A method for correcting aerodynamic loads on the wing of a straight-wing aircraft

Abstract

The invention discloses a method for correcting aerodynamic loads on a wing of a straight-wing aircraft, comprising: S1: obtaining relevant information of a wing airfoil used by the straight-wing aircraft; S2: determining a shape equation of the wing airfoil according to the relevant information of the wing airfoil; S3: modeling the straight-wing aircraft according to the shape equation and the span of the wing to obtain a simulation model; S4: calculating an angle α between the aerodynamic load and the vertical direction according to the shape equation. X The cosine value of cosα X ; S5: Establish the modified maximum aerodynamic load expression #imgabs0# S6: Based on #imgabs1# and cosine value cosα X , construct the component expression of the aerodynamic load in front of the main beam along the vertical direction #imgabs2# and the component expression of the aerodynamic load behind the main beam along the vertical direction #imgabs3#S7: According to #imgabs4# and #imgabs5#, establish the equilibrium equation of the aerodynamic load on the wing along the vertical direction; S8: According to the equilibrium equation, obtain the result of the aerodynamic load correction coefficient δ; S9: According to the result of the aerodynamic load correction coefficient δ, correct the aerodynamic load of the wing of the simulation model.

Description

Technical Field

[0001] The present invention relates to the technical field of aerodynamic load analysis of aircraft wings, and particularly to a method for correcting aerodynamic loads on the wings of a straight-wing aircraft. Background Art

[0002] The strength analysis of an aircraft wing is of crucial value for measuring the working performance of the wing. An effective means for analyzing the strength of an aircraft wing is to establish its finite element simulation model with the aid of finite element simulation software and calculate its working conditions under actual working conditions based on the finite element simulation software.

[0003] During flight, the wing of an aircraft is subjected to aerodynamic loads, gravitational loads, and thrust loads. In the case of wing-mounted equipment, the wing is also subjected to concentrated loads. Generally, assuming that all the lift during flight comes from the wing and the influence of concentrated loads is not considered, the resultant force of the aerodynamic loads on the wing in the vertically upward direction is equal to the product of the aircraft maneuvering overload and the aircraft mass. For a straight-wing aircraft, the aerodynamic load distribution on a single wing can be approximated as an elliptical distribution along the spanwise direction and a triangular distribution along the chordwise direction, from which the aerodynamic load distribution result on the wing can be obtained. This aerodynamic load distribution result assumes that the direction of the applied aerodynamic load is vertically upward and the resultant force magnitude is half of the product of the aircraft maneuvering overload and the aircraft mass. However, when directly applying aerodynamic loads to the wing using finite element simulation software such as Abaqus, the direction of the applied aerodynamic load is perpendicular to the wing skin. Therefore, directly using the above aerodynamic load distribution result to apply aerodynamic loads to the wing will cause the resultant force in the vertical direction borne by the aircraft to be non-zero, thereby introducing analysis errors. Summary of the Invention

[0004] The purpose of the present invention is to provide a method for correcting aerodynamic loads on the wings of a straight-wing aircraft to correct the aerodynamic load distribution and combine the information on the angle between the aerodynamic load and the vertical direction, so that the force balance condition in the vertical direction can be satisfied after directly applying aerodynamic loads to the wing using finite element simulation software such as Abaqus.

[0005] The technical solution of the present invention for solving the above technical problems is as follows:

[0006] The present invention provides a method for correcting aerodynamic loads on the wings of a straight-wing aircraft, and the method for correcting aerodynamic loads on the wings of a straight-wing aircraft includes:

[0007] S1: Obtain relevant information on the airfoil used by the straight-wing aircraft;

[0008] S2: Determine the shape equation of the airfoil according to the relevant information on the airfoil;

[0009] S3: Model the straight-wing aircraft according to the shape equation of the wing airfoil and the wingspan of the wing to obtain a simulation model;

[0010] S4: Calculate the cosine value cosα of the angle α between the aerodynamic load and the vertical direction according to the shape equation of the wing airfoil X ; X ;

[0011] S5: Introduce an aerodynamic load correction coefficient δ into the maximum aerodynamic load expression P max to establish a modified maximum aerodynamic load expression

[0012] S6: Construct the vertical component force expression of the aerodynamic load in front of the main beam and the vertical component force expression of the aerodynamic load behind the main beam according to the maximum aerodynamic load expression X ; and the vertical component force expression of the aerodynamic load behind the main beam

[0013] S7: Establish the balance equation of the aerodynamic load on the wing in the vertical direction according to the vertical component force expression of the aerodynamic load in front of the main beam and the vertical component force expression of the aerodynamic load behind the main beam ;

[0014] S8: Obtain the result of the aerodynamic load correction coefficient δ according to the balance equation of the aerodynamic load on the wing in the vertical direction;

[0015] S9: Correct the aerodynamic load of the wing of the simulation model according to the result of the aerodynamic load correction coefficient δ.

