A duct fairing and method of designing the same

Abstract

The application relates to the technical field of aero-engine design, and discloses a pipeline fairing and a design method thereof. A plurality of fairings are distributed in the circumferential direction of an outer duct flow channel and are arranged in the radial direction. The hollow structure in the inside can wrap the outer duct pipeline, realizes the integration of pipeline storage and fairing functions, reduces the splicing error and airflow leakage risk of the split structure, replaces the traditional suspension pipeline mode, avoids the interference of the direct exposure of the pipeline in the outer duct flow channel to the airflow, improves the uniformity of the outer duct flow field, and reduces the flow field distortion. In addition, the radial section of the fairing adopts a conical structure with curvature, which can form a smooth transition with the front support plate, reduces the sudden expansion and sudden contraction of the outer duct airflow, and reduces the risk of airflow separation. The pipeline fairing can significantly reduce the aerodynamic loss, effectively improve the total pressure recovery coefficient of the outer duct, and further improve the engine thrust and reduce the specific fuel consumption.

Description

TECHNICAL FIELD

[0001] The application relates to the technical field of aero-engine design, and discloses a pipeline fairing and a design method thereof. BACKGROUND

[0002] The advantages and disadvantages of an aero-engine outer channel design will closely affect important performance parameters such as engine thrust and fuel consumption rate, especially for a large-bypass-ratio engine, which cannot bear the cost of high fuel consumption rate. Therefore, in order to improve the performance of the outer channel, the outer channel flow passage needs to be finely designed, the sudden expansion or sudden contraction structure needs to be reduced, the gas flow damage needs to be reduced, and the total pressure recovery coefficient of the outer channel needs to be improved to achieve the purpose of improving the engine performance.

[0003] There are many pipelines, cables and other structures penetrating the outer channel of an aero-engine, such as bleed air pipes, oil supply pipes, oil return pipes, ignition cables and the like. These pipelines penetrating the outer channel will seriously affect the outer channel flow field and pose a challenge to the design of the outer channel of the engine.

[0004] At present, for a small or medium bypass ratio engine, the pipelines penetrating the outer channel are generally suspended in the outer channel without treatment, which reduces the performance of the outer channel flow. For a large-bypass-ratio engine, a simple fairing is generally designed to wrap the pipelines penetrating the outer channel, which reduces the loss of the outer channel. However, since the fairing is a thin-walled part with low inherent frequency, there is a risk of fatigue failure caused by resonance with the engine and airflow excitation. In addition, under the installation stress or external airflow excitation, the outer shape of the thin-walled part is difficult to maintain, which has a bad influence on the outer channel flow field and causes great performance loss of the outer channel and the whole engine. SUMMARY

[0005] The application aims to provide a pipeline fairing and a design method thereof, which can effectively identify the occurrence of non-integer order vibration of a blade.

[0006] In order to achieve the above technical effects, the technical scheme adopted by the application is as follows:

[0007] A design method of a pipeline fairing, the fairing is arranged along the radial direction of an aero-engine, and the radial cross section of the fairing is in a conical structure with curvature, the conical opening end of the fairing is fixedly connected with an intermediate nacelle support plate, and the conical part of the fairing is located at a downstream position of the intermediate nacelle support plate; the inside of the fairing is provided as a hollow structure for the pipeline to pass through; the design method comprises the following steps:

[0008] establishing a finite element analysis model of the fairing according to the size parameters of the outer channel flow passage of the aero-engine and the profile structure of the radial cross section of the fairing;

[0009] performing simulation analysis on the finite element analysis model to obtain the inherent frequency of the fairing;

[0010] determining whether the natural frequency is within a 20% margin of the high-pressure rotor fundamental frequency of the aero-engine and within a 20% margin of the low-pressure rotor fundamental frequency, if not, the corresponding pipeline fairing structure meets the design requirements, if yes, adjusting the taper end angle of the taper structure until the natural frequency of the fairing exceeds the 20% margin of the high-pressure rotor fundamental frequency of the aero-engine, and the natural frequency of the fairing exceeds the 20% margin of the low-pressure rotor fundamental frequency.

[0011] Further, the spline curve of the taper type with curvature on the radial section of the fairing is , wherein is the width of the tail of the intermediate nacelle support plate, is the taper end angle of the fairing, is the longitudinal coordinate of the plane rectangular coordinate system, which is established with the taper of the taper structure as the coordinate origin, the tangent line at the position of the taper as the X-axis, and the straight line passing through the taper and perpendicular to the X-axis as the Y-axis, is the abscissa, , the value range of which is , is the circumferential width of the tail of the intermediate nacelle support plate in front of the fairing.

