A design method for asymmetric guide vanes suitable for vibration reduction of high-pressure turbine rotor blades

Abstract

A design method for asymmetric guide vanes suitable for vibration reduction of high-pressure turbine rotor blades is proposed. The guide vanes within the high-pressure turbine guide vane guide are circumferentially divided into multiple sectors. The single-channel circumferential guide vane spacing angles between adjacent guide vanes within a single sector are equal. The number of guide vanes in two adjacent sectors is different, and the single-channel circumferential guide vane spacing angles between the guide vanes in the two adjacent sectors are different, forming an asymmetric guide vane arrangement structure. By dividing the guide vanes within the high-pressure turbine guide vane guide, the number of guide vanes in each sector and the single-channel circumferential guide vane spacing angles between the guide vanes are designed to form an asymmetric guide vane arrangement. This effectively reduces the excitation amplitude, avoids resonant frequencies, and does not generate new high-energy excitations. Aerodynamic and vibration simulations show that the guide vane wake excitation amplitude is reduced by approximately 75%, and the vibration amplitude of the rotor blades under this frequency excitation is reduced by approximately 82%. The performance cost is a reduction of approximately 0.4% in the high-pressure turbine efficiency, with little impact on the low-pressure turbine.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engine high-pressure turbine guide vane design, and in particular to a method for designing an asymmetric guide vane suitable for high-pressure turbine rotor blade vibration reduction. Background Art

[0002] Turbine components are key to aircraft engines, significantly impacting overall engine performance and reliability. The high-pressure turbine guide vane, located between the combustion chamber and the high-pressure turbine rotor, is a crucial component of the turbine. Its primary function is to rectify, accelerate, and redirect the airflow at the combustion chamber outlet, creating optimal inlet airflow conditions for the turbine rotor. This allows the high-pressure turbine blades to achieve maximum performance while ensuring safe, reliable, and efficient operation of the turbine components.

[0003] The circumferential layout of the guide vanes of the high-pressure turbine guide vanes of aircraft engines is usually symmetrical, which inevitably generates wake excitation of the guide vane blades at several times the frequency. When designing high-pressure turbine blades, while ensuring high aerodynamic performance, it is also necessary to ensure a high strength reserve and lifespan. To do this, it is necessary to avoid the resonant excitation caused by the guide vanes. When it is impossible to avoid resonant excitation due to structural limitations, there are generally two measures that can be taken: one is to increase damping to reduce or limit the amplitude of the blades under resonant excitation, thereby reducing vibration stress; the other is to re-match the design of the guide vanes so that the wake excitation avoids the low-order vibration frequency of the blades, thereby reducing the vibration stress of the blades.

[0004] In the second case of re-matching and designing the guide vanes, a high-pressure turbine moving blade has a low-order resonance frequency under wake excitation. How to design the circumferential spacing angle of the high-pressure guide vanes without changing the guide vane blade shape, number, meridian flow channel size, outlet airflow angle and flow capacity, change the aerodynamic excitation frequency and amplitude distribution generated by the gas after high guide, and avoid the resonance point of the high-pressure turbine moving blades, is a technical problem that needs to be solved in the process of re-matching and designing the guide vanes. Summary of the Invention

[0005] The main purpose of the present invention is to propose an asymmetric guide vane design method suitable for high-pressure turbine rotor blade vibration reduction, aiming to solve the above technical problems.

[0006] To achieve the above-mentioned objectives, the present invention proposes an asymmetric guide vane design method suitable for vibration reduction of high-pressure turbine rotor blades, wherein the guide vanes in the high-pressure turbine guide vane are divided into multiple sectors in the circumferential direction; the single-channel circumferential guide vane spacing angles between adjacent guide vanes in a single sector are equal; the number of guide vanes in two adjacent sectors is different, and the single-channel circumferential guide vane spacing angles between the guide vanes in the two adjacent sectors are different, thereby forming an asymmetric guide vane arrangement structure.

[0007] Preferably, the maximum and minimum values ​​of the single-channel circumferential guide vane spacing angle are designed based on a margin of ±10% of the resonance frequency, wherein the single-channel circumferential guide vane spacing angle within one or more sectors is the maximum value; and the single-channel circumferential guide vane spacing angle within one or more sectors is the minimum value.

[0008] Preferably, the guide vane transition interval angle is designed based on a ±5% margin of the resonance frequency, wherein the single-channel circumferential guide vane interval angle in one or more sectors is the transition interval angle.