[0016] Optionally, in step S2, the shape equation y N (x N ) of the wing airfoil is:

[0017]

[0018] where x N represents the distance along the horizontal axis between the point on the airfoil and the leading edge in the relevant information of the wing airfoil.

[0019] Optionally, step S4 includes:

[0020] S41: Differentiate the shape equation of the wing airfoil to obtain the slope k X expression of the tangent line of the wing airfoil at X;

[0021] S42: According to the slope k X, obtain the included angle β between the tangent line at X and the horizontal axis X ;

[0022] S43: Calculate the included angle α between the aerodynamic load and the vertical direction according to the included angle β X and the aircraft elevation angle θ X ;

[0023] S44: Calculate the sine value sinβ X of the included angle β X and the cosine value cosβ X ;

[0024] S45: According to the sine value sinβ X of the included angle β X and the cosine value cosβ X , as well as the sine and cosine values of the aircraft elevation angle, obtain the cosine value cosα X of the included angle α between the aerodynamic load and the vertical direction X .

[0025] Optionally, in the step S41, the slope k X of the tangent line of the airfoil at X

[0026]

[0027] Optionally, in the step S43, the included angle α X between the aerodynamic load and the vertical direction

[0028]

[0029] wherein, when the slope k X of the tangent line of the airfoil at X X satisfies k X ≥0, the included angle between the tangent line and the positive direction of the X-axis is denoted as β X ; when the slope k X of the tangent line of the airfoil at X X <0, the included angle between the tangent line and the negative direction of the X-axis is denoted as β

[0030] Optionally, the step S44 includes:

[0031] S441: Calculate the value of cos2β X using the universal formula;

[0032] S442: According to the value of cos2β X , use the half-angle formula to obtain the cosine value cosβ X of the included angle β X ;

[0033] S443: Calculate the sine value sinβ of the included angle β based on the cosine value cosβ of the included angle β X X X X .

[0034] Optionally, in the step S45, the cosine value cosα of the included angle α between the aerodynamic load and the vertical direction X is: X

[0035]

[0036] Optionally, in the step S5, the maximum aerodynamic load expression P max is:

[0037]

[0038] The modified maximum aerodynamic load expression is:

[0039]

[0040] where n0 represents the aircraft maneuver overload, m represents the aircraft mass, l represents the wingspan of the aircraft wing, c represents the chord length of the airfoil, π represents the pi, and Z represents the position coordinate along the spanwise direction.

[0041] Optionally, in the step S6, the vertical component force expression of the aerodynamic load in front of the main beam is:

[0042]

[0043] The vertical component force expression of the aerodynamic load behind the main beam is:

[0044]

[0045] where is the modified maximum aerodynamic load expression, X is the position coordinate along the chordwise direction, cosα X is the cosine value of the included angle α between the aerodynamic load and the vertical direction X , c represents the chord length of the airfoil, κ represents the percentage of the part in front of the main beam in the chord length, and Z represents the position coordinate along the spanwise direction.

[0046] Optionally, in the step S7, the balance equation of the aerodynamic load on the wing in the vertical direction is:

[0047]

[0048] ​​​​Among them, is the expression of the vertical component of the aerodynamic load on the front main beam, is the expression of the vertical component of the aerodynamic load on the rear main beam. Here, l represents the wingspan of the aircraft wing, c represents the chord length of the airfoil, κ represents the percentage of the part in front of the main beam in the chord length, X is the position coordinate along the chord direction, and Z represents the position coordinate along the span direction.

[0049] The present invention has the following beneficial effects:

[0050] By introducing an aerodynamic load correction coefficient and combining the force balance equation and the shape equation of the airfoil, the present invention can effectively correct the aerodynamic load on the wing, solve the problem that the resultant force in the vertical direction is not zero when directly loading the aerodynamic load through a finite element simulation software, and reduce the analysis error. Description of the Drawings

[0051] Figure 1 is the flowchart of the method for correcting the aerodynamic load on the wing of a straight-wing aircraft according to the present invention;

[0052] Figure 2 is the NACA0012 symmetric airfoil diagram of the present invention;

[0053] Figure 3 is the finite element simulation model of the straight wing of the present invention;

[0054] Figure 4 is the schematic diagram of applying the aerodynamic load on the straight wing of the present invention;

[0055] Figure 5 is the effect diagram of applying the aerodynamic load on the straight wing of the present invention. Detailed Embodiments

[0056] The principles and features of the present invention will be described below with reference to the drawings. The examples given are only for explaining the present invention and are not intended to limit the scope of the present invention.

