Gas turbine engine rotor blade and gas turbine engine
By designing the torsion angle of the first edge plate on the rotor blade of the turbine engine, a complex airflow path is formed, the problem of poor rim sealing effect in the prior art is solved, and a more efficient sealing effect is achieved, and the safety and stability of the engine are improved.
Patent Information
- Application Number
- CN202311556134.7
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2023-11-21
- Publication Date
- 2025-05-23
AI Technical Summary
In existing aircraft engines, the tightening effect of the rim is not ideal, which causes gas to flow back into the disk cavity, causing ablation, and it is difficult to use the tightening gas, which affects the efficiency and fuel consumption of the whole machine.
A turbine engine rotor blade is designed, which includes a first edge plate protruding forward along the turbine axial direction. The extension curve of the first edge plate is at a non-zero angle with the circumferential direction of the turbine disc in its position, forming a complex air flow path and increasing the resistance to gas intrusion.
By improving the rim sealing structure, the resistance to gas intrusion is increased, the sealing effect of the original rim sealing device is strengthened, the efficiency of rim sealing of the turbine disc rim is improved, and the safety and stability of engine operation is increased.
Smart Images
Figure CN120026966A_ABST
Abstract
Description
Technical Field
[0001] The invention relates to the technical field of aeroengines, and in particular to a turbine rotor blade capable of improving the sealing effect of a wheel rim. Background Art
[0002] Existing aircraft engines use gas turbine engines to convert the internal energy of high-temperature combustion gas into kinetic energy through the rotor. In order to prevent the high-temperature combustion gas flowing through the mainstream of the turbine from flowing back into the disc cavity and causing ablation of the turbine disc, a secondary air system is generally used to seal the disc cavity. Usually, the secondary air system draws air from the compressor, and the sealing gas flow path and the complex cooling structure arranged along the flow path are used to achieve the effect of sealing and cooling the disc cavity. In the prior art, fish mouths, double-layer seals, comb teeth seals and other structures are set on the edge of the disc, and these structures cooperate with the introduced sealing gas to achieve the sealing effect of the disc cavity.
[0003] The amount of air bleed from the compressor needs to be strictly controlled. Too little air bleed may cause mainstream gas backflow; too much air bleed will reduce the efficiency of the whole machine, increase fuel consumption, and affect the overall performance of aircraft engines and gas turbines. Therefore, it is an important part of engine design to reasonably arrange the amount of air used for sealing and improve the sealing effect. Summary of the invention
[0004] One of the purposes of the present invention is to improve the sealing structure, thereby further improving the sealing efficiency.
[0005] To this end, the present invention provides a turbine engine rotor blade, comprising a blade body, a root, and a connecting section connecting the blade body and the root, wherein the root is used to be embedded in an axial groove fixed to a turbine disk, and the connecting section comprises a first edge plate for wheel rim sealing, which protrudes forward along the axial direction of the turbine from the connecting section body and constitutes a part of the sealing mechanism, and an extension curve of the first edge plate forms a non-zero angle with the circumferential direction of the turbine disk at which it is located.
[0006] It should be understood that the so-called "axial" refers to the axial direction of the engine; and "front and rear" are defined according to the airflow of the turbine engine along the axial direction of the engine, "front" refers to the upstream direction, and "rear" refers to the downstream direction. The "extension curve" of the first edge plate refers to the curve defined by the main direction of extension of the first edge plate, and also includes a straight line. The first edge plate according to the present invention does not extend in a circumferential direction on a plane perpendicular to the axial direction of the engine, but forms a non-zero angle with the circumferential direction, such as a straight line intersecting the circumference, or an arc intersecting the circumference and having a different curvature and / or direction, or any other non-circular curve.
[0007] It should be understood that when the rotor blade is mounted on the turbine disk, it cooperates with the rearward protruding edge plate provided on the upstream intermediate support to form a sealing structure. For example, two protruding edge plates are provided on the intermediate support at different radial positions, namely the second edge plate and the third edge plate, which generally extend in the circumferential direction. The first edge plate of the rotor blade is inserted between the second edge plate and the third edge plate, for example, to form a sealing structure. However, since the rotor blade rotates around the axis and has relative movement with the stationary intermediate support, there must be a gap between these edge plates. According to the present invention, the first edge plate forms a non-zero angle with the circumferential direction, so the distance between it and the second edge plate (or the third edge plate) extending in the circumferential direction is not constant, but varies with different circumferential positions. It can be understood that where the gap between the first edge plate and the second edge plate (or the third edge plate) is reduced, the resistance to gas intrusion will increase. In other words, this arrangement of the first edge plate causes the gas path of the gas to become more complicated, increases the resistance to gas intrusion, and strengthens the sealing effect of the original sealing device. Therefore, this technical solution improves the rim sealing effect without increasing the sealing air volume, improves the sealing efficiency of the turbine disk rim, and thereby increases the safety and stability of the engine operation.
