EP3591237A1 - Structural module for a compressor of a turbomachine - Google Patents

Abstract

The invention relates to a structural assembly for a compressor of a turbomachine, which has an impeller (6) with a plurality of blades (60), which extend radially in a flow path (25) of the turbomachine, and a compressor housing (4), which has a flow path boundary (410) forms, which delimits the flow path (25) through the turbomachine radially on the outside. The compressor housing (4) has a housing structure (8) adjacent to the impeller (6). The compressor housing (4) has a plurality of peripheral segments (41, 42) which extend in the peripheral direction, only one or only some of the peripheral segments (41) forming a housing structure (8).

Description

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

It is generally 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 structuring, which are assigned to the two main groups in a circumferentially symmetrical manner (for example in the form of circumferential grooves) or discretely in circumference (for example in the form of axial grooves). The aim of structuring the housing is to ensure the stable, i.e. to expand the stall-free or pump-free working range of the compressor.

It is known, for example, that the blades of compressors in an engine experience non-symmetrical vibrations. One phenomenon that occurs is known as rotating stall (also known as stall flutter). With the rotating detachment, unstable local cells form in the blade tips of the rotor blades, in which the flow breaks off locally. These cells can rotate in the circumferential direction against the direction of rotation Paddle wheel hiking. The rotating detachment disadvantageously excites the individual blades to oscillations or vibrations, as a result of which the service life of the blades is reduced. Blade failure due to resonance is also possible if the periodic excitations are in the region 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 invention is based on the object of providing a structural assembly for a compressor of a turbomachine which effectively realizes a housing structuring.

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

The invention then considers a structural assembly for a compressor of a turbomachine that has an impeller with a plurality of blades that 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. Housing structuring structures the flow path boundary, i. H. the wall area of the compressor housing, which delimits the flow path, it being possible in principle to use any housing structures known in the prior art.

According to the present invention it is provided that the compressor housing has a plurality of circumferential segments which extend in the circumferential direction. Only one or only some of the peripheral segments form a housing structure, while the other peripheral segments are designed without a housing structure. A circumferential segment is a segment of the compressor housing which extends in the circumferential direction and which adjoins the flow path. The individual circumferential segments adjoin one another in the circumferential direction. At least one scope segment, but not all Circumferential segments form a housing structure, so that the compressor housing has variations with respect to the housing structure.

The invention effectively counteracts rotating detachment. The variation of the housing structuring in the circumferential direction provided by the invention in such a way that only one or only some of the circumferential segments have a housing structuring has the effect that the coherence of the rotating demolition 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. This significantly increases the stable working range of the compressor and the turbomachine overall.

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

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

It is pointed out that the statement that the compressor housing has an upper housing half and a lower housing half is merely a geometric statement that does not say anything about the structure of the compressor housing of the housing half. The upper half of the housing is the upper region of the compressor housing and the lower half of the housing is the lower region of the compressor housing. It can only be imaginary areas, 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 in the gravitational field of the earth, which automatically results in a vertical direction is defined. 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 peripheral segment has a housing structure, the peripheral segments with the housing structure can have different housing structures, and of course an identical housing structure can also be provided. Exemplary embodiments provide that a plurality of adjoining circumferential segments are formed with different housing structuring, for example in the upper housing half, or that a plurality of circumferential segments with housing structuring are separated by regions without housing structuring.

One embodiment of the invention provides that the peripheral segments have the same extension angle in the peripheral direction. However, this is not 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 extent angle.

The sequence of the circumferential segments can furthermore 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 during a rotation, i.e. leads to the same overall structuring.

