CN109214014B - Method, system and equipment for acquiring residual track life of near-earth track space object - Google Patents

Abstract

The present invention provides a method for obtaining the remaining orbit life of a space object in low-Earth orbit, which includes: obtaining the semi-major axis change rate, eccentricity change rate, orbital elements at the current moment and the atmosphere area-to-mass ratio of the space object according to the effect of atmospheric resistance ; Use the numerical integration algorithm to numerically integrate the rate of change of the semi-major axis and the rate of change of the eccentricity with a predetermined integration step, and obtain the atmospheric density, and then integrate from the current moment to the nth predetermined integration step to obtain the nth predetermined integral The semi-major axis and eccentricity of the orbit where the space object is located in the step length, wherein the predetermined integral step is the ratio of the semi-major axis step length to the semi-major axis change rate; calculate the product between the semi-major axis and the eccentricity and determine the semi-major axis The difference between the axis and the product, the perigee height of the space object is obtained as the difference between the difference and the radius of the earth; when the perigee height is less than the preset perigee height, the remaining orbital life of the space object is obtained as the time corresponding to the nth predetermined integration step difference from the initial time.

Description

Method, system and device for obtaining the remaining orbital life of a space object in near-Earth orbit

Technical Field

The present invention relates to the field of spacecraft orbital dynamics, and in particular to a method, system and device for obtaining the remaining orbital life of a space object in a near-Earth orbit.

Background Art

Orbital lifetime refers to the time a space object remains in orbit. It is the time from when a space object enters orbit to when it falls. Some space objects in orbit are mainly affected by atmospheric drag, and their actual orbits present a spiral line that continues to descend, causing the perigee height to gradually decrease, eventually causing the space object to fall into the dense atmosphere (100 kilometers from the ground) and burn up. Larger space objects may not be completely burned up in the dense atmosphere, and thus fall to the ground, causing harm to the safety of people and property on the ground.

The traditional calculation of the remaining orbital lifetime of space objects in low-Earth orbit has a large gap compared with the actual observation data. The orbital lifetime obtained is not accurate enough and the calculation speed is slow.

Summary of the invention

Based on this, it is necessary to provide a method, system and device for obtaining the remaining orbital life of a space object in low-Earth orbit with fast calculation speed and accuracy in response to the above technical problems, wherein the method comprises:

According to the atmospheric drag effect, the semi-major axis change rate and the eccentricity change rate of the space object are obtained, and the orbital elements and the atmospheric area mass ratio of the space object at the current moment are calculated and obtained;

numerically integrating the semi-major axis change rate and the eccentricity change rate with a predetermined integration step using a numerical integration algorithm, and obtaining the atmospheric density, and then integrating from the current moment to the nth predetermined integration step to obtain the semi-major axis and eccentricity of the orbit of the space object at the nth predetermined integration step, wherein the predetermined integration step is the ratio of the semi-major axis step to the semi-major axis change rate;

Calculating the product of the semi-major axis and the eccentricity and determining the difference between the semi-major axis and the product, obtaining the perigee height of the space object as the difference between the difference and the radius of the earth; and

When the perigee height is less than a preset perigee height, the remaining orbital life of the space object is obtained as the difference between the time corresponding to the nth predetermined integration step and the initial time.

In one embodiment, obtaining the atmospheric density includes:

Obtain a preset atmospheric density model and space environment parameters, perform numerical integration with the true near angle of the orbit of the space object as a variable, and obtain the atmospheric density corresponding to the nth predetermined integration step, wherein the solar activity parameter F10.7 and the geomagnetic index AP corresponding to the true near angle after the variable changes are obtained.

In one embodiment, in the step of obtaining the semi-major axis change rate and the eccentricity change rate of the space object according to the atmospheric drag effect, and obtaining the orbital elements and the atmospheric area mass ratio of the space object at the current moment, obtaining the atmospheric area mass ratio includes:

Obtaining the orbital elements of the space object within a preset epoch time range, sorting the orbital elements according to the epoch time, and calculating the semi-major axis difference and epoch time difference between two adjacent orbital elements;

According to the atmospheric density model, obtaining atmospheric density parameters and space environment parameters corresponding to the epoch time, so as to obtain a first rate of change of the orbital semi-major axis caused by the mass ratio per unit atmospheric area;

According to the solar light pressure perturbation model and the cylindrical earth shadow model, the positional relationship between the sun, the earth and the space object is obtained to obtain the second rate of change of the orbital semi-major axis caused by the unit light pressure area mass ratio;

According to the semi-major axis difference, time difference, first change rate and second change rate, the atmosphere area-to-mass ratio and the light pressure area-to-mass ratio are obtained by the least square method.

