Attitude maneuvering imaging task planning method and system based on satellite autonomous energy management

Abstract

The invention provides an attitude maneuvering imaging task planning method and system based on satellite autonomous energy management. The attitude maneuvering imaging task planning method comprises the steps of S1, dividing a maneuvering imaging task hotspot area range; s2, calculating satellite related parameters respectively; s3, further calculating the energy balance state within a certain time according to the calculation result of the satellite related parameters; and S4, when the satellite is in the task hotspot area range of the maneuvering imaging, performing whole-satellite energy early warning according to the energy balance state, so that the satellite autonomously completes the observation task and records the execution state of the task. According to the method, the energy state can be autonomously monitored in real time, the illumination duration of next 24 hours is calculated according to orbit parameters, the energy state of the whole satellite is evaluated, and a plurality of maneuvering observation tasks are autonomously completed; the energy of the whole satellite is guaranteed to the sun through inertia when no observation task exists, the satellite for earth observation is achieved through attitude maneuver when the observation task exists, and good flexibility is achieved.

Description

Technical Field

[0001] The present invention relates to the field of satellite imaging technology, and in particular to a method and system for planning satellite autonomous energy management attitude maneuvering imaging tasks. Background Art

[0002] According to demand, satellite on-orbit observation tasks are increasing, especially for hot spot observation and specific revisited areas. Satellites are required to complete imaging tasks autonomously on-orbit and transmit remote sensing data in a timely manner. Traditional satellites usually need to calculate the area covered by the satellite's sub-satellite point on the ground, issue instructions through the ground remote control station, wait for the satellite to pass through the observation, and the satellite's energy status is reviewed and monitored by the ground. After the task is completed, the next mission observation is carried out through the ground-instructed ground. In the context of high-frequency observation tasks, the observation timeliness and efficiency are not high.

[0003] Patent document CN106324631A discloses a remote sensing satellite energy balance constraint analysis system and method, which uses the discharge depth safety threshold as the energy constraint condition of the mission planning system, but does not introduce the production basis of the imaging mission, does not involve the real-time calculation method of power and the statistical method of illumination duration, and does not involve the energy consumption statistics of the entire satellite and the energy status judgment method.

[0004] Patent document CN106599334B discloses a short-term payload work planning method for improving satellite energy efficiency, introducing the model of energy constraint conditions, modeling of energy constraint conditions and input conditions that affect satellite energy conditions, and battery pack charging and discharging process model. However, it is only for a single work task, and does not consider the impact of long-term energy acquisition conditions and the entire satellite lighting conditions and long-term energy balance.

[0005] Patent document CN104361210A discloses a method for estimating energy of a cube satellite with a body-mounted battery array, which calculates the total energy of the satellite within an orbital period based on the changes in the normal vectors of each face of the cube satellite and whether the satellite is in the shadow of the earth. It is applicable to observation mission satellites without solar panels, but only considers the lighting conditions and does not involve attitude maneuvers.

[0006] Patent document CN107579587B discloses an energy system and control method suitable for LEO satellites, which generates a drive signal based on the output voltage signal and output current signal of the solar array module and the voltage signal and current signal of the battery pack to perform closed-loop control on the corresponding MPPT circuit, thereby achieving maximum power tracking of the solar array module and charging management of the battery pack. However, it does not involve the estimation of the energy balance of the entire satellite, nor does it involve the satellite attitude maneuver observation mission.

[0007] Patent document CN106972581B discloses an onboard energy scheduling method for remote sensing satellites based on energy status, which uses real-time acquisition of energy system related parameters, calculates the power supply status of the energy system for the injected satellite working mode, and schedules whether the working mode can be executed. However, it does not involve the calculation of lighting conditions and the 24-hour energy balance status. Summary of the invention

[0008] In view of the defects in the prior art, the purpose of the present invention is to provide a method and system for planning an attitude maneuver imaging mission based on satellite autonomous energy management.

[0009] A satellite autonomous energy management attitude maneuver imaging mission planning method provided by the present invention includes:

[0010] Step S1: dividing the mission hotspot area range of mobile imaging;

[0011] Step S2: Calculate satellite related parameters respectively; the satellite related parameters include satellite power parameters, satellite sub-satellite parameters and illumination duration within a certain period of time; when the satellite is not within the mission hotspot area of ​​the mobile imaging, calculate the total power consumption of the satellite;

[0012] Step S3: further calculating the energy balance state within a certain period of time according to the calculation results of satellite related parameters;

[0013] Step S4: When the satellite is within the mission hotspot area of ​​the mobile imaging, an energy early warning for the entire satellite is performed according to the energy balance state, so that the satellite can autonomously complete the observation mission and record the execution status of the mission.

[0014] Preferably, step S1 includes demarcating the mission hotspot area of ​​mobile imaging into a square area, setting the observation time of the hotspot area, and the maximum time does not exceed the satellite's mobile observation capability; according to actual conditions, recording, clearing or replacing multiple mobile imaging mission hotspot areas on the satellite; if there is no hotspot area, the satellite mobile imaging mission power consumption is recorded as zero, and the observation time interval between adjacent hotspot areas is greater than the recovery time of the last satellite mobile observation.

[0015] Preferably, the satellite power parameter calculation process includes:

[0016] Instruct the onboard computer software to perform remote measurement and collection of battery discharge current and charge current at a fixed period;

[0017] Calculate the current discharge and charge capacity of the battery:

[0018]

[0019] Among them, Q 充i Indicates the current charging power; Q放i Indicates the current discharge capacity; Q 充(i-1) Indicates the last charging level; Q 放(i-1) Indicates the discharge power of the previous beat; I 充i Indicates the current charging current; I 放i Indicates the current discharge current; K 充 Indicates the charging capacity adjustment factor; K 放 Indicates the discharge capacity adjustment coefficient; I 充i and I 放i are the charging current and discharging current collected by the onboard computer respectively; T is the current collection time interval; and the charging power sets the saturation threshold.

