A method for determining the distributed load of an aerodynamic surface with any shape on a helicopter

Abstract

Technical solution of the present invention: A method for determining the distributed load of an arbitrary-shaped aerodynamic surface of a helicopter, the method comprising: establishing a spanwise distribution function based on a local coordinate system according to the aerodynamic load distribution law, and calculating a spanwise distributed load coefficient k according to the spanwise distribution function; establishing a chordwise distribution function based on a local coordinate system according to the aerodynamic load distribution law, and calculating a chordwise load distribution parameter h according to the chordwise distribution function; obtaining a surface distribution function of the aerodynamic load according to the spanwise distributed load coefficient k and the chordwise load distribution parameter h; converting the surface distribution function based on the local coordinate system into a surface distribution function based on the body coordinate system.

Description

Technical Field

[0001] This patent relates to the field of aircraft strength design and verification, and relates to a method for determining the distributed load of an aerodynamic surface with any shape on a helicopter. Background Art

[0002] The aerodynamic surface structures of a helicopter fuselage mainly include a horizontal tail, a vertical tail, and short wings, etc. When a helicopter is flying forward, it should ensure certain static stability of the angle of attack, static stability of speed, static stability of heading, dihedral effect, and angular velocity damping. Therefore, it is necessary to install a horizontal tail surface, etc. to improve the static instability of the angle of attack. With the continuous development of new configurations such as coaxial dual rotors and tiltrotors, the aerodynamic surface design and functions of traditional helicopters have gradually become diversified. Structures such as the wings of tiltrotors and the tail control surfaces of fixed-wing aerodynamic stabilizers have emerged. Different from traditional helicopters, these functions pose new challenges to the strength and stiffness design of aerodynamic surfaces with any shape. Therefore, more accurate load input is required for static strength and dynamic analysis based on refined modeling.

[0003] At present, the calculation of aerodynamic loads on domestic helicopters mainly relies on the regulations in the helicopter load manual. Based on the results of fluid mechanics or CFD calculations, the position of the center of pressure and the total equivalent concentrated load are given, and the load coefficient of the field is determined through simulation. However, this method fails to establish relevant theories and lacks sufficient verification of the results. Summary of the Invention

[0004] The present invention provides a method for determining the distributed load of an aerodynamic surface with any shape on a helicopter to solve or at least alleviate the problem of accurately defining the distributed load of an aerodynamic surface with any shape in the background art.

[0005] The technical solution of the present invention: A method for determining the distributed load of an aerodynamic surface with any shape on a helicopter, the method includes:

[0006] According to the aerodynamic load distribution law, establish a spanwise distribution function based on a local coordinate system, and calculate the spanwise distribution load coefficient k according to the spanwise distribution function;

[0007] According to the aerodynamic load distribution law, establish a chordwise distribution function based on a local coordinate system, and calculate the chordwise load distribution parameter h according to the chordwise distribution function;

[0008] According to the spanwise distribution load coefficient k and the chordwise load distribution parameter h, obtain the surface distribution function of the aerodynamic load;

[0009] Convert the surface distribution function based on the local coordinate system to the surface distribution function based on the airframe coordinate system.

[0010] Specifically, the local coordinate system has an arbitrary wing root endpoint of the two-dimensional projection plane of any aerodynamic surface of the helicopter as the origin, with the chordwise direction as the X-axis and the spanwise direction as the Y-axis.

[0011] Specifically, according to the aerodynamic load distribution law, a spanwise distribution function based on the local coordinate system is established, and the spanwise distribution load coefficient k is calculated according to the spanwise distribution function, which specifically includes:

[0012] According to the aerodynamic load distribution law, the total load is proportional to the chord length of the aerodynamic surface along the spanwise direction, and the spanwise distribution function P(y) = kb(y) is obtained, where b(y) is the chord length and k is the spanwise load distribution coefficient;

[0013] Integrate the spanwise distribution function along the spanwise direction to obtain the total load P of the aerodynamic surface;

[0014] According to the total load P of the aerodynamic surface and the total area A of the aerodynamic surface with any shape, use the formula Calculate the spanwise distribution load coefficient k.

