Rapid calculation method for structural quality characteristics in conceptual design stage of high-speed aircraft

Abstract

The invention relates to a quick calculation method for structural quality characteristics of a high-speed aircraft in a conceptual design stage, and belongs to the field of aircraft design. The structural mass characteristics comprise aircraft structural mass, structural mass center and rotational inertia. The method comprises the following steps: defining basic composition of a high-speed aircraft structure; giving a geometric shape of the high-speed aircraft; generating a high-speed aircraft surface grid; importing grids and identifying the grids of the fuselage, the wings and the empennage; respectively calculating the mass, the mass center and the rotational inertia of each component of the body, the wings and the empennage; respectively calculating the mass, the mass center and the rotational inertia of the body, the wings and the empennage; and calculating the total mass, center of mass and rotational inertia of the high-speed aircraft structure. The method can quantitatively analyze the influence of the density, thickness and other parameters of the structural material on the structural quality of the aircraft, can reduce the time cost of multidisciplinary optimization design in the conceptual design stage of the high-speed aircraft, and improves the calculation efficiency.

Description

Technical Field

[0001] The invention provides a method for quickly calculating structural quality characteristics in the conceptual design stage of a high-speed aircraft, and belongs to the field of aircraft design. Background Art

[0002] Hypersonic vehicles refer to aircraft with a flight speed exceeding 5 Ma. They can not only greatly reduce the time of transcontinental transportation of personnel and cargo, but also can be used as hypersonic weapons to quickly strike long-range targets. They are a current and future research hotspot in the aerospace field. Hypersonic vehicle geometry, aerodynamics, structural quality and other disciplines are tightly coupled with each other, and there are complex nonlinear relationships. The linear design method of traditional aerospace aircraft is difficult to meet the design requirements. Therefore, a multidisciplinary optimization design method for aircraft is proposed. In the conceptual design stage of hypersonic aircraft, the overall performance of many disciplines such as aircraft geometry, aerodynamics, and structural quality is considered as much as possible.

[0003] The structural mass characteristics of a high-speed aircraft are an important part of the total takeoff mass of the aircraft. The calculation of the structural mass characteristics of a high-speed aircraft depends on the geometric layout and needs to be considered at the beginning of the geometric layout design.

[0004] When calculating the structural quality characteristics of high-speed aircraft, the reference geometric layout must be determined first, and then the structural quality estimation adopts the component method. According to the empirical formula, the geometric shape parameters of the fuselage, wing, and tail are input to calculate the structural quality of the aircraft fuselage, wing, and tail, thereby obtaining the structural quality characteristics of the high-speed aircraft. However, during the aircraft design process, parameters such as the density of the aircraft structure material and the thickness of the structure will affect the structural quality characteristics of the aircraft. The current calculation method of the structural quality characteristics of the aircraft cannot reflect the relationship between the structural quality characteristics of the aircraft and the density of the structural material and the thickness of the structure, which will have an adverse effect on the aircraft optimization design. Summary of the invention

[0005] The problem solved by the technology of the present invention is: to overcome the shortcomings of the prior art and provide a method for quickly calculating the structural quality characteristics in the conceptual design stage of a high-speed aircraft, which can quantitatively analyze the influence of parameters such as structural material density and thickness on the structural quality of the aircraft, and at the same time can reduce the time cost of multidisciplinary optimization design in the conceptual design stage of a high-speed aircraft and improve calculation efficiency.

[0006] The technical solution of the present invention is:

[0007] A method for rapid calculation of structural mass characteristics in the conceptual design stage of a high-speed aircraft includes:

[0008] The basic structure of the high-speed aircraft is clarified, and the high-speed aircraft includes a fuselage, wings, and tail fins, wherein the fuselage includes a heat insulation layer, a skin, a bulkhead, and a truss; the wings and tail fins both include a heat insulation layer, a skin, a wing rib, and a wing spar;

[0009] Given the high-speed aircraft geometry;

[0010] Generate high-speed aircraft surface mesh;

[0011] Import the mesh and identify the fuselage, wing, and tail meshes;

[0012] Calculate the mass, center of mass, and moment of inertia of each component of the fuselage, wing, and tail respectively;

[0013] Calculate the mass, center of mass, and moment of inertia of the fuselage, wing, and tail wing respectively;

[0014] Calculate the total mass, center of mass, and moment of inertia of high-speed aircraft structures.

[0015] Preferably, the generated high-speed aircraft surface mesh is a quadrilateral structured mesh, and the surface mesh can be used for calculating the aerodynamic performance of the aircraft.

[0016] Preferably, importing the grid refers to importing all grid points of the aircraft surface grid.

[0017] Preferably, the method for identifying the fuselage, wing, and tail grids is as follows:

[0018] The aircraft reference point is set as the head vertex, the X axis is within the aircraft longitudinal symmetry plane and points to the tail, the Y axis is perpendicular to the aircraft longitudinal symmetry plane and points to the right, and the Z axis satisfies the right-hand rule and points upward;

[0019] The wing has the largest span-wise dimension, and the average Y coordinate of the wing block grid points is the largest. The tail is close to the tail of the fuselage, and the average X coordinate of the tail block grid points is the largest. The fuselage, wing, and tail grids can be identified by comparing the average coordinates of the grid points of each grid block.

