A launch vehicle flight static load checking method

Abstract

This invention relates to a method for verifying the static load of a launch vehicle in flight, comprising: analyzing the factors affecting the flight load of the launch vehicle, determining the overall parameter deviation terms and deviation ranges; generating random numbers for each deviation using the Monte Carlo method; establishing a six-degree-of-freedom trajectory simulation model containing six dynamic motion differential equations, and performing six-degree-of-freedom simulation of the trajectory; screening the obtained trajectory to select the load calculation conditions; using d'Alembert's principle to solve for the verification static load on each section of the rocket under each load calculation condition; and using the equivalent axial force method to determine the difference between the verification static load calculation results and the initial design flight static load calculation results, and verifying the initial design flight static load calculation results. This invention proposes a method for verifying the static load of a launch vehicle in flight, which considers the influencing factors of the flight static load more comprehensively, meeting the requirements of simultaneity and coverage of flight load design conditions.

Description

Technical Field

[0001] This invention belongs to the field of overall launch vehicle technology, specifically relating to a method for verifying the static load of a launch vehicle during flight. Background Technology

[0002] Load design is one of the fundamental tasks in aircraft development. Its main purpose is to select load conditions and perform load calculations by analyzing the aircraft's mission profile, providing a basis for the structural design and strength and stability calculations of each component. Generally, aircraft structural design is primarily based on the loads required during flight, while other load conditions can be used to check and verify the structural strength. Throughout the entire flight process, except for a few short-duration phases such as ignition and takeoff, stage separation, engine start-up and shutdown, and transonic flight, the loads on the rocket body can be calculated as static loads for most of the time. Therefore, flight static load calculation is the basic and primary task of load analysis. For a given operating condition, the method for calculating static loads is relatively uniform, usually referring to industry standards for launch vehicle load design. However, static load design requires not only load calculations but also the selection of the maximum load condition, which serves as a crucial basis for the rocket's structural design and the strength analysis of its various sections. Static load calculations need to be performed at the initial stage of rocket design. However, at this stage, due to the lack of parameters such as the rocket body's elastic coefficients, the ballistic parameters used for selecting flight load conditions are typically simplified calculations based on rigid body motion equations after considering wind interference from a standard trajectory. This method does not comprehensively consider the factors affecting the launch vehicle's load during flight. Therefore, it is necessary to address how to comprehensively consider the factors affecting the launch vehicle's load during flight and to verify the static loads. Summary of the Invention

[0003] This invention addresses the static load calculation results during the preliminary design phase of rocket design and proposes a static load verification scheme for launch vehicles based on a six-degree-of-freedom simulated trajectory. The scheme includes formulating pull-off conditions, six-degree-of-freedom simulation, selecting load conditions, calculating loads, and verifying the initial design flight load calculation results.

[0004] The present invention provides a method for verifying the static load of a launch vehicle during flight, comprising the following steps:

[0005] The factors affecting the flight payload of the launch vehicle are analyzed to determine the overall parameter deviation items and deviation ranges;

[0006] The Monte Carlo method is adopted, considering the probability deviation distribution of different overall parameters, generating random numbers for each deviation, and considering the various forces and torques acting on the rocket during the motion. A six-degree-of-freedom simulation model of the trajectory containing six dynamic motion differential equations is established, and the trajectory is simulated in six degrees of freedom.

[0007] The ballistics obtained from the six-degree-of-freedom simulation are screened to remove those with deviations exceeding the set deviation range. Based on the screened ballistics, load calculation conditions are selected, including the maximum point of the product of dynamic pressure and total angle of attack, the maximum point of lateral overload, the maximum point of dynamic pressure, and the maximum point of axial overload.

[0008] Based on the mass substation data and distributed aerodynamic load data of the selected load calculation conditions, the verification static loads on each section of the rocket under each load calculation condition are solved using general mechanics methods and d'Alembert's principle.

[0009] The results of the static load verification calculation are compared and analyzed with the results of the initial design flight load calculation. The equivalent axial force method is used to determine the difference between the results of the static load verification calculation and the results of the initial design flight load calculation, thereby verifying the results of the initial design flight load calculation.

