EP3591237B1 - Structural module for a compressor of a turbomachine - Google Patents

Description

The invention relates to a structural assembly for a compressor of a turbo machine according to the preamble of claim 1.

It is basically known to provide fans and axial compressors of turbomachines with a housing structure, also referred to as "casing treatment". A large number of designs are known for such a housing structure, which are assigned to the two main groups circumferentially symmetrically (for example in the form of circumferential grooves) or circumferentially discrete (for example in the form of axial grooves). The aim of structuring the housing is to expand the stable, i.e. stall or pump-free working area of the compressor.

It is known that the blades of the compressors of an engine experience non-symmetrical vibrations. A phenomenon occurring in this case is known as rotating detachment ("rotating stall") (also referred to as tear-off flutter - "stall flutter"). During the rotating detachment, unstable local cells form at the blade tips of the rotor blades, in which the flow locally breaks off. These cells can in the rotating reference system in the circumferential direction opposite to the direction of rotation of the Wander the paddle wheel. The rotating separation disadvantageously excites the individual blades to oscillate or vibrate, which reduces the life of the blades. Blade failure due to resonance is also possible if the periodic excitations are in the range of the natural vibrations of the blades.

From the CN 201190695 Y it is known to provide different housing structures along the circumference of a compressor housing.

The DE 25 04 073 A1 discloses an inlet duct of a fan of a jet engine which is provided with a sound-absorbing acoustic lining. Variants of the embodiment provide that the sound-absorbing acoustic cladding is only formed in one or more peripheral segments of the inlet channel.

The GB 2,477,745 A discloses a generic structural assembly.

The invention is based on the object of providing a structural assembly for a compressor of a turbomachine which effectively realizes a structuring of the housing.

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

The invention then contemplates a structural assembly for a compressor of a turbomachine, which has an impeller with a plurality of blades which extend radially in a flow path of the turbomachine. A compressor housing is provided which forms a flow path boundary which delimits the flow path through the turbomachine radially on the outside. The compressor housing has a housing structure adjacent to the impeller. A housing structuring structures the flow path boundary, ie the wall area of the compressor housing which delimits the flow path, it being possible in principle to use any housing structuring known in the prior art.

It is further provided that the compressor housing has a plurality of circumferential segments which extend in the circumferential direction. In this case, only one or only some of the circumferential segments form a housing structure, while the other circumferential segments are formed without a housing structure. A circumferential segment is a segment of the compressor housing that extends in the circumferential direction and adjoins the flow path. The individual circumferential segments adjoin one another in the circumferential direction. At least one circumferential segment, but not all circumferential segments, form a housing structure, so that the compressor housing has variations in relation to the housing structure. It is further provided that at least one circumferential segment of the compressor housing has an inlet coating.

According to the invention it is provided that only one or only some of the circumferential segments have an inlet coating, the circumferential segments having two different housing radii (the housing radius relating to the housing wall of the actual compressor housing and not to the radius of the inlet coating). The housing radius of a circumferential segment that has a run-in coating is larger than the housing radius of a circumferential segment that does not have a run-in coating. By means of circumferential segments with different radii it can be achieved that, despite the fact that an inlet coating is only formed in partial segments and not over 360 ° in the compressor housing, a uniform radius of the radially outer flow path limitation is realized.

The invention effectively counteracts rotating separation. The variation of the housing structuring in the circumferential direction provided by the invention to the effect that only one or only some of the circumferential segments have a housing structuring has the effect that the coherence of the rotating tear-off patterns that are formed is disturbed. This counteracts a locally occurring flow separation at the respective blade tips. The invention thus suppresses the formation of coherent separation processes at the blade tip and thus of rotor vibrations. As a result, the stable working range of the compressor and the turbomachine is significantly increased overall.

One embodiment of the invention provides that the compressor housing has an upper housing half and a lower housing half, which each extend in the circumferential direction over a circumferential region of 180 °, and that the The circumferential segment or the circumferential segments that form a housing structure are formed in the upper housing half of the compressor housing.

This aspect of the invention thus provides for the casing to be structured only in one or more peripheral segments which are formed in the upper casing half of the compressor casing. This ensures safe and effective operation of the compressor into which the structural assembly is integrated. Because by avoiding a structuring of the housing in the lower housing half of the compressor housing, the risk of the structuring of the housing being blocked by ice, in which case there is an at least temporary loss of functionality, is avoided.

