Fuselage design method and system based on pressure gradient line tracking

Abstract

The invention discloses a fuselage design method and system based on pressure gradient line tracking. The method comprises the steps that the source intensity of a sound source and the incident sound pressure pin of the sound source at a ground receiving point are obtained through calculation; enabling the scattering sound pressure psc of the equivalent source at the ground receiving point to be equal to-pin, and calculating to obtain the source intensity of the equivalent source; obtaining a pressure gradient vector of the fuselage wall surface based on the source intensity of the sound source and the source intensity of the equivalent source; and performing pressure gradient line tracking on the fuselage wall surface on the basis of the pressure gradient vector on the basis of the fuselage symmetric section, so that a point product result of a normal vector of the fuselage wall surface and the pressure gradient vector is 0, obtaining a fuselage half-model, generating a fuselage full-model after mirror symmetry of the fuselage half-model, and completing fuselage design. The method is applied to the field of aircraft design, the fuselage wall surface is generated through the equivalent negative incident sound pressure and the tracking pressure gradient line, the purpose of reducing noise can be achieved, and the dynamic characteristics of the aircraft can be effectively prevented from being affected.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft design, and in particular to a fuselage design method and system based on pressure gradient line tracing. Background Art

[0002] With the rise of the low-altitude economy, noise issues from low-altitude aircraft such as helicopters, quadrotors, and multirotors are becoming increasingly prominent, impacting residents' physical and mental health and quality of life. Reducing noise has become crucial for increasing aircraft payload, density, and flight frequency, and has long been a research focus and hot topic in the field of aeroacoustics.

[0003] Noise reduction generally starts from two aspects: sound source and sound propagation.

[0004] From the perspective of the sound source, the noise of low-altitude aircraft mainly comes from power-generating devices such as engines and rotors. The team of Professor Zhang Xin of the Hong Kong University of Science and Technology has developed a design method for low-noise rotors and carried out a large number of unsteady simulations and experimental verifications. Teams such as Han Zhonghua of Northwestern Polytechnical University, Zhao Qijun of Nanjing University of Aeronautics and Astronautics, and Yang Cheng of Shanghai Jiaotong University have also carried out a large number of innovative research. In general, the noise reduction schemes for controlling sound sources are basically to reduce the intensity of the sound source and change the directionality of the sound source by optimizing the number, speed, position, size, airfoil, material and other parameters of the rotors, thereby reducing the impact of noise on residents. However, the load noise emitted by the sound source is often related to the weight of the entire aircraft. When the load noise is reduced, the lift characteristics will also decrease. The present invention only changes the fuselage and will not affect the power of the aircraft.

[0005] From the perspective of sound propagation, the most common and effective applications are noise-reducing materials and structures, such as acoustic liners, porous media, and metasurfaces, which achieve noise reduction by increasing dissipation or controlling the direction of sound propagation. Controlling the phase of multi-axis rotor rotation can achieve phase cancellation noise reduction; optimizing flight trajectory can also reduce the impact of noise on the ground to a certain extent. Noise reduction solutions that manipulate fuselage materials absorb sound by changing the properties and distribution of wall materials. This sound absorption method generally reduces scattered sound pressure, but does not take into account incident sound pressure, making it difficult to achieve or approach the goal of zero total sound pressure. Summary of the Invention

[0006] In response to the above-mentioned deficiencies in the existing technology, the present invention provides a fuselage design method and system based on pressure gradient line tracking. By generating the fuselage wall surface through equivalent negative incident sound pressure and tracking pressure gradient lines, it can not only achieve the purpose of reducing noise, but also effectively avoid affecting the dynamic characteristics of the aircraft because only the surface structure of the fuselage is changed.

