Spacecraft energy model construction method supporting multi-degree-of-freedom moving parts

Abstract

The embodiment of the application provides a spacecraft energy model construction method supporting a multi-degree-of-freedom movable component. A three-dimensional model of a body of a spacecraft to be analyzed and three-dimensional models of each sailboard assembly corresponding to each sailboard are constructed and combined. A solar cell array patch model is set on each three-dimensional model of the sailboard. The rotation angle of each three-dimensional model of the sailboard relative to the three-dimensional model of the body is set to obtain a spacecraft model. Orbit information, attitude information and epoch time information of the spacecraft to be analyzed are input into the spacecraft model, a spacecraft on-orbit model corresponding to the spacecraft to be analyzed is constructed, and a spacecraft energy model for the spacecraft on-orbit model is constructed. The spacecraft energy model is used for energy balance analysis of the spacecraft to be analyzed to obtain an energy balance state of the spacecraft to be analyzed in real time.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of spacecraft energy system, and particularly relates to a spacecraft energy model construction method supporting multi-degree-of-freedom movable components. BACKGROUND

[0002] The spacecraft energy system (referred to as energy system) is one of the key subsystems of a spacecraft, and the state of the energy system is directly related to the operation safety of the spacecraft and the success of the on-orbit mission. In the energy system, a solar sail is an essential component, and its function is to convert solar energy into electric energy to charge the battery of the spacecraft or directly provide power support for various devices (for example, communication devices, navigation devices, attitude and orbit control systems, and scientific application payloads).

[0003] At present, except for micro-nano spacecraft, the solar sail of a spacecraft is generally configured as an unfolded sail. The unfolded sail is unfolded after the spacecraft enters space. Through attitude control of the spacecraft and rotation of the unfolded sail, the sun can be tracked to ensure the maximum power generation efficiency of the solar sail.

[0004] However, the solar sail of the spacecraft is affected by the body and sail configuration, the orbit position, and the motion attitude, and at some moments, the solar sail is blocked by the body or other sails and cannot be in the best power generation state in real time. Especially when the spacecraft encounters some irresistible factors during launching or operation, the power generation capacity of the spacecraft is greatly affected in the case that part of the sail is damaged, thereby seriously restricting the working capacity of the spacecraft and shortening the service life of the spacecraft. In the face of multiple factors such as self and external factors, especially when the sail is damaged or damaged, how to accurately evaluate the blocked area and effective power generation area of the solar sail becomes the key to estimating the safety of the spacecraft energy system. SUMMARY

[0005] The technical problem to be solved by the present application is to overcome the shortcomings of the prior art, and to provide a spacecraft energy model construction method supporting multi-degree-of-freedom movable components, which is fast, high-precision, and suitable for any orbit of a spacecraft, any central body blocking, and any state of a solar sail, and can simulate and analyze the power generation state and energy balance of the spacecraft energy system in real time.

[0006] The technical solution of the present application to solve the above technical problem is as follows:

[0007] On the one hand, the present application provides a spacecraft energy model construction method supporting multi-degree-of-freedom movable components, comprising:

[0008] determining a spacecraft to be analyzed. The spacecraft to be analyzed includes a sail structure and a non-sail structure, and the sail structure includes a plurality of sail components and a plurality of sail connecting components.

[0009] Construct a three-dimensional model of the body corresponding to the spacecraft to be analyzed based on the non-sailboard structure.

[0010] For any sailboard assembly in the plurality of sailboard assemblies, construct a three-dimensional model of the sailboard corresponding to the sailboard assembly.

[0011] Based on the position layout of the sailboard structure and the non-sailboard structure, combine the three-dimensional model of the body and the three-dimensional models of the sailboards.

[0012] For any one of the three-dimensional models of the sailboards, based on the position of the solar cell array patch on the sailboard assembly corresponding to the three-dimensional model of the sailboard, set a solar cell array patch model on the three-dimensional model of the sailboard.

[0013] Set the rotation angle of each three-dimensional model of the sailboard relative to the three-dimensional model of the body to obtain a spacecraft model.

[0014] Input the orbit information, attitude information and epoch time information of the spacecraft to be analyzed into the spacecraft model to construct a spacecraft on-orbit model corresponding to the spacecraft to be analyzed.

[0015] Construct a spacecraft energy model for the spacecraft on-orbit model. The spacecraft energy model is used to perform a preset task, which includes:

[0016] Based on the position distribution between the sun model corresponding to the sun, the celestial body model corresponding to the target celestial body, and the spacecraft on-orbit model, determine whether the celestial body model exists to block the spacecraft on-orbit model.

[0017] In the case where the celestial body model does not exist to block the spacecraft on-orbit model, for any one of the three-dimensional models of the sailboards, calculate a target angle corresponding to the three-dimensional model of the sailboard. The target angle is the angle between a first vector and a second vector. The first vector is a pointing vector from the position of the three-dimensional model of the sailboard to the position of the sun model. The second vector is a normal vector corresponding to the solar cell array patch model set on the three-dimensional model of the sailboard.

[0018] In the case where the target angle corresponding to any one of the three-dimensional models of the sailboards is greater than or equal to 90°, determine that the number of power generation units of the three-dimensional model of the sailboard is 0.

[0019] In the case where the target angle corresponding to any one of the three-dimensional models of the sailboards is less than 90°, assign the value of 1 to the three-dimensional model of the sailboard, and assign the values of 0 to the three-dimensional model of the body and the target three-dimensional model of the sailboard. The target three-dimensional model of the sailboard is a three-dimensional model of the sailboard different from any one of the three-dimensional models of the sailboards.

[0020] A shadow area corresponding to any sailboard three-dimensional model on a solar cell array patch model is determined. The shadow area is formed based on a structure of any sailboard three-dimensional model and / or an occlusion of each target sailboard three-dimensional model to any sailboard three-dimensional model.

[0021] A surface of the solar cell array patch model corresponding to any sailboard three-dimensional model is segmented into a plurality of sub-areas based on a region segmentation rule.

[0022] A first sub-area and a second sub-area in the plurality of sub-areas are determined. The first sub-area is an area coinciding with the shadow area, and the second sub-area is an area not coinciding with the shadow area.

[0023] A sum of areas of each second sub-area corresponding to any sailboard three-dimensional model is determined as an effective power generation area of any sailboard three-dimensional model.

[0024] Based on the effective power generation area of any sailboard three-dimensional model, a solar radiation degree of a position where any sailboard three-dimensional model is located, an angle between any sailboard three-dimensional model and a solar ray emitted by a solar model, an operating temperature of a sailboard assembly corresponding to any sailboard three-dimensional model, and an aging rate of the sailboard assembly corresponding to any sailboard three-dimensional model, a power generation power of any sailboard three-dimensional model is determined.

[0025] A sum of the effective power generation areas of each sailboard three-dimensional model is determined as an effective power generation area of a spacecraft on-orbit model, and a sum of the power generation powers of each sailboard three-dimensional model is determined as a power generation power of the spacecraft on-orbit model.

[0026] Power consumptions of each load device included in a spacecraft to be analyzed are obtained. For any load device in each load device included in the spacecraft to be analyzed, the power consumption of the any load device is determined based on an operating state of the any load device. The operating state includes a power-on state and a power-off state. In addition, the operating state can also include a working mode and a time-varying parameter related to the power consumption.

[0027] Power consumptions of each power distribution board and power consumptions of each bus included in the spacecraft to be analyzed are obtained.