It is pointed out that the statement that the compressor housing has an upper housing half and a lower housing half is merely a geometrical statement that does not say anything about the structure of the compressor housing of the housing half (s). The upper housing half is the upper area of the compressor housing and the lower housing half is the lower area of the compressor housing. The areas can only be imaginary, the two areas being separated from one another by a horizontal plane. The terms "upper" and "lower" take into account that the structural assembly and the compressor housing are located in the earth's gravitational field, which automatically defines a vertical direction. Starting from a vector that points downwards according to the gravitational field and defines the angle of 0 °, the upper housing half extends in an angular range between 90 ° and 270 ° and the lower housing half in an angular range between 270 and 90 °.

It is further pointed out that, in the event that more than one circumferential segment has a housing structure, the circumferential segments with the housing structure can have different housing structures, whereby an identical housing structure can of course also be provided. Embodiments provide that several adjoining circumferential segments are formed with different housing structures, for example in the upper housing half, or that several circumferential segments separated by areas without a housing structure are formed with housing structures.

One embodiment of the invention provides that the circumferential segments have the same extension angle in the circumferential direction. However, this is not the case necessarily the case. In embodiments of the invention it can be provided that the extent of the circumferential segments varies in the circumferential direction, that is to say the angle of extent.

The sequence of the circumferential segments can also be circumferentially symmetrical or circumferentially asymmetrical, whereby circumferential asymmetry means that apart from the angles 0 ° and 360 ° there are no angles at which the sequence of the circumferential segments is mapped onto itself when rotated, i.e. leads to the same overall structure .

One embodiment of the invention provides that the compressor housing has exactly two circumferential segments, an upper circumferential segment being formed in the upper housing half and a lower circumferential segment being formed in the lower housing half. The structure of the housing is formed exclusively in the upper circumferential segment. A variant of this embodiment can provide that both circumferential segments each extend over a circumferential angle of 180 °. However, this is not necessarily the case. For example, it can be provided that the upper circumferential segment, which forms a structure of the housing, extends over more or less than 180 ° in the circumferential direction, while the lower circumferential segment accordingly extends over less or more than 180 ° in the circumferential direction.

If both circumferential segments each extend over a circumferential angle of 180 °, the two circumferential segments of the compressor housing are designed at least approximately as half cylinders.

According to one embodiment of the invention, the compressor housing is formed by a two-part housing which forms two parts which each extend over 180 ° in the circumferential direction. Such a compressor casing is also referred to as a "split casing". If the compressor housing is designed as a two-part housing, it is particularly advisable for the compressor housing to have two circumferential segments with a circumferential angle of 180 ° each, the dividing plane between the two parts of the compressor housing at the same time being the boundary between the two circumferential segments with and without Represents housing structuring. However, it can also be provided that two circumferential segments are formed with a circumferential angle of 180 ° each in the case of compressor housings that are not divided into two.

The invention provides that at least one peripheral segment of the compressor housing has an inlet coating. The inlet coating forms the flow path boundary of the compressor housing. In the context of the present invention, the run-in coating is part of the compressor housing. The use of a running-in coating, also referred to as an abradable coating or "liner", enables narrow running gaps between the tips of the rotor blades and the surrounding housing, as a result of which good compressor performance values can be achieved.

A large number of variants are possible. According to a variant, at least one circumferential segment that has a housing structure has a run-in coating, the housing structure being formed in the run-in coating. This is associated with the advantage that the housing structuring cannot take place on the actual compressor housing (that is, in the metallic housing wall of the compressor housing), but on the inlet coating. A housing structuring can hereby be provided in a simple manner.

In an embodiment not according to the invention, it can be provided that the upper circumferential segment has a running-in coating which forms a housing structure, and the lower circumferential segment forms a running-in coating without a casing structure. As a result, narrow running gaps between the blade tips and the flow path delimitation are achieved along the entire circumference of the compressor housing.

Another variant provides that only the circumferential segment or only the circumferential segments that do not form a housing structure have an inlet coating, while the housing structure is formed on the actual compressor housing, i.e. in the housing wall of the compressor housing.

The structure of the housing is designed circumferentially discrete in one embodiment and in this embodiment has, for example, circumferential grooves which each extend in the circumferential direction, the circumferential grooves being spaced apart in the axial direction.

According to a further embodiment, the structure of the housing is designed to be circumferentially discrete. In this embodiment, it has, for example, axial grooves which each extend over a defined length in the axial direction, the axial grooves being spaced apart in the circumferential direction. A variant of this embodiment provides that the Housing structuring is designed in the form of half-heart axial grooves. A housing structure in the form of half-heart axial grooves is for example from the DE 10 2007 056 953 A1 known.

In further variants, the circumferential grooves or axial grooves are provided with a rectangular or parallelogram-shaped cross section. It can also be provided that the housing is structured via recirculation channels instead of grooves. It is provided that a recirculation channel connects two openings to one another at the flow path boundary, namely a removal opening with a supply opening provided further upstream. Such circulation channels are for example from the DE 10 2008 037 154 A1 known.

