RU2596691C2 - Method of dispersion of gas turbine engine compression stage and stage of dispersion for application - Google Patents

Abstract

FIELD: machine building.

SUBSTANCE: invention can be used in centrifugal compressors. Disclosed are nonaxisymmetric shapes in the flow direction and in tangential direction. According to the implementation version, the stage of dispersion of radial or combined centrifugal axial compressor of gas turbine engine comprises a wheel formed by two disks (9), between which in the centrifugal manner or inclined from the centre to the periphery the gaseous substance flows. Grid blades (60) are located between disks (9) to direct the gaseous substance movement between front edges (6a) of those blades (60) from the centre side and rear edges from the periphery side. At least one of disks (9) has inboard side (9i) containing at least zone (Z1, Z2) with alternating curvature with recess (91) and convexity (92), between two adjacent blades (60), at least in one of two essentially perpendicular directions, namely in the direction of flow (F) along blades (60) and in the inter-blade tangential direction.

EFFECT: invention is aimed at implementation of the air flow by installation of a disc of optimised shape.

9 cl, 9 dwg

Description

FIELD OF THE INVENTION

The invention relates to a method for scattering air flow in a compressor stage of a gas turbine engine, as well as to a scattering stage, which can practically implement this method.

The scope of the invention relates to improving the performance and surge limit of centrifugal compressors and combined centrifugal-axial compressors of the combination of structural elements of the dispersion of the considered stage. The objective of this set of structural elements of dispersion is the conversion of the kinetic energy of a gaseous substance created at the output of a centrifugal wheel forming a step into static pressure. The operation should proceed with a minimum loss of total pressure, while maintaining compressor stability sufficient to maintain the surging margin acceptable for the operation of the turbomachine.

A centrifugal compressor comprises at least one radial compression stage, i.e. capable of practically realizing the movement of air perpendicular to the central axis of the compressor. The combined centrifugal-axial compressor comprises at least one compression stage mounted obliquely with respect to said central axis.

The set of structural elements of the dispersion of the compression stage consists of a wheel formed by two disks between which a gaseous substance flows out centrifugally or inclined from the center to the periphery. The blades are placed between the discs along the entire length of the wheel. These blades form an outflow lattice between the leading edges of these blades from the center side and the trailing edges from the outside.

Prior art

The disks of the radial and combined sets of structural elements of scattering are usually flat, and the cross sections for the passage of gaseous matter between the blades are preferably variable. Changes in the cross sections for gas passage are determined by the passage section of the narrow passage of the diffuser and the level of deceleration between the front and rear edges of the blade lattice.

In other designs, axisymmetric disks are provided in conjunction with variable current tubes for additional control of the cross section for gas passage and, thus, optimization of scattering in the grating.

Such solutions allow monitoring only when measuring changes in the cross section for gas passage. No control of the tangential heterogeneity of the outflow of gaseous matter between the two blades is not carried out. However, this control allows you to adjust and optimize the outflow.

Disclosure of invention

The invention is directed to such an outflow by installing form-optimized discs, since these discs have the largest area “washed” by the flow. Thus, non-axisymmetric shapes in the direction of motion are proposed tangentially.

An object of the present invention is to provide a method for scattering an air stream in a compression stage of a gas turbine engine comprising a plurality of scattering components consisting of a wheel formed by two disks between which a gaseous substance flows out centrifugally or inclined from the center to the periphery. The lattice blades are placed between the disks along the entire length of the wheel to direct the gaseous substance between the leading edges of these blades on the center side and the trailing edges on the periphery side. According to this method, at least one of the disks contains at least one alternation of curved parts, concave and convex, in at least one of two essentially perpendicular directions, namely, in the direction of movement along the blades and in the tangent tangent direction.

In these modes, the three-dimensional shape of the gaseous substance jet allows you to redistribute and homogenize its movement in this jet: the secondary outflows, which are the generators of pressure loss, are substantially reduced. The position of the impacts in the transonic vanes is changed, and their intensity is weakened. In addition, aerodynamic blocking at the inlet of the combustion chamber following the compression stage is also significantly reduced.

The invention also relates to a dispersion stage of a radial or combined centrifugal-axial compressor of a gas turbine engine for implementing this method. Such a step comprises a wheel formed by two disks, between which a gaseous substance flows out in a centrifugal manner or inclined from the center to the periphery. The grating blades are distributed between the disks along the entire length of the wheel to direct the gaseous substance between the leading edges of these blades on the center side and the trailing edges on the periphery side. At least one of the disks contains an inner side containing at least one zone with alternating (with a recess and convex) curved parts between two adjacent vanes in at least one of two essentially perpendicular directions, namely the direction of movement along the blades and in the interscapular tangential direction.

