Decoupling Method and Device for Atmospheric Drag and Maneuvering Force of Low-Orbit Continuous Maneuvering Targets

Abstract

The present invention belongs to the technical field of aerospace data decoupling processing. Specifically, it relates to a method and device for decoupling atmospheric drag and maneuvering force of a low-orbit continuously maneuvering target, including: Step 1) Select a certain low-orbit continuously maneuvering target and obtain continuous orbit tracking and monitoring data under different space environments; Step 2) Divide the continuous orbit tracking data in each space environment into n segments equally to obtain n segments of data; Step 3) Construct an orbit dynamics model of the maneuvering target; Step 4) For each equally divided segment of data, deduct external factors and only retain the factors of maneuvering force and atmospheric damping force to obtain the change in the semi-major axis of the orbit and the change in the argument of latitude of each segment of data; Step 5) Construct an optimization objective function for the semi-major axis of the orbit and an optimization objective function for the argument of latitude; Use the optimal estimation method to obtain the component of the maneuvering force along the orbit direction and the radial component of the maneuvering force for each segment of data; and integrate them respectively.

Description

Technical Field

[0001] The present invention belongs to the technical field of aerospace data decoupling processing, and specifically relates to a method and device for decoupling atmospheric drag and maneuvering force of a continuously maneuvering target in a low orbit. Background Art

[0002] Ion propulsion technology has the characteristics of high specific impulse, high efficiency and small thrust. In recent years, this ion propulsion technology has been widely used in aspects such as spacecraft attitude control, position keeping and orbit maneuvering. At the same time, in applications such as small-thrust continuous maneuvering orbit raising of constellation targets, the propulsion force is deeply coupled with the atmospheric drag of the target, making it difficult to accurately model the components, especially for the target situation, which brings difficulties to the orbit determination, prediction and cataloging work of space targets.

[0003] At present, since the application of electric propulsion engines in spacecraft is in the development stage, the decoupling of electric thrust and atmospheric drag force has not been considered in the existing orbit determination and prediction processing, resulting in problems such as large orbit determination prediction errors. Summary of the Invention

[0004] In order to solve the problems existing in the prior art, the present invention proposes a method for decoupling atmospheric drag and maneuvering force of a continuously maneuvering target in a low orbit. Based on the different orbital dynamics characteristics of atmospheric drag and continuous small thrust, the orbit measurement data for a certain period is selected, and by using a combined orbital change model and differential correction method, the two acting forces of the target atmospheric drag and maneuvering force can be effectively decoupled and accurately modeled respectively, so as to meet the requirements of orbit determination, prediction and cataloging work of space targets.

[0005] The present invention provides a method for decoupling atmospheric drag and maneuvering force of a continuously maneuvering target in a low orbit, and the method includes:

[0006] Step 1) Select a continuously maneuvering target in a low orbit and obtain continuous orbit tracking and monitoring data under different space environments; wherein, the continuously maneuvering target is a continuously maneuvering non-cooperative target;

[0007] Step 2) Divide the continuous orbit tracking data in each space environment into n segments equally to obtain n segments of data; wherein, n≥5, and the duration of each segment of data is not less than 12h;

[0008] Step 3) Construct an orbital dynamics model of the maneuvering target;

[0009] Step 4) According to the constructed orbital dynamics model, for each equally divided segment of data, deduct external factors and only retain the factors of maneuvering force and atmospheric drag force, so as to obtain the change amount of the semi-major axis of the orbit and the change amount of the argument of latitude of each segment of data;

[0010] Step 5) According to the change in the semi-major axis of the orbit for each segment of data, construct an optimization objective function for the semi-major axis of the orbit. According to the change in the argument of the ascending node for each segment of data, construct an optimization objective function for the argument of the ascending node; adopt the optimal estimation method to obtain the component of the maneuvering force along the orbit direction and the component of the maneuvering force in the radial direction for each segment of data; and perform integration separately to achieve the decoupling of the maneuvering force and the atmospheric drag.

[0011] As one of the improvements to the above technical solution, step 4) specifically includes:

[0012] Construct the orbital dynamics model of the maneuvering target;

[0013]

[0014] Among them, is the central gravitational force of the earth on the satellite;

[0015] is the perturbing force of other celestial bodies on the earth except the earth; including the solar and lunar gravitational forces and the planetary gravitational forces;

[0016] is the perturbing force of the non-spherical part of the earth on the satellite, denoted as the high-order gravitational force;

[0017] is the gravitational force of the earth on the satellite caused by the earth's tidal factors, denoted as the tidal force;

[0018] is the force of the relativistic effect on the satellite motion;

[0019] is the radiation pressure of the sun on the satellite, denoted as the light pressure;

[0020] is the pressure of the earth's infrared radiation and reflected sunlight on the satellite;

[0021] is the damping force of the atmosphere on the satellite;

[0022] is the orbital maneuvering force;

[0023] According to the constructed orbital dynamics model, solve for each equally divided segment of data,

[0024]

