A fast calculation method for spacecraft load design

Abstract

The present invention discloses a method for rapid calculation of aerospace vehicle load design, comprising the following steps: S1, selecting aerospace vehicle load design working condition; S2, dividing aerospace vehicle load calculation sections and allocating mass; S3, calculating aerodynamic forces of each section and normal forces on a servo mechanism under each working condition; S4, calculating the design load of each section under each working condition, wherein the design load includes axial force, transverse shear force and transverse bending moment. The present invention can obtain a more realistic and reasonable cross-sectional load distribution, and has simple engineering application and reasonable method, while greatly reducing the difficulty of aircraft structure design.

Description

Technical Field

[0001] The present invention relates to the technical field of aerospace vehicle load design, and in particular to a rapid calculation method for aerospace vehicle load design. Background Art

[0002] Spacecraft load design is a major task in spacecraft design. Its main purpose is to analyze the spacecraft mission profile, select load conditions, and perform load calculations, so as to provide a basis for the spacecraft structure design and the strength calculation of sections (including stability and dynamic characteristics).

[0003] At present, the calculation method for aerospace vehicle load design is to calculate the pressure on the vehicle surface through CFD, and then carry out load design and strength analysis comprehensively through surface pressure loading according to the stress state of the vehicle. However, there are many shortcomings in long-term model design:

[0004] (1) The pressure on the aircraft surface is uneven. There is a deviation when loading with equal values ​​over a large area. When loading with surface distribution nodes, there is a mismatch between the strength calculation finite element nodes and the CFD calculation nodes. There is also a pressure deviation, which leads to a deviation in the final aerodynamic loading results.

[0005] (2) The load design calculation based on pressure loading is relatively complex for strength analysis modeling, requires a large amount of calculation work, and is not convenient for analyzing the load status of a single compartment. Summary of the invention

[0006] The present invention aims to provide a rapid calculation method for aerospace vehicle load design, so as to solve the problem that the existing aerospace vehicle load design calculation method is complicated and inconvenient for strength analysis.

[0007] In order to achieve the above object, the present invention provides the following technical solutions:

[0008] A rapid calculation method for aerospace vehicle load design comprises the following steps:

[0009] S1. Select the load design condition of the aerospace vehicle;

[0010] S2. Divide the spacecraft load calculation section and distribute the mass;

[0011] S3. Calculate the aerodynamic force of each section and the normal force on the servo mechanism under each working condition;

[0012] S4. Calculate the design load of each section under each working condition, where the design load includes axial force, transverse shear force and transverse bending moment.

[0013] Furthermore, in step S1, the load design conditions of the aerospace vehicle include the maximum axial overload, the maximum lateral overload and the maximum dynamic pressure point.

[0014] Further, the step S2 is specifically:

[0015] A coordinate system is established with the theoretical cusp of the spacecraft as the origin o, the longitudinal axis of the spacecraft pointing to the tail as the x-axis, the axis perpendicular to ox upward as the y-axis, the z-axis and the x- and y-axes follow the right-hand rule, and the spacecraft is divided into s calculation sections along the axial direction, and the total mass of the missile is distributed to each section according to the mass distribution of each section.

[0016] Further, the step S4 is specifically:

[0017] (1) Calculation of cross-sectional axial force, including:

[0018] A. Calculation of axial force in active section. The engine thrust action section is the engine front frame section, set as i F , then:

[0019] Axial force at the nth node in front of the engine thrust action section:

[0020]

[0021] Engine thrust acting cross section i F Axial force:

[0022]

[0023] Axial force at the nth node behind the engine thrust action section:

[0024]

[0025] B. Calculation of axial force in passive section, without engine thrust, axial force at the nth node:

[0026]

[0027] (2) Calculation of transverse shear force of each section:

[0028]

[0029] (3) Calculation of transverse bending moment of each section:

[0030]

[0031] In the above formula, i=1,2,···,s;n=1,2,···,s;G i is the mass of the i-th section, x i is the centroid of the i-th section, X i is the axial aerodynamic force distributed in the i-th section, Y iis the normal aerodynamic force distributed on the i-th section, n xi is the axial overload coefficient of the i-th section, n xi =n x0 , n yi is the normal overload coefficient of the i-th section, Y r is the normal force on the servo mechanism, x n is the centroid of the nth section, x r is the center of mass of the servo mechanism, and F is the engine thrust.

[0032] Compared with the prior art, the present invention has the following beneficial effects:

[0033] The invention discloses a rapid calculation method for aerospace vehicle load design, which can obtain a more realistic and reasonable cross-sectional load distribution, and has simple engineering application, reasonable and reliable method, and reduces the difficulty of vehicle structure design to a great extent.

[0034] At the same time, this method can conveniently give the axial force, shear force and bending moment of different cross-sections of the aircraft. The calculation results reflect the comprehensive force results of the aircraft during the flight process. The calculation results can be directly applied to carry out strength analysis of the aircraft structure more conveniently. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Figure 1 A flowchart of a method for rapid calculation of aerospace vehicle load design provided by an embodiment of the present invention;

[0036] Figure 2 A schematic diagram of the cross-section division for spacecraft load calculation provided in an embodiment of the present invention. DETAILED DESCRIPTION

[0037] The present invention is further described in detail below in conjunction with the accompanying drawings and embodiments:

[0038] like Figure 1 As shown, a rapid calculation method for aerospace vehicle load design includes the following steps:

[0039] S1. Select the load design condition of the aerospace vehicle;

[0040] Specifically, according to the mission profile of the spacecraft, the main ballistic characteristic point conditions are selected, which are generally divided into the active section (with engine thrust) and the passive section (without engine thrust). The maximum axial overload, maximum lateral overload and maximum dynamic pressure point are selected as the load calculation conditions.