The impeller of the structural assembly considered according to the invention can be a fan, the impeller of a low-pressure compressor, the impeller of a medium-pressure compressor or the impeller of a high-pressure compressor. It can be formed by the first stage (compressor inlet stage) or an embedded stage of the compressor.

It can be provided that the impeller has a BLISK design. In the case of impellers of the BLISK design, problems arise in particular due to a rotating detachment, which the present invention counteracts.

One embodiment provides that the impeller is a BLISK design fan.

Another exemplary embodiment provides that the impeller is an impeller of a compressor input stage of a compressor, designed in BLISK design. Such a compressor input stage further comprises a stator with stator blades adjustable in the stagger angle, which is arranged in front of the first rotor of the compressor. Such a stator is called an inlet guide vane or inlet guide vane or IGV (IGV - Inlet Guide Vane). Inlet guide vanes increase the swirl in the flow and improve the working range of a compressor. In connection with the provision of a circumferential structuring only in circumferential segments, the working area of the compressor is expanded to a particular extent.

Another embodiment of the invention provides that the housing structure is formed adjacent to the leading edge of the rotor blades in the compressor housing. It extends in an area that is based on the axial direction in front of the The leading edge of the blades begins and ends behind the leading edge of the blades.

• einen Triebwerkskern, der eine Turbine, einen Verdichter mit einer erfindungsgemäßen Strukturbaugruppe und eine die Turbine mit dem Verdichter verbindende, als Hohlwelle ausgebildete Turbinenwelle umfasst;
• einen Fan, der stromaufwärts des Triebwerkskerns positioniert ist, wobei der Fan mehrere Fanschaufeln umfasst; und
• ein Getriebe, das einen Eingang von der Turbinenwelle empfängt und Antrieb für den Fan zum Antreiben des Fans mit einer niedrigeren Drehzahl als die Turbinenwelle abgibt.
  Eine Ausgestaltung hierzu kann vorsehen, dass
• die Turbine eine erste Turbine ist, der Verdichter ein erster Verdichter ist und die Turbinenwelle eine erste Turbinenwelle ist;
• der Triebwerkskern ferner eine zweite Turbine, einen zweiten Verdichter und eine zweite Turbinenwelle, die die zweite Turbine mit dem zweiten Verdichter verbindet, umfasst; und
• die zweite Turbine, der zweite Verdichter und die zweite Turbinenwelle dahingehend angeordnet sind, sich mit einer höheren Drehzahl als die erste Turbinenwelle zu drehen.

The present disclosure also relates to a gas turbine engine, in particular for an aircraft, with a structural assembly according to the invention. It can be provided that the gas turbine engine has:

• an engine core which comprises a turbine, a compressor with a structural assembly according to the invention and a turbine shaft designed as a hollow shaft and connecting the turbine to the compressor;
• a fan positioned upstream of the engine core, the fan including a plurality of fan blades; and
• a transmission that receives an input from the turbine shaft and provides drive for the fan to drive the fan at a lower speed than the turbine shaft.
  A configuration for this can provide that
• the turbine is a first turbine, the compressor is a first compressor, and the turbine shaft is a first turbine shaft;
• the engine core further comprises a second turbine, a second compressor, and a second turbine shaft connecting the second turbine to the second compressor; and
• the second turbine, the second compressor, and the second turbine shaft are arranged to rotate at a higher speed than the first turbine shaft.

It should be noted that the present invention is described with reference to a cylindrical coordinate system which has the coordinates x, r and ϕ. Here x indicates the axial direction, r the radial direction and ϕ the angle in the circumferential direction. The axial direction is identical to the machine axis of a gas turbine engine in which the structural assembly is arranged. From the x axis starting out, the radial direction points radially outwards. Terms such as "in front", "behind", "front" and "rear" relate to the axial direction or the direction of flow in the engine. Terms like "outer" or "inner" refer to the radial direction.

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

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

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

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

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

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

The or each compressor (for example the first compressor and the second compressor as described above) can comprise any number of stages, for example a plurality of stages. Each stage can include a series of rotor blades and a series of stator blades, which can be variable stator blades (in that their angle of attack can be variable). The row of rotor blades and the row of stator blades can be axially offset from one another.

The or each turbine (for example the first turbine and the second turbine as described above) can comprise any number of stages, for example multiple stages. Each stage can include a number of rotor blades and a number of stator blades. The row of rotor blades and the row of stator blades can be axially offset from one another.