[0025] After deducting the external factors including and only retain the maneuvering force and the atmospheric damping force factors, and then organize; obtain the change in the semi-major axis of the orbit for the i-th segment of data as Δai , and the variation Δu of the ascending intersection angle distance of the i-th segment of data i ;

[0026] Δa i = Δa Ci + Δa Di

[0027] Δu Ci = Δω Ci + Δf i

[0028] where, Δa Ci is the variation of the component of the motive force of the i-th segment of data in the orbital direction having a long-term variation effect on the semi-major axis;

[0029]

[0030] where, F ti is the component of the motive force of the i-th segment of data in the orbital direction; e is the average value of the orbital eccentricity in the i-th segment; n is the average value of the orbital angular rate of change in the i-th segment; t1 and t2 are the start time and end time of the i-th segment respectively;

[0031] Δa Di is the variation of the atmospheric damping force of the i-th segment of data having a long-term variation effect on the semi-major axis; integrating the rate of change of the orbital semi-major axis drag under the action of the atmospheric damping force F to obtain the part of the change in the semi-major axis caused by atmospheric damping in the time period from t1 to t2

[0032]

[0033] where,

[0034] Integrating the above formula gives

[0035]

[0036] Denote the change in the orbital semi-major axis of the i-th segment as: Δa Di After arrangement, the following expression is obtained:

[0037]

[0038] where, ρ is the atmospheric density of the i-th segment of the real-time orbit; v is the real-time target velocity of the i-th segment; BC is the ballistic coefficient; a is the semi-major axis of the i-th segment of the real-time orbit; v r is the velocity vector of the i-th segment of the real-time target relative to the atmosphere; μ is the gravitational constant of the earth;

[0039]

[0040] Among them, F si is the component of the maneuvering force of the i-th segment of data in the radial direction;

[0041] Among them, Δf i is the change in the true anomaly of the i-th segment of data;

[0042] Repeat the above process to obtain the change in the semi-major axis of the orbit and the change in the argument of latitude for each segment of data.

[0043] As one of the improvements to the above technical solution, step 5) specifically includes:

[0044] Construct an optimization objective function f(x) for the semi-major axis of the orbit based on the change in the semi-major axis of the orbit for each segment of data:

[0045]

[0046] Among them, Δa i is the change in the semi-major axis of the orbit of the i-th segment of data, which is a calculated value; Δa′ i is the measured value of the change in the semi-major axis of the orbit of the i-th segment of data;

[0047] Construct an optimization objective function f(x) for the argument of latitude based on the change in the argument of latitude for each segment of data 1 ;

[0048]

[0049] Among them, Δu i is the change in the argument of latitude of the i-th segment of data, which is a calculated value; Δu′ i is the measured value of the change in the argument of latitude of the i-th segment of data;

[0050] According to the constructed optimization objective function f(x) for the semi-major axis of the orbit, use the least squares method to take the partial derivative of F t :

[0051]

[0052] Among them, Δt = (t2 - t1);

[0053] Use an iterative method to continuously iterate until convergence is reached to obtain the component F of the maneuvering force of the i-th segment of data along the orbit direction t ;

[0054] According to the constructed optimization objective function f(x) for the argument of latitude 1 , use the least squares method to take the partial derivative of F s :

[0055]

[0056] An iterative method is adopted to continuously iterate until convergence is achieved, and the radial component F of the maneuvering force of the i-th segment of data is obtained. s .

[0057] The present invention also provides a decoupling device for atmospheric drag and maneuvering force of a low-orbit continuously maneuvering target, and the device includes:

[0058] A data acquisition module, which is used to select a certain low-orbit continuously maneuvering target and acquire continuous orbit tracking and monitoring data under different space environments;

[0059] A segmentation module, which is used to equally divide the continuous orbit tracking data under each space environment into n segments to obtain n segments of data; where n≥5, and the duration of each segment of data is not less than 12h;

[0060] A model construction module, which is used to construct an orbit dynamics model of the maneuvering target; according to the constructed orbit dynamics model, for each equally divided segment of data, external influencing factors are deducted, and only the maneuvering force and atmospheric drag force factors are retained, so as to obtain the change amount of the orbit semi-major axis and the change amount of the argument of latitude of each segment of data; and

[0061] A decoupling module, which is used to construct an orbit semi-major axis optimization objective function according to the change amount of the orbit semi-major axis of each segment of data, and construct an argument of latitude optimization objective function according to the change amount of the argument of latitude of each segment of data; an optimal estimation method is adopted to obtain the component of the maneuvering force in the orbital direction and the radial component of the maneuvering force of each segment of data; and integration is carried out respectively to realize the decoupling of the maneuvering force and the atmospheric drag.

[0062] As an improvement of the above technical solution, the model construction module includes: a model construction unit and a data processing unit;

[0063] The model construction unit is used to construct an orbit dynamics model of the maneuvering target;

[0064] The data processing unit is used to deduct external influencing factors from each equally divided segment of data according to the constructed orbit dynamics model, and only retain the maneuvering force and atmospheric drag force factors, so as to obtain the change amount of the orbit semi-major axis and the change amount of the argument of latitude of each segment of data.

