Spacecraft mass estimation method and device, storage medium and electronic equipment

Abstract

The application discloses a spacecraft mass estimation method and device, a storage medium and an electronic device. The method comprises the following steps: obtaining a measured comprehensive drag coefficient of a target orbit in which a spacecraft is located and an atmospheric drag coefficient initial value of the target orbit, wherein the measured comprehensive drag coefficient is an actually measured comprehensive drag coefficient; determining a nominal comprehensive drag coefficient of the target orbit based on the atmospheric drag coefficient initial value and a preset mass of the spacecraft, wherein the nominal comprehensive drag coefficient is a theoretically calculated comprehensive drag coefficient; determining an atmospheric drag coefficient target value based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; and estimating a target mass of the spacecraft based on the atmospheric drag coefficient initial value, the preset mass and the atmospheric drag coefficient target value. The application solves the technical problem of low efficiency in estimating the mass of the spacecraft in the related art.

Description

Technical Field

[0001] This invention relates to the aerospace field, and more specifically, to a method, apparatus, storage medium, and electronic device for estimating the mass of a spacecraft. Background Technology

[0002] Spacecraft mass is one of the fundamental physical properties of a spacecraft and a crucial physical parameter for its mission. Precise mass measurement is performed before launch. Typically, after a spacecraft enters orbit, its mass changes due to fuel consumption and the release of test vehicles. Fuel consumption can be measured using telemetry parameters, and the mass of released test vehicles is also precisely measured before launch. Therefore, the changes in spacecraft mass under these circumstances can be accurately calculated.

[0003] In some cases, near-Earth spacecraft frequently employ attitude control jetting based on mission requirements, resulting in significant changes in mass. These changes do not refer to the mass of fuel consumed by attitude control jetting, but rather to the mass exchange caused by actions such as refueling and cargo transport after rendezvous and docking. When the mass changes after spacecraft separation and evacuation cannot be accurately assessed, it has a certain impact on the control and management of subsequent flight missions, meaning that the efficiency of estimating the mass of spacecraft is low.

[0004] There is currently no effective solution to the above problems. Summary of the Invention

[0005] This invention provides a method, apparatus, storage medium, and electronic device for estimating the mass of a spacecraft, thereby addressing at least the technical problem of low efficiency in estimating the mass of spacecraft in related technologies.

[0006] According to one aspect of the present invention, a method for estimating the mass of a spacecraft is provided, comprising: obtaining a measured composite drag coefficient of a target orbit in which the spacecraft is located and an initial value of the atmospheric drag coefficient of the target orbit, wherein the measured composite drag coefficient is a composite drag coefficient obtained by actual measurement; determining a nominal composite drag coefficient of the target orbit based on the initial value of the atmospheric drag coefficient and a preset mass of the spacecraft, wherein the nominal composite drag coefficient is a composite drag coefficient obtained by theoretical calculation; determining a target value of the atmospheric drag coefficient based on the measured composite drag coefficient and the nominal composite drag coefficient; and estimating the target mass of the spacecraft based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient.

[0007] Optionally, estimating the target mass of the spacecraft based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient includes: determining a first ratio between the preset mass and the initial value of the atmospheric drag coefficient; and determining a first product of the first ratio and the target value of the atmospheric drag coefficient as the target mass.

[0008] Optionally, the target value of the atmospheric drag coefficient is determined based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient, including: determining the coefficient error of the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; and determining the target value of the atmospheric drag coefficient based on the coefficient error and the initial value of the atmospheric drag coefficient.

[0009] Optionally, determining the target value of the atmospheric drag coefficient based on the coefficient error and the initial value of the atmospheric drag coefficient includes: if the coefficient error is greater than or equal to a preset threshold, determining the target value of the atmospheric drag coefficient based on the first sum of the initial value of the atmospheric drag coefficient and the preset threshold; if the coefficient error is less than the preset threshold, determining the initial value of the atmospheric drag coefficient as the target value of the atmospheric drag coefficient.

[0010] Optionally, determining the target value of the atmospheric drag coefficient based on the initial value of the atmospheric drag coefficient and a first sum of a preset threshold includes: determining the first sum as the initial value of the atmospheric drag coefficient, and performing orbit determination on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft to obtain the nominal comprehensive drag coefficient of the target orbit, until the coefficient error is less than the preset threshold, and determining the initial value of the atmospheric drag coefficient as the target value of the atmospheric drag coefficient.

[0011] Optionally, the nominal comprehensive drag coefficient of the target orbit is determined based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft, including: determining the atmospheric drag acceleration of the target orbit based on the initial value of the atmospheric drag coefficient; performing orbit determination on the atmospheric drag acceleration and the preset mass to obtain the attitude control jet acceleration of the spacecraft in the target orbit; and determining the nominal comprehensive drag coefficient based on the attitude control jet acceleration, the initial value of the atmospheric drag coefficient, and the atmospheric drag acceleration.

[0012] Optionally, the nominal comprehensive drag coefficient is determined based on the attitude control jet acceleration, the initial value of the atmospheric drag coefficient, and the atmospheric drag acceleration, including: determining a second ratio of the attitude control jet acceleration and the atmospheric drag acceleration; determining a second sum of the preset value and the second ratio; and determining a second product of the second sum and the initial value of the atmospheric drag coefficient as the nominal comprehensive drag coefficient.

[0013] According to another aspect of the present invention, a spacecraft mass estimation device is also provided, comprising: an acquisition module for acquiring a measured comprehensive drag coefficient of a target orbit in which the spacecraft is located and an initial value of the atmospheric drag coefficient of the target orbit, wherein the measured comprehensive drag coefficient is a comprehensive drag coefficient obtained by actual measurement; a first determination module for determining a nominal comprehensive drag coefficient of the target orbit based on the initial value of the atmospheric drag coefficient and a preset mass of the spacecraft, wherein the nominal comprehensive drag coefficient is a comprehensive drag coefficient obtained by theoretical calculation; a second determination module for determining a target value of the atmospheric drag coefficient based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; and an estimation module for estimating the target mass of the spacecraft based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient.

[0014] According to another aspect of the present invention, a computer-readable storage medium is also provided, the computer-readable storage medium including a stored program, wherein, when the program is executed, the aforementioned spacecraft mass estimation method is executed in the processor of the device.