[0041] S2. Divide the spacecraft load calculation section and distribute the mass;

[0042] Specifically, refer to Figure 2, establish a coordinate system with the theoretical cusp of the spacecraft as the origin o, take the longitudinal axis of the spacecraft pointing to the tail as the x-axis, and the axis perpendicular to ox upward as the y-axis. The z-axis and the x- and y-axes follow the right-hand rule, and divide the spacecraft into s calculation sections along the axial direction. In this embodiment, s=29, and distribute the total mass to each section according to the mass distribution of each section. The weight of each section is recorded as G i (i=1~s), the corresponding axial coordinates are recorded as x i (i=1~s).

[0043] S3. Calculate the aerodynamic force of each section and the normal force on the servo mechanism under each working condition;

[0044] Specifically, the distributed axial aerodynamic force X i , normal aerodynamic force Y i , lateral aerodynamic force Z i and the normal force Y on the servo mechanism r These parameters are all calculated by aerodynamic professionals and are known state parameters.

[0045] S4. Calculate the design load of each section under each working condition, wherein the design load includes axial force, transverse shear force and transverse bending moment;

[0046] Specifically:

[0047] (1) Calculation of overall parameters of aerospace vehicles, including:

[0048] Total weight of spacecraft:

[0049]

[0050] Spacecraft center of mass:

[0051]

[0052] The total axial aerodynamic force and normal aerodynamic force of aerospace vehicle:

[0053]

[0054] Spacecraft pressure center:

[0055]

[0056] Spacecraft moment of inertia:

[0057]

[0058] Overload factor at the center of mass of the spacecraft:

[0059] n y0 =(YY r ) / G,nx0 =(FX) / G (6)

[0060] (2) Calculation of cross-sectional axial force, including:

[0061] A. Calculation of axial force in active section. The engine thrust action section is the engine front frame section, set as i F , then:

[0062] Axial force at the nth node in front of the engine thrust action section:

[0063]

[0064] Engine thrust acting cross section i F Axial force:

[0065]

[0066] Axial force at the nth node behind the engine thrust action section:

[0067]

[0068] B. Calculation of axial force in passive section, without engine thrust, axial force at the nth node:

[0069]

[0070] (3) Calculation of transverse shear force of each section:

[0071]

[0072] (4) Calculation of transverse bending moment of each section:

[0073]

[0074] In the above formula, i=1,2,···,s;n=1,2,···,s;G i is the mass of the i-th section, x i is the centroid of the i-th section, X i is the axial aerodynamic force distributed in the i-th section, Y i is the normal aerodynamic force distributed on the i-th section, n xi is the axial overload coefficient of the i-th section, n xi =n x0 , n yi is the normal overload coefficient of the i-th section, x n is the centroid of the nth section, x r is the center of mass of the servo mechanism, and F is the engine thrust.

[0075] The above is only an embodiment of the present invention, and the common knowledge such as the known specific technical solutions and / or characteristics in the solution is not described in detail here. It should be pointed out that for those skilled in the art, without departing from the technical solution of the present invention, several modifications and improvements can be made, which should also be regarded as the protection scope of the present invention, and these will not affect the effect of the implementation of the present invention and the practicality of the patent. The scope of protection required by this application shall be based on the content of its claims, and the specific implementation methods and other records in the specification can be used to interpret the content of the claims.

Claims

1. A rapid calculation method for aerospace vehicle load design, characterized in that: The following steps are involved: S1. Select the load design condition of the aerospace vehicle; S2. Divide the spacecraft load calculation section and distribute the mass; S3. Calculate the aerodynamic force of each section and the normal force on the servo mechanism under each working condition; S4. Calculate the design load of each section under each working condition, wherein the design load includes axial force, transverse shear force and transverse bending moment; The step S4 is specifically: (1) Calculation of cross-sectional axial force, including: A. Calculation of axial force in active section. The engine thrust action section is the engine front frame section, set as i F , then: Axial force at the nth node in front of the engine thrust action section: Engine thrust acting cross section i F Axial force: Axial force at the nth node behind the engine thrust action section: B. Calculation of axial force in passive section, without engine thrust, axial force at the nth node: (2) Calculation of transverse shear force of each section: (3) Calculation of transverse bending moment of each section: In the above formula, i=1,2,···,s;n=1,2,···,s;G i is the mass of the i-th section, x i is the centroid of the i-th section, X i is the axial aerodynamic force distributed in the i-th section, Y i is the normal aerodynamic force distributed on the i-th section, n xi is the axial overload coefficient of the i-th section, n xi =n x0 , n yi is the normal overload coefficient of the i-th section, Y r is the normal force on the servo mechanism, x n is the centroid of the nth section, x r is the center of mass of the servo mechanism, and F is the engine thrust.

2. The method for rapid calculation of aerospace vehicle load design according to claim 1, characterized in that: In step S1, the load design conditions of the aerospace vehicle include the maximum axial overload, the maximum lateral overload and the maximum dynamic pressure point.

3. The method for rapid calculation of aerospace vehicle load design according to claim 2, characterized in that: The step S2 is specifically: A coordinate system is established with the theoretical cusp of the spacecraft as the origin o, the longitudinal axis of the spacecraft pointing to the tail as the x-axis, the axis perpendicular to ox upward as the y-axis, the z-axis and the x- and y-axes follow the right-hand rule, and the spacecraft is divided into s calculation sections along the axial direction, and the total mass of the missile is distributed to each section according to the mass distribution of each section.

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