[0065] As an improvement of the above technical solution, the specific process of the data processing unit includes:

[0066] Construct an orbit dynamics model of the maneuvering target;

[0067]

[0068] Among them, is the total external force acting on the unit mass of the space satellite; is the central gravitational force of the Earth on the satellite;

[0069] is the disturbing force of other celestial bodies on the Earth except the Earth; including the gravitational forces of the sun, moon and planets;

[0070] is the disturbing force of the non-spherical part of the Earth on the satellite, denoted as the high-order gravitational force;

[0071] is the gravitational force of the Earth on the satellite caused by the Earth's tidal factors, denoted as the tidal force;

[0072] is the force of the relativistic effect on the satellite motion;

[0073] is the radiation pressure of the sun on the satellite, denoted as the light pressure;

[0074] is the pressure of the Earth's infrared radiation and reflected sunlight on the satellite;

[0075] is the damping force of the atmosphere on the satellite;

[0076] is the orbital vehicle power;

[0077] According to the constructed orbital dynamics model, each segment of the equally divided data is solved.

[0078]

[0079] After deducting the external factors including and , only the vehicle power and the atmospheric damping force factors are retained and sorted out; the change in the semi-major axis of the orbit for the i-th segment of data is Δa i , and the change in the right ascension of the ascending node for the i-th segment of data is Δu i ;

[0080] Δa i = Δa Ci + Δa Di

[0081] Δu Ci = Δω Ci + Δf i

[0082] Among them, Δa Ci is the change in the long-term variation effect of the vehicle power component along the orbit direction on the semi-major axis for the i-th segment of data;

[0083]

[0084] Among them, F ti is the component of the maneuvering force of the i-th segment of data in the orbital direction; e is the average value of the orbital eccentricity in the i-th segment; n is the average value of the orbital angular rate of change in the i-th segment; t1 and t2 are the starting time and ending time of the i-th segment respectively;

[0085] Δa Di is the change amount of the long-term change effect of the atmospheric damping force on the semi-major axis of the i-th segment of data; for the atmospheric damping force F drag under the action, the change rate of the orbital semi-major axis is integrated to obtain the part of the change caused by atmospheric damping of the semi-major axis in the time period from t1 to t2

[0086]

[0087] Among them,

[0088] Integrating the above formula can obtain

[0089]

[0090] The change amount of the orbital semi-major axis of the i-th segment is denoted as: Δa Di After sorting, the following expression is obtained:

[0091]

[0092] Among them, ρ is the atmospheric density of the real-time orbit of the i-th segment; v is the real-time target velocity of the i-th segment; BC is the ballistic coefficient; a is the real-time orbit semi-major axis of the i-th segment; v r is the velocity vector of the real-time target relative to the atmosphere of the i-th segment; μ is the gravitational constant of the earth;

[0093]

[0094] Among them, F si is the component of the maneuvering force of the i-th segment of data in the radial direction;

[0095] Among them, Δf i is the change amount of the true anomaly of the i-th segment of data;

[0096] Repeat the above process to obtain the change amount of the orbital semi-major axis and the change amount of the argument of latitude of each segment of data.

[0097] As one of the improvements of the above technical solution, the specific process of the decoupling module includes:

[0098] Construct an optimization objective function \(f(x)\) for the semi-major axis of the orbit based on the change in the semi-major axis of the orbit for each segment of data:

[0099]

[0100] where \(\Delta a\) i is the change in the semi-major axis of the orbit for the \(i\)-th segment of data and is the calculated value; \(\Delta a'\) i is the measured value of the change in the semi-major axis of the orbit for the \(i\)-th segment of data;

[0101] Construct an optimization objective function \(f(x)\) for the argument of the ascending node based on the change in the argument of the ascending node for each segment of data 1 ;

[0102]

[0103] where \(\Delta u\) i is the change in the argument of the ascending node for the \(i\)-th segment of data and is the calculated value; \(\Delta u'\) i is the measured value of the change in the argument of the ascending node for the \(i\)-th segment of data;

[0104] According to the constructed optimization objective function \(f(x)\) for the semi-major axis of the orbit, use the least squares method to take the partial derivative of \(F\) t with respect to:

[0105]

[0106] where \(\Delta t=(t2 - t1)\);

[0107] Adopt an iterative method and continuously iterate until convergence to obtain the component \(F\) of the maneuvering force along the orbit direction for the \(i\)-th segment of data t ;

[0108] According to the constructed optimization objective function \(f(x)\) for the argument of the ascending node 1 , use the least squares method to take the partial derivative of \(F\) s with respect to:

[0109]

[0110] Adopt an iterative method and continuously iterate until convergence to obtain the component \(F\) of the maneuvering force in the radial direction for the \(i\)-th segment of data s .