In one embodiment, the steps of obtaining the orbital elements of the orbit of the space object within a preset epoch time range, sorting the orbital elements according to the epoch time, and calculating the semi-major axis difference and epoch time difference between two adjacent orbital elements include:

Obtaining the number of orbital elements of the orbit of the space object within a preset epoch time range;

Sorting the orbital elements according to epoch time;

Calculating the epoch time difference between two adjacent orbital elements; and

When the epoch time difference is greater than or equal to 1 day, the semi-major axis difference between two adjacent orbital elements is calculated.

In one embodiment, the method for obtaining the remaining orbital lifetime of a low-Earth orbit space object further comprises:

When the perigee height is greater than or equal to the preset perigee height, n+1 is assigned to n, and the step of numerically integrating the semi-major axis change rate and the eccentricity change rate using the numerical integration algorithm with the predetermined integration step is returned to obtain the atmospheric density, and then integrating from the current moment to the nth predetermined integration step to obtain the semi-major axis and eccentricity of the orbit of the space object at the nth predetermined integration step, wherein the predetermined integration step is the ratio of the semi-major axis step to the semi-major axis change rate.

The present invention also provides a system for obtaining the remaining orbital life of a space object, comprising:

A parameter acquisition module, used to obtain the semi-major axis change rate and eccentricity change rate of the space object according to the atmospheric drag effect, and to obtain the orbital elements and the atmospheric area mass ratio of the space object at the current moment;

an integration module, configured to numerically integrate the semi-major axis change rate and the eccentricity change rate with a predetermined integration step length by using a numerical integration algorithm, and obtain the atmospheric density, and then integrate from the current moment to the nth predetermined integration step length to obtain the semi-major axis and eccentricity of the orbit of the space object at the nth predetermined integration step length, wherein the predetermined integration step length is the ratio of the semi-major axis step length to the semi-major axis change rate;

a perigee height obtaining module, configured to calculate the product of the semi-major axis and the eccentricity and determine the difference between the semi-major axis and the product, and obtain the perigee height of the space object as the difference between the difference and the radius of the earth; and

The remaining life acquisition module is used to obtain the remaining orbital life of the space object as the difference between the time corresponding to the nth predetermined integration step and the initial time when the perigee height is less than the preset perigee height.

In one embodiment, the integration module further includes:

The atmospheric density acquisition module is used to obtain a preset atmospheric density model and space environment parameters, perform numerical integration with the true near angle of the orbit of the space object as a variable, and obtain the atmospheric density corresponding to the nth predetermined integration step, wherein the solar activity parameter F10.7 and the geomagnetic index AP corresponding to the true near angle after the variable changes are obtained.

In one embodiment, in the parameter acquisition module, acquiring the atmosphere area mass ratio comprises:

A semi-major axis difference obtaining module is used to obtain the orbital elements of the orbit of the space object within a preset epoch time range, sort the orbital elements according to the epoch time, and when the epoch time difference is greater than or equal to 1 day, calculate the semi-major axis difference and epoch time difference between two adjacent orbital elements;

The first change rate acquisition module of the orbital semi-major axis is used to obtain the atmospheric density parameters and the space environment parameters corresponding to the epoch time according to the atmospheric density model, and then obtain the first change rate of the orbital semi-major axis caused by the mass ratio per unit atmospheric area;

a module for obtaining a second change rate of the orbital semi-major axis, for obtaining the positional relationship between the sun, the earth and the space object according to a solar light pressure perturbation model and a cylindrical earth shadow model, and further obtaining a second change rate of the orbital semi-major axis caused by a unit light pressure area mass ratio; and

The atmospheric area mass ratio and light pressure area mass ratio obtaining module is used to obtain the atmospheric area mass ratio and light pressure area mass ratio by the least square method according to the semi-major axis difference, time difference, first change rate and second change rate.

In one of the embodiments, the remaining life acquisition module is also used to assign n+1 to n when the perigee altitude is greater than or equal to a preset perigee altitude, return to the step of numerically integrating the semi-major axis change rate and the eccentricity change rate using a numerical integration algorithm with a predetermined integration step size, and obtain the atmospheric density, and then integrate from the current moment to the nth predetermined integration step size to obtain the semi-major axis and eccentricity of the orbit of the space object at the nth predetermined integration step size, wherein the predetermined integration step size is the ratio of the semi-major axis step size to the semi-major axis change rate.