One embodiment of the invention provides that the compressor housing has exactly two peripheral segments, an upper peripheral segment being formed in the upper housing half and a lower peripheral segment being formed in the lower housing half. The housing structure is formed exclusively in the upper peripheral segment. An embodiment variant 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 housing structure, 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 at least approximately designed 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 housing is also drawn as a "split casing". In the case of the compressor housing being formed by a two-part housing, it is particularly advantageous that the compressor housing has two circumferential segments with a circumferential angle of 180 ° in each case, the division plane between the two parts of the compressor housing simultaneously the boundary between the two circumferential segments with and without Representing housing structuring. However, it can also be provided that two circumferential segments with a circumferential angle of 180 ° are formed in the case of compressor housings that are not divided into two.

A further embodiment of the invention provides that at least one peripheral segment of the compressor housing has an inlet lining. The inlet lining forms the boundary of the flow path of the compressor housing. In the sense of the present invention, the inlet lining is part of the compressor housing. The use of a run-in pad, also known as a rub-on pad 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 an inlet lining, the housing structure being formed in the inlet lining. This is associated with the advantage that the casing structuring cannot take place on the actual compressor casing (ie in the metallic casing wall of the compressor casing), but on the inlet lining. In this way, a housing structure can be provided in a simple manner.

It can be provided that the upper circumferential segment has an inlet lining which forms a housing structure and the lower circumferential segment forms an inlet surface without a housing structure. This will along the The entire circumference of the compressor housing has narrow running gaps between the blade tips and the flow path limitation.

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

In the event that only one or only some of the peripheral segments have an inlet lining, it can be provided that the peripheral segments have two different housing radii (the housing radius referring to the housing wall of the actual compressor housing and not to the radius of the inlet lining). The radius of the housing of a peripheral segment that has an inlet lining is larger than the housing radius of a peripheral segment that has no inlet lining. By means of circumferential segments with different radii, it is possible to achieve a uniform radius of the radially outer flow path limitation despite the fact that an inlet lining is only formed in partial segments and not over 360 ° in the compressor housing.

In one configuration, the housing structuring is designed to be circumferentially discrete and, in this configuration, has circumferential grooves, for example, which each extend in the circumferential direction, the circumferential grooves being spaced apart in the axial direction.

According to a further embodiment, the housing structuring 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. One embodiment variant provides that the housing structure is designed in the form of half-heart axial grooves. A case structure in the form of half-heart axial grooves is, for example, from the DE 10 2007 056 953 A1 known, to which extent reference is made.

In further variants, the circumferential grooves or axial grooves with a rectangular or parallelogram cross section are provided. 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 at the flow path boundary, namely one removal opening with another upstream feed opening. Such circulation channels are for example from the DE 10 2008 037 154 A1 known, to which extent reference is made.

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 input stage) or an embedded stage of the compressor.

It can be provided that the impeller is designed in a BLISK design. In the case of BLISK-type wheels, problems arise due to a rotating detachment, which are counteracted by the present invention.

One embodiment provides that the impeller is a BLISK-designed fan.

Another exemplary embodiment provides that the impeller is a BLISK-type impeller of a compressor input stage of a compressor. Such a compressor input stage further comprises a stator with stator blades adjustable in staggering angle, which is arranged in front of the first rotor of the compressor. Such a stator is referred to as an inlet guide vane or a guide vane or as an IGV (IGV - Inlet Guide Vane). Inlet idlers increase the swirl in the flow and improve the working range of a compressor. In connection with the provision of a circumferential structure only in circumferential segments, the work 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 front edge of the blades in the compressor housing. It extends in a region that begins in front of the front edge of the blades in relation to the axial direction and ends behind the front edge of the blades.

• ein in BLISK-Bauweise ausgebildetes Laufrad mit einer Mehrzahl von Schaufeln, die sich in einem Strömungspfad der Strömungsmaschine radial erstrecken,
• ein Verdichtergehäuse, das eine Strömungspfadberandung ausbildet, die den Strömungspfad durch die Strömungsmaschine radial außen begrenzt, wobei
• das Verdichtergehäuse eine obere Gehäusehälfte und eine untere Gehäusehälfte aufweist, die sich jeweils in Umfangsrichtung über einen Umfangsbereich von 180° erstrecken,
• das Verdichtergehäuse angrenzend an das Laufrad eine Gehäusestrukturierung aufweist,
• das Verdichtergehäuse zwei Umfangssegmente aufweist, die sich jeweils über einen Umfangswinkel von 180° erstrecken, wobei ein oberes Umfangssegment in der oberen Gehäusehälfte und ein unteres Umfangssegment in der unteren Gehäusehälfte ausgebildet ist,
• wobei die Gehäusestrukturierung ausschließlich in dem oberen Umfangssegment ausgebildet ist.