[0028] Based on the power consumptions of each load device, the power consumptions of each power distribution board, and the power consumptions of each bus, a total power consumption of the spacecraft to be analyzed is determined.

[0029] Operating state information of a battery included in the spacecraft to be analyzed is obtained. The operating state information of the battery includes temperature information, voltage information, and power information of the battery.

[0030] The power supply mode of the to-be-analyzed spacecraft is determined based on effective power generation area, power generation power, total power consumption of the to-be-analyzed spacecraft, and working state information of the battery. The effective power generation area of the to-be-analyzed spacecraft is determined based on the effective power generation area of the on-orbit model of the spacecraft, and the power generation power of the to-be-analyzed spacecraft is determined based on the power generation power of the on-orbit model of the spacecraft. The power supply mode of the to-be-analyzed spacecraft includes a battery power supply mode, a sailboard assembly power supply mode, and a joint power supply mode. The joint power supply mode is a mode in which the battery and the sailboard assembly supply power at the same time.

[0031] In a case where the power supply mode of the to-be-analyzed spacecraft is determined to be the battery power supply mode or the joint power supply mode, the discharge depth and the discharge voltage of the battery are acquired.

[0032] In a case where the discharge depth of the battery is detected to be greater than a discharge depth threshold value, and / or the discharge voltage of the battery is out of a preset discharge voltage range, prompt information is generated. The prompt information is used to prompt power supply abnormity of the to-be-analyzed spacecraft.

[0033] In a case where the power supply mode of the to-be-analyzed spacecraft is determined to be the sailboard assembly power supply mode, if the total power consumption of the to-be-analyzed spacecraft is greater than a power consumption threshold value, prompt information is generated.

[0034] On the basis of the above technical solutions, the application can be further improved as follows.

[0035] Further, in a case where the on-orbit model of the spacecraft is shielded by the celestial body model, the effective power generation area of the on-orbit model of the spacecraft is determined to be 0, and the power generation power of the on-orbit model of the spacecraft is determined to be 0.

[0036] Further, the connection mode between the first sailboard three-dimensional model and the body three-dimensional model is a three-degree-of-freedom movable connection. The first sailboard three-dimensional model is a sailboard three-dimensional model that is directly connected to the body three-dimensional model among the sailboard three-dimensional models. The connection mode between the second sailboard three-dimensional model and the first sailboard three-dimensional model is a one-degree-of-freedom movable connection. The second sailboard three-dimensional model is a sailboard three-dimensional model that is directly or indirectly connected to the first sailboard three-dimensional model among the sailboard three-dimensional models.

[0037] Further, in setting the rotation angle of each sailboard three-dimensional model relative to the body three-dimensional model, the positions of the sailboard three-dimensional models relative to the body three-dimensional model are calculated according to the topological relationship among the body three-dimensional model, the first sailboard three-dimensional model, and the second sailboard three-dimensional model.

[0038] Further, in a case where the power supply mode of the to-be-analyzed spacecraft is determined to be the sailboard assembly power supply mode, and the battery is in a charging state, if the charging current of the battery is greater than a charging current threshold value, information for prompting abnormality of the charging state of the battery is generated.

[0039] Further, any sailboard three-dimensional model comprises a first sticking surface and a second sticking surface. A solar cell array patch model is arranged on the first sticking surface comprised in any sailboard three-dimensional model.

[0040] Further, the declination and hour angle corresponding to the sun are obtained with the earth as a reference object, and the declination and hour angle corresponding to the target celestial body are obtained. A position vector of the sun relative to the earth is determined based on the radius of the earth, the declination and hour angle corresponding to the sun. A position vector of the target celestial body relative to the earth is determined based on the radius of the earth, the declination and hour angle corresponding to the target celestial body. A position vector of the spacecraft to be analyzed relative to the earth is calculated based on the orbital height, the orbital inclination, the ascending node right ascension, the perigee amplitude angle and the true perigee angle of the spacecraft to be analyzed relative to the earth. The position distribution among the earth, the sun and the target celestial body is determined based on the position vector of the sun relative to the earth, the position vector of the target celestial body relative to the earth and the position vector of the spacecraft to be analyzed relative to the earth. The position distribution among the sun model, the celestial body model and the spacecraft in-orbit model is determined based on the position distribution among the earth, the sun and the target celestial body.

[0041] Further, if the product of the first vector and the second vector corresponding to any sailboard three-dimensional model is less than or equal to 0, it is determined that the target included angle corresponding to any sailboard three-dimensional model is greater than or equal to 90°.

[0042] If the product of the first vector and the second vector corresponding to any sailboard three-dimensional model is greater than 0, it is determined that the target included angle corresponding to any sailboard three-dimensional model is less than 90°.

[0043] Further, the spacecraft energy model performs a preset task based on a preset frequency.

[0044] The beneficial effects of the present application are:

[0045] The present application considers that the spacecraft sailboard has multiple configurations and may have failure or damage at the initial modeling, and can simulate the attitude of each part of the solar sailboard in any state. Then, by constructing the spacecraft in-orbit model, the shielding of the spacecraft by any central celestial body is analyzed, which is suitable for near-earth spacecraft and deep space spacecraft analysis. The energy system power generation state of the spacecraft at a certain time can be calculated, and the continuous power generation simulation analysis of the spacecraft energy system can be realized by inputting the time sequence data of the real orbit, attitude and sailboard attitude of the spacecraft. After combining the high-precision battery, power supply and load model, the spacecraft energy balance analysis can be performed.

[0046] The whole process is to calculate the energy system power generation state of the spacecraft at a certain time, and by inputting the time sequence data of the real orbit, attitude and sailboard attitude of the spacecraft, the continuous power generation simulation analysis of the spacecraft energy system can be realized. After combining the high-precision battery, power supply and load model, the spacecraft energy balance analysis can be performed.

[0047] In another aspect, the present application provides an electronic device, comprising: a memory, one or more processors; the memory and the processor are coupled; wherein the memory stores computer program codes, the computer program codes comprise computer instructions, when the computer instructions are executed by the processor, the electronic device executes the method of any one of the above first aspect.

[0048] It can be understood that the beneficial effects of the electronic device provided above can refer to the beneficial effects of the first aspect and any possible design thereof, which will not be repeated here. BRIEF DESCRIPTION OF DRAWINGS

[0049] Figure 1 A flowchart of a spacecraft energy model construction method supporting a multi-degree-of-freedom moving part provided by the present application is shown in the figure;

[0050] Figure 2 A structural diagram of a spacecraft model provided by the present application is shown in the figure;

[0051] Figure 3 A flowchart of a preset task provided by the present application is shown in the figure;

[0052] Figure 4 A schematic diagram of the central celestial body umbra and penumbra provided by the present application is shown in the figure;

[0053] Figure 5 A schematic diagram of the spacecraft to be analyzed in the shadow cone provided by the present application is shown in the figure;

[0054] Figure 6 A schematic diagram of the target included angle provided by the present application is shown in the figure;

[0055] Figure 7 A schematic diagram of the shadow area provided by the present application is shown in the figure;

[0056] Figure 8 A statistical diagram of the effective power generation area included in the shadow area provided by the present application is shown in the figure;

[0057] Figure 9 A schematic diagram of the inertial system attitude quaternion curve corresponding to the spacecraft to be analyzed provided by the present application is shown in the figure;

[0058] Figure 10 A schematic diagram of the effective power generation area ratio curve of the spacecraft to be analyzed provided by the present application is shown in the figure;

[0059] Figure 11 A schematic diagram of the power generation curve of the spacecraft to be analyzed provided by the present application is shown in the figure;