[0015] According to another aspect of the present invention, an electronic device is also provided, comprising one or more processors; a storage device for storing one or more programs; and when the one or more programs are executed by the one or more processors, causing the one or more processors to perform the aforementioned spacecraft mass estimation method.

[0016] In this embodiment of the invention, the measured comprehensive drag coefficient of the target orbit and the initial value of the atmospheric drag coefficient of the target orbit are obtained, wherein the measured comprehensive drag coefficient is the comprehensive drag coefficient obtained by actual measurement; the nominal comprehensive drag coefficient of the target orbit is determined based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft, wherein the nominal comprehensive drag coefficient is the comprehensive drag coefficient obtained by theoretical calculation; the target value of the atmospheric drag coefficient is determined based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; the target mass of the spacecraft is estimated based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient. This invention establishes a dynamic model for near-Earth spacecraft, compares the comprehensive drag coefficient calculated by the theoretical model with the actual calculated value, and when the error is large, iteratively approximates the actual calculated value of the comprehensive drag coefficient by using the deviation correction initial value of the atmospheric drag coefficient, thereby achieving accurate calculation of the atmospheric drag coefficient; based on the characteristics of the atmospheric drag calculated by the orbit, the parameter changes in the analytical formula are analyzed, and finally the proportional relationship between the atmospheric drag coefficient and the mass is analyzed. Using the known conditions, the actual mass of the spacecraft is finally obtained, thereby solving the technical problem of low efficiency in estimating the mass of spacecraft in related technologies. Attached Figure Description

[0017] The accompanying drawings, which are included to provide a further understanding of the invention and form part of this invention, illustrate exemplary embodiments of the invention and are used to explain the invention, but do not constitute an undue limitation of the invention. In the drawings:

[0018] Figure 1 This is a flowchart of a spacecraft mass estimation method according to an embodiment of the present invention;

[0019] Figure 2 This is a schematic diagram of the windward area of ​​an optional spacecraft for atmospheric drag, according to an embodiment of the present invention.

[0020] Figure 3 This is a flowchart of an optional method for calculating spacecraft mass using a dynamic model according to an embodiment of the present invention;

[0021] Figure 4 This is a schematic diagram of an optional spacecraft atmospheric density and atmospheric drag acceleration according to an embodiment of the present invention;

[0022] Figure 5 This is a schematic diagram of a spacecraft mass estimation device according to an embodiment of the present invention. Detailed Implementation

[0023] To enable those skilled in the art to better understand the present invention, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings of the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort should fall within the scope of protection of the present invention.

[0024] It should be noted that the terms "first," "second," etc., in the specification, claims, and accompanying drawings of this invention are used to distinguish similar objects and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged where appropriate so that the embodiments of the invention described herein can be implemented in orders other than those illustrated or described herein. Furthermore, the terms "comprising" and "having," and any variations thereof, are intended to cover a non-exclusive inclusion; for example, a process, method, system, product, or apparatus that comprises a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units not explicitly listed or inherent to such processes, methods, products, or apparatus.

[0025] Example 1

[0026] According to an embodiment of the present invention, an embodiment of a spacecraft mass estimation method is provided. It should be noted that the steps shown in the flowchart in the accompanying drawings can be executed in a computer system such as a set of computer-executable instructions. Furthermore, although a logical order is shown in the flowchart, in some cases, the steps shown or described may be executed in a different order than that shown here.

[0027] Figure 1 This is a spacecraft mass estimation method according to an embodiment of the present invention, such as... Figure 1 As shown, the method includes the following steps:

[0028] Step S102: Obtain the measured comprehensive drag coefficient of the target orbit where the spacecraft is located and the initial value of the atmospheric drag coefficient of the target orbit.

[0029] Among them, the measured comprehensive resistance coefficient is the comprehensive resistance coefficient obtained by actual measurement.

[0030] The aforementioned spacecraft can refer to any spacecraft that conducts scientific research, exploration, or manned flight in outer space. It can include various types of spacecraft such as satellites, spacecraft, space stations, and probes. In this embodiment, it mainly refers to near-Earth spacecraft, that is, on-orbit spacecraft with an orbital altitude of less than 1,000 kilometers.

[0031] The target orbit mentioned above can refer to the actual orbit of the spacecraft.

[0032] Since atmospheric drag only has a decelerating and dragging effect on the spacecraft along its track, this application can consider only the acceleration and deceleration forces mainly acting along the track in the attitude control jet calculations. This can be done based on the differential orbit correction solution for orbital and dynamic parameters. The aforementioned measured composite drag coefficient can refer to the actual composite drag coefficient, which can be denoted as Cd. tz The actual overall drag coefficient can refer to the overall drag coefficient experienced by a spacecraft when it is running in the atmosphere. This coefficient is affected by a variety of factors, including the shape of the spacecraft, surface materials, speed, air density, etc.

[0033] In an alternative embodiment, the measured comprehensive drag coefficient Cd can be calculated through orbit determination. tzSpecifically, we can first collect measured data of the spacecraft in orbit, including its orbital parameters, speed, and mass. Using this data, we can analyze the changes in the spacecraft's orbit, including changes in altitude and speed, through mathematical models and calculation methods. Based on the analyzed orbital changes, we can use aerodynamic theory and numerical simulation methods to calculate the overall drag experienced by the spacecraft in orbit. This drag typically includes atmospheric drag, solar radiation pressure, and other external disturbances. Finally, by comparing the calculated results with the measured data, we can verify the accuracy of the calculation and improve and optimize the calculation method. Through these steps, we can calculate the measured overall drag coefficient through orbit determination, providing important reference and support for the spacecraft's flight in orbit.

[0034] The initial value of the atmospheric drag coefficient mentioned above can refer to the empirical value of the atmospheric drag coefficient. The empirical value of the atmospheric drag coefficient refers to the empirical value used in the aerospace field to calculate the magnitude of the drag experienced by an aircraft in the atmosphere. This coefficient is usually obtained through experiments and observations and can be determined based on the shape, speed and environmental conditions of the aircraft.