[0009] Preferably, the number of channels in the sector corresponding to the transition interval angle of the single-channel circumferential guide vane interval angle is m; the number of channels in the sector corresponding to the maximum value of the single-channel circumferential guide vane interval angle is n; the number of channels in the sector corresponding to the minimum value of the single-channel circumferential guide vane interval angle is q; wherein m, n, and q satisfy: m>n>q.

[0010] Preferably, the full ring of the high-pressure turbine guide vane is divided into six sectors in the circumferential direction, namely the first sector, the second sector, the third sector, the fourth sector, the fifth sector and the sixth sector;

[0011] The single-channel circumferential guide vane interval angles corresponding to the first sector, the third sector, and the fifth sector are minimum values;

[0012] The single-channel circumferential guide vane spacing angle corresponding to the sixth sector is the maximum value;

[0013] There are two values ​​for the transition interval angle, corresponding to the second sector and the fourth sector respectively.

[0014] Preferably, the number of guide vanes in the high-pressure turbine guide vane is 29, which are divided into the six sectors; wherein, the first sector is 2 channels, and the single-channel circumferential guide vane interval angle in the first sector accounts for 1 / 33.4 of the entire ring; the second sector is 9 channels, and the single-channel circumferential guide vane interval angle in the second sector accounts for 1 / 27.7 of the entire ring; the third sector is 2 channels, and the single-channel circumferential guide vane interval angle in the third sector accounts for 1 / 33.4 of the entire ring; the fourth sector is 9 channels, and the single-channel circumferential guide vane interval angle in the fourth sector accounts for 1 / 30.7 of the entire ring; the fifth sector is 2 channels, and the single-channel circumferential guide vane interval angle in the fifth sector accounts for 1 / 33.4 of the entire ring; the sixth sector is 5 channels, and the single-channel circumferential guide vane interval angle in the sixth sector accounts for 1 / 24.7 of the entire ring.

[0015] Due to the adoption of the above technical solution, the beneficial effects of the present invention are as follows:

[0016] The present invention divides the guide vanes in the high-pressure turbine guide vane into multiple sectors in the circumferential direction, and forms an asymmetric guide vane arrangement structure by designing the number of guide vanes in each sector and the single-channel circumferential guide vane spacing angle between the guide vanes. This can effectively reduce the excitation amplitude, avoid the resonant frequency, and not generate new high-energy excitation. Through aerodynamic simulation and vibration simulation, the guide vane wake excitation amplitude is reduced by about 75%, and the vibration amplitude of the moving blades under this frequency excitation is reduced by about 82%. The performance cost is: the high-pressure turbine efficiency is reduced by about 0.4%, and there is basically no impact on the low-pressure turbine. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying any creative work.

[0018] Figure 1 Schematic diagram of circumferentially symmetrical distribution of high-pressure guide blades;

[0019] Figure 2 Schematic diagram of the circumferential asymmetric distribution of high-pressure guide blades in the present invention;

[0020] Figure 3 This is the vibration spectrum of the moving blade under the excitation of the guide vane wake when the high-pressure guide vanes are circumferentially symmetrically distributed;

[0021] Figure 4 This is a vibration spectrum diagram of the moving blade under the excitation of the guide vane wake when the high-pressure guide blades are circumferentially asymmetrically distributed in the present invention. DETAILED DESCRIPTION

[0022] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. All other embodiments obtained by ordinary technicians in this field based on the embodiments of the present invention without making any creative efforts shall fall within the scope of protection of the present invention.

[0023] It should be noted that all directional indications in the embodiments of the present invention (such as up, down, left, right, front, back, etc.) are only used to explain the relative position relationship, movement status, etc. between the various components under a certain specific posture (as shown in the accompanying drawings). If the specific posture changes, the directional indication will also change accordingly.

[0024] In addition, the terms "first," "second," and so on, used in this disclosure are for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly indicating the number of the technical features indicated. Furthermore, the technical solutions of the various embodiments may be combined with one another, but this must be based on the ability of a person of ordinary skill in the art to implement them. If a combination of technical solutions contradicts or cannot be implemented, such combination of technical solutions shall be deemed non-existent and outside the scope of protection claimed by this disclosure.