In a further aspect of the invention, the invention provides a structural assembly for a compressor of a turbomachine, which has:

• a BLISK-type impeller with a plurality of blades that extend radially in a flow path of the turbomachine,
• a compressor housing that forms a flow path boundary that delimits the flow path through the turbomachine radially on the outside, wherein
• the compressor housing has an upper housing half and a lower housing half, each of which extends in the circumferential direction over a circumferential region of 180 °,
• the compressor housing has a housing structure adjacent to the impeller,
• the compressor housing has two circumferential segments, each of which extends over a circumferential angle of 180 °, an upper circumferential segment being formed in the upper half of the housing and a lower circumferential segment in the lower half of the housing,
• wherein the housing structure is formed exclusively in the upper peripheral segment.

• 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.

In a further aspect of the invention, the invention 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 which is designed as a hollow shaft and connects the turbine to the compressor;
• a fan positioned upstream of the engine core, the fan comprising a plurality of fan blades; and
• a transmission that receives an input from the turbine shaft and drives the fan to drive the fan at a lower speed than the turbine shaft.

• 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.

An embodiment of 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 includes 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 is pointed out that the present invention is described with reference to a cylindrical coordinate system which has the coordinates x, r and ϕ. 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 from, the radial direction points radially outwards. Terms such as "in front", "behind", "front" and "rear" refer to the axial direction or the direction of flow in the engine. Terms such as "outer" or "inner" refer to the radial direction.

As stated 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 by a transmission. Accordingly, the gas turbine engine may include a transmission that receives an input from the core shaft and drives 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 compressor so that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine described and / or claimed herein can have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting turbines and compressors, for example one, two or three shafts. For example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may 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 may be positioned axially downstream of the first compressor. The second compressor can be arranged To receive flow from the first compressor (for example, to take up directly, for example through a generally annular channel).

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

In a gas turbine engine described and / or claimed herein, a combustion chamber may be provided axially downstream of the blower and the compressor (s). For example, the combustion chamber can be located directly downstream of the second compressor (for example at the outlet thereof) if a second compressor is provided. As another example, the flow at the outlet of the compressor can be supplied 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) may include any number of stages, for example multiple stages. Each stage can include a series of rotor blades and a series 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 that extends from a foot (or hub) at a radially inner gas-flowed location or at a 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade on the hub to the radius of the fan blade on 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 limited by two of the values in the previous sentence (i.e. the values can form upper or lower limits). These ratios can be commonly 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 front edge portion (or the axially most forward edge) of the blade. The hub-to-tip ratio, of course, refers to the portion of the fan blade over which gas flows, i.e. H. the section that is radially outside of any platform.

The radius of the fan can be measured between the center line of the engine and the tip of the fan blade at its front edge. The diameter of the blower (which can simply be twice the radius of the blower) 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, About 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 limited by two of the values in the previous sentence (i.e. the values can be upper or lower limits).