[0060] Figure 12 A schematic diagram of the spacecraft total power consumption curve to be analyzed provided by the present application is shown in the figure;

[0061] Figure 13 A schematic diagram of a battery power curve of a spacecraft to be analyzed is provided. DETAILED DESCRIPTION

[0062] The technical solutions in the embodiments of the present application will be described below with reference to the drawings in the embodiments of the present application. In the description of the present application, unless otherwise specified, “ / ” represents a “or” relationship of the objects before and after the “ / ”, for example, A / B can represent A or B; “and / or” in the present application is only a description of the relationship between the associated objects, which means that there can be three relationships, for example, A and / or B, which can represent the following three cases: A exists alone, A and B exist together, and B exists alone. In addition, in the description of the present application, unless otherwise specified, “multiple” means two or more than two. “At least one of the following” or similar expressions means any combination of the items, including any combination of single item or multiple items. For example, at least one of a, b, or c can represent a, b, c, a-b, a-c, b-c, or a-b-c, where a, b, and c can be single or multiple. In addition, in order to clearly describe the technical solutions in the embodiments of the present application, in the embodiments of the present application, “first”, “second”, and the like are used to distinguish the same items or similar items with basically the same function and role. Those skilled in the art can understand that “first”, “second”, and the like do not limit the quantity and execution order, and “first”, “second”, and the like do not necessarily mean different. At the same time, in the embodiments of the present application, “exemplary” or “for example” means to serve as an example, illustration or description. Any embodiment or design scheme described as “exemplary” or “for example” in the embodiments of the present application should not be interpreted as more preferred or more advantageous than other embodiments or design schemes.

[0063] The present application has the following assumptions:

[0064] (1) When calculating the shielding of the sailboard by the shield, the sunlight is parallel light;

[0065] (2) When calculating the size of the central celestial body's umbra, both the sun and the central celestial body are standard spherical celestial bodies;

[0066] (3) All shielding objects are not transparent, i.e. sunlight cannot penetrate the shielding objects;

[0067] (4) The sailboard is a plane.

[0068] The orbit elements, spacecraft attitude, sun, earth and central celestial body positions of the present application are all in the geocentric J2000 coordinate system.

[0069] Reference Figure 1The application provides a spacecraft energy model construction method supporting a multi-degree-of-freedom movable component, comprising the following steps S101-S108:

[0070] S101: determining a spacecraft to be analyzed.

[0071] The spacecraft to be analyzed comprises a sailboard structure and a non-sailboard structure, and the sailboard structure comprises a plurality of sailboard assemblies and a plurality of sailboard connecting components.

[0072] S102: constructing a three-dimensional model of the body corresponding to the spacecraft to be analyzed based on the non-sailboard structure.

[0073] In some embodiments, the three-dimensional model of the body corresponding to the spacecraft to be analyzed can be constructed in actual size by using 3dsmax software.

[0074] According to the principle of minimizing modeling, the three-dimensional model of the body is constructed to be limited to the external structure of the body of the spacecraft to be analyzed, which can cause shielding to the sailboard, such as a communication antenna, a star sensor, a rotating joint, a mechanical arm and the like, and the model only needs to construct an envelope of the external structure, and the connecting screws between components can be omitted in the modeling process, so as to reduce the model volume as much as possible.

[0075] S103: constructing a three-dimensional model of any sailboard corresponding to any sailboard assembly in the plurality of sailboard assemblies.

[0076] S104: combining the three-dimensional model of the body and the three-dimensional models of the sailboards based on the position layout of the sailboard structure and the non-sailboard structure.

[0077] In some embodiments, the connection mode between the first sailboard three-dimensional model and the three-dimensional model of the body is a three-degree-of-freedom movable connection. The first sailboard three-dimensional model is a sailboard three-dimensional model directly connected to the three-dimensional model of the body among the three-dimensional models of the sailboards. The connection mode between the second sailboard three-dimensional model and the first sailboard three-dimensional model is a one-degree-of-freedom movable connection. The second sailboard three-dimensional model is a sailboard three-dimensional model directly or indirectly connected to the first sailboard three-dimensional model among the three-dimensional models of the sailboards.

[0078] Taking a common spacecraft and a common solar sailboard as examples, the sailboards are distributed in the +Y and -Y directions, a single-sided sailboard configuration can be composed of a plurality of sailboard assemblies, connected by hinges and folded up; after being launched into orbit, the single-sided sailboard is unfolded and returned to the 0 position, and the position is taken as the initial position of the solar sailboard.

[0079] Taking a +Y single-sided sailboard as an example, the sailboard is disconnected at the connection with the body, and the sailboard is disconnected at the folding position, and n sailboard assemblies can be obtained by disassembling the single-sided solar sailboard.

[0080] The sailboard connected to the body after disassembly is a proximal sailboard assembly, and a coordinate system of the proximal sailboard assembly is set as follows: +y1 -X+ y1 Y +y1 Z +y1 (subscript +y1 indicates the first component of the +Y side sail), wherein the coordinate origin O +y1 is the connection point of the sail and the body, X +y1 , Y +y1 , Z +y1 The coordinate axis direction is consistent with the X, Y, Z coordinate axis direction of the spacecraft body.

[0081] If the sail is a foldable multi-section sail, the sail around the hinge folding point is a distal sail component, and the coordinate system of the distal sail component is set as follows: O +y2 -X +y2 Y +y2 Z +y2 (subscript +y2 indicates the second component of the +Y side sail), wherein the coordinate origin O +y2 is the midpoint of the sail folding hinge, X +y2 , Y +y2 , Z +y2 The coordinate axis direction is consistent with the X, Y, Z coordinate axis direction of the spacecraft body.

[0082] The setting method of the distal sail component coordinate system is repeated until the coordinate system of all sail components of the single-sided sail is set.

[0083] The rotational freedom and connection topology of the single-sided sail are set.

[0084] The proximal sail component is connected to the body system, and can be set to rotate around the X +y1 , Y +y1 , Z +y1 axis with O +y1 as the center.

[0085] The distal sail component is connected to the proximal sail component or other distal sail components, and can be set to rotate around the X +y2 axis with O +y2 as the center. In abnormal conditions, a multi-degree-of-freedom model of the distal sail component can also be established according to the situation.

[0086] The sail connection topology is set to confirm the sail deployment sequence, and the relative proximal sail has no effect on the attitude motion of the distal sail. When the relative proximal sail attitude moves, it can drive the entire distal sail to change attitude. The single-sided sail is a model set, and when the spacecraft body attitude moves, it drives the entire single-sided sail to change attitude.

[0087] According to the single-sided sail modeling method, the modeling of the other side sail is completed.

[0088] Among them, refer to Figure 2In (a) of FIG. 3, the area framed by the blue line on the right side of the body three-dimensional model represents the sailboard three-dimensional model on the +Y proximal side. See (b) of FIG. 3. Figure 2 In (c) of FIG. 3, the area framed by the blue line on the right side of the body three-dimensional model represents the sailboard three-dimensional model on the +Y distal side. See (d) of FIG. 3. Figure 2 In (d) of FIG. 3, the area framed by the blue line on the left side of the body three-dimensional model represents the sailboard three-dimensional model on the -Y proximal side. See (e) of FIG. 3. Figure 2 In (e) of FIG. 3, the area framed by the blue line on the left side of the body three-dimensional model represents the sailboard three-dimensional model on the -Y distal side. Figure 2

[0089] S105: For any sailboard three-dimensional model, based on the position of the solar cell array patch on the sailboard assembly corresponding to the sailboard three-dimensional model, a solar cell array patch model is set on the sailboard three-dimensional model.