[0025] Figure 1 The figure shows a schematic diagram of the circumferentially symmetrical distribution of high-pressure guide vanes for the guide vane guide of a certain aircraft engine's high-pressure turbine. The blades number 29 and are arranged symmetrically around the circumference. To address the fourth-order resonance of the high-pressure turbine blades in this aircraft engine caused by the guide vane wake, this paper proposes an asymmetric guide vane design method suitable for vibration reduction of high-pressure turbine rotor blades. By varying the excitation frequency and amplitude of the guide vane wake on the moving blades, the vibration stress level of the moving blades is reduced, thereby improving the blade's operational safety and reliability.

[0026] Specifically, combined Figure 2 The figure shows a schematic diagram of the asymmetric circumferential distribution of high-pressure guide vanes in the present invention. To achieve the goal of changing the excitation frequency and amplitude of the high-turbine wake on the rotor blades and reducing the vibration stress level of the high-vortex rotor blades, an asymmetric guide vane design method suitable for high-pressure turbine rotor blade vibration reduction is proposed, without changing the guide vane blade shape, number, meridian flow channel size, outlet airflow angle, and high-turbine flow capacity. The circumferential distribution of the guide vanes is changed from the prototype's symmetrical distribution, where each channel has the same circumferential spacing angle of (360 / 29) degrees, to a method where the guide vanes within the high-pressure turbine guide vane are circumferentially divided into multiple sectors; the single-channel circumferential guide vane spacing angles between adjacent guide vanes within a single sector are equal; the number of guide vanes within two adjacent sectors is different, and the single-channel circumferential guide vane spacing angles between the guide vanes within the two adjacent sectors are different, forming an asymmetric guide vane arrangement structure.

[0027] In this embodiment, the maximum and minimum values ​​of the single-channel circumferential guide vane spacing angle are designed based on a ±10% margin of the resonant frequency to ensure that the rotor blades avoid the resonant excitation frequency. The single-channel circumferential guide vane spacing angle within one or more sectors is the maximum value; and the single-channel circumferential guide vane spacing angle within one or more sectors is the minimum value.

[0028] In this embodiment, the guide vane transition interval angle is designed based on a ±5% margin of the resonance frequency, wherein the single-channel circumferential guide vane interval angle in one or more sectors is the transition interval angle.

[0029] A higher guide vane number weight is assigned to the sector corresponding to the transition interval angle to form a larger number of channels to ensure higher flow field quality at the guide vane outlet and guarantee turbine performance.

[0030] The sector corresponding to the minimum value of the single-channel circumferential guide vane interval angle is assigned a smaller guide vane number weight, resulting in a smaller number of channels.

[0031] In this embodiment, the number of channels in the sector corresponding to the transition interval angle of the single-channel circumferential guide vane interval angle is m; the number of channels in the sector corresponding to the maximum value of the single-channel circumferential guide vane interval angle is n; the number of channels in the sector corresponding to the minimum value of the single-channel circumferential guide vane interval angle is q; where m, n, and q satisfy: m>n>q.

[0032] Specifically, the full ring of the high-pressure turbine guide vane is divided into six sectors in the circumferential direction, namely the first sector 1, the second sector 2, the third sector 3, the fourth sector 4, the fifth sector 5 and the sixth sector 6; the single-channel circumferential guide vane interval angles corresponding to the first sector 1, the third sector 3, and the fifth sector 5 are minimum values; the single-channel circumferential guide vane interval angle corresponding to the sixth sector 6 is maximum value; the transition interval angle has two values ​​(one is between the design value and the maximum value, and the other is between the design value and the minimum value), and the corresponding sectors are the second sector 2 and the fourth sector 4 respectively.

[0033] Furthermore, the number of guide vanes in the high-pressure turbine guide vane is 29, which are divided into the six sectors; specifically: the first sector 1 has 2 channels, and the single-channel circumferential guide vane interval angle in the first sector 1 accounts for 1 / 33.4 of the entire ring; the second sector 2 has 9 channels, and the single-channel circumferential guide vane interval angle in the second sector 2 accounts for 1 / 27.7 of the entire ring; the third sector 3 has 2 channels, and the single-channel circumferential guide vane interval angle in the third sector 3 accounts for 1 / 33.4 of the entire ring; the fourth sector 4 has 9 channels, and the single-channel circumferential guide vane interval angle in the fourth sector 4 accounts for 1 / 30.7 of the entire ring; the fifth sector 5 has 2 channels, and the single-channel circumferential guide vane interval angle in the fifth sector 5 accounts for 1 / 33.4 of the entire ring; the sixth sector 6 has 5 channels, and the single-channel circumferential guide vane interval angle in the sixth sector 6 accounts for 1 / 24.7 of the entire ring.