[0010] Furthermore, the static load verification of each section of the rocket body includes axial force, shear force, and bending moment. The calculation method for the verification load is as follows: based on the mass and center of mass characteristics of each section of the rocket, the rocket is reduced to multiple equivalent mass points, and the rocket is divided into stations, thereby discretizing the rocket into a single beam composed of a finite number of mass points with different masses. Each mass point is called a discretized station. The aerodynamic force and inertial force on the rocket body are distributed to each station. Starting from the apex of the rocket nose, the axial force, shear force, and bending moment of each section are calculated station by station. The axial force of the current station section is the cumulative value of the axial force of each station before this station section, the shear force of the current station section is the cumulative value of the normal force of each station before this station section, and the bending moment of the current station section is the cumulative value of the product of the bending moment generated by each station on this station section, i.e., the normal force and the distance.

[0011] The shear force and bending moment generated by aerodynamic forces during rocket flight can be calculated using the following formulas.

[0012] (1)

[0013] (2)

[0014] (3)

[0015] (4)

[0016] In the formula ，Q yn 、Q zn Represents the shear forces in the y and z directions generated by aerodynamic forces.

[0017] M yn 、M znIndicates the bending moments in the y and z directions generated by aerodynamic forces.

[0018] q Indicates flight pressure

[0019] s Represents aerodynamic reference area

[0020] CN i CZ i CA i Aerodynamic coefficients representing normal, lateral, and axial directions.

[0021] CN, CZ, CA This represents the sum of the aerodynamic coefficients in the normal, lateral, and axial directions.

[0022] x T 、x y Indicates the rocket's center of mass and center of aerodynamic pressure.

[0023] Jy、Jz Represents the longitudinal and lateral rotational inertia of the rocket.

[0024] i Represents the site serial number, with a value ranging from 1 to... n natural numbers

[0025] mi and xi This indicates the quality and location of each substation.

[0026] M It is the total mass of the entire rocket

[0027] xy and xz They represent y and z aerodynamic pressure center 。

[0028] Furthermore, the shear force and bending moment caused by the rocket's control forces during flight can be calculated using the following formulas.

[0029] (5)

[0030] (6)

[0031] (7)

[0032] (8)

[0033] In the formula, Q yn Q zn This represents the y- and z-axis shear forces generated by the control force.

[0034] M yn M zn Indicates the bending moments in the y and z directions generated by the control force.

[0035] R c Indicates control force

[0036] x R Indicates the position where the control force is applied.

[0037] Q y(i-1) It is y to the first i -1 shear force at a cross section.

[0038] Furthermore, the shear force and bending moment in the y and z directions of each load calculation section of the entire rocket can be combined using the following formula.

[0039] (9)

[0040] (10)

[0041] The axial force during rocket flight can be calculated using the following formula.

[0042] (11)

[0043] (12)

[0044] In the formula, T n Indicates rocket axial force

[0045] P represents engine thrust.

[0046] CAi Indicates the aerodynamic coefficient in the axial direction

[0047] CA The sum of aerodynamic coefficients in the axial direction

[0048] x f Indicates the position of engine thrust application

[0049] g It is gravitational acceleration. x n Indicates the first n Location of each section of the rocket body;

[0050] Represents the step function, i.e.

[0051] (13).

[0052] Furthermore, the formula for calculating the equivalent axial force is shown below.

[0053] (14)

[0054] In the formula, T eq This represents the equivalent axial force; a positive value represents the equivalent axial compression, and a negative value represents the equivalent axial tension.

[0055] D This indicates the diameter of the cross section used for calculating the loads on the rocket body.

[0056] Furthermore, based on the equivalent axial force comparison results, if the equivalent axial force of the initial design load is less than or does not exceed 5% of the verification static load calculation result obtained by this method, then the initial design load calculation result is considered correct and valid; otherwise, the verification static load obtained by this method is used as the basis for structural strength verification and the structural design is carried out again.

[0057] On the other hand, the present invention also provides a computer-readable storage medium comprising a stored program, wherein the program executes the above-described method for verifying the static load of a launch vehicle during flight.

[0058] On the other hand, the present invention also provides an electronic device including a memory and a processor, wherein the memory stores a computer program and the processor is configured to execute the aforementioned launch vehicle flight static load verification method through the computer program.

