A method for calculating unsteady flow and noise of a high-speed aircraft

Abstract

This invention discloses a method for calculating unsteady flow and noise in high-speed aircraft, relating to the field of aircraft design. The method includes: S1, selecting reference physical quantities based on the actual flight environment of the high-speed aircraft, and obtaining corresponding dimensionless physical quantities through these reference physical quantities; S2, defining the control equations for calculating unsteady flow and noise in the high-speed aircraft based on the dimensionless physical quantities, and completing the dimensionless transformation of the control equations; S3, optimizing the computational region and computational network; S4, setting the cavity wall as the solid boundary of the flow field, and calculating the sound pressure level at each measurement point at each frequency under corresponding speed conditions in the high-speed aircraft; S5, comparing and correcting the sound pressure data and spectrum data calculated in S4 with wind tunnel test results until the calculation results stabilize; S6, characterizing the flow topology and vortex morphology near the cavity wall of the aircraft using the streamline morphology corresponding to the calculation results in S5.

Description

Technical Field

[0001] This invention relates to the field of aircraft design. More specifically, this invention relates to a method for calculating unsteady flow and noise in high-speed aircraft. Background Technology

[0002] High-speed aircraft face complex flow-induced vibration and noise environments during their development. To shorten the development cycle of high-speed aircraft, it is necessary to conduct research on high-precision calculation methods and analysis models for complex flow and noise in high-speed environments, as well as fluid-structure-acoustic coupling mechanisms and flow-induced oscillation acoustic modal prediction models. This research aims to fill the gaps in our understanding of the fluid-structure-acoustic coupling mechanisms and their effects on aircraft walls or cavities under high-speed conditions. Therefore, it is essential to establish a high-precision calculation method for unsteady flow and noise during high-speed aircraft flight, analyze the flow topology in typical aircraft configurations, including circulation, flow vortices, horseshoe vortices, sidewall flow vortices, and tornado structures within cavities, analyze the main vortex morphology, and explore the dominant vortex structures on the aircraft surface and within typical structures at different speeds. Summary of the Invention

[0003] One object of the present invention is to solve at least the above-mentioned problems and / or defects, and to provide at least the advantages described below.

[0004] To achieve these objectives and other advantages of the present invention, a method for calculating unsteady flow and noise in a high-speed aircraft is provided, comprising:

[0005] S1. Select reference physical quantities based on the actual flight environment of high-speed aircraft, and obtain the corresponding dimensionless physical quantities through the reference physical quantities;

[0006] S2. Based on dimensionless physical quantities, define the control equations for unsteady flow and noise calculation of high-speed aircraft, and complete the dimensionless transformation of the control equations.

[0007] S3. Perform three-dimensional numerical simulation calculations on the cavity structure in the aircraft based on the control equations, and optimize the computational region and computational network during the calculation process;

[0008] S4. In the cavity structure of the aircraft, the cavity wall is assumed to be the solid boundary of the flow field. The FW-H equation for predicting flow-induced noise is simplified, and the simplified formula is integrated along the cavity wall based on the Green's function. Then, the total sound pressure level at each measuring point at each frequency under the corresponding speed condition in the high-speed aircraft is obtained. SPL ( f The following formula is used to obtain:

[0009]

[0010] In the above formula, pref For reference pressure, P( f The sound pressure amplitude is obtained by performing a Fast Fourier Transform on the sound pressure time series at the measurement point location. ;

[0011] S5. Compare and correct the sound pressure data and spectrum data calculated in S4 with the wind tunnel test results until the calculation results are stable.

[0012] S6. The streamline morphology corresponding to the calculation results of S5 is used to characterize the flow topology and vortex morphology near the wall of the aircraft cavity structure.

[0013] Preferably, in S1, the reference physical quantity includes: reference length. Reference speed Reference temperature Reference density Reference viscous dynamic coefficient The relevant dimensionless physical quantities include: coordinate components in three directions under an orthogonal rectangular coordinate system. Velocity components in three directions in an orthogonal rectangular coordinate system Pressure p, temperature T, total energy E, density ρ, speed of sound a, Reynolds number Re, Mach number Ma, Prandtl number P r Time t, static temperature T m , Kinetic viscosity coefficient µ.