According to preferred embodiments, the scattering stage comprises alternating zones (with a recess and convexity) between the blades, in particular up to substantially 80% (preferably substantially up to 50%) of the chord of the blade, at the leading edge of the vanes with a beginning in front of the leading edge and / or at the trailing edge, to the trailing edge of the trailing edge. These alternating zones (with a notch and a bulge) can be used on one and / or the other of two centrifugal (radial) and combined centrifugal-axial scattering disks, in particular, symmetrically with respect to the central symmetrical plane of the disks or in parallel if two disks are considered steps.

Description of the attached drawing figures

Other characteristics, features and advantages of the present invention are explained in the following description, which is not restrictive, given with reference to the accompanying drawing figures, including:

- figure 1 depicts a partial sectional view of a gas turbine engine containing an air diffuser;

- figa and 2b depict isometric views of the scattering stage with blades containing one or two disks;

- figa-3c - schematic views of the profiles of the disk in the direction of the outflow of air along the blade with two zones of alternating curvature, respectively, along the blade on the trailing edge to the rear of the blade and on the leading edge of the front of the blade;

- figure 4 depicts a schematic partial view in the interscapular tangential direction with a disk containing two zones of alternating curvature;

- figure 5 is an interscapular schematic view in isometry on the leading edge with a disk containing a zone of alternating curvature;

- Fig.6 is a schematic isometric view on the trailing edge of two blades with a flat disk and a disk containing a zone of alternating curvature.

DETAILED DESCRIPTION OF THE INVENTION

The terms "rear" and "front" mean the location relative to the expiration of the air flow. In all the figures of the drawing, the same digital positions make reference to paragraphs of description, in which structural elements corresponding to these digital positions are defined.

As shown in a schematic partial sectional view of the gas turbine engine 1 of the helicopter shown in FIG. 1, the air flow F is first sucked into the intake air intake duct 2, then is compressed between the blades 3 of the impeller 4 of the centrifugal compressor 5 and the cover 10. The turbine has axial symmetry around X′X axis.

In this case, the compressor 5 is centrifugal, and the compressed flow F, thus radially flowing out of the impeller 4. When the compressor is a combined centrifugal-axial flow, the flow flows inclined at an angle of 0 ° to 90 ° relative to the radial direction, perpendicular X′X axis.

Thus, the flow F passes through a diffuser or wheel 6, located at the outlet of the compressor 4, for straightening and directing to the input channels 7 of the combustion chamber 8.

To accomplish this straightening, the wheel 6 consists of a plurality of curved blades 60 mounted between two disks 9 around the circumference of the impeller 4 (in this case radially) and, thus, rotating around the X′X axis.

Figure 2 shows a more accurate isometric view of the diffuser 6 with vanes 60 rigidly connected to two disks 9. As shown in Fig. 2a, on which one disk is removed for clarity, each vane 60, as you know, contains a side 6e, the so-called back, and side 6i, the so-called trough. In the presented example, these sides are connected by a tapering front edge 6a and a rounded rear edge 6f in the direction of air flow. In the transverse backs and troughs directions, each blade 60 comprises flat side surfaces 6p rigidly connected to the disks 9.

The disks 9 shown in FIGS. 2a and 2b are generally flat. According to the invention, at least one of these disks 9 contains in the space E, which is defined between them, at least one zone of alternating curvature between the two blades 60.

In Fig. 3a, such a disk 9 is shown laterally in the direction of flow of air F along the blades 60 from the entrance to the leading edge 6a of the blade to the exit channel 7 to the combustion chamber. Two zones of alternating curvature Z1 and Z2 are made in the disk 9, along the blades 60. Each of these zones, on the flow side F and relative to the flat side portion 9p of the disk 9, contains a recess section 91 and a convex section 92. In the presented example, which does not have restrictive, zones Z1 and Z2 are generally elongated by approximately 80% of the length of the chord 6c of the blades 60.

As shown in the figures, the disk 9 has a very small thickness to simplify the design, but in reality it has a certain thickness, and alternating curvature zones are formed on the inner side 9i of the disk where the air flow F. The outer side 9e of the disk 9 can remain flat or also repeat the forms of alternating curvature - convex and notched, the reverse in this example, the form with a notch and convexity of the inner side 9i. In the first case, the disk has a variable thickness, and in the second case, the thickness is constant. The shape of the disks may depend, in particular, on the method used to implement zones of alternating curvature: by milling, EDM, molding, stamping, using a laser, etc.

