Aero-engine fan rotor blade and noise reduction structure and test method thereof

Abstract

The invention belongs to the technical field of aero-engine fan design, and particularly relates to an aero-engine fan rotor blade, a noise reduction structure of the aero-engine fan rotor blade and a test method.In the aero-engine fan rotor blade, the front edge of the fan rotor blade is divided into a front edge lower sweepforward area A1, a front edge upper sweepback area A2 and a front edge tip sweepback area A3 according to the height percentage; the height percentage of the high-density polyethylene is 0-40%, 40-85% and 85-100%, and the swept angles of the high-density polyethylene are concentrated at-20 degrees to-35 degrees, 0-10 degrees and-30 degrees to-35 degrees; the rear edge of the fan rotor blade is divided into a rear edge lower sweepback area B1, a rear edge middle sweepback area B2, a rear edge upper transition area B3 and a rear edge tip sweepforward area B4 according to the height percentage, the height percentage of the rear edge lower sweepback area B1, the height percentage of the rear edge middle sweepback area B2, the height percentage of the rear edge upper transition area B3 and the height percentage of the rear edge tip sweepforward area B4 are 0-45%, 45%-60%, 60%-85% and 85%-100%, and the sweep angles are concentrated at 10-30 degrees, 30 degrees, 0-30 degrees and-10-30 degrees.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of fan design of an aero-engine, and particularly relates to an aero-engine fan rotor blade, a noise reduction structure thereof, and a test method. BACKGROUND

[0002] Most of the noise of an airplane comes from an aero-engine, and fan noise accounts for 60-70% of the noise of the aero-engine. More than 70% of the thrust of a high-bypass-ratio aero-engine is generated by a fan. The fan generates a huge thrust, but also generates a huge noise. How to reduce the fan noise becomes a key to airworthiness certification of a large airplane.

[0003] The maximum diameter of a fan component can reach 3.5 meters, and the thrust of an aero-engine can be more than 40 tons. The fan component of a high-bypass-ratio aero-engine is composed of a fan rotor unit body and an intermediate nacelle unit body connected in sequence. Under the premise of ensuring the aerodynamic performance of the fan, reducing the fan noise is an unremitting pursuit in the development of a high-bypass-ratio aero-engine.

[0004] Currently, the secondary noise generated in the propagation process is controlled, the fan is passively noise-reduced from the propagation path in the intermediate nacelle unit body, and the noise reduction mode is single, so it is difficult to effectively control the fan noise.

[0005] The fan rotor blade in the fan rotor unit body is composed of a blade profile stack. The blade profile is composed of a front pressure surface and a rear suction surface. The chord length extends from the leading edge to the trailing edge, and the radial height extends from the blade root to the blade tip. The aerodynamic noise generated by the rotation of the fan rotor blade is the source of the fan noise. However, the current design of the fan rotor blade often only considers the aerodynamic performance of the fan, and does not pay much attention to how to reduce the aerodynamic noise generated by the rotation of the fan rotor blade through the design of the fan rotor blade to actively reduce the noise of the fan from the source.

[0006] Under the premise of not affecting the flow and efficiency of the fan, the fan rotor blade is designed, a general blade configuration technology is developed, a general blade configuration design method is formed, the fan is actively noise-reduced, and the multiple combination noise reduction, superposition noise reduction, and active / passive simultaneous noise reduction are combined, so that the fan noise can be more effectively controlled, and the technology can be effectively applied to a high-bypass-ratio aero-engine, and has great technical potential and commercial value.

[0007] According to the development needs of a high-bypass-ratio aero-engine, the design and experimental technology of the fan needs to consider noise reduction design and noise testing. The existing high-bypass-ratio fan test technology lacks experience in noise reduction evaluation in an open-air condition. Therefore, a multifunctional fan test technology is needed.

[0008] In view of the above, the application is proposed. SUMMARY

[0009] The purpose of this application is to provide an aero-engine fan rotor blade and its noise reduction structure and test method to overcome or mitigate at least one of the known technical defects.

[0010] The technical solution of this application is:

[0011] An aircraft engine fan rotor blade, comprising:

[0012] The leading edge of the fan rotor blades is divided into three regions based on height percentage: the lower leading edge swept region A1, the upper leading edge swept region A2, and the leading edge tip swept region A3. The height percentages are 0-40%, 40%-85%, and 85%-100%, respectively, and the sweep angles are concentrated in -20°~-35°, 0°~10°, and -30°~-35°.

