DE102019216635A1 - ROTATING BLADE FOR A FLOW MACHINE - Google Patents

Abstract

The present invention relates to a moving blade (10) to be arranged in a gas duct (4) of a turbo-engine (1), with a blade (20) which extends from a blade root (25) to a blade tip (26), the blade ( 20) has a front and a rear edge (21,22), as well as a suction side surface (23) and a pressure side surface (24), and wherein the airfoil (20) has a first channel (32) in its interior for a first fluid to flow through (52) which extends away from a first suction opening (32.1) provided for sucking the first fluid (52) out of the gas channel (4), the first suction opening (32.1) being arranged at the blade tip (26).

Description

Technical area

The present invention relates to a rotor blade for arrangement in a gas duct of a turbomachine.

State of the art

The turbo engine can be, for example, a jet engine, e.g. B. a turbofan engine. The flow machine is functionally divided into a compressor, combustion chamber and turbine. In the case of the jet engine, for example, the air that is sucked in is compressed by the compressor and burned in the downstream combustion chamber with the added kerosene. The resulting hot gas, a mixture of combustion gas and air, flows through the downstream turbine and is expanded in the process. The turbine also partially extracts energy from the hot gas in order to drive the compressor.

The present subject matter is directed to a rotor blade for arrangement in the gas duct of the turbomachine, in particular in the compressor gas duct. The rotor blade has a rotor blade which extends from a blade root radially inward to a blade tip radially outward. In relation to the flow around the gas channel, the airfoil has a leading and a trailing edge, and the suction and pressure side surfaces of the airfoil extend between the leading and trailing edges.

Presentation of the invention

The present invention is based on the technical problem of specifying an advantageous rotor blade.

This is achieved according to the invention with the rotor blade according to claim 1. The interior of its blade is traversed by a first channel which has a suction opening assigned to the gas duct of the turbomachine (“first suction opening”). During operation, a fluid (“first fluid”) can be sucked off through the first channel via the first suction opening, which is arranged on the tip of the Schuafel.

In the turbomachine or a module thereof, a duct wall is arranged radially outside of the airfoil which delimits the gas duct of the turbomachine radially outward. In this case, a small gap between the channel wall and the blade tip, which during operation as a result of the blade rotation, brushes along the channel wall, cannot be avoided. In this gap, during operation, a gap flow directed from the pressure side to the suction side can form, which has a flow pulse directed against the main flow in the gas channel. If the gap and main flow interact, a gap vortex and a blade tip blockage can develop.

As a result of its positioning at the blade tip, the first suction opening is assigned to the gap between the blade tip and the gas channel wall (it points radially outwards). By sucking off splitting fluid through the first suction opening, for example, the growth of the blade tip blockage can then be reduced or prevented, which creates relief in the blade tip area.

The flow stabilization is to a certain extent integrated into the rotor blade or the airfoil itself, which can enable use not only in the geometrically larger, front compressor stages, but also in the smaller compressor stages axially further back. Especially with the latter, for example, the use of structure structures for flow stabilization would already be geometrically difficult to implement, especially since filigree structure structures also generally mean increased costs and can also be critical in terms of operational safety due to the risk of contamination.

Preferred embodiments can be found in the dependent claims and the entire description, although a distinction is not always made between the different claim categories in detail when the features are presented; the disclosure is in any case to be read implicitly at the same time for device or process and usage aspects. If, for example, a rotor blade suitable for a specific operation or manufactured in a specific manner is described, this is to be read as a disclosure of a corresponding use or manufacture (and vice versa).

The information “axial”, “radial” and “circumferential”, as well as the associated directions (axial direction, etc.), relate in the context of the present disclosure to the axis of rotation about which the rotor blade rotates during operation (and which is typically with a longitudinal axis of the Module or the flow machine coincides). “Front” and “rear” refer to the flow around the airfoil in the gas duct, ie, in the case of the compressor module, to the flow of the compressor fluid, unless stated otherwise. “Vome” means upstream, “rear” means downstream.

The axial chord length is considered in a tangential plane at the corresponding radial height, in the case of the second injection and first suction openings that is, at the blade tip (at their radially outer end). In detail, the axial chord length is then taken as the axial distance between a tangent placed axially at the front and an axially rearward on the blade profile (the tangents are each perpendicular to the axial direction). The position is determined from front to back, i.e. between 0% for the front tangent and 100% for the rear tangent.