As shown in FIGS. 3b and 3c, two zones of alternating curvature Z1 and Z2 are formed respectively on the trailing edge 6f of the blades 60 and to the rear of the blades at the entrance of the channel 7 and / or on the leading edge 6a from the front entrance of the blades 60.

Fig. 4 is a schematic partial view of the scattering stage in the interscapular "tangential direction" 6t, i.e. between the two blades 60. The arrow F indicates the direction of air flow. The inner side 9i of the disk 9 contains two zones of alternating curvature Z1 and Z2, which are mainly made between the inner 6i and outer 6e of the surfaces of the two blades 60.

The interscapular isometric view of the leading edge shown in FIG. 5 shows, in particular, a disk 9 with an alternating curvature zone Z1 between two blades 60. The zone contains a curvature section with a recess 91 on the trough 6i of the blade 60 and a section with a bulge 92 on the back 6e another blade 60. This alternation of curvature allows pressure equalization between the trough and the back of each blade 60. The cross section of the narrow passage 6s between the front edges 6a is maintained.

At the trailing edges 6f, the homogenization of the air flow F is shown in perspective in FIG. 6. With a flat disk 90 (shaded in FIG.) Between two vanes 60, an aerodynamic blockage is formed in zone Z0 with very little movement. In this configuration, based on the prior art, the movement of primary air (arrow F) occurs with a very large Mach number. In contrast, with a disc 9 with alternating curvature, according to the invention, the aerodynamic blocking zone is eliminated, and the air movement F is homogenized, occupying the entire provided flow section, with a lower Mach number.

The invention is not limited to practical examples that have been described and depicted. Thus, in the examples presented, the curved parts alternate in the direction of outflow of the gaseous substance F, but also in the tangential direction 6t. In other embodiments, zones of alternating curvature may be adjacent so that surface areas with the same curvature, with a recess or bulge, can border each other.

Claims (9)

1. The method of dispersing the outflow of air (F) in the compression stage (5) of a gas turbine engine (1) containing a set of diffusion structural elements formed by a wheel (6) formed by two disks (9), between which are centrifugal or inclined from the center to the periphery gaseous substance expires; moreover, the blades (60) of the grating are placed between the disks (9) along the entire length of the wheel (6) to direct the outflow of gaseous matter between the leading edges (6a) of these blades (6) from the center side and the trailing edges (6f) from the periphery side, characterized in that at least one of the disks (9) is provided with at least one set of alternating curved parts, concave (91) and convex (92), in the direction of flow (F) along the blades (6). 2. The method according to claim 1, according to which at least one of the disks (9) is also provided with at least one set of alternating curved parts, concave (91) and convex (92), in the interscapular tangential direction (6t ), essentially perpendicular to the direction of flow (F) along the blades (6). 3. The scattering stage of a radial or combined centrifugal-axial compressor of a gas turbine engine for implementing the method according to claim 1, comprising a wheel (6) formed by two disks (9), between which a gaseous substance flows out centrifugally or inclined from the center to the periphery; moreover, the blades (60) of the lattice are placed between the disks (9) along the entire length of the wheel (6) to direct the outflow of gaseous matter between the leading edges (6a) of these blades (60) from the center side and the trailing edges (6f) from the periphery side, characterized in that at least one of the disks (9) contains an inner side (91) containing at least one zone (Z1, Z2) of alternating curvature, with a recess (91) and a bulge (92), between two adjacent shoulder blades (60). 4. The scattering stage according to claim 3, in which at least one of the disks (9) contains the inner side (91), also containing at least one zone (Z1, Z2) of alternating curvature, with a recess (91 ) and convexity (92), between two adjacent vanes (60) in the interscapular tangential direction (6t). 5. The scattering stage according to claim 3, characterized in that it comprises alternating zones (Z1, Z2) with a recess (91) and a bulge (92) between the blades, in particular, essentially up to 80%, preferably up to 50%, of the chord ( 6c) vanes (60). 6. The scattering stage according to claim 4, characterized in that it comprises zones with a recess and convexity (Z1, Z2) on the leading edge (6a) of the blades (60) with a beginning in the front of the leading edge (6a). 7. The scattering stage according to claim 4, characterized in that it comprises zones with a recess and convexity (Z1, Z2) on the trailing edge (6f) of the blades (60) to the trailing edge of the trailing edge (6f). 8. The scattering stage according to claim 4, in which these alternating zones with a recess and a bulge (Z1, Z2) are used on one and / or the other of the two discs (9) of centrifugal or combined scattering. 9. The scattering stage according to the preceding paragraph, in which these alternating zones (Z1, Z2) are applied on the disks (9) symmetrically with respect to the plane of central symmetry of the disks (9) or in parallel.

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