The present invention also provides a device for obtaining the remaining orbital life of a near-Earth orbit space object, comprising a processor, a memory, and a computer program stored in the memory, characterized in that the computer program implements the steps of any one of the aforementioned methods when processed by the processor.

Since the semi-major axis change rate varies with the perigee height, in a method, system and device for obtaining the remaining orbital lifetime of a low-Earth orbit space object provided by the present invention, the ratio of the semi-major axis step size to the semi-major axis change rate is used as the predetermined integration step size when performing numerical integration, which can make it faster and more accurate to obtain the remaining orbital lifetime of a low-Earth orbit space object.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG1 is a flow chart of a method for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit according to an embodiment of the present invention;

FIG2 is a flow chart of obtaining the atmosphere area mass ratio in the method for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit according to an embodiment of the present invention;

3 is a flow chart of obtaining the semi-major axis difference and epoch time difference of two adjacent orbital elements in the method for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit according to an embodiment of the present invention;

FIG4 is a flow chart of a method for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit in another embodiment of the present invention;

FIG5 is a structural block diagram of a system for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit according to an embodiment of the present invention;

FIG6 is a structural block diagram of obtaining the atmosphere area mass ratio in a system for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit according to an embodiment of the present invention;

FIG. 7 is a comparison of the calculation results of an embodiment of the present invention and the results published by space-track.org.

DETAILED DESCRIPTION

In order to make the purpose, technical solution and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention.

FIG1 is a flow chart of a method for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit according to an embodiment of the present invention. The method comprises:

Step S100, obtaining the semi-major axis change rate and eccentricity change rate of the space object according to the atmospheric drag effect, and obtaining the orbital elements and the atmospheric area-mass ratio of the space object at the initial moment.

Specifically, according to the atmospheric drag effect, the semi-major axis change rate and the eccentricity change rate of the space object are obtained, and the semi-major axis change rate and the eccentricity change rate are expressed as:

Among them, σ＝C _D A/2 is the ballistic coefficient, C _D is the drag coefficient of the space object, in this embodiment, C _D =2.2, A is the area-mass ratio of the space object, ρ is the atmospheric density at the height, v is the speed of the space object, r is the radial distance of the space object, a is the semi-major axis of the orbit of the space object, e is the eccentricity of the orbit of the space object, μ is the earth's gravitational constant, and θ is the true anomaly of the orbit of the space object. When integrating the semi-major axis change rate and eccentricity change rate equations over time, it is necessary to obtain the orbital elements of the space object at the initial moment, that is, the semi-major axis and eccentricity at the initial moment and related parameters such as the atmospheric area-mass ratio. And the orbital semi-major axis and eccentricity in the orbital elements at the current moment should meet the low-Earth orbit (LEO) conditions: a＜8378km and e＜0.1km.

Step S200, numerically integrate the semi-major axis change rate and the eccentricity change rate using a numerical integration algorithm with a predetermined integration step, and obtain the atmospheric density, and then integrate from the current moment to the nth predetermined integration step to obtain the semi-major axis and eccentricity of the orbit of the space object at the nth predetermined integration step, wherein the predetermined integration step is the ratio of the semi-major axis step to the semi-major axis change rate.

Specifically, the predetermined integration step length may be a fixed length value. In one embodiment, the predetermined integration step length is preferably the ratio of the semi-major axis step length to the semi-major axis change rate. In this embodiment, the numerical integration algorithm is the Runge-Kutta fourth-order integration method. The semi-major axis change rate and the eccentricity change rate are integrated from the current moment to the nth predetermined integration step length to obtain the semi-major axis and eccentricity of the space object at the nth predetermined integration step length.

Since the semi-major axis change rate varies with the perigee height, in the method for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit provided by the present invention, the ratio of the semi-major axis step size to the semi-major axis change rate is used as the predetermined integration step size during numerical integration, so that the remaining orbital lifetime of the space object in a low-Earth orbit can be obtained faster and more accurately.

Step S300, calculating the product of the semi-major axis and the eccentricity and determining the difference between the semi-major axis and the product, and obtaining the perigee height of the space object as the difference between the difference and the radius of the earth.

Specifically, the perigee height of the space object is obtained according to the semi-major axis and eccentricity obtained in step S200, and the perigee height of the space object is obtained by the perigee height formula r _p =a*(1-e) _-Re , where _Re is the radius of the earth.

