runs transversely to its longitudinal axis.

One embodiment of the invention provides that the L-shaped sealing element is fastened to the radially extending wall of the outer platform by means of at least two fastening elements. For this purpose, it can be provided, for example, that the fastening elements are designed as metal tabs which are each fastened to the first wall with an elongated section and, starting from the elongated section, form a substantially closed loop that extends through a recess in the first leg with axial play and radial play , wherein the recess is formed in a region of the L-shaped leg which protrudes axially forwards with respect to the sealing edge. In this way, a stable connection is provided in a simple manner, which enables axial play and radial play of the fastening.

Another embodiment of the invention provides that the L-shaped sealing element is fastened to the radially extending wall of the adjoining component via at least two fastening elements. For this purpose, one embodiment variant provides that the fastening elements each reach through a recess formed in the second leg with axial play and radial play and are connected to the second wall. The fastening elements are designed, for example, as pins, bolts or rivets, it being possible for the respective fastening element to have a head that is larger than the recess so that the second leg is axially secured to the second wall.

A further embodiment provides that the outer platforms of the turbine guide vane segments each form an axially running groove on their end faces and a sealing strip is inserted into the grooves of two adjacent platforms, which serves to seal two adjacent outer platforms. Such sealing strips are known per se. It is further provided that the sealing strip forms a fastening section in the downstream direction which protrudes from the grooves, the fastening section forming a holding element via which the first leg is fixed in the radial direction on the first wall. According to this embodiment, the L-shaped sealing element is thus fastened to the first wall via an extension of the sealing strip which protrudes from the grooves and is designed as a fastening section.

In this embodiment, a radial play can be provided in that the holding element of the fastening section is designed to be resilient and thus exerts a spring force on the first leg of the L-shaped sealing element this presses while providing a radial play from above onto the sealing edge of the first wall.

It is provided that the fastening section of the sealing strip forms two holding elements, via which the two L-shaped sealing elements of two adjacent outer platforms are each fastened at one of their ends to the first wall of the respective outer platform. In each case one of the two holding elements that the fastening section of a sealing strip forms is used to fasten one end of an L-shaped sealing element. It can be provided that the fastening section is overall roof-shaped, with the two holding elements extending at an angle of less than 180 ° to one another. The roof-shaped configuration provides a spring force which the holding elements exert on the respective sealing element in a simple manner.

The holding elements of the fastening section of the respective sealing strip are, for example, rectangular.

Another embodiment provides that the first legs each form a widened area at their ends, on which the respective holding element rests. This ensures a reliable connection between the holding element of the fastening section and the leg.

It can further be provided that the ends of the first legs each have a slot through which the fastening section runs, the slot having a greater thickness than the fastening section to provide axial play. The fastening section secures the holding element in the axial direction with respect to the first wall. Because the slot through which the fastening section runs has a width which is greater than the thickness of the fastening section, the fastening section can move in the axial direction in this slot, so that axial play is provided.

According to a further embodiment, the L-shaped sealing element is part of a triangular carrier, the L-shaped sealing element forming the base of the triangular carrier and the two other sides of the triangular carrier being formed by two connecting webs which extend obliquely from the two ends of the L-shaped sealing element extend axially forward and which are attached at their connection point with an axial play on a fastening element of the outer platform. This fastening element is, for example, a fastening element which is used to fasten sealing plates which are arranged between the outer platform and an upstream combustion chamber. This configuration is associated with the advantage that a very stable axial mounting of the L-shaped sealing element can be provided.

A further variant of the invention combines this exemplary embodiment with the exemplary embodiment in which the sealing strip is inserted between two outer platforms adjoining one another in the circumferential direction for fastening the L-shaped sealing element. According to this variant of the invention, the L-shaped sealing element is fastened to the fastening element of the outer platform via the triangular support with axial play and to the first wall with radial play using retaining tabs of the fastening sections.

According to a further exemplary embodiment, at least one of the legs of the L-shaped sealing element has one or more recesses. These can serve, for example, to keep cooling air bores that are formed in the adjacent component free.

Further refinements of the invention can also provide for a play in the fastening of the L-shaped sealing element in the circumferential direction.

The turbine guide vane ring designed according to the present invention is, according to one embodiment, the turbine guide vane ring of the first turbine stage of the high-pressure turbine arranged behind the combustion chamber of the gas turbine engine. However, the present invention is not reduced to this, since the problem of sealing gaps also exists between other components of a turbine.

The adjoining component is, for example, a liner segment connected to the outer housing of the high pressure turbine, which delimits the flow path through the high pressure turbine in the area of the rotor of the first turbine stage radially on the outside. Such a liner segment typically has a starting material adjacent to the flow path, into which the rotor blades of the rotor blade ring of the turbine stage can work in order to minimize a gap.

According to a further aspect of the invention, the present invention relates to a gas turbine engine with a device according to claim 1.