[0020] Preferably, the calculation of the satellite sub-satellite point parameters includes:

[0021] Calculate geographic longitude λ w :

[0022] λ w =arctan(r y1 / r x1 );

[0023] Among them, r y1 , r x1 are the Y-axis and X-axis components of the satellite position vector in the WGS-84 coordinate system;

[0024] Calculate the geocentric latitude φ' w :

[0025]

[0026] Among them, r z1 , is the Z-axis component of the satellite position vector in the WGS-84 coordinate system;

[0027] Calculate geographic latitude φ w :

[0028]

[0029] Among them, R e is the equatorial radius of the Earth; R p is the polar radius of the Earth.

[0030] Preferably, step S2 includes calculating the total power consumption of the satellite based on the attitude and orbit control subsystem, and the calculation process includes:

[0031] Calculation of power consumption increase of attitude maneuver of attitude and orbit control subsystem: When the attitude and orbit control subsystem is in the solar mode during observation missions without hotspot areas, it guarantees the energy supply of the whole satellite. When in maneuvering missions, it performs attitude maneuvers through the propulsion subsystem to complete the observation mission of hotspot areas on the ground. Estimation of power consumption increase of attitude and orbit control subsystem during attitude maneuvers wz1 Calculated as:

[0032] w z1 =k w1 t1;

[0033] Among them, t1 is the working time of the thruster, k w1 is the power consumption per unit time of thruster operation;

[0034] Calculation of power consumption increase when the payload is turned on: The attitude and orbit control subsystem is in standby mode when there is no hotspot observation mission to save the energy supply of the entire satellite. When there is a hotspot observation mission, it is turned on by the on-board computer program control command. The default value of the payload power-on time is set to t2; after the power-on time ends, it is shut down by the on-board computer program control command; the power consumption increase when the payload is turned on is estimated to be w z2 Calculated as:

[0035] w z2 =k w2 t2;

[0036] Among them, t2 is the working time of the payload, k w2 The power consumption per unit time of the payload when it is powered on;

[0037] Satellite total power consumption w z Calculated as:

[0038] w z =k1w z1 +k2w z2 +k3w z3 ;

[0039] w z3 =k w3 (t1+t2);

[0040] where k w3 is the satellite steady-state average power consumption per unit time, w z3 is the steady-state power consumption of the satellite, and k1, k2, and k3 represent the adjustment coefficients of the number of on-orbit injections.

[0041] Preferably, the calculation of the illumination duration within the certain period of time includes:

[0042] The angle between the earth, star and sun vector α r calculate:

[0043] α r =cos -1 (x s x w1 +y s y w1 +z s z w1 );

[0044] Among them, [x w1 ,y w1 , z w1 ] is the position vector of the satellite in the inertial system at time T, [x s ,y s , z s ] is the sun position vector at time T;

[0045] Calculation of light duration within a certain period of time:

[0046]

[0047] Among them, β is the angle between the earth-satellite-sun vector when the satellite enters the shadow area from the light;

[0048] When α r When ≥β, the satellite is in the shadow area; when α r When <β, the satellite is in the illumination area;

[0049] The on-board computer recursively calculates the illumination α that meets the condition within a certain period of time according to the fixed interval. r The accumulated time of ≥β is the total illumination duration T g .

[0050] Preferably, step S3 includes the following sub-steps:

[0051] Step S3.1: Calculate the total discharge capacity Q 总放i :

[0052]

[0053] Where n = [0, 1, 2, 3 ... n], is the number of autonomous imaging area tasks defined according to geographic longitude and latitude, V z The bus voltage can be adjusted;

[0054] Step S3.2: Calculate the total charge Q 总充i :

[0055] Q 总充i =Q 充i +I g (T g -nt2);

[0056] Among them, I g is the full power charging current;

[0057] Step S3.3: Calculation of charge-discharge ratio D:

[0058]

[0059] Wherein, p is the charge-discharge ratio adjustment coefficient.

[0060] Preferably, step S4 includes the following sub-steps:

[0061] Step S4.1: Whole-star energy warning:

[0062] When 0≤D<DOD1, it means that the energy is sufficient to perform the mobile imaging task;

[0063] When DOD1≤D<DOD2, it indicates that the energy is tight and the mobile imaging mission can be performed;

[0064] When DOD2≤D, it means that the energy is insufficient and the motorized imaging task cannot be performed;

[0065] Among them, DOD1 and DOD2 represent the charge and discharge ratio coefficients;

[0066] Step S4.2: Display and record the status of the imaging task during execution.

[0067] A satellite autonomous energy management attitude maneuver imaging mission planning system provided by the present invention includes:

[0068] Module M1: Divide the mission hotspot area of ​​mobile imaging;

[0069] Module M2: Calculate satellite related parameters respectively; the satellite related parameters include satellite power parameters, satellite sub-satellite parameters and illumination duration within a certain period of time; when the satellite is not in the mission hotspot area of ​​the mobile imaging, calculate the total power consumption of the satellite;

[0070] Module M3: Based on the calculation results of satellite-related parameters, further calculate the energy balance state within a certain period of time;

[0071] Module M4: When the satellite is in the mission hotspot area of ​​the mobile imaging, an energy early warning for the entire satellite is performed according to the energy balance state, so that the satellite can autonomously complete the observation mission and record the execution status of the mission.

[0072] Preferably, the module M1 includes demarcating the mission hotspot area of ​​mobile imaging into a square area, setting the observation time of the hotspot area, and the maximum time does not exceed the satellite's mobile observation capability; according to actual conditions, recording, clearing or replacing multiple mobile imaging mission hotspot areas on the satellite; if there is no hotspot area, the satellite mobile imaging mission power consumption is recorded as zero, and the observation time interval between adjacent hotspot areas is greater than the recovery time of the last satellite mobile observation.