[0020] Preferably, the fuselage components include a fuselage heat insulation layer, a fuselage skin, a fuselage bulkhead, and a fuselage girders;

[0021] The mass of the fuselage heat insulation layer is obtained by multiplying the sum of the areas of the grid units constituting the fuselage surface by the thickness of the fuselage heat insulation layer and the material density of the fuselage heat insulation layer; the center of mass of the fuselage heat insulation layer is obtained by dividing the sum of the products of the mass of the fuselage heat insulation layer grid units and the coordinates of the center points of the grid units by the mass of the fuselage heat insulation layer; the moment of inertia of the fuselage heat insulation layer is the sum of the moments of inertia of all the fuselage heat insulation layer grid units relative to the rotation axis of the center of mass of the fuselage heat insulation layer, and the moment of inertia of any fuselage heat insulation layer grid unit relative to the rotation axis of the center of mass of the fuselage heat insulation layer is obtained by multiplying the square of the vertical distance between the center point of the fuselage heat insulation layer grid unit and the rotation axis corresponding to the center of mass of the fuselage heat insulation layer by the mass of the fuselage heat insulation layer grid unit;

[0022] The mass of the fuselage skin is obtained by multiplying the sum of the areas of the grid units constituting the fuselage surface by the thickness of the fuselage skin and the density of the fuselage skin material; the core of the fuselage skin is obtained by dividing the sum of the products of the mass of the fuselage skin grid unit and the coordinates of the center point of the grid unit by the mass of the fuselage skin; the moment of inertia of the fuselage skin is the sum of the moments of inertia of all the fuselage skin grid units relative to the rotation axis of the core of the fuselage skin, and the moment of inertia of any fuselage skin grid unit relative to the rotation axis of the core of the fuselage skin is obtained by multiplying the square of the vertical distance between the center point of the fuselage skin grid unit and the corresponding rotation axis of the core of the fuselage skin by the mass of the fuselage skin grid unit;

[0023] The mass of the fuselage bulkhead frame is obtained by multiplying the sum of the lengths of the fuselage bulkhead frame grid units by the cross-sectional area of ​​the fuselage bulkhead frame and the material density of the fuselage bulkhead frame; the center of mass of the fuselage bulkhead frame is obtained by dividing the sum of the product of the mass of the fuselage bulkhead frame grid units and the coordinates of the center points of the fuselage bulkhead frame length units by the mass of the fuselage bulkhead frame; the moment of inertia of the fuselage bulkhead frame is the sum of the moments of inertia of all the fuselage bulkhead frame grid units relative to the rotation axis of the center of mass of the fuselage bulkhead frame, and the moment of inertia of any fuselage bulkhead frame grid unit relative to the rotation axis of the center of mass of the fuselage bulkhead frame is obtained by multiplying the square of the vertical distance between the center point of the fuselage bulkhead frame grid unit and the rotation axis corresponding to the center of mass of the fuselage bulkhead frame by the mass of the fuselage bulkhead frame grid unit;

[0024] The mass of the fuselage truss is obtained by multiplying the sum of the lengths of the fuselage truss grid units by the cross-sectional area of ​​the fuselage truss and the material density of the fuselage truss; the center of mass of the fuselage truss is obtained by dividing the sum of the product of the mass of the fuselage truss grid unit and the coordinates of the center point of the fuselage truss length unit by the mass of the fuselage truss; the moment of inertia of the fuselage truss is the sum of the moments of inertia of all fuselage truss grid units relative to the rotation axis of the center of mass of the fuselage truss, and the moment of inertia of any fuselage truss grid unit relative to the rotation axis of the center of mass of the fuselage truss is obtained by multiplying the square of the vertical distance between the center point of the fuselage truss grid unit and the rotation axis corresponding to the center of mass of the fuselage truss by the mass of the fuselage truss grid unit.

[0025] Preferably, the wing components include a wing heat insulation layer, a wing skin, a wing rib, and a wing spar;

[0026] The mass of the wing insulation layer is obtained by multiplying the sum of the areas of the grid units constituting the wing surface by the thickness of the wing insulation layer and the material density of the wing insulation layer; the centroid of the wing insulation layer is obtained by dividing the sum of the products of the mass of the wing insulation layer grid units and the coordinates of the center points of the grid units by the mass of the wing insulation layer; the moment of inertia of the wing insulation layer is the sum of the moments of inertia of all the wing insulation layer grid units relative to the rotation axis of the centroid of the wing insulation layer, and the moment of inertia of any wing insulation layer grid unit relative to the rotation axis of the centroid of the wing insulation layer is obtained by multiplying the square of the vertical distance between the center point of the wing insulation layer grid unit and the rotation axis corresponding to the centroid of the wing insulation layer by the mass of the wing insulation layer grid unit;

[0027] The mass of the wing skin is obtained by multiplying the sum of the areas of the grid units constituting the wing surface by the thickness of the wing skin and the density of the wing skin material; the core of the wing skin is obtained by dividing the sum of the products of the mass of the wing skin grid unit and the coordinates of the center point of the grid unit by the mass of the wing skin; the moment of inertia of the wing skin is the sum of the moments of inertia of all the wing skin grid units relative to the rotation axis of the core of the wing skin, and the moment of inertia of any wing skin grid unit relative to the rotation axis of the core of the wing skin is obtained by multiplying the square of the vertical distance between the center point of the wing skin grid unit and the corresponding rotation axis of the core of the wing skin by the mass of the wing skin grid unit;

[0028] The mass of the wing rib is obtained by multiplying the sum of the areas of the wing rib grid units by the thickness of the wing rib and the density of the wing rib material; the centroid of the wing rib is obtained by dividing the sum of the products of the mass of the wing rib grid units and the coordinates of the center points of the wing rib grid units by the mass of the wing rib; the moment of inertia of the wing rib is the sum of the moments of inertia of all the wing rib grid units relative to the rotation axis of the wing rib centroid, and the moment of inertia of any wing rib grid unit relative to the rotation axis of the wing rib centroid is obtained by multiplying the square of the vertical distance between the center point of the wing rib grid unit and the rotation axis corresponding to the wing rib centroid by the mass of the wing rib grid unit;

[0029] The mass of the wing spar is obtained by multiplying the sum of the lengths of the wing spar grid units by the cross-sectional area of ​​the wing spar and the material density of the wing spar; the center of mass of the wing spar is obtained by dividing the sum of the products of the mass of the wing spar grid units and the coordinates of the center points of the wing spar grid units by the total mass of the wing spar; the moment of inertia of the wing spar is the sum of the moments of inertia of all wing spar grid units relative to the rotation axis of the center of mass of the wing spar, and the moment of inertia of any wing spar grid unit relative to the rotation axis of the center of mass of the wing spar is obtained by multiplying the square of the vertical distance between the center point of the wing spar grid unit and the rotation axis corresponding to the center of mass of the wing spar by the mass of the wing spar grid unit.