[0059] Compared with the prior art, the method of the present invention can achieve the following beneficial effects:

[0060] This paper proposes a flight static load verification scheme for launch vehicles based on a six-degree-of-freedom simulated trajectory, addressing the flight static load calculation results from the preliminary design phase of rocket design. After completing the preliminary design of the launch vehicle, the Monte Carlo method can be used to simulate the trajectory using six degrees of freedom, considering the probability deviation distribution of different overall parameters. Load conditions are selected based on the calculation results of the simulated trajectory, and load calculations are then performed for the selected conditions. Finally, the initial design flight load is verified. The method of this invention comprehensively considers the influencing factors of the flight static load, meeting the requirements of simultaneity and coverage of flight load design conditions. Attached Figure Description

[0061] Figure 1 This is a schematic diagram of the process according to the method of the present invention;

[0062] Figure 2 This diagram compares the static load calculation results of the method of this invention with those of the preliminary design stage (Tn+4Mn / D-equivalent axial compression); point O in the diagram is the tip of the arrow; current load - static load calculation result of the method of this invention, original load - static load calculation result of the preliminary design stage;

[0063] Figure 3 This diagram compares the load calculation results of the method of this invention with those of the preliminary design stage (Tn-4Mn / D-equivalent shaft tension); point O in the diagram is the tip of the arrow; current load - static load calculation result of the method of this invention, original load - static load calculation result of the preliminary design stage. Detailed Implementation

[0064] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0065] like Figure 1 The flowchart shown in this embodiment provides a method for verifying the static load of a launch vehicle during flight, including the following steps.

[0066] Step 1:

[0067] The factors affecting the flight payload of the launch vehicle are analyzed to determine the overall parameter deviation items and deviation ranges. The overall parameter deviation items to be considered should include: engine performance deviation, rocket body elasticity deviation, structural mass deviation, aerodynamic performance deviation, wind field interference, atmospheric density deviation, atmospheric pressure deviation, etc. The deviation range needs to be determined according to the specific circumstances of different rockets.

[0068] Step 2:

[0069] The Monte Carlo method is adopted, considering the probability deviation distribution of different overall parameters, generating random numbers for each deviation, and considering the various forces and torques acting on the rocket during the motion. A six-degree-of-freedom simulation model of the trajectory is established, which includes six dynamic motion differential equations (including three degrees of freedom for instantaneous position and three degrees of freedom for determining the instantaneous attitude of the rigid body, combining guidance and attitude control). The trajectory is simulated with six degrees of freedom. To ensure the reliability of the results, there are generally no fewer than 10,000 trajectories.

[0070] Step 3:

[0071] First, the trajectories obtained from the six-degree-of-freedom simulation are screened to remove those with deviations exceeding the set deviation range. Based on the screened trajectories, load calculation conditions such as the maximum point of the product of dynamic pressure and total angle of attack, the maximum point of lateral overload, the maximum point of dynamic pressure, and the maximum point of axial overload are selected. The parameters of the selected load calculation conditions include parameters such as flight time, flight altitude, rocket mass, engine thrust, Mach number, angle of attack, sideslip angle, dynamic pressure, and rocket overload.

[0072] Step 4: Based on the mass substation data and distributed aerodynamic load data of the selected load calculation conditions, use general mechanics methods and d'Alembert's principle to solve for the static loads on each section of the rocket under each load calculation condition.

[0073] The loads on each section of the rocket body refer to the internal forces within the section, including axial force, shear force, and bending moment. The load calculation method is as follows: Based on the mass and center-of-mass characteristics of each section of the rocket, the rocket is reduced to multiple equivalent point masses. The rocket is then discretized into a single beam composed of a finite number of point masses of different masses, and each point mass is called a discretized station. Then, the aerodynamic forces and inertial forces on the rocket body are distributed to each station. Starting from the tip of the arrowhead, the axial force, shear force, and bending moment of each section (each station is on that section) are calculated station by station. The axial force is the cumulative value of the axial forces of all stations before that station section, the shear force is the cumulative value of the normal forces of all stations before that station section, and the bending moment is the cumulative value of the bending moment (the product of normal force and distance) generated by all stations on that section section.

[0074] The shear force and bending moment generated by aerodynamic forces during rocket flight can be calculated using the following formulas.