[0013] The trailing edge of the fan rotor blades is divided into the lower trailing edge sweeping area B1, the middle trailing edge sweeping area B2, the upper trailing edge transition area B3, and the tip trailing edge forward sweeping area B4 according to the height percentage, with the height percentages being 0-45%, 45%-60%, 60%-85%, and 85%-100%, respectively. The sweep angles are concentrated in 10°~30°, 30°, 0°~30°, and -10°~-30°.

[0014] In the aforementioned aero-engine fan rotor blades, CH1 / CH2 = 0.65~0.8, where CH1 is the chord length of the tip region and CH2 is the chord length of the middle region.

[0015] A noise reduction structure for an aircraft engine fan includes a fan rotor unit and an intermediate casing unit connected in sequence.

[0016] The fan rotor blades in the fan rotor unit are the aforementioned aero-engine fan rotor blades M1;

[0017] The intermediate casing unit is equipped with an acoustic liner S1, and the outlet stator blades in the intermediate casing unit are swept fan stator blades M2 with a sweep angle controlled between 30° and 45°, so that the sound waves propagating through the path are perpendicularly incident on the acoustic liner S1.

[0018] In the aforementioned aircraft engine fan noise reduction structure, an acoustic liner S2 is installed at the engine nacelle outlet.

[0019] A test method for an aircraft engine fan, used to evaluate the noise reduction effect of the aforementioned aircraft engine fan noise reduction structure, includes:

[0020] Based on the fan unit design, noise reduction test specimens were constructed, including control group specimens and test group specimens;

[0021] In the control group test piece, the fan rotor blades in the fan rotor unit body adopt the current common configuration of fan rotor blade Z1, and the outlet stator blades in the intermediate casing unit body adopt the current common configuration of stator blade Z2.

[0022] The fan rotor blades in the fan rotor unit of the test piece adopt the aero-engine fan rotor blade M1 as described in claim 2, and the outlet stator blades in the intermediate casing unit adopt swept fan stator blades M2.

[0023] An acoustic sensor P2 and a far-field microphone P1 were added after the stator blade at the outlet of the intermediate casing unit of the control group test piece to measure the noise reduction effect of the control group test piece under different working conditions.

[0024] An acoustic sensor P2 and a far-field microphone P1 were added after the stator blade at the outlet of the intermediate casing unit of the test piece to measure the noise reduction effect of the test piece under different working conditions.

[0025] The noise reduction effect of the experimental group test and the control group test was compared to evaluate the noise reduction effect of the aircraft engine fan noise reduction structure. Attached Figure Description

[0026] Figure 1 This is a schematic diagram of the configuration of the aero-engine fan rotor blades provided in an embodiment of this application;

[0027] Figure 2 This is a distribution diagram of the leading and trailing edge sweep angles of the aero-engine fan rotor blades provided in an embodiment of this application;

[0028] Figure 3 This is a schematic diagram of the noise reduction structure for an aircraft engine fan provided in an embodiment of this application;

[0029] Figure 4 This is a schematic diagram of the test method for an aircraft engine fan provided in an embodiment of this application.

[0030] To better illustrate this embodiment, some content in the accompanying drawings may be omitted, enlarged, or reduced. They are for illustrative purposes only and should not be construed as limiting the scope of this application. Detailed Implementation

[0031] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings, and other related parts can be referred to the general design.

[0032] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The word "comprising" as used in this application description indicates that the concept preceding the word encompasses the concepts listed following the word and their equivalents, without excluding other related concepts.

[0033] Furthermore, the terms indicating location used in the description of this application are only used to indicate relative directions or positional relationships. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. It should also be noted that, unless otherwise explicitly specified and limited, terms such as "installation" and "connection" used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection or a detachable connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0034] A type of aircraft engine fan rotor blade, with its leading edge tip tapering backward along the fan axis, such as... Figure 1 As shown.

[0035] The leading edge of the fan rotor blades is divided into three regions based on height percentage: the lower leading edge swept region A1, the upper leading edge swept region A2, and the leading edge tip swept region A3. The height percentages are 0-40%, 40%-85%, and 85%-100%, respectively. The sweep angles are concentrated in -20°~-35°, 0°~10°, and -30°~-35°, with smooth transitions between them.