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

The radius of the fan can be measured between the centerline of the engine and the tip of the fan blade at its leading edge. The diameter of the fan (which can be simply twice the radius of the fan) can be greater than (or on the order of): 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), or 390 cm (about 155 inches). The fan diameter can be in an inclusive range bounded by two of the values in the preceding sentence (i.e., the values can be upper or lower limits).

The speed of the fan can vary with use. In general, the speed is lower for fans with a larger diameter. By way of non-limiting example only, the speed of the fan under constant speed conditions may be less than 2500 rpm, for example less than 2300 rpm. Merely as a further non-limiting example, the speed of the fan under constant speed conditions for an engine with a fan diameter in the range from 250 cm to 300 cm (for example 250 cm to 280 cm) in the range from 1700 rpm to 2500 rpm, for example in the range from 1800 rpm to 2300 rpm, for example in the range from 1900 rpm to 2100 rpm. Merely as a further non-limiting example, the speed of the fan under constant speed conditions for an engine with a fan diameter in the range from 320 cm to 380 cm in the range from 1200 rpm to 2000 rpm, for example in the range from 1300 rpm min to 1800 rpm, for example in the range from 1400 rpm to 1600 rpm.

When the gas turbine engine is in use, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tip of the fan blade to move _{at a speed U tip.} The work done by the fan blades on the flow results in an increase in the enthalpy dH of the flow. _{A fan peak} load can be defined as dH / U ^{peak 2} , where dH is the enthalpy increase (e.g. the average 1-D enthalpy increase) across the fan and U _{peak is} the (translational) speed of the fan tip, 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 (where all units in this section are Jkg ^-1 K ^-1 / (ms ^-1 ) ² ). The fan peak load can be in an inclusive range bounded by two of the values in the preceding sentence (ie, the values can be upper or lower limits).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, the bypass ratio being defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at constant velocity conditions. In some arrangements, the bypass ratio can be more than (on the order of): 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5 , 16, 16.5, or 17. The bypass ratio can be in an inclusive range bounded by two of the values in the preceding sentence (ie the values can form 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 herein can be defined as the ratio of the back pressure upstream of the fan to the back pressure at the exit of the super high pressure compressor (before the entrance to the combustion chamber). As a non-limiting example, the total pressure ratio of a gas turbine engine described herein at constant speed may be 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 (i.e., the values can be upper or lower limits).

The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow through the engine. Under constant speed conditions, the specific thrust of an engine described here can be less than (or on the order 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 (lie). The specific thrust can lie in an inclusive range which is limited by two of the values in the previous sentence (ie the values can form upper or lower limits). Such engines can be particularly efficient compared to conventional gas turbine engines.

A gas turbine engine described herein can have any maximum thrust desired. As a non-limiting example, a gas turbine described here 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 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, can be measured. 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 (ie the values can form upper or lower limits). The maximum TET when the engine is in use can be at least (or on the order of): 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K, for example. The maximum TET can be in an inclusive range bounded by two of the values in the preceding sentence (ie 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 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 made using titanium or a titanium-based alloy, for example. Thus, by way of example only, the fan blade may have a carbon fiber or aluminum based body (such as an aluminum-lithium alloy) with a leading edge made of titanium.

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

The gas turbine engines described here may or may not have a VAN (Variable Area Nozzle). Such a nozzle with a variable cross-section can allow the exit cross-section of the bypass channel to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

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

As used herein, constant speed conditions may mean constant speed conditions of an aircraft on which the gas turbine engine is mounted. Such constant speed conditions can conventionally be defined as the conditions during the middle part of the flight, for example the conditions to which the aircraft and / or the engine are exposed between (in terms of time and / or distance) the end of the climb and the start of the descent. become.

By way of example only, the forward speed under the constant speed condition may be at any point in the range of Mach 0.7 to 0.9, e.g. 0.75 to 0.85, e.g. 0.76 to 0.84, e.g. 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example in the order of Mach 0.8, in the order of Mach 0.85 or in the range of 0 .8 to 0.85. Any speed within these ranges can be the constant travel condition. In some aircraft, the cruise control conditions may be outside of these ranges, for example below Mach 0.7 or above Mach 0.9.

By way of example only, the constant velocity conditions may standard atmospheric conditions at an altitude that is in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (about 38,000 feet), for example in Range from 10,500 m to 11,500 m, for example in the range from 10,600 m to 11,400 m, for example in the range from 10,700 m (about 35,000 feet) to 11,300 m, for example in the range from 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example in the order of 11,000 m, correspond. The constant velocity conditions can correspond to standard atmospheric conditions at any given altitude in these areas.