[0015] Specifically, according to the aerodynamic load distribution law, a chordwise distribution function based on the local coordinate system is established, and the chordwise load distribution parameter h is calculated according to the chordwise distribution function, which specifically includes:

[0016] According to the chordwise distribution law of the aerodynamic load, its chordwise load distribution along the chordwise direction can list the chordwise load distribution function of any spanwise micro-segment:

[0017]

[0018] According to the total load P(y) and chord length b(y) of the spanwise micro-segment, use the formula Calculate the chordwise load distribution parameter h.

[0019] Specifically, according to the spanwise distribution load coefficient k and the chordwise load distribution parameter h, the surface distribution function of the aerodynamic load is obtained, which specifically includes:

[0020] According to the spanwise distribution load coefficient k and the chordwise load distribution parameter h, the surface distribution function of the aerodynamic load is obtained:

[0021]

[0022] Specifically, the spanwise load distribution law is specifically: the one-dimensional distribution of the integral result of the total load along the chordwise direction along the spanwise direction is proportional to the chord length value corresponding to any spanwise direction.

[0023] Specifically, the chordwise load distribution law is specifically: the load of any chordwise section is uniformly distributed from the leading edge to 1 / 10 of the chord length and linearly distributed from 1 / 10 of the chord length to the trailing edge.

[0024] Specifically, after obtaining the surface distribution function based on the body coordinate system, the method further includes:

[0025] Calculating the surface distribution function using the PCL method.

[0026] In summary, based on the principle of aerodynamic force action, by constructing a spatial field function, the present invention establishes a definition method for the distributed load of an aerodynamic surface of any shape. At the same time, taking the horizontal tail of a certain helicopter as an example for finite element simulation, the definition method for the distributed load of an aerodynamic surface of any shape is verified, and a general method for loading the distributed load of an aerodynamic surface of any shape is established. Brief Description of the Drawings

[0027] Figure 1 is a flowchart of a method for determining the distributed load of an aerodynamic surface of any shape of a helicopter according to the present invention;

[0028] Figure 2 is a schematic diagram of the distribution of the load along the chord direction provided by the present invention;

[0029] Figure 3 is a local coordinate system provided by the present invention;

[0030] Figure 4 is a curve of the spanwise distribution of the aerodynamic surface pressure provided by the present invention;

[0031] Figure 5 is a curve of the chordwise distribution of the aerodynamic surface pressure provided by the present invention. Detailed Embodiment

[0032] Based on the principle of aerodynamic force action, by constructing a spatial field function, the present invention establishes a definition method for the distributed load of an aerodynamic surface of any shape. At the same time, taking the horizontal tail of a certain helicopter as an example for finite element simulation, the definition method for the distributed load of an aerodynamic surface of any shape is verified, and a general method for loading the distributed load of an aerodynamic surface of any shape is established.

[0033] The present invention provides a method for determining the distributed load of an aerodynamic surface of any shape of a helicopter to solve or at least mitigate the problem of accurately defining the distributed load of an aerodynamic surface of any shape in the background art.

[0034] Embodiment 1

[0035] The technical solution adopted by the present invention is: providing a method for determining the distributed load of an aerodynamic surface of any shape of a helicopter, including the following steps:

[0036] Step 101: According to the aerodynamic load distribution law, establish a spanwise distribution function based on the local coordinate system, and calculate the spanwise distribution load coefficient k according to the spanwise distribution function;

[0037] Among them, the local coordinate system takes an arbitrary wing root endpoint of the two-dimensional projection plane of any aerodynamic surface of the helicopter as the origin, with the chord direction as the X-axis and the span direction as the Y-axis. The spanwise load distribution law of the aerodynamic load distribution law is specifically that the one-dimensional distribution of the total load along the span direction obtained by integrating along the chord direction is proportional to the chord length value corresponding to any span.