[0073] Preferably, the satellite power parameter calculation process includes:

[0074] Instruct the onboard computer software to perform remote measurement and collection of battery discharge current and charge current at a fixed period;

[0075] Calculate the current discharge and charge capacity of the battery:

[0076]

[0077]

[0078] Among them, Q 充i Indicates the current charging power; Q 放i Indicates the current discharge capacity; Q 充(i-1) Indicates the last charging level; Q 放(i-1) Indicates the discharge power of the previous beat; I 充i Indicates the current charging current; I 放i Indicates the current discharge current; K 充 Indicates the charging capacity adjustment factor; K 放 Indicates the discharge capacity adjustment coefficient; I 充i and I 放i are the charging current and discharging current collected by the onboard computer respectively; T is the current collection time interval; and the charging power sets the saturation threshold.

[0079] Preferably, the calculation of the satellite sub-satellite point parameters includes:

[0080] Calculate geographic longitude λ w :

[0081]

[0082] Among them, r y1 , r x1 are the Y-axis and X-axis components of the satellite position vector in the WGS-84 coordinate system;

[0083] Calculate the geocentric latitude φ' w :

[0084]

[0085] Among them, r z1 , is the Z-axis component of the satellite position vector in the WGS-84 coordinate system;

[0086] Calculate geographic latitude φ w :

[0087]

[0088] Among them, R e is the equatorial radius of the Earth; R p is the polar radius of the Earth.

[0089] Preferably, the module M2 includes calculating the total power consumption of the satellite based on the attitude and orbit control subsystem, and the calculation process includes:

[0090] Calculation of power consumption increase of attitude maneuver of attitude and orbit control subsystem: When the attitude and orbit control subsystem is in the solar mode during observation missions without hotspot areas, it guarantees the energy supply of the whole satellite. When in maneuvering missions, it performs attitude maneuvers through the propulsion subsystem to complete the observation mission of hotspot areas on the ground. Estimation of power consumption increase of attitude and orbit control subsystem during attitude maneuvers w z1 Calculated as:

[0091] w z1 =k w1 t1;

[0092] Among them, t1 is the working time of the thruster, k w1 is the power consumption per unit time of thruster operation;

[0093] Calculation of power consumption increase when the payload is turned on: The attitude and orbit control subsystem is in standby mode when there is no hotspot observation mission to save the energy supply of the entire satellite. When there is a hotspot observation mission, it is turned on by the on-board computer program control command. The default value of the payload power-on time is set to t2; after the power-on time ends, it is shut down by the on-board computer program control command; the power consumption increase when the payload is turned on is estimated to be w z2 Calculated as:

[0094] w z2 =k w2 t2;

[0095] Among them, t2 is the working time of the payload, k w2 The power consumption per unit time of the payload when it is powered on;

[0096] Satellite total power consumption w z Calculated as:

[0097] w z =k1w z1 +k2w z2 +k3w z3 ;

[0098] w z3 =k w3 (t1+t2);

[0099] where k w3 is the satellite steady-state average power consumption per unit time, w z3 is the steady-state power consumption of the satellite, and k1, k2, and k3 represent the adjustment coefficients of the number of on-orbit injections.

[0100] Preferably, the calculation of the illumination duration within the certain period of time includes:

[0101] The angle between the earth, star and sun vector α r calculate:

[0102] α r =cos -1 (xs x w1 +y s y w1 +z s z w1 );

[0103] Among them, [x w1 ,y w1 , z w1 ] is the position vector of the satellite in the inertial system at time T, [x s ,y s , z s ] is the sun position vector at time T;

[0104] Calculation of light duration within a certain period of time:

[0105]

[0106] Among them, β is the angle between the earth-satellite-sun vector when the satellite enters the shadow area from the light;

[0107] When α r When ≥β, the satellite is in the shadow area; when α r When <β, the satellite is in the illumination area;

[0108] The on-board computer recursively calculates the illumination α that meets the condition within a certain period of time according to the fixed interval. r The accumulated time of ≥β is the total illumination duration T g .

[0109] Preferably, the module M3 includes the following submodules:

[0110] Module M3.1: Calculate the total discharge power Q 总放i :

[0111]

[0112] Where n = [0, 1, 2, 3 ... n], is the number of autonomous imaging area tasks defined according to geographic longitude and latitude, V z The bus voltage can be adjusted;

[0113] Module M3.2: Calculate the total charge Q 总充i :

[0114] Q 总充i =Q 充i +I g (T g -nt2);

[0115] Among them, I g is the full power charging current;

[0116] Module M3.3: Charge and discharge ratio calculation D:

[0117]

[0118] Wherein, p is the charge-discharge ratio adjustment coefficient.

[0119] Preferably, the module M4 includes the following submodules:

[0120] Module M4.1: Whole-star energy warning:

[0121] When 0≤D<DOD1, it means that the energy is sufficient to perform the mobile imaging task;

[0122] When DOD1≤D<DOD2, it indicates that the energy is tight and the mobile imaging mission can be performed;

[0123] When DOD2≤D, it means that the energy is insufficient and the motorized imaging task cannot be performed;

[0124] Among them, DOD1 and DOD2 represent the charge and discharge ratio coefficients;

[0125] Module M4.2: Display and record the status of the imaging task during execution.

[0126] Compared with the prior art, the present invention has the following beneficial effects:

[0127] 1. The present invention can autonomously monitor the energy status in real time, calculate the illumination duration of the next 24 hours according to the orbital parameters, evaluate the energy status of the entire satellite, and autonomously complete multiple mobile observation tasks; when there is no observation task, it can perform inertial solar alignment to ensure the energy of the entire satellite, and when there is an observation task, it can realize the satellite of earth observation through attitude maneuvers, which has good flexibility.