The speed of the fan can vary in use. In general, the speed is lower for fans with a larger diameter. For example only, as a non-limiting example, the fan speed may be less than 2500 rpm, for example less than 2300 rpm, under constant speed conditions. Just as another, 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. Just as another 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 can be 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 using the gas turbine engine, the blower (with associated blower 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 performed by the fan blades on the flow results in an increase in the enthalpy dH of the flow. A blower _tip load can be defined as dH / U _tip ² , where dH is the enthalpy increase (e.g. the average 1-D enthalpy increase) across the blower and U _{tip is} the (translation) speed of the blower _tip, e.g. at the front edge of the tip , (which can be defined as the blower tip radius at the front edge multiplied by the angular velocity). The blower peak load under 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 are (lie) (all units in this section being Jkg ^-1 K ^-1 / (ms ^-1 ) ² ). The blower peak load can be in an inclusive range limited by two of the values in the previous 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 channel to the mass flow rate of flow through the core at constant speed 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 limited by two of the values in the previous sentence (ie the values can form upper or lower limits). The bypass channel can be essentially ring-shaped. The bypass channel can be located radially outside of the engine core. The radially outer surface of the bypass duct can be defined by an engine nacelle and / or a blower 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 output of the supercharger (prior to entering the combustion chamber). As a non-limiting example, the total pressure ratio of a gas turbine engine described and / or claimed herein at constant speed may be more than (or on the order of): 35, 40, 45, 50, 55, 60, 65, 70, 75 (lie). The total pressure ratio can be in an inclusive range limited by two of the values in the previous sentence (i.e. the values can be upper or lower limits).

The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow through the engine. Under constant speed conditions, the specific thrust of an engine described and / or claimed herein may 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. The specific thrust can be in an inclusive range limited by two of the values in the previous sentence (ie the values can be upper or lower limits). Such engines can be particularly efficient compared to conventional gas turbine engines.

A gas turbine engine described and / or claimed herein can have any desired maximum thrust. By way of non-limiting example only, a gas turbine described and / or claimed herein can produce 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 limited by two of the values in the previous sentence (i.e. the values can be upper or lower limits). The thrust referred to above can 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) for static engines.

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, are measured. At constant speed, the TET can be at least (or in the order of magnitude): 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The constant velocity TET can be in an inclusive range limited by two of the values in the previous sentence (ie the values can be upper or lower limits). For example, the maximum TET in use of the engine may 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 limited by two of the values in the previous sentence (ie the values can be upper or lower limits). The maximum TET can occur, for example, in a condition of high thrust, for example an MTO condition (MTO - maximum take-off thrust - maximum start 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 part 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, for example. B. carbon fiber. As another example, at least a portion of the fan blade and / or the blade may be at least partially made of a metal, such as. B. a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two areas made using different materials. For example, the fan blade may have a protective front edge that is made using a material that is more resistant to impact (e.g., birds, ice, or other material) than the rest of the blade. Such a leading edge can be produced, for example, using titanium or a titanium-based alloy. Thus, as an example only, the fan blade may include a carbon fiber or aluminum based body (such as an aluminum-lithium alloy) with a titanium leading edge.

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 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). Such a fixing device in the form of a dovetail, which can be inserted and / or brought into engagement with a corresponding slot in the hub / disc for fixing the fan blade, can be present only as an example. As another example, the fan blades can be integrally formed with a central portion. Such an arrangement can be referred to as a blisk or a bling. Any suitable method can be used to make such a blisk or bling. For example, at least some of the fan blades can be machined out of a block and / or at least some of the fan blades can be welded, e.g. B. linear friction welding, attached to the hub / disc.

The gas turbine engines described and / or claimed herein may or may not be provided with a VAN (Variable Area Nozzle - nozzle with a variable cross-section). Such a variable cross-section nozzle can allow the output 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 blower of a gas turbine, which is described and / or claimed here, can have any desired number of blower blades, for example 16, 18, 20 or 22 blower blades.

As used herein, constant speed conditions may mean constant speed conditions of an aircraft to which the gas turbine engine is attached. 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 is exposed between (in terms of time and / or distance) the end of the climb and the start of the descent. become.

For example only, the forward speed in the constant speed condition at any point may range from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 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. For some aircraft, constant speed conditions may be outside of these ranges, for example below Mach 0.7 or above Mach 0.9.

For example only, the constant speed conditions may be 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 (approximately 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 (approximately 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 speed conditions can correspond to standard atmospheric conditions at any given altitude in these areas.