[0007] To achieve the above object, the present invention provides a fuselage design method based on pressure gradient line tracing, comprising the following steps:

[0008] Step 1: Obtain the sound source integral surface information of the sound-emitting components on the aircraft, and calculate the source intensity of the sound source and the incident sound pressure p of the sound source at the ground receiving point based on the sound source integral surface information. in ;

[0009] Step 2: Arrange an equivalent source in the fuselage area of ​​the aircraft and set the scattered sound pressure p of the equivalent source at the ground receiving point sc =-p in , and calculate the source strength of the equivalent source;

[0010] Step 3: Obtain the sound pressure gradient of the sound source at the receiving point on the fuselage wall based on the source intensity of the sound source, and obtain the sound pressure gradient of the equivalent source at the receiving point on the fuselage wall based on the source intensity of the equivalent source;

[0011] Step 4: synthesize the sound pressure gradients of the sound source and the equivalent source at the receiving point on the fuselage wall to obtain the pressure gradient vector of the fuselage wall;

[0012] Step 5: Based on the symmetrical cross-section of the fuselage, the pressure gradient line of the fuselage wall is traced based on the pressure gradient vector, so that the dot product of the normal vector of the fuselage wall and the pressure gradient vector is 0, thereby obtaining a fuselage half mold. The fuselage half mold is mirrored to generate a full fuselage mold, thus completing the fuselage design.

[0013] In one embodiment, in step 1, the sound source integral surface information of the sound-generating components on the aircraft is obtained through unsteady numerical simulation.

[0014] In one embodiment, in step 1, the height of the ground receiving point is 0 m to 2 m above the ground.

[0015] In one embodiment, in step 2, the source strength calculation process of the equivalent source is:

[0016] G(a,b,ω)·q(a)=p sc (b)

[0017] Where q(a) is the source strength of the equivalent source, G(a,b,ω) is the transfer matrix between the equivalent source and the ground receiving point, a is the coordinate of the equivalent source, b is the coordinate of the ground receiving point, ω is the circular frequency, and p sc (b) is the scattered sound pressure of the equivalent source at the ground receiving point.

[0018] In one embodiment, in step 4, the pressure gradient vector of the fuselage wall is:

[0019]

[0020] Among them, G in is the incident sound pressure of the sound source at the receiving point on the fuselage wall, G scis the scattered sound pressure of the equivalent source at the receiving point on the fuselage wall, The incident sound pressure G in The pressure gradient components in the x, y, and z directions, are the scattered sound pressure G sc Pressure gradient components in the x, y, and z directions;

[0021] The dot product of the normal vector of the fuselage wall and the pressure gradient vector is 0, which is:

[0022]

[0023] in, are the projection components of the normal vector of the fuselage wall in the x, y, and z directions respectively.

[0024] In one embodiment, in step 5, the pressure gradient line tracking of the fuselage wall is performed based on the pressure gradient vector on the basis of the symmetrical cross section of the fuselage, specifically:

[0025] The velocity vectors in three directions in the flow direction tracking technology are replaced by pressure gradient vectors to form pressure gradient line tracking, and pressure gradient line tracking is performed on the fuselage wall based on the pressure gradient vector on the basis of the symmetrical cross-section of the fuselage.

[0026] To achieve the above object, the present invention further provides a fuselage design system based on pressure gradient line tracing, which uses the above method to design the fuselage of an aircraft. The fuselage design system includes:

[0027] The sound source calculation unit is used to obtain the sound source integral surface information of the sound-emitting components on the aircraft, and calculate the source intensity of the sound source and the incident sound pressure p of the sound source at the ground receiving point based on the sound source integral surface information. in ;

[0028] The equivalent source calculation unit is used to arrange the equivalent source in the fuselage area of ​​the aircraft and make the scattered sound pressure p of the equivalent source at the ground receiving point sc =-p in , and calculate the source strength of the equivalent source;

[0029] a pressure gradient calculation unit, configured to obtain the sound pressure gradient of the sound source at the receiving point on the fuselage wall according to the source intensity of the sound source, and obtain the sound pressure gradient of the equivalent source at the receiving point on the fuselage wall according to the source intensity of the equivalent source, and then combine the sound pressure gradients of the sound source and the equivalent source at the receiving point on the fuselage wall to obtain the pressure gradient vector on the fuselage wall;

[0030] a pressure gradient line tracking unit, configured to perform pressure gradient line tracking on the fuselage wall surface based on the pressure gradient vector on the basis of a symmetrical cross-section of the fuselage, so that a dot product of a normal vector of the fuselage wall surface and the pressure gradient vector is 0, thereby obtaining a fuselage half mold;

[0031] The mirror symmetry unit is used to generate a full fuselage mold after mirroring the fuselage half mold to complete the fuselage design.