[0057] The present invention provides a method for correcting the aerodynamic load on the wing of a straight-wing aircraft. Referring to Figure 1 shown, the method for correcting the aerodynamic load on the wing of the straight-wing aircraft includes:

[0058] S1: Obtain relevant information about the airfoil used by the straight-wing aircraft;

[0059] The airfoil of the straight-wing aircraft selected in the present invention is the NACA0012 symmetric airfoil, with a chord length c = 1000 mm, a maximum thickness of 0.12c, and the maximum thickness position is at 0.3c from the leading edge.

[0060] S2: Determine the shape equation of the airfoil according to the relevant information of the airfoil;

[0061] The airfoil of the straight-wing aircraft of the present invention is as follows Figure 2 shown, and the shape equation of the airfoil is y N (x N ) is as follows. Set the wingspan l of the aircraft wing to 4000 mm and establish a finite element simulation model of the straight wing based on finite element analysis software, as Figure 3 shown (part of the skin is hidden to show the internal structure).

[0062] The shape equation of the airfoil y N (x N ) is:

[0063]

[0064] where x N represents the distance along the horizontal axis between the point on the airfoil and the leading edge in the relevant information of the airfoil. Among them, x N ∈[0, c], and y N (x N ) represents half of the airfoil thickness at x N .

[0065] In some embodiments, if there is no expression of the shape equation of the airfoil, a surrogate model expression of the airfoil shape can be constructed based on the characteristic point information on the airfoil and in combination with the Kriging surrogate model.

[0066] S3: Model the straight-wing aircraft according to the shape equation of the airfoil and the wingspan of the wing to obtain a simulation model;

[0067] The present invention uses finite element analysis software for simulation. Those skilled in the art can select appropriate software for simulation according to the present invention, and the present invention does not make specific limitations.

[0068] S4: Calculate the cosine value cosα X of the angle α X between the aerodynamic load and the vertical direction according to the shape equation of the airfoil;

[0069] Optionally, the step S4 includes:

[0070] S41: Perform a derivative process on the shape equation of the airfoil to obtain the slope k X expression of the tangent line of the airfoil at X;

[0071] The slope k X expression of the tangent line of the airfoil at X is:

[0072]

[0073] S42: According to the slope k of the tangent line of the airfoil at X X , obtain the angle β between the tangent line at X and the X-axis X ;

[0074] S43: Calculate the angle α between the aerodynamic load and the vertical direction according to the angle β X and the aircraft elevation angle θ X ;

[0075] When the flight elevation angle is θ, the aerodynamic load of the wing structure is applied as Figure 4 shown. The angle α between the aerodynamic load and the vertical direction X is:

[0076]

[0077] where, when the slope k of the tangent line of the airfoil at X X satisfies k X ≥0, denote the angle between the tangent line and the positive direction of the X-axis as β X ; when the slope k of the tangent line of the airfoil at X X satisfies k X <0, denote the angle between the tangent line and the negative direction of the X-axis as β X .

[0078] S44: Calculate the sine value sinβ X and cosine value cosβ X of the angle β X ;

[0079] Optionally, the step S44 includes:

[0080] S441: Calculate the value of cos2β X using the universal formula;

[0081] That is:

[0082] S442: According to the value of cos2β X , use the half-angle formula to obtain the cosine value cosβ X of the angle β X ;

[0083]

[0084] S443: According to the cosine value cosβ X of the angle β X , calculate the sine value sinβ X of the angle β X

[0085] That is:

[0086] S45: According to the sine value sinβ X and cosine value cosβ X of the included angle β X , as well as the sine and cosine values of the aircraft elevation angle, obtain the cosine value cosα X of the included angle α between the aerodynamic load and the vertical direction X .