As an 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" can mean the aerodynamic design point. Such an aerodynamic design point (or ADP - Aerodynamic Design Point) can correspond to the conditions (including, for example, the Mach number, ambient conditions and thrust requirement) for which the fan operation is designed. This can mean, for example, the conditions in which the blower (or the gas turbine engine) has the optimal efficiency by design.

In use, a gas turbine engine described and / or claimed herein may be operated at the constant speed conditions defined elsewhere herein. Such constant speed conditions may differ from the constant speed conditions (e.g. conditions during the middle part of the flight) of an aircraft on which at least one (For example, 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 one of the above aspects can be applied to any other aspect, unless they are 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, if they are not mutually exclusive.

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 using several exemplary embodiments. Show it:

Figure 1

a sectional side view of a gas turbine engine;

Figure 2

a side sectional large 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 Figure 4 ; and

Figure 6

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

Figure 1 represents a gas turbine engine 10 with a main axis of rotation 9. The engine 10 comprises an air inlet 12 and a thrust blower or fan 23, which generates two air streams: a core air stream A and a bypass air stream B. The gas turbine engine 10 comprises a core 11, which is the core air stream A. receives. The engine core 11 comprises, in axial flow order, 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 thruster nozzle 18. The bypass air flow B flows through the bypass duct 22. The fan 23 is connected to the low-pressure turbine 19 via a shaft 26 and an epicycloid gear 30 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 discharged from the high pressure compressor 15 is conducted into the combustion device 16, where it is mixed with fuel and the mixture is burned. The resulting hot combustion products then propagate through and drive the high pressure and low pressure turbines 17, 19 before being expelled through the nozzle 20 to provide some thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 through a suitable connecting shaft 27. The blower 23 generally provides most of the thrust. The epicycloid gear 30 is a reduction gear.

An exemplary arrangement for a gear blower 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 epicycloid gear arrangement 30. A plurality of planet gears 32, which are coupled to one another by a planet carrier 34, are located radially on the outside of the sun gear 28 and mesh with it. The planet carrier 34 limits the planet gears 32 to orbit synchronously around the sun gear 28 while allowing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the blower 23 via linkages 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 linkages 40, is located radially outside of the planet gears 32 and meshes therewith.

It is noted that the terms "low pressure turbine" and "low pressure compressor" as used herein can be understood to mean the lowest pressure turbine stage and the lowest pressure compressor stage (ie, they are not the blower 23) and / or the turbine and compressor stage connected by the lowest speed link shaft 26 in the engine (ie, not including the transmission output shaft that drives fan 23). In some documents, the "low pressure turbine" and the "low pressure compressor" referred to herein will alternatively be known as the "medium pressure turbine" and "medium pressure compressor". When using such alternative nomenclature, the blower 23 can be referred to as a first compression stage or compression stage with the lowest pressure.

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

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

It is understood that the in Figure 2 and 3 The arrangement shown is merely exemplary and various alternatives are within the scope of the present disclosure. Any suitable arrangement for positioning the gear 30 in the engine 10 and / or for connecting the gear 30 to the engine 10 can be used merely by way of example. As another example, the connections (e.g., linkages 36, 40 in the example of Figure 2 ) between the transmission 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) have some degree of rigidity or flexibility. As another example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (e.g., between the input and output shafts of the transmission and the defined 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 the person skilled in the art that the arrangement of the output and support rods and bearing positions in a star arrangement (described above) of the transmission 30 generally differs from that which is exemplified in FIG Figure 2 would be shown.

Accordingly, the present disclosure extends to a gas turbine engine with an arbitrary arrangement of the transmission types (for example star-shaped or planet-like), support structures, input and output shaft arrangement and bearing positions.