[0032] To achieve the above object, the present invention further provides a terminal device, wherein the terminal device is provided with:

[0033] Memory, used to store programs;

[0034] The processor is configured to execute the program stored in the memory. When the program is executed, the processor is configured to execute the method described above.

[0035] To achieve the above objectives, the present invention further provides a computer-readable storage medium, wherein the computer-readable storage medium stores computer-executable instructions; when the computer-executable instructions are executed by a processor, they are used to implement the above method.

[0036] Compared with the prior art, the present invention has the following beneficial technical effects:

[0037] 1. This invention generates the fuselage wall surface by equivalent negative incident sound pressure and tracing pressure gradient lines, taking minimum total sound pressure as the criterion. By manipulating the wall surface shape, the scattered sound pressure can be controlled in a novel way, thereby achieving the purpose of noise reduction. At the same time, since only the fuselage surface structure is changed, the dynamic characteristics of the aircraft can be effectively avoided.

[0038] 2. The aircraft fuselage designed using the design method of the present invention can create a relatively quiet area below the aircraft, reducing the impact of noise on pedestrians on the ground. Although noise still exists in areas above the height of a normal person, it is inaudible to the human ear.

[0039] 3. The present invention can be realized by numerical simulation and has better universality than theoretical methods. It can be applied to sound sources of different frequencies, different motion modes and different intensities, and it is only necessary to wrap the sound source within the integral surface. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] 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.

[0041] Figure 1 Flowchart of the fuselage design method based on pressure gradient line tracing in Example 1 of the present invention;

[0042] Figure 2 This is a schematic diagram of the numerical structure of the fuselage based on the minimum total sound pressure criterion in Example 1 of the present invention;

[0043] Figure 3 Schematic diagram of the sound source in Example 1 of the present invention;

[0044] Figure 4 Schematic diagram of sound radiation in Example 1 of the present invention;

[0045] Figure 5 Schematic diagram of the numerical equivalent process in Example 1 of the present invention;

[0046] Figure 6 Schematic diagram of the numerical matching process in Example 1 of the present invention;

[0047] Figure 7 Schematic diagram of the symmetrical cross-sectional shape of the fuselage in Example 1 of the present invention;

[0048] Figure 8 Schematic diagram of a fuselage half mold based on pressure gradient line tracking in Example 1 of the present invention;

[0049] Figure 9 This is a structural block diagram of a fuselage design system based on pressure gradient line tracking in Example 2 of the present invention;

[0050] Figure 10 This is a structural block diagram of the terminal device in Example 3 of the present invention.

[0051] Figure numbers: sound source 1, sound source integration surface 2, pedestrian 3, ground receiving point 4, ground 5, equivalent source 6, fuselage wall point 7, fuselage 8, fuselage wall receiving point 9, fuselage symmetrical cross-sectional shape 10, pressure gradient line 11.

[0052] The purpose, features and advantages of the present invention will be further described with reference to the accompanying drawings and in conjunction with the embodiments. DETAILED DESCRIPTION

[0053] 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.

[0054] In addition, the technical solutions between the various embodiments of the present invention can be combined with each other, but it must be based on the fact that ordinary technicians in this field can implement it. When the combination of technical solutions is mutually contradictory or cannot be implemented, it should be deemed that such a combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0055] Example 1

[0056] This embodiment discloses a fuselage design method based on pressure gradient line tracing, in which the design of the fuselage profile is completed with the criterion of minimizing the total sound pressure, thereby reducing the impact of aircraft noise on pedestrians on the ground. The total sound pressure is divided into two parts: incident sound pressure and scattered sound pressure. The incident sound pressure refers to the sound pressure directly from the sound source to the ground receiving point, and the scattered sound pressure refers to the sound pressure from the sound source to the fuselage and then to the ground receiving point. When the total sound pressure is zero, the sound source will not affect the ground receiving point. Specifically, in the fuselage design method of this embodiment, by controlling the fuselage shape of the aircraft so that the scattered sound pressure is equal to the negative incident sound pressure, the effect of zero total sound pressure can be achieved, thereby achieving the purpose of noise reduction. The actual aircraft fuselage has a larger size and more degrees of freedom. Noise reduction achieved by designing the fuselage can also effectively avoid affecting the dynamic characteristics of the aircraft.