[0012] Further, the taper end angle of the fairing is .

[0013] Further, the wall thickness of the fairing is 1.0mm-2.0mm.

[0014] Further, the fairing is provided with an axial mounting hole, the axial mounting hole is fixedly connected with the mounting edge of the intermediate nacelle, and the axial mounting hole is a waist-shaped hole extending along the radial direction of the aero-engine.

[0015] Further, the top surface of the fairing is fixed to the inner wall surface of the outer nacelle through a mounting flange, and the mounting hole on the mounting flange is a waist-shaped hole extending in the axial direction.

[0016] Further, a sealing rubber gasket is arranged at the position where the fairing cooperates with the inner wall surface of the outer nacelle.

[0017] In order to achieve the above technical effects, the application further provides a pipeline fairing, which is obtained by the design method.

[0018] Compared with the prior art, the pipeline fairing of the present application has the beneficial effects that: the pipeline fairing of the present application realizes the integration of pipeline storage and fairing function, reduces the splicing error and air leakage risk of split structure, replaces the traditional suspension pipeline mode, avoids the interference of the pipeline directly exposed in the outer duct flow channel to the airflow, improves the uniformity of the outer duct flow field, and reduces the flow field distortion; in addition, the radial section of the fairing adopts a tapered structure with curvature, which can form a smooth transition with the front support plate, reduce the sudden expansion and sudden contraction of the outer duct airflow, and reduce the risk of airflow separation. The pipeline fairing of the present application can significantly reduce the aerodynamic loss, effectively improve the total pressure recovery coefficient of the outer duct, and further improve the engine thrust and reduce the fuel consumption rate. BRIEF DESCRIPTION OF DRAWINGS

[0019] Figure 1 It is a schematic diagram of the installation of the pipeline fairing in the outer duct flow channel in the embodiment;

[0020] Figure 2 It is a schematic diagram of the tapered structure with curvature of the radial section of the pipeline fairing in the embodiment;

[0021] Figure 3 It is Figure 1 a structural schematic diagram of the section A-A;

[0022] Figure 4 It is a perspective view of the pipeline fairing in the embodiment;

[0023] Figure 5 It is Figure 1 an enlarged schematic diagram of the local part B;

[0024] Among them, 1, fairing; 2, outer duct flow channel; 3, intermediate nacelle support plate; 4, pipeline; 5, axial mounting hole; 6, mounting flange; 7, sealing gasket. DETAILED DESCRIPTION

[0025] The present application will be further described in detail below in conjunction with the embodiments and the drawings. However, it should not be understood that the scope of the above-mentioned subject matter of the present application is limited to the following embodiments only, and any technology realized based on the content of the present application belongs to the scope of the present application.

[0026] EMBODIMENT

[0027] Referring to Figures 1 to 5 , a pipeline fairing, the fairing 1 is multiple in number, multiple said fairing 1 is distributed in the outer duct flow channel 2 of the aero-engine in a ring shape, each said fairing 1 is arranged along the radial direction of the aero-engine, and the radial section of each said fairing 1 is in a tapered structure with curvature, the tapered open end of each said fairing 1 is fixedly connected with the intermediate nacelle support plate 3, and the tapered part of each said fairing 1 is located at a downstream position of the intermediate nacelle support plate 3; the inside of the fairing 1 is provided as a hollow structure through which the pipeline 4 can pass.

[0028] In this embodiment, multiple fairings 1 are circumferentially distributed around the bypass duct 2 and arranged radially. The hollow internal structure can enclose the bypass pipe 4, realizing the integration of pipe 4 storage and rectification functions. This reduces the splicing error and airflow leakage risk of split structures, while replacing the traditional method of hanging pipelines. It avoids the interference of pipelines directly exposed in the bypass duct 2, improves the uniformity of the bypass flow field, and reduces flow field distortion. The fixing method of the conical structure and the support plate provides a reasonable spatial layout for the subsequent "internal cooling and external heating" pipeline 4 arrangement and hollow heat dissipation design, which helps to reduce internal thermal stress and coordinate installation deformation. In addition, the radial section of the fairing 1 adopts a conical structure with curvature, which can form a smooth transition with the front support plate, reducing the sudden expansion and contraction of the bypass airflow and the risk of airflow separation, significantly reducing aerodynamic losses, effectively improving the total pressure recovery coefficient of the bypass duct, and thus improving engine thrust and reducing fuel consumption.

[0029] Figure 1 The dashed line represents the shaft of an aircraft engine.

