A rapid evaluation method for fan shock wave-related noise

Abstract

This application belongs to the field of aero-engine noise prediction and control, and particularly relates to a method for rapidly evaluating fan shock-wave-related noise. The method comprises the following steps: Step S1: obtaining aerodynamic parameters and geometric parameters of the fan aerodynamic design results; Step S2: invoking a shock wave model and, based on the aerodynamic parameters and geometric parameters, calculating the pressure change caused by the shock wave; and Step S3: invoking a shock wave noise propagation model and, based on the aerodynamic parameters and geometric parameters, calculating the propagation of the shock-wave-related noise along the fan path. This method utilizes minimal computational effort and is time-efficient. It can rapidly evaluate the intensity of shock-wave-related noise during the fan aerodynamic design phase, optimize design parameters, and achieve a low-noise fan design.

Description

Technical Field

[0001] The present application relates to the field of aero-engine noise prediction and control, and in particular to a method for rapidly evaluating fan shock wave-related noise. Background Art

[0002] As the bypass ratio of civil turbofan engines continues to increase, the jet noise is significantly reduced, and the proportion of fan noise increases. During engine takeoff conditions, the fan shock wave-related noise is the most prominent in the total noise. Therefore, the evaluation of shock wave-related noise is particularly important and remains an important focus of civil aviation.

[0003] Existing noise prediction methods related to fan shock waves include analytical methods based on weak shock wave theory and numerical methods based on computational fluid dynamics (CFD). In terms of analytical methods, Fisher et al. (Fisher MJ, Tester BJ, Schwaller PJ G. Supersonic fan tone noise prediction [R]. AIAA-98-2249.) proposed a time domain numerical solution (TDNS) for shock wave propagation in hard-walled ducts based on the Riemann equation of weak shock waves. Subsequently, McAlpine and Fisher (McAlpine A., Fisher MJ, On the prediction of "buzz-saw" noise in aero-engine inlet ducts [J], Journal of Sound and Vibration, 248 (1) (2001): 123-149.) proposed a frequency domain numerical solution (FDNS) that can take into account the influence of soft wall surfaces such as acoustic lining. In terms of numerical methods, Rolls-Royce and Boeing jointly developed a complete CFD-CAA hybrid method (Coupland J, etc. Demonstration of a CFD-CAA Methodology to predict Buzz-saw Noise Propagation to the Aircraft[R]. AIAA-98-2249.) to predict the shock wave noise of the engine fan.

[0004] The existing assessment of shock-related noise has the following shortcomings: 1. In terms of analytical methods, both the TDNS and FDNS methods require a given initial shock wave intensity and only predict the noise propagation process. They cannot evaluate the sound source intensity of shock-related noise and cannot consider the influence of fan blade sweep.

[0005] 2. Although numerical methods can obtain the sound source information of shock wave related noise, due to the large amount of calculation required and the long time period, it is not conducive to the evaluation of shock wave related noise during the fan aerodynamic design stage. Summary of the Invention

[0006] To address the above issues, this application provides a method for rapidly evaluating fan shock wave-related noise, including:

[0007] Step S1: Obtaining aerodynamic parameters and geometric parameters of fan aerodynamic design results;

[0008] Step S2: calling the shock wave model and calculating the pressure change caused by the shock wave based on the aerodynamic parameters and the geometric parameters;

[0009] Step S3: calling the shock wave noise propagation model, and calculating the propagation of shock wave related noise along the path based on the aerodynamic parameters and the geometric parameters.

[0010] Preferably, before step S2, coordinate transformation is performed on the geometric parameters and the aerodynamic parameters are corrected.

[0011] Preferably, correcting the aerodynamic parameters includes correcting the blade profile inlet relative Mach number using the sweep angle, and the corrected inlet relative Mach number is the initial inlet relative Mach number multiplied by the cosine value of the sweep angle.

[0012] Preferably, performing coordinate transformation on the geometric parameters includes transforming streamline surface blade profile data into R-θ coordinate blade profile data.

[0013] Preferably, the aerodynamic parameters include: an inlet airflow angle β1 and a relative Mach number M1; the geometric parameters include streamline profile coordinates generated by blade shape design; and a sweep angle λ.

[0014] Preferably, the shock wave model includes a normal shock wave model, a single shock wave model or a double shock wave model.

[0015] Preferably, the calculation method for calculating the pressure change caused by the shock wave using the double shock wave model is:

[0016]

[0017] Among them, ψ Sis the average Mach angle of the detached shock wave, which is the Mach angle at the intersection of the sound velocity line behind the wave and the detached shock wave; γ is the specific heat ratio, M 1S is the Mach number of the blade front flow M1 and the Mach number of point B M B The average value of the Mach number at point B can be obtained from the incoming flow Mach number M1 and the turning angle μ from the leading edge of the blade to point B. B Determined by iteratively solving the Prandtl-Meyer formula; Point B is the point where the vertical track center streamline of the blade leading edge falls on the adjacent blade;

[0018] The static pressure ratio after the shock wave is P2 / P1, where P1 is the static pressure before the shock wave and P2 is the static pressure after the shock wave:

[0019]

[0020] The normal Mach number M 1s,norm =M 1s sinψ S .

[0021] Preferably, the calculation formula for calculating the propagation of shock wave related noise along the path includes:

[0022]

[0023] in p0 is the reference pressure, T is the “flight time”, τ is the blade consistency, λ is the wavelength of the sound wave, c0 is the speed of sound, t is the propagation time, z is the axial distance of propagation, M z is the axial Mach number of the airflow, where T is calculated as:

[0024]

[0025] Where α is the angle between the shock wave and the tangent line of the blade centerline at the leading edge of the blade. β1 is the angle between M1 and M z The vector angle.

[0026] This application has a small amount of calculation and is short in time. It can quickly evaluate the noise intensity related to the shock wave during the fan aerodynamic design stage, optimize the design parameters, and make the designed fan have a lower noise level. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 Schematic diagram of the double shock wave model.

[0028] Figure 2 Schematic diagram of the shock wave propagation model.

[0029] Figure 3 This is an example of the noise sound pressure level in front of the fan calculated by the present invention. The figure also shows the evaluation results using the CFD numerical method. DETAILED DESCRIPTION

[0030] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the implementation of this application will be described in more detail below in conjunction with the drawings in the implementation of this application. In the drawings, the same or similar numbers throughout represent the same or similar elements or elements with the same or similar functions. The described implementation is a part of the implementation of this application, not all of the implementations. The implementation described below with reference to the drawings is exemplary and is intended to be used to explain this application, and should not be understood as a limitation on this application. Based on the implementation in this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The implementation of this application is described in detail below in conjunction with the drawings.

[0031] The basic principle of the present invention is to use the shock wave model formed in the two-dimensional forward problem research process to calculate the pressure change when passing through the shock wave based on the results of flow calculation and blade shape design during the fan / compressor aerodynamic design, and then use the analytical model of shock wave propagation to calculate the propagation of noise along the way, so that the evaluation of shock wave-related noise can be quickly completed during the aerodynamic design process.

[0032] The present invention provides a rapid assessment method for shock wave-related noise applicable to the fan aerodynamic scheme design stage, which mainly includes the following steps:

[0033] This application provides a method for quickly evaluating fan shock wave-related noise, including:

[0034] Step S1: Obtaining aerodynamic parameters and geometric parameters from the fan aerodynamic design results. The data read in are derived from the results of the flow field calculation and blade shape design during the fan aerodynamic design. The aerodynamic parameters read in include the inlet airflow angle and relative Mach number, and the geometric parameters include the streamline profile coordinates (Cartesian coordinates) and sweep angle generated by the blade shape design. (Definitions are provided in: Smith, L.H. and Yeh, H., 1963, “Sweep and Dihedral Effect in Axial FloTurbomachinery”, ASME Journal of Basic Engineering [J], Vol. 85.)

[0035] Step S2: calling the shock wave model and calculating the pressure change caused by the shock wave based on the aerodynamic parameters and the geometric parameters;

[0036] Step S3: calling the shock wave noise propagation model, and calculating the propagation of shock wave related noise along the path based on the aerodynamic parameters and the geometric parameters.

[0037] Preferably, before step S2, coordinate transformation is performed on the geometric parameters and the aerodynamic parameters are corrected.

[0038] Preferably, correcting the aerodynamic parameters includes correcting the blade profile inlet relative Mach number using the sweep angle, and the corrected inlet relative Mach number is the initial inlet relative Mach number multiplied by the cosine value of the sweep angle.

[0039] Preferably, performing coordinate transformation on the geometric parameters includes transforming streamline surface blade profile data into R-θ coordinate blade profile data.

[0040] Preferably, the shock wave model includes a normal shock wave model, a single shock wave model or a double shock wave model.

[0041] Preferably, in order to calculate the pressure change caused by the shock wave, a normal shock wave model, a single shock wave model or a double shock wave model can be used. Here is a schematic diagram of an embodiment of the double shock wave model, as shown in FIG. Figure 1 As shown in the figure, the Mach angle at the intersection of the post-wave sonic line and the detached shock wave is taken as the average Mach angle of the detached shock wave, and the calculation formula is:

[0042]

[0043] Among them, ψ S is the average Mach angle of the detached shock wave, which is the Mach angle at the intersection of the sound velocity line behind the wave and the detached shock wave; γ is the specific heat ratio, M 1S is the Mach number of the blade front flow M1 and the Mach number of point B M B The average value of the Mach number at point B can be obtained from the incoming flow Mach number M1 and the turning angle μ from the leading edge of the blade to point B. B Determined by iteratively solving the Prandtl-Meyer formula; Point B is the point where the vertical track center streamline of the blade leading edge falls on the adjacent blade;

[0044] The static pressure ratio after the shock wave is P2 / P1, where P1 is the static pressure before the shock wave and P2 is the static pressure after the shock wave:

[0045]

[0046] The normal Mach number M 1s,norm =M 1s sinψ S .

[0047] Preferably, the calculation formula for calculating the propagation of shock wave related noise along the path includes:

[0048]

[0049] in p0 is the reference pressure, T is the “flight time”, τ is the blade consistency, λ is the wavelength of the sound wave, c0 is the speed of sound, t is the propagation time, z is the axial distance of propagation, M z is the axial Mach number of the airflow, where T is calculated as:

[0050]

[0051] Where α is the angle between the shock wave and the tangent line of the blade centerline at the leading edge of the blade. β1 is the angle between M1 and M z The vector angle of Figure 2 shown.

[0052] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A method for rapid assessment of fan shock wave related noise, characterized in that: include: Step S1: Obtaining aerodynamic parameters and geometric parameters of fan aerodynamic design results; Step S2: calling the shock wave model and calculating the pressure change caused by the shock wave based on the aerodynamic parameters and the geometric parameters; Step S3: calling the shock wave noise propagation model, and calculating the propagation of shock wave related noise along the path based on the aerodynamic parameters and the geometric parameters; The calculation formula for calculating the propagation of shock wave related noise along the path includes: in, p0 is the reference pressure, T is the flight time, and λ is the wavelength of the sound wave. The calculation formula for T is: Among them, ɑ is the angle between the shock wave and the tangent line of the blade centerline at the leading edge of the blade, and β1 is the angle between M1 and M z The vector angle, τ is the blade density, t is the propagation time, z is the axial distance of propagation, c0 is the speed of sound, M z is the axial Mach number of the airflow, P1 is the static pressure before the shock wave, P2 is the static pressure after the shock wave, and M1 is the incoming flow Mach number.

2. The method for rapid evaluation of fan shock wave related noise according to claim 1, characterized in that: Before step S2, coordinate transformation is performed on the geometric parameters and the aerodynamic parameters are corrected.

3. The method for rapid evaluation of fan shock wave related noise according to claim 2, wherein: Correcting the aerodynamic parameters includes correcting the blade profile inlet relative Mach number using the sweep angle, wherein the corrected inlet relative Mach number is the initial inlet relative Mach number multiplied by the cosine value of the sweep angle.

4. The method for rapid evaluation of fan shock wave related noise according to claim 1, wherein: The coordinate conversion of the geometric parameters includes converting the streamline surface blade profile data into R-θ coordinate blade profile data.

5. The method for rapid evaluation of fan shock wave related noise according to claim 1, wherein: Aerodynamic parameters include: inlet airflow angle β1 and relative Mach number M1; geometric parameters include streamline blade profile coordinates generated by blade shape design; sweep angle λ.

6. The method for rapid evaluation of fan shock wave related noise according to claim 1, wherein: The shock wave model includes a normal shock wave model, a single shock wave model or a double shock wave model.

7. The method for rapid evaluation of fan shock wave related noise according to claim 6, characterized in that: The calculation method for calculating the pressure change caused by the shock wave using the double shock wave model is: Among them, ψ S is the average Mach angle of the detached shock wave, which is the Mach angle at the intersection of the sound velocity line behind the wave and the detached shock wave; γ is the specific heat ratio, M 1S is the Mach number of the blade front flow M1 and the Mach number of point B M B The average value of the Mach number at point B can be obtained from the incoming flow Mach number M1 and the turning angle μ from the leading edge of the blade to point B. B Determined by iteratively solving the Prandtl-Meyer formula; Point B is the point where the vertical track center streamline of the blade leading edge falls on the adjacent blade; Static pressure ratio P2 / P1 after shock wave: The normal Mach number M 1s,norm =M 1s sinψ S .

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