[0128] 2. The present invention enables the satellite to pass through the designated imaging mission hotspot area, autonomously complete the observation mission according to the energy situation of the entire satellite, and record the mission execution status, thereby completing the satellite's on-orbit autonomous imaging mission management, which has good practicality.

[0129] Other beneficial effects of the present invention will be explained in the specific implementation manner through the introduction of specific technical features and technical solutions. Through the introduction of these technical features and technical solutions, those skilled in the art should be able to understand the beneficial technical effects brought about by the technical features and technical solutions. BRIEF DESCRIPTION OF THE DRAWINGS

[0130] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings:

[0131] Figure 1 The figure is a flow chart of the method of the present invention.

[0132] Figure 2 This is a flow chart of the electric quantity calculation in the present invention. DETAILED DESCRIPTION

[0133] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0134] Reference Figure 1 As shown, a satellite autonomous energy management attitude maneuver imaging mission planning method includes:

[0135] Step 1: Delineate the hotspot areas for mobile imaging missions:

[0136] Step 1.1: The hotspot area of ​​the mobile imaging mission is delineated in the form of a square area, that is, the geographical longitude and latitude range of the hotspot area is given;

[0137] Step 1.2: Set the observation time of the hotspot area, and the maximum time shall not exceed the satellite's mobile observation capability;

[0138] Step 1.3: Multiple hot spots of mobile imaging missions can be recorded on the satellite, and the number of hot spots can be cleared and replaced. If the hot spot area is 0, it means that there is no hot spot area currently being observed, and the power consumption of the satellite mobile imaging mission is recorded as zero;

[0139] Step 1.4: The observation time interval between adjacent hotspot areas must be greater than the recovery time of the last satellite observation maneuver.

[0140] Step 2: Reference Figure 2 As shown, the fuel gauge is calculated in real time on track:

[0141] Step 2.1: The onboard computer software performs remote measurement and acquisition of battery discharge current and charge current at a period of 0.1 s;

[0142] Step 2.2: Calculate the current discharge and charge capacity of the battery;

[0143]

[0144] Among them, Q 充i Indicates the current charging power; Q 放i Indicates the current discharge capacity; Q 充(i-1) Indicates the last charging level; Q 放(i-1) Indicates the discharge power of the previous beat; I 充i Indicates the current charging current; I 放i Current beat discharge current; K 充Charging capacity adjustment factor; K 放 Discharge capacity adjustment factor; I 充i and I 放i are the charging current and discharging current collected by the onboard computer; T is the current collection time interval; and the charging power sets the saturation threshold.

[0145] Step 3: Calculation of satellite sub-satellite point:

[0146] Step 3.1 Geographic longitude λ w calculate:

[0147] λ w =arctan(r y1 / r x1 );

[0148] where r y1 , r x1 are the Y-axis and X-axis components of the satellite position vector in the WGS-84 coordinate system.

[0149] Step 3.2 Geocentric latitude φ' w calculate:

[0150]

[0151] where r z1 , which is the Z-axis component of the satellite position vector in the WGS-84 coordinate system.

[0152] Step 3.3 Geographic latitude φ w calculate:

[0153]

[0154] Among them, R e is the equatorial radius of the Earth; R p is the polar radius of the Earth.

[0155] Step 4: Estimation of total power consumption of satellite maneuver imaging mission:

[0156] Step 4.1: Estimation of power consumption of attitude and orbit control subsystem during attitude maneuvers: The attitude and orbit control subsystem is in the solar mode during observation missions without hotspot areas to ensure the energy supply of the entire satellite. During maneuvering missions, the propulsion subsystem performs attitude maneuvers to complete observation missions of hotspot areas on the ground. Estimation of power consumption of attitude and orbit control subsystem during attitude maneuvers z1 calculate:

[0157] w z1 =k w1 t1;

[0158] Among them, t1 is the working time of the thruster, k w is the power consumption per unit time of the thruster.

[0159] Step 4.2: Estimation of power consumption when the payload is turned on: The attitude and orbit control subsystem is in standby mode when there is no hotspot observation mission to save the energy supply of the entire satellite. When there is a hotspot observation mission, it is turned on by the onboard computer program control command. The default value of the payload power-on time is set to t2, which can be adjusted by the orbit number. After the power-on time is over, it is shut down by the onboard computer program control command. Estimation of power consumption when the payload is turned on z2 calculate:

[0160] w z2 =k w2 t2;

[0161] Among them, t2 is the working time of the payload, k w2 It is the power consumption per unit time of the payload when it is turned on.

[0162] Step 4.3: Satellite total power consumption w z calculate:

[0163] w z =k1w z1 +k2w z2 +k3w z3 ;

[0164] w z3 =k w3 (t1+t2);

[0165] Among them, k w3 is the satellite steady-state average power consumption per unit time, w z3 is the steady-state power consumption of the satellite, and k1, k2, and k3 represent the adjustment coefficients of the number of on-orbit injections.

[0166] Step 5: Calculation of 24-hour light duration:

[0167] Step 5.1: Earth-planet-sun vector angle α r calculate:

[0168] α r =cos -1 (x s x w1 +y s y w1 +z s z w1 );

[0169] Where: w1 ,y w1 , z w1 ] is the position vector of the satellite in the inertial system at time T, [x s ,y s , z s] is the sun position vector at time T.

[0170] Step 5.2: Calculation of light duration within 24 hours:

[0171]

[0172] β is the angle between the Earth, the satellite and the sun when the satellite enters the shadow area from the light

[0173] When α r When ≥β, the satellite is in the shadow area; when α r When <β, the satellite is in the illuminated area.