[0090] In some embodiments, any sailboard three-dimensional model comprises a first pasting surface and a second pasting surface, and the solar cell array patch model is set on the first pasting surface comprised by the sailboard three-dimensional model. That is, the solar cell array patch model is attached to only one side of each sailboard three-dimensional model, as the power generation surface of the sailboard three-dimensional model, to simulate the case that the solar sailboard can only generate power on one side.

[0091] S106: The rotation angle of each sailboard three-dimensional model relative to the body three-dimensional model is set, to obtain a spacecraft model.

[0092] S107: The orbit information, attitude information and epoch time information of the spacecraft to be analyzed are input to the spacecraft model, to construct the on-orbit spacecraft corresponding to the spacecraft to be analyzed.

[0093] The orbit information of the spacecraft to be analyzed at least comprises spacecraft orbit elements.

[0094] In some embodiments, before step S107 is performed, the following step can also be performed:

[0095] The spacecraft orbit elements are obtained, and the spacecraft orbit position is set.

[0096] In the entire process of sailboard shielding calculation, Kepler orbit elements (semi-major axis a, eccentricity e, orbital inclination i, ascending node right ascension Ω, perihelion amplitude ω0 and true anomaly ) are required to participate in calculation. Therefore, if the obtained orbit elements are Cartesian orbit elements (position vector and velocity vector ), they need to be converted into Kepler orbit elements.

[0097] ​The conversion steps are as follows:

[0098] (1) Calculate the angular momentum vector and the orbital energy E

[0099] Angular momentum vector: The components of the angular momentum vector are obtained as h x = yv z -zv x , h y = zv x -xv z , h z = xv y -yv x .

[0100] Orbital energy where μ is the gravitational constant of the central gravitational body (for example, for the Earth, μ = GM, where G is the gravitational constant and M is the mass of the Earth).

[0101] (2) Calculate the semi-major axis a

[0102] The energy equation can be converted to from which it follows that

[0103] (3) Calculate the eccentricity e

[0104] First, calculate the eccentricity vector The components of which can be calculated from the components of and .

[0105] Eccentricity

[0106] (4) Calculate the orbital inclination i

[0107] where

[0108] (5) Calculate the ascending node right ascension Ω

[0109] First, calculate the ascending node vector ( is the unit vector of the z-axis), from which the components of the ascending node vector are obtained as n x = h y , n y = -h x , n z = 0.

[0110] Ascending node right ascension when n y ≥ 0 or when n y<0), wherein

[0111] (6) Calculate the argument of perigee ω0

[0112] (when e z ≥ 0) or (when e z < 0).

[0113] (7) Calculate the true anomaly

[0114] (when ) or (when ).

[0115] Obtain the spacecraft attitude parameters, and set the spacecraft attitude.

[0116] To avoid the singularity problem and the restrictions brought by the rotation sequence, the spacecraft attitude parameters are in the form of quaternions: q = (q0, q1, q2, q3), wherein q0 is the real part, and q1, q2, and q3 are the imaginary parts.

[0117] Obtain the spacecraft epoch time, and set the reference as UTC time.

[0118] S108: Construct a spacecraft energy model for the on-orbit model of the spacecraft.

[0119] In some embodiments, when setting the rotation angle of each sailboard three-dimensional model relative to the body three-dimensional model, the position of each sailboard three-dimensional model relative to the body three-dimensional model can be calculated according to the topological relationship among the body three-dimensional model, the first sailboard three-dimensional model, and the second sailboard three-dimensional model.

[0120] Specifically, C++ can be used for secondary development of 3dsmax, and an interface can be written to set the angle of the sailboard assembly according to the previously set model degrees of freedom. The sailboard assembly rotation angle setting interface can be used to set the attitude of the sailboard assembly. When the sailboard is in an abnormal state, the actual attitude data of the sailboard is used to set the rotation angle of each sailboard assembly through the sailboard assembly rotation angle setting interface to simulate the real attitude of the spacecraft sailboard in orbit. When the sailboard is in a normal state, the actual SADA motor rotation angle data is used to drive the sailboard attitude, and the sailboard assembly rotation angle setting data does not work.

[0121] In some embodiments, the spacecraft energy model is used to perform a preset task, and the preset task can be performed based on a preset frequency.

[0122] Referring to Figure 3 , the preset task includes the following steps S301-S316.

[0123] S301: Determine whether the celestial body model exists occlusion to the spacecraft on-orbit model based on the position distribution between the sun model corresponding to the sun, the celestial body model corresponding to the target celestial body and the spacecraft on-orbit model.

[0124] In some embodiments, the declination and hour angle of the sun corresponding to the earth are obtained, and the declination and hour angle of the target celestial body corresponding to the earth are obtained. The position vector of the sun relative to the earth is determined based on the radius of the earth, the declination and hour angle of the sun corresponding to the earth. The position vector of the target celestial body relative to the earth is determined based on the radius of the earth, the declination and hour angle of the target celestial body corresponding to the earth. The position vector of the spacecraft to be analyzed relative to the earth is calculated based on the orbital height, the orbital inclination, the ascending node right ascension, the perigee amplitude and the true perigee angle of the spacecraft to be analyzed relative to the earth. The position distribution among the earth, the sun and the target celestial body is determined based on the position vector of the sun relative to the earth, the position vector of the target celestial body relative to the earth, and the position vector of the spacecraft to be analyzed relative to the earth. The position distribution between the sun model, the celestial body model and the spacecraft on-orbit model is determined based on the position distribution among the earth, the sun and the target celestial body.

[0125] Specifically, only the earth is taken as the central celestial body as an example here, when other celestial bodies are taken as the central celestial body, only the position vector of the earth pointing to the other celestial body is obtained, the vector of the other celestial body pointing to the sun and the vector of the other celestial body pointing to the spacecraft are converted, and the same calculation method of the spacecraft-earth-sun spatial relationship is used for calculation.

[0126] The sun vector is calculated, assuming that the sun declination is δ, the hour angle is ω, and the earth radius is R E The coordinates of the earth pointing to the sun can be expressed as:

[0127] Among them, S x , S y , S z can be expressed as follows respectively:

[0128] S x = R E cosδcosω

[0129] S y = R E cossinω

[0130] S z = R E sinδ

[0131] The spacecraft vector is calculated, assuming that the spacecraft orbital height is H, the orbital inclination i, the ascending node right ascension Ω, the perigee amplitude ω0 and the true perigee angle The coordinates of the satellite position vector can be expressed as:

[0132] wherein r x , r y , r z may be respectively expressed as follows:

[0133]

[0134] Calculate the angle between sunlight and the position vector of the spacecraft:

[0135] By judging the angle θ, the spatial relationship between the spacecraft, the central celestial body and the sun can be determined; whether the spacecraft is in the shadow area also needs to be combined with the size of the shadow area.

[0136] When θ≤90°, the spacecraft is in the sunlit area; when θ>90°, it needs to be combined with the spacecraft position and the calculation result of the central celestial body shadow area to confirm whether the spacecraft is blocked.

[0137] S302: In the case where the celestial body model does not block the spacecraft on-orbit model, for any sailboard three-dimensional model, calculate the target angle corresponding to any sailboard three-dimensional model.

[0138] Wherein the target angle is the angle between the first vector and the second vector. The first vector is the pointing vector from the position of any sailboard three-dimensional model to the position of the sun model. The second vector is the normal vector corresponding to the solar cell array patch model set on any sailboard three-dimensional model.