[0008] According to some embodiments of the present invention, the first edge plate is at a torsion angle with the circumferential direction of the turbine disk to form an impeller structure for extracting the sealing gas in the disk cavity when the turbine rotates. In other words, the first edge plates arranged along the circumferential direction in the prior art are replaced with blade shapes with a torsion angle, i.e., with airfoil features. When the rotor blades are all installed on the rotor turbine disk, multiple first edge plates are evenly arranged along the circumference of the turbine disk, thereby forming an impeller structure that rotates around the engine axis. The torsion angle of the first edge plate is set according to the rotation direction of the turbine rotor so that the formed impeller structure extracts the sealing gas in the disk cavity when the rotor rotates. In other words, the impeller structure formed by the first edge plate increases the external pressure of the sealing gas at the sealing structure, making it more difficult for the fuel gas to enter the disk cavity. As a result, the structure can further enhance the sealing effect and improve the sealing efficiency.
[0009] According to some embodiments of the present invention, the length of the first edge plate in the circumferential direction of the turbine disk is greater than the length of the connecting section. In other words, the first edge plate extends beyond the connecting section in the circumferential direction and overlaps radially with the first edge plates on adjacent rotor blades. It can be understood that overlapping two adjacent first edge plates will result in a smaller gap between them and the second and third edge plates, as well as a more complex gas intrusion path, thereby increasing the resistance to gas intrusion, thereby further strengthening the sealing effect and improving the sealing efficiency.
[0010] According to some embodiments of the present invention, the two ends of the first edge plate in the circumferential direction of the turbine disk respectively exceed the two ends of the connecting section. Therefore, each first edge plate extends to the circumferential inside of the two adjacent rotor blades, extending the circumferential length of the radially overlapping part, and also extending the length of the part with reduced gap, further increasing the resistance to gas intrusion, strengthening the sealing effect, and improving the sealing efficiency.
[0011] The present invention also relates to a turbine engine, which comprises: a turbine engine rotor blade as described above; an intermediate support having a second edge plate and a third edge plate arranged along the radial direction of the turbine, wherein the first edge plate is inserted between the second and third edge plates in the radial direction of the turbine, and there is a gap between the three in the radial direction of the turbine. As described above, the first edge plate, the second edge plate, and the third edge plate are stacked in the radial direction, and since the turbine rotor disk rotates while the intermediate support is fixed, the first edge plate rotates relative to the second and third edge plates. As described above, the shape of the first edge plate is such that there is a portion with a reduced gap between it and the second edge plate and the third edge plate, and the resistance to the intrusion of the gas therein will increase. Therefore, this technical solution improves the effect of the wheel rim sealing without increasing the sealing gas volume, improves the sealing efficiency of the turbine disk wheel rim, and thereby increases the safety and stability of the engine operation.
[0012] It should be understood that the above general description and the following detailed description are exemplary and explanatory only and cannot limit the present invention. Other features, objects and advantages of the present invention will become apparent from the specification, drawings and claims. BRIEF DESCRIPTION OF THE DRAWINGS
[0013] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for describing the embodiments. It is easy for those skilled in the art to understand that these drawings are only for illustrative purposes and are not intended to limit the scope of protection of the present invention. For illustrative purposes, these drawings may not be drawn completely to scale.
[0014] Figure 1 It is a partial axial cross-sectional view of a turbine engine in the prior art.
[0015] Figure 2 It is a partial stereoscopic view of three adjacent rotor blades after installation in the prior art, in which the rotor turbine disk is omitted.
[0016] Figure 3 is similar to Figure 2 , showing a partial perspective view of three adjacent rotor blades after installation according to an embodiment of the present invention.
[0017] Figure 4 is similar to Figure 2, showing a partial stereoscopic view of three adjacent rotor blades after installation according to another embodiment of the present invention.
[0018] List of reference numerals:
[0019] 10, 10': rotor blade; 12: blade body; 14, 14': first edge plate; 16: connecting section; 18: tenon; 2: intermediate support; 22: second edge plate; 23: third edge plate; 30: baffle; 32: fourth edge plate; 50: air duct hole; 52: sealed air path. DETAILED DESCRIPTION
[0020] Exemplary embodiments will be described in detail herein, examples of which are shown in the accompanying drawings. When the following description refers to the drawings, the same numbers in different drawings represent the same or similar elements unless otherwise indicated. The embodiments described in the following exemplary embodiments do not represent all embodiments according to the present invention. Instead, they are merely examples of devices consistent with some aspects of the present invention as detailed in the appended claims.