[0036] The trailing edge of the fan rotor blades is divided into the lower trailing edge sweeping area B1, the middle trailing edge sweeping area B2, the upper trailing edge transition area B3, and the tip trailing edge forward sweeping area B4 according to the height percentage, with the height percentages being 0-45%, 45%-60%, 60%-85%, and 85%-100%, respectively. The sweep angles are concentrated in 10°~30°, 30°, 0°~30°, and -10°~-30°, with smooth transitions between them.

[0037] The sweep angle distribution of the following regions can be specifically designed: lower leading edge sweep area A1, upper leading edge sweep area A2, leading edge tip sweep area A3, lower trailing edge sweep area B1, middle trailing edge sweep area B2, upper trailing edge transition area B3, and trailing edge tip sweep area B4. Figure 2 As shown.

[0038] In the fan rotor blades, CH1 / CH2 = 0.65~0.8, where CH1 is the chord length of the blade tip region, that is, the chord length at 100% of the blade height; CH2 is the chord length of the blade middle region, that is, the chord length at 50% of the blade height.

[0039] The aero-engine fan rotor blades disclosed in the above embodiments, by changing the sweep distribution ratio of the leading and trailing edge regions, are adjusted to a fan noise reduction configuration. The swept A1 region and swept A2 region can be used to reduce the Mach number of the fan inlet airflow, which can significantly reduce fan noise without affecting the aerodynamic performance of the fan.

[0040] Experimental data show that the aero-engine fan rotor blades disclosed in the above embodiments can reduce noise by more than 10 decibels in specific frequency bands, such as the noise range of 2000-3000Hz, compared with the current conventional fan rotor blades, and can improve the aerodynamic performance of the fan rotor blades to a certain extent. They can be used as a general and structured fan rotor blade configuration.

[0041] A noise reduction structure for an aircraft engine fan includes a fan rotor unit and an intermediate casing unit connected in sequence, such as... Figure 3 As shown.

[0042] The fan rotor blades in the fan rotor unit adopt the aero-engine fan rotor blades M1 disclosed in the above embodiments to actively reduce noise from the source.

[0043] An acoustic liner S1 is provided in the intermediate casing unit to absorb sound along the propagation path and passively reduce noise from the fan. An acoustic liner S2 can also be provided at the engine nacelle outlet to further absorb sound along the propagation path and passively reduce noise from the fan.

[0044] The outlet stator blades in the intermediate casing unit adopt swept fan stator blades M2 with a sweep angle controlled between 30° and 45°. This can cause the sound waves on the propagation path to be incident perpendicularly towards the acoustic liner S1, which can reduce the secondary noise generated during the propagation process and achieve the best noise reduction effect. Experimental data shows that this design can reduce fan noise by about 1 decibel.

[0045] The aircraft engine fan noise reduction structure disclosed in the above embodiments integrates aerodynamic configuration, noise reduction, and intermediate casing load-bearing frame design. It adopts the aircraft engine fan rotor blade M1, acoustic liner S1, and swept fan stator blade M2 ​​disclosed in the above embodiments to construct a multi-layer superimposed noise reduction mode. It comprehensively reduces noise from the noise source and propagation path, combining active and passive noise reduction. It can effectively reduce fan noise while ensuring the aerodynamic performance of the fan.

[0046] A test method for an aircraft engine fan is provided to evaluate the noise reduction effect of the aircraft engine fan noise reduction structure disclosed in the above embodiments.

[0047] Based on the fan unit design and aerodynamic testing technology, noise reduction test specimens were constructed, including control group test specimens and test group test specimens.

[0048] In the control group test piece, the fan rotor blades in the fan rotor unit adopt the currently common configuration of fan rotor blades Z1, and the outlet stator blades in the intermediate casing unit adopt the currently common configuration of stator blades Z2. An acoustic liner S1 can be installed in the intermediate casing unit.

[0049] In the test piece, the fan rotor blades in the fan rotor unit are the aero-engine fan rotor blades M1 disclosed in the above embodiment, and the outlet stator blades in the intermediate casing unit are swept fan stator blades M2. An acoustic liner S1 can be provided in the intermediate casing unit.

[0050] The test set of test pieces can be obtained by replacing the ordinary configuration fan rotor blades Z1 and stator blades Z2 in the test piece with the aero-engine fan rotor blades M1 and swept-back fan stator blades M2 disclosed in the above embodiments, such as... Figure 4 As shown.