[0174] The onboard computer recursively calculates the number of hours of illumination that will meet the condition α in the next 24 hours every 0.5 seconds. r The accumulated time of ≥β is the total duration of 24 hours of light T g .

[0175] Step 6: Energy balance within 24 hours:

[0176] Step 6.1: Calculate the total discharge capacity Q 总放i :

[0177]

[0178] Where n = [0, 1, 2, 3 ... n], is the number of autonomous imaging area tasks defined according to geographic longitude and latitude, V z The bus voltage can be adjusted.

[0179] Step 6.2: Calculate the total charge Q 总充i :

[0180] Q 总充i =Q 充i +I g (T g -nt2);

[0181] Among them I g is the full power charging current, which can be changed in rail note number.

[0182] Step 6.3: Calculation of charge-discharge ratio D:

[0183]

[0184] p is the charge-discharge ratio adjustment coefficient.

[0185] Step 7: Recording of the execution status of the whole-satellite energy warning and mobile imaging mission:

[0186] Step 7.1: Whole-star energy warning:

[0187] When 0≤D<DOD1, it means that the energy is sufficient and the mobile imaging mission can be performed;

[0188] When DOD1≤D<DOD2, energy is tight and mobile imaging tasks can be performed;

[0189] When DOD2≤D, the energy is insufficient and the motorized imaging mission cannot be performed;

[0190] Among them, DOD1 and DOD2 represent the charge and discharge ratio coefficients, which can be adjusted on the track.

[0191] Step 7.2: Record the execution status of the imaging task:

[0192] The time, energy status, and geographical longitude and latitude of the designated imaging area of ​​the completed imaging observation tasks are displayed and recorded through telemetry, and up to 10 items can be displayed in chronological order; the time, energy status, and geographical longitude and latitude of the designated imaging area of ​​the imaging observation tasks that have been instructed but not yet executed are displayed and recorded through telemetry, and up to 10 items can be displayed in chronological order.

[0193] The above is a basic embodiment of the present invention. The technical solution of the present invention is further described below through a preferred embodiment.

[0194] The present invention can autonomously monitor the energy status in real time, calculate the illumination duration of the next 24 hours according to the orbital parameters, evaluate the energy status of the entire satellite, and autonomously complete multiple mobile observation tasks; when there is no observation task, it can perform inertial solar alignment to ensure the energy of the entire satellite, and when there is an observation task, it can realize the satellite of earth observation through attitude maneuvers, and has good flexibility.

[0195] Example 1

[0196] Reference Figure 1 As shown, step 1 includes:

[0197] The hotspot area of ​​the mobile imaging mission is delineated in the form of a square area, that is, the geographical longitude and latitude range of the hotspot area is given, as shown in Table 1;

[0198] Table 1 Observation area table

[0199]

[0200] Multiple hot spots of mobile imaging missions can be recorded on the satellite, and the number of hot spots can be cleared and replaced. If the hot spot area is 0, it means that there is no hot spot area currently being observed, and the power consumption of the satellite mobile imaging mission is recorded as zero. The n in the table represents the mission sequence number. The maximum observation time of the hot spot area is set to 20 minutes. The satellite completes the attitude maneuver and returns autonomously according to the time. The observation time interval of adjacent hot spots must be greater than the recovery time of the last satellite observation maneuver, that is, 20 minutes.

[0201] Step 2 includes:

[0202] The onboard computer software performs telemetry acquisition of battery discharge current and charge current at a cycle of 0.1s;

[0203] The current discharge and charge capacity of the battery is calculated using the following formula;

[0204]

[0205]

[0206] The charging capacity is set to a saturation threshold of 180Ah.

[0207] Step 3 includes:

[0208] Geographic longitude λ w calculate:

[0209] λ w =arctan(r y1 / r x1 );

[0210] where r y1 , r x1 are the Y-axis and X-axis components of the satellite position vector in the WGS-84 coordinate system.

[0211] Geocentric latitude φ' w calculate:

[0212]

[0213] where r z1 , which is the Z-axis component of the satellite position vector in the WGS-84 coordinate system.

[0214] Geographic latitude φ w calculate:

[0215]

[0216] Among them, R e is the equatorial radius of the Earth; R p is the polar radius of the Earth.

[0217] Step 4 includes:

[0218] Estimation of additional power consumption of attitude maneuvering of attitude and orbit control subsystem z1 calculate:

[0219] w z1 =k w1 t1;

[0220] Among them, t1 is the working time of the thruster, k wis the power consumption per unit time of the thruster.

[0221] Estimated power consumption when the payload is turned on: The attitude and orbit control subsystem is in standby mode when there is no hotspot observation mission to save the energy supply of the entire satellite. When there is a hotspot observation mission, it is turned on by the onboard computer program control command. The default value of the payload power-on time is set to t2, which can be adjusted by the orbit number. After the power-on time is over, it is shut down by the onboard computer program control command. Estimated power consumption when the payload is turned on z2 calculate:

[0222] w z2 =k w2 t2;

[0223] Among them, t2 is the working time of the payload, k w2 It is the power consumption per unit time of the payload when it is turned on.

[0224] Satellite total power consumption w z calculate:

[0225] w z =k1w z1 +k2w z2 +k3w z3 ;

[0226] w z3 =k w3 (t1+t2);

[0227] Among them, k w3 is the satellite steady-state average power consumption per unit time, w z3 is the steady-state power consumption of the satellite, and k1, k2, and k3 are the in-orbit adjustment coefficients.

[0228] Step 5 includes:

[0229] The angle between the earth, star and sun vector α r calculate:

[0230] α r =cos -1 (x s x w1 +y s y w1 +z s z w1 );

[0231] Where: [x w1 ,y w1 , z w1 ] is the position vector of the satellite in the inertial system at time T, [x s ,y s , z s ] is the sun position vector at time T.