It should be noted that the present invention is described with reference to a cylindrical coordinate system having 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 through defines the engine axis of the gas turbine engine in which the present invention is implemented, with the axial direction pointing from the engine entrance towards the engine exit. 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 comprise an engine core comprising 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 gearbox 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 can 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. For example only, the turbine connected to the core shaft can be a first turbine, the compressor connected to the core shaft can be a first compressor and the core shaft can be a first core shaft. The 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 example above). For example, the transmission can be arranged to be driven only by the core shaft which is configured to rotate (e.g. in use) at the lowest speed (e.g. only the first core shaft and not the second core shaft in the above example) will. 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 that is described and / or claimed here, a combustion chamber can 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 several 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 (e.g. the first turbine and the second turbine as described above) can comprise any number of stages, e.g. multiple stages. Each stage can include a number of rotor blades and a number of stator blades. The row of rotor blades and the row of stator blades can be axially offset from one another.

Each fan blade may be defined with a radial span extending from a root (or hub) on a radially inward of Gas overflow point or extends at a position of a span of 0% to a peak at a position of a span of 100%. 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 (ie, 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 the axially most forward edge) of the blade. The hub-to-tip ratio relates, of course, to the portion of the fan blade overflowing with gas, ie the portion which is located 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 during 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, e.g. at the front edge of the tip , (which can be defined as the fan tip radius at the front edge multiplied by the angular velocity). The fan peak load at constant speed conditions can be more than (or on the order of): 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38 , 0.39, or 0.4 (all units in this section are Jkg ^-1 K ^-1 / (ms ^-1 ) ² ). The fan peak load can be in 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 flow through the bypass duct to the mass flow rate of flow through the core at constant velocity conditions. In some arrangements, the bypass ratio can be more than (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 are (lie). 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, may the total pressure ratio of a gas turbine engine described and / or claimed herein at constant speed is greater than (or on the order of): 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio 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).

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 that is 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 (lying). 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 as described and / or claimed herein can have any maximum thrust desired. As a non-limiting example only, a gas turbine described and / or claimed herein can be used to generate a maximum thrust of at least (or on the order of): 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 form 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 of materials. 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 a composite with an organic matrix, such as e.g. B. carbon fiber. As another example, at least a portion of the fan blade and / or the blade can be at least in part made of a metal, such as metal. A titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy), or a steel-based material. The fan blade 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 manufactured using titanium or a titanium-based alloy, for example. Thus, as an example only, the fan blade may comprise 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 to fix the fan blade to the hub / disc. As another example For example, the fan blades may 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 that are described and / or claimed here may or may not be provided with a VAN (Variable Area Nozzle - nozzle with variable cross section). Such a nozzle with a variable cross section can allow the 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 , 18th , 20th or 22nd Fan blades, have.

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. will.

By way of example only, the forward speed under the constant speed condition may be at any point in the range of Mach 0.7-0.9, e.g. 0.75-0.85, e.g. 0.76-0.84, e.g. 0.77-0 .83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example in the order of Mach 0.8, in the order of Mach 0.85 or in the range from 0.8 to 0, 85 lie. Any speed within these ranges can be the constant travel condition. In some aircraft, the cruise control conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

By way of example only, the constant velocity conditions may mean 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 herein can be applied to any aspect and / or combined with any other feature or parameter described herein, if they can not mutually exclusive.