[0035] In one alternative embodiment, extensive experiments and observations can be conducted to determine empirical values ​​for the atmospheric drag coefficient. This can be achieved by utilizing flight experiment data of aircraft or spacecraft in the atmosphere, analyzing the motion and forces acting on the aircraft at different speeds and altitudes, and thus deriving empirical values ​​for the atmospheric drag coefficient. Alternatively, empirical values ​​for the atmospheric drag coefficient can be determined through numerical simulation and calculation. Using tools such as computational fluid dynamics, the motion and forces acting on the aircraft in the atmosphere can be simulated to obtain empirical values ​​for the atmospheric drag coefficient. Combining these methods, empirical values ​​for the atmospheric drag coefficient can be determined by combining experimental data and numerical simulation results with theoretical derivation. These empirical values ​​can serve as important reference data for spacecraft design and flight control.

[0036] Step S104: Determine the nominal comprehensive drag coefficient of the target orbit based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft.

[0037] The nominal comprehensive resistance coefficient is the comprehensive resistance coefficient obtained by theoretical calculation.

[0038] The aforementioned preset mass can refer to the initial mass of the spacecraft before the mass change, which can be denoted as m1.

[0039] The aforementioned nominal comprehensive resistance coefficient can be an index-based comprehensive resistance coefficient, denoted as Cd. tz0 In the aerospace field, this can refer to the coefficient of comprehensive drag experienced by an aircraft under specific conditions. This coefficient takes into account factors such as the shape, surface roughness, and speed of the aircraft and is used to calculate the motion state and performance of the aircraft in the atmosphere.

[0040] In an alternative embodiment, the initial value of the atmospheric drag coefficient Cd can be used. drag Calculate the atmospheric drag acceleration a at the current orbit. drag Using the initial value of atmospheric drag coefficient Cd drag The attitude control jet acceleration a obtained from orbit determination of other model parameters. wol Calculate the nominal combined drag coefficient Cd, which includes atmospheric drag and attitude control jet propulsion. tz0 The calculation formula can be:

[0041] Cd tz0 =(1+a) wol / a drag )·Cd drag ;

[0042] Among them, a wol For the attitude control jet acceleration to be solved for orbit determination, Cd drag Let a be the initial value of the atmospheric drag coefficient for the current orbit. drag Given the current orbit and initial value Cd drag The corresponding atmospheric drag acceleration.

[0043] Step S106: Determine the target value of the atmospheric drag coefficient based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient.

[0044] The aforementioned target value of atmospheric drag coefficient can refer to the true value of atmospheric drag coefficient, or the actual value of atmospheric drag coefficient when a spacecraft or aircraft is affected by atmospheric drag. This value can be obtained through experiments or simulation calculations and is used to measure the degree to which air hinders the movement of an aircraft.

[0045] In one optional embodiment, after obtaining the measured comprehensive drag coefficient Cd tz and nominal combined drag coefficient Cd tz0 Then, the measured comprehensive drag coefficient Cd can be calculated. tz and nominal combined drag coefficient Cd tz0 The error between them is Δ=Cd tz -Cd tz0 Compare the error with the convergence threshold lim = 0.05. When Δ ≥ lim, Cd drag =Cd drag +Δ, repeat the above steps, and when Δ < lim, it is considered that the target value of the atmospheric drag coefficient has been obtained, that is, the true value of the atmospheric drag coefficient Cd. drag0 =Cd drag .

[0046] Step S108: Estimate the target mass of the spacecraft based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient.

[0047] The target mass mentioned above can refer to the mass of the spacecraft after the mass changes, and can be denoted as m0.

[0048] In an optional embodiment, for the atmospheric drag acceleration formula, if the orbit determination residual fitting tends to white noise, then it can be considered that the calculated value of the total acceleration in orbital dynamics is constant and approximates the true value. Correspondingly, the calculated value of the atmospheric drag acceleration is also constant, thus allowing the use of the Cd obtained from the orbit determination solution based on mass m1. drag1 Meanwhile, Cd is obtained by using mass m0 for orbit determination. drag0 This allows us to obtain the target mass of the spacecraft and estimate its mass.

[0049] In an optional embodiment, the target mass estimation method for a spacecraft can first calculate the spacecraft orbit and the actual combined drag coefficient, including: inputting parameters and initializing them; initializing the spacecraft orbit integral; reading in orbit measurement data; calculating theoretical observations, partial derivatives, and residuals; determining whether the orbit measurement data has been completely read; if not, repeating the process of reading in orbit measurement data; if the data has been completely read, solving the equations; calculating correction parameters and correcting the state variables; determining whether the state variables have converged; if the state variables have not converged, repeating the initialization of the spacecraft orbit integral.

[0050] Furthermore, when the state variables converge, the spacecraft orbit and attitude control jet acceleration are calculated, including initializing the atmospheric drag coefficient, calculating the atmospheric drag perturbation acceleration, initializing the spacecraft orbit integral, reading in the orbit measurement data, calculating the theoretical observations, partial derivatives, and residuals, and determining whether the orbit measurement data has been completely read. If the orbit measurement data has not been completely read, the process of reading in the orbit measurement data is repeated. When the orbit measurement data has been completely read, the normal equations are solved, the correction parameters are calculated, and the state variables are corrected. If the state variables have not converged, the process of initializing the spacecraft orbit integral is repeated. When the state variables converge, the nominal comprehensive drag coefficient is calculated, the comprehensive drag coefficient deviation is compared, and it is determined whether the deviation has converged. If the deviation has not converged, the process of initializing the atmospheric drag coefficient is repeated. When the deviation converges, the target mass of the spacecraft can be calculated.

[0051] In this embodiment of the invention, the measured comprehensive drag coefficient of the target orbit and the initial value of the atmospheric drag coefficient of the target orbit are obtained, wherein the measured comprehensive drag coefficient is the comprehensive drag coefficient obtained by actual measurement; the nominal comprehensive drag coefficient of the target orbit is determined based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft, wherein the nominal comprehensive drag coefficient is the comprehensive drag coefficient obtained by theoretical calculation; the target value of the atmospheric drag coefficient is determined based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; the target mass of the spacecraft is estimated based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient. This invention establishes a dynamic model for near-Earth spacecraft, compares the comprehensive drag coefficient calculated by the theoretical model with the actual calculated value, and when the error is large, iteratively approximates the actual calculated value of the comprehensive drag coefficient by using the deviation correction initial value of the atmospheric drag coefficient, thereby achieving accurate calculation of the atmospheric drag coefficient; based on the characteristics of the atmospheric drag calculated by the orbit, the parameter changes in the analytical formula are analyzed, and finally the proportional relationship between the atmospheric drag coefficient and the mass is analyzed. Using the known conditions, the actual mass of the spacecraft is finally obtained, thereby solving the technical problem of low efficiency in estimating the mass of spacecraft in related technologies.