Step S400, when the perigee height is less than a preset perigee height, the remaining orbital life of the space object is obtained as the difference between the time corresponding to the nth predetermined integration step and the initial time.

Specifically, the remaining orbital life of the space object is determined by the perigee height, and when the perigee height is less than a preset perigee height, the remaining orbital life of the space object can be obtained. Preferably, when r _p <100 km, the remaining orbital life of the space object is obtained as the difference between the time corresponding to the nth predetermined integration step and the initial time.

In one embodiment, step S200 also includes obtaining a preset atmospheric density model and space environment parameters, performing numerical integration with the true near angle of the orbit of the space object as a variable, and obtaining the atmospheric density corresponding to the nth predetermined integration step, wherein the solar activity parameter F10.7 and the geomagnetic index AP corresponding to the true near angle after the change of the variable are obtained.

Specifically, in this embodiment, the atmospheric density is a function of the true near angle and is related to the solar activity parameters and the geomagnetic index. The atmospheric density is also related to the position of the space object at the time corresponding to the nth predetermined integral step. The preset atmospheric density model and space environment parameters give the corresponding values of these related parameters. The atmospheric density at the time corresponding to the nth predetermined integral step can be finally obtained by numerically integrating the true near angle. The preset atmospheric model is the NRLMSISE2000 standard earth atmosphere model, and the preset space environment parameters are space weather parameters. The solar extreme ultraviolet effect parameters corresponding to the true near angle of the orbit where the space object is located under different true near angles, namely the solar activity parameter F10.7, the geomagnetic index AP, and the orbital height are obtained, and then the atmospheric density is given according to the NRLMSISE2000 standard earth atmosphere model. When the conventional calculation of the remaining life of the orbit uses a constant for the atmospheric density or does not calculate the atmospheric density according to the prediction results of the space environment parameters, the final calculation result is inaccurate. The present invention is more scientific to obtain the atmospheric density according to different true near angles at different times, and also makes the remaining life of the orbit finally obtained more accurate.

As shown in FIG. 2 , in one embodiment, in the step S100 of obtaining the semi-major axis change rate and the eccentricity change rate of the space object according to the atmospheric drag effect, and obtaining the orbital elements and the atmospheric area mass ratio of the space object at the current moment, the obtaining of the atmospheric area mass ratio includes:

Step S110, obtaining the orbital elements of the orbit of the space object within a preset epoch time range, sorting the orbital elements according to the epoch time, and obtaining the semi-major axis difference and epoch time difference between two adjacent orbital elements.

Specifically, the orbital element number of the orbit of the space object within a period of epoch time is selected. This orbital element number is the average orbital element number. The orbital elements within this period of epoch time are sorted according to the epoch time, and then the semi-major axis difference and epoch time difference between two adjacent orbital elements are calculated.

As shown in FIG. 3 , in one embodiment, step S110 includes:

Step S111, obtaining the orbital elements of the orbit of the space object within a preset epoch time range;

Step S112, sorting the orbital elements according to epoch time;

Step S113, calculating and obtaining the epoch time difference between two adjacent orbital elements; and

Step S114, when the epoch time difference is greater than or equal to 1 day, calculating the semi-major axis difference between two adjacent orbital elements.

Step S120, obtaining the atmospheric density parameters and the space environment parameters corresponding to the epoch time according to the atmospheric density model, and then obtaining the first change rate of the orbit semi-major axis caused by the mass ratio per unit atmospheric area.

Specifically, according to the orbital elements at the corresponding epoch, the atmospheric density model NRLMSISE 2000 and the space environment parameter space weaher, according to the atmospheric drag model, the equation for the average rate of change of the semi-major axis over time caused by the mass ratio per unit atmospheric area is as follows:

Wherein, _CD is the drag coefficient of the space object, in this embodiment _CD = 2.2, v is the velocity of the space object, r is the radial distance of the space object, a is the semi-major axis of the orbit of the space object, μ is the earth's gravitational constant, and θ is the true anomaly angle of the orbit of the space object.

Step S130, according to the solar radiation pressure perturbation model and the cylindrical earth shadow model, the positional relationship between the sun, the earth and the space object is obtained, and then the second change rate of the orbital semi-major axis caused by the unit radiation pressure area mass ratio is obtained.