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

As used throughout, "constant speed" or "constant speed conditions" may 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 herein can be operated at the constant speed conditions defined elsewhere herein. Such constant speed conditions can vary from the constant speed conditions (for example, the conditions during the middle part of the flight) of an aircraft on which at least one (e.g. 2 or 4) gas turbine engine can be attached to provide thrust.

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

Figur 1

eine Seitenschnittansicht eines Gasturbinentriebwerks;

Figur 2

eine Seitenschnittgroßansicht eines stromaufwärtigen Abschnitts eines Gasturbinentriebwerks;

Figur 3

eine zum Teil weggeschnitte Ansicht eines Getriebes für ein Gasturbinentriebwerk;

Figur 4

schematisch eine Strukturbaugruppe, die ein Eintrittsleitrad, ein Laufrad und ein Verdichtergehäuse mit einer Gehäusestrukturierung aufweist;

Figur 5

einen Schnitt senkrecht zur axialen Richtung entlang der Linie D-D der Figur 4; und

Figur 6

ein Ausführungsbeispiel einer Axialnut, die in Form einer Halbherz-Nut ausgebildet 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:

Figure 1

a side sectional view of a gas turbine engine;

Figure 2

Fig. 3 is a side sectional close-up view of an upstream portion of a gas turbine engine;

Figure 3

a partially cut away view of a transmission for a gas turbine engine;

Figure 4

schematically, a structural assembly which has an inlet stator, an impeller and a compressor housing with a housing structure;

Figure 5

a section perpendicular to the axial direction along the line DD of the Figure 4 ; and

Figure 6

an embodiment of an axial groove which is designed in the form of a half-heart groove.

Figure 1 Figure 10 illustrates a gas turbine engine 10 with a main axis of rotation 9. The engine 10 includes an air inlet 12 and a thrust fan 23 that generates two air flows: a core air flow A and a bypass air flow B. The gas turbine engine 10 includes a core 11 which contains the core air flow A. records. The engine core 11 comprises, in axial flow sequence, a low-pressure compressor 14, a high-pressure compressor 15, a combustion device 16, a high-pressure turbine 17, a low pressure turbine 19 and a core thrust nozzle 20. An engine nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass thrust nozzle 18. The bypass air flow B flows through the bypass duct 22. The fan 23 is via a shaft 26 and an epicycloid gear 30 on the low pressure turbine 19 attached and is driven by this.

In use, the core air flow A is accelerated and compressed by the low pressure compressor 14 and passed into the high pressure compressor 15 where further compression takes place. The compressed air expelled from the high pressure compressor 15 is directed into the combustion device 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 17, 19 and thereby drive them before they are ejected through the nozzle 20 to provide a certain thrust. The high pressure turbine 17 drives the high pressure compressor 15 through a suitable connecting shaft 27. The fan 23 generally provides the majority of the thrust. The epicycloidal gear 30 is a reduction gear.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG Figure 2 shown. The low pressure turbine 19 (see Figure 1 ) drives the shaft 26, which is coupled to a sun gear 28 of the epicycloidal gear assembly 30. Several planet gears 32, which are coupled to one another by a planet carrier 34, are located radially outward from the sun gear 28 and mesh with it. Planet carrier 34 constrains planet gears 32 to orbit synchronously about sun gear 28 while allowing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the fan 23 via linkage 36 to drive its rotation about the engine axis 9. An outer gear or ring gear 38, which is coupled to a stationary support structure 24 via linkage 40, is located radially on the outside of the planet gears 32 and meshes with them.

It is noted that the terms "low-pressure turbine" and "low-pressure compressor" as used herein can be understood to mean the turbine stage with the lowest pressure and the compressor stage with the lowest pressure (ie that it is not the fan 23) and / or the turbine and compressor stages which are connected to one another by the connecting shaft 26 with the lowest speed in the engine (ie that it does not include the transmission output shaft which drives the fan 23). In some documents the "low-pressure turbine" and the "low-pressure compressor" may be referred to here will, 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 compression stage having the lowest pressure.

The epicycloidal gear 30 is shown in Figure 3 shown in more detail as an example. The sun gear 28, planet gears 32 and ring gear 38 each include teeth around their periphery for meshing with the other gears. However, for the sake of clarity, only exemplary sections of the teeth are shown in FIG Figure 3 shown. Although four planet gears 32 are shown, it will be apparent to those skilled in the art that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicycloidal gear 30 generally include at least three planet gears 32.