[0052] Optionally, estimating the target mass of the spacecraft based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient includes: determining a first ratio between the preset mass and the initial value of the atmospheric drag coefficient; and determining a first product of the first ratio and the target value of the atmospheric drag coefficient as the target mass.

[0053] When the residuals calculated from two different parameters are constant, the spacecraft's position and velocity can be considered to be the same, that is, the magnitude of the spacecraft's velocity relative to the atmosphere, v. r Similarly, the direction of the spacecraft's velocity relative to the atmosphere is e v The atmospheric density ρ calculated using empirical models can also be considered the same. When the spacecraft is in a three-axis Earth-oriented flight attitude, the drag equivalent area can be divided into the body cross-sectional area and the solar panel area. The body cross-sectional area remains unchanged, while the solar panel normal changes under continuous rotation towards the sun. The rotation period of the near-Earth spacecraft's solar panels is consistent with the orbital period.

[0054] Figure 2 This is a schematic diagram of the windward area of ​​an optional spacecraft's atmospheric drag according to an embodiment of the present invention. The windward area when the solar panel rotates at a certain angle is shown in the diagram. Figure 2 As shown, the specific equivalent area can be expressed as:

[0055] S = S body +S plane ;

[0056]

[0057] Therefore, based on the above analysis, the following formula is obtained:

[0058]

[0059]

[0060] Since orbit determination uses GNSS positioning data, and the residual fitting of the calculated drag coefficient and empirical force is basically consistent, the atmospheric drag acceleration calculated by the two different model parameters can be considered equal. When the area is equal, the atmospheric drag coefficient is proportional to the mass. drag1 ,m1,Cd drag0 Given this information, we can obtain the actual mass m0 of the spacecraft.

[0061] Optionally, the target value of the atmospheric drag coefficient is determined based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient, including: determining the coefficient error of the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; and determining the target value of the atmospheric drag coefficient based on the coefficient error and the initial value of the atmospheric drag coefficient.

[0062] The aforementioned coefficient error can refer to the difference between the measured comprehensive resistance coefficient and the nominal comprehensive resistance coefficient.

[0063] In one optional embodiment, the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient can be calculated, with the coefficient error Δ=Cd. tz -Cd tz0 The target value of the atmospheric drag coefficient can be determined based on the coefficient error and the initial value of the atmospheric drag coefficient.

[0064] Optionally, determining the target value of the atmospheric drag coefficient based on the coefficient error and the initial value of the atmospheric drag coefficient includes: if the coefficient error is greater than or equal to a preset threshold, determining the target value of the atmospheric drag coefficient based on the first sum of the initial value of the atmospheric drag coefficient and the preset threshold; if the coefficient error is less than the preset threshold, determining the initial value of the atmospheric drag coefficient as the target value of the atmospheric drag coefficient.

[0065] The aforementioned preset threshold can refer to a pre-set convergence threshold lim. In this embodiment, the preset threshold lim can be 0.05. The preset threshold can be set according to actual needs, and is not limited here.

[0066] In an alternative embodiment, the coefficient error can be compared with the convergence threshold lim = 0.05. When Δ ≥ lim, Cd can be determined. drag =Cd drag +Δ; when Δ < lim, the true value of the atmospheric drag coefficient Cd is considered to be obtained. drag0 =Cddrag .

[0067] Optionally, determining the target value of the atmospheric drag coefficient based on the initial value of the atmospheric drag coefficient and a first sum of a preset threshold includes: determining the first sum as the initial value of the atmospheric drag coefficient, and performing orbit determination on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft to obtain the nominal comprehensive drag coefficient of the target orbit, until the coefficient error is less than the preset threshold, and determining the initial value of the atmospheric drag coefficient as the target value of the atmospheric drag coefficient.

[0068] In an alternative embodiment, the coefficient error can be compared with a convergence threshold lim = 0.05, and when Δ ≥ lim, Cd drag =Cd drag +Δ, repeat the steps of solving for the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft to obtain the nominal comprehensive drag coefficient of the target orbit, until Δ < lim, at which point the true value of the atmospheric drag coefficient Cd is considered to have been obtained. drag0 =Cd drag .

[0069] Optionally, the nominal comprehensive drag coefficient of the target orbit is determined based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft, including: determining the atmospheric drag acceleration of the target orbit based on the initial value of the atmospheric drag coefficient; performing orbit determination on the atmospheric drag acceleration and the preset mass to obtain the attitude control jet acceleration of the spacecraft in the target orbit; and determining the nominal comprehensive drag coefficient based on the attitude control jet acceleration, the initial value of the atmospheric drag coefficient, and the atmospheric drag acceleration.

[0070] In one optional embodiment, determining the nominal composite drag coefficient may include the following steps: initializing the atmospheric drag coefficient, calculating the atmospheric drag perturbation acceleration, initializing the spacecraft orbital integral, reading in the orbit measurement data, calculating the theoretical observations, partial derivatives, and residuals, determining whether the orbit measurement data has been completely read, repeating the reading of orbit measurement data if the orbit measurement data has not been completely read, solving the normal equations when the orbit measurement data has been completely read, calculating the correction parameters and correcting the state variables, repeating the spacecraft orbital integral initialization if the state variables have not converged, and calculating the nominal composite drag coefficient when the state variables have converged.

[0071] Optionally, the nominal comprehensive drag coefficient is determined based on the attitude control jet acceleration, the initial value of the atmospheric drag coefficient, and the atmospheric drag acceleration, including: determining a second ratio of the attitude control jet acceleration and the atmospheric drag acceleration; determining a second sum of the preset value and the second ratio; and determining a second product of the second sum and the initial value of the atmospheric drag coefficient as the nominal comprehensive drag coefficient.