[0030] Preferably, the tail component includes a tail heat insulation layer, a tail skin, a tail rib, and a tail spar;

[0031] The mass of the tail insulation layer is obtained by multiplying the sum of the areas of the grid units constituting the tail surface by the thickness of the tail insulation layer and the density of the material of the tail insulation layer; the center of mass of the tail insulation layer is obtained by dividing the sum of the products of the mass of the tail insulation layer grid units and the coordinates of the center points of the grid units by the mass of the tail insulation layer; the moment of inertia of the tail insulation layer is the sum of the moments of inertia of all the tail insulation layer grid units relative to the axis of rotation of the center of mass of the tail insulation layer, and the moment of inertia of any tail insulation layer grid unit relative to the axis of rotation of the center of mass of the tail insulation layer is obtained by multiplying the square of the vertical distance between the center point of the tail insulation layer grid unit and the axis corresponding to the center of mass of the tail insulation layer by the mass of the tail insulation layer grid unit;

[0032] The mass of the tail skin is obtained by multiplying the sum of the areas of the grid units constituting the tail surface by the thickness of the tail skin and the density of the tail skin material; the center of the tail skin is obtained by dividing the sum of the products of the mass of the tail skin grid unit and the coordinates of the center point of the grid unit by the mass of the tail skin; the moment of inertia of the tail skin is the sum of the moments of inertia of all the tail skin grid units relative to the axis of rotation of the center of the tail skin, and the moment of inertia of any tail skin grid unit relative to the axis of rotation of the center of the tail skin is obtained by multiplying the square of the vertical distance between the center point of the tail skin grid unit and the axis corresponding to the center of the tail skin by the mass of the tail skin grid unit;

[0033] The mass of the tail rib is obtained by multiplying the sum of the areas of the tail rib grid units by the thickness of the tail rib and the density of the tail rib material; the center of mass of the tail rib is obtained by dividing the sum of the products of the mass of the tail rib grid unit and the coordinates of the center point of the tail rib grid unit by the total mass of the tail rib; the moment of inertia of the tail rib is the sum of the moments of inertia of all the tail rib grid units relative to the axis of rotation of the center of mass of the tail rib, and the moment of inertia of any tail rib grid unit relative to the axis of rotation of the center of mass of the tail rib is obtained by multiplying the square of the vertical distance between the center point of the tail rib grid unit and the axis corresponding to the center of mass of the tail rib by the mass of the tail rib grid unit;

[0034] The mass of the tail wing beam is obtained by multiplying the sum of the lengths of the tail wing beam grid units by the cross-sectional area of ​​the tail wing beam and the material density of the tail wing beam; the center of mass of the tail wing beam is obtained by dividing the sum of the products of the mass of the tail wing beam grid units and the coordinates of the center points of the tail wing beam grid units by the mass of the tail wing beam; the moment of inertia of the tail wing beam is the sum of the moments of inertia of all the tail wing beam grid units relative to the axis of rotation of the center of mass of the tail wing beam, and the moment of inertia of any tail wing beam grid unit relative to the axis of rotation of the center of mass of the tail wing beam is obtained by multiplying the square of the vertical distance between the center point of the tail wing beam grid unit and the axis corresponding to the center of mass of the tail wing beam by the mass of the tail wing beam grid unit.

[0035] Preferably, the fuselage mass is the sum of the masses of the fuselage heat insulation layer, the fuselage skin, the fuselage bulkhead, and the fuselage trusses; the fuselage center of mass is the sum of the products of the masses of the fuselage heat insulation layer, the fuselage skin, the fuselage bulkhead, and the fuselage trusses and their respective centers of mass divided by the fuselage mass; the fuselage moment of inertia is the sum of the products of the masses of the fuselage heat insulation layer, the fuselage skin, the fuselage bulkhead, and the fuselage trusses and the squares of the perpendicular distances between their respective centers of mass and the fuselage center of mass coordinate axis, where the fuselage center of mass coordinate axis is the rotation axis of the fuselage moment of inertia.

[0036] Preferably, the wing mass is the sum of the masses of the wing thermal insulation layer, wing skin, wing ribs, and wing spar; the wing center of mass is the sum of the products of the masses of the wing thermal insulation layer, wing skin, wing ribs, and wing spar and their respective centers of mass divided by the wing mass; the wing moment of inertia is the sum of the products of the masses of the wing thermal insulation layer, wing skin, wing ribs, and wing spar and the squares of the perpendicular distances between their respective centers of mass and the wing center of mass coordinate axis, where the wing center of mass coordinate axis is the rotation axis of the wing moment of inertia.

[0037] Preferably, the mass of the tail wing is the sum of the masses of the tail wing heat insulation layer, the tail wing skin, the tail wing ribs, and the tail wing beam; the center of mass of the tail wing is the sum of the products of the masses of the tail wing heat insulation layer, the tail wing skin, the tail wing ribs, and the tail wing beam and their respective centers of mass divided by the mass of the tail wing; the moment of inertia of the tail wing is the sum of the products of the masses of the tail wing heat insulation layer, the tail wing skin, the tail wing ribs, and the tail wing beam and the squares of the perpendicular distances between the masses of the tail wing heat insulation layer, the tail wing skin, the tail wing ribs, and the tail wing beam and their respective centers of mass and the coordinate axis of the center of mass of the tail wing, where the coordinate axis of the center of mass of the tail wing is the axis of rotation of the moment of inertia of the tail wing.

[0038] Preferably, the total mass of the high-speed aircraft structure is the sum of the fuselage, wings, and tail; the center of mass of the high-speed aircraft structure is the sum of the product of the masses of the fuselage, wings, and tail and their respective centers of mass divided by the total mass of the aircraft structure; the moment of inertia of the high-speed aircraft structure is the sum of the products of the masses of the fuselage, wings, and tail and the squares of the perpendicular distances between their respective centers of mass and the coordinate axis of the center of mass of the aircraft structure, where the coordinate axis of the center of mass of the aircraft structure is the axis of rotation of the aircraft structure moment of inertia.

