RU2796182C1 - Blade platform, bladed row, impeller disk and gas turbine engine - Google Patents

Abstract

FIELD: gas turbine engines.

SUBSTANCE: platforms (6, 7, 8) blades, blade ring (100), impeller disk and gas turbine engine) are proposed. The platform of the blade contains a pair of side edges (71, 72) and a pair of end edges (73, 74), and the first side edge (71) in the direction of its extension contains a linear section (700) and a curved section (800), while the linear section (700) contains the first linear section (701) and the second linear section (702), and the curved section (800) contains the first curved section (801) and the second curved section (802). The first and second linear sections (701, 702) are connected to two ends of the curved section (800) and extend tangentially to the curved section (800), while the bends of the first and second curved sections (801, 802) extend in the same direction, and the first and second curved portions (801), 802) are tangentially connected on the outside.

EFFECT: effective performance.

9 cl, 12 dwg

Description

FIELD OF TECHNOLOGY

The invention relates to the field of turbomachines, in particular to the blade platform, blade ring, impeller disk and gas turbine engine.

BACKGROUND OF THE INVENTION

In a turbomachine such as the high pressure compressor of a gas turbine engine shown in FIG. 1-5, the blades of the row of blades 10 can be installed in the circumferential direction by placing herringbone shanks in the grooves of the compressor disc 2, with the conventional position of the herringbone shanks and grooves corresponding to the reference line T, as shown in FIG. 5, and there is a contact area M between the herringbone shank and the groove, so that an impeller is formed. The blade contains a platform configured to be aligned with the flow channel, and during engine operation a certain circumferential clearance must be provided between the platforms of the blades so that the fir-tree shank of the blade adjoins the working surface of the disc groove to transfer the load due to centrifugal force. The circumferential clearances of the blades on the blade ring 10 will accumulate, for example, as shown in FIG. 1, wherein the calculated circumferential clearance W is provided between the platform 1 of the first blade and the platform 5 of the last blade, which are the two ends of the blade ring 10 in the circumferential direction.

Based on the requirements of a gas turbine engine for a high-pressure compressor with high efficiency and a large margin of safety, it is necessary to provide a larger number of blades, i.e. increase the density of the blades, while it is necessary to increase the angle of twist to improve the performance of the rotor blades. A blade platform with high density, high set angle, and circumferential herringbone liners typically requires a sloped edge platform design, such as the parallelogram design shown in FIG. 1 and FIG. 2 and the platform structure shown in FIG. 2, in which the angle a of the platform cannot be a right angle of 90° in order to be able to install a sufficient number of blades and provide a large installation angle.

However, as shown in FIG. 2-4, by creating a ramped edge platform during the course of the present invention, the inventor has found that under the action of an external force, such as when performing a rotor unbalance test, the blade is twisted through a twisting angle B, with adjacent blades pressed against each other, and a pair of forces N (N=H⋅cosα⋅F, where H is the width of the platform, α is the angle of the platform, F is the pushing force between the platforms of adjacent blades) relative to the center P of the herringbone shank, so that the blades will rotate around the center P of the herringbone shank type, with a gap C between the platforms of adjacent blades. Since the platform is shown as a straight line with a constant slope k (k=tanα=constant), the platforms of the blades will not intersect with each other after twisting and can move in the same direction S, so that the circumferential clearance between the first blade platform 1 and the last blade platform 5 will increase to the actual circumferential clearance W'=W+n⋅C (n is the number of blades).

Since the actual circumferential clearance W' and the calculated circumferential clearance W are different, the measurement accuracy of the platform circumferential clearance and dynamic balance will be compromised, for example, the residual dynamic unbalance of the rotor cannot be accurately reflected by the rotor dynamic balance test, resulting in large potential safety hazards in the operation of the motor. .

Thus, in order to obtain a test design that will accurately reflect the dynamic unbalance of the rotor, one way of testing according to the prior art is to repair or replace the blades with different platform widths to eliminate the effect of gap C, so that even if twisting occurs, the actual circumferential clearance W' will be equal to the design circumferential clearance W. However, during the course of this invention, the inventor found that this solution could result in insufficient circumferential clearance W during actual operation of the blade ring 10, causing adjacent blade platforms to press against each other. to each other, and the fir-tree shank of the blade will not fully adhere to the working surface of the groove of the disk, while the platform is twisted at an angle under the combined effect of centrifugal force and aerodynamic force. First, the contact between the herringbone root of the blade and the groove changes from surface contact to point or line contact, which adversely affects the force applied to the disc groove; at the same time, the blade angle changes, which leads to a change in the flow area and affects the aerodynamic characteristics; as the rotation speed increases, the blade platforms are adjacent and pressed against each other, and also stuck under the influence of the working environment of high temperature and high pressure, and the blade platforms often get stuck during disassembly, so that a large external force is required to knock out the blade platform to loosen it and then dismantle.