[0232] Calculation of light duration within 24 hours:

[0233]

[0234] β is the angle between the Earth-satellite-sun vector when the satellite enters the shadow area from the light area;

[0235] When α r When ≥β, the satellite is in the shadow area; when α r When <β, the satellite is in the illuminated area.

[0236] The onboard computer recursively calculates the number of hours of illumination that will meet the condition α in the next 24 hours every 0.5 seconds. r The accumulated time of ≥β is the total duration of 24 hours of light T g .

[0237] Step 6 includes:

[0238] Calculate the total discharge capacity Q 总放i

[0239]

[0240] Where n = [0, 1, 2, 3 ... n], is the number of autonomous imaging area tasks defined according to geographic longitude and latitude, V z The bus voltage can be adjusted.

[0241] Calculate the total charge Q 总充i :

[0242] Q 总充i =Q 充i +I g (T g -nt2);

[0243] Among them, I g is the full power charging current, which can be changed in rail note number.

[0244] Calculation of charge-discharge ratio D:

[0245]

[0246] p is the charge-discharge ratio adjustment coefficient, and its default value is 1.

[0247] Step 7 includes:

[0248] Whole-star energy warning

[0249] When 0≤D<DOD1, it means that the energy is sufficient and the mobile imaging mission can be performed;

[0250] When DOD1≤D<DOD2, energy is tight and mobile imaging tasks can be performed;

[0251] When DOD2≤D, the energy is insufficient and the motorized imaging mission cannot be performed;

[0252] DOD1 and DOD2 are set at 20% and 40%, respectively, and the number of injections can be adjusted on track.

[0253] The time, energy status, and geographical longitude and latitude of the designated imaging area of ​​the completed imaging observation tasks are displayed and recorded through telemetry, and up to 10 items can be displayed in chronological order; the time, energy status, and geographical longitude and latitude of the designated imaging area of ​​the imaging observation tasks that have been instructed but not yet executed are displayed and recorded through telemetry, and up to 10 items can be displayed in chronological order.

[0254] The present invention also provides a satellite-based autonomous energy management attitude maneuverable imaging mission planning system. The satellite-based autonomous energy management attitude maneuverable imaging mission planning system can be implemented by executing the process steps of the satellite-based autonomous energy management attitude maneuverable imaging mission planning method, that is, those skilled in the art can understand the satellite-based autonomous energy management attitude maneuverable imaging mission planning method as a preferred implementation of the satellite-based autonomous energy management attitude maneuverable imaging mission planning system.

[0255] Specifically, a satellite autonomous energy management attitude maneuver imaging mission planning system includes:

[0256] Module M1: Divide the mission hotspot area of ​​mobile imaging;

[0257] Module M2: Calculate satellite related parameters respectively; the satellite related parameters include satellite power parameters, satellite sub-satellite parameters and illumination duration within a certain period of time; when the satellite is not in the mission hotspot area of ​​the mobile imaging, calculate the total power consumption of the satellite;

[0258] Module M3: Based on the calculation results of satellite-related parameters, further calculate the energy balance state within a certain period of time;

[0259] Module M4: When the satellite is in the mission hotspot area of ​​the mobile imaging, an energy early warning for the entire satellite is performed according to the energy balance state, so that the satellite can autonomously complete the observation mission and record the execution status of the mission.

[0260] The module M1 includes demarcating the mission hotspot area of ​​mobile imaging into a square area, setting the observation time of the hotspot area, and the maximum time does not exceed the satellite's mobile observation capability; recording, clearing or replacing multiple mobile imaging mission hotspot areas on the satellite according to actual conditions; if there is no hotspot area, the satellite mobile imaging mission power consumption is recorded as zero, and the observation time interval between adjacent hotspot areas is greater than the recovery time of the last satellite mobile observation.

[0261] The satellite power parameter calculation process includes:

[0262] Instruct the onboard computer software to perform remote measurement and collection of battery discharge current and charge current at a fixed period;

[0263] Calculate the current discharge and charge capacity of the battery:

[0264]

[0265] Among them, Q 充i Indicates the current charging power; Q 放i Indicates the current discharge capacity; Q 充(i-1) Indicates the last charging level; Q 放(i-1) Indicates the discharge power of the previous beat; I 充i Indicates the current charging current; I 放i Indicates the current discharge current; K 充 Indicates the charging capacity adjustment factor; K 放 Indicates the discharge capacity adjustment coefficient; I 充i and I 放i are the charging current and discharging current collected by the onboard computer respectively; T is the current collection time interval; and the charging power sets the saturation threshold.

[0266] The calculation of the satellite sub-satellite point parameters includes:

[0267] Calculate geographic longitude λ w :

[0268] λ w =arctan(r y1 / r x1 );

[0269] Among them, r y1 , r x1 are the Y-axis and X-axis components of the satellite position vector in the WGS-84 coordinate system;

[0270] Calculate the geocentric latitude φ' w :

[0271]

[0272] Among them, r z1 , is the Z-axis component of the satellite position vector in the WGS-84 coordinate system;

[0273] Calculate geographic latitude φ w :

[0274]

[0275] Among them, R eis the equatorial radius of the Earth; R p is the polar radius of the Earth.