[0030] Based on the same inventive concept, this embodiment also provides a design method for a pipeline shroud, including:

[0031] The first step is to determine the number of pipes (4) that run through the engine's inner and outer ducts. Information such as the flow radius and spatial position of each pipe 4, based on the circumferential width of the rear end of the intermediate casing support plate 3 in front of the fairing 1. Preliminary estimate of the axial length of fairing 1 ,in This is a coefficient, with a value range of 1.2 to 1.8. Indicates the first Root canal 4 flow radius, This represents the area coefficient, ranging from 3.0 to 5.0.

[0032] The second step is to determine the circumferential width of the rear end of the intermediate casing support plate 3. And the preliminary estimated width and axial length of fairing 1 Smooth cubic or higher-order spline curves are plotted to ensure a smooth transition at the junction of fairing 1 and the front support plate. Using finite element simulation software such as ANSYS, the aerodynamic losses of the blade surface are simulated under the condition of limiting the engine inlet and outlet parameters (under typical operating conditions). Multiple iterations are used to obtain the ideal aerodynamic profile with the minimum flow loss, ensuring that the flow loss of a single fairing 1 is within 0.02%.

[0033] The spline curve selected in this embodiment is ,in The width of the rear of the intermediate casing support plate 3, the taper end angle of the cone portion of the fairing 1, the longitudinal coordinate of a plane rectangular coordinate system, which is established with the cone portion of the conical structure as the coordinate origin, the tangent line at the position of the cone portion as the X-axis, and the straight line passing through the cone portion and perpendicular to the X-axis as the Y-axis, the abscissa, the value range is , the circumferential width of the tail portion of the intermediate nacelle support plate 3 in front of the fairing 1. The spline curve takes into account the influence of the taper end angle of the cone portion, generates a smooth and continuous flow passage surface, obtains the profile structure of the fairing 1 in the radial section, avoids the turbulence and vortex phenomenon caused by the traditional line or circular arc splicing, further optimizes the fluid flow velocity distribution, significantly reduces the fluid resistance, and improves the outer-duct aerodynamic efficiency. In the embodiment, the taper end angle of the cone portion of the fairing 1 , can reduce the pressure loss caused by airflow separation, and further improve the outer-duct aerodynamic efficiency.

[0034] Step 3, according to the size parameters of the aero-engine outer-duct flow passage 2 and the profile structure of the fairing 1 in the radial section, a finite element analysis model of the fairing 1 is established;

[0035] Step 4, the finite element analysis model is simulated and analyzed to obtain the natural frequency of the fairing 1; it is judged whether the natural frequency is within the 20% margin range of the high-pressure rotor base frequency of the aero-engine and within the 20% margin range of the low-pressure rotor base frequency, if not, the corresponding fairing 1 structure meets the design requirements; if yes, adjust the taper end angle of the cone portion of the conical structure until the natural frequency of the fairing 1 exceeds the 20% margin range of the high-pressure rotor base frequency of the aero-engine, and the natural frequency of the fairing 1 exceeds the 20% margin range of the low-pressure rotor base frequency.

[0036] Step 5, in the finally obtained fairing 1 aerodynamic profile, the pipelines are uniformly arranged. The arrangement principle is "cold inside and hot outside". "Cold inside and hot outside" means that the fluid with low temperature in the pipeline 4 is arranged close to the middle of the fairing 1, and the hot one is close to the wall surface of the fairing 1. Since the fairing 1 is located in the outer-duct, the fluid flowing outside the fairing 1 has high speed and relatively low temperature, which is convenient for taking away the heat of the wall surface of the fairing 1, so as to reduce the problem of excessive thermal stress inside the fairing 1, and solve the problem of deformation coordination of the fairing 1.

[0037] In the sixth step, the hollowed-out weight-reducing design is used to ensure that the wall thickness of the pipeline 4 and the outer surface is 1.0mm-2.0mm, and a reinforcing rib is added between the adjacent two pipelines 4, and a reinforcing rib is added between the pipeline 4 and the wall surface of the fairing 1, the wall thickness of the reinforcing rib is 4-6mm, the intersection between the reinforcing rib and the wall surface and the pipeline 4 needs to be rounded, the rounding radius is not less than R1, and the rest of the positions are designed to be hollowed out. The reinforcing rib can enhance the overall rigidity of the fairing 1 and improve the overall heat conduction effect; the hollowed-out design can further ensure the heat dissipation of the high-temperature pipeline 4 in the fairing 1 and solve the deformation coordination problem of the fairing 1 during installation.