[0039] The beneficial technical effects of the present invention are as follows:

[0040] (1) When the method of the present invention is used for conceptual design of high-speed aircraft, the structural density, thickness and other parameters of the aircraft skin, thermal insulation layer, etc. can be used as design variables to calculate the aircraft structural quality characteristics related to the structural density and thickness parameters, which is more in line with the actual design.

[0041] (2) Based on the present invention, the mass characteristics of the entire aircraft can be quickly calculated after the mass characteristics of the equipment, fuel and propulsion system are given;

[0042] (3) The method of the present invention enables the calculation of the aircraft structure mass characteristics and the calculation of the aircraft aerodynamic performance to share a set of surface grids, thus achieving "one grid for two uses". The aircraft structure mass characteristics can be calculated while the aircraft aerodynamic performance can be calculated, thereby improving the efficiency of multidisciplinary optimization calculations in the conceptual design stage and reducing time costs.

[0043] (4) The invention provides a rapid calculation method for the structural mass characteristics of an aircraft in the conceptual design stage. The calculation method for the structural mass characteristics of an aircraft is simple, easy to program and implement, and can be conveniently integrated into a multidisciplinary optimization design. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Figure 1 Flowchart of the rapid calculation method of structural mass characteristics in the conceptual design phase of high-speed aircraft for this application.

[0045] Figure 2 This is the basic structure diagram of the high-speed aircraft during the conceptual design phase of this application. DETAILED DESCRIPTION

[0046] The present application is described in further detail below in conjunction with the accompanying drawings and specific embodiments so that those skilled in the art can implement the application by referring to the description.

[0047] The present invention discloses a method for quickly calculating the structural quality characteristics of an aircraft in the conceptual design stage. Figure 1 As shown, the method at least includes:

[0048] Step 1: clarify the basic structure of the high-speed aircraft. The basic structure of the high-speed aircraft is as follows: Figure 2 As shown, the aircraft structure is divided into three parts: fuselage, wings and tail. The fuselage is mainly composed of insulation layer, skin, bulkhead and truss, and the wings and tail are mainly composed of insulation layer, skin, wing ribs and wing spars.

[0049] Step 2: Given the geometric shape of a high-speed aircraft. The geometric shape of the aircraft is generally generated by 3D modeling software, such as CATIA, etc. The aircraft reference point is set as the head vertex, the X-axis is in the longitudinal symmetry plane of the aircraft, pointing to the tail, the Y-axis is perpendicular to the longitudinal symmetry plane of the aircraft, pointing to the right, and the Z-axis satisfies the right-hand rule and points upward.

[0050] Step 3, generate the surface grid of the high-speed aircraft. The surface grid is a quadrilateral structure grid, which can be generated by grid division software such as Pointwise. When dividing the surface grid, due to the complexity of the aircraft shape, it is often difficult to divide the aircraft surface grid into a single N×M grid. Generally, the aircraft grid is divided into blocks according to the composition of the aircraft to form a multi-block grid consisting of a fuselage grid block, a wing grid block, and a tail grid block. The three-dimensional grid point coordinates of each grid block of the aircraft are defined in a Cartesian coordinate system. The coordinates of each grid point are in the form of [x, y, z]. The three-dimensional point grid data is sorted in rows and columns, corresponding to the position of the grid points on the surface of the aircraft. There are n points in each row and m points in each column. The number of grid points is n×m, and the number of grid units is (n-1)×(m-1); the grid point data is stored one by one and row by row in a file format such as .data, and finally it can be seen that there are three data in each row, with a total of (n×m) rows.

[0051] Step 4: Import the grid and identify the fuselage, wing, and tail grids. The grid identification method is as follows: the wing has the largest span-wise size, the wing block grid point has the largest Y coordinate average value, the tail is close to the tail of the fuselage, and the tail block grid point has the largest X coordinate average value. The fuselage, wing, and tail grids can be identified by comparing the average coordinate values ​​of the grid points of each grid block.

[0052] Step 5: Calculate the mass, center of mass, and moment of inertia of each component of the fuselage, wing, and tail. The fuselage is composed of a heat insulation layer, a skin, a bulkhead, and a truss. The mass, center of mass, and moment of inertia of the heat insulation layer, the skin, the bulkhead, and the truss can be calculated.

[0053] The calculation method of the fuselage insulation layer mass is as follows:

[0054]

[0055] Where: M 机身隔热层 is the mass of fuselage thermal insulation layer, m 机身隔热层 (i,j) is the mass of the fuselage insulation grid unit, n jg , m jg is the number of rows and columns of the fuselage surface grid points. The mass of the fuselage thermal insulation layer grid unit is calculated as follows:

[0056] m 机身隔热层 (i,j)=S 机身隔热层 (i,j)×TH 机身隔热层 ×ρ 机身隔热层 (2)

[0057] Where: S 机身隔热层 (i,j) is the grid unit area of ​​the fuselage thermal insulation layer, TH 机身隔热层 is the thickness of the fuselage insulation layer, ρ 机身隔热层 is the density of the fuselage insulation layer.

[0058] The calculation method of the centroid of the fuselage insulation layer is as follows:

[0059]

[0060] Where: comx 机身隔热层 ,comy 机身隔热层 ,comz 机身隔热层 They are respectively the X coordinate, Y coordinate, and Z coordinate of the center of mass of the fuselage insulation layer. 机身隔热层 (i,j),y 机身隔热层 (i,j),z 机身隔热层 (i, j) are the X-coordinate, Y-coordinate, and Z-coordinate of the center point of the fuselage insulation layer grid unit, respectively.