[0021] Figure 1 FIG. 1 is a cross-sectional view along the axial direction of the engine, showing a schematic diagram of a sealing structure in a turbine engine of the prior art. Figure 1 As shown, an intermediate support 2 and a rotor turbine disk 1 are sequentially arranged along the engine axial direction A from upstream (left in the figure) to downstream (right in the figure), and a plurality of rotor blades 10' are installed on the circumferential edge of the turbine disk 1. In order to prevent the high-temperature combustion gas (its flow path is schematically represented by F in the figure) flowing through the mainstream of the turbine from flowing back into the disk cavity, a secondary air system is used to seal the disk cavity. The secondary air system draws air from the air duct hole 50, and cools the disk cavity through the flow path of the sealing gas 52 and the complex cooling structure arranged along the flow path. In order to achieve the sealing effect of the disk cavity, a sealing structure will be set on the edge of the wheel disk. Figure 1 A double-layer sealing structure is shown, which includes a first edge plate 14' arranged on the rotor blade 10' and extending forward, a second edge plate 22 and a third edge plate 23 arranged on the intermediate support 2 and extending backward, and a fourth edge plate 32 arranged on the rotor baffle 30. These edge plates are arranged in sequence in the radial direction and staggered and overlapped in the axial direction to form a complex airflow path, so as to increase the resistance of the main flow of gas entering the disc cavity and obtain a better sealing effect.
[0022] Figure 2A partial stereoscopic view of three adjacent rotor blades after installation in the prior art is shown, and the viewing angle is roughly perpendicular to the axial direction of the engine from upstream to downstream. As shown in the figure, each rotor blade 10' includes a blade body 12 and a tenon 18, and a connecting section 16 connecting them. The shape of the tenon 18 is complementary to the tenon groove (not shown) of the rotor turbine disk 1, and the assembly of each rotor blade 10' is achieved by embedding and fixing the tenon 18 into the tenon groove. The connecting section 16 includes a first edge plate 14' located at the front end, and the so-called "front" refers to the side where the airflow reaches the blade, and the "rear" is the opposite side, that is, the side where the airflow leaves the blade. The turbine engine also includes an axial stop device (not shown) to maintain the axial position of the rotor blade after installation, such as an annular baffle or a stop ring.
[0023] like Figure 2 As shown, the connecting section 16 of the rotor blade 10' is provided with a first edge plate 14' protruding forward, and in the solution of the prior art, it extends along the circumferential direction C. It should be understood that when all the rotor blades 10' are installed on the turbine disk 1, the first edge plates 14' are assembled into a ring, that is, arranged along the same ring. Similarly, the second edge plate 22 and the third edge plate 23 protruding backward provided on the intermediate support also extend in the circumferential direction. In this way, the gaps between the first edge plate 14' and the second edge plate 22 and the third edge plate 23 are respectively constant.
[0024] Figure 3 A rotor blade according to an embodiment of the present invention is shown, and Figure 2 Similarly, a partial perspective view of three adjacent blades after installation is also shown. Figure 2 As in the prior art shown, each rotor blade 10 includes a blade body 12 and a tenon 18, and a connecting section 16 connecting them. The connecting section 16 includes a first edge plate 14 located at the front end, and the so-called "front" refers to the side where the airflow reaches the blade. It can be seen that the first edge plate 14 of the rotor blade 10 according to the present invention does not extend along the circumferential direction C of the turbine disk, but extends along the curve BB, and forms a non-zero angle with the circumferential direction C of the turbine disk where it is located. As a result, the gap between the first edge plate 14 and the second edge plate 22 (or the third edge plate 23) can be partially reduced, thereby increasing the resistance to gas intrusion and strengthening the sealing effect.