[0057] refer to Figure 1 In this embodiment, the fuselage design method based on pressure gradient line tracking includes the following steps:

[0058] Step 1: Obtain the sound source integral surface information of the sound-emitting components on the aircraft, and calculate the source intensity of the sound source and the incident sound pressure p of the sound source at the ground receiving point based on the sound source integral surface information. in ;

[0059] Step 2: Arrange an equivalent source in the fuselage area of ​​the aircraft and set the scattered sound pressure p of the equivalent source at the ground receiving point sc =-p in , and calculate the source strength of the equivalent source;

[0060] Step 3: Obtain the sound pressure gradient of the sound source at the receiving point on the fuselage wall based on the source intensity of the sound source, and obtain the sound pressure gradient of the equivalent source at the receiving point on the fuselage wall based on the source intensity of the equivalent source;

[0061] Step 4: synthesize the sound pressure gradients of the sound source and the equivalent source at the receiving point on the fuselage wall to obtain the pressure gradient vector of the fuselage wall;

[0062] Step 5: Based on the symmetrical cross-section of the fuselage, the pressure gradient line of the fuselage wall is traced based on the pressure gradient vector, so that the dot product of the normal vector of the fuselage wall and the pressure gradient vector is 0, and the fuselage half mold is obtained. The fuselage half mold is mirrored and symmetrically generated to generate the full fuselage mold, completing the fuselage design.

[0063] refer to Figure 2 This is a schematic diagram of the numerical construction principle of the fuselage based on the minimum total sound pressure criterion in this embodiment. A sound source 1 on the aircraft can be stationary or moving. The sound waves it emits reach the sound source integration surface 2 and then radiate outward from the sound source integration surface 2 to a ground receiving point 4, representing the impact of the noise on a pedestrian 3 on the ground 5. Next, assuming that the scattered sound pressure at the ground receiving point 4 is equal to the negative incident sound pressure, an equivalent source 6 within the fuselage 8 is used to equate the scattered sound pressure at the ground receiving point 4. The source intensity of the equivalent source 6 is then matched to the sound source integration surface 2 to obtain a usable fuselage wall point 7, thereby constructing the fuselage. Because both the equivalence and matching processes are implemented numerically, the fuselage design method in this embodiment is a numerical construction method. Compared to theoretical methods, it is more universal and can be applied to sound sources of different frequencies, different motion modes, and different intensities, simply by enclosing the sound source within the integration surface.

[0064] In this embodiment, the sound source 1 has various forms, such as Figure 3 As shown, different sound-generating components on the aircraft can be covered. Through unsteady numerical simulation, information on the sound source integral surface 2 that wraps the sound source can be obtained, such as pressure disturbance, density disturbance and velocity disturbance, etc. Among them, the sound source integral surface 2 refers to a closed surface defined around the sound source 1, and the flow field parameters (such as pressure and velocity) on the closed surface are used to calculate the far-field sound field without directly solving the complex sound propagation equation. As for obtaining the information on the sound source integral surface 2 through unsteady numerical simulation, it belongs to conventional technical means and will not be described in detail in this embodiment.

[0065] refer to Figure 4 The height of a normal pedestrian 3 is generally less than 2 meters. Therefore, in this embodiment, the ground receiving point 4 is arranged within a range of about 0m to 2m from the ground 5. By substituting the information recorded on the sound source integration surface 2 into the classic FW-H equation, the incident sound pressure p of the sound source 1 at the ground receiving point 4 can be predicted. in and the source strength of the sound source.