[0087] The cosine value cosα X of the included angle α between the aerodynamic load and the vertical direction X is:

[0088]

[0089] Table 1 shows the results of cosα X calculated at different positions X and elevation angles θ

[0090] Table 1 Results of cosα X at different positions X and elevation angles θ

[0091]

[0092]

[0093] S5: Introduce the aerodynamic load correction coefficient δ into the maximum aerodynamic load expression P max to establish the corrected maximum aerodynamic load expression

[0094] The maximum aerodynamic load expression P max is:

[0095]

[0096] The corrected maximum aerodynamic load expression is:

[0097]

[0098] where n0 represents the aircraft maneuvering overload, m represents the aircraft mass, l represents the wingspan of the aircraft wing, c represents the chord length of the airfoil, π represents the pi, and Z represents the position coordinate along the spanwise direction

[0099] S6: According to the maximum aerodynamic load expression and the cosine value cosα X , construct the vertical component force expression of the aerodynamic load in front of the main beam and the vertical component force expression

[0100] The expression of the vertical component of the aerodynamic load on the front of the main beam is as follows:

[0101]

[0102] The expression of the vertical component of the aerodynamic load on the rear of the main beam is as follows:

[0103]

[0104] Among them, is the expression of the corrected maximum aerodynamic load, X is the position coordinate along the chord direction, and cosα X is the cosine value of the angle α between the aerodynamic load and the vertical direction X c represents the chord length of the airfoil of the wing, κ represents the percentage of the part in front of the main beam in the chord length, and Z represents the position coordinate along the span direction.

[0105] S7: According to the expression of the vertical component of the aerodynamic load on the front of the main beam and the expression of the vertical component of the aerodynamic load on the rear of the main beam establish the balance equation of the aerodynamic load on the wing in the vertical direction;

[0106] The balance equation of the aerodynamic load on the wing in the vertical direction is as follows:

[0107]

[0108] Among them, is the expression of the vertical component of the aerodynamic load on the front of the main beam, is the expression of the vertical component of the aerodynamic load on the rear of the main beam, l represents the span of the aircraft wing, c represents the chord length of the airfoil of the wing, κ represents the percentage of the part in front of the main beam in the chord length, X is the position coordinate along the chord direction, and Z represents the position coordinate along the span direction.

[0109] S8: According to the balance equation of the aerodynamic load on the wing in the vertical direction, obtain the result of the aerodynamic load correction coefficient δ;

[0110] Solving the above balance equation, the results of the aerodynamic load correction coefficient δ at different elevation angles θ are shown in Table 2.

[0111] Table 2 Results of the aerodynamic load correction coefficient δ at different elevation angles θ

[0112] θ / degree Correction coefficient δ 0 1.00358 1 1.00420 2 1.00512 3 1.00636 4 1.00791 5 1.00977 6 1.01194 7 1.01444 8 1.01726 9 1.02040 10 1.02388

[0113] The table shows that the value of the correction coefficient is the smallest at an elevation angle of 0°, and as the elevation angle increases, the value of the correction coefficient also increases. This indicates that in the case of a small elevation angle, the component of the aerodynamic load applied to the wing in the vertical direction accounts for a relatively large proportion overall. Therefore, only a small correction coefficient is required to correct the aerodynamic load. However, as the elevation angle increases, the proportion of the component of the aerodynamic load applied to the wing in the vertical direction decreases overall. Therefore, a larger correction coefficient is required to correct it.

[0114] S9: Correct the aerodynamic load on the wing of the simulation model according to the result of the aerodynamic load correction coefficient δ.

[0115] Assume that the aircraft maneuvering overload is n0 = 1.25g, where g = 9800 mm / s 2 represents the acceleration due to gravity, the mass of the aircraft is m = 20 tons, and the distribution and magnitude of the aerodynamic load on the upper and lower surfaces of the wing are the same. Then the aerodynamic loads on the front upper and lower wing surfaces of the main beam are respectively The aerodynamic loads on the rear upper and lower wing surfaces of the main beam are respectively If the aircraft elevation angle is taken as θ = 0°, then the correction coefficient is δ = 1.00358, and the maximum corrected aerodynamic loads at the positions of the main beam on the upper and lower wing surfaces of the wing root section are respectively Fix the displacement degrees of freedom and rotation degrees of freedom in three directions of the wing rib at the wing root as boundary conditions, and apply aerodynamic loads with magnitudes of on the front and rear of the main beam on the upper and lower wing surfaces respectively. The direction is upward perpendicular to the skin. The effect diagram after applying the load is shown in Figure 5 .

[0116] The above are only the preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention shall be included within the protection scope of the present invention.