[0038] In practical applications, according to the helicopter load manual, the aerodynamic load distribution law can be obtained.

[0039] Specifically, step 101 includes:

[0040] Step 1011: According to the aerodynamic load distribution law, since the total load along the span is proportional to the chord length of the aerodynamic surface, the spanwise distribution function P(y) satisfies:

[0041] P(y) = kb(y) (1)

[0042] In the formula, b(y) is the chord length, and k is the spanwise load distribution coefficient.

[0043] Step 1012: Integrate the spanwise distribution function along the span to obtain the total load P of the aerodynamic surface;

[0044] Step 1013: According to the total load P of the aerodynamic surface and the total area A of the aerodynamic surface with any shape, use the formula to calculate the spanwise distributed load coefficient k.

[0045] Step 102: According to the aerodynamic load distribution law, establish a chordwise distribution function based on the local coordinate system, and calculate the chordwise load distribution parameter h according to the chordwise distribution function;

[0046] See Figure 2 As shown, the aerodynamic load distribution law is that the load of any section is distributed along the chord of the section where it is located. The chordwise load distribution law is specifically that the load of any chordwise section is uniformly distributed from the leading edge to 1 / 10 of the chord length and linearly distributed from 1 / 10 of the chord length to the trailing edge.

[0047] Specifically, step 102 includes:

[0048] Step 1021: According to the chordwise distribution law of the aerodynamic load, obtain the chordwise load distribution function of any spanwise micro-segment:

[0049]

[0050] Step 1022: According to the total load P(y) and chord length b(y) of the spanwise micro-segment, use the formula to calculate the chordwise load distribution parameter h.

[0051] Step 103: Obtain the surface distribution function of the aerodynamic load according to the spanwise distributed load coefficient k and the chordwise load distribution parameter h:

[0052]

[0053] Step 104: Convert the surface distribution function based on the local coordinate system to the surface distribution function based on the airframe coordinate system:

[0054]

[0055] It should be noted that assuming that under the airframe coordinate system, the coordinates of the four endpoints of the aerodynamic surface are as Figure 3 shown, where the 0-1 side is the leading edge, taking point 0 as the origin, a local coordinate system is established as shown in Figure 3 .

[0056] Then the following relationships exist:

[0057] x = |X - X 01 (Y)|

[0058] b(y) = |X 23 (Y) - X 01 (Y)| (5)

[0059] In the formula, X and Y are the coordinate values under the airframe coordinate system, and X 01 (Y) and X 23 (Y) are the inverse functions of the line functions of the leading and trailing edge sides in the airframe coordinate system respectively.

[0060] Substitute formula (5) into formula (4), perform coordinate system conversion, and obtain the field function under the airframe coordinate system as:

[0061]

[0062] Preferably, after obtaining the surface distribution function based on the airframe coordinate system, the method further includes: calculating the surface distribution function using the PCL method.

[0063] In the finite element software, the field function is established using the PCL method, or the respective space field functions are established according to different regions.

[0064] In summary, the present application proposes a method for loading distributed loads on an aerodynamic surface of any shape of a helicopter, constructs a space field function. At the same time, through finite element simulation, the aerodynamic load loading of a certain helicopter horizontal tail is analyzed, and the results show that the distributed load defined by this method is in good agreement with the theoretical value, and it can be applied to the aerodynamic surface structures of any shape of helicopters, with high universality.

[0065] Example 2

[0066] In the embodiment of this application, the left horizontal tail of a certain type of helicopter is selected as the research object to verify the above method.

[0067] Step 1: Geometric parameters of the aerodynamic surface of the left horizontal tail of a certain type of helicopter

[0068] This horizontal tail is a trapezoidal aerodynamic surface, and the geometric parameters of the aerodynamic surface are shown in Table 1.