This in Figure 2 and 3 The epicycloidal gear 30 shown as an example is a planetary gear in which the planet carrier 34 is coupled to an output shaft via linkage 36, the ring gear 38 being fixed. However, any other suitable type of epicycloidal gear 30 can be used. As another example, the epicycloidal gear 30 may be a star configuration in which the planet carrier 34 is held in place while allowing the ring gear (or outer gear) 38 to rotate. With such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the transmission 30 may be a differential gear that allows both the ring gear 38 and the planetary carrier 34 to rotate.

It goes without saying that the in Figure 2 and 3 The arrangement shown is only exemplary and various alternatives are possible. Any suitable arrangement for positioning the transmission 30 in the engine 10 and / or for connecting the transmission 30 to the engine 10 can be used merely as an example. As another example, the links (e.g., linkages 36, 40 in the example of FIG Figure 2 ) have some degree of rigidity or flexibility between the transmission 30 and other parts of the engine 10 (such as the input shaft 26, output shaft, and fixed structure 24). As another example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (e.g., between the input and output shafts of the transmission and the fixed structures such as the transmission housing) can be used and the disclosure is not on the exemplary arrangement of Figure 2 limited. For example, it is readily apparent to a person skilled in the art that the arrangement of the output and support rods and bearing positioning in a star arrangement (described above) of the transmission 30 generally differs from those exemplified in FIG Figure 2 would differ.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gear types (e.g., star or planetary), support structures, input and output shaft arrangements, and bearing positions.

Optionally, the transmission can drive secondary and / or alternative components (e.g. the medium-pressure compressor and / or a booster).

Other gas turbine engines to which the present disclosure may find application may have alternative configurations. For example, such engines can have an alternative number of compressors and / or turbines and / or an alternative number of connecting shafts. As another example, the Figure 1 The gas turbine engine shown has a split flow nozzle 20, 22, which means that the flow through the bypass duct 22 has its own nozzle which is separate from the engine core nozzle 20 and radially outward therefrom. However, this is not limiting and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are before (or upstream) a single nozzle, which can be referred to as a mixed flow nozzle, mixed or combined. One or both nozzles (whether mixed or split flow) can have a fixed or variable range. For example, while the example described relates to a turbo fan engine, the disclosure may be applied to any type of gas turbine engine such as a gas turbine engine. B. in an open rotor (in which the fan stage is not surrounded by an engine nacelle) or a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

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 (which is aligned with the axis of rotation 9), a radial direction (in the downward direction up in Figure 1 ) and a circumferential direction (perpendicular to the view in Figure 1 ) includes. The axial, radial and circumferential directions are perpendicular to one another.

In the context of the present invention, the formation of a housing structure in the compressor housing is important. The invention can basically be implemented in the fan stage, in a low-pressure compressor, a medium-pressure compressor (if present) and / or a high-pressure compressor.

The Figure 4 FIG. 4 shows a sectional view of a structural assembly that defines a flow path 25 and comprises an inlet stator 5, a rotor 6 and flow path boundaries. The flow path 25 directs the core air flow A according to FIG Figure 1 through the core engine.

The flow path 25 is delimited radially on the inside by inner wall or hub structures 95 which form an inner flow path boundary 950. The flow path 25 is delimited radially on the outside by a compressor housing 4 that forms a radially outer flow path boundary 410. In the illustrated embodiment, but not necessarily, the structural assembly is located in the area of the first stage of a compressor. In this case, the compressor comprises the inlet stator 5, which has stator blades 50 which can be adjusted in terms of the stagger angle. The swirl in the flow is increased by the inlet guide wheel 5 and, as a result, the following rotor 6 is flown against in a more effective manner.

The rotor 6 comprises a series of rotor blades or rotor blades 60, which extend radially in the flow path 25. The rotor blades 60 have a leading edge 61, a trailing edge 62 and a blade tip 63. A gap is implemented between the blade tip 63 and the compressor housing 4.

To minimize this gap, an inlet lining 7 is integrated into the compressor housing 4. The housing wall of the compressor housing 4 facing the flow path 25 forms a corresponding recess 45 for this purpose. On its side facing the flow path 25, the inlet coating 7 forms the flow path boundary 410. The running-in lining 7 is fastened in the recess 45 on its radially outer side.

A housing structure 8 is integrated into the run-in coating 7. In the illustrated embodiment, this is implemented by axial grooves 81 in the shape of a half-heart. A such axial groove 81 in half-heart shape 81 is in the Figure 4 shown schematically. However, the illustration should not be understood to mean that the housing structure 8 would protrude from the run-in coating 7 in the radial direction. The illustration only serves to show the axial course of the axial groove 81 in half-heart shape.

A more detailed understanding of the structure of the housing structure 8 results from the Figure 5 which shows a sectional view through a compressor housing 4 and a run-in lining 7 in a plane perpendicular to the axial direction which is defined by the machine axis 9 (along the line DD of FIG Figure 4 ). The direction of rotation of the rotor 6 is marked with u.