[0075] ........................(1)

[0076] ........................(2)

[0077] ........................(3)

[0078] ........................(4)

[0079] In the formula ，

[0080] Q yn 、Q zn Represents the shear forces in the y and z directions generated by aerodynamic forces.

[0081] M yn 、M zn Indicates the bending moments in the y and z directions generated by aerodynamic forces.

[0082] q Indicates flight pressure

[0083] s Represents aerodynamic reference area

[0084] CN i CZ i CA i Aerodynamic coefficients representing normal, lateral, and axial directions.

[0085] CN, CZ, CA This represents the sum of the aerodynamic coefficients in the normal, lateral, and axial directions.

[0086] x T 、x y Indicates the rocket's center of mass and center of aerodynamic pressure.

[0087] J y 、J z Represents the longitudinal and lateral rotational inertia of the rocket.

[0088] i Represents the site serial number, with a value ranging from 1 to... n natural numbers

[0089] m i and x i This indicates the quality and location of each substation.

[0090] M It is the total mass of the entire rocket

[0091] x y and x z These represent the aerodynamic pressure centers in the y and z directions, respectively.

[0092] The shear force and bending moment caused by the rocket's control forces during flight can be calculated using the following formulas.

[0093] (5)

[0094] (6)

[0095] (7)

[0096] (8)

[0097] In the formula, Q yn 、Q zn Indicates the y and z directions of the force generated by the control. Shear force

[0098] M yn 、M zn Indicates the bending moments in the y and z directions generated by the control force.

[0099] R c Indicates control force

[0100] x R Indicates the position where the control force is applied.

[0101] Q y(i-1) It is y to the first i -1 shear force at a cross section.

[0102] The shear force and bending moment in the y and z directions at each load calculation section of the entire rocket can be combined using the following formula:

[0103] (9)

[0104] (10)

[0105] The axial force during rocket flight can be calculated using the following formula.

[0106] (11)

[0107] (12)

[0108] In the formula, T n Indicates rocket axial force

[0109] P represents engine thrust.

[0110] CAi Indicates the aerodynamic coefficient in the axial direction

[0111] CA The sum of aerodynamic coefficients in the axial direction

[0112] xf Indicates the position of engine thrust application

[0113] g It is gravitational acceleration. x n Indicates the first n Location of each section of the rocket body;

[0114] Represents the step function, i.e.

[0115] (13).

[0116] Step 5: Compare and analyze the static load calculation results obtained by this method with the initial design flight load calculation results. Due to the difference in the ballistic characteristic point parameters selected by the two load conditions, the static load calculation results obtained by this method and the initial design flight load calculation results may show different magnitudes in the comparison of axial force and bending moment. At this time, the equivalent axial force method is used to judge the difference between the two static loads, so as to verify the initial design flight load calculation results.

[0117] The formula for calculating the equivalent axial force is shown below.

[0118] (14)

[0119] In the formula, T eq Indicates the equivalent axial force. T eq A positive value indicates equivalent axial compression, and a negative value indicates equivalent axial tension.

[0120] D This indicates the diameter of the cross section used for calculating the loads on the rocket body.

[0121] The equivalent axial force described in the method of this invention takes into account the combined effect of axial force and bending moment, where bending moment is the accumulation of the product of shear force and distance.

[0122] Based on the equivalent axial force comparison results, if the equivalent axial force of the initial design load is less than or does not exceed 5% of the static load calculation result obtained by this method, the initial design load calculation result is considered correct and valid; otherwise, the static load obtained by this method is used as the basis for structural strength verification and the structural design is carried out again. Figure 2 and Figure 3 This is an example of the comparison results of static loads obtained by the two methods.