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

Other gas turbine engines to which the present disclosure may apply 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, in Figure 1 Gas turbine engine shown a split stream 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 outside thereof. However, this is not limitative, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass passage 22 and the flow through the core 11 are in front of (or upstream) a single nozzle, which may be referred to as a mixed flow nozzle, can be mixed or combined. One or both nozzles (whether mixed or dividing stream) can have a fixed or variable range. For example, although the example described relates to a turbofan engine, the disclosure may be applicable to any type of gas turbine engine, such as a. B. in an open rotor (in which the blower stage is not surrounded by an engine nacelle) or a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a transmission 30.

The geometry of the gas turbine engine 10 and components thereof is 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 direction from the bottom to the bottom) 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 each other.

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, in a medium-pressure compressor (if present) and / or in a high-pressure compressor.

The Figure 4 shows a sectional view of a structural assembly that defines a flow path 25 and includes an inlet stator 5, a rotor 6 and flow path boundaries. The flow path 25 conducts 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 in the area of the first stage of a compressor. The compressor comprises the inlet guide wheel 5, which has stator blades 50 that can be adjusted in a staggered angle. The swirl in the flow is increased by the inlet guide wheel 5, and the subsequent rotor 6 is thereby striven for 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 blades 60 have a front edge 61, a rear edge 62 and a blade tip 63. A gap is realized between the blade tip 63 and the compressor housing 4.

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

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

A further 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 an inlet 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, it is indicated terminologically 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 structures.

Like in the Figure 5 The compressor housing comprises an upper peripheral segment 41 and a lower peripheral segment 42. The peripheral segments 41, 42 adjoin the flow path 25. In the exemplary embodiment shown, the upper peripheral segment 41 extends, but not necessarily, in the region of the upper compressor housing 410. The lower peripheral segment 42 extends in the region 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 thus provided that only the upper peripheral segment 41 has a housing structure 8, while the lower peripheral segment 42 has no housing structure.

As for the Figure 4 explained, the housing structure in the illustrated embodiment is, but not necessarily in the inlet lining. Accordingly, the housing 4 has an upper inlet lining 71, in which a housing structure 8 is realized, and a lower inlet lining 72, which is realized without a housing structure. The two inlet linings 71, 72 each extend over 180 ° in the circumferential direction and adjoin one another in the horizontal plane 100. It can be provided that the housing 4 is a split housing which has two housing halves 411, 421, the division 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 inlet linings 71, 72.

The upper inlet lining 71, in which a housing structure 8 is implemented, and the lower inlet lining 72, which is realized without a housing 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 601 NS 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 nickel chrome aluminum bentonite mixture.

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

The radial thickness of the upper inlet lining 71 and the radial thickness of the lower inlet lining 72 are in any case identical in exemplary embodiments of the invention.

According to the terminology used, the inlet 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 shown in FIG Figure 5 formed by the inlet linings 71, 72 or in any case comprise these inlet linings 71, 72. Alternatively, it can be provided that the peripheral 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 housing structuring, the corresponding structures, such as axial grooves, are formed directly in the housing wall.

In the Figure 5 the axial grooves 81 are shown in half-heart shape in a sectional view. As explained, the axial course of these axial grooves 81 is in the half-heart shape 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 embodiment of the Figures 4 and 5 an inlet covering 72 is provided only in the lower circumferential segment 72. In this case, the casing structuring 8 is not realized in an upper inlet covering, but in an upper circumferential segment 41, which is through the casing wall of the compressor casing is realized. So that the flow path 25 does not have any cross-sectional change in such a case, it can be 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 region in which it receives the lower inlet lining 72.

In a further modification of the embodiment of the Figures 4 and 5 neither the upper housing half 410 nor the lower housing half 420 has an inlet lining. In the upper half of the housing, a peripheral segment with a structuring of the housing is formed. A peripheral segment without housing structuring is formed in the lower housing half.

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 front edge 61, a rear 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 the Figure 6 a housing structure 8 is formed directly in the housing wall of the compressor housing 4 without the use of an inlet lining. As in the exemplary embodiment of FIG Figures 4 and 5 formed by half-heart axial grooves 81. Such half-heart axial grooves are in the DE 10 2007 056 953 A1 described, to which extent reference is made.