First of all, terminologically, it is pointed out that the compressor housing 4 has an upper compressor housing 410 and a lower compressor housing 420. The upper compressor housing 410 and the lower compressor housing 420 indicate spatial areas of the compressor housing 4. These are separated from one another by the horizontal plane 100. The upper compressor housing 410 and the lower compressor housing 420 can in principle be formed by any desired structures.

Like in the Figure 5 As shown, the compressor housing comprises an upper circumferential segment 41 and a lower circumferential segment 42. The circumferential segments 41, 42 adjoin the flow path 25. In the example shown, the upper circumferential segment 41 extends, but not necessarily, in the area of the upper compressor housing 410. The lower circumferential segment 42 extends in the area of the lower compressor housing 420.

The circumferential segments 41, 42 differ in any case in that only one of the circumferential segments forms a housing structure. It is provided that only the upper peripheral segment 41 has a housing structure 8, while the lower peripheral segment 42 has no housing structure.

As in terms of the Figure 4 explained, the structure of the housing is formed in the illustrated embodiment, but not necessarily in the run-in coating. Correspondingly, the housing 4 has an upper run-in coating 71, in which a housing structuring 8 is implemented, and a lower run-in coating 72, which is implemented without a housing structuring. The two run-in linings 71, 72 extend in the example not according to the invention in FIG Figure 5 in the circumferential direction in each case over 180 ° and adjoin one another in the horizontal plane 100. It can be provided that the housing 4 is a divided housing which has two housing halves 411, 421, the dividing plane (ie the horizontal plane 100) between the two housing halves 411, 421 also representing the boundary between the two circumferential segments 41, 42 or run-in coatings 71, 72. one or only some of the circumferential segments have a running-in coating.

The upper run-in coating 71, in which a casing structure 8 is implemented, and the lower run-in coating 72, which is implemented without a casing structure, consist, according to one embodiment, of the material Metco 601NS, Metco 320NS or Metco 314NS from Oerlikon Metco Switzerland in 8808 Pfäffikon, Switzerland . Metco 601NS is a mixture of silicon-aluminum powders and polyester powders. Metco 320NS is an aluminum-silicon boron nitride powder. Metco 314NS is a thermal spray powder made from a mixture of nickel chromium aluminum bentonite.

Alternatively, high-temperature plastics, porous materials or metallic honeycomb structures are used as the material for the run-in linings 71, 72.

The radial thickness of the upper run-in coating 71 and the radial thickness of the lower run-in coating 72 are identical in any case in exemplary embodiments of the invention.

According to the terminology used, the run-in linings 71, 72 are part of the compressor housing 4. The peripheral segments 41, 42, which are only partially provided with a housing structure, are in FIG Figure 5 formed by the inlet linings 71, 72 or at least include these inlet linings 71, 72. Alternatively, it can be provided that the circumferential segments 41, 42 are formed by the actual compressor housing, ie the wall of the compressor housing facing the flow path 25. In the case of a structuring of the housing, the corresponding structures such as axial grooves are formed directly in the housing wall.

In the Figure 5 the axial grooves 81 are also shown in a sectional view in half-heart shape. As explained, the axial course of these axial grooves 81 is in the shape of a half-heart in FIG Figure 4 shown. According to the Figure 5 the axial grooves 81 are slightly inclined with respect to the radial direction. This can be done in one or the other circumferential direction. The axial grooves 81 can also run exactly in the radial direction.

In a modification of the example of the Figures 4 and 5 and according to the invention, a run-in coating 72 is only provided in the lower circumferential segment 72. In this case, the structure of the housing 8 is not implemented in an upper run-in coating, but in an upper circumferential segment 41 that extends through the housing wall of the compressor housing is realized. So that the flow path 25 does not have a change in cross section in such a case, it is provided that the housing 4 has a larger housing radius in the lower compressor housing 420 than in the upper compressor housing 410 along the axial area in which it receives the lower inlet coating 72.

In a further modification of the example of FIG Figures 4 and 5 neither the upper housing half 410 nor the lower housing half 420 has a running-in coating. A circumferential segment with a housing structure is formed in the upper half of the housing. In the lower housing half, a circumferential segment is formed without a housing structure.

The Figure 6 shows a further embodiment of a structural assembly according to the present invention. According to the description of the Figure 4 the structural assembly comprises a compressor housing 4 which forms a flow path boundary 410. A rotor 6 comprises rotor blades 60, each of which has a leading edge 61, a trailing edge 62 and a blade tip 63. It can be seen that a gap is formed between the blade tip 63 and the flow path boundary 410. The machine axis 9 is also shown.