[0139] In some embodiments, in the case where the celestial body model blocks the spacecraft on-orbit model, the effective power generation area of the spacecraft on-orbit model is determined to be 0, and the power generation of the spacecraft on-orbit model is determined to be 0.

[0140] In some embodiments, the model is not only suitable for near-earth spacecraft, but also suitable for shadow calculation of deep space exploration spacecraft, so it is necessary to calculate the shadow generated by the target celestial body (also referred to as the central celestial body in the embodiments of the present application). The central celestial body generates a shadow including the umbra and the penumbra, as shown in Figure 4 .

[0141] In the model, only whether the spacecraft enters the umbra is considered, the spacecraft can still receive sunlight in the penumbra, but the radiation intensity will be reduced, which needs to be considered separately when calculating the power generation of the solar sailboard.

[0142] Calculate the radius of the umbra of the central celestial body:

[0143] According to the principle of similar triangles, the radius of the umbra is wherein R umbra is the radius of the penumbra, and R ER is the radius of the Earth (non-geocentric, the radius of the central celestial body), R S ρ is the radius of the Sun, ρ S is the distance between the Earth and the Sun (non-geocentric, the distance between the Sun and the central celestial body).

[0144] Here, R E Considering the atmospheric attenuation effect, the value is 6398km (other central celestial bodies do not consider the atmospheric attenuation effect, and the value is the radius size); R S The value is 696000km; ρ S The value is 1.496×10 8 km, R umbra is calculated to be 1387961.172km.

[0145] The half-cone angle of the umbra region is calculated as Substituting R E and R umbra , σ is 0.264°

[0146] Combined with the angle θ between the sunlight and the spacecraft position vector, it can be known that when θ≤90°+σ, the spacecraft is in the sunlight area; when θ>90°+σ, it needs to be combined with the spacecraft position and the calculation result of the central celestial body shadow area conical surface to confirm whether the spacecraft is blocked.

[0147] In the shadow conical coordinate system, the central celestial body shadow area conical surface can be expressed as:

[0148] F(x s h,y s h,z s h)=y s h 2 +z s h 2 -(R umbra -x s h) 2 tan 2 σ=0 (Formula 1).

[0149] The parameter expression of the spacecraft in the orbit plane coordinate system can be expressed as:

[0150]

[0151] Where r=R E +H, representing the distance from the spacecraft to the center of the central celestial body; is the true anomaly.

[0152] The conical surface expression needs to be converted from the shadow conical coordinate system to the orbit plane coordinate system through a series of unit conversion matrix multiplication transformation, and the transformed vector can be expressed as:

[0153] r0=Az (ω0)A x (i)A z (Ω)A x (-ε)A z (π-L)r s h。

[0154] wherein, A n (α) represents a rotation matrix rotating α angle around the coordinate axis n, ω0 is the argument of perigee, Ω is the right ascension of the ascending node, i is the orbital inclination, ε is the ecliptic angle, and L is the solar right ascension.

[0155] Since the intersection of the spacecraft orbit plane and the shadow cone surface has z0=0, the transformed cone surface vector component can be expressed as:

[0156]

[0157] After unifying the coordinate system, equations 1, 2, and 3 can be solved to determine whether the spacecraft is in the shadow cone, as shown in equation 4. Figure 5

[0158] In some embodiments, the vector of all sail panels of the spacecraft pointing to the sun can be equivalent to the vector of the spacecraft pointing to the sun

[0159] The nth block of solar sail panel assembly power generation plane to be analyzed can be expressed as a n x+b n y+c n z+d n =0, (a n ,b n ,c n ,d n are constants), and the sail panel assembly plane normal vector can be expressed as

[0160] The angle between the sunlight and the sail panel assembly power generation plane normal vector (target angle) can be calculated, as shown in equation 6. Figure 6

[0161]

[0162] When ∈<90°, the sail panel assembly can be irradiated by sunlight, and when ∈≥90°, the sail panel assembly cannot be irradiated by sunlight.

[0163] After the spacecraft energy model performs S302, based on the size of the target angle corresponding to any sail panel three-dimensional model, steps S303 or S304 are continued.

[0164] ​​S303: In a case where the target included angle corresponding to any sailboard three-dimensional model is greater than or equal to 90°, it is determined that the number of power generation units of any sailboard three-dimensional model is 0.

[0165] In some embodiments, in a case where the product of the first vector and the second vector corresponding to any sailboard three-dimensional model is less than or equal to 0, it is determined that the target included angle corresponding to any sailboard three-dimensional model is greater than or equal to 90°.

[0166] After the spacecraft energy model performs step S303, it continues to perform steps S310-S316.

[0167] S304: In a case where the target included angle corresponding to any sailboard three-dimensional model is less than 90°, it is determined that any sailboard three-dimensional model is assigned a value of 1, and the body three-dimensional model and the target sailboard three-dimensional model are both assigned a value of 0.

[0168] The target sailboard three-dimensional model is a sailboard three-dimensional model that is different from any sailboard three-dimensional model.

[0169] In some embodiments, in a case where the product of the first vector and the second vector corresponding to any sailboard three-dimensional model is greater than 0, it is determined that the target included angle corresponding to any sailboard three-dimensional model is less than 90°.

[0170] After the spacecraft energy model performs step S304, it continues to perform steps S305-S316.

[0171] S305: Determine the shadow area on the solar cell array patch model corresponding to any sailboard three-dimensional model.

[0172] The shadow area is formed based on the structure of any sailboard three-dimensional model and / or the blocking of any sailboard three-dimensional model by each target sailboard three-dimensional model.

[0173] Specifically, when calculating the shadow area, the sunlight can be regarded as parallel light, so the coordinate components of the sunlight can be represented as (l x ,l y ,l z ), and the surface coordinates of the spacecraft body and the solar sailboard three-dimensional model can be represented as (M x ,M y ,M z ), and the coordinates of the projection point in the a n x+b n y+c n z+d n =0 plane can be represented as (P x ,P y ,P z ).

[0174] The specific coordinates can be represented as:

[0175]

[0176] The process of obtaining the shadow area of the three-dimensional model uses CUDA to call the GPU processor to perform operations to speed up the operation, as shown in Figure 7 .

[0177] The shadow area of the occluder is represented as:

[0178] C = (C x , C y , C z , 0) (Formula 5).

[0179] The shadow area of the sailboard assembly to be analyzed is represented as:

[0180] B = (B x , B y , B z , 1) (Formula 6).

[0181] (C x , C y , C z ) and (B x , B y , B z ) are calculated by Formula 4.

[0182] S306: Based on the region segmentation rule, the surface of the solar cell array patch model corresponding to any sailboard three-dimensional model is segmented into multiple sub-regions.

[0183] In some embodiments, the shadow area can be rasterized. In order to balance the accuracy and efficiency of the occlusion calculation, the shadow area of the occluder and the sailboard assembly to be analyzed can be segmented at a grid size of 1 cm 2 The shadow area of the occluder is represented by Formula 5, and the shadow area of the sailboard assembly to be analyzed is represented by Formula 6.

[0184] The occluder and sailboard assembly to be analyzed are obtained respectively containing j and k elements, and the projection point data queue is recorded as: C all = (C1, C2, …, C j ) and B all = (B1, B2, …, B k ).

[0185] The Euclidean distance of the data set (C1, C2, …, C j ) and (B1, B2, …, B k ) is calculated.

[0186] Convert D to cm units.