According to the cylindrical earth shadow model, the positional relationship between the sun, the earth and space objects, and the light pressure perturbation model, the equation for the average rate of change of the semi-major axis over time caused by the unit light pressure area mass ratio at the corresponding epoch time under the influence of the solar light pressure perturbation is calculated as follows:

P_SR＝4.57×10^-6N/m²为太阳辐射压常数，C_R为太阳光压常数(本实施例中取1.40)，R_sat2sun为卫星指向太阳的位置矢量。参考圆柱形地球阴影模型，当卫星运行在地球的背面，没有太阳光照射时(即阴影区)，卫星受到的太阳光压摄动影响为0。Among them, λ is expressed as:

P _SR =4.57×10 ^-6 N/m ² is the solar radiation pressure constant, _CR is the solar light pressure constant (taken as 1.40 in this embodiment), and R _sat2sun is the position vector of the satellite pointing to the sun. Referring to the cylindrical earth shadow model, when the satellite is operating on the back of the earth without sunlight (i.e., in the shadow area), the solar light pressure perturbation effect on the satellite is 0.

Step S140, obtaining the atmosphere area-to-mass ratio and the light pressure area-to-mass ratio by the least square method according to the semi-major axis difference, the time difference, the first change rate and the second change rate.

Specifically, according to the semi-major axis difference Δa _i , the time difference Δt _i , the first change rate and the second change rate, the atmosphere area mass ratio A ₁ and the light pressure area mass ratio A ₂ of the space object within this range, the equation can be established:

Where i = 1, 2...N-1 corresponds to each epoch time difference. The least square method is used to calculate the above equation to obtain the optimal values of the atmosphere area mass ratio _A1 and the light pressure area mass ratio _A2 .

Therefore, in the method for obtaining the remaining orbital lifetime of a space object in a low-Earth orbit provided by the present invention, the method for obtaining the optimal values of the atmosphere area mass ratio _A1 and the light pressure area mass ratio _A2 is adopted when performing numerical integration, so that the remaining orbital lifetime of the space object in a low-Earth orbit is obtained more accurately.

As shown in FIG4 , in one embodiment, the method for obtaining the remaining orbital lifetime of a low-Earth orbit space object further includes:

Step S500, when the perigee height is greater than or equal to the preset perigee height, assign n+1 to n, return to the step of numerically integrating the semi-major axis change rate and the eccentricity change rate using the numerical integration algorithm with the predetermined integration step, and obtain the atmospheric density, and then integrate from the current moment to the nth predetermined integration step to obtain the semi-major axis and eccentricity of the orbit of the space object at the nth predetermined integration step, wherein the predetermined integration step is the ratio of the semi-major axis step to the semi-major axis change rate.

Specifically, generally speaking, one integration step cannot satisfy the condition that the perigee height is less than the preset perigee height. At this time, it is necessary to integrate to the next integration step. Therefore, it is necessary to assign n+1 to n and return to the step S200 to continue the integration calculation.

The present invention also provides a system for obtaining the remaining orbital life of a space object in a near-Earth orbit.

FIG5 is a structural block diagram of a system for obtaining the remaining orbital lifetime of a space object in a near-Earth orbit according to an embodiment of the present invention. The system includes:

The parameter acquisition module 100 is used to obtain the semi-major axis change rate and the eccentricity change rate of the space object according to the atmospheric drag effect, and obtain the orbital elements and the atmospheric area mass ratio of the space object at the current moment;

An integration module 200 is used to numerically integrate the semi-major axis change rate and the eccentricity change rate with a predetermined integration step length by using a numerical integration algorithm, and obtain the atmospheric density, and then integrate from the current moment to the nth predetermined integration step length to obtain the semi-major axis and eccentricity of the orbit of the space object at the nth predetermined integration step length, wherein the predetermined integration step length is the ratio of the semi-major axis step length to the semi-major axis change rate;

a perigee height obtaining module 300, configured to calculate the product of the semi-major axis and the eccentricity and determine the difference between the semi-major axis and the product, and obtain the perigee height of the space object as the difference between the difference and the radius of the earth; and

The remaining life acquisition module 600 is used to obtain the remaining orbital life of the space object as the difference between the time corresponding to the nth predetermined integration step and the initial time when the perigee height is less than the preset perigee height.

In one embodiment, the integration module 200 further includes:

The atmospheric density acquisition module is used to obtain a preset atmospheric density model and space environment parameters, perform numerical integration with the true near angle of the orbit of the space object as a variable, and obtain the atmospheric density corresponding to the nth predetermined integration step, wherein the solar activity parameter F10.7 and the geomagnetic index AP corresponding to the true near angle after the variable changes are obtained.