In the embodiment of Figure 6 a housing structure 8 is formed directly in the housing wall of the compressor housing 4 without the use of an inlet coating. As in the example of the Figures 4 and 5 formed by axial grooves 81 in half-heart shape. Such half-heart shaped axial grooves are in the DE 10 2007 056 953 A1 described.

The axial groove 81 with a half-heart shape can then be further defined by two angles α, β. These span the starting area and the end area of the cross-sectional curve given by the half heart shape. The angle legs are arranged tangentially to the curve profile at the beginning or at the end. The angle α is, for example, in the range between 20 ° and 70 ° to the wall of the compressor housing and the angle β is, for example, in the range between 30 ° and 80 ° to the wall of the compressor housing.

The Figure 6 further shows that, in exemplary embodiments of the invention, the housing structure 8 in the area of the leading edge 61 of the rotor blades 60 of the rotor 6 is formed. The axial extent of the axial grooves 81 is selected in such a way that the axial grooves 81 extend from the leading edge 630 of the blade tip 63 by a certain extension length opposite to the axial direction and starting from the leading edge 630 by a certain extension length in the axial direction. The extension length in the two directions mentioned is, for example, a maximum of 50% of the axial length of the blade tip 63. Such an axial extension of the housing structuring can also be provided in the case of housing structurings designed in a different way.

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. For example, a different segmental division of the compressor housing can be provided.

Claims (12)

1. Structural subassembly for a compressor of a turbomachine, which has:

   - a rotor (6) having a plurality of blades (60), which extend radially in a flow path (25) of the turbomachine, and

   - a compressor casing (4), which forms a flow path boundary (410), which delimits the flow path (25) through the turbomachine radially on the outside, wherein

   - the compressor casing (4) has casing structuring (8) adjoining the rotor (6), which casing structuring counteracts rotating stall,

   - the compressor casing (4) has a plurality of circumferential segments (41, 42), which extend in a circumferential direction, wherein only one or only some of the circumferential segments (41) forms or form the casing structuring (8), and

   - at least one circumferential segment (41, 42) of the compressor casing (4) has an abradable lining (7, 71, 72),

   characterized
   in that only one or only some of the circumferential segments (41) has or have an abradable lining (71), wherein the circumferential segments (41, 42) have two different casing radii, wherein the casing radius of a circumferential segment which has an abradable lining (7, 71, 72) is larger than the casing radius of a circumferential segment which does not have an abradable lining.
2. Structural subassembly according to Claim 1, characterized in that the compressor casing (4) has an upper casing half (410) and a lower casing half (420), which each extend over a circumferential range of 180° in the circumferential direction, and in that the circumferential segment (41) or the circumferential segments which forms or form the casing structuring (8) is or are formed in the upper casing half (410) of the compressor casing (4).
3. Structural subassembly according to Claim 2, characterized in that the compressor casing (4) has precisely two circumferential segments (41, 42), wherein an upper circumferential segment (41) is formed in the upper casing half and a lower circumferential segment (42) is formed in the lower casing half, and wherein the casing structuring (8) is formed in the upper circumferential segment (41).
4. Structural subassembly according to any one of the preceding claims, characterized in that the circumferential segments (41, 42) have the same angle of extent in the circumferential direction.
5. Structural subassembly according to any one of the preceding claims, characterized in that the sequence of circumferential segments (41, 42) is circumferentially asymmetrical in the circumferential direction.
6. Structural subassembly according to any one of the preceding claims, characterized in that at least one circumferential segment (41) which has the casing structuring (8) has an abradable lining (71), wherein the casing structuring (8) is formed in the abradable lining (71) .
7. Structural subassembly according to Claim 3, characterized in that only the upper circumferential segment (41) or only the lower circumferential segment (42) has an abradable lining and the two circumferential segments (41, 42) have a different casing radius.
8. Structural subassembly according to any one of the preceding claims, characterized in that the compressor casing (4) is formed by a casing divided into two, which has two parts (411, 421) that each extend over 180° in the circumferential direction.
9. Structural subassembly according to any one of the preceding claims, characterized in that the casing structuring (8) is of circumferentially discrete design.
10. Structural subassembly according to any one of the preceding claims, characterized in that the casing structuring (8) is in the form of half-heart-shaped axial grooves (81).
11. Structural subassembly according to any one of the preceding claims, characterized in that the casing structuring (8) is formed in the compressor casing (4) adjoining the leading edge of the rotor blades.
12. Structural subassembly according to any one of the preceding claims, characterized in that the rotor (6) is of BLISK-type construction.

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