[0014] Preferably, in S2, the dimensionless governing equations are characterized by the following equation:

[0015]

[0016] In the above formula, U is a conserved variable, x, y, and z are the coordinate components in three directions under an orthogonal rectangular coordinate system, and F... v G v H v All are non-viscous flux, (F v G v H v () represents viscous flux.

[0017] Preferably, in S3, the three-dimensional numerical simulation calculation uses a simplified model;

[0018] Among them, the optimization of the computational grid refers to: using a structured stretched grid, with the grid being denser near the wall and gradually stretched in the far field region;

[0019] The optimization of the computational region refers to: taking the depth D of the aircraft cavity structure as the reference, setting the upper corner of the leading edge of the aircraft cavity structure as the origin of the coordinate system, then in the three directions of x, y, and z, the range of the computational region is limited to: −16D≤x≤20D, −1D≤y≤12D and −3D≤z≤5D, and the aircraft cavity wall is set as an insulating wall.

[0020] Preferably, in S4, the total sound pressure level is... SPL ( f The process for obtaining ) is as follows:

[0021] S40. When only considering boundary layer noise generated by wall pressure fluctuations, the sound source integral surface is also selected only from the solid wall of the cavity. f ca， The simplified form of the FW-H equation is represented by the following equation:

[0022]

[0023] In the above formula, c The velocity of sound in the far field of the flow field. p For far-field sound pressure, t For time, u n Let be the fluid velocity component perpendicular to the given sound source surface. u j For x i Fluid velocity components in the direction, v n For the corresponding x i The tangential fluid component of the sound source surface in the direction of the sound source. P ij The stress tensor defined for Lighthill For the Dirac function, f ca The solid wall of the cavity. n j For x j The unit normal vector in the direction;

[0024] S41. Based on the Green's function, the simplified expression of S40 is obtained by integrating along the cavity wall:

[0025]

[0026] In the above formula, ρ 0 represents the incoming flow density. S ca The area of ​​the cavity wall is 1. ret Where p is the net flux, p is the far-field sound pressure, and R is the gas constant. p 0 represents the incoming static pressure, and... NThis represents the total number of time series data collection points. p s,i The pressure value at the sampling point. v s It is the wall motion speed. n It is the outward normal vector of the wall, and t is time. dS The area of ​​the infinitesimal element;

[0027] S42. Based on the sound pressure time series at the measurement point location, and the sound pressure amplitude P after the fast Fourier transform, f This allows us to obtain the total sound pressure level at each measurement point at each frequency. SPL ( f ).

[0028] This invention offers at least the following advantages: Addressing the challenges of high-speed aircraft development, such as the need for more precise calculations of high-speed, complex flow-induced vibrations and noise environments, this invention, based on the real flight environment of high-speed aircraft, defines and dimensionlessly renders the governing equations for high-precision calculations of unsteady flow and noise in high-speed aircraft. It also designs the calculation region and mesh for typical high-speed aircraft configurations. Furthermore, it obtains the total sound pressure level at key points under typical speed conditions and compares and corrects this level with wind tunnel test results. Finally, it interprets the wall flow state of typical aircraft structures using flow results, ultimately summarizing a high-precision calculation method for unsteady flow and noise in high-speed aircraft.

[0029] Other advantages, objectives and features of the present invention will become apparent in part from the following description, and in part from those skilled in the art through study and practice of the invention. Attached Figure Description

[0030] Figure 1 This is a simplified model diagram used in the present invention for high-precision calculation of unsteady flow and noise in high-speed aircraft;

[0031] Figure 2 This is the mesh configuration of the central cross-section of the cavity structure of the aircraft of the present invention;

[0032] Figure 3 Comparison of total sound pressure level calculation and experimental results at measuring points on the front and rear walls and bottom of the central section of the cavity structure of the aircraft of this invention;

[0033] Figure 4 This invention describes the time-averaged streamline morphology of the spacecraft cavity structure wall at a Mach number of 2.0.

[0034] Figure 5 This invention provides the time-averaged streamline morphology of the spacecraft cavity structure wall at a Mach number of 2.5.