[0276] The module M2 includes calculating the total power consumption of the satellite based on the attitude and orbit control subsystem, and the calculation process includes:

[0277] Calculation of power consumption increase of attitude maneuver of attitude and orbit control subsystem: When the attitude and orbit control subsystem is in the solar mode during observation missions without hotspot areas, it guarantees the energy supply of the whole satellite. When in maneuvering missions, it performs attitude maneuvers through the propulsion subsystem to complete the observation mission of hotspot areas on the ground. Estimation of power consumption increase of attitude and orbit control subsystem during attitude maneuvers w z1 Calculated as:

[0278] w z1 =k w1 t1;

[0279] Among them, t1 is the working time of the thruster, k w1 is the power consumption per unit time of thruster operation;

[0280] Calculation of power consumption increase when the payload is turned on: The attitude and orbit control subsystem is in standby mode when there is no hotspot observation mission to save the energy supply of the entire satellite. When there is a hotspot observation mission, it is turned on by the on-board computer program control command. The default value of the payload power-on time is set to t2; after the power-on time ends, it is shut down by the on-board computer program control command; the power consumption increase when the payload is turned on is estimated to be w z2 Calculated as:

[0281] w z2 =k w2 t2;

[0282] Among them, t2 is the working time of the payload, k w2 The power consumption per unit time of the payload when it is powered on;

[0283] Satellite total power consumption w z Calculated as:

[0284] w z =k1w z1 +k2w z2 +k3w z3 ;

[0285] w z3 =k w3 (t1+t2);

[0286] where k w3 is the satellite steady-state average power consumption per unit time, w z3 is the steady-state power consumption of the satellite, and k1, k2, and k3 represent the adjustment coefficients of the number of on-orbit injections.

[0287] The calculation of the illumination duration within a certain period of time includes:

[0288] The angle between the earth, star and sun vector α r calculate:

[0289] α r =cos -1 (x s x w1 +y s y w1 +z s z w1 );

[0290] Among them, [x w1 ,y w1 , z w1 ] is the position vector of the satellite in the inertial system at time T, [x s ,y s , z s ] is the sun position vector at time T;

[0291] Calculation of light duration within a certain period of time:

[0292]

[0293] Among them, β is the angle between the earth-satellite-sun vector when the satellite enters the shadow area from the light;

[0294] When α r When ≥β, the satellite is in the shadow area; when α r When <β, the satellite is in the illumination area;

[0295] The on-board computer recursively calculates the illumination α that meets the condition within a certain period of time according to the fixed interval. r The accumulated time of ≥β is the total illumination duration T g .

[0296] The module M3 includes the following submodules:

[0297] Module M3.1: Calculate the total discharge power Q 总放i :

[0298]

[0299] Where n = [0, 1, 2, 3 ... n], is the number of autonomous imaging area tasks defined according to geographic longitude and latitude, V z The bus voltage can be adjusted;

[0300] Module M3.2: Calculate the total charge Q 总充i :

[0301] Q 总充i =Q 充i +I g (Tg -nt2);

[0302] Among them, I g is the full power charging current;

[0303] Module M3.3: Charge and discharge ratio calculation D:

[0304]

[0305] Wherein, p is the charge-discharge ratio adjustment coefficient.

[0306] The module M4 includes the following submodules:

[0307] Module M4.1: Whole-star energy warning:

[0308] When 0≤D<DOD1, it means that the energy is sufficient to perform the mobile imaging task;

[0309] When DOD1≤D<DOD2, it indicates that the energy is tight and the mobile imaging mission can be performed;

[0310] When DOD2≤D, it means that the energy is insufficient and the motorized imaging task cannot be performed;

[0311] Among them, DOD1 and DOD2 represent the charge and discharge ratio coefficients;

[0312] Module M4.2: Display and record the status of the imaging task during execution.

[0313] Those skilled in the art know that, in addition to realizing the system and its various devices, modules, and units provided by the present invention in a purely computer-readable program code, it is entirely possible to realize the same functions in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered as a hardware component, and the devices, modules, and units included therein for realizing various functions can also be regarded as structures within the hardware component; the devices, modules, and units for realizing various functions can also be regarded as both software modules for realizing the method and structures within the hardware component.

[0314] In the description of the present application, it should be understood that the terms "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the present application.

[0315] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A satellite autonomous energy management attitude maneuver imaging mission planning method, characterized in that: include: Step S1: Divide the mission hotspot area range of mobile imaging; Step S2: Calculate satellite related parameters respectively; the satellite related parameters include satellite power parameters, satellite sub-satellite point parameters and illumination duration within a certain period of time; When the satellite is not within the mission hotspot area of ​​the mobile imaging, calculating the total power consumption of the satellite; Step S3: further calculating the energy balance state within a certain period of time according to the calculation results of satellite related parameters; Step S4: When the satellite is within the mission hotspot area of ​​the mobile imaging, an energy early warning for the entire satellite is performed according to the energy balance state, so that the satellite can autonomously complete the observation mission and record the execution status of the mission.

2. The satellite autonomous energy management attitude maneuver imaging mission planning method according to claim 1 is characterized in that: The step S1 includes demarcating the mission hotspot area of ​​the mobile imaging into a square area, setting the observation time of the hotspot area, and the maximum time does not exceed the mobile observation capability of the satellite; according to the actual situation, recording, clearing or replacing multiple mission hotspot areas of the mobile imaging on the satellite; If there is no hotspot area, the power consumption of the satellite maneuver imaging mission is recorded as zero, and the observation time interval between adjacent hotspot areas is greater than the recovery time of the last satellite maneuver observation.

3. The satellite autonomous energy management attitude maneuver imaging mission planning method according to claim 1 is characterized in that: The satellite power parameter calculation process includes: Instruct the onboard computer software to perform remote measurement and collection of battery discharge current and charge current at a fixed period; Calculate the current discharge and charge capacity of the battery: Among them, Q 充i Indicates the current charging power; Q 放i Indicates the current discharge capacity; Q 充(i-1) Indicates the last charging level; Q 放(i-1) Indicates the discharge power of the previous beat; I 充i Indicates the current charging current; I 放i Indicates the current discharge current; K 充 Indicates the charging capacity adjustment factor; K 放 Indicates the discharge capacity adjustment coefficient; I 充i and I 放i are the charging current and discharging current collected by the onboard computer respectively; T is the current collection time interval; and the charging power sets the saturation threshold.