[0111] The beneficial effects of the present invention compared with the prior art are:

[0112] 1. The method of the present invention respectively constructs the orbital dynamic effects of the maneuvering force and the atmospheric drag, selects the equal division of continuous orbit tracking data to obtain statistical correction data. Construct an objective function to perform optimal estimation on the two components of the maneuvering force, and successfully decouple the atmospheric drag and the maneuvering force effects;

[0113] 2. Based on the orbit measurement data of non - cooperative continuously maneuvering targets, by using the different orbital dynamics characteristics and statistical theories of atmospheric drag and continuous small thrust, decouple the two acting forces of target atmospheric drag and maneuvering force, and construct an accurate orbit prediction model to meet the requirements of orbit determination, prediction and cataloging of space targets. Brief Description of the Drawings

[0114] Figure 1 It is the flowchart of the method for decoupling atmospheric drag and maneuvering force of continuously maneuvering targets in low - earth orbit of the present invention. Detailed Embodiment

[0115] The present invention will be further described in conjunction with the accompanying drawings.

[0116] As Figure 1 shown, the present invention provides a method for decoupling atmospheric drag and maneuvering force of continuously maneuvering targets in low - earth orbit, and the method includes:

[0117] Step 1) Select a certain continuously maneuvering target in low - earth orbit and obtain continuous orbit tracking and monitoring data under different space environments; wherein, the continuously maneuvering target is a continuously maneuvering non - cooperative target;

[0118] Step 2) Divide the continuous orbit tracking data under each space environment into n segments equally to obtain n segments of data; wherein, n≥5, and the duration of each segment of data is not less than 12 h;

[0119] Step 3) Construct an orbit dynamics model of the maneuvering target; wherein, the orbit dynamics model includes but is not limited to high - order gravitational field, atmospheric drag, maneuvering force, maneuvering model with unknown parameters, tides and solar - lunar gravity;

[0120] Step 4) According to the constructed orbit dynamics model, for each equally - divided segment of data, deduct the factors of high - order gravitational field, maneuvering model with unknown parameters, tides and solar - lunar gravity, and only retain the factors of maneuvering force and atmospheric drag force, to obtain the change amount of the semi - major axis of the orbit and the change amount of the argument of latitude of each segment of data.

[0121] Specifically, construct an orbit dynamics model of the maneuvering target;

[0122]

[0123] The above are the forces acting on the unit mass of the satellite, and they are respectively expressed as:

[0124] Wherein, is the total external force acting on the unit mass of the space satellite;

[0125] is the perturbing force of other celestial bodies on the earth except the earth; including solar - lunar gravity and planetary gravity;

[0126] The disturbing force of the non-spherical part of the Earth on the satellite, denoted as the high-order gravity;

[0127] The gravitational force of the Earth on the satellite caused by the Earth's tidal factors, denoted as the tidal force;

[0128] The force of the influence of the relativistic effect on the satellite motion;

[0129] The radiation pressure of the Sun on the satellite, denoted as the light pressure;

[0130] The pressure on the satellite due to the Earth's infrared radiation and reflected sunlight;

[0131] The damping force of the atmosphere on the satellite;

[0132] The orbital vehicle power;

[0133] According to the constructed orbital dynamics model, each segment of the equally divided data is solved.

[0134]

[0135] After deducting and factors, only the vehicle power and the atmospheric damping force factors are retained, and then sorted out; the change in the semi-major axis of the orbit for the i-th segment of data is Δa i , and the change in the argument of latitude for the i-th segment of data is Δu i ; Considering the small eccentricity orbit, the determination deviation of the argument of perigee ω is relatively large. Here, the argument of latitude u = ω + f is used, where ω and f represent the argument of perigee and the true anomaly in the orbital elements respectively; since the argument of latitude is not a direct orbital six-element, it needs to be obtained indirectly through ω and f.

[0136] Δa i = Δa Ci + Δa Di

[0137] Δu Ci = Δω Ci + Δf i

[0138] Among them, Δa Ci is the change in the long-term variation effect of the vehicle power component along the orbit direction on the semi-major axis for the i-th segment of data.

[0139]

[0140] Among them, F ti is the component of the maneuvering force of the i-th segment of data in the orbital direction; e is the average value of the orbital eccentricity in the i-th segment; n is the average value of the orbital angular rate of change in the i-th segment; t1 and t2 are the start time and end time of the i-th segment respectively;

[0141] Δa Di is the change amount of the long-term variation effect of the atmospheric damping force on the semi-major axis for the i-th segment of data; for the atmospheric damping force F drag under the action, the change rate of the orbital semi-major axis is integrated to obtain the part of the change in the semi-major axis caused by atmospheric damping in the time period from t1 to t2

[0142]

[0143] Among them,

[0144] Integrating the above formula can obtain

[0145]

[0146] After arrangement, the following expression is obtained:

[0147]

[0148] Among them, ρ is the atmospheric density of the real-time orbit of the i-th segment; v is the real-time target speed of the i-th segment; BC is the ballistic coefficient; a is the semi-major axis of the real-time orbit of the i-th segment; v r is the velocity vector of the real-time target relative to the atmosphere in the i-th segment; μ is the gravitational constant of the earth;

[0149]

[0150] Among them, F si is the component of the maneuvering force of the i-th segment of data in the radial direction;

[0151] Among them, Δf i is the change amount of the true anomaly of the i-th segment of data;