[0066] refer to Figure 5 In this embodiment, based on the minimum total sound pressure criterion, the scattered sound pressure p of the equivalent source 6 at the ground receiving point 4 is sc =-p in , assuming that the coordinate of the equivalent source 6 is a, the coordinate of the ground receiving point 4 is b, the transfer matrix between the equivalent source 6 and the ground receiving point 4 is G(a, b, ω), and the source strength of the equivalent source is q(a), then the following equivalent relationship exists:

[0067] G(a,b,ω)·q(a)=p sc (b)

[0068] The transfer matrix G(a, b, ω) can be directly calculated based on the coordinates a of the equivalent source 6, the coordinates b of the ground receiving point 4, and the circular frequency ω. The scattered sound pressure p of the equivalent source at the ground receiving point is sc (b) is known, so the source intensity of the equivalent source q(a) = G can be obtained by truncated singular value decomposition method or regularization method. -1 (a,b,ω)·p sc (b).

[0069] refer to Figure 6 Since this embodiment is based on the minimum total sound pressure criterion (i.e., the zero total sound pressure criterion), the boundary conditions required for the sound scattering problem in this embodiment are:

[0070]

[0071] Where i is the imaginary unit, ρ0 is the air density, Z is the impedance, G in 、 They are respectively the incident sound pressure and pressure gradient of the sound source 1 at the receiving point 9 on the fuselage wall predicted by the sound source integration surface 2, G sc 、 are the scattered sound pressure and pressure gradient of the equivalent source 6 at the receiving point 9 on the fuselage wall;

[0072] When impedance is not considered, the fuselage is a rigid wall (i.e., impedance Z is infinite), and the boundary condition is changed to Its expanded form is:

[0073]

[0074] in, The incident sound pressure G in The pressure gradient components in the x, y, and z directions, are the scattered sound pressure G sc The pressure gradient components in the x, y, and z directions, are the projection components of the normal vector of the fuselage wall in the x, y, and z directions respectively;

[0075] The above can be regarded as the pressure gradient vector of the fuselage wall Normal vector to the fuselage wall Therefore, by tracking the pressure gradient line of the fuselage wall based on the pressure gradient vector on the basis of the symmetrical cross-section of the fuselage, so that the point product of the normal vector of the fuselage wall and the pressure gradient vector is 0, the design purpose can be achieved. In the specific implementation process, this embodiment draws on the streamline tracking technology commonly used in the design of conventional air inlets and nozzles, replaces the velocity vectors in the three directions of the streamline tracking technology with pressure gradient vectors, and forms a pressure gradient line tracking technology. Then, in Figure 7 Based on the given fuselage symmetrical cross-sectional shape 10 shown in FIG, a pressure gradient line tracing technique is used to generate a Figure 8 The pressure gradient line 11 is shown. Since the pressure gradient line 11 is perpendicular to the normal vector of the fuselage wall, the fuselage half mold can be constructed directly using the pressure gradient line 11. Finally, since the aircraft fuselage is generally a mirror-symmetrical structure, the fuselage half mold is mirrored to generate the full fuselage mold, thus completing the fuselage design.

[0076] Example 2

[0077] Based on the fuselage design method based on pressure gradient line tracking in Example 1, this embodiment discloses a fuselage design system based on pressure gradient line tracking, referring to Figure 9 The fuselage design system includes a sound source calculation unit, an equivalent source calculation unit, a pressure gradient calculation unit, a pressure gradient line tracking unit, and a mirror symmetry unit. It can design the fuselage based on the pressure gradient line tracking of the sound pressure to achieve noise reduction. Specifically:

[0078] The sound source calculation unit is used to obtain the sound source integral surface information of the sound-emitting components on the aircraft, and calculate the source intensity of the sound source and the incident sound pressure p of the sound source at the ground receiving point based on the sound source integral surface information. in ;

[0079] The equivalent source calculation unit is used to arrange the equivalent source in the fuselage area of ​​the aircraft, so that the scattered sound pressure p of the equivalent source at the ground receiving point is sc =-p in , and calculate the source strength of the equivalent source;

[0080] The pressure gradient calculation unit is used to obtain the sound pressure gradient of the sound source at the receiving point on the fuselage wall according to the source intensity of the sound source, and obtain the sound pressure gradient of the equivalent source at the receiving point on the fuselage wall according to the source intensity of the equivalent source, and then combine the sound pressure gradients of the sound source and the equivalent source at the receiving point on the fuselage wall to obtain the pressure gradient vector of the fuselage wall;

[0081] The pressure gradient line tracking unit is used to track the pressure gradient line of the fuselage wall based on the pressure gradient vector on the basis of the symmetrical cross section of the fuselage, so that the dot product of the normal vector of the fuselage wall and the pressure gradient vector is 0, thereby obtaining a fuselage half mold;

[0082] The mirror symmetry unit is used to generate the full fuselage mold after mirroring the fuselage half mold to complete the fuselage design.