• 1 eine Seitenschnittansicht eines Gasturbinentriebwerks, in dem die vorliegende Erfindung realisierbar ist;
• 2 in Schnittdarstellung einen Abschnitt eines Gasturbinentriebwerks, der den stromabwärtigen Bereich der Brennkammer und die sich daran anschließende erste Turbinenstufe der Hochdruckturbine umfasst, wobei die erste Turbinenstufe einen Turbinen-Leitschaufelkranz mit einer äußeren Plattform und ein in Strömungsrichtung an die äußere Plattform angrenzendes Linersegment aufweist, wobei ein sich zwischen diesen beiden Komponenten radial erstreckender Spalt durch ein L-förmiges Dichtelement abgedichtet ist;
• 3 in schematischer und perspektivischer Darstellung eine Ausführungsvariante eines L-förmigen Dichtelements und dessen Anbringung an einer sich radial erstreckende Wand einer äußeren Schaufelplattform;
• 4 die Ausführungsvariante der 3 in einer Seitenansicht;
• 5 in schematischer und perspektivischer Darstellung eine weitere Ausführungsvariante eines L-förmigen Dichtelements und dessen Anbringung an einer sich radial erstreckende Wand eines Linersegments;
• 6 die Ausführungsvariante der 5 in einer Seitenansicht;
• 7 in perspektivischer Darstellung ein Ausführungsbeispiel, bei dem ein L-förmiges Dichtelement gemäß den 3 und 4 zur Abdichtung eines Spalts zwischen einer sich radial erstreckende Wand einer äußeren Schaufelplattform und einer sich radial erstreckende Wand eines Linersegments vorgesehen ist;
• 8 in perspektivischer Darstellung ein weiteres Ausführungsbeispiel, bei dem ein L-förmiges Dichtelement zur Abdichtung eines Spalts zwischen einer sich radial erstreckende Wand einer äußeren Schaufelplattform und einer sich radial erstreckende Wand eines Linersegments vorgesehen ist, wobei das L-förmige Dichtelement durch einen Befestigungsabschnitt eines Dichtungsstreifens an der sich radial erstreckende Wand der äußeren Schaufelplattform befestigt ist;
• 9 das Ausführungsbeispiel der 8 in einer seitlichen Schnittansicht;
• 10 das Ausführungsbeispiel der 8 mit einer vergrößerten Darstellung des L-förmigen Dichtelements und der Komponenten, an dem dieses befestigt ist;
• 11 die Darstellung der 10 in einer seitlichen Schnittansicht;
• 12 eine perspektivische Ansicht in Umfangsrichtung des L-förmigen Dichtelements und von Befestigungsabschnitten gemäß der 8, mittels derer dieses an der sich radial erstreckenden Wand der äußeren Schaufelplattform befestigt ist;
• 13 eine perspektivische Ansicht von vorne des L-förmigen Dichtelements und eines Befestigungsabschnitts gemäß der 8, mittels dessen dieses an der sich radial erstreckenden Wand der äußeren Schaufelplattform befestigt ist;
• 14 eine weitere perspektivische Ansicht des L-förmigen Dichtelements und eines Befestigungsabschnitts gemäß der 8, mittels dessen dieses an der sich radial erstreckenden Wand der äußeren Schaufelplattform befestigt ist;
• 15 eine weitere Darstellung gemäß der 10 in einer perspektivischen, teilweise geschnittenen Ansicht;
• 16 eine perspektivische Darstellung der Verbindung zwischen einem Halteelement eines Befestigungsabschnitts eines Dichtungsstreifens und einem vergrößerten Endbereich des ersten Schenkels des L-förmigen Dichtelements;
• 17 in perspektivischer Darstellung ein Ausführungsbeispiel gemäß der 8, wobei zwei in Umfangsrichtung aneinander angrenzende Turbinen-Leitschaufelsegmente dargestellt sind;
• 18 ein Ausführungsbeispiel eines L-förmigen Dichtelements in perspektivischer Ansicht;
• 19 ein weiteres Ausführungsbeispiel eines L-förmigen Dichtelements in perspektivischer Ansicht;
• 20 ein weiteres Ausführungsbeispiel eines L-förmigen Dichtelements in perspektivischer Ansicht;
• 21 ein weiteres Ausführungsbeispiel eines L-förmigen Dichtelements in perspektivischer Ansicht;
• 22 ein weiteres Ausführungsbeispiel eines L-förmigen Dichtelements in perspektivischer Ansicht; und
• 23 in perspektivischer Darstellung ein weiteres Ausführungsbeispiel, bei dem ein L-förmiges Dichtelement zur Abdichtung eines Spalts zwischen einer sich radial erstreckenden Wand einer äußeren Schaufelplattform und einer sich radial erstreckenden Wand eines Linersegments vorgesehen ist, wobei das L-förmige Dichtelement Bestandteil eines Dreiecksträgers ist.

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 Fig. 3 is a side sectional view of a gas turbine engine in which the present invention may be implemented;
• 2 a section of a gas turbine engine comprising the downstream region of the combustion chamber and the adjoining first turbine stage of the high-pressure turbine, the first turbine stage having a turbine guide vane ring with an outer platform and a liner segment adjoining the outer platform in the direction of flow, wherein a between these two components the radially extending gap is sealed by an L-shaped sealing element;
• 3 a schematic and perspective representation of an embodiment variant of an L-shaped sealing element and its attachment to a radially extending wall of an outer blade platform;
• 4th the variant of the 3 in a side view;
• 5 in a schematic and perspective illustration a further embodiment of an L-shaped sealing element and its attachment to a radially extending wall of a liner segment;
• 6th the variant of the 5 in a side view;
• 7th in a perspective view an embodiment in which an L-shaped sealing element according to the 3 and 4th for sealing a gap between a radially extending wall there is provided an outer paddle platform and a radially extending wall of a liner segment;
• 8th a perspective view of a further embodiment in which an L-shaped sealing element is provided for sealing a gap between a radially extending wall of an outer blade platform and a radially extending wall of a liner segment, the L-shaped sealing element being attached by a fastening section of a sealing strip the radially extending wall of the outer vane platform is attached;
• 9 the embodiment of 8th in a side sectional view;
• 10 the embodiment of 8th with an enlarged view of the L-shaped sealing element and the components to which it is attached;
• 11 the representation of the 10 in a side sectional view;
• 12 a perspective view in the circumferential direction of the L-shaped sealing element and of fastening sections according to FIG 8th by means of which it is attached to the radially extending wall of the outer vane platform;
• 13 a perspective view from the front of the L-shaped sealing element and a fastening portion according to FIG 8th by means of which it is attached to the radially extending wall of the outer vane platform;
• 14th a further perspective view of the L-shaped sealing element and a fastening portion according to FIG 8th by means of which it is attached to the radially extending wall of the outer vane platform;
• 15th a further representation according to the 10 in a perspective, partially sectioned view;
• 16 a perspective view of the connection between a holding element of a fastening portion of a sealing strip and an enlarged end region of the first leg of the L-shaped sealing element;
• 17th in a perspective view an embodiment according to FIG 8th Two turbine guide vane segments adjoining one another in the circumferential direction are shown;
• 18th an embodiment of an L-shaped sealing element in a perspective view;
• 19th a further embodiment of an L-shaped sealing element in a perspective view;
• 20th a further embodiment of an L-shaped sealing element in a perspective view;
• 21st a further embodiment of an L-shaped sealing element in a perspective view;
• 22nd a further embodiment of an L-shaped sealing element in a perspective view; and
• 23 a perspective view of a further embodiment in which an L-shaped sealing element is provided to seal a gap between a radially extending wall of an outer blade platform and a radially extending wall of a liner segment, the L-shaped sealing element being part of a triangular support.