In a preferred embodiment, the first suction opening is between 20% and 100% of the axial chord length, preferably in an axially central area between 20% and 60%, particularly preferably between 20% and 50%, i.e. in an area of maximum profile curvature (due to the Area with particularly strong crevice flow, which is extracted by the suction). The first suction opening can extend over the entire axially central area or only over part of it.

According to a preferred embodiment, the interior of the airfoil is additionally traversed by a second channel which opens into a second injection opening. A fluid guided in the second channel (“second fluid”) can be blown into the gas channel of the turbomachine via the second injection opening. The second injection opening is preferably also arranged in the blade tip, particularly preferably in front of the first suction opening. With the injection through the second injection opening, in particular the formation of the split vortex can be prevented. The inventor has observed that the core of the split vortex (core vortex) lies in the front 20% of the chord length, and this core vortex can be suppressed by the injection.

In this way, the overall aerodynamic stability can be improved, in particular in the case of a compressor (axial compressor). Otherwise, with increasing throttling and thus increasing gap flow, the gap vortex could grow and split up until it hits the blade pressure side of the downstream rotor blade and migrates to its leading edge with further throttling. This can ultimately lead to a so-called “spill forward”. Conversely, the lower splitting or suppression of the gap vortex and the reduced blade tip blockage allow the (safe) operating range to be expanded, i.e. the compressor can be operated at a higher pressure ratio (advantageous in terms of efficiency and size, and thus also in terms of material costs, integrability and fuel consumption) .

The second injection opening is therefore preferably in an axially front area between 0% and 20% of the axial chord length. It can completely fill this area, that is, it can have an elongated slot shape, or it can only extend over a section of the axially front area, as shown in the exemplary embodiment. In general, the injection and suction openings discussed in the context of this disclosure can each be provided as precisely one contiguous opening or also be subdivided into several partial openings. The same applies to the channels, each of which can also be subdivided into several sub-channels or can branch accordingly.

According to a preferred embodiment, the second channel, where it opens into the second injection opening, is inclined towards the pressure side surface of the airfoil. In a section perpendicular to the airfoil chord at the blade tip, the second channel at the injection opening is tilted in the direction of the pressure side. A central axis of the channel, viewed in this section, can be inclined, for example, by at least 10 °, 20 °, 30 ° or 40 ° relative to the radial direction to the pressure side (possible upper limits can be at most 80 °, 70 ° or 60 °, for example ). The described suppression or deflection of the core vortex can be achieved particularly well if the second fluid is blown in as tangentially as possible to the surface of the blade tip, for example in the direction of the blade chord and / or in the direction of the blade pressure side. Preferably, the second channel can also be inclined axially rearward in the direction of the blade chord in the direction of the pressure side surface.

In a preferred embodiment, the first channel at the first suction opening is inclined in the direction of the pressure side, i.e., in a section perpendicular to the blade chord at the blade tip, a center axis of the first channel forms an angle with the radial direction, compare the angle information in the previous paragraph (also may be preferred in the case of the first channel). In relation to the blade thickness, the first suction opening is preferably not located in the center of the profile, but rather offset towards the pressure side. Correspondingly, in a section perpendicular to the blade chord at the blade tip, an edge proximal to the suction side surface, i.e. closer to the suction side surface, can be spaced from the suction side surface by at least 50% of the blade thickness taken in this section.

In a preferred embodiment, the first suction opening even extends partially into the pressure side surface. In a section perpendicular to the blade chord at the blade tip, a radially outer edge of the pressure side surface is offset inward, for example by up to 15% or 10% of the radial blade height (with possible lower limits at, for example, at least 3% or 5% ). These and those described in the preceding paragraphs Orientation or configuration of the first suction opening can promote the suction of fission fluid, which would otherwise fill up blockage zones.

In general, the concept described up to now could also be implemented with an active channel structure, that is to say with an active application of pressure (second channel) or negative pressure generation (first channel) outside the blade or gas channel of the turbomachine. However, the channel or channels of the blade are preferably passive, so the respective other opening is also assigned to the gas channel of the turbomachine and is part of the rotor blade, see in detail below.