[0034] Based on the above limited parameters, three-dimensional flow field calculations were performed. The results show that the amplitude of the K = 29th frequency can be reduced from 16.3kPa to 4.1kPa, a reduction of 75%. The excitation amplitudes of K = 27, 28, 30, 31, and 32nd frequencies do not exceed 10.6kPa. K is a multiple of the fundamental frequency.

[0035] Depend on Figure 3 and Figure 4 It can be seen that the asymmetric guide vane design can effectively reduce the vibration amplitude of the moving blade under the frequency excitation near K=29 by about 82% without increasing additional vibration.

[0036] The above description is only a preferred embodiment of the present invention and does not limit the patent scope of the present invention. Any equivalent structural transformation made by using the contents of the present invention specification and drawings under the inventive concept of the present invention, or directly / indirectly applied in other related technical fields, is included in the patent protection scope of the present invention.

Claims

1. A method for designing asymmetric guide vanes suitable for vibration reduction of high-pressure turbine rotor blades, characterized by: The guide vanes in the high-pressure turbine guide vane are divided into multiple sectors in the circumferential direction; the single-channel circumferential guide vane spacing angles between adjacent guide vanes in a single sector are equal; the number of guide vanes in two adjacent sectors is different, and the single-channel circumferential guide vane spacing angles between the guide vanes in the two adjacent sectors are different, forming an asymmetric guide vane arrangement structure; The high-pressure turbine guide vane ring is divided into six sectors in the circumferential direction, namely the first sector (1), the second sector (2), the third sector (3), the fourth sector (4), the fifth sector (5) and the sixth sector (6); The single-channel circumferential guide vane interval angles corresponding to the first sector (1), the third sector (3), and the fifth sector (5) are minimum values; the single-channel circumferential guide vane interval angle corresponding to the sixth sector (6) is maximum value; the transition interval angle has two values, corresponding to the second sector (2) and the fourth sector (4) respectively; The number of guide vanes in the high-pressure turbine guide vane is 29, divided into the six sectors mentioned above; The first sector (1) has two channels, and the ratio of the circumferential guide vane spacing angle of a single channel in the first sector (1) to the entire ring is 1 / 33.4; The second sector (2) consists of 9 channels, and the ratio of the circumferential guide vane spacing angle of a single channel in the second sector (2) to the entire ring is 1 / 27.7; The third sector (3) has two channels, and the ratio of the circumferential guide vane spacing angle of a single channel in the third sector (3) to the entire ring is 1 / 33.4; The fourth sector (4) has 9 channels, and the ratio of the circumferential guide vane spacing angle of a single channel in the fourth sector (4) to the entire ring is 1 / 30.7; The fifth sector (5) has two channels, and the ratio of the circumferential guide vane spacing angle of a single channel in the fifth sector (5) to the entire ring is 1 / 33.4; The sixth sector (6) has five channels, and the ratio of the circumferential guide vane interval angle of a single channel in the sixth sector (6) to the entire ring is 1 / 24.

7.

2. The asymmetric guide vane design method for high-pressure turbine rotor blade vibration reduction according to claim 1, characterized in that: The maximum and minimum values ​​of the single-channel circumferential guide vane interval angle are designed based on a margin of ±10% of the resonance frequency, wherein the single-channel circumferential guide vane interval angle within one or more sectors is the maximum value; the single-channel circumferential guide vane interval angle within one or more sectors is the minimum value.

3. The asymmetric guide vane design method for high-pressure turbine rotor blade vibration reduction according to claim 2, characterized in that: The guide vane transition interval angle is designed based on the ±5% margin of the resonance frequency, wherein the single-channel circumferential guide vane interval angle in one or more sectors is the transition interval angle.

4. The asymmetric guide vane design method for high-pressure turbine rotor blade vibration reduction according to claim 3, characterized in that: The number of channels in the sector corresponding to the transition interval angle of the single-channel circumferential guide vane is m; the number of channels in the sector corresponding to the maximum value of the single-channel circumferential guide vane interval angle is n; the number of channels in the sector corresponding to the minimum value of the single-channel circumferential guide vane interval angle is q; where m, n, and q satisfy: m>n>q.

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