1 represents a gas turbine engine 10 with a main axis of rotation 9 represent. The engine 10 includes an air inlet 12 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 21st 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 an epicycloidal gear 30th on the low pressure turbine 19th attached and is driven by this.

In use, the core air flow is A 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 combustion products then propagate through the high pressure and low pressure turbines 17th , 19th and thereby propel them before they are used to provide a certain thrust through the nozzle 20th be expelled. The high pressure turbine 17th drives the high pressure compressor 15th through a suitable connecting shaft 27 at. The blower 23 generally provides the majority of the thrust. The epicycloidal gear 30th is a reduction gear.

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 writings, the “low pressure turbine” and “low pressure compressor” referred to here may alternatively be known as the “medium pressure turbine” and “medium pressure compressor”. Using such alternative nomenclature, the fan 23 may be referred to as a first compression stage or the lowest pressure compression stage.

Other gas turbine engines to which the present disclosure may find application may have alternative configurations. For example, such engines can have an alternative number of compressors and / or turbines and / or an alternative number of connecting shafts. As another example, the 1 gas turbine engine shown has a split flow nozzle 20th , 22nd on, which means that the flow through the bypass duct 22nd has its own nozzle, that of the engine core nozzle 20th is separate and radially outside thereof. 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 turbo-prop 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.

In the context of the present invention, the sealing of a gap is important in the first turbine stage of the high pressure turbine 17th runs between the outer platform of the turbine guide vane segments of the turbine guide vane ring and the liner segments arranged downstream.

The 2 shows a sectional view of a section of the core engine of a gas turbine engine, the illustrated section - in relation to the flow direction - the rear section of a combustion chamber 16 , a turbine guide vane segment 4th one to the combustion chamber 16 directly adjacent turbine guide vane ring 400 as well as a blade 31 a turbine rotor blade ring 310 shows. The Guide vane ring 400 and the blade ring 310 a turbine stage of the high pressure compressor 17th .

The combustion chamber 16 includes an outer combustion chamber wall 161 and an inner combustion chamber wall 162 , where the terms “outer” and “inner” refer to the main flow path 24 refer, which runs through the core engine. To protect against the flow of hot gas from the combustion chamber 16 is the outer combustion chamber wall 161 with a plurality of heat shingles 163 provided, which is located on the outer combustion chamber wall 161 prop up. The inner combustion chamber wall is also in a corresponding manner 162 with a plurality of heat shingles 164 provided, which is located on the inner combustion chamber wall 162 prop up.

The outer combustion chamber wall 161 forms part of an outer combustion chamber housing, of which another wall structure 165 is shown. The inner combustion chamber wall 162 forms part of an inner combustion chamber housing that also includes further wall structures.

The turbine guide vane ring 400 is segmented and includes a plurality of turbine vane segments 4th that adjoin one another in the circumferential direction. Each turbine vane segment 4th of the turbine guide vane ring 400 includes an outer platform 41 which is the main flow path 24 bounded radially outside by the core engine, an inner platform 42 which is the main flow path 24 limited radially on the inside by the core engine, and at least one guide vane 43 that is between the inner platform 42 and the outer platform 41 extends. The individual turbine guide vane segments 4th border on their platforms 41 , 42 frontally against each other in the circumferential direction.