[0187] The size of a single image grid cell is 1 cm*1 cm, and the radius of the circumscribed circle is taken as as the judgment threshold.

[0188] When , it indicates that the area of the sailboard assembly to be analyzed is blocked by an occluder, and the grid data is (B xk , B yk , B zk , 0), and is removed from the projection point data queue of the sailboard assembly to be analyzed, and does not participate in the next iteration calculation to reduce the calculation amount.

[0189] Loop calculation until the judgment of all projection points is completed, at which time the number of elements remaining in the projection point data queue of the sailboard assembly to be analyzed is the unblocked area, and the removed elements are the blocked area.

[0190] S307: Determine a first sub-region and a second sub-region in a plurality of sub-regions.

[0191] Among them, the first sub-region is the area coinciding with the shadow area, and the second sub-region is the area not coinciding with the shadow area.

[0192] S308: Determine the sum of the areas of each second sub-region corresponding to any sailboard three-dimensional model as the effective power generation area of any sailboard three-dimensional model.

[0193] Specifically, the number of elements remaining in the projection point data queue of the sailboard assembly to be analyzed can be counted, at which time all elements in the queue represent (B x , B y , B z , 1), and divided by 10000, the unit is converted to square meters, to obtain the effective power generation area A N of the sailboard three-dimensional model. As shown in Figure 8 , the area of the non-occluded / shadow area is the effective power generation area.

[0194] S309: Determine the power generation of any sailboard three-dimensional model based on the effective power generation area of any sailboard three-dimensional model, the solar radiation of the position where any sailboard three-dimensional model is located by the solar model, the angle between any sailboard three-dimensional model and the sunlight emitted by the solar model, the working temperature of the sailboard assembly corresponding to any sailboard three-dimensional model, and the aging rate of the sailboard assembly corresponding to any sailboard three-dimensional model.

[0195] In some embodiments, the power generation of each sailboard three-dimensional model can be calculated based on the formula P=η×A N ×I×cos∈×(1-α×ΔT)×C.

[0196] Among them, η is the photoelectric conversion efficiency of the solar sailboard, and η takes the value of the sailboard [0, 1];

[0197] A N is the effective area of the corresponding sail three-dimensional model, in square meters (m 2 ).

[0198] I is the irradiance of the spacecraft to be analyzed from the sun, I0 takes the solar constant, indicating the solar irradiance outside the earth's atmosphere, which is 1361 W / m 2 , r is the distance from the spacecraft to be analyzed to the center of the sun (in astronomical units AU, 1 AU is the average distance from the earth to the sun, taking 1.496 x 10 8 km).

[0199] cos∈ is the correction factor of the effective irradiance, and ∈ is the angle between the sunlight and the normal vector of the sail assembly power generation surface. When ∈ = 0°, the sunlight is perpendicular to the sail, and the correction factor is equal to 1. As ∈ increases, the correction factor decreases, and the effective irradiance received by the sail decreases. Here, since the case of ∈ ≥ 90° is excluded in advance, the value range of cos∈ is (0, 1].

[0200] 1-α×ΔT is the working temperature correction term, and α represents the power temperature coefficient, with a unit of 1 / ℃, and different sail materials have different values; ΔT = T-T ref , T is the actual working temperature, and T ref is the reference temperature, both in ℃, and T ref usually takes a value of 25℃.

[0201] C is the sail power generation correction factor considering the aging and damage factors of the sail, which can be obtained according to the actual on-orbit monitoring data or ground test data, or calculated according to the sail aging model, and the value range is [0, 1].

[0202] S310: Determine the sum of the effective power generation areas of all sail three-dimensional models as the effective power generation area of the spacecraft on-orbit model, and determine the sum of the power generation powers of all sail three-dimensional models as the power generation power of the spacecraft on-orbit model.

[0203] Exemplarily, the sum of the effective power generation areas of all sail three-dimensional models and the sum of the power generation powers obtain the energy system power generation state (effective power generation area and power generation power) of the spacecraft on-orbit model at the moment, that is, obtain the energy system power generation state of the spacecraft to be analyzed at the moment.

[0204] S311: Obtain the power consumption of each load device included in the spacecraft to be analyzed.

[0205] Wherein, for any one of the load devices included in the spacecraft to be analyzed, the power consumption of the any one load device is determined based on the working state of the any one load device. The working state can include a power-on state and a power-off state. In addition, the working state can also include a working mode and a time-varying parameter related to power consumption.

[0206] Specifically, the on-off state of each load device can be obtained from a spacecraft model (an external model, not belonging to the energy model of the spacecraft) or receive the on-off instruction of each load device, and the power consumption and current of each load device during working are calculated according to the working mode of each load device, the corresponding power consumption table, and the time-varying parameter related to power consumption.

[0207] S312: Obtain the power consumption of each power distribution panel and the power consumption of each bus included in the spacecraft to be analyzed.

[0208] S313: Determine the total power consumption of the spacecraft to be analyzed based on the power consumption of each load device, the power consumption of each power distribution panel, and the power consumption of each bus.

[0209] Specifically, according to the corresponding relationship between the power distribution panel, the bus and each load unit, the current and power consumption corresponding to each power distribution panel and bus can be calculated by using the power supply and distribution model, while considering the loss of each power distribution panel and bus itself during power supply and distribution, and all data are summarized to obtain the total power consumption of the spacecraft.

[0210] S314: Obtain the working state information of the battery included in the spacecraft to be analyzed.

[0211] Wherein, the working state information of the battery includes temperature information, voltage information and power information of the battery.

[0212] S315: Determine the power supply mode of the spacecraft to be analyzed based on the effective power generation area, the power generation power, the total power consumption of the spacecraft to be analyzed, and the working state information of the battery.

[0213] Wherein, the effective power generation area of the spacecraft to be analyzed is determined based on the effective power generation area of the spacecraft on-orbit model, and the power generation power of the spacecraft to be analyzed is determined based on the power generation power of the spacecraft on-orbit model. The power supply mode of the spacecraft to be analyzed includes a battery power supply mode, a sailboard assembly power supply mode and a joint power supply mode. The joint power supply mode is a mode in which the battery and the sailboard assembly supply power at the same time.

[0214] And when the spacecraft is blocked by the sailboard central celestial body (i.e. in the central celestial body umbra region), the power supply mode of the spacecraft to be analyzed is the battery power supply mode;

[0215] When the sailboard power generation power (the sum of the power supply array and the charging array power) is less than the total power consumption (which can also be referred to as the total power consumption of the spacecraft in the embodiments of the present application), the power supply mode of the spacecraft to be analyzed is the joint power supply mode;

[0216] When the sailboard power generation power is greater than or equal to the total power, and the battery is full, the power supply mode of the spacecraft to be analyzed is the sailboard assembly power supply mode;

[0217] When the sailboard power generation power is greater than or equal to the total power, and the battery is not full, the power supply mode of the spacecraft to be analyzed is the sailboard assembly power supply mode, and the charging array is the battery power supply.

[0218] S316: Based on the power supply mode of the spacecraft to be analyzed, energy balance analysis is performed on the spacecraft to be analyzed.

[0219] The energy balance analysis includes:

[0220] In the case where it is determined that the power supply mode of the spacecraft to be analyzed is the battery power supply mode or the joint power supply mode, the discharge depth and discharge voltage of the battery are obtained. In the case where it is detected that the discharge depth of the battery is greater than the discharge depth threshold, and / or the discharge voltage of the battery is out of the preset discharge voltage range, a prompt information is generated. The prompt information is used to prompt the power supply anomaly of the spacecraft to be analyzed.