The prior art solutions also include a connection zone between the compressor disc and the blade platform, as shown in FIG. 5-7, which is a protruding limiting part made on one or both sides of the edge part of the blade platform or the edge part of the compressor disk, for example, in the form of the first protruding limiting part 3, which is formed by the protruding part 31 of the first groove located on one side of the edge part of the groove 21 of the compressor disk, and the protruding part 32 of the first platform located on the edge part of the platform 11 of the blades, as shown in FIG. 6, or the first protruding limiting part 3 and the second protruding limiting part 4, which are located separately on both sides, as shown in FIG. 7, wherein the clearance G between the blade platform and the lobe should be minimized as much as possible to limit the pitch angle of the blade. However, the actual circumferential clearance W' and the calculated circumferential clearance W are still different due to the clearance G.

The prior art solutions also include an intermediate obturator and a structure between platforms of adjacent blades to provide a fastening connection, which further complicates the design of the blade ring, and a high-strength intermediate obturator is required to ensure the reliability of the fastening.

Thus, there is a need in the art for a blade with a design that would overcome problems such as poor accuracy of blade circumferential clearance testing caused by blade angular rotation and poor accuracy of dynamic balance test results, and would also prevent the platforms from pressing against each other. to each other and their sticking when the crown of blades is in working order to improve the stability and reliability of the compressor and gas turbine engine.

SUMMARY OF THE INVENTION

The aim of the invention is to create a blade platform.

The aim of the invention is to create a blade crown.

The aim of the invention is to create an impeller disk.

The aim of the invention is to create a gas turbine engine.

According to one aspect of the invention, the platform of the blade contains a pair of side edges and a pair of end edges, and the side edge contains a linear section and a curved section, while the linear section contains the first linear section and the second linear section, and the curved section contains the first curved section and the second curved section; moreover, the specified first linear section and the specified second linear section are connected to two ends of the curved section and pass tangentially to the curved section, while the bends of the said first and second curved sections pass in the same direction, and the said first and second curved sections pass along the circumference.

In one or more embodiments of the blade platform, said first curved section extends tangentially to said first linear section, and said second curved section extends tangentially to said second linear section.

In one or more embodiments of the blade platform, said first and second linear sections are perpendicular to the end edges respectively connected.

In one or more embodiments of the blade platform, the edge section of the blade platform and the body section of the blade platform are in the same plane, forming a flat panel structure.

According to one aspect of the invention, the blade ring comprises blades, each of which contains any of the above platforms, while the side edges of a pair of side edges have the same design, and the side edges of the platform of adjacent blades have the same design.

In one or more embodiments of the blade ring, adjacent blades are configured to rotate through the same angle around the centers of their own ridges under the action of the centrifugal force of the blade ring, so that the said first and second curved sections, made on the side edges of the platforms of adjacent blades, are separated, forming a non-contact structure with a non-uniform gap between the parts of the non-contact structure, wherein said first and second linear sections, made on the side edges of the platforms of adjacent blades, are attached to provide a connection without an intermediate locking part between the platforms of the blades at the two ends of the blade ring in the circumferential direction.

According to one aspect of the invention, the impeller disk comprises a blade ring and a blade disk, the blade ring and the blade disk being tongue-and-groove, the blade ring being any of the above.

In one or more embodiments of the impeller disk, the edge section of the blade disk groove and the body section of the blade disk groove are in the same plane.

According to one aspect of the invention, there is provided a gas turbine engine comprising a fan, a compressor and a turbine, the fan and/or compressor and/or turbine comprising any of the impeller disks mentioned above.