Both in the radially outer platform 41 as well as in the radially inner platform 42 is a groove running essentially in the axial direction in the end face of the platform 411 , 421 educated. The grooves 411 , 421 serve to each have a sealing strip (in the 2 not shown), which is also essentially in the axial direction in the grooves 411 , 421 runs and thereby two radially inner platforms 42 or two radially outer platforms 41 , which abut one another at the front, sealed against one another. Such grooves 411 , 421 and sealing strips arranged in these are known per se.

To seal a radial gap between the combustion chamber 16 and the turbine vane segments 4th is formed are a plurality of sealing plates 32 provided, which are each elongated and form an arc. The sealing plates 32 are for example via rivets that are attached to the fastening projections 44 the respective platform 41 , 42 are attached, held and provided with a contact pressure via spring elements.

The blades 31 of the turbine blade ring 310 are driven by the high pressure shaft of the gas turbine engine. In the area of the turbine blade ring 310 becomes the flow path 24 through the high pressure turbine 17th radially outside through so-called liner segments 6th limited to the downstream of the outer platforms 41 of the turbine guide vane segment 4th are arranged. The liner segments 6th are corresponding to the turbine guide vane segments 4th segmented and also adjoin one another in the circumferential direction. The liner segments 6th are with an outer casing 33 the high pressure turbine 17th connected and are structurally held by this.

It can be provided that the liner segments 6th a starting material 60 have that the flow path 24 through which the high pressure turbine 17th limited. It is in the design of the 2 provided that the blades 31 of the rotor blade ring 310 the high pressure turbine 17th are formed without a shroud, so that the blade tips are spaced apart by a gap opposite the outer flow path limitation. By providing a start-up material 60 , into which the blade tips can work, this gap can be minimized.

In the context of the present invention it is important that the outer platform 41 a radially extending first wall at its downstream end 5 which is provided at its radially outer end with a sealing edge running in a straight line in the circumferential direction, as will be explained in more detail with reference to the further figures. It also includes the liner segment 6th a radially extending second wall at its upstream end 7th , with the two radially extending walls 5 , 7th directly adjoin one another in the axial direction.

Between the outer platform 41 and the liner segment 6th namely between the radially extending wall 5 the outer platform 41 and the also radially extending wall 7th of the liner segment 6th a gap also extends in the radial direction 57 . This gap connects the main flow path 24 through the high pressure turbine 17th with a flow area, also known as a “secondary flow area”, that is located radially outside the main flow path 24 is located and flows in the bleed air that cools components of the high-pressure turbine 17th serves. It applies here that the pressure of the bleed air is significantly higher than the pressure of the gas in the main flow path 24 . The pressure difference can be up to 10 bar, for example. Accordingly lies over the gap 57 a very high pressure. When the gap 57 is not completely sealed, there is a significant leakage flow into the main flow path 24 thereby reducing the efficiency of the gas turbine engine.

To provide an effective seal for the gap 57 is a longitudinally extending sealing element with an L-shaped cross section 8th provided with axial play and radial play on the outer platform 41 or its radially extending wall 5 or with axial play and radial play on the liner segment 6th or its radially extending wall 7th is attached. Embodiments of this are in the 3 to 23 shown.

The 3 and 4th show, in a perspective view and in a side view, schematically a first exemplary embodiment for sealing a gap 57 that is between a first radially extending wall 5 and a second radially extending wall 7th runs, with the walls 5 , 7th according to the 2 Part of the outer platform 41 a turbine guide vane segment 4th or a liner segment 6th are.

The wall 5 has a sealing edge at its radially outer end 51 on. The sealing edge 51 extends over the length of the turbine guide vane segment in a straight line in the circumferential direction. It is therefore not curved. This is to create a straight line of contact with the adjacent wall 7th provide around which a tilting movement can take place. The adjacent wall 7th points to the wall 5 a flat surface 70 on. Between the wall 5 and the wall 7th runs the gap 57 that is to be sealed.

To seal the gap 57 is a longitudinally extending sealing element with an L-shaped cross section 8th intended. This comprises two longitudinally extending legs 81 , 82 which are arranged perpendicular to each other. The two legs 81 , 82 have a different width, but this is not necessarily the case. So is the leg in the illustrated embodiment 81 formed wider and includes this one area 811 in which the thigh 81 on the sealing edge 51 rests, and an area 812 in which the thigh 81 opposite the sealing edge 51 stands out.

The L-shaped sealing element 8th is along its length at two spaced attachment points P by means of metal tabs 90 on the first wall 5 attached. The metal tabs 90 each have an elongated section 91 on the one with the wall 5 attached, for example is welded to this. Starting from the elongated section 91 forms the metal tab 90 a substantially closed loop 92 off that one recess 83 in the first leg 81 of the sealing element 8th takes action. The recess 83 is at least partially in the area 812 of the thigh 81 formed, the opposite of the sealing edge 51 axially protrudes forward. Here is the recess 83 designed such that it has a width in the axial direction that is greater than the thickness of the metal ash 90 . This causes the L-shaped sealing element 8th in the axial direction opposite the metal tab 90 is movable, in other words with axial play on the wall 5 is attached.