[0072] In an alternative embodiment, the initial value of the atmospheric drag coefficient Cd can be used. drag Calculate the atmospheric drag acceleration a at the current orbit.drag Using the initial value of atmospheric drag coefficient Cd drag The attitude control jet acceleration a obtained from orbit determination of other model parameters. wol Calculate the nominal combined drag coefficient Cd, which includes atmospheric drag and attitude control jet propulsion. tz0 The calculation formula can be:

[0073] Cd tz0 =(1+a) wol / a drag )·Cd drag ;

[0074] Among them, a wol For the attitude control jet acceleration to be solved for orbit determination, Cd drag Let a be the initial value of the atmospheric drag coefficient for the current orbit. drag Given the current orbit and initial value Cd drag The corresponding atmospheric drag acceleration.

[0075] Spacecraft mass is one of the fundamental physical properties of a spacecraft and an important physical parameter during its flight mission. A precise mass measurement process is involved before launch. Typically, after a spacecraft enters orbit, changes in its mass refer to fuel consumption and the release of test vehicles. Fuel consumption can be measured using telemetry parameters, and the mass of the released vehicles is also precisely measured before launch. Under these circumstances, changes in spacecraft mass can be accurately calculated.

[0076] In some cases, near-Earth spacecraft frequently employ attitude control jetting based on mission requirements, resulting in significant changes in mass. This mass change refers not to the mass of fuel consumed by attitude control jetting, but rather to the mass exchange caused by refueling, cargo transport, and other actions after rendezvous and docking. When the mass change after spacecraft separation and departure cannot be accurately assessed, it can impact the control and management of subsequent flight missions. In such situations, short-term engine pulse operation can be used to analyze changes in the spacecraft's orbit, and the current mass of the spacecraft can be calculated using the momentum theorem. This invention proposes a novel spacecraft mass assessment method for the independent flight process after spacecraft departure, combining the characteristics of attitude control jetting forces and atmospheric drag.

[0077] This application proposes a method for estimating the mass of near-Earth spacecraft. Utilizing the similar effects of atmospheric drag and attitude control jet forces on the trajectory in near-Earth spacecraft orbital dynamics, a differential correction equation is established to obtain the attitude control jet acceleration and the combined drag coefficient of the two forces. Then, based on the relationship between acceleration and drag coefficient, an iterative algorithm is constructed to obtain the atmospheric drag force coefficient and the final mass of the spacecraft. Specifically, the method includes the following steps: designing a near-Earth spacecraft orbital dynamics model, constructing a method for calculating the atmospheric drag coefficient, calculating the spacecraft mass using the atmospheric drag model, and performing spacecraft mass estimation and analysis.

[0078] An optional embodiment of the present invention will now be described in detail.

[0079] Step 1: Design of Orbital Dynamics Model for Near-Earth Spacecraft

[0080] Near-Earth spacecraft mainly refer to spacecraft in orbit with an altitude below 1000 kilometers. The position, velocity, and total acceleration of a spacecraft at time T can be expressed as: Total acceleration is expressed as the sum of the accelerations of the principal force and all types of perturbing forces. The statistics of accelerations of various forces are shown in Table 1:

[0081] Table 1. Statistics of Orbital Dynamics Models for Near-Earth Spacecraft

[0082]

[0083]

[0084] The atmospheric drag acceleration can be expressed as:

[0085]

[0086] Among them, Cd drag Here, S is the atmospheric drag coefficient, m is the spacecraft drag equivalent area, ρ is the spacecraft mass, and v is the atmospheric density. Empirical models can be used, but this method recommends using the MSIS2000 model. r e represents the velocity modulus of the spacecraft relative to the atmosphere. v Let be the direction of the spacecraft's velocity relative to the atmosphere. The partial derivative of acceleration with respect to the drag coefficient is:

[0087]

[0088] The acceleration of the attitude control jet force along the RTN direction in the orbital coordinate system can be expressed as:

[0089]

[0090] Its model partial derivatives are:

[0091]

[0092] Among them, [a x a y a z [a] represents the acceleration component in the inertial frame of reference. R a T a N [] represents the acceleration component in the orbital frame, and [RTN] represents the transformation matrix from the inertial frame to the orbital frame.

[0093] Step 2: Constructing a method for calculating atmospheric drag coefficient

[0094] The spacecraft's orbit and other state parameters can be expressed as:

[0095]

[0096] In the formula, P represents other dynamic or geometric parameters to be estimated, which should typically include atmospheric drag coefficient, attitude control jet force, etc.

[0097] remember X(t0) = X0, the system state equation can be expressed by the following equation:

[0098]

[0099] Record reference state X * (t), satisfying

[0100]

[0101] F in reference state X * Linearize at (t) and denote x(t) = X(t) - X * (t), we obtain the linearized state equation, which is expressed by the following equation:

[0102]

[0103] in, The solution to the above equation can be described in the form of the following equation.

[0104] x(t) = Φ(t,t0)x(t0);

[0105] Where Φ(t,t0) is the state transition matrix, which is obtained by integrating the following formula:

[0106]

[0107] In the above formula, I is the identity matrix.

[0108] t i Observation Y at time i With state variable Xi The functional relationship between them is described by the following formula:

[0109] Y i =G(X) i ,t i )+ε i ;

[0110] Where Xi, Yi, and εi represent the state, observation, and observation noise at time ti, respectively. Let y i =Y i -G(X i * ,t i The linearized observation equation is described by the following formula:

[0111] y i =H i x0+ε i ;

[0112] Where, x0 = x(t0), remember The overall observation equation is expressed by the following formula:

[0113] y = Hx0 + ε;

[0114] A statistical solution is performed based on least squares, and the above process is repeated using Newton-Raphson iteration. The latest best estimate is used each time. As the standard value for linearization, the iteration continues until the differential correction vector x0 becomes small. Using the above method, we can obtain not only the spacecraft's orbital information, but also the atmospheric drag coefficient and the force acceleration of the attitude control jet.

[0115] Step 3: Calculate the spacecraft mass using an atmospheric drag model.