Claims

1. A method for correcting aerodynamic loads on the wing of a straight-wing aircraft, characterized in that, The method for correcting the aerodynamic load on the wing of a straight-wing aircraft includes: S1: Obtain the relevant information of the airfoil used by the straight-wing aircraft; S2: Determine the shape equation of the airfoil according to the relevant information of the airfoil; S3: Model the straight-wing aircraft according to the shape equation of the airfoil and the wingspan of the wing to obtain a simulation model; S4: Calculate the cosine value cosα of the angle α between the aerodynamic load and the vertical direction according to the shape equation of the wing airfoil. X X ; S5: Introduce the aerodynamic load correction coefficient δ into the maximum aerodynamic load expression P max to establish the corrected maximum aerodynamic load expression The maximum aerodynamic load expression P max is as follows: The expression of the corrected maximum aerodynamic load is as follows: where n0 represents the aircraft maneuvering overload, m represents the aircraft mass, l represents the wingspan of the aircraft wing, c represents the chord length of the airfoil, π represents pi, and Z represents the position coordinate along the spanwise direction; S6: According to the maximum aerodynamic load expression and the cosine value cosα X , construct the expression for the vertical component of the aerodynamic load on the front of the main beam and the expression for the vertical component of the aerodynamic load on the rear of the main beam The expression of the vertical component of the aerodynamic load in front of the main beam is as follows: The expression of the vertical component of the aerodynamic load behind the main beam is as follows: Among them, is the corrected maximum aerodynamic load expression, X is the position coordinate along the chord direction, and cosα X is the angle α between the aerodynamic load and the vertical direction X is the cosine value, c represents the chord length of the airfoil, κ represents the percentage of the part in front of the main beam in the chord length, and Z represents the position coordinate along the span direction; S7: Establish the balance equation of the aerodynamic load on the wing in the vertical direction according to the expression of the vertical component force of the front aerodynamic load of the main beam and the expression of the vertical component force of the rear aerodynamic load of the main beam ​ The equilibrium equation of the aerodynamic load on the wing in the vertical direction is: Among them, is the expression of the vertical component of the aerodynamic load in front of the main beam, is the expression of the vertical component of the aerodynamic load behind the main beam. l represents the wingspan of the aircraft wing, c represents the chord length of the airfoil, κ represents the percentage of the part in front of the main beam occupying the chord length, X is the position coordinate along the chord direction, and Z represents the position coordinate along the span direction; S8: Obtain the result of the aerodynamic load correction coefficient δ according to the equilibrium equation of the aerodynamic load on the wing in the vertical direction; S9: Correct the aerodynamic load on the wing of the simulation model according to the result of the aerodynamic load correction coefficient δ.

2. The method for correcting the aerodynamic load on the wing of a straight-wing aircraft according to claim 1, wherein In the step S2, the shape equation y N (x N ) is as follows: where x N represents the distance along the horizontal axis between a point on the airfoil and the leading edge in the relevant information of the airfoil of the wing.

3. The method for correcting the aerodynamic load on the straight-wing aircraft wing according to claim 1, wherein The step S4 includes: S41: Differentiate the shape equation of the airfoil to obtain the slope k of the tangent line of the airfoil at X X Expression S42: According to the slope k of the tangent line of the airfoil at X X , obtain the angle β between the tangent line at X and the X-axis X ; S43: Calculate the angle α between the aerodynamic load and the vertical direction according to the included angle β X and the aircraft elevation angle θ X ; S44: Calculate the included angle β X whose sine value is sinβ X and cosine value is cosβ X ; S45: According to the sine value sinβ X and the cosine value cosβ X of the included angle β, as well as the sine value and cosine value of the elevation angle of the aircraft, obtain the cosine value cosα X of the included angle α between the aerodynamic load and the vertical direction X . X .

4. The method for correcting the aerodynamic load on the wing of a straight-wing aircraft according to claim 3, characterized in that In the step S41, the slope k of the tangent line of the airfoil at X X The expression is as follows:

5. The method for correcting the aerodynamic load on the straight-wing aircraft wing according to claim 3, wherein In the step S43, the included angle α between the pneumatic load and the vertical direction X is Among them, when the slope k of the tangent line of the airfoil at X X satisfies k X ≥0, the angle between the tangent line and the positive direction of the X-axis is denoted as β X ; when the slope k of the tangent line of the airfoil at X X satisfies k X <0, the angle between the tangent line and the negative direction of the X-axis is denoted as β X .

6. The method for correcting the aerodynamic load on the straight-wing aircraft wing according to claim 3, wherein The step S44 includes: S441: Calculate the value of cos2β using the universal formula X ; S442: According to the value of the said cos2β, use the half - angle formula to obtain the cosine value cosβ of the said included angle β X ; X ; X ; S443: Calculate the sine value sinβ of the included angle β according to the cosine value cosβ of the included angle β X X X X .​​​ 7. The method for correcting the aerodynamic load on the wing of a straight-wing aircraft according to claim 3, characterized in that, In the step S45, the cosine value cosα of the included angle α between the pneumatic load and the vertical direction X is X as follows:

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