In addition, a radial play of the fastening is provided in that the loop 92 some radial movement of the leg 81 inside the loop 92 allowed. The attachment via the metal straps 90 thus provides both axial play and radial play of the fastening.

In operation of the gas turbine engine, in which the components according to the 2 to 4th are arranged is above the gap 57 as explained before a high pressure drop. This pressure drop leads to the L-shaped sealing element 8th with his thigh 82 against the plane surface 70 the Wall 7th and with his thigh 81 against the also flat sealing edge 51 the Wall 5 is pressed. This will make the gap 57 securely locked. Relative movements between the wall 5 and the wall 7th can occur due to the loose attachment of the sealing element 8th be balanced. For example, if the wall 7th moved axially backwards, this can be compensated for by the axial play of the fastening in the fastening points P.

The 5 and 6th show, in a perspective view and in a side view, a further exemplary embodiment for sealing a gap 57 that is between a first radially extending wall 5 and a second radially extending wall 7th runs, with the walls 5 , 7th according to the 2 Part of the outer platform 41 a turbine guide vane segment 4th or a liner segment 6th are.

The difference to the embodiment of the 3 and 4th is that the L-shaped sealing element 8th along its length at two spaced attachment points P by fasteners on the second wall 7th is attached, the attachment in the embodiment of 5 and 6th over the second leg 82 he follows. The second leg is in this embodiment 82 wider than the first leg 81 .

To connect the leg 82 with the wall 7th are in the thigh 82 two elongated ones extending in the radial direction Recesses 84 intended. The recesses 84 extends through an elongated fastening element 95 , which is designed for example as a bolt or rivet. The fastener 95 has a head 96 whose diameter is greater than the width of the recess 84 so that the thigh 82 not off the wall despite axial play 70 can solve.

In order to provide axial play, it is provided that the head 96 at a certain axial distance from the wall 70 is arranged. To provide a radial play is provided that the elongated recess 84 has a length that is greater than the diameter of the fastener 95 , so that the leg can move radially 82 and thus the L-shaped sealing element 8th in the elongated recess 84 given is.

The 7th shows a perspective view of an embodiment in which an L-shaped sealing element 8th in the design of 3 and 4th to seal the gap between the outer platform 41 and the liner segment 6th is arranged, wherein to explain the other components on the description of 2 is referred. It is provided according to this exemplary embodiment that the second leg 82 in its middle area a recess 85 trains. This way they stay in the liner segment 6th trained cooling air holes 65 free and not covered.

The 8th to 17th show a further exemplary embodiment, which has the same structure of the radially outer blade platform 41 and the liner segment 6th like the 2 and 7th another way of fastening the L-shaped sealing element 8th provides. They show 8th and 9 in a perspective view and in a sectional view the overall arrangement, while the 10 to 17th show the type of attachment in an enlarged view and in numerous perspective views.

Referring first to the 8th and 9 can be seen that in the on the front sides of the platforms 41 , 42 formed axially extending grooves 411 , 421 Sealing strips 90 are used, which serve two circumferentially adjacent platforms 41 , 42 to seal against each other. It is provided that the sealing strip 90 that is in the outer platform 41 extends, a fastening portion in the downstream direction 91 that from the grooves 411 protrudes. This fastening section 91 is used to attach the L-shaped sealing element 8th . The weatherstrip 90 thus fulfills two functions, on the one hand the function of sealing two adjacent platforms 41 and on the other hand in its fastening section 91 the function of fastening the L-shaped sealing element 8th . To explain the other components of the 8th and 9 will refer to the description of the 2 Referenced.

The 10 and 11 show the fastening section in a perspective view and in a sectional view 91 and the L-shaped sealing element 8th in an enlarged view. Thereafter, the fastening section comprises 91 two substantially rectangular holding elements projecting in opposite directions 92 which together form a roof-shaped structure in the axial view from the front. A holding element exercises 92 in each case a spring force on the end of a first leg 81 an L-shaped sealing element 8th out. The two holding elements 92 a fastening section 91 hold the ends of two sealing elements 8th that are in adjacent turbine vane segments 4th are arranged.

The holding elements 92 are designed to be resilient so that they exert a spring force on the first leg 81 exercise, thereby the thigh 81 providing a radial play from above onto the sealing edge 51 the Wall 5 is pressed. The size of the radial play is determined by the extent to which the holding element 92 can be bent maximally radially outwards.

It should be noted that in the 10 and 11 and also sealing elements 61 , 62 of the liner segment 6th are shown. However, these are of no importance for the present invention, so that they are not explained in detail.

It should also be noted that the L-shaped sealing element 8th also an area that is folded down radially 83 that - based on the length of the sealing element 8th - in the middle area of the sealing element 8th is trained. This can, for example, be a stiffening of the sealing element 8th serve. In the middle of the thigh 82 is the same as in the exemplary embodiment 7th a recess 85 intended.