[0116] Since atmospheric drag only has a decelerating and dragging effect on the spacecraft along its track, this method only considers the acceleration and deceleration forces mainly acting along the track direction when calculating attitude control jets. Based on the differential orbit correction solution for orbit and dynamic parameters, the following processing flow is designed:

[0117] Actual combined drag coefficient Cd tz The estimate can be obtained by calculating the measured comprehensive drag coefficient Cd through orbit determination. tz Nominal combined drag coefficient Cd tz0 The estimate can be made using the initial value of the atmospheric drag coefficient Cd. drag Calculate the atmospheric drag acceleration a at the current orbit. drag The initial value of the atmospheric drag coefficient Cd can be used. drag The attitude control jet acceleration a obtained from orbit determination of other model parameters. wolCalculate the nominal combined drag coefficient Cd, which includes atmospheric drag and attitude control jet propulsion. tz0 =(1+a) wol / a drag )·Cd drag , where a wol For the attitude control jet acceleration to be solved for orbit determination, Cd drag Let a be the initial value of the atmospheric drag coefficient for the current orbit. drag Given the current orbit and initial value Cd drag The corresponding atmospheric drag acceleration; calculate the error Δ=Cd between the measured and nominal combined drag coefficient. tz -Cd tz0 Comparing the error with the convergence threshold lim = 0.05, when Δ ≥ lim, Cd drag =Cd drag +Δ, repeat the above steps, and when Δ < lim, it is considered that the true value of the atmospheric drag coefficient Cd has been obtained. drag0 =Cd drag .

[0118] In estimating the actual mass, for the atmospheric drag acceleration formula, if the orbit determination residual fit tends to white noise, then it can be assumed that the calculated total acceleration in orbital dynamics is constant and approximates the true value. Correspondingly, the calculated atmospheric drag acceleration is also constant. If Cd is obtained using the orbit determination solution based on mass m1... drag1 Meanwhile, the Cd obtained by orbit determination using mass m0 is... drag0 .

[0119] When the residuals calculated from two different parameters are constant, the spacecraft's position and velocity can be considered to be the same, that is, the magnitude of the spacecraft's velocity relative to the atmosphere, v. r Similarly, the direction of the spacecraft's velocity relative to the atmosphere is e v The atmospheric density ρ calculated using empirical models can also be considered the same. When the spacecraft is in a three-axis Earth-oriented flight attitude, the drag equivalent area can be divided into the body cross-sectional area and the solar panel area. The body cross-sectional area remains constant, while the solar panel normal changes as it continuously rotates towards the sun. The rotation period of the near-Earth spacecraft's solar panels coincides with the orbital period. The windward area when the solar panels rotate at a certain angle is as follows: Figure 2 As shown, the specific equivalent area can be expressed as:

[0120] S = S body +S plane ;

[0121]

[0122] Therefore, based on the above analysis, the following formula is obtained:

[0123]

[0124] Ultimately, we can obtain:

[0125]

[0126] Since orbit determination uses GNSS positioning data, the residual fitting obtained by solving for the drag coefficient and the empirical force is basically consistent, so the atmospheric drag acceleration calculated by the two different model parameters can be considered equal. When the area is equal, the atmospheric drag coefficient is proportional to the mass. If Cd drag1 ,m1,Cd drag0 Given this information, we can obtain the actual mass m0.

[0127] Figure 3 This is a flowchart illustrating an optional method for calculating spacecraft mass using a dynamic model, according to an embodiment of the present invention. Figure 3 As shown, the spacecraft mass calculation process is initiated to calculate the spacecraft orbit and actual combined drag coefficient. This includes inputting and initializing parameters, initializing the spacecraft orbit integral, reading in orbit measurement data, calculating theoretical observations, partial derivatives, and residuals, determining whether the orbit measurement data has been completely read, and repeating the process if the data has not been completely read. If the data has been completely read, the process is repeated. If the data has been completely read, the method equations are solved, correction parameters are calculated, and the state variables are corrected. The process is then determined whether the state variables have converged. If the state variables have not converged, the spacecraft orbit integral initialization process is repeated.

[0128] Furthermore, when the state variables converge, the spacecraft orbit and attitude control jet acceleration are calculated, including initializing the atmospheric drag coefficient, calculating the atmospheric drag perturbation acceleration, initializing the spacecraft orbit integral, reading in the orbit measurement data, calculating the theoretical observations, partial derivatives, and residuals, and determining whether the orbit measurement data has been completely read. If the orbit measurement data has not been completely read, the process of reading in the orbit measurement data is repeated. When the orbit measurement data has been completely read, the normal equations are solved, the correction parameters are calculated, and the state variables are corrected. If the state variables have not converged, the process of initializing the spacecraft orbit integral is repeated. When the state variables converge, the nominal comprehensive drag coefficient is calculated, the deviation of the comprehensive drag coefficient is compared, and it is determined whether the deviation has converged. If the deviation has not converged, the process of initializing the atmospheric drag coefficient is repeated. When the deviation has converged, the spacecraft mass can be calculated, and the spacecraft mass estimation process ends.

[0129] Step 4: Spacecraft Mass Estimation Analysis

[0130] Based on the above steps, a simulation is performed to calculate the orbit and mass of a set of spacecraft in orbit, with an orbital altitude of 390 kilometers and the nominal value of the atmospheric drag coefficient Cd for the current spacecraft's orbit. drag0 =2.0, the initial value of the atmospheric drag coefficient Cd is used in the calculation. drag=1.6, nominal mass m = 11000 kg, initial mass used in the calculation m1 = 9500 kg, equivalent area 28 m² 2 The equivalent perturbation acceleration a of attitude control jet wol = -1.32 × 10 -6 m / s 2 The above orbit determination solution can be used. Figure 4 This is a schematic diagram of an optional spacecraft atmospheric density and atmospheric drag acceleration according to an embodiment of the present invention. The atmospheric drag density and acceleration are calculated using an empirical model as follows: Figure 4 As shown, Figure 4 The upper and lower graphs represent atmospheric drag density calculated using empirical models. The horizontal axis represents time (t), in hours; the vertical axis represents the corresponding atmospheric drag density (Rou), in kg / m³. 3 ; Figure 4 The lower figure shows the atmospheric drag acceleration calculated using an empirical model. The horizontal axis represents time t (in hours), and the vertical axis represents the corresponding atmospheric drag acceleration Acc (in m / s²). 2 ,based on Figure 4 Cd can be calculated drag1 =1.705, which is the true value of the orbit determination calculation iteration convergence. When calculating atmospheric drag using an empirical model, the calculated value is: m0 = 11182.69 kg. The deviation from the simulated nominal value is 182.69 kg, with a relative deviation of 1.6%.