[0051] An acoustic sensor P2 and a far-field microphone P1 were added after the stator blade at the outlet of the intermediate casing unit of the control group test piece. The far-field microphone P1 was more than 30 meters away from the rotor blade inlet of the control group test piece. The noise reduction effect of the control group test piece under different operating conditions was measured, that is, the difference between the noise intensity detected by the acoustic sensor P2 and the noise intensity emitted by the far-field microphone P1.

[0052] The test group was tested, referring to the control group test. An acoustic sensor P2 was added after the stator blade at the outlet of the intermediate casing unit of the test piece in the test group, and a far-field microphone P1 was set. The far-field microphone P1 was more than 30 meters away from the rotor blade inlet of the test piece in the test group. The noise reduction effect of the test piece in the test group under different working conditions was measured.

[0053] The noise reduction effect of the experimental group test and the control group test was compared to evaluate the noise reduction effect of the aircraft engine fan noise reduction structure.

[0054] When conducting experimental group and control group tests, the aerodynamic performance of the fan can be tested simultaneously.

[0055] The test method for aero-engine fans disclosed in the above embodiments, based on aerodynamic test technology, proposes a noise measurement scheme for secondary comparison of fan rotor and stator blades with noise reduction configuration, and develops a multifunctional fan test scheme that can conduct aerodynamic performance and noise reduction tests simultaneously. It is convenient and fast to conduct test group tests and control group tests simply by changing different fan rotor blade configurations and stator blade configurations.

[0056] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the scope of protection of this application.

Claims

1. An aero-engine fan rotor blade, characterized in that, include: The leading edge of the fan rotor blades is divided into three regions according to the height percentage: the lower leading edge swept area A1, the upper leading edge swept area A2, and the leading edge tip swept area A3, with the height percentages being 0-40%, 40%-85%, and 85%-100%, respectively. The sweep angles are concentrated in -20°~-35°, 0°~10°, and -30°~-35°, respectively. The trailing edge of the fan rotor blades is divided into the lower trailing edge sweeping area B1, the middle trailing edge sweeping area B2, the upper trailing edge transition area B3, and the tip trailing edge forward sweeping area B4 according to the height percentage, with the height percentages being 0-45%, 45%-60%, 60%-85%, and 85%-100%, respectively. The sweep angles are concentrated in 10°~30°, 30°, 0°~30°, and -10°~-30°.

2. The aero-engine fan rotor blade according to claim 1, characterized in that, In the fan rotor blades, CH1 / CH2 = 0.65~0.8, where CH1 is the chord length of the tip region and CH2 is the chord length of the middle region.

3. A noise reduction structure for an aircraft engine fan, characterized in that, Includes a fan rotor unit and an intermediate casing unit connected in sequence; The fan rotor blades in the fan rotor unit are the aero-engine fan rotor blades M1 as described in claim 2; The intermediate casing unit is equipped with an acoustic liner S1, and the outlet stator blades in the intermediate casing unit are swept fan stator blades M2 with a sweep angle controlled between 30° and 45°, so that the sound waves propagating through the path are perpendicularly incident on the acoustic liner S1.

4. The aircraft engine fan noise reduction structure according to claim 3, characterized in that, The engine nacelle outlet is equipped with an acoustic liner S2.

5. A test method for an aircraft engine fan, used to evaluate the noise reduction effect of the aircraft engine fan noise reduction structure of claim 4, characterized in that, include: Based on the fan unit design, noise reduction test specimens were constructed, including control group specimens and test group specimens; In the control group test piece, the fan rotor blades in the fan rotor unit body adopt the current common configuration of fan rotor blade Z1, and the outlet stator blades in the intermediate casing unit body adopt the current common configuration of stator blade Z2. The fan rotor blades in the fan rotor unit of the test piece adopt the aero-engine fan rotor blade M1 as described in claim 2, and the outlet stator blades in the intermediate casing unit adopt swept fan stator blades M2. An acoustic sensor P2 and a far-field microphone P1 were added after the stator blade at the outlet of the intermediate casing unit of the control group test piece to measure the noise reduction effect of the control group test piece under different working conditions. An acoustic sensor P2 and a far-field microphone P1 were added after the stator blade at the outlet of the intermediate casing unit of the test piece to measure the noise reduction effect of the test piece under different working conditions. The noise reduction effect of the experimental group test and the control group test was compared to evaluate the noise reduction effect of the aircraft engine fan noise reduction structure.

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