[0221] In the case where it is determined that the power supply mode of the spacecraft to be analyzed is the battery power supply mode or the joint power supply mode, the battery is in a discharging state at this time, and the discharge current, battery voltage, battery capacity and discharge depth of the battery can be calculated according to the charge-discharge curve of the battery, the power generation state and the total spacecraft power consumption.

[0222] In the case where it is determined that the power supply mode of the spacecraft to be analyzed is the sailboard assembly power supply mode, if the total power consumption of the spacecraft to be analyzed is greater than the power consumption threshold, a prompt information is generated.

[0223] In the case where it is determined that the power supply mode of the spacecraft to be analyzed is the sailboard assembly power supply mode, since the battery is in a full state at this time, only the battery capacity and voltage parameters at the current time are obtained, and the discharge depth is 0 at this time.

[0224] In the case where it is determined that the power supply mode of the spacecraft to be analyzed is the sailboard assembly power supply mode, and the battery is in a charging state, if the charging current of the battery is greater than the charging current threshold, an information for prompting the abnormal charging state of the battery is generated.

[0225] In the case where it is determined that the power supply mode of the spacecraft to be analyzed is the sailboard assembly power supply mode, and the battery is in a charging state, the charging current, battery voltage, battery capacity and discharge depth under the current temperature and capacity can be calculated in combination with the battery charge-discharge curve, the total spacecraft power consumption, the discharge array power generation power, the charging array power generation power and the battery state.

[0226] The following is a specific embodiment:

[0227] Assume that a deep space exploration satellite performs orbit control task in the condition of damaged solar panels. The rotation angle of +Y near-end solar panel is set as (15°, 45°, 30°), the rotation angle of +Y far-end solar panel is set as 30°; the rotation angle of -Y near-end solar panel is set as (-15°, 10°, -30°), and the rotation angle of +Y far-end solar panel is set as 30°.

[0228] Input the initial orbit elements of the spacecraft, and based on the constructed on-orbit model of the spacecraft and the simulation task action sequence, the attitude angle (such as shown in Figure 9 ) and the effective power generation area proportion (such as shown in Figure 10 ), power generation (such as shown in Figure 11 ), whole spacecraft power consumption (such as shown in Figure 12 ), and battery power (such as shown in Figure 13 ) of the two side solar panels in the whole process of the spacecraft performing the corresponding orbit control task are obtained.

[0229] In this example, the scenario of deep space spacecraft solar panel being damaged and locked is simulated, and the attitude of the solar panel does not change in the whole process, and the relative relationship with the sun only changes with the orbit position and the attitude of the spacecraft. As can be seen from Figure 9 and Figure 10 , the inflection point positions of the effective power generation area proportion curves of the two side solar panels are highly consistent with the positions of the attitude fluctuation of the spacecraft, which is consistent with the theory.

[0230] On the basis of confirming the effective power generation area and power generation of the two side solar panels, the energy storage battery model, power supply and distribution model, and spacecraft load model are jointly simulated, and finally the energy balance modeling and simulation analysis of the spacecraft in any state (such as shown in Figures 11 to 13 ) are realized.

[0231] By analogy with this method, it can be applied to the modeling of other spacecraft energy systems, and a high-precision three-dimensional model of the spacecraft and an on-orbit attitude and orbit model need to be established during modeling. If the whole process of the spacecraft on-orbit task is to be analyzed, a high-precision orbit dynamics model, attitude dynamics model, and star service model also need to be used.

[0232] In some aspects, multiple embodiments of the present application can be combined, and the combined aspects can be implemented. Optionally, some operations in the flow of various method embodiments are combined, and / or the order of some operations is changed. Also, the order of execution or performance of the operations in the flows between various steps is not limited to the order described. Numerous representations of the operations between the steps are permitted. The execution or performance of the operations in the various steps can not be implemented in the exact order as shown, and that the specific order or hierarchy of steps in the flows are represented only as an example. The use of the term "processing" does not necessarily require that the steps be performed in the order described. Various actions depicted can occur at different times and need not occur at the same time. Also, various actions can occur concurrently, not necessarily in the order shown. Furthermore, various activities performed can be repeated, not only once, but can also be repeated an unspecified number of times. The descriptions herein are not intended to be interpreted as a limitation to the steps illustrated. The steps illustrated can be combined or divided. Various embodiments can be used to implement the steps in the flows.

[0233] In addition, some of the steps in the method embodiments can be replaced by other possible steps. Or, some of the steps in the method embodiments can be optional, and can be deleted in some use scenarios. Or, other possible steps can be added to the method embodiments. Also, the method embodiments can be implemented individually, or in combination.

[0234] The above merely provides a detailed description of the embodiments of the present application, but the protection scope of the present application is not limited thereto. Any changes or replacements within the technical scope disclosed in the present application should be covered within the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