[0131] The above analysis verifies the correctness of this method. This method is applicable to the quality assessment of cargo spacecraft after separation and evacuation from the space station.

[0132] This invention proposes a method for estimating the mass of near-Earth spacecraft. Addressing the challenges of significant mass variations and frequent attitude control jet propulsion in near-Earth spacecraft, a refined dynamic model is established, incorporating atmospheric drag and attitude control jet forces. Based on the model, differential correction equations are constructed to calculate the actual attitude control jet force acceleration and the overall drag coefficient. The calculated overall drag coefficient is then compared with the actual calculated value. When the error is large, the initial value of the atmospheric drag coefficient is iteratively approximated using a deviation correction, achieving accurate calculation of the atmospheric drag coefficient. Based on the characteristics of atmospheric drag calculated from the orbit, the parameter variations in the analytical formula are analyzed, ultimately revealing the proportional relationship between the atmospheric drag coefficient and the mass. Using known conditions, the actual spacecraft mass is finally obtained.

[0133] Example 2

[0134] According to another aspect of the present invention, a spacecraft mass estimation apparatus is also provided. This apparatus can execute the spacecraft mass estimation method of the above embodiments. The specific implementation method and preferred application scenarios are the same as those of the above embodiments, and will not be described in detail here.

[0135] Figure 5 This is a schematic diagram of a spacecraft mass estimation device according to an embodiment of the present invention, such as... Figure 5 As shown, the device includes the following: an acquisition module 502, a first determination module 504, a second determination module 506, and an estimation module 508.

[0136] The system includes: an acquisition module 502 for acquiring the measured comprehensive drag coefficient of the target orbit and the initial value of the atmospheric drag coefficient of the target orbit, wherein the measured comprehensive drag coefficient is the comprehensive drag coefficient obtained by actual measurement; a first determination module 504 for determining the nominal comprehensive drag coefficient of the target orbit based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft, wherein the nominal comprehensive drag coefficient is the comprehensive drag coefficient obtained by theoretical calculation; a second determination module 506 for determining the target value of the atmospheric drag coefficient based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; and an estimation module 508 for estimating the target mass of the spacecraft based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient.

[0137] In the above embodiments of this application, the estimation module includes: a first determining unit and a second determining unit.

[0138] The first determining unit is used to determine a first ratio between the preset mass and the initial value of the atmospheric drag coefficient; the second determining unit is used to determine the first product of the first ratio and the target value of the atmospheric drag coefficient as the target mass.

[0139] In the above embodiments of this application, the second determining module includes: a third determining unit and a fourth determining unit.

[0140] The third determining unit is used to determine the coefficient error between the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; the fourth determining unit is used to determine the target value of the atmospheric drag coefficient based on the coefficient error and the initial value of the atmospheric drag coefficient.

[0141] In the above embodiments of this application, the fourth determining unit includes: a first determining subunit and a second determining subunit.

[0142] The first determining subunit is used to determine the target value of the atmospheric drag coefficient based on the initial value of the atmospheric drag coefficient and the first sum of the preset threshold if the coefficient error is greater than or equal to the preset threshold; the second determining subunit is used to determine the initial value of the atmospheric drag coefficient as the target value of the atmospheric drag coefficient if the coefficient error is less than the preset threshold.

[0143] The first determining subunit is also used to determine the first sum as the initial value of the atmospheric drag coefficient, and to perform the orbit determination solution based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft to obtain the nominal comprehensive drag coefficient of the target orbit, until the coefficient error is less than the preset threshold, and to determine the initial value of the atmospheric drag coefficient as the target value of the atmospheric drag coefficient.

[0144] In the above embodiments of this application, the first determining module includes: a fifth determining unit, a solving unit, and a sixth determining unit.

[0145] The fifth determining unit is used to determine the atmospheric drag acceleration of the target orbit based on the initial value of the atmospheric drag coefficient; the solving unit is used to solve the orbit determination problem based on the atmospheric drag acceleration and the preset mass to obtain the attitude control jet acceleration of the spacecraft in the target orbit; the sixth determining unit is used to determine the nominal comprehensive drag coefficient based on the attitude control jet acceleration, the initial value of the atmospheric drag coefficient and the atmospheric drag acceleration.

[0146] In the above embodiments of this application, the sixth determining unit includes: a third determining subunit, a fourth determining subunit, and a fifth determining subunit.

[0147] The third determining subunit is used to determine the second ratio of attitude control jet acceleration to atmospheric drag acceleration; the fourth determining subunit is used to determine the second sum of the preset value and the second ratio; and the fifth determining subunit is used to determine the second product of the second sum and the initial value of the atmospheric drag coefficient as the nominal comprehensive drag coefficient.

[0148] Example 3

[0149] According to another aspect of the present invention, a computer-readable storage medium is also provided, the computer-readable storage medium including a stored program, wherein, when the program is executed, the aforementioned spacecraft mass estimation method is executed in the processor of the device.

[0150] The computer storage medium mentioned in the above steps can be a medium used in computer memory to store certain discontinuous physical quantities. Computer storage media mainly include semiconductors, magnetic cores, magnetic drums, magnetic tapes, and laser discs. Computer-readable storage media includes stored programs, which can be a set of instructions that a computer can recognize and execute, running on an electronic computer to meet certain human needs—an information tool.

[0151] Example 4

[0152] According to another aspect of the present invention, an electronic device is also provided, comprising one or more processors; a storage device for storing one or more programs; and when the one or more programs are executed by the one or more processors, causing the one or more processors to perform the aforementioned spacecraft mass estimation method.

[0153] The storage device in the above steps can be a type of sequential logic circuit, a memory component used to store data and instructions, mainly used to store programs and data; the processor can be a functional unit that interprets and executes instructions, and it has a unique set of operation commands, which can be called the processor's instruction set, such as store, load, etc.; the storage device stores computer programs, which can be a set of instructions that a computer can recognize and execute, running on an electronic computer, and is an information tool that meets people's certain needs.