[0083] In this embodiment, the specific working processes and working principles of the sound source calculation unit, equivalent source calculation unit, pressure gradient calculation unit, pressure gradient line tracking unit, and mirror symmetry unit are the same as those in Example 1, and therefore will not be described in detail in this embodiment. Each unit module can be implemented in whole or in part through software, hardware, or a combination thereof. Each unit module can be embedded in or independent of a processor in a computer device in hardware form, or can be stored in a memory in a computer device in software form, so that the processor can call and execute the corresponding operations of each of the above unit modules.

[0084] Example 3

[0085] like Figure 10 The terminal device disclosed in this embodiment includes a transmitter, a receiver, a memory, and a processor. The transmitter is used to send instructions and data, the receiver is used to receive instructions and data, the memory is used to store computer-executable instructions, and the processor is used to execute the computer-executable instructions stored in the memory to implement the method in the above-mentioned embodiment 1.

[0086] It should be noted that the above memory can be independent or integrated with the processor. When the memory is independently provided, the terminal device further includes a bus for connecting the memory and the processor.

[0087] Example 4

[0088] This embodiment discloses a computer-readable storage medium, in which computer-executable instructions are stored. When a processor executes the computer-executable instructions, the method in the above-mentioned embodiment 1 is implemented.

[0089] Those skilled in the art will appreciate that all or part of the processes in the above-mentioned embodiments can be implemented by instructing the relevant hardware through a computer program. The computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the embodiments of the above-mentioned methods. Among them, any reference to memory, storage, database or other media used in the embodiments provided in this application can include non-volatile and / or volatile memory. Non-volatile memory can include read-only memory (ROM), programmable ROM (PROM), electrically programmable ROM (EPROM), electrically erasable programmable ROM (EEPROM) or flash memory. Volatile memory can include random access memory (RAM) or external cache memory. By way of illustration and not limitation, RAM is available in various forms, such as static RAM (SRAM), dynamic RAM (DRAM), synchronous DRAM (SDRAM), double data rate SDRAM (DDRSDRAM), enhanced SDRAM (ESDRAM), synchronous link (Synchlink) DRAM (SLDRAM), memory bus (Rambus) direct RAM (RDRAM), direct memory bus dynamic RAM (DRDRAM), and memory bus dynamic RAM (RDRAM).

[0090] The above description is only a preferred embodiment of the present invention and does not limit the scope of protection of the present invention. All equivalent structural transformations made by using the contents of the present invention description and drawings under the inventive concept of the present invention, or direct / indirect application in other related technical fields are included in the scope of protection of the present invention.

Claims

1. A fuselage design method based on pressure gradient line tracing, characterized in that: The steps include: Step 1: Obtain the sound source integral surface information of the sound-emitting components on the aircraft, and calculate the source intensity of the sound source and the incident sound pressure p of the sound source at the ground receiving point based on the sound source integral surface information. in ; Step 2: Arrange an equivalent source in the fuselage area of ​​the aircraft and set the scattered sound pressure p of the equivalent source at the ground receiving point sc =-p in , and calculate the source strength of the equivalent source; Step 3: Obtain the sound pressure gradient of the sound source at the receiving point on the fuselage wall based on the source intensity of the sound source, and obtain the sound pressure gradient of the equivalent source at the receiving point on the fuselage wall based on the source intensity of the equivalent source; Step 4: synthesize the sound pressure gradients of the sound source and the equivalent source at the receiving point on the fuselage wall to obtain the pressure gradient vector of the fuselage wall; Step 5: Based on the symmetrical cross-section of the fuselage, the pressure gradient line of the fuselage wall is traced based on the pressure gradient vector, so that the dot product of the normal vector of the fuselage wall and the pressure gradient vector is 0, thereby obtaining a fuselage half mold. The fuselage half mold is mirrored to generate a full fuselage mold, thus completing the fuselage design.

