Zero-momentum satellite lunar calibration observation control method and system

Abstract

The application provides a zero-momentum satellite moon calibration observation control method and system, which increases an attitude periodic bias function when a satellite pitch attitude is maneuvered, and increases a sinusoidal change bias on a pitch attitude reference of moon calibration observation; a position of the moon in a satellite body coordinate system is calculated and determined by satellite orbit measurement, attitude measurement and a star-moon vector of a satellite center of mass pointing to a moon center of mass, so as to be used as a reference of moon calibration observation, and the star-moon vector is always located in a YbObZb plane of the satellite body coordinate system through attitude control. The application can effectively avoid the risk caused by a large deviation between a satellite attitude reference and a load optical axis pointing, and ensure that the effective load realizes moon calibration observation and the observation time meets the requirements.

Description

Technical Field

[0001] This invention relates to the fields of satellite attitude control and on-orbit payload calibration, specifically to a method and system for lunar calibration observation and control of zero-momentum satellites. More specifically, it relates to a method for lunar calibration observation and control of satellite payloads. Background Technology

[0002] Earth observation satellites are playing an increasingly important role in monitoring land, ocean, and atmosphere. As the crown jewel of aerospace high technology, the function and performance of the payload determine the breadth and depth of satellite applications. Currently, researchers are working hard on research into optical, infrared, and other payloads, with applications covering meteorological and oceanographic observation, land resource exploration, and other fields.

[0003] Due to the impact and vibration experienced by the instruments during launch and changes in the on-orbit operating environment, the calibration parameters of the payload remote sensing instruments may deviate from the results of the pre-launch calibration tests. Engineering experience shows that the sensitivity of the visible light and infrared channel detectors of meteorological satellites decreases to some extent every year. In order to monitor the changes in the calibration parameters of the on-orbit detectors, it is necessary to conduct on-orbit calibration of the remote sensing instruments.

[0004] Currently, satellite on-orbit calibration sources mainly use on-board blackbodies