[0061] The calculation method of the moment of inertia of the fuselage thermal insulation layer is as follows:

[0062]

[0063] Where: Ix 机身隔热层 ,Iy 机身隔热层 ,Iz 机身隔热层 are the moments of inertia of the fuselage thermal insulation layer around the X-axis, Y-axis, and Z-axis respectively.

[0064] Fuselage skin mass M 机身蒙皮 , centroid comx 机身蒙皮 ,comy 机身蒙皮 ,comz 机身蒙皮 , moment of inertia Ix 机身蒙皮 ,Iy 机身蒙皮,Iz 机身蒙皮 The calculation method is the same as that of the fuselage insulation layer, and only parameters can be replaced during calculation.

[0065] The calculation method of fuselage bulkhead mass is as follows:

[0066]

[0067] Where: M 机身隔框 is the mass of the fuselage bulkhead, m 机身隔框 (i,j) is the mass of the fuselage frame grid unit, n jk , m jk are the number of grid points of bulkhead outline and the number of bulkheads respectively. The mass of fuselage bulkhead grid unit is calculated as follows:

[0068] m 机身隔框 (i,j)=L 机身隔框 (i,j)×S 机身隔框 ×ρ 机身隔框 (6)

[0069] Where: L 机身隔框 (i,j) is the fuselage bulkhead grid unit length, S 机身隔框 is the cross-sectional area of ​​the fuselage bulkhead, ρ 机身隔框 is the density of the fuselage bulkhead.

[0070] The calculation method of the center of mass of the fuselage frame is as follows:

[0071]

[0072] Where: comx 机身隔框 ,comy 机身隔框 ,comz 机身隔框 They are respectively the X coordinate, Y coordinate, and Z coordinate of the center of mass of the fuselage frame. 机身隔框 (i,j),y 机身隔框 (i,j),z 机身隔框 (i, j) are the X-coordinate, Y-coordinate and Z-coordinate of the center point of the fuselage bulkhead length unit respectively.

[0073] The calculation method of the fuselage bulkhead moment of inertia is as follows:

[0074]

[0075] Where: Ix 机身隔框 ,Iy 机身隔框 ,Iz 机身隔框 They are the moments of inertia of the fuselage frame around the X-axis, Y-axis, and Z-axis respectively.

[0076] Fuselage girder mass M 机身桁梁 , centroid comx 机身桁梁 ,comy 机身桁梁 ,comz机身桁梁 , moment of inertia Ix 机身桁梁 ,Iy 机身桁梁 ,Iz 机身桁梁 The calculation method is the same as that of the fuselage bulkhead, and only parameters need to be replaced during calculation.

[0077] The wing is composed of a heat insulation layer, a skin, a wing rib, and a wing beam. The mass of the heat insulation layer of the wing is M 机翼隔热层 , centroid comx 机翼隔热层 ,comy 机翼隔热层 ,comz 机翼隔热层 , moment of inertia Ix 机翼隔热层 ,Iy 机翼隔热层 ,Iz 机翼隔热层 The calculation method is the same as that of the fuselage thermal insulation layer; the wing skin mass M 机翼蒙皮 , centroid comx 机翼蒙皮 ,comy 机翼蒙皮 ,comz 机翼蒙皮 , moment of inertia Ix 机翼蒙皮 ,Iy 机翼蒙皮 ,Iz 机翼蒙皮 The calculation method is the same as that of the fuselage skin; the wing spar mass M 机翼翼梁 , centroid comx 机翼翼梁 ,comy 机翼翼梁 ,comz 机翼翼梁 , moment of inertia Ix 机翼翼梁 ,Iy 机翼翼梁 ,Iz 机翼翼梁 The calculation method is the same as that of the fuselage truss, and only parameters need to be replaced during calculation.

[0078] The wing rib mass is calculated as follows:

[0079]

[0080] Where: M 机翼翼肋 is the mass of the wing rib, m 机翼翼肋 (i,j) is the wing rib mass unit, n jl , m jl are the number of mesh points of the wing rib profile and the number of ribs respectively. The wing rib mass unit is calculated as follows:

[0081] m 机翼翼肋 (i,j)=S 机翼翼肋 (i,j)×TH 机翼翼肋 ×ρ 机翼翼肋 (10)

[0082] Where: S 机翼翼肋 (i,j) is the wing rib area unit, TH 机翼翼肋 is the wing rib thickness, ρ 机翼翼肋 is the wing rib density.

[0083] The calculation method of the wing rib centroid is as follows:

[0084]

[0085] Where: comx 机身隔热层 ,comy 机身隔热层 ,comz 机身隔热层 are the X, Y and Z coordinates of the wing rib centroid, respectively. 机翼翼肋 (i,j),y 机翼翼肋 (i,j),z 机翼翼肋 (i, j) are the X-coordinate, Y-coordinate, and Z-coordinate of the center point of the wing rib grid unit, respectively.

[0086] The calculation method of the wing rib moment of inertia is as follows:

[0087]

[0088] Where: Ix 机翼翼肋 ,Iy 机翼翼肋 ,Iz 机翼翼肋 are the moments of inertia of the wing ribs around the X-axis, Y-axis, and Z-axis respectively.

[0089] The tail is composed of a heat insulation layer, a skin, ribs, and a wing beam. The heat insulation layer of the tail has a mass M 尾翼隔热层 , centroid comx 尾翼隔热层 ,comy 尾翼隔热层 ,comz 尾翼隔热层 , moment of inertia Ix 尾翼隔热层 ,Iy 尾翼隔热层 ,Iz 尾翼隔热层 The calculation method is the same as that of the wing thermal insulation layer; the tail skin mass M 尾翼蒙皮 , centroid comx 尾翼蒙皮 ,comy 尾翼蒙皮 ,comz 尾翼蒙皮 , moment of inertia Ix 尾翼蒙皮 ,Iy 尾翼蒙皮 ,Iz 尾翼蒙皮 The calculation method is the same as that of the wing skin; the mass of the tail wing rib M 尾翼翼肋 , centroid comx 尾翼翼肋 ,comy 尾翼翼肋 ,comz 尾翼翼肋 , moment of inertia Ix 尾翼翼肋 ,Iy 尾翼翼肋 ,Iz 尾翼翼肋 The calculation method is the same as that of the wing ribs; the mass of the tail wing beam M 尾翼翼梁 , centroid comx 尾翼翼梁 ,comy 尾翼翼梁 ,comz 尾翼翼梁 , moment of inertia Ix 尾翼翼梁 ,Iy尾翼翼梁 ,Iz 尾翼翼梁 The calculation method is the same as that of the wing spar, and only parameters need to be replaced during calculation.