According to a preferred embodiment, the second channel has a second suction opening through which the second fluid is sucked out of the gas channel of the turbomachine (and via the second channel reaches the second injection opening at the blade tip). The second suction opening can ideally be positioned in an area with the highest possible pressure; it is preferably arranged in the pressure side surface of the airfoil. It is preferably placed there in an area on the front edge, for example in the axially front 15% or 10% of the axial chord length (viewed at the radial height at which the suction opening is located).

Positioning on the pressure side at the front edge can be advantageous, for example, as the stagnation point moves towards the pressure side with increasing throttling, so that the amount of fluid sucked in there and correspondingly blown out at the second injection opening increases with increasing throttling. With increasing aerodynamic blade loading, the amount of fluid routed via the second channel then automatically increases as a result of the increasing pressure gradient, which conversely can result in a largely reduced fluid routing via the second channel at the aerodynamic design point (ADP). This allows possible negative effects on the efficiency to be limited, the channel structure represents a passive stabilization mechanism. The amount of fluid passed through the second channel and thus the intensity of the injection via the second injection opening can of course also be adjusted via the channel cross-section, which is also the case with an efficiency optimization can find input.

According to a preferred embodiment, the first channel opens into the gas channel of the turbomachine in a first injection opening, so the first fluid extracted at the first suction opening and guided through the first channel is then blown into the gas channel again at another point. The injection opening of the first channel lies relative to its suction opening at the blade tip in an area of lower pressure, preferably in the suction side surface of the airfoil.

The injection preferably takes place via the first channel as tangential as possible to the airfoil profile. The first channel can, for example, open into the suction side surface at an angle of at most 30 °, preferably at most 20 ° or 10 ° (with possible lower limits at, for example, 3 ° or 5 °). The angle between a central axis of the channel in the first injection opening and a plane placed in the suction side surface at the injection opening is considered here. A corresponding injection can, for example, also be advantageous with regard to energizing the boundary layer and thus prevent detachment areas as the throttling increases.

In the case of the first channel, too, the intensity of the injection can be set via the channel cross-section and this can be used for an efficiency optimization. The first injection opening is preferably arranged in a radially inner half of the airfoil. In relation to an airfoil height taken from radially inside to radially outside, it is accordingly between 0% and 50%, preferably between 0% and 30% of the airfoil height.

The invention also relates to a compressor module with a rotor blade described here, in particular a compressor module of an aircraft engine. The subject matter of the invention can advantageously be used both in the axially front and in the axially rear stages for stabilizing the flow of axial compressor rotors that are critical to the blade tips.

The invention also relates to a production method in which at least the airfoil is built up generatively. The blade is thus, for example, generated layer by layer using a data model, e.g. B. in a powder bed process. The channel structure can thus be implemented particularly well inside. The rotor blade as a whole can also be produced conventionally in parts, for example the blade root by casting or forging, but on the other hand the entire rotor blade can also be built up generatively.

Brief description of the drawings

In the following, the invention is explained in more detail with the aid of exemplary embodiments, with the individual features within the framework of the independent claims also being essential to the invention in other combinations and furthermore no distinction is made between the different claim categories.

• 1 in schematischer Darstellung ein Mantelstromtriebwerk in einem Axialschnitt;
• 2 ein Laufschaufelblatt mit Kanalstruktur in einer Schrägansicht;
• 3 eine strömungsmechanische Illustration zu der Ansicht gemäß 2;
• 4 eine zweite Einblasöffnung des Schaufelblatts gemäß 2 in einem Schnitt;
• 5 die Positionierung einer zweiten Absaugöffnung des Schaufelblatts gemäß 2 in einem Tangentialschnitt;
• 6 eine zweite Möglichkeit zur Ausgestaltung einer ersten Absaugöffnung des Schaufelblatts gemäß 2;
• 7 eine erste Möglichkeit zur Ausgestaltung einer ersten Absaugöffnung des Schaufelblatts gemäß 2;
• 8 eine erste Einblasöffnung des Schaufelblatts gemäß 2 in einem Tangential schnitt.

In detail shows

• 1 a schematic representation of a turbofan engine in an axial section;
• 2 a rotor blade with a channel structure in an oblique view;
• 3 a fluid mechanical illustration of the view according to FIG 2 ;
• 4th a second injection opening of the airfoil according to FIG 2 in one cut;
• 5 the positioning of a second suction opening of the airfoil according to FIG 2 in a tangential section;
• 6th a second possibility for designing a first suction opening of the airfoil according to FIG 2 ;
• 7th a first possibility for designing a first suction opening of the blade according to FIG 2 ;
• 8th a first injection opening of the airfoil according to FIG 2 cut in a tangential.