In general, the advantages of the present invention include, but are not limited to, the following:

since the lateral edge contains a linear section and a curved section, the bends of these first and second curved sections extend in the same direction and are tangent to each other when the blade rotates at an angle around the center of the ridge P under the action of an external force, since the bending of the curved section is variable at different locations, the gap between adjacent platforms will be uneven, resulting in the movement of adjacent blades along the same direction, so that adjacent platforms can only have a fixed angular position; thus, the actual circumferential clearance W' can be matched to the calculated circumferential clearance W in the circumferential clearance test and the dynamic balance test, with the problems of the platforms being pressed against each other and stuck in operation, as in the prior art, or associated with insufficient accuracy of the results of blade circumferential clearance tests and dynamic balance tests, can be prevented, and a snug fit and pressing of the tab on the blade to the working surface of the disk groove for effective load transfer, while the blade angle meets the design requirements, which provides the required high efficiency and high margin of safety of the compressor in the gas turbine engine, thereby increasing the efficiency and reliability of the gas turbine engine.

BRIEF DESCRIPTION OF THE DRAWINGS

The above and other features, features and advantages of the present invention will become more apparent from the following description in conjunction with the accompanying drawings and embodiments. It should be noted that the drawings are examples only and are not drawn to scale and should not be taken as limiting the scope of the protection claimed by this invention.

On the drawings:

Fig. 1 schematically depicts a fragment of a blade ring known from the prior art;

Fig. 2 schematically depicts a blade platform according to the prior art;

Fig. 3 and FIG. 4 schematically depicts a blade ring under the action of an external force according to the prior art;

Fig. 5 and FIG. 6 schematically shows a blade and a blade disc with a protruding limiting part made on one side of the groove, according to the prior art;

Fig. 7 schematically shows a blade and a blade disk with a protruding limiting part made on both sides of the groove, according to the prior art;

Fig. 8 schematically shows a fragment of a blade ring according to one embodiment;

Fig. 9 schematically shows the blade of the blade ring according to the embodiment shown in FIG. 8;

Fig. 10 and FIG. 11 is a schematic representation of a blade of a row of blades according to the embodiment shown in FIG. 8;

Fig. 12 schematically depicts, for comparison, the construction of a tongue-and-groove connection of a blade and a blade disc according to one embodiment and a boundary portion according to the prior art shown in FIG. 6.

DETAILED DESCRIPTION OF THE INVENTION

Further, various embodiments or examples of implementation of technical solutions according to this invention are disclosed. To simplify the description, the following describes specific embodiments of each element and device, which are only examples and do not limit the scope of legal protection of this invention. The expressions "one embodiment", "an embodiment" and/or "some embodiments" refer to a specific feature, design or characteristic that is associated with at least one embodiment of the present invention. Thus, it should be emphasized and noted that the expressions "an embodiment", or "one embodiment", or "alternative embodiment", mentioned twice or more times in different places in this description, do not necessarily refer to the same embodiment. . In addition, certain features, structures, or characteristics in one or more embodiments of this invention may be combined as needed.

In the following embodiments, a high pressure compressor impeller disk of a gas turbine engine is given as an example, but the invention is not limited to this embodiment. For example, the invention can also be applied to a turbine disk of a gas turbine engine, and even to a fan disk, and the invention is also not limited to these embodiments, provided that the blade row blade and disk are connected by a herringbone shank and tongue and groove.

As shown in FIG. 8 and FIG. 9, in some embodiments, the blade ring 100 is mounted on the blade disk 2 using a tongue and groove connection. The crown 100 includes blades, the platforms of adjacent blades being the first platform 7 and the second platform 8 as examples, the platforms at both ends of the crown 100 in the circumferential direction being the first platform 7 and the third platform 6.

As shown in FIG. 9, the first blade deck 7, taken as an example, comprises a pair of side edges 71, 72 and a pair of end edges 73, 74, i. the first side edge 71, the second side edge 72, the first end edge 73 and the second end edge 74. The first side edge 71, taken as an example, contains a linear section 700 and a curved section 800, and the linear section 700 contains the first linear section 701 and the second a linear portion 702 and a curved portion 800 comprising a first curved portion 801 and a second curved portion 802 centered respectively at the center O1 of the first curved portion and the center 02 of the second curved portion; thus, the centers of the curved portions of the second side edge 72 are the center O1' of the third curved portion and the center O2' of the fourth curved portion. The compressive force acting on the first side edge 71 from the adjacent platform is the first compressive force F1, and the compressive force acting on the second side edge 72 from the adjacent platform is the second compressive force F1', with l being the length of the platform. The first linear portion 701 and the second linear portion 702 are connected to two ends of the curved portion 800 and are tangent to the curved portion 800, for example, as shown in FIG. 9, wherein the first linear portion 701 is connected to and extends tangentially to one end of the first curved portion 801, and the second linear portion 702 is connected to and extends tangentially to one end of the second curved portion 802. The design of the curved portion 800 is such that the curves of the first curved portion 801 and the second curved portion 802 are in the same direction, with the first curved portion 801 and the second curved portion 802 extending circumferentially. The bends of the first curved section 801 and the second curved section 802 are in the same direction, so that:

k1=ƒ(x)=(Arc1)≤0, k2=ƒ(x)=(Arc2)≤0

or

k1=ƒ(x)=(Arc1)≥0, k2=ƒ(x)=(Arc2)≥0,

where Arc1 is the first curve 801, k1 is the curve of the first curve 801, Arc2 is the second curve 802, and k2 is the curve of the second curve 802.

Referring to FIG. 10 and FIG. 11, the advantages of the above embodiments are that the problem shown in FIG. 3 and FIG. 4, when the actual circumferential clearance W' differs from the calculated circumferential clearance W due to the parallelogram-shaped inclined edge platform can be solved, so that the problem that the platforms are pressed against each other and stuck in the running state, which caused by the repair or replacement of blades with different platform widths can be eliminated, while ensuring that the ledge on the blade fits snugly and the disc groove is pressed against the working surface for efficient load transfer, so that the blade angle meets the design requirements, which ensures the required high efficiency and high margin of safety of the compressor in the gas turbine engine, thereby increasing the efficiency and reliability of the gas turbine engine. There is no need for an intermediate locking portion provided between the blade platforms at the two ends in the circumferential direction to provide the fastening connection used in prior art solutions, which can simplify the design of the blade ring and reduce the manufacturing cost.

The principle of achieving the above advantages is shown in FIG. 9-11, namely, under the action of the centrifugal force of the blade ring, adjacent platforms, namely the first platform 7 and the second platform 8, will rotate at the same angle around the center of the ridge, so that said first and second curved sections made on the side edges the first platform 7 and the second platform 8 blades will be separated, forming a non-contact structure. Since the curvature of the curved portion is variable at different locations, when the curved portion of the first platform 7 rotates around the center of the ridge P, the curved portions of adjacent platforms, namely the first platform 7 and the second platform 8, will be separated to form a non-contact structure. Since said curved sections are made, the gap in the non-contact structure between adjacent platforms, namely the first platform 7 and the second platform 8, is uneven, and the said first and second linear sections, made on the side edges of the first platform 7 and the second platform 8, are adjacent to ensure connection , and they are adjacent at location E and location F to provide a connection, as shown in FIG. 10 and FIG. 11, so that the gap in the non-contact structure between the first platform 7 and the second platform 8 will be uneven, resulting in the movement of the first platform 7 and the second platform 8 along the same direction S, as shown in FIG. 3 and FIG. 4, thereby eliminating the difference between the actual circumferential clearance W' and the calculated circumferential clearance W. Since angular torsion to form a larger circumferential clearance does not occur, there is no need for an intermediate locking portion to be placed between the blade platforms at the two ends in the circumferential direction to provide a fastening connection.

Those skilled in the art will appreciate that the above curved portion is to be understood in a broad sense, not limited to a narrow interpretation, such as the arc shape shown in FIG. 8-11, including all variations in which the curvature of the curved portion will be variable at different locations.

Referring to FIG. 8-11, in one or more embodiments, the blade platform may be of a particular design, in which the first linear section 701 and the second linear section 702 of the linear section 700 are perpendicular to the first end edge 73 and the second end edge 74 connected, respectively, so that the connections at the locations E and F are more stable. Like the prior art, as shown in FIG. 9, the angles between the end edge and the connecting lines L1, L2 of the end points of the first end edge 73 and the second end edge 74 of the blade platform on one side are sharp, providing an effect similar to the prior art sloped edge platform, which improves the thickness of the blades and the angle installation. Also referring to FIG. 8-11, for one blade of the crown 100, each side edge of a pair of side edges has the same design, while for several blades, the side edges of each platform have the same design, which simplifies the implementation process and reduces design complexity.