As shown in FIG6 , in one embodiment, in the parameter acquisition module 100, the atmosphere area mass ratio includes:

The semi-major axis difference obtaining module 110 is used to obtain the orbital elements of the orbit of the space object within a preset epoch time range, sort the orbital elements according to the epoch time, and when the epoch time difference is greater than or equal to 1 day, calculate the semi-major axis difference and epoch time difference between two adjacent orbital elements;

The first change rate obtaining module 120 of the orbital semi-major axis is used to obtain the atmospheric density parameters and the space environment parameters corresponding to the epoch time according to the atmospheric density model, and then obtain the first change rate of the orbital semi-major axis caused by the mass ratio per unit atmospheric area;

The second change rate obtaining module 130 of the orbital semi-major axis is used to obtain the positional relationship between the sun, the earth and the space object according to the solar light pressure perturbation model and the cylindrical earth shadow model, and then obtain the second change rate of the orbital semi-major axis caused by the unit light pressure area mass ratio;

The atmosphere area mass ratio and light pressure area mass ratio obtaining module 140 is used to obtain the atmosphere area mass ratio and light pressure area mass ratio by least square method according to the semi-major axis difference, time difference, first change rate and second change rate.

In one embodiment, the remaining life acquisition module 400 is also used to assign n+1 to n when the perigee height is greater than or equal to a preset perigee height, return to the step of numerically integrating the semi-major axis change rate and the eccentricity change rate using a numerical integration algorithm with a predetermined integration step, and obtain the atmospheric density, and then integrate from the current moment to the nth predetermined integration step to obtain the semi-major axis and eccentricity of the orbit of the space object at the nth predetermined integration step, wherein the predetermined integration step is the ratio of the semi-major axis step to the semi-major axis change rate.

The present invention also provides a device for obtaining the remaining orbital life of a near-Earth orbit space object, comprising a processor, a memory, and a computer program stored in the memory, characterized in that the computer program implements the steps of any one of the aforementioned methods when processed by the processor.

According to the above specific embodiment, a specific application case is given as follows:

The orbital elements of Cosmos 1151 (international number 11671) from April to July 2015 are used to estimate its atmospheric area mass ratio and photometric area mass ratio using this method, and then its remaining orbital lifetime is calculated. The method is verified and explained in detail, and the actual fall period of the object 11671 is given. The target orbital elements used are all from the TLE file downloaded from the Internet, and the space environment parameters are from the space-weather.txt file downloaded from the Internet.

a) Orbital element sorting and screening. Using the orbital elements of Cosmos 1151 from April to July 2015, after sorting and screening, the semi-major axis difference and time difference results are calculated as shown in the following table:

b) Calculation of the average change rate of the semi-major axis caused by the mass ratio per unit atmospheric area and the mass ratio per unit light pressure. The average change of the semi-major axis caused by the mass ratio per unit atmospheric area is calculated using the space environment parameters, orbital elements and the atmospheric model NRLMSISE 2000; the average change of the semi-major axis caused by the mass ratio per unit light pressure is calculated using the cylindrical earth shadow model and orbital elements. The results are shown in the following table:

c) Calculation of the atmospheric and light pressure area mass ratios: According to the calculation contents in the above table, the least square method was used to calculate the atmospheric area mass ratio and light pressure area mass ratio of 11671, which were 0.00616 (m2/kg) and 36.45092 (m2/kg), respectively.

(2) Estimation of the remaining orbital lifetime of Cosmos 1151

The orbital elements of Cosmos 1151 at 15 different times were selected, and the remaining orbital lifetime was estimated using the atmospheric area mass ratio of 0.00616 (square meters/kg) obtained in step (1). The results are shown in the following table:

Forecast initial epoch Predicted fall period Predicted remaining orbital lifetime (days) Actual fall period 2014-4-23 2014-7-31 98 2014-8-4 2014-5-14 2014-7-29 74 2014-8-4 2014-5-22 2014-7-28 66 2014-8-4 2014-5-28 2014-7-28 61 2014-8-4 2014-6-9 2014-7-30 51 2014-8-4 2014-6-11 2014-7-30 48 2014-8-4 2014-6-18 2014-7-30 42 2014-8-4 2014-6-25 2014-7-31 36 2014-8-4 2014-7-2 2014-8-1 30 2014-8-4 2014-7-9 2014-8-1 twenty three 2014-8-4 2014-7-18 2014-8-1 13 2014-8-4 2014-7-31 2014-8-4 4 2014-8-4 2014-8-1 2014-8-4 3 2014-8-4 2014-8-2 2014-8-4 2 2014-8-4 2014-8-3 2014-8-4 1 2014-8-4

As shown in Figure 7, the calculation results of an embodiment of the present invention are compared with the results published by space-track.org. This figure shows that the present invention is closer to the actual observation results and more accurate.