Claims

1. A spacecraft energy model construction method supporting a multi-degree-of-freedom moving component, characterized by, The method comprises the following steps: determining a spacecraft to be analyzed; the spacecraft to be analyzed comprises a sailboard structure and a non-sailboard structure; the sailboard structure comprises a plurality of sailboard assemblies and a plurality of sailboard connecting components; constructing a three-dimensional model of the body corresponding to the spacecraft to be analyzed based on the non-sailboard structure; constructing a three-dimensional model of the sailboard corresponding to any sailboard assembly in the plurality of sailboard assemblies; combining the three-dimensional model of the body and each three-dimensional model of the sailboard based on the position layout of the sailboard structure and the non-sailboard structure; for any three-dimensional model of the sailboard in each three-dimensional model of the sailboard, setting a solar array patch model on the any three-dimensional model of the sailboard based on the position of the solar array patch on the sailboard assembly corresponding to the any three-dimensional model of the sailboard; setting the rotation angle of each three-dimensional model of the sailboard relative to the three-dimensional model of the body to obtain a spacecraft model; inputting the orbit information, attitude information and epoch time information of the spacecraft to be analyzed into the spacecraft model to construct a spacecraft on-orbit model corresponding to the spacecraft to be analyzed; constructing a spacecraft energy model for the spacecraft on-orbit model; the spacecraft energy model is used to perform a preset task, and the preset task comprises: determining whether the celestial body model exists occlusion to the spacecraft on-orbit model based on the position distribution between the sun model corresponding to the sun, the celestial body model corresponding to the target celestial body and the spacecraft on-orbit model; in the case that the celestial body model does not exist occlusion to the spacecraft on-orbit model, for the any three-dimensional model of the sailboard, calculating a target included angle corresponding to the any three-dimensional model of the sailboard; the target included angle is the included angle between a first vector and a second vector; the first vector is a pointing vector from the position of the any three-dimensional model of the sailboard to the position of the sun model; the second vector is a normal vector corresponding to the solar array patch model set on the any three-dimensional model of the sailboard; in the case that the target included angle corresponding to the any three-dimensional model of the sailboard is greater than or equal to 90°, determining that the power generation unit number of the any three-dimensional model of the sailboard is 0; in the case that the target included angle corresponding to the any three-dimensional model of the sailboard is less than 90°, assigning the any three-dimensional model of the sailboard as 1, and assigning the three-dimensional model of the body and a target three-dimensional model of the sailboard as 0; the target three-dimensional model of the sailboard is a three-dimensional model of the sailboard different from the any three-dimensional model of the sailboard in each three-dimensional model of the sailboard; determining a shadow area on the solar array patch model corresponding to the any three-dimensional model of the sailboard; the shadow area is formed based on the structure of the any three-dimensional model of the sailboard and / or the occlusion of each target three-dimensional model of the sailboard to the any three-dimensional model of the sailboard; based on a region segmentation rule, the surface of the solar array patch model corresponding to the any three-dimensional model of the sailboard is segmented into a plurality of sub-regions; determining a first sub-region and a second sub-region in the plurality of sub-regions; the first sub-region is a region coinciding with the shadow area; the second sub-region is a region not coinciding with the shadow area; determine a sum of areas of each of the second sub-regions corresponding to the any sailboard three-dimensional model as an effective power generation area of the any sailboard three-dimensional model; determine a power generation of the any sailboard three-dimensional model based on the effective power generation area of the any sailboard three-dimensional model, a solar radiation degree of a location where the any sailboard three-dimensional model is located, an angle between the any sailboard three-dimensional model and a sun ray emitted by the sun model, a working temperature of a sailboard assembly corresponding to the any sailboard three-dimensional model, and an aging rate of the sailboard assembly corresponding to the any sailboard three-dimensional model; determine a sum of the effective power generation areas of each of the sailboard three-dimensional models as an effective power generation area of the spacecraft on-orbit model, and determine a sum of the power generations of each of the sailboard three-dimensional models as a power generation of the spacecraft on-orbit model; obtain power consumptions of each of load devices included in the spacecraft to be analyzed; for any one of the load devices included in the spacecraft to be analyzed, the power consumption of the any one of the load devices is determined based on a working state of the any one of the load devices; the working state includes a powered-on state and a powered-off state; obtain power consumptions of each of power distribution boards and power consumptions of each of bus bars included in the spacecraft to be analyzed; determine a total power consumption of the spacecraft to be analyzed based on the power consumptions of each of the load devices, the power consumptions of each of the power distribution boards, and the power consumptions of each of the bus bars; obtain working state information of a battery included in the spacecraft to be analyzed; the working state information of the battery includes temperature information, voltage information, and power information of the battery; determine a power supply mode of the spacecraft to be analyzed based on the effective power generation area, the power generation, and the total power consumption of the spacecraft to be analyzed, and the working state information of the battery; the effective power generation area of the spacecraft to be analyzed is determined based on the effective power generation area of the spacecraft on-orbit model, and the power generation of the spacecraft to be analyzed is determined based on the power generation of the spacecraft on-orbit model; the power supply mode of the spacecraft to be analyzed includes a battery power supply mode, a sailboard assembly power supply mode, and a joint power supply mode; the joint power supply mode is a mode in which the battery and the sailboard assembly supply power at the same time; in a case where it is determined that the power supply mode of the spacecraft to be analyzed is the battery power supply mode or the joint power supply mode, obtain a discharge depth and a discharge voltage of the battery; in a case where it is detected that the discharge depth of the battery is greater than a discharge depth threshold and / or the discharge voltage of the battery is out of a preset discharge voltage range, generate prompt information; the prompt information is used to prompt power supply abnormity of the spacecraft to be analyzed; in a case where it is determined that the power supply mode of the spacecraft to be analyzed is the sailboard assembly power supply mode, generate the prompt information if the total power consumption of the spacecraft to be analyzed is greater than a power consumption threshold.

2. The method of claim 1, wherein, before the determination of the power supply mode of the spacecraft to be analyzed based on the effective power generation area, the power generation, and the total power consumption of the spacecraft to be analyzed, and the working state information of the battery, the method further includes: In a case where the celestial body model is occluded by the spacecraft on-orbit model, the effective power generation area of the spacecraft on-orbit model is determined to be 0, and the power generation of the spacecraft on-orbit model is determined to be 0.

3. The method of claim 2, wherein, The connection mode between the first sailboard three-dimensional model and the body three-dimensional model is a three-degree-of-freedom movable connection; the first sailboard three-dimensional model is a sailboard three-dimensional model directly connected to the body three-dimensional model among the sailboard three-dimensional models; the connection mode between the second sailboard three-dimensional model and the first sailboard three-dimensional model is a one-degree-of-freedom movable connection; the second sailboard three-dimensional model is a sailboard three-dimensional model directly or indirectly connected to the first sailboard three-dimensional model among the sailboard three-dimensional models.

4. The method of claim 3, wherein, Further comprising: When setting the rotation angle of each sailboard three-dimensional model relative to the body three-dimensional model, the positions of each sailboard three-dimensional model relative to the body three-dimensional model are calculated according to the topological relationship among the body three-dimensional model, the first sailboard three-dimensional model, and the second sailboard three-dimensional model.

5. The method of claim 4, wherein, After determining the power supply mode of the spacecraft to be analyzed based on the effective power generation area, the power generation, the total power consumption of the spacecraft to be analyzed, and the working state information of the battery, the method further comprises: In a case where the power supply mode of the spacecraft to be analyzed is the sailboard assembly power supply mode, and the battery is in a charging state, if the charging current of the battery is greater than a charging current threshold, information for prompting an abnormal charging state of the battery is generated.

6. The method of claim 5, wherein, The any sailboard three-dimensional model comprises a first pasting surface and a second pasting surface, and setting a solar cell array patch model on the any sailboard three-dimensional model based on the position of the solar cell array patch on the sailboard assembly corresponding to the any sailboard three-dimensional model comprises: Setting the solar cell array patch model on the first pasting surface comprised in the any sailboard three-dimensional model.

7. The method of claim 6, wherein, Before determining whether the celestial body model is occluded by the spacecraft on-orbit model based on the position distribution among the sun corresponding to the sun model, the target celestial body corresponding to the celestial body model, and the spacecraft on-orbit model, the method further comprises: Obtaining the declination and hour angle of the sun corresponding to the earth as a reference, and obtaining the declination and hour angle of the target celestial body corresponding to the earth as a reference; Determining the position vector of the sun relative to the earth based on the radius of the earth, the declination and hour angle of the sun corresponding to the earth; Determining the position vector of the target celestial body relative to the earth based on the radius of the earth, the declination and hour angle of the target celestial body corresponding to the earth; Calculating the position vector of the spacecraft to be analyzed relative to the earth based on the orbital height, the orbital inclination, the ascending node right ascension, the perigee amplitude, and the true perigee angle of the spacecraft to be analyzed relative to the earth; Determining the position distribution among the earth, the sun, and the target celestial body based on the position vector of the sun relative to the earth, the position vector of the target celestial body relative to the earth, and the position vector of the spacecraft to be analyzed relative to the earth; Determine a position distribution between the sun model, the celestial body model, and the spacecraft on-orbit model based on a position distribution among the Earth, the sun, and the target celestial body.

8. The method of claim 7, wherein, The target angle corresponding to the any sailboard three-dimensional model is calculated in the case that the spacecraft on-orbit model is not blocked by the celestial body model, including: If the product of the first vector and the second vector corresponding to the any sailboard three-dimensional model is less than or equal to 0, it is determined that the target angle corresponding to the any sailboard three-dimensional model is greater than or equal to 90°; If the product of the first vector and the second vector corresponding to the any sailboard three-dimensional model is greater than 0, it is determined that the target angle corresponding to the any sailboard three-dimensional model is less than 90°.

9. The method of claim 8, wherein, The spacecraft energy model performs the preset task based on a preset frequency.

10. An electronic device, comprising: Including: A memory and one or more processors; the memory is coupled with the processor; wherein the memory has computer program code stored therein, the computer program code includes computer instructions, when the computer instructions are executed by the processor, the electronic device executes a spacecraft energy model construction method supporting multi-degree-of-freedom active components according to any one of claims 1-9.

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