Claims

1. A method of flight static load verification of a launch vehicle, characterized by Includes the following steps: The factors affecting the flight payload of the launch vehicle are analyzed to determine the overall parameter deviation items and deviation ranges; Using the Monte Carlo method, random numbers are generated for each deviation based on the probability deviation distribution of different overall parameters; for the various forces and torques acting on the rocket during motion, a six-degree-of-freedom simulation model of the trajectory containing six dynamic motion differential equations is established to simulate the trajectory in six degrees of freedom. The ballistics obtained from the six-degree-of-freedom simulation are screened to remove those with deviations exceeding the set deviation range. Based on the screened ballistics, load calculation conditions are selected, including the maximum point of the product of dynamic pressure and total angle of attack, the maximum point of lateral overload, the maximum point of dynamic pressure, and the maximum point of axial overload. Based on the mass substation data and distributed aerodynamic load data of the selected load calculation conditions, the verification static loads on each section of the rocket under each load calculation condition are solved using general mechanics methods and d'Alembert's principle. The calculation results of the verification static load are compared and analyzed with the calculation results of the initial design flight static load. The difference between the calculation results of the verification static load and the calculation results of the initial design flight static load is determined by the equivalent axial force method, thereby verifying the calculation results of the initial design flight static load. The static load verification of each section of the rocket body includes axial force, shear force, and bending moment. The calculation method of the verification load is as follows: Based on the mass and center of mass characteristics of each section of the rocket, the rocket is reduced to multiple equivalent mass points. The rocket is divided into stations, thus discretizing the rocket into a single beam composed of a finite number of mass points of different masses. Each mass point is called a discretized station. The aerodynamic force and inertial force on the rocket body are distributed to each station. Starting from the tip of the rocket head, the axial force, shear force, and bending moment of each section are calculated station by station. The axial force of the current station section is the cumulative value of the axial force of each station before this section. The shear force of the current station section is the cumulative value of the normal force of each station before this section. The bending moment of the current station section is the cumulative value of the product of the bending moment generated by each station on this station section, i.e., the normal force and the distance. The shear force and bending moment generated by aerodynamic forces during rocket flight can be calculated using the following formulas. .......................（1） .......................（2） .......................（3） .......................（4） In the formula ， Q yn 、Q zn Represents the shear forces in the y and z directions generated by aerodynamic forces. M yn 、M zn Indicates the bending moments in the y and z directions generated by aerodynamic forces. q Indicates flight pressure s Represents aerodynamic reference area CN i CZ i CA i Aerodynamic coefficients representing normal, lateral, and axial directions. CN, CZ, CA This represents the sum of the aerodynamic coefficients in the normal, lateral, and axial directions. x T 、x y Indicates the rocket's center of mass and aerodynamic pressure center J y 、J z Represents the longitudinal and lateral rotational inertia of the rocket. i Represents the site serial number, with a value ranging from 1 to... n natural numbers m i and x i This indicates the quality and location of each substation. M It is the total mass of the entire rocket x y and x z These represent the aerodynamic center of pressure in the y and z directions, respectively; The shear force and bending moment caused by the rocket's control forces during flight can be calculated using the following formulas. （5） （6） （7） （8） In the formula, Q yn 、Q zn Indicates the y- and z-axis shear generated by the control force force M yn 、M zn Indicates the bending moments in the y and z directions generated by the control force. R c Indicates control force x R Indicates the position where the control force is applied. Q y(i-1) It is y to the first i -1 section shear force; The shear forces and bending moments in the y and z directions of each load calculation section of the entire rocket can be combined using the following formula: （9） （10） The axial force during rocket flight can be calculated using the following formula. （11） （12） In the formula, T n Indicates rocket axial force P represents engine thrust. CAi Indicates the aerodynamic coefficient in the axial direction CA The sum of aerodynamic coefficients in the axial direction x f Indicates the position of engine thrust application g It is gravitational acceleration. x n Indicates the first n Location of each section of the rocket body; Represents the step function, i.e. （13）； The formula for calculating the equivalent axial force is as follows: （14） In the formula, T eq This represents the equivalent axial force; a positive value represents the equivalent axial compressive force, and a negative value represents the equivalent axial tensile force. D Indicates the diameter of the cross section used for calculating each load on the rocket body; According to the equivalent axial force comparison results, if the equivalent axial force of the initial design load is less than or does not exceed 5% of the verification static load calculation result obtained by this method, the initial design load calculation result is considered to be correct and valid; otherwise, the verification static load obtained by this method is used as the basis for structural strength verification and the structural design is carried out again.

2. A computer-readable storage medium, characterized in that, The computer-readable storage medium includes a stored program, wherein the program, when executed, performs the launch vehicle flight static load verification method according to claim 1.

3. An electronic device comprising a memory and a processor, characterized in that, The memory stores a computer program, and the processor is configured to execute the launch vehicle flight static load verification method according to claim 1 through the computer program.

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