[0038] In the seventh step, the airfoil mounting surface is designed. The fairing 1 is in the outer channel flow passage 2, and bears the aerodynamic load of the outer channel airflow, the larger thermal stress and thermal deformation of the high-temperature pipeline 4 inside the fairing 1, and the vibration load of the whole machine. In order to solve the above problems, the fairing 1 is provided with a flange edge for axial and radial fixation. It is particularly pointed out that the radial mounting edge is designed in the form of an airfoil, the wall thickness is ensured to be within 2mm, and the cantilever is 20-30mm, so that the fairing 1 can play a damping buffer role under the action of the larger overall load of the machine, and the vibration stress is reduced. The top surface of the fairing 1 is fixed to the inner wall surface of the outer channel casing through the radial mounting flange 6, and the mounting hole in the mounting flange 6 is a waist-shaped hole extending in the axial direction. Axial displacement is allowed, and the problem of axial deformation incoordination can be solved.

[0039] The axial mounting hole 5 is arranged on the axial mounting edge of the fairing 1, the axial mounting hole 5 is fixedly connected with the intermediate casing mounting edge, and the axial mounting hole 5 is a waist-shaped hole extending in the radial direction of the aero-engine. Radial displacement is allowed, and by arranging the axial mounting hole 5 as a waist-shaped hole extending in the radial direction, the problem of radial deformation incoordination caused by the high temperature of the internal pipeline 4 can be avoided.

[0040] In the eighth step, the maximum temperature difference between the inside and outside of the fairing 1 and the maximum aerodynamic force are selected to evaluate the strength, the deformation and stress distribution of the fairing 1 are evaluated, and the long-term allowable stress requirement of the material is ensured to be met, otherwise the second step to the eighth step are iterated.

[0041] In the ninth step, since the pipeline 4 needs to be inserted into the outer channel, the fairing 1 and the outer channel casing must have mechanical contact, and there is a risk of outer channel gas leakage. In order to improve the air tightness of the mounting structure of the fairing 1 and ensure the aerodynamic performance of the outer channel, a sealing gasket 7 is added to the cooperation surface between the fairing 1 and the casing mounting edge.

[0042] In the tenth step, the integral casting technology or 3D printing technology is used to complete the processing and manufacturing.

[0043] The above is only a preferred embodiment of the present application, and is not intended to limit the present application. Any modification, equivalent replacement and improvement made within the spirit and principle of the present application shall be included in the protection scope of the present application.

Claims

1. A design method for a pipeline fairing, wherein the fairing is arranged radially along an aero-engine, and the radial cross-section of the fairing is a conical structure with curvature, the conical opening end of the fairing is fixedly connected to an intermediate casing support plate, and the conical part of the fairing is located downstream of the intermediate casing support plate. The fairing is internally configured as a hollow structure that allows pipes to pass through; characterized in that, Design methods include: Based on the dimensional parameters of the aero-engine bypass duct and the surface structure of the fairing in the radial section, a finite element analysis model of the fairing is established. The natural frequency of the fairing is obtained by performing simulation analysis on the finite element analysis model. Determine whether the natural frequency is within the 20% margin range of both the high-pressure rotor and low-pressure rotor fundamental frequencies of the aero-engine. If not, the corresponding pipeline fairing structure meets the design requirements. If so, adjust the tapered end angle of the conical structure until the natural frequency of the fairing exceeds the 20% margin range of both the high-pressure rotor and low-pressure rotor fundamental frequencies. The spline curve of the conical shape with curvature on the radial section of the fairing is... ,in The width of the rear end of the intermediate casing support plate. The cone-shaped tail angle of the fairing. Let y be the ordinate of a Cartesian coordinate system. This Cartesian coordinate system is established with the cone-shaped structure's cone portion as the origin, the tangent at the cone portion as the X-axis, and the line passing through the cone portion and perpendicular to the X-axis as the Y-axis. The x-axis is... The range of values ​​is , The circumferential width of the rear end of the intermediate casing support plate in front of the fairing.

2. The design method according to claim 1, characterized in that, The cone-shaped tail angle of the fairing .

3. The design method according to claim 1, characterized in that, The fairing has a wall thickness of 1.0~2.0mm.

4. The design method according to claim 1, characterized in that, The fairing is provided with an axial mounting hole, which is fixedly connected to the mounting edge of the intermediate casing, and the axial mounting hole is an oblong hole extending radially along the aero-engine.

5. The design method according to claim 1, characterized in that, The top surface of the fairing is fixed to the inner wall of the outer bypass casing by a mounting flange, and the mounting hole on the mounting flange is an axially extending waist-shaped hole.

6. The design method according to claim 5, characterized in that, A sealing gasket is provided at the position where the fairing mates with the inner wall of the outer bypass casing.

7. A pipeline shroud, characterized in that, The fairing is obtained by the design method described in any one of claims 1-6.

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