Preferred embodiment of the invention

1 shows a turbomachine 1 in section, specifically a jet engine (turbofan engine). The turbo machine 1 is functionally divided into compressors 1a , Combustion chamber 1b and turbine 1c . Both the compressor 1a as well as the turbine 1c are each made up of several modules, the compressor 1a present from a low pressure 1aa and a high pressure compressor module 1ab . Any compressor module 1aa , 1ab is for its part made up of several stages, each stage usually consists of a rotor blade ring and a subsequent guide vane ring. The compressor is in operation 1a based on a longitudinal axis 2 axially from the compressor gas 3 , in this case air, flows through, in a compressor gas channel 4th . The compressor gas 3 compressed, in the combustion chamber 1b kerosene is then mixed in and this mixture is burned.

2 shows a rotor blade 10 in an oblique view, in particular the airfoil 20th . This has a leading edge 21st and a trailing edge 22nd on, as well as a suction side surface 23 and a print face 24 . It extends radially from a blade root 25th the blade 10 up to the tip of a shovel 26th . Inside is the shovel blade 20th from a first channel 32 and a second channel 31 streaked.

The second channel 31 extends between a second suction opening 31.1 and a second injection port 31.2 , the first channel 32 between a first suction opening 32.1 and a first injection port 32.2 . The second injection port 31.2 and the first suction opening 32.1 are in the tip of the shovel 26th arranged, the second injection port 31.2 in an axially front area (between 0% and 20% of the axial chord length) and the first suction opening in an axially central area (between 20% and 50% of the axial chord length). The second channel 31 is from a second fluid 51 flows through the first channel 32 from a first fluid 52 .

The function of the arrangement becomes apparent 3 that are analogous in a view 2 shows fluid mechanical details. In general, the object is aimed at a reduction or stabilization with regard to a split fluid 35 which during operation via the bucket tip 26th flows away from the pressure to the suction side. The fission fluid 35.1 in an axially front area forms the core 36 a cleft vortex, which can split up with increasing throttling until a spill-forward occurs (compare the introduction to the description in detail). The gap flow 35.2 in the axially central area there is a blockage area 37 and has as the mainstream 38 opposite gap flow has a negative effect on aerodynamic stability. With the in 2 illustrated measures, both effects can be limited or prevented, namely the front part of the split fluid 35.1 distracted and the blockage area 37 nourishing part of the fission fluid 35.2 sucked off.

4th shows the second injection port 31.2 of the second channel 31 in a section perpendicular to the blade chord. A central axis 40 of the second channel 31 is at the second injection port 31.2 to the print face 24 inclined towards. In addition, the second channel 31 be inclined axially rearward along the airfoil chord (not shown).

5 illustrates the position of the second suction opening 31.1 , this is in an area on the leading edge 21st and at the same time in the print side area 24 arranged. A stagnation point area 50 namely migrates into the pressure side surface with increasing throttling 24 into it so that the amount of through the second channel 31 guided second fluids 51 automatically increases.

The 6th and 7th illustrate two variants of the first suction opening 32.1 . The first channel 32 is there in both cases to the print side surface 24 inclined, compare the inclination of the central axis 60 . Furthermore, there is the first suction opening 32.1 in both cases to the print face 24 offset towards, namely one of the suction side surface lies 23 proximal edge 32.1.1 the opening 32.1 by at least 50% of the blade thickness 61 from the suction side surface 23 spaced. With the variant according to 7th extends the first suction opening 32.1 even in the print face 24 into it, correspondingly is a radially outer edge 24.1 the print face 24 opposite a radially outer edge 23.1 the suction side surface 23 in the area of the opening 32.1 offset radially inwards.

8th illustrates the shovel blade 10 in a tangential section and shows the first injection opening 32.2 . The first channel 32 opens at an angle 80 from around 15 ° to the suction side surface 23 . This profile section also shows the axial chord length 81 illustrated in the axial direction 82 between a front tangent 83 and a back tangent 84 is taken.