4. The satellite autonomous energy management attitude maneuver imaging mission planning method according to claim 1 is characterized in that: The calculation of the satellite sub-satellite point parameters includes: Calculate geographic longitude λ w : λ w =arctan(r y1 / r x1 ); Among them, r y1 , r x1 are the Y-axis and X-axis components of the satellite position vector in the WGS-84 coordinate system; Calculate the geocentric latitude φ' w : Among them, r z1 , is the Z-axis component of the satellite position vector in the WGS-84 coordinate system; Calculate geographic latitude φ w : Among them, R e is the equatorial radius of the Earth; R p is the polar radius of the Earth.

5. The satellite autonomous energy management attitude maneuver imaging mission planning method according to claim 3 is characterized in that: The step S2 includes calculating the total power consumption of the satellite based on the attitude and orbit control subsystem, and the calculation process includes: Calculation of power consumption increase of attitude maneuver of attitude and orbit control subsystem: When the attitude and orbit control subsystem is in the solar mode during observation missions without hotspot areas, it guarantees the energy supply of the whole satellite. When in maneuvering missions, it performs attitude maneuvers through the propulsion subsystem to complete the observation mission of hotspot areas on the ground. Estimation of power consumption increase of attitude and orbit control subsystem during attitude maneuvers w z1 Calculated as: w z1 =k w1 t1; Among them, t1 is the working time of the thruster, k w1 is the power consumption per unit time of thruster operation; Calculation of power consumption increase when the payload is turned on: The attitude and orbit control subsystem is in standby mode when there is no hotspot observation mission to save the energy supply of the entire satellite. When there is a hotspot observation mission, it is turned on by the on-board computer program control command. The default value of the payload power-on time is set to t2; after the power-on time ends, it is shut down by the on-board computer program control command; the power consumption increase when the payload is turned on is estimated to be w z2 Calculated as: w z2 =k w2 t2; Among them, t2 is the working time of the payload, k w2 The power consumption per unit time of the payload when it is powered on; Satellite total power consumption w z Calculated as: In z =k1w z1 +k2w z2 +k3w z3 ; w z3 =k w3 (t1+t2); where k w3 is the satellite steady-state average power consumption per unit time, w z3 is the steady-state power consumption of the satellite, and k1, k2, and k3 represent the adjustment coefficients of the number of on-orbit injections.

6. The satellite autonomous energy management attitude maneuver imaging mission planning method according to claim 5 is characterized in that: The calculation of the illumination duration within a certain period of time includes: The angle between the earth, star and sun vector α r calculate: α r <cos -1 (x s x w1 +y s y w1 +z s z w1 ): Among them, [x w1 ,y w1 , z w1 ] is the position vector of the satellite in the inertial system at time T, [x s ,y s , z s ] is the sun position vector at time T; Calculation of light duration within a certain period of time: Among them, β is the angle between the Earth-satellite-sun vector when the satellite enters the shadow area from the light; When α r When ≥β, the satellite is in the shadow area; when α r When <β, the satellite is in the illumination area; The on-board computer recursively calculates the illumination α that meets the condition within a certain period of time according to the fixed interval. r The accumulated time of ≥β is the total illumination duration T g .

7. The satellite autonomous energy management attitude maneuver imaging mission planning method according to claim 6 is characterized in that: The step S3 comprises the following sub-steps: Step S3.1: Calculate the total discharge capacity Q 总放i : Where n = [0, 1, 2, 3 ... n], is the number of autonomous imaging area tasks defined according to geographic longitude and latitude, V z The bus voltage can be adjusted; Step S3.2: Calculate the total charge Q 总充i : Q 总充i =Q 充i +I g (T g -nt2); Among them, I g is the full power charging current; Step S3.3: Calculation of charge-discharge ratio D: Wherein, p is the charge-discharge ratio adjustment coefficient.

8. The satellite autonomous energy management attitude maneuver imaging mission planning method according to claim 7 is characterized in that: The step S4 comprises the following sub-steps: Step S4.1: Whole-star energy warning: When 0≤D<DOD1, it means that the energy is sufficient to perform the mobile imaging task; When DOD1≤D<DOD2, it indicates that the energy is tight and the mobile imaging mission can be performed; When DOD2≤D, it means that the energy is insufficient and the motorized imaging task cannot be performed; Among them, DOD1 and DOD2 represent the charge and discharge ratio coefficients; Step S4.2: Display and record the status of the imaging task during execution.

9. A satellite autonomous energy management attitude maneuver imaging mission planning system, characterized in that: include: Module M1: Divide the mission hotspot area of ​​mobile imaging; Module M2: Calculate satellite related parameters respectively; the satellite related parameters include satellite power parameters, satellite sub-satellite point parameters and illumination duration within a certain period of time; When the satellite is not within the mission hotspot area of ​​the mobile imaging, calculating the total power consumption of the satellite; Module M3: Based on the calculation results of satellite-related parameters, further calculate the energy balance state within a certain period of time; Module M4: When the satellite is in the mission hotspot area of ​​the mobile imaging, an energy early warning for the entire satellite is performed according to the energy balance state, so that the satellite can autonomously complete the observation mission and record the execution status of the mission.

10. The satellite autonomous energy management attitude maneuver imaging mission planning system according to claim 9 is characterized in that: The module M1 includes demarcating the mission hotspot area of ​​mobile imaging into a square area, setting the observation time of the hotspot area, and the maximum time does not exceed the satellite's mobile observation capability; according to actual conditions, recording, clearing or replacing multiple mission hotspot areas of mobile imaging on the satellite; If there is no hotspot area, the power consumption of the satellite maneuver imaging mission is recorded as zero, and the observation time interval between adjacent hotspot areas is greater than the recovery time of the last satellite maneuver observation.

Images