Referring to FIG. 6 and 12, in some embodiments, the protruding boundary shown in FIG. 6 may be optional for the platform in the above embodiment and is shown by the dashed line in FIG. 12. Compared to FIG. 6, the platform edge portion 110 and the platform body portion 111 are in the same plane, so that the blade platform is like a flat panel structure, and the groove edge portion 210 of the disc 2 and the groove body portion 211 of the blade disc 2 can also be in the same plane, so that the cost manufacturing of the blade platform and blade disk can be reduced.

As can be seen from the above, in the above embodiment, when testing the rotor for unbalance, such a test can be carried out directly on the impeller disk with the blade platform. It is not necessary to perform the steps of repairing or replacing blades with platforms of a different width during the test, as in the prior art, in order to maintain the actual circumferential clearance W' according to the calculated circumferential clearance W, so using the impeller disk in the above embodiment to perform the unbalance test rotor can not only provide accurate test results, but also simplify the test behavior process.

In essence, the advantages of using the blade platform, blade ring and impeller disk according to the above embodiments are that since the side edge contains a linear section and a curved section, the bends of these first and second curved sections extend in the same direction and are tangent to each other, so that when the blade rotates at an angle around the center of the ridge P under the action of an external force, since the bending of the bent section is variable at different locations, the gap between adjacent platforms will not be uniform, which causes the adjacent blades to move in the same direction, so that adjacent platforms may have only fixed angular position, so the actual circumferential clearance W' can be equivalent to the calculated circumferential clearance W in the circumferential clearance test and dynamic balance test, with the problems of the platforms being pressed against each other and stuck in running condition as in designs of the prior art, or associated with insufficient accuracy of the results of tests of the circumferential gap of the blades and tests of dynamic balance, can be eliminated, and a snug fit of the blade protrusion is ensured and the disk groove is pressed against the working surface for effective load transfer, while the blade installation angle corresponds to the design requirements, thereby providing the required high efficiency and a large margin of safety of the compressor in the gas turbine engine, thereby increasing the efficiency and reliability of the gas turbine engine.

Although the present invention has been disclosed by means of the above embodiments, they do not limit the present invention, and any person skilled in the art can make possible changes and modifications within the spirit and scope of the invention. Thus, any modifications, equivalent changes and modifications made to the above embodiments according to the technical essence of the invention, without departing from the content of the technical solutions of the invention, are included in the scope of legal protection defined by the claims.

Claims (11)

1. The blade platform, containing a pair of side edges and a pair of end edges, moreover, the side edge contains a linear section and a curved section, while the linear section contains the first linear section and the second linear section, and the curved section contains the first curved section and the second curved section, wherein the first and second linear sections are connected to the two ends of the curved section and run tangentially to the specified curved section, and the bends of the first and second curved sections are in the same direction, and the first and second curved sections are circumferential. 2. The platform of claim 1, wherein the first curved portion extends tangentially to the first linear portion and the second curved portion extends tangent to the second linear portion. 3. Platform according to claim 1, in which the first and second linear sections are perpendicular, respectively, to the attached end edges. 4. The platform according to claim. 1, in which the edge section of the blade platform and the body of the blade platform are in the same plane, forming a flat panel structure. 5. Blade crown containing blades, each blade contains a platform according to any one of paragraphs. 1-4, wherein the side edges of said pair of side edges are of the same design and the side edges of adjacent blade platforms are of the same design. 6. The blade ring according to claim 5, in which adjacent blades are made with the possibility of turning at the same angle around the centers of their crests under the action of the centrifugal force of the specified crown, so that the said first and second curved sections located on the side edges of the platforms of adjacent blades are separated, forming a non-contact structure with an uneven gap in said non-contact structure, wherein said first and second linear sections located on the side edges of the platforms of adjacent blades are fixed to provide a connection without an intermediate locking part between the platforms of the blades at the two ends of the blade ring in the circumferential direction. 7. The impeller disk containing the blade ring and the blade disk, and the blade ring and the blade disk are connected by a tongue and groove connection, while the blade ring is made according to claim 5 or 6. 8. The disk according to claim 7, in which the edge section of the groove of the blade disk and the body section of the groove of the blade disk are in the same plane. 9. A gas turbine engine containing a fan, a compressor and a turbine, and the fan and/or compressor and/or turbine contain an impeller disk made according to claim 7 or 8.

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