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Long-range rocket launching initial state error spreading estimation method

A rocket launch and error propagation technology, applied in computing, special data processing applications, instruments, etc., can solve the problems of inability to fully analyze the propagation mechanism, low efficiency of landing point calculation, and speed deviation

Active Publication Date: 2015-11-18
HARBIN INST OF TECH
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

[0004] The purpose of the present invention is to solve the low calculation efficiency of the shutdown point position deviation, speed deviation, vertical deviation and lateral deviation of the landing point caused by the existing launch initial state error, and the inability to fully analyze the propagation of the launch initial state error in the trajectory design process. Mechanism, and proposed a long-range rocket launch initial state error propagation estimation method

Method used

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  • Long-range rocket launching initial state error spreading estimation method
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  • Long-range rocket launching initial state error spreading estimation method

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specific Embodiment approach 1

[0025] Specific implementation mode one: combine figure 1 Description of this embodiment, a method for estimating the initial state error propagation of a long-range rocket launch, is characterized in that a method for estimating the initial state error propagation of a long-range rocket launch is specifically carried out according to the following steps:

[0026] Step 1, establishing the kinetic perturbation equation;

[0027] Step 2: According to the dynamic perturbation equation obtained in step 1, solve the thrust acceleration deviation of the remote rocket, the aerodynamic acceleration deviation of the remote rocket, the normal gravitational acceleration deviation of the remote rocket, the Coriolis acceleration deviation of the remote rocket and the centrifugal acceleration deviation of the remote rocket;

[0028] Step 3. According to the dynamic perturbation equation obtained in step 1 and the long-range rocket thrust acceleration deviation, long-range rocket aerodynami...

specific Embodiment approach 2

[0029] Specific embodiment two: the difference between this embodiment and specific embodiment one is: the kinetic perturbation equation is established in the step one; the specific process is:

[0030] Let the nominal launch coordinate system be o 1 -x 1 the y 1 z 1 , o 1 is the origin of the nominal launch coordinate system, x 1 is the nominal emission coordinate system x-axis, y 1 is the y-axis of the nominal launch coordinate system, z 1 It is the z-axis of the nominal emission coordinate system, and the actual emission coordinate system is o 2 -x 2 the y 2 z 2 , o 2 is the origin of the actual launch coordinate system, x 2 is the actual emission coordinate system x-axis, y 2 is the y-axis of the actual launch coordinate system, z 2 is the z-axis of the actual emission coordinate system, such as figure 2 shown;

[0031] The difference between the nominal launch coordinate system and the actual launch coordinate system reflects the initial launch error, whic...

specific Embodiment approach 3

[0037] Specific embodiment three: the difference between this embodiment and specific embodiment one or two is that in the step two, according to the dynamic perturbation equation obtained in step one, the long-range rocket thrust acceleration deviation, the long-range rocket aerodynamic acceleration deviation, the long-range rocket aerodynamic acceleration deviation, and the remote The normal gravitational acceleration deviation of the rocket, the Coriolis acceleration deviation of the long-range rocket and the centrifugal acceleration deviation of the long-range rocket; the specific process is:

[0038] (1) Thrust acceleration deviation and aerodynamic acceleration deviation

[0039] is the launch initial state error vector Δ P → s = Δλ 0 ...

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Abstract

The invention provides a long-range rocket launching initial state error spreading estimation method, and relates to the long-range rocket launching initial state error spreading estimation method. The method aims at solving the problems that the calculation efficiency of burnout point position deviation, speed deviation, landing point longitudinal deviation and transverse deviation is low due to the existing launching initial state error, so that the spreading mechanism of the launching initial state error in the trajectory design process cannot be sufficiently analyzed. The goal of the method is achieved through the following technical scheme that the method comprises the following steps of: 1, building a kinetics perturbation equation; 2, solving long-range rocket thrust acceleration deviation, pneumatic acceleration deviation, normal gravitational acceleration deviation, Coriolis acceleration deviation and centrifugal acceleration deviation; and 3, obtaining similar analytical solutions of the burnout point position deviation and the speed deviation due to the long-range rocket launching initial state error and similar analytical solutions of the landing point longitudinal deviation and transverse deviation according to the first step and the second step. The method is applied to the field of long-range rocket or carrier rocket flight dynamics.

Description

technical field [0001] The invention relates to a method for estimating the initial state error propagation of a long-distance rocket launch. Background technique [0002] Long-range rockets are generally established in the ground launch coordinate system in ballistic design and calculation. However, due to the existence of launch initial state errors, the actual launch coordinate system deviates from the nominal launch coordinate system. The initial launch error includes initial positioning error (geographic latitude and longitude deviation, elevation deviation) and initial orientation error (perpendicular deviation, launch azimuth deviation). The error of the launch initial state will bring the deviation of each acceleration in the launch coordinate system, that is, the deviation of gravitational acceleration, thrust acceleration, aerodynamic acceleration, Coriolis acceleration and centrifugal acceleration, which in turn will cause the state deviation of the shutdown point...

Claims

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Application Information

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IPC IPC(8): G06F19/00
Inventor 荆武兴郑旭高长生常晓华
Owner HARBIN INST OF TECH
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