[0090] Step 6: Calculate the mass, center of mass, and moment of inertia of the fuselage, wing, and tail wing respectively. The method for calculating the fuselage mass is as follows:

[0091] M 机身 =M 机身隔热层 +M 机身蒙皮 +M 机身隔框 +M 机身桁梁 (13)

[0092] Where: M 机身 For the quality of the fuselage.

[0093] The calculation method of the fuselage center of mass is as follows:

[0094]

[0095] Where: comx 机身 ,comy 机身 ,comz 机身 They are the X coordinate, Y coordinate, and Z coordinate of the center of mass of the fuselage.

[0096] The calculation method of the fuselage moment of inertia is as follows:

[0097]

[0098] Where: Ix 机身 ,Iy 机身 ,Iz 机身 They are the moments of inertia of the fuselage around the X-axis, Y-axis, and Z-axis respectively.

[0099] The wing and tail mass M 机翼 、M 尾翼 , centroid comx 机翼 ,comy 机翼 ,comz 机翼 ,comx 尾翼 ,comy 尾翼 ,comz 尾翼 , moment of inertia Ix 机翼 ,Iy 机翼 ,Iz 机翼 , Ix 尾翼 ,Iy 尾翼 ,Iz 尾翼 The calculation method is the same as that of the fuselage, and you only need to replace the parameters during calculation.

[0100] Step 7: Calculate the total mass, center of mass, and moment of inertia of the high-speed aircraft structure. The total mass calculation method of the aircraft structure is as follows:

[0101] M 飞行器结构 =M 机身 +M 机翼 +M 尾翼 (16)

[0102] Where: M 飞行器结构 For the quality of the fuselage.

[0103] The calculation method of the center of mass of the aircraft structure is as follows:

[0104]

[0105] Where: comx 飞行器结构 ,comy 飞行器结构 ,comz 飞行器结构 They are the X coordinate, Y coordinate, and Z coordinate of the center of mass of the fuselage.

[0106] The calculation method of the aircraft structure moment of inertia is as follows:

[0107]

[0108] Where: Ix 飞行器结构 ,Iy 飞行器结构 ,Iz 飞行器结构 They are the moment of inertia of the aircraft around the X-axis, Y-axis, and Z-axis respectively.

[0109] For the multidisciplinary design optimization problems faced by high-speed aircraft design, it is necessary to comprehensively consider the overall performance of various disciplines such as aircraft geometry, aerodynamics, and structural quality as much as possible at the beginning of the design. When using the method of the present invention to perform conceptual design of high-speed aircraft, after the geometric layout is determined, the calculation of the aircraft structural mass characteristics and the calculation of the aircraft aerodynamic performance share a set of surface grids, realizing "one network for two purposes". While calculating the structural mass characteristics, the aircraft aerodynamic performance can also be calculated, which improves the efficiency of multidisciplinary optimization calculations in the conceptual design stage and reduces the time cost of multidisciplinary optimization design.

[0110] Parts of the present invention that are not described in detail belong to common knowledge among those skilled in the art.

Claims

1. A method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft, characterized in that: include: The basic structure of the high-speed aircraft is clarified, and the high-speed aircraft includes a fuselage, wings, and tail fins, wherein the fuselage includes a heat insulation layer, a skin, a bulkhead, and a truss; the wings and tail fins both include a heat insulation layer, a skin, a wing rib, and a wing spar; Given the high-speed aircraft geometry; Generate high-speed aircraft surface mesh; Import the mesh and identify the fuselage, wing, and tail meshes; Calculate the mass, center of mass, and moment of inertia of each component of the fuselage, wing, and tail respectively; Calculate the mass, center of mass, and moment of inertia of the fuselage, wing, and tail wing respectively; Calculate the total mass, center of mass, and moment of inertia of high-speed aircraft structures.

2. A method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1, characterized in that: The generated surface mesh of the high-speed aircraft is a quadrilateral structured mesh, and the surface mesh can be used for the calculation of the aerodynamic performance of the aircraft.

3. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1 is characterized in that: Importing a mesh means importing all mesh points of the aircraft surface mesh.

4. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 3 is characterized in that: The method to identify the fuselage, wing, and tail meshes is as follows: The aircraft reference point is set as the head vertex, the X axis is within the longitudinal symmetry plane of the aircraft and points to the tail, the Y axis is perpendicular to the longitudinal symmetry plane of the aircraft and points to the right, and the Z axis satisfies the right-hand rule and points upward; The wing has the largest span-wise dimension, and the average Y coordinate of the wing block grid points is the largest. The tail is close to the tail of the fuselage, and the average X coordinate of the tail block grid points is the largest. The fuselage, wing, and tail grids can be identified by comparing the average coordinates of the grid points of each grid block.

5. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1 is characterized in that: The fuselage components include fuselage thermal insulation layer, fuselage skin, fuselage bulkhead, and fuselage girders; The mass of the fuselage heat insulation layer is obtained by multiplying the sum of the areas of the grid units constituting the fuselage surface by the thickness of the fuselage heat insulation layer and the material density of the fuselage heat insulation layer; the center of mass of the fuselage heat insulation layer is obtained by dividing the sum of the products of the mass of the fuselage heat insulation layer grid units and the coordinates of the center points of the grid units by the mass of the fuselage heat insulation layer; the moment of inertia of the fuselage heat insulation layer is the sum of the moments of inertia of all the fuselage heat insulation layer grid units relative to the rotation axis of the center of mass of the fuselage heat insulation layer, and the moment of inertia of any fuselage heat insulation layer grid unit relative to the rotation axis of the center of mass of the fuselage heat insulation layer is obtained by multiplying the square of the vertical distance between the center point of the fuselage heat insulation layer grid unit and the rotation axis corresponding to the center of mass of the fuselage heat insulation layer by the mass of the fuselage heat insulation layer grid unit; The fuselage skin mass is obtained by multiplying the sum of the areas of the mesh units constituting the fuselage surface by the thickness of the fuselage skin and the density of the fuselage skin material; the fuselage skin center is obtained by dividing the sum of the products of the fuselage skin mesh unit mass and the coordinates of the center point of the mesh unit by the fuselage skin mass; The moment of inertia of the fuselage skin is the sum of the moments of inertia of all fuselage skin grid units relative to the rotation axis of the fuselage skin center. The moment of inertia of any fuselage skin grid unit relative to the rotation axis of the fuselage skin center is obtained by multiplying the square of the vertical distance between the center point of the fuselage skin grid unit and the corresponding rotation axis of the fuselage skin center by the mass of the fuselage skin grid unit. The mass of the fuselage bulkhead frame is obtained by multiplying the sum of the lengths of the fuselage bulkhead frame grid units by the cross-sectional area of ​​the fuselage bulkhead frame and the density of the fuselage bulkhead frame material; the center of mass of the fuselage bulkhead frame is obtained by dividing the sum of the product of the mass of the fuselage bulkhead frame grid units and the coordinates of the center point of the fuselage bulkhead frame length unit by the mass of the fuselage bulkhead frame; The moment of inertia of the fuselage bulkhead frame is the sum of the moments of inertia of all fuselage bulkhead frame grid units relative to the rotation axis of the fuselage bulkhead frame mass center. The moment of inertia of any fuselage bulkhead frame grid unit relative to the rotation axis of the fuselage bulkhead frame mass center is obtained by multiplying the square of the vertical distance between the center point of the fuselage bulkhead frame grid unit and the rotation axis corresponding to the fuselage bulkhead frame mass center by the mass of the fuselage bulkhead frame grid unit. The mass of the fuselage truss is obtained by multiplying the sum of the lengths of the fuselage truss grid units by the cross-sectional area of ​​the fuselage truss and the material density of the fuselage truss; the center of mass of the fuselage truss is obtained by dividing the sum of the product of the mass of the fuselage truss grid unit and the coordinates of the center point of the fuselage truss length unit by the mass of the fuselage truss; the moment of inertia of the fuselage truss is the sum of the moments of inertia of all fuselage truss grid units relative to the rotation axis of the center of mass of the fuselage truss, and the moment of inertia of any fuselage truss grid unit relative to the rotation axis of the center of mass of the fuselage truss is obtained by multiplying the square of the vertical distance between the center point of the fuselage truss grid unit and the rotation axis corresponding to the center of mass of the fuselage truss by the mass of the fuselage truss grid unit.

6. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1 is characterized in that: The wing components include wing insulation layer, wing skin, wing ribs, and wing spars; The mass of the wing insulation layer is obtained by multiplying the sum of the areas of the grid units constituting the wing surface by the thickness of the wing insulation layer and the material density of the wing insulation layer; the centroid of the wing insulation layer is obtained by dividing the sum of the products of the mass of the wing insulation layer grid units and the coordinates of the center points of the grid units by the mass of the wing insulation layer; the moment of inertia of the wing insulation layer is the sum of the moments of inertia of all the wing insulation layer grid units relative to the rotation axis of the centroid of the wing insulation layer, and the moment of inertia of any wing insulation layer grid unit relative to the rotation axis of the centroid of the wing insulation layer is obtained by multiplying the square of the vertical distance between the center point of the wing insulation layer grid unit and the rotation axis corresponding to the centroid of the wing insulation layer by the mass of the wing insulation layer grid unit; The mass of the wing skin is obtained by multiplying the sum of the areas of the grid units constituting the wing surface by the thickness of the wing skin and the density of the wing skin material; the core of the wing skin is obtained by dividing the sum of the products of the mass of the wing skin grid unit and the coordinates of the center point of the grid unit by the mass of the wing skin; the moment of inertia of the wing skin is the sum of the moments of inertia of all the wing skin grid units relative to the rotation axis of the core of the wing skin, and the moment of inertia of any wing skin grid unit relative to the rotation axis of the core of the wing skin is obtained by multiplying the square of the vertical distance between the center point of the wing skin grid unit and the corresponding rotation axis of the core of the wing skin by the mass of the wing skin grid unit; The mass of the wing rib is obtained by multiplying the sum of the areas of the wing rib grid units by the thickness of the wing rib and the density of the wing rib material; the centroid of the wing rib is obtained by dividing the sum of the products of the mass of the wing rib grid units and the coordinates of the center points of the wing rib grid units by the mass of the wing rib; the moment of inertia of the wing rib is the sum of the moments of inertia of all the wing rib grid units relative to the rotation axis of the wing rib centroid, and the moment of inertia of any wing rib grid unit relative to the rotation axis of the wing rib centroid is obtained by multiplying the square of the vertical distance between the center point of the wing rib grid unit and the rotation axis corresponding to the wing rib centroid by the mass of the wing rib grid unit; The mass of the wing spar is obtained by multiplying the sum of the lengths of the wing spar grid units by the cross-sectional area of ​​the wing spar and the material density of the wing spar; the center of mass of the wing spar is obtained by dividing the sum of the products of the mass of the wing spar grid units and the coordinates of the center points of the wing spar grid units by the total mass of the wing spar; the moment of inertia of the wing spar is the sum of the moments of inertia of all wing spar grid units relative to the rotation axis of the center of mass of the wing spar, and the moment of inertia of any wing spar grid unit relative to the rotation axis of the center of mass of the wing spar is obtained by multiplying the square of the vertical distance between the center point of the wing spar grid unit and the rotation axis corresponding to the center of mass of the wing spar by the mass of the wing spar grid unit.

7. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1 is characterized in that: The tail components include tail insulation layer, tail skin, tail ribs, and tail beams; The mass of the tail insulation layer is obtained by multiplying the sum of the areas of the grid units constituting the tail surface by the thickness of the tail insulation layer and the density of the material of the tail insulation layer; the center of mass of the tail insulation layer is obtained by dividing the sum of the products of the mass of the tail insulation layer grid units and the coordinates of the center points of the grid units by the mass of the tail insulation layer; the moment of inertia of the tail insulation layer is the sum of the moments of inertia of all the tail insulation layer grid units relative to the axis of rotation of the center of mass of the tail insulation layer, and the moment of inertia of any tail insulation layer grid unit relative to the axis of rotation of the center of mass of the tail insulation layer is obtained by multiplying the square of the vertical distance between the center point of the tail insulation layer grid unit and the axis corresponding to the center of mass of the tail insulation layer by the mass of the tail insulation layer grid unit; The mass of the tail skin is obtained by multiplying the sum of the areas of the grid units constituting the tail surface by the thickness of the tail skin and the density of the tail skin material; the center of the tail skin is obtained by dividing the sum of the products of the mass of the tail skin grid unit and the coordinates of the center point of the grid unit by the mass of the tail skin; the moment of inertia of the tail skin is the sum of the moments of inertia of all the tail skin grid units relative to the axis of rotation of the center of the tail skin, and the moment of inertia of any tail skin grid unit relative to the axis of rotation of the center of the tail skin is obtained by multiplying the square of the vertical distance between the center point of the tail skin grid unit and the axis corresponding to the center of the tail skin by the mass of the tail skin grid unit; The mass of the tail rib is obtained by multiplying the sum of the areas of the tail rib grid units by the thickness of the tail rib and the density of the tail rib material; the center of mass of the tail rib is obtained by dividing the sum of the products of the mass of the tail rib grid unit and the coordinates of the center point of the tail rib grid unit by the total mass of the tail rib; the moment of inertia of the tail rib is the sum of the moments of inertia of all the tail rib grid units relative to the axis of rotation of the center of mass of the tail rib, and the moment of inertia of any tail rib grid unit relative to the axis of rotation of the center of mass of the tail rib is obtained by multiplying the square of the vertical distance between the center point of the tail rib grid unit and the axis corresponding to the center of mass of the tail rib by the mass of the tail rib grid unit; The mass of the tail wing beam is obtained by multiplying the sum of the lengths of the tail wing beam grid units by the cross-sectional area of ​​the tail wing beam and the material density of the tail wing beam; the center of mass of the tail wing beam is obtained by dividing the sum of the products of the mass of the tail wing beam grid units and the coordinates of the center points of the tail wing beam grid units by the mass of the tail wing beam; the moment of inertia of the tail wing beam is the sum of the moments of inertia of all the tail wing beam grid units relative to the axis of rotation of the center of mass of the tail wing beam, and the moment of inertia of any tail wing beam grid unit relative to the axis of rotation of the center of mass of the tail wing beam is obtained by multiplying the square of the vertical distance between the center point of the tail wing beam grid unit and the axis corresponding to the center of mass of the tail wing beam by the mass of the tail wing beam grid unit.

8. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1 is characterized in that: The fuselage mass is the sum of the masses of the fuselage insulation layer, fuselage skin, fuselage bulkheads, and fuselage girders; the fuselage center of mass is the sum of the products of the masses of the fuselage insulation layer, fuselage skin, fuselage bulkheads, and fuselage girders and their respective centers of mass divided by the fuselage mass; the fuselage moment of inertia is the sum of the products of the masses of the fuselage insulation layer, fuselage skin, fuselage bulkheads, and fuselage girders and the squares of the perpendicular distances from their respective centers of mass to the fuselage center of mass coordinate axis, where the fuselage center of mass coordinate axis is the axis of rotation of the fuselage moment of inertia.

9. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1 is characterized in that: The wing mass is the sum of the masses of the wing insulation layer, wing skin, wing ribs, and wing spar; the wing center of mass is the sum of the products of the masses of the wing insulation layer, wing skin, wing ribs, and wing spar and their respective centers of mass divided by the wing mass; the wing moment of inertia is the sum of the products of the masses of the wing insulation layer, wing skin, wing ribs, and wing spar and the squares of the perpendicular distances between their respective centers of mass and the wing center of mass coordinate axis, where the wing center of mass coordinate axis is the axis of rotation of the wing moment of inertia.

10. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1, characterized in that: The mass of the tail wing is the sum of the masses of the tail wing insulation layer, tail wing skin, tail wing ribs, and tail wing beam; the center of mass of the tail wing is the sum of the products of the masses of the tail wing insulation layer, tail wing skin, tail wing ribs, and tail wing beam and their respective centers of mass divided by the mass of the tail wing; the moment of inertia of the tail wing is the sum of the products of the masses of the tail wing insulation layer, tail wing skin, tail wing ribs, and tail wing beam and the squares of the perpendicular distances between their respective centers of mass and the coordinate axis of the center of mass of the tail wing, where the coordinate axis of the center of mass of the tail wing is the axis of rotation of the moment of inertia of the tail wing.

11. The method for rapid calculation of structural quality characteristics in the conceptual design stage of a high-speed aircraft according to claim 1, characterized in that: The total mass of the high-speed aircraft structure is the sum of the fuselage, wings, and tail; the center of mass of the high-speed aircraft structure is the sum of the product of the mass of the fuselage, wings, and tail and their respective center of mass divided by the total mass of the aircraft structure; the moment of inertia of the high-speed aircraft structure is the sum of the product of the mass of the fuselage, wings, and tail and the square of the perpendicular distance between their respective center of mass and the coordinate axis of the center of mass of the aircraft structure, where the coordinate axis of the center of mass of the aircraft structure is the axis of rotation of the aircraft structure moment of inertia.

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