[0069] Table 1 Aerodynamic surface parameters

[0070]

[0071] Step 2: Definition of aerodynamic surface pressure

[0072] In the body coordinate system, from the coordinates of the four endpoints of the airfoil, the line functions of the leading edge and the trailing edge can be calculated. Using Equation (6), the load distribution field function of the aerodynamic surface is obtained. In the Patran software, the field functions are established respectively and loaded into different regions of the aerodynamic surface to complete the definition of aerodynamic loads.

[0073] Step 3: Distribution effect of aerodynamic loads along the spanwise direction

[0074] According to the overall load calculation results, the total load of the working condition is defined, and all degrees of freedom of any node on the wing root line are constrained for static calculation to obtain the structural load and internal force distribution. The results are processed, and the distribution curve of the aerodynamic load along the spanwise direction is as Figure 4 shown.

[0075] Step 4: Distribution effect of aerodynamic loads along the chordwise direction

[0076] The results are processed, and the chordwise distribution curve of the aerodynamic load is as Figure 5 shown.

[0077] Step 5: Comparison of calculation results with theoretical values

[0078] Using the distributed load defined by the above field function, the simulation results are basically consistent with the overall distribution law of the aerodynamic load, that is, linearly distributed along the spanwise direction and wedge-shaped distributed along the chordwise direction. The calculation result data are read, the total load, center of pressure position and spanwise distribution coefficient of the aerodynamic surface are calculated respectively, and the chordwise distribution coefficient of the load is calculated at the semi-span section position. The comparison of the calculation results with the theoretical values is shown in Table 2.

[0079] Table 2 Comparison of calculation results

[0080]

[0081] In summary, the present application proposes a method for loading distributed loads on the aerodynamic surfaces of arbitrary shapes of a helicopter and constructs a spatial field function. At the same time, the aerodynamic load loading of a certain helicopter horizontal tail is analyzed through finite element simulation. The results show that the distributed loads defined by this method are in good agreement with the theoretical values, can be applied to the aerodynamic surface structures of arbitrary shapes of helicopters, and have high universality.

Claims

1. A method for determining the distributed load of an arbitrary-shaped aerodynamic surface of a helicopter, characterized in that, the method includes: According to the aerodynamic load distribution law, obtain the spanwise distribution function P(y) = kb(y), where b(y) is the chord length and k is the spanwise load distribution coefficient; Integrate the spanwise distribution function along the span to obtain the total load P of the aerodynamic surface; According to the total aerodynamic surface load P and the total area A of the aerodynamic surface with any shape, the spanwise distribution load coefficient k is calculated using the formula ​ According to the chordwise distribution law of the aerodynamic load, list the chordwise load distribution function of any spanwise micro-segment: According to the total load P(y) and chord length b(y) of the spanwise micro-segment, use the formula to calculate the chordwise load distribution parameter h; Based on the spanwise distributed load coefficient k and the chordwise load distribution parameter h, obtain the surface distribution function of the aerodynamic load; Convert the surface distribution function based on the local coordinate system to the surface distribution function based on the airframe coordinate system; The local coordinate system takes any wing root endpoint of the two-dimensional projection surface of the arbitrary-shaped aerodynamic surface of the helicopter as the origin, the chordwise direction as the X-axis, and the spanwise direction as the Y-axis.

2. The method according to claim 1, characterized in that, the spanwise load distribution law is specifically: The one-dimensional distribution of the total load along the span obtained by chordwise integration is proportional to the chord length value corresponding to any span.

3. The method according to claim 1, characterized in that, the chordwise load distribution law is specifically: The load of any chordwise section is uniformly distributed from the leading edge to 1 / 10 of the chord length and linearly distributed from 1 / 10 of the chord length to the trailing edge.

4. The method according to claim 1, characterized in that, after obtaining the surface distribution function based on the airframe coordinate system, the method further includes: Using the PCL (Patran Command Language) method to calculate the surface distribution function.

Images