2. The fuselage design method based on pressure gradient line tracing according to claim 1, characterized in that: In step 1, the sound source integral surface information of the sound-generating components on the aircraft is obtained through unsteady numerical simulation.

3. The fuselage design method based on pressure gradient line tracing according to claim 1, characterized in that: In step 1, the height of the ground receiving point is 0m to 2m above the ground.

4. The fuselage design method based on pressure gradient line tracing according to claim 1, 2 or 3, characterized in that: In step 2, the source strength calculation process of the equivalent source is: G(a,b,ω)·q(a)=p sc (b) Where q(a) is the source strength of the equivalent source, G(a,b,ω) is the transfer matrix between the equivalent source and the ground receiving point, a is the coordinate of the equivalent source, b is the coordinate of the ground receiving point, ω is the circular frequency, and p sc (b) is the scattered sound pressure of the equivalent source at the ground receiving point.

5. The fuselage design method based on pressure gradient line tracing according to claim 1, 2 or 3, characterized in that: In step 4, the pressure gradient vector of the fuselage wall is: Among them, G in is the incident sound pressure of the sound source at the receiving point on the fuselage wall, G sc is the scattered sound pressure of the equivalent source at the receiving point on the fuselage wall, The incident sound pressure G in The pressure gradient components in the x, y, and z directions, are the scattered sound pressure G sc Pressure gradient components in the x, y, and z directions; The dot product of the normal vector of the fuselage wall and the pressure gradient vector is 0, which is: in, are the projection components of the normal vector of the fuselage wall in the x, y, and z directions respectively.

6. The fuselage design method based on pressure gradient line tracing according to claim 1, 2 or 3, characterized in that: In step 5, the pressure gradient line is tracked on the fuselage wall based on the pressure gradient vector on the basis of the symmetrical cross section of the fuselage, specifically: The velocity vectors in three directions in the flow direction tracking technology are replaced by pressure gradient vectors to form pressure gradient line tracking, and pressure gradient line tracking is performed on the fuselage wall based on the pressure gradient vector on the basis of the symmetrical cross-section of the fuselage.

7. A fuselage design system based on pressure gradient line tracing, characterized in that: The method according to any one of claims 1 to 6 is used to design the fuselage of an aircraft, wherein the fuselage design system comprises: The sound source calculation unit is used to obtain the sound source integral surface information of the sound-emitting components on the aircraft, and calculate the source intensity of the sound source and the incident sound pressure p of the sound source at the ground receiving point based on the sound source integral surface information. in ; The equivalent source calculation unit is used to arrange the equivalent source in the fuselage area of ​​the aircraft and make the scattered sound pressure p of the equivalent source at the ground receiving point sc =-p in , and calculate the source strength of the equivalent source; a pressure gradient calculation unit, configured to obtain the sound pressure gradient of the sound source at the receiving point on the fuselage wall according to the source intensity of the sound source, and obtain the sound pressure gradient of the equivalent source at the receiving point on the fuselage wall according to the source intensity of the equivalent source, and then combine the sound pressure gradients of the sound source and the equivalent source at the receiving point on the fuselage wall to obtain the pressure gradient vector on the fuselage wall; a pressure gradient line tracking unit, configured to perform pressure gradient line tracking on the fuselage wall surface based on the pressure gradient vector on the basis of a symmetrical cross-section of the fuselage, so that a dot product of a normal vector of the fuselage wall surface and the pressure gradient vector is 0, thereby obtaining a fuselage half mold; The mirror symmetry unit is used to generate a full fuselage mold after mirroring the fuselage half mold to complete the fuselage design.

8. A terminal device, characterized in that: The terminal device is provided with: Memory, used to store programs; A processor is configured to execute the program stored in the memory, wherein when the program is executed, the processor is configured to execute the method according to any one of claims 1 to 6.

9. A computer-readable storage medium, characterized in that The computer-readable storage medium stores computer-executable instructions; when the computer-executable instructions are executed by a processor, they are used to implement the method according to any one of claims 1 to 6.

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