The axial groove 81 with a half-heart shape can then be further defined by two angles α, β. These span the start and end of the cross-section curve specified by half the heart shape. The angle legs are arranged tangentially to the curve at the beginning or 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 furthermore shows that, in exemplary embodiments of the invention, the housing structure 8 in the region of the front edge 61 of the rotor blades 60 of the rotor 6 is formed. The axial extent of the axial grooves 81 is selected such that the axial grooves 81 extend from the front edge 630 of the blade tip 63 by a certain extension length opposite to the axial direction and from the front 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 structuring designed in a different way.

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

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

Claims (15)

Structural assembly for a compressor of a turbomachine, which has: - An impeller (6) with a plurality of blades (60) which extend radially in a flow path (25) of the turbomachine, and - A compressor housing (4) which forms a flow path boundary (410) which delimits the flow path (25) through the turbomachine radially on the outside, wherein - The compressor housing (4) has a housing structure (8) adjacent to the impeller (6), characterized,
that the compressor housing (4) has a plurality of peripheral segments (41, 42) which extend in the peripheral direction, only one or only some of the peripheral segments (41) forming a housing structure (8). Structural assembly according to claim 1, characterized in that the compressor housing (4) has an upper housing half (410) and a lower housing half (420), which each extend in the circumferential direction over a circumferential area of 180 °, and that the circumferential segment (41) or the peripheral segments, which form a housing structure (8), are formed in the upper housing half (410) of the compressor housing (4). Structural assembly according to claim 2, characterized in that the compressor housing (4) has exactly two peripheral segments (41, 42), an upper peripheral segment (41) being formed in the upper housing half and a lower peripheral segment (42) in the lower housing half, and wherein the housing structure (8) is formed in the upper peripheral segment (41). Structural assembly according to one of the preceding claims, characterized in that the circumferential segments (41, 42) have the same extension angle in the circumferential direction. Structural assembly according to one of the preceding claims, characterized in that the sequence of the circumferential segments (41, 42) is circumferentially asymmetrical in the circumferential direction. Structural assembly according to one of the preceding claims, characterized in that at least one peripheral segment (41, 42) of the compressor housing (4) has an inlet lining (7, 71, 72). Structural assembly according to one of the preceding claims, characterized in that at least one peripheral segment (41) which has a housing structure (8) has an inlet lining (71), the housing structure (8) being formed in the inlet lining (71). Structural assembly according to claim 7, as far as referred back to claim 3, characterized in that the upper peripheral segment (41) has an inlet lining (71) which forms a housing structure (8), and the lower peripheral segment (42) an inlet surface (72) without housing structure (8) trains. Structural assembly according to one of the preceding claims, characterized in that in the event that only one or only some of the peripheral segments (41) have an inlet lining (71), the peripheral segments (41, 42) have two different housing radii, the housing radius of a peripheral segment , which has an inlet lining, is larger than the housing radius of a peripheral segment that has no inlet lining. Structural assembly according to claim 9, insofar as it refers back to claim 3, characterized in that only the upper peripheral segment (41) or only the lower peripheral segment (42) has an inlet lining and the two peripheral segments (41, 42) have a different housing radius. Structural assembly according to one of the preceding claims, characterized in that the compressor housing (4) is formed by a two-part housing that has two parts (411, 421) which each extend over 180 ° in the circumferential direction. Structural assembly according to one of the preceding claims, characterized in that the housing structuring (8) is circumferentially discrete. Structural assembly according to one of the preceding claims, characterized in that the housing structuring (8) is designed in the form of half-heart axial grooves (81). Structure assembly according to one of the preceding claims, characterized in that the housing structure (8) is formed adjacent to the front edge of the blades in the compressor housing (4). Structural assembly according to one of the preceding claims, characterized in that the impeller (6) is constructed in a BLISK design.

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