To provide an axial connection of the L-shaped sealing element 8th with the fastening section 91 while providing an axial play, the leg 81 a slot at each end 814 on, through which the fastening section 91 runs before the two holding elements arranged in a roof-like manner to one another at its end 92 trains. This is in the perspective detailed representations of the 12-14 to recognize. The slot has 814 a width in the axial direction which is greater than the thickness of the fastening portion 91 . This causes the thigh 81 in the axial direction opposite the Fastening section 91 is displaceable, the width of the slot 814 specifies the maximum axial displaceability.

Over the slot 814 in the first leg 81 of the L-shaped sealing element 8th and the holding elements 92 of the fastening section 91 becomes the L-shaped sealing element 8th on his thigh 81 thus providing both an axial play and a radial play on the wall 5 the outer platform 41 attached.

With reference to the 12 to 14th it is also pointed out that the leg 81 a widespread area at both ends 810 forms on which the respective holding element 92 rests. By widening the area 810 can be a comparatively wide retaining element 92 be used that has sufficient structural strength to support the thigh 81 and thus the L-shaped sealing element 8th safe on the wall 5 to fix.

The 15th and 16 additionally show a sectional view of the overall arrangement and a perspective detailed view of the connection between the holding element 92 and the common area 810 of the thigh 81 of the sealing element 8th .

The 17th shows a perspective view of two turbine guide vane segments adjoining one another in the circumferential direction 4th , 4a and liner segments adjoining it in the direction of flow 6th , 6a , with two L-shaped sealing elements 8th , 8a serve in the manner described to seal a gap between the respective components. It can be seen that the two holding elements 92 of a fastening section of a sealing strip, the respective end of two adjacent sealing elements 8th , 8a on the wall 5 establish.

The 18th to 22nd show different design variants of an L-shaped sealing element 8th . The basic structure with a first leg extending in a straight line 81 and a second straight leg 82 that to the first leg 81 is arranged vertically, is always given.

The structure of the sealing element 8th the 18th corresponds to the structure of the sealing element 8th in the the 8th to 16 . The thigh 82 widens towards its middle, with an additional recess in the middle area 85 is provided. This serves to keep cooling air holes in the liner segment free. The thigh 81 has slots at its ends 814 for axial fixing of the sealing element 8th on. The leg points further 81 in the central area a vertically protruding section pointing radially downwards 83 on.

In the 19th is the first leg 81 compared to the embodiment of 18th narrower and without the vertically protruding area 83 educated. A recess 86 in the middle of the thigh 82 interrupts the thigh 82 not completely, but is designed as an opening that is radially outward by a web 821 is limited. In the 20th becomes the recess 86 through a plurality of holes 87 in the thigh 82 replaced, which form a kind of perforation. In the 21st is the thigh 82 Formed over the entire surface, without recesses or perforations.

The 22nd shows an embodiment in which the L-shaped sealing element 8th Part of a triangular beam 800 is. The triangular beam 800 includes in addition to the L-shaped sealing element 8th that is the base of the triangular beam 800 forms, two connecting webs 810 , 820 extending from the two ends of the first leg 81 of the L-shaped sealing element 8th extend obliquely axially forward and thereby at a connection point 830 are connected to one another by a fastening element 835 trains.

The 23 shows the arrangement of such a triangular beam 800 in the outer platform 41 . It is provided that the fastening element 835 of the triangular beam 800 about a rivet 45 or another fastening means with axial play on a fastening projection 44 the outer platform 41 is connected, which is also used to attach sealing plates, compare 2 . This configuration enables a particularly stable axial fixing of the L-shaped sealing element 8th . The design of slots 814 in the thigh 81 for axially fixing the L-shaped sealing element 8th according to the 12 , 14th , 16 and 18th to 22nd is still possible, but no longer necessary.

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. Furthermore, any of the features can be used separately or in combination with any other features, provided that they are not mutually exclusive, and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein. If areas are defined, these include all values within these areas as well as all sub-areas that fall into one area.

Claims (20)