[0154] In the above embodiments of the present invention, the descriptions of each embodiment have different focuses. For parts not described in detail in a certain embodiment, please refer to the relevant descriptions of other embodiments.

[0155] In the several embodiments provided by this invention, it should be understood that the disclosed technical content can be implemented in other ways. The device embodiments described above are merely illustrative; for example, the division of units can be a logical functional division, and in actual implementation, there may be other division methods. For example, multiple units or components may be combined or integrated into another system, or some features may be ignored or not executed. Furthermore, the displayed or discussed mutual coupling, direct coupling, or communication connection can be through some interfaces; the indirect coupling or communication connection of units or modules can be electrical or other forms.

[0156] The units described as separate components may or may not be physically separate. The components shown as units may or may not be physical units; that is, they may be located in one place or distributed across multiple units. Some or all of the units can be selected to achieve the purpose of this embodiment according to actual needs.

[0157] Furthermore, the functional units in the various embodiments of the present invention can be integrated into one processing unit, or each unit can exist physically separately, or two or more units can be integrated into one unit. The integrated unit can be implemented in hardware or as a software functional unit.

[0158] If the integrated unit is implemented as a software functional unit and sold or used as an independent product, it can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present invention, in essence, or the part that contributes to the prior art, or all or part of the technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and includes several instructions to cause a computer device (which may be a personal computer, server, or network device, etc.) to execute all or part of the steps of the methods described in the various embodiments of the present invention. The aforementioned storage medium includes various media capable of storing program code, such as USB flash drives, read-only memory (ROM), random access memory (RAM), portable hard drives, magnetic disks, or optical disks.

[0159] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A method for estimating the mass of a spacecraft, characterized in that, include: The measured comprehensive drag coefficient of the target orbit where the spacecraft is located and the initial value of the atmospheric drag coefficient of the target orbit are obtained, wherein the measured comprehensive drag coefficient is the comprehensive drag coefficient obtained by actual measurement; The nominal comprehensive drag coefficient of the target orbit is determined based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft, wherein the nominal comprehensive drag coefficient is a comprehensive drag coefficient obtained by theoretical calculation. The target value of the atmospheric drag coefficient is determined based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient. The target mass of the spacecraft is estimated based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient.

2. The method according to claim 1, characterized in that, Estimate the target mass of the spacecraft based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient, including: Determine a first ratio between the preset mass and the initial value of the atmospheric drag coefficient; The first product of the first ratio and the target value of the atmospheric drag coefficient is determined as the target mass.

3. The method according to claim 1, characterized in that, Determining the target value of the atmospheric drag coefficient based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient includes: Determine the coefficient error between the measured comprehensive resistance coefficient and the nominal comprehensive resistance coefficient; The target value of the atmospheric drag coefficient is determined based on the coefficient error and the initial value of the atmospheric drag coefficient.

4. The method according to claim 3, characterized in that, Determining the target value of the atmospheric drag coefficient based on the coefficient error and the initial value of the atmospheric drag coefficient includes: If the coefficient error is greater than or equal to a preset threshold, the target value of the atmospheric drag coefficient is determined based on the initial value of the atmospheric drag coefficient and the first sum of the preset threshold. If the coefficient error is less than the preset threshold, the initial value of the atmospheric drag coefficient is determined to be the target value of the atmospheric drag coefficient.

5. The method according to claim 4, characterized in that, Determining the target value of the atmospheric drag coefficient based on the initial value of the atmospheric drag coefficient and the first sum of the preset threshold includes: The first sum is determined to be the initial value of the atmospheric drag coefficient, and the orbit determination is performed on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft to obtain the nominal comprehensive drag coefficient of the target orbit, until the coefficient error is less than the preset threshold, and the initial value of the atmospheric drag coefficient is determined to be the target value of the atmospheric drag coefficient.

6. The method according to claim 1, characterized in that, The nominal comprehensive drag coefficient of the target orbit is determined based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft, including: The atmospheric drag acceleration of the target orbit is determined based on the initial value of the atmospheric drag coefficient. The atmospheric drag acceleration and the preset mass are used to solve for orbit determination, thereby obtaining the attitude control jet acceleration of the spacecraft in the target orbit; The nominal combined drag coefficient is determined based on the attitude control jet acceleration, the initial value of the atmospheric drag coefficient, and the atmospheric drag acceleration.

7. The method according to claim 6, characterized in that, The nominal combined drag coefficient is determined based on the attitude control jet acceleration, the initial value of the atmospheric drag coefficient, and the atmospheric drag acceleration, including: Determine a second ratio between the attitude control jet acceleration and the atmospheric drag acceleration; Determine the second sum of the preset value and the second ratio; The second product of the second sum and the initial value of the atmospheric drag coefficient is determined to be the nominal comprehensive drag coefficient.

8. A spacecraft mass estimation device, characterized in that, include: The acquisition module is used to acquire the measured comprehensive drag coefficient of the target orbit where the spacecraft is located and the initial value of the atmospheric drag coefficient of the target orbit, wherein the measured comprehensive drag coefficient is the comprehensive drag coefficient obtained by actual measurement; The first determining module is used to determine the nominal comprehensive drag coefficient of the target orbit based on the initial value of the atmospheric drag coefficient and the preset mass of the spacecraft, wherein the nominal comprehensive drag coefficient is a comprehensive drag coefficient obtained by theoretical calculation; The second determining module is used to determine the target value of the atmospheric drag coefficient based on the measured comprehensive drag coefficient and the nominal comprehensive drag coefficient; An estimation module is used to estimate the target mass of the spacecraft based on the initial value of the atmospheric drag coefficient, the preset mass, and the target value of the atmospheric drag coefficient.

9. A computer-readable storage medium, characterized in that, The computer-readable storage medium includes a stored program, wherein, when the program is executed, it controls the execution of the spacecraft mass estimation method according to any one of claims 1 to 7 in the processor of the device.

10. An electronic device, characterized in that, include: One or more processors; Storage device for storing one or more programs; When the one or more programs are executed by the one or more processors, the one or more processors perform the spacecraft mass estimation method according to any one of claims 1 to 7.

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