[0152] Repeat the above process to obtain the change amount of the orbital semi-major axis and the change amount of the argument of latitude for each segment of data. Step 5) Construct an orbital optimization objective function according to the change amount of the orbital semi-major axis of each segment of data, and construct an argument-of-latitude optimization objective function according to the change amount of the argument of latitude of each segment of data; adopt the optimal estimation method to obtain the component of the maneuvering force in the along-orbital direction and the component of the maneuvering force in the radial direction;

[0153] Specifically, according to the change amount of the orbital semi-major axis of each segment of data, construct an orbital semi-major axis optimization objective function f(x):

[0154]

[0155] Among them, Δa i is the change in the semi-major axis of the orbit for the i-th segment of data, which is a calculated value; Δa′ i is the measured value of the change in the semi-major axis of the orbit for the i-th segment of data;

[0156] According to the change in the argument of latitude for each segment of data, construct the optimization objective function f(x) of the argument of latitude 1 ;

[0157]

[0158] Among them, Δu i is the change in the argument of latitude for the i-th segment of data, which is a calculated value; Δu′ i is the measured value of the change in the argument of latitude for the i-th segment of data;

[0159] Adopt the optimal estimation method to obtain the component of the maneuvering force along the orbit direction and the component of the maneuvering force in the radial direction for the i-th segment of data;

[0160] Among them, the optimal estimation method can be methods such as least squares or Kalman filtering. Preferably, the least squares method is taken as an example for illustration.

[0161] According to the constructed optimization objective function f(x) of the semi-major axis of the orbit, adopt the least squares method to take the partial derivative of F t :

[0162]

[0163] Among them, Δt = (t2 - t1);

[0164] Because the partial derivative function of the objective with respect to the estimator is non-linear, generally an iterative method is adopted, and continuous iteration is carried out until convergence is achieved to obtain the component of the maneuvering force along the orbit direction F t ;

[0165] According to the constructed optimization objective function f(x) of the argument of latitude 1 , adopt the least squares method to take the partial derivative of F s :

[0166]

[0167] Adopt the iterative method, and continuously iterate until convergence is achieved to obtain the component of the maneuvering force in the radial direction F s .

[0168] Step 6) Repeat Step 4) - Step 5) to obtain the component of the maneuvering force of each data segment in the orbital direction and the component of the maneuvering force in the radial direction, and integrate them separately to achieve the decoupling of the maneuvering force and the atmospheric drag.

[0169] The present invention also provides a device for decoupling the atmospheric drag and the maneuvering force of a low-orbit continuously maneuvering target, which includes: a data acquisition module, a segmentation module, a model construction module, and a decoupling module;

[0170] The data acquisition module is used to select a certain low-orbit continuously maneuvering target and obtain continuous orbital tracking and monitoring data under different space environments; wherein, the continuously maneuvering target is a continuously maneuvering non-cooperative target;

[0171] The segmentation module is used to equally divide the continuous orbital tracking data under each space environment into n segments to obtain n segments of data; wherein, n≥5, and the duration of each segment of data is not less than 12h;

[0172] The model construction module is used to construct the orbital dynamics model of the maneuvering target; according to the constructed orbital dynamics model, for each equally divided segment of data, external influencing factors are deducted, and only the factors of the maneuvering force and the atmospheric drag force are retained to obtain the change amount of the semi-major axis of the orbit and the change amount of the argument of latitude of each segment of data;

[0173] Specifically, the model construction module includes: a model construction unit and a data processing unit;

[0174] The model construction unit is used to construct the orbital dynamics model of the maneuvering target;

[0175] The data processing unit is used to, according to the constructed orbital dynamics model, for each equally divided segment of data, deduct external influencing factors and only retain the factors of the maneuvering force and the atmospheric drag force to obtain the change amount of the semi-major axis of the orbit and the change amount of the argument of latitude of each segment of data.

[0176] Specifically, the specific process of the data processing unit includes:

[0177] Construct the orbital dynamics model of the maneuvering target;

[0178]

[0179] wherein, is the total external force acting on the unit mass of the space satellite;

[0180] is the central gravitational force of the earth on the satellite;

[0181] is the perturbing force of other celestial bodies on the earth except the earth; including the solar and lunar gravitational forces and the planetary gravitational forces;

[0182] The disturbing force of the non-spherical part of the Earth on the satellite, denoted as the high-order gravity;

[0183] The gravitational force of the Earth on the satellite caused by the Earth's tidal factors, denoted as the tidal force;

[0184] The force of the relativistic effect on the satellite's motion;

[0185] The radiation pressure of the Sun on the satellite, denoted as the light pressure;

[0186] The pressure on the satellite due to the Earth's infrared radiation and reflected sunlight;

[0187] The damping force of the atmosphere on the satellite;

[0188] The orbital vehicle power;

[0189] According to the constructed orbital dynamics model, each segment of the equally divided data is solved,

[0190]

[0191] After deducting the external factors including and , only the vehicle power and the atmospheric damping force factors are retained, and then sorted out; the change in the semi-major axis of the orbit for the i-th segment of data is Δa i , and the change in the argument of latitude for the i-th segment of data is Δu i ;

[0192] Δa i = Δa Ci + Δa Di

[0193] Δu Ci = Δω Ci + Δf i

[0194] where, Δa Ci is the change in the semi-major axis due to the long-term change effect of the vehicle power component along the orbit direction for the i-th segment of data;

[0195]

[0196] where, F tiis the component of the maneuvering force of the i-th segment of data in the orbital direction; e is the average value of the orbital eccentricity in the i-th segment; n is the average value of the orbital angular rate of change in the i-th segment; t1 and t2 are the start time and end time of the i-th segment respectively;

[0197] Δa Di is the change amount of the long-term variation effect of the atmospheric damping force on the semi-major axis for the i-th segment of data; for the atmospheric damping force F drag under the action, the change rate of the orbital semi-major axis Integrate to obtain the part of the change in the semi-major axis caused by atmospheric damping in the time period from t1 to t2

[0198]

[0199] Among them,

[0200] Integrating the above formula can obtain

[0201]

[0202] Denote the change amount of the orbital semi-major axis of the i-th segment as: Δa Di After arrangement, the following expression is obtained:

[0203]

[0204] Among them, ρ is the atmospheric density of the real-time orbit of the i-th segment; v is the real-time target speed of the i-th segment; BC is the ballistic coefficient; a is the real-time orbital semi-major axis of the i-th segment; v r is the velocity vector of the real-time target relative to the atmosphere in the i-th segment; μ is the gravitational constant of the earth;

[0205]

[0206] Among them, F si is the component of the maneuvering force of the i-th segment of data in the radial direction;

[0207] Among them, Δf i is the change amount of the true anomaly of the i-th segment of data;

[0208] Repeat the above process to obtain the change amount of the orbital semi-major axis and the change amount of the argument of latitude for each segment of data.

[0209] The decoupling module is used to construct an optimization objective function for the orbital semi-major axis according to the change amount of the orbital semi-major axis of each segment of data, and construct an optimization objective function for the argument of latitude according to the change amount of the argument of latitude of each segment of data; adopt the optimal estimation method to obtain the component of the maneuvering force along the orbital direction and the component of the maneuvering force in the radial direction of each segment of data; and perform integration respectively to realize the decoupling of the maneuvering force and the atmospheric damping.

[0210] Specifically, according to the change in the semi-major axis of the orbit for each segment of data, an optimization objective function f(x) for the semi-major axis of the orbit is constructed:

[0211]

[0212] where Δa i is the change in the semi-major axis of the orbit for the i-th segment of data and is a calculated value; Δa' i is the measured value of the change in the semi-major axis of the orbit for the i-th segment of data;

[0213] According to the change in the argument of latitude for each segment of data, an optimization objective function f(x) for the argument of latitude is constructed 1 ;

[0214]

[0215] where Δu i is the change in the argument of latitude for the i-th segment of data and is a calculated value; Δu' i is the measured value of the change in the argument of latitude for the i-th segment of data;

[0216] According to the constructed optimization objective function f(x) for the semi-major axis of the orbit, using the least squares method, the partial derivative of F t is obtained:

[0217]

[0218] where Δt = (t2 - t1);

[0219] An iterative method is adopted to continuously perform iterations until convergence is achieved, and the component of the maneuvering force of the i-th segment of data in the orbital direction F t is obtained;

[0220] According to the constructed optimization objective function f(x) for the argument of latitude 1 , using the least squares method, the partial derivative of F s is obtained:

[0221]

[0222] An iterative method is adopted to continuously perform iterations until convergence is achieved, and the component of the maneuvering force of the i-th segment of data in the radial direction F s is obtained.

[0223] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and not to limit them. Although the present invention has been described in detail with reference to the embodiments, those of ordinary skill in the art should understand that any modification or equivalent replacement of the technical solutions of the present invention does not depart from the spirit and scope of the technical solutions of the present invention, and they should all be covered within the scope of the claims of the present invention.

Claims

1. A method for decoupling atmospheric drag and maneuvering force of a continuously maneuvering target in a low orbit, the method comprising: Step 1) Select a continuously maneuvering target in a low orbit and obtain continuous orbit tracking and monitoring data under different space environments; Step 2) Divide the continuous orbit tracking data under each space environment into n segments equally to obtain n segments of data; where n≥5 and the duration of each segment of data is not less than 12 h; Step 3) Construct an orbit dynamics model of the maneuvering target; Step 4) According to the constructed orbit dynamics model, for each equally divided segment of data, deduct external factors and only retain the maneuvering force and atmospheric damping force factors to obtain the change amount of the semi-major axis of the orbit and the change amount of the argument of latitude for each segment of data; Step 5) According to the change amount of the semi-major axis of the orbit for each segment of data, construct an optimization objective function for the semi-major axis of the orbit, and according to the change amount of the argument of latitude for each segment of data, construct an optimization objective function for the argument of latitude; adopt the optimal estimation method to obtain the component of the maneuvering force along the orbit direction and the radial component of the maneuvering force for each segment of data; and integrate them respectively to achieve the decoupling of the maneuvering force and atmospheric damping.

2. The method for decoupling atmospheric drag and maneuvering force of a low-orbit continuously maneuvering target according to claim 1, wherein The specific content of the said Step 4) includes: Construct an orbit dynamics model of the maneuvering target; Among them, is the total external force acting on the unit mass of the space satellite; is the central gravitational force of the Earth on the satellite; The disturbing force exerted on the Earth by celestial bodies other than the Earth, including the gravitational forces of the sun, moon and planets; The disturbing force of the non-spherical part of the Earth on the satellite is denoted as the high-order gravity; The gravitational force exerted by the Earth on the satellite due to the Earth's tidal factors is denoted as the tidal force; the force due to the relativistic effect on the satellite motion; is the solar radiation pressure on the satellite, denoted as the light pressure; The pressure on the satellite due to the infrared radiation of the earth and the reflected sunlight; is the damping force of the atmosphere on the satellite; For the power of the orbiter; According to the constructed orbit dynamics model, solve for each equally divided segment of data, After deducting external factors including and , only the maneuvering force and the atmospheric damping force factors are retained, and then sorted out; the change in the semi-major axis of the orbit for the i-th segment of data is Δa i , and the change in the argument of latitude for the i-th segment of data is Δu i ; Δa i = Δa Ci + Δa Di Δu Ci = Δω Ci + Δf i where, Δa Ci is the variation amount of the maneuvering force of the i-th segment of data in the component along the orbital direction having a secular variation effect on the semi-major axis; Among them, F ti is the component of the maneuvering force of the i-th segment of data in the orbital direction; e is the average value of the orbital eccentricity in the i-th segment; n is the average value of the orbital angular rate of change in the i-th segment; t1 and t2 are the start time and end time of the i-th segment respectively; Δa Di is the change in the semi-major axis due to the long-term variation effect of the atmospheric drag force on the i-th segment of data; for the atmospheric drag force F drag under the action, the change rate of the orbital semi-major axis Integrate to obtain the part of the change in the semi-major axis caused by atmospheric drag in the time period from t1 to t2 Among them, Integrating the above formula can obtain Denote the change in the semi-major axis of the i-th segment as: Δa Di After arrangement, the following expression is obtained: where ρ is the real-time orbital atmospheric density of the i-th segment; v is the real-time target velocity of the i-th segment; BC is the ballistic coefficient; a is the semi-major axis of the real-time orbit of the i-th segment; v r is the velocity vector of the real-time target relative to the atmosphere; μ is the gravitational constant of the Earth; Among them, F si is the component of the motive force of the i-th segment of data in the radial direction; where Δf i is the change in the true anomaly of the data for the i-th segment; Repeat the above process to obtain the change amount of the semi-major axis of the orbit and the change amount of the argument of latitude for each segment of data.

3. The method for decoupling atmospheric drag and maneuvering force of a low-orbit continuously maneuvering target according to claim 1, wherein The specific content of the said Step 5) includes: According to the change amount of the semi-major axis of the orbit for each segment of data, construct an optimization objective function f(x) for the semi-major axis of the orbit: Among them, Δa i is the change in the semi-major axis of the orbit for the i-th segment of data, which is a calculated value; Δa' i is the measured value of the change in the semi-major axis of the orbit for the i-th segment of data; Construct the ascending node right ascension optimization objective function f(x) according to the change in the ascending node right ascension angle distance for each data segment 1 ; where, Δu i is the change in the ascending node right ascension for the i-th data segment, which is a calculated value; Δu' i is the measured value of the change in the ascending node right ascension for the i-th data segment; Optimize the objective function f(x) according to the constructed semi-major axis of the orbit, and use the least squares method to find the partial derivative of F t with respect to: Where Δt=(t2 - t1); Using an iterative method, continuously perform iterations until convergence is achieved, and obtain the component F of the maneuvering force of the i-th segment of data in the orbital direction t ; According to the constructed ascending crossing angle distance optimization objective function f(x) 1 , using the least squares method, for F s Find the partial derivative: An iterative method is adopted to continuously perform iterations until convergence is achieved, and the radial component F of the maneuvering force of the i-th segment of data is obtained s .

4. A device for decoupling atmospheric drag and maneuvering force of a continuously maneuvering target in low orbit, characterized in that, The device includes: A data acquisition module, which is used to select a continuously maneuvering target in a low orbit and obtain continuous orbit tracking and monitoring data under different space environments; A segmentation module, which is used to divide the continuous orbit tracking data under each space environment into n segments equally to obtain n segments of data; where n≥5 and the duration of each segment of data is not less than 12 h; A model construction module, which is used to construct an orbit dynamics model of the maneuvering target; according to the constructed orbit dynamics model, for each equally divided segment of data, deduct external influencing factors and only retain the maneuvering force and atmospheric damping force factors to obtain the change amount of the semi-major axis of the orbit and the change amount of the argument of latitude for each segment of data; and A decoupling module, which is used to construct an optimization objective function for the semi-major axis of the orbit according to the change amount of the semi-major axis of the orbit for each segment of data, and construct an optimization objective function for the argument of latitude according to the change amount of the argument of latitude for each segment of data; adopt the optimal estimation method to obtain the component of the maneuvering force along the orbit direction and the radial component of the maneuvering force for each segment of data; and integrate them respectively to achieve the decoupling of the maneuvering force and atmospheric damping.

5. The atmospheric drag and maneuvering force decoupling device for a continuously maneuvering low-orbit target according to claim 4, characterized in that The said model construction module includes: a model construction unit and a data processing unit; The said model construction unit is used to construct an orbit dynamics model of the maneuvering target; The said data processing unit is used to deduct external influencing factors for each equally divided segment of data according to the constructed orbit dynamics model and only retain the maneuvering force and atmospheric damping force factors to obtain the change amount of the semi-major axis of the orbit and the change amount of the argument of latitude for each segment of data.

6. The atmospheric drag and maneuvering force decoupling device for a low-orbit continuously maneuvering target according to claim 5, characterized in that, The specific process of the said data processing unit includes: Construct an orbit dynamics model of the maneuvering target; wherein, is the total external force acting on the unit mass of the space satellite; is the central gravitational force of the Earth on the satellite; The disturbing force exerted on the Earth by celestial bodies other than the Earth; including the gravitational forces of the sun, moon and planets; The disturbing force exerted by the non-spherical part of the Earth on the satellite, denoted as the high-order gravity; The gravitational force exerted by the Earth on the satellite due to the Earth's tidal factors is denoted as the tidal force; the force due to the relativistic effect on the satellite motion; is the radiation pressure of the sun on the satellite, denoted as the optical pressure; For the pressure of the Earth's infrared radiation and reflected sunlight on the satellite; is the damping force of the atmosphere on the satellite; For the power of the orbiter; According to the constructed orbital dynamics model, solve each segment of equally divided data. After deducting external factors including and , only the maneuvering force and the atmospheric damping force are retained, and after organizing; the change in the semi-major axis of the orbit for the i-th segment of data is Δa i , and the change in the argument of latitude for the i-th segment of data is Δu i ; Δa i = Δa Ci + Δa Di Δu Ci = Δω Ci + Δf i where Δa Ci is the change in the component of the maneuvering force of the i-th segment of data in the orbital direction that has a secular variation effect on the semi-major axis; Among them, F ti is the component of the maneuvering force of the i-th segment of data in the orbital direction; e is the average value of the orbital eccentricity in the i-th segment; n is the average value of the orbital angular rate of change in the i-th segment; t1 and t2 are the start time and end time of the i-th segment, respectively; Δa Di is the change amount of the long-term variation effect of the atmospheric damping force on the semi-major axis for the i-th segment of data; for the atmospheric damping force F drag under the action, the change rate of the orbital semi-major axis Integrate to obtain the part of the change caused by atmospheric damping of the semi-major axis in the time period from t1 to t2 Among them, Integrating the above equation gives Denote the change in the semi-major axis of the i-th segment as: Δa Di After arrangement, the following expression is obtained: where ρ is the real-time orbital atmospheric density of the i-th segment; v is the real-time target velocity of the i-th segment; BC is the ballistic coefficient; a is the semi-major axis of the real-time orbit of the i-th segment; v r is the velocity vector of the real-time target relative to the atmosphere of the i-th segment; μ is the gravitational constant of the Earth; Among them, F si is the component of the maneuvering force of the i-th segment of data in the radial direction; where, Δf i is the change in the true anomaly of the data for the i-th segment; Repeat the above process to obtain the change in the semi-major axis of the orbit and the change in the argument of the ascending node for each segment of data.

7. The atmospheric drag and maneuvering force decoupling device for a low-orbit continuously maneuvering target according to claim 4, wherein The specific process of the decoupling module includes: According to the change in the semi-major axis of the orbit for each segment of data, construct an optimization objective function f(x) for the semi-major axis of the orbit: where, Δa i is the change in the semi-major axis of the orbit for the i-th segment of data, which is a calculated value; Δa′ i is the measured value of the change in the semi-major axis of the orbit for the i-th segment of data; Construct the ascending node right ascension optimization objective function \(f(x)\) according to the change amount of the ascending node right ascension for each data segment 1 ; where, Δu i is the change in the ascending node right ascension for the i-th segment of data, which is a calculated value; Δu' i is the measured value of the change in the ascending node right ascension for the i-th segment of data; According to the constructed objective function \(f(x)\) for optimizing the semi-major axis of the orbit, the least squares method is used to find the partial derivative of \(F\) with respect to t : where Δt = (t2 - t1); An iterative method is adopted to continuously perform iterations until convergence is achieved, and the component of the maneuvering force of the i-th segment of data in the orbital direction F is obtained t ; According to the constructed optimization objective function \(f(x)\) of the ascending intersection angle distance 1 , the least squares method is used to find the partial derivative of \(F\) with respect to s : Using an iterative method, continuous iteration is carried out until convergence is achieved, and the radial component F of the maneuvering force of the i-th segment of data is obtained s .

Images