Device for sealing a gap (57) between two components (4, 6) of a turbine of a gas turbine engine, the device comprising: - a turbine guide vane ring (400) which comprises a plurality of turbine guide vane segments (4), each turbine - Guide vane segment (4) comprises an outer platform (41) and at least one guide vane (43), wherein the outer platform (41) has a radially extending first wall (5) which at its radially outer end with a straight in the circumferential direction Sealing edge (51) is provided, - a component (6) adjoining the outer platform (41) of the turbine guide vane segment (4) downstream, which has a radially extending second wall (7) which in the axial direction directly adjoins the first wall (6) adjoins the outer platform (41) and forms a flat surface (70) towards the first wall (5), - the gap (57) to be sealed between the radially extending first W and (5) of the outer platform (41) and the radially extending second wall (7) of the adjoining component (6), characterized by a longitudinally extending sealing element (8) which is L-shaped in cross section, the two each extending in a straight line and first and second legs (81, 82) arranged perpendicular to one another, the sealing element (8) with its first leg (81) with axial play and radial play with the outer platform (41) or with its second leg (82) with axial play and Radial play is connected to the adjoining component (6), the first leg (81) resting on the sealing edge (51) of the first wall (5), which extends in a straight line in the circumferential direction, and the second leg (82) resting on it to seal the gap (57) the flat surface (70) of the second wall (7) rests. Device according to Claim 1 , characterized in that the L-shaped sealing element (8) is connected along its length at several fastening points with axial play and radial play with the outer platform (42) or with the adjacent component (6). Device according to Claim 1 or 2 , characterized in that the L-shaped sealing element (8) is elastic in such a way that elastic deformations of the first wall (5) and / or the second wall (7), which can occur during operation of the gas turbine engine, are compensated. Device according to one of the preceding claims, characterized in that the L-shaped sealing element (8) is fastened to the first wall (5) by means of at least two fastening elements (90). Device according to Claim 4 , characterized in that the fastening elements (90) are designed as metal tabs, each of which is fastened to the first wall (5) by an elongated section (91) and, starting from the elongated section (91), forms a substantially closed loop (92) form which passes through a recess (83) in the first leg (81) with axial play and radial play, the recess (83) being formed in an area (812) of the leg (81) which is axially forward opposite the sealing edge (51) stands out. Device according to one of the preceding claims, characterized in that the L-shaped sealing element (8) is connected to the second wall (7) via at least two fastening elements (95). Device according to Claim 6 , characterized in that the fastening elements (95) each reach through a recess (84) formed in the second leg (82) with axial play and radial play and are connected to the second wall (7). Device according to one of the preceding claims, characterized in that the outer platforms (41) of the turbine guide vane segments (4) each form an axially extending groove (411) on their end faces and into the grooves (411) of two adjacent platforms (41) a sealing strip (90) is used, which serves to seal two adjacent outer platforms (41), the sealing strip (90) forming a fastening section (91) in the downstream direction which protrudes from the grooves (411), the fastening section ( 91) forms a holding element (92), via which the first leg (81) is fixed in the radial direction on the first wall (5). Device according to Claim 8 , characterized in that the holding element (92) is resilient and exerts a spring force on the first leg (81), which presses it from above onto the sealing edge (51) of the first wall (5) while providing radial play. Device according to Claim 8 or 9 , characterized in that the fastening section (91) of the sealing strip (90) forms two holding elements (92), via which the L-shaped sealing elements (8) of two adjacent outer platforms (41) each at one of their ends on the first wall ( 5) are attached. Device according to Claim 10 , characterized in that the fastening portion (91) is roof-shaped. Device according to one of the Claims 8 to 11 , characterized in that the holding elements (92) are essentially rectangular. Device according to one of the Claims 8 to 12 , characterized in that the first legs (81) each form a widened area (810) at their ends on which the respective holding element (92) rests. Device according to one of the Claims 8 to 13 , characterized in that the first legs (81) each have a slot (814) at their ends through which the fastening section (91) runs, the slot (814) having a greater thickness than the fastening section (to provide axial play) ( 91). Device according to one of the preceding claims, characterized in that the L-shaped sealing element (8) is part of a triangular support (800), the L-shaped sealing element (8) forming the base of the triangular support (800) and the two other sides of the Triangular support (800) are formed by two connecting webs (810, 820) which extend from the two ends of the L-shaped sealing element (8) obliquely axially forward and which at their connection point (830) with an axial play on a fastening element ( 44) of the outer platform (41) are attached. Device according to Claim 15 , as far as referenced Claim 8 , characterized in that the L-shaped sealing element (8) in each case over the triangular support (800) with an axial play on the fastening element (44) of the outer platform (41) and over holding elements (92) of the fastening sections (91) of two sealing strips ( 90) is attached to the first wall (5) with radial play. Device according to one of the preceding claims, characterized in that the L-shaped sealing element (8) is designed as a sheet metal. Device according to one of the preceding claims, characterized in that at least one of the legs (81, 82) of the L-shaped sealing element (8) has one or more recesses (85-87). Device according to one of the preceding claims, characterized in that the turbine guide vane ring (400), the turbine guide vane ring of the first turbine stage of the high-pressure turbine (17) arranged behind the combustion chamber (16) of the gas turbine engine and that the adjacent component (6) is connected to the The outer casing (33) of the high-pressure turbine (17) is connected to the liner segment, which delimits the flow path (24) through the high-pressure turbine (17) in the area of the rotor blade ring (310) of the first turbine stage radially on the outside. Gas turbine engine with a device according to Claim 1 .

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