List of reference symbols

1

Turbo machine

1a

compressor

1aa

Low pressure compressor module

1ab

High pressure compressor module

1b

Combustion chamber

1c

turbine

2

Longitudinal axis

3

Compressor gas

4th

Gas duct

10

Blade

20th

Shovel blade

21st

Leading edge

22nd

Trailing edge

23

Suction side surface

23.1

Radially outer edge of the suction side surface

24

Print face

24.1

Radially outer edge of the print side surface

25th

Blade root

26th

Shovel tip

31

Second channel

31.1

Second suction opening

31.2

Second injection opening

32

First channel

32.1

First suction opening

32.1.1

Proximal edge

32.2

First injection port

35

Fission fluid

35.1

Splitting fluid in the axially front area

35.2

Fissure fluid in the part nourishing the blockage area

36

Core of the cleft vertebra

37

Blockage area

38

Mainstream

40

Central axis

50

Stagnation point area

51

Second fluid

52

First fluid

60

Central axis

61

Blade thickness

80

angle

81

Axial chord length

82

Axial direction

83

Front tangent

84

Back tangent

Claims (15)

Rotor blade (10) for arrangement in a gas duct (4) of a turbo-engine (1), with a blade (20) which extends from a blade root (25) to a blade tip (26), wherein the airfoil (20) has a leading and a trailing edge (21,22), as well as a suction side surface (23) and a pressure side surface (24), and wherein the airfoil (20) has a first channel (32) in its interior for throughflow with a first fluid (52) which extends away from a first suction opening (32.1) provided for sucking the first fluid (52) out of the gas channel (4), wherein the first suction opening (32.1) is arranged on the blade tip (26). Blade (10) after Claim 1 , in which the first suction opening (32.1) is arranged on the blade tip (26) between 20% and 100% of an axial chord length (81) taken from front to back. Blade (10) after Claim 1 or 2 , in which the airfoil (20) has a second channel (31) for a second fluid (51) to flow through which opens into a second injection opening (31.2) provided for blowing the second fluid (51) into the gas channel (4). Blade (10) after Claim 3 , in which the second injection opening (31.2) on the Blade tip (26) is arranged, specifically in front of the first suction opening (32.1), preferably between 0% and 20% of an axial chord length (81) taken from front to back. Blade (10) after Claim 4 , in which the second channel (31) at the second injection opening (31.2) is inclined towards the pressure side surface (24). Blade (10) after one of the Claims 3 to 5 , in which the second channel (31) extends away from a second suction opening (31.1) provided for sucking the second fluid out of the gas duct (4), which is preferably arranged in the pressure side surface (24) in an area on the front edge (21) is. Rotating blade (10) according to one of the preceding claims, in which the first channel (32) at the first suction opening (32.1) is inclined towards the pressure side surface (24). Rotating blade (10) according to one of the preceding claims, in which, viewed in a section perpendicular to the airfoil chord at the blade tip (26), an edge (32.1.1) of the first suction opening (32.1) which is proximal to the suction side surface (23), is spaced from the suction side surface (23) by at least 50% of a blade thickness (61) taken in the cut at the blade tip (26). Rotating blade (10) according to one of the preceding claims, in which the first suction opening (32.1) extends into the pressure side surface (24) so that a radially outer edge (24.1) of the pressure side surface viewed in a section perpendicular to the blade chord at the blade tip (26) (24) is offset radially inward relative to a radially outer edge (23.1) of the suction side surface (23). Rotating blade (10) according to one of the preceding claims, in which the first channel (32) opens into a first injection opening (32.2) provided for blowing the first fluid (52) into the gas channel (4), which is preferably in the suction side surface (23). is arranged. Blade (10) after Claim 10 , in which the first channel (32) at the first injection opening (32.2) opens into the suction side surface (23) at an angle (80) of at most 30 °. Blade (10) after Claim 10 or 11 , in which the first injection opening (32.2) is arranged between 0% and 50% of a radial blade height taken from radially inside to radially outside. Compressor module (1aa, 1ab) with a rotor blade (10) according to one of the preceding claims. Method for producing a rotor blade (10) according to one of the Claims 1 to 12 or a compressor module (1aa, 1ab) Claim 13 , in which at least the blade (20) of the rotor blade (10) is built up generatively. Use of a blade (10) according to one of the Claims 1 to 12 or a compressor module (1aa, 1ab) Claim 13 in a turbomachine (1), the first fluid (52) being sucked off at the first suction opening (32.1) and flowing through the first channel (32).

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