[0035] Figure 6This invention describes the time-averaged streamline morphology of the spacecraft cavity structure wall at a Mach number of 3.0.

[0036] Figure 7 This invention describes the time-averaged streamline morphology of the spacecraft cavity structure wall at a Mach number of 3.5. Detailed Implementation

[0037] The present invention will now be described in further detail with reference to the accompanying drawings, so that those skilled in the art can implement it based on the description.

[0038] Step 1: Based on the real flight environment of high-speed aircraft, the control equations for the high-precision calculation method of unsteady flow and noise of high-speed aircraft are defined and dimensionless.

[0039] To address the unsteady flow and noise issues on the walls or characteristic structures of high-speed aircraft, this invention designs a three-dimensional numerical simulation method based on the cavity-type structure most widely used in aircraft. The dimensions of the aircraft cavity structure are referenced to the C201 standard model, with a length, width, and depth ratio of L:W:D=6:2:1. Based on the actual flight environment of high-speed aircraft, the control equations for the high-precision calculation method of unsteady flow and noise of high-speed aircraft are defined and dimensionless.

[0040] The governing equations are the three-dimensional Navier-Stokes equations without external forces. During numerical solutions, the equations need to be dimensionless. By selecting reference physical quantities, the relevant dimensionless physical quantities are as follows:

[0041]

[0042] In the above formula, the superscript "∗" indicates a dimensional physical quantity, and the top "—" indicates a selected reference physical quantity. ρ , T , a and µ These represent the density, temperature, sound velocity, and dynamic viscosity at infinity, respectively. L is the characteristic length, which needs to be set according to the specific physical problem being calculated. This is a reference speed, chosen here as the speed of sound at infinity. a ,and γ is the specific heat ratio, and R is the gas constant. and These represent the coordinate components and velocity components in three directions within an orthogonal rectangular coordinate system. p , T and E These are density, pressure, temperature, and total energy, respectively. t For time, Re , Ma and Pr These are the Reynolds number, Mach number, and Prandtl number, respectively.

[0043] The dimensionless Navier-Stokes equations in their conservation form are as follows:

[0044]

[0045] Among them, the conserved variables U for:

[0046] .

[0047] Non-viscous flux F , G , H They are respectively:

[0048]

[0049] Viscous flux (F) v G v H v )for:

[0050]

[0051] Viscous stress term , , , , , for:

[0052]

[0053] heat flux q x , q y , q z They are respectively:

[0054]

[0055] Here, turbulence simulation is used to conduct numerical calculations of the flow in the three-dimensional cavity structure of the aircraft. At this time, the governing equations need to be replaced with the following equations based on the Navier-Stokes equations:

[0056]

[0057] in, and The thermal conductivity coefficients corresponding to laminar and turbulent flow, respectively. For isobaric specific heat capacity, and The viscosity coefficients corresponding to laminar and turbulent flow, respectively; laminar viscosity coefficient. Calculations are performed using the Sutherland formula:

[0058]

[0059] In the above formula, C represents the temperature constant;

[0060] The turbulent viscosity coefficient was calculated using the Spalart-Allmaras turbulence model combined with the DDES method. l =0.7 and Pr t =0.9 corresponds to the Prandtl number for laminar and turbulent flow, respectively.

[0061] The total energy E is:

[0062]

[0063] In the above formula, u , v , w These represent velocities in the x, y, and z directions, respectively.

[0064] At the same time, based on the ideal gas assumption, we have:

[0065]

[0066] In step one, the coupling effects between various physical quantities during the flight of a high-speed aircraft are considered, so that they participate in the calculation as much as possible during the simulation process to ensure the calculation effect.

[0067] Step 2: After the control equations are constructed, the computational domain and mesh are designed for typical high-speed aircraft configurations.