[0025] In this embodiment, the first edge plate 14 also forms a twist angle with the circumferential direction C of the turbine disk, that is, the whole is in the shape of a blade with airfoil characteristics. Therefore, when the rotor blades are installed on the rotor turbine disk, multiple first edge plates 14 are evenly arranged along the circumference of the rotor disk, thereby forming an impeller structure rotating around the engine axis A. Figure 3As shown, the rotation direction R of the turbine rotor disk is counterclockwise as shown in the figure, and the first edge plate 14 is set to be an arc with a lower left and a higher right, that is, the blade shape is formed by rotating counterclockwise relative to the circumferential direction C with the leading edge (left end) of the first edge plate 14 as the starting point, for example, as shown Figure 3 The pre-twisted shape shown enables the impeller structure formed to extract the sealing gas in the disc cavity when the rotor rotates. Thus, the impeller structure formed by the first edge plate 14 can further increase the external pressure of the sealing gas at the sealing structure, making it more difficult for the gas to enter the disc cavity.
[0026] Figure 4 A rotor blade according to another embodiment of the present invention is shown. Figure 2 Similarly, a partial perspective view of three adjacent blades after installation is also shown. Figure 3 The embodiment shown differs in that Figure 4 The width of the first edge plate 14 of the rotor blade 10 shown (i.e., the length along the circumference of the turbine disk) exceeds the width of the connecting section 16, and the two circumferential ends 141, 142 respectively exceed the two ends of the connecting section 16 in the circumferential direction. In this way, when each rotor blade 10 is assembled on the rotor turbine disk as shown in the figure, adjacent first edge plates 14 will overlap each other in the radial direction, for example, the second end 142 of the left first edge plate overlaps the first end 141 of the right first edge plate. The first edge plate 14 with an extended width results in a smaller gap between it and the second and third edge plates, as well as a more complex gas intrusion path. This can increase the resistance to gas intrusion, thereby further strengthening the sealing effect and improving the sealing efficiency. Of course, it is also possible to set the first edge plate 14 so that only one end extends out of the connecting section 16 and overlaps with the adjacent first edge plate.
[0027] Unless otherwise defined, the technical and scientific terms used herein have the same meanings as those generally understood by those skilled in the art of the technical field of the present invention. The terms used in this application are only for the purpose of describing specific embodiments, but are not intended to limit the present invention. The singular forms of "a", "an" and "the" used in this application are also intended to include plural forms, unless the context clearly indicates other meanings. It should also be understood that the term "and / or" that may be used herein refers to and includes any or all possible combinations of one or more associated listed items.
[0028] Terms such as "upper", "lower", "left", "right", "front", "back", "thickness", "radial", "axial" and the like that may be used herein to indicate relative positions in space are for the purpose of convenience of description to describe the relationship of one feature relative to another feature as shown in the accompanying drawings, and are not limited to one position or one spatial orientation. It is understood that, depending on the placement of the product, the terms of relative positions in space may be intended to include different orientations in addition to the orientation shown in the drawings, and should not be understood as limiting. Words such as "include" or "comprise" and the like used herein mean that the elements or objects appearing before "include" or "comprises" include the elements or objects listed after "include" or "comprises" and their equivalents, and do not exclude other elements or objects.
[0029] It should be understood that the terms “first”, “second” and similar terms used in the present application specification do not indicate any order, quantity or importance, but are only used to distinguish different components.
[0030] The above are only preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.
Claims
1. A gas turbine engine rotor blade (10), comprising a blade body (12), a root (20), and a connecting section (16) connecting the blade body (12) and the root (20), wherein the root (20) is used to be embedded in an axial groove fixed to a turbine disk (1), and the connecting section (16) comprises a first edge plate (14) for sealing the wheel rim, which protrudes forward from the connecting section (16) along the turbine axial direction. It is characterized in that The extension curve of the first edge plate (14) forms a non-zero angle with the circumferential direction of the turbine disk where it is located.
2. The gas turbine engine rotor blade (10) according to claim 1, It is characterized in that The first edge plate (14) forms a twist angle with the circumferential direction of the turbine disk, so as to form an impeller structure for extracting the sealed gas in the disk cavity when the turbine rotates.
3. A gas turbine engine rotor blade (10) according to claim 1 or 2, It is characterized in that The length of the first edge plate (14) in the circumferential direction of the turbine disk is greater than the length of the connecting section (16).
4. The gas turbine engine rotor blade (10) according to claim 3, It is characterized in that Both ends of the first edge plate (14) in the circumferential direction of the turbine disk respectively extend beyond both ends of the connecting section (16).
5. A gas turbine engine, include: - a plurality of turbine engine rotor blades (10) as claimed in any one of claims 1 to 4, mounted on a turbine disk (1); - an intermediate support (2) having a second edge plate (22) and a third edge plate (23) arranged in the radial direction of the turbine, The first edge plate (14) is inserted between the second edge plate (22) and the third edge plate (23) in the radial direction of the turbine, and a gap exists between the three in the radial direction of the turbine.