The above embodiments only express several implementation methods of the present invention, and the descriptions thereof are relatively specific and detailed, but they cannot be understood as limiting the scope of the invention patent. It should be pointed out that, for those of ordinary skill in the art, several variations and improvements can be made without departing from the concept of the present invention, and these all belong to the protection scope of the present invention. Therefore, the protection scope of the patent of the present invention shall be subject to the attached claims.

Claims (10)

1. A method of acquiring remaining orbital life of an object in a near-earth orbital space, the method comprising:

according to the atmospheric resistance, the semi-long axis change rate and the eccentricity change rate of the space object are obtained, and the track number of the space object at the initial moment and the track number of the space object in the preset epoch moment range are obtained;

obtaining an atmospheric area mass ratio by a least square method according to the semi-long axis difference of two adjacent track numbers, the epoch time difference of the two adjacent track numbers, the first change rate of the semi-long axis of the track and the second change rate of the semi-long axis of the track;

numerical integration is carried out on the semi-long axis change rate and the eccentricity ratio change rate by a numerical integration algorithm according to a preset integration step length, the atmospheric density is obtained, and then the semi-long axis and the eccentricity ratio of the track where the space object is located when the preset integration step length is obtained from the initial moment to the nth preset integration step length, wherein the preset integration step length is the ratio of the semi-long axis step length to the semi-long axis change rate;

calculating the product between the semi-long axis and the eccentricity ratio, and determining the difference between the semi-long axis and the product to obtain the near-place height of the space object as the difference between the difference and the earth radius; and

and when the near-ground point height is smaller than the preset near-ground point height, obtaining the residual track life of the space object as the difference value between the corresponding time of the nth preset integral step length and the initial time.

2. The method of acquiring the remaining orbital life of a near-earth orbital space object of claim 1, wherein said acquiring the atmospheric density comprises:

and acquiring a preset atmospheric density model and space environment parameters, and carrying out numerical integration by taking a true near angle of a track where the space object is located as a variable to acquire the atmospheric density corresponding to the nth preset integration step length, wherein a solar activity parameter F10.7 and a geomagnetic index AP corresponding to the true near angle after the variable change are acquired.

3. The method for obtaining the remaining track life of the near-earth track space object according to claim 1, wherein the obtaining the atmospheric area mass ratio by the least square method based on the semi-major axis difference of the adjacent two track numbers, the epoch time difference of the adjacent two track numbers, the first change rate of the semi-major axis of the track, and the second change rate of the semi-major axis of the track comprises:

acquiring the track number of the track where the space object is located in a preset epoch time range, sequencing the track numbers according to epoch time, and calculating to obtain the semi-long axis difference and epoch time difference of two adjacent track numbers;

according to the atmospheric density model, acquiring an atmospheric density parameter and a space environment parameter corresponding to the epoch time so as to obtain a first change rate of a semi-long axis of the orbit caused by a unit atmospheric area mass ratio;

acquiring the position relation of the sun, the earth and the space object according to the solar pressure perturbation model and the cylindrical earth shadow model so as to obtain a second change rate of the orbit semi-long axis caused by the unit light pressure area mass ratio;

and obtaining the air area mass ratio and the light pressure area mass ratio by a least square method according to the semi-long axis difference, the epoch time difference, the first change rate and the second change rate.

4. The method for obtaining the remaining track life of the near-earth track space object according to claim 3, wherein the step of obtaining the track number of the track where the space object is located in the preset epoch time range, sorting the track numbers according to epoch time, and calculating to obtain the semi-major axis difference and epoch time difference of two adjacent track numbers comprises:

acquiring the track number of the track where the space object is located in a preset epoch time range;

sequencing the track number according to epoch time;

calculating to obtain epoch time differences of the adjacent two track numbers; and

and when the epoch time difference is more than or equal to 1 day, calculating the semi-long axis difference of the adjacent two track numbers.

5. The method of acquiring the remaining orbital life of a near earth orbital space object of claim 1, further comprising:

when the near-place height is greater than or equal to the preset near-place height, n+1 is assigned to n, the numerical integration algorithm is adopted to carry out numerical integration on the semi-long axis change rate and the eccentricity ratio by the preset integration step length, the atmospheric density is obtained, and then the semi-long axis and the eccentricity ratio of the orbit where the space object is located when the preset integration step length is obtained from the initial moment to the n preset integration step length, wherein the preset integration step length is the ratio of the semi-long axis step length to the semi-long axis change rate.

6. A system for capturing the remaining track life of a near-earth track space object, comprising:

the parameter acquisition module is used for acquiring the semi-long axis change rate and the eccentricity change rate of the space object according to the atmospheric resistance, and acquiring the track number of the space object at the initial moment and the track number of the track where the space object is located in the preset epoch moment range;

the atmospheric area mass ratio and light pressure area mass ratio obtaining module is used for obtaining the atmospheric area mass ratio through a least square method according to the semi-long axis difference of two adjacent track numbers, the epoch moment difference of the two adjacent track numbers, the first change rate of the semi-long axis of the track and the second change rate of the semi-long axis of the track;

the integration module is used for carrying out numerical integration on the semi-long axis change rate and the eccentricity ratio change rate by adopting a numerical integration algorithm according to a preset integration step length, obtaining the atmospheric density, and further integrating the initial time to the nth preset integration step length to obtain the semi-long axis and the eccentricity ratio of the orbit where the space object is located when the nth preset integration step length is obtained, wherein the preset integration step length is the ratio of the semi-long axis step length to the semi-long axis change rate;

the near-place height obtaining module is used for calculating the product between the semi-long axis and the eccentricity ratio and determining the difference value between the semi-long axis and the product to obtain the near-place height of the space object as the difference between the difference value and the earth radius; and

and the remaining life obtaining module is used for obtaining the remaining track life of the space object as the difference value between the corresponding time of the nth preset integral step and the initial time when the near-place height is smaller than the preset near-place height.

7. The system for acquiring remaining orbital life of a geostationary orbital space object of claim 6 wherein the integration module further comprises:

the atmospheric density obtaining module is used for obtaining a preset atmospheric density model and a space environment parameter, carrying out numerical integration by taking a true near angle of a track where the space object is located as a variable, and obtaining the atmospheric density corresponding to the nth preset integration step length, wherein a solar activity parameter F10.7 and a geomagnetic index AP corresponding to the true near angle after the variable change are obtained.

8. The system for acquiring the remaining orbital life of a near-earth orbital space object of claim 6, wherein the atmospheric area to mass ratio and optical pressure area to mass ratio acquisition module comprises:

the semi-long axis difference obtaining module is used for obtaining the track root number of the track where the space object is located in the preset epoch time range, sequencing the track root numbers according to epoch time, and calculating to obtain the semi-long axis difference and epoch time difference of two adjacent track root numbers when the epoch time difference is more than or equal to 1 day;

the first change rate obtaining module is used for obtaining the atmospheric density parameter and the space environment parameter corresponding to the epoch moment according to the atmospheric density model, so as to obtain the first change rate of the semi-long axis of the orbit caused by the unit atmospheric area mass ratio;

the second change rate obtaining module is used for obtaining the position relation of the sun, the earth and the space object according to the solar pressure perturbation model and the cylindrical earth shadow model, so as to obtain the second change rate of the orbit semi-long axis caused by the unit light pressure area mass ratio; and

the atmospheric area mass ratio and light pressure area mass ratio obtaining module is used for obtaining the atmospheric area mass ratio and the light pressure area mass ratio through a least square method according to the semi-long axis difference, the epoch moment difference, the first change rate and the second change rate.

9. The system for obtaining the remaining orbital life of a near-earth orbit space object according to claim 6, wherein said remaining life obtaining module is further configured to assign n+1 to n when said near-earth altitude is equal to or greater than a preset near-earth altitude, return said step of numerically integrating said semimajor axis change rate and said eccentricity change rate with a predetermined integration step size using a numerical integration algorithm, and obtain an atmospheric density, and further integrate from an initial time to an nth of said predetermined integration step size, to obtain a semimajor axis and an eccentricity of an orbit in which said space object is located at the nth of said predetermined integration step sizes, wherein said predetermined integration step size is a ratio of the semimajor axis step size to the semimajor axis change rate.

10. An apparatus for obtaining a remaining track life of a near earth track space object, comprising a processor, a memory and a computer program stored on the memory, characterized in that the computer program when processed by the processor implements the steps of the method of any one of claims 1-5.

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