[0068] The dimensions of the three-dimensional aircraft cavity structure calculation model are referenced from the standard model designed in the experiment. The length (L), width (W), and depth (d) of the aircraft cavity structure are 200mm, 66.7mm, and 33.3mm, respectively, satisfying L:W:D=6:2:1. The calculation uses... Figure 1 The simplified model shown ( Figure 1 L i L is the distance between the leading edges of the cavity. o W is the distance from the trailing edge of the cavity. b (This refers to the distance between the left and right edges of the cavity). The computational mesh is a structurally stretched mesh, with the mesh being finer near the wall and gradually stretched in the far field. Some meshes are as follows: Figure 2 As shown. The minimum grid size for the wall is 1.0 × 10⁻⁶. -4The total mesh size is approximately 11 million. The origin of the coordinate system is set at the upper corner of the leading edge of the spacecraft cavity structure, with the X, Y, and Z directions representing the flow direction, longitudinal direction, and spanwise direction, respectively. Based on the depth D of the spacecraft cavity structure, the computational domain is defined as −16D≤x≤20D, −1D≤y≤12D, and −3D≤z≤5D, with the walls being adiabatic.

[0069] Step 3: Obtain the total sound pressure level and spectrum of key measurement points under typical speed conditions of high-speed aircraft, and compare and correct them with the wind tunnel test results.

[0070] Assuming the cavity wall of the aircraft is the solid boundary of the flow field, the classical FW-H equations can be used to predict flow-induced noise, such as boundary layer noise, shear layer noise, and turbulence noise. For considering the cavity wall, the differential form of the FW-H equations is shown below:

[0071]

[0072] in, c The far-field sound velocity of the flow field. p This refers to far-field sound pressure, which is small relative to fluid pressure. and The integral surface of the sound source includes the solid walls of the cavity. f ca and penetrable walls f pe Two parts, u j for x i directional fluid velocity components, u n Let be the fluid velocity component perpendicular to the given sound source surface. v n For the corresponding x i The tangential fluid component of the sound source surface in the direction of the sound. For the Dirac function, =0 indicates the given sound source integral surface, and H(f) is the Heaviside function constructed from the sine function, used to automatically determine the integral surface. P ij The stress tensor defined by Lighthill represents the sound sources inside the fluid. The first term on the right-hand side of the formula is the volumetric noise source, caused by the complex flow of the fluid itself, such as turbulent stress; the second term is the dipole sound source, which comes from wall pressure fluctuations and mainly generates boundary layer noise; the third term is the monopole sound source, which originates from the normal motion of the wall, such as noise caused by wall vibration.

[0073] This step mainly focuses on the dipole sound source caused by the wall; the sound source integration surface only needs to be the cavity wall. f caTherefore, FW-H can be simplified to:

[0074]

[0075] Based on the Green's function, integrating the above formula along the cavity wall yields:

[0076]

[0077] in, v s It is the wall motion speed. n It is the outward normal vector of the wall.

[0078] The sound source is extracted based on the computational fluid dynamics (CFD) simulation results, and the reference value for force pulsation is calculated according to the quasi-steady results. Because the pressure inside the cavity differs significantly from the incoming static pressure, calculating based on the incoming static pressure would increase the error in the overall sound pressure level calculation. Therefore, in this step, the incoming static pressure p0 is set as the time-averaged value at various points in the quasi-steady flow field (i.e., the reference value for calculating force pulsation), as shown in the following expression:

[0079]

[0080] Where N represents the total number of time series data collection points. p s,i The pressure value at the sampling point. The wall pulsating pressure value is extracted based on the unsteady calculation results. p .

[0081] Based on the sound pressure time series at the measurement point location, a fast Fourier transform is performed to obtain the sound pressure amplitude, as shown in the following expression:

[0082]

[0083] The formula for calculating the total sound pressure level at each frequency of the measuring point is as follows:

[0084]

[0085] In the formula, p ref =2×10 −5 Pa represents the reference pressure.

[0086] In this step, the incoming flow temperature is assumed to be 160K, the incoming flow velocity to be Mach 2.0, and the Reynolds number based on the depth of the spacecraft cavity structure to be 8.0 × 10⁻⁶. 5 Once the calculations stabilized, pressure pulsation monitoring was performed at measuring points on the bottom and front and rear walls of the aircraft cavity structure (monitoring numbers B1-B7). The monitoring locations are shown in Table 1.

[0087] Table 1: Coordinates of the locations of pulsating pressure measurement points within the aircraft cavity structure

[0088]

[0089] Two measurements were conducted in the corresponding wind tunnel test. In one test, the measurement hole position was consistent with the calculation, while in the other test, the measurement hole position was slightly different from the calculation result. These measurements can be used for comparison.

[0090] Figure 3 This paper presents a comparison between numerical and experimental results of the total sound pressure level at different locations on the bottom wall of the center plane of the aircraft cavity structure (where COM represents the calculated result, EXP1 represents experimental result 1, and EXP2 represents experimental result 2). Figure 3 The results show that the calculated results agree well with the overall experimental results. Compared with the experiment at the same measurement location, the maximum and minimum errors of the total sound pressure level in the calculated results are 2.56 dB and 0.74 dB, respectively, with an average error of 1.60 dB, indicating high calculation accuracy.

[0091] Step 4: Obtain and analyze the flow results of key configurations under typical speed conditions of high-speed aircraft.

[0092] The flow type within an aircraft cavity structure is primarily related to the ratio of its length (L) to its depth (d) and the incoming Mach number. When the cavity is relatively deep, i.e., L / D < 6–8 in subsonic conditions or L / D < 10 in supersonic conditions, the flow within the aircraft cavity structure is open. In this case, the shear layer at the cavity opening impacts the rear wall of the aircraft cavity structure.

[0093] Figures 4-7 The flow states obtained through the above calculation method are shown, respectively illustrating the streamline results near the cavity structure wall of the aircraft at Mach numbers of 2.0, 2.5, 3.0, and 3.5. Figure 4 For example, from Figure 4 The results show that a vortex core structure formed by the recirculation zone can be seen on the left and right side walls. Figure 4 This is clearly evident in the flow structure at the center section of the aircraft cavity. This recirculation zone occupies the vast majority of the aircraft cavity structure, a significant characteristic of flow in open aircraft cavity structures. In the description of the flow results in two-dimensional aircraft cavity structures, the recirculation zone is formed due to the impact of the shear layer on the rear wall, causing some fluid to enter the aircraft cavity structure and flow in the reverse direction.

[0094] exist Figure 4On the near-side wall section, a clockwise focal point formed by the recirculation zone is visible, indicating the presence of a vortex core originating from the side wall within the recirculation zone. A counter-clockwise rotating leading vortex can also be observed near the forward wall and side wall regions of the weapon compartment. In the two-dimensional description, the leading vortex is interpreted as being induced by the recirculation zone; in the three-dimensional case, it exhibits distinct crossflow characteristics. The leading vortex is generated by fluid moving upstream along the side wall and separating near the forward edge, transporting the fluid medium from both sides of the weapon compartment to the central region. This separation characteristic is reflected in the flow pattern in the corresponding area of ​​the side wall. At the front end of the bottom wall of the weapon compartment, there is a pair of counter-rotating stable focal points, formed by a tornado-like vortex located at the front of the weapon compartment. The tail vortex is located downstream of the recirculation zone, near the rear and side walls, and rotates clockwise. The tail vortex is directly generated by the shear layer impact process, hence its high intensity. The fluid medium moves upstream with the tail vortex and is then deflected, flowing out at high speed from both sides of the weapon compartment. From the central section towards the side wall, the tail vortex gradually spirals outward.

[0095] The above solution is merely an illustration of a preferred example and is not limited thereto. When implementing this invention, appropriate substitutions and / or modifications can be made according to the user's needs.

[0096] Although embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and embodiments. It can be applied to various fields suitable for the present invention. Other modifications can be readily made by those skilled in the art. Therefore, without departing from the general concept defined by the claims and their equivalents, the present invention is not limited to the specific details and examples shown and described herein.

Claims

1. A method for calculating unsteady flow and noise in a high-speed aircraft, characterized in that, include: S1. Select reference physical quantities based on the actual flight environment of high-speed aircraft, and obtain the corresponding dimensionless physical quantities through the reference physical quantities; S2. Define the control equations for unsteady flow and noise calculation of high-speed aircraft based on dimensionless physical quantities, and complete the dimensionless transformation of the control equations. S3. Perform three-dimensional numerical simulation calculations on the cavity structure in the aircraft based on the control equations, and optimize the computational region and computational network during the calculation process; S4. In the cavity structure of the aircraft, the cavity wall is assumed to be the solid boundary of the flow field. The FW-H equation for predicting flow-induced noise is simplified, and the simplified formula is integrated along the cavity wall based on the Green's function. Then, the total sound pressure level at each measuring point at each frequency under the corresponding speed condition in the high-speed aircraft is obtained. SPL ( f The following formula is used to obtain: In the above formula, p ref For reference pressure, P( f The sound pressure amplitude is obtained by performing a Fast Fourier Transform on the sound pressure time series at the measurement point location. ; S5. Compare and correct the sound pressure data calculated in S4 with the wind tunnel test results until the calculation results are stable. S6. The streamline morphology corresponding to the calculation results of S5 is used to characterize the flow topology and vortex morphology near the wall of the aircraft cavity structure.

2. The method for calculating unsteady flow and noise of a high-speed aircraft as described in claim 1, characterized in that, In S1, the reference physical quantity includes: reference length. Reference speed Reference temperature Reference density Reference viscous dynamic coefficient The relevant dimensionless physical quantities include: coordinate components in three directions under an orthogonal rectangular coordinate system. Velocity components in three directions in an orthogonal rectangular coordinate system Pressure p, temperature T, total energy E, density ρ, speed of sound a, Reynolds number Re, Mach number Ma, Prandtl number P r Time t, static temperature T m kinetic viscosity coefficient µ .

3. The method for calculating unsteady flow and noise of a high-speed aircraft as described in claim 1, characterized in that, In S2, the dimensionless governing equations are characterized by the following equation: In the above formula, U is a conserved variable, x, y, and z are the coordinate components in three directions under an orthogonal rectangular coordinate system, and F... v G v H v All are non-viscous flux, (F v G v H v () represents viscous flux.

4. The method for calculating unsteady flow and noise of a high-speed aircraft as described in claim 1, characterized in that, In S3, the three-dimensional numerical simulation calculation adopts a simplified model; Among them, the optimization of the computational grid refers to: using a structured stretched grid, with the grid being denser near the wall and gradually stretched in the far field region; The optimization of the computational region refers to: taking the depth D of the aircraft cavity structure as the reference, setting the upper corner of the leading edge of the aircraft cavity structure as the origin of the coordinate system, then in the three directions of x, y, and z, the range of the computational region is limited to: −16D≤x≤20D, −1D≤y≤12D and −3D≤z≤5D, and the aircraft cavity wall is set as an insulating wall.

5. The method for calculating unsteady flow and noise of a high-speed aircraft as described in claim 1, characterized in that, In S4, the total sound pressure level SPL ( f The process for obtaining ) is as follows: S40. When only considering boundary layer noise generated by wall pressure fluctuations, the sound source integral surface is also selected only from the solid wall of the cavity. f ca Then the simplified form of the FW-H equation is characterized by the following equation: In the above formula, c The velocity of sound in the far field of the flow field. p For far-field sound pressure, t For time, u n Let be the fluid velocity component perpendicular to the given sound source surface. u j For x i Fluid velocity components in the direction, v n For the corresponding x i The tangential fluid component of the sound source surface in the direction of the sound source. P ij The stress tensor defined for Lighthill For the Dirac function, f ca The solid wall of the cavity. n j For x j The unit normal vector in the direction; S41. Based on the Green's function, the simplified expression of S40 is obtained by integrating along the cavity wall: In the above formula, ρ 0 represents the incoming flow density. S ca The area of ​​the cavity wall is 1. ret Net flux, p Where R is the far-field sound pressure, and R is the gas constant. p 0 represents the incoming static pressure, and , N This represents the total number of time series data collection points. p s,i The pressure value at the sampling point. v s It is the wall motion speed. n It is the outward normal vector of the wall, and t is time. dS The area of ​​the infinitesimal element; S42. Based on the sound pressure time series at the measurement point location, and the sound pressure amplitude P after the fast Fourier transform, f This allows us to obtain the total sound pressure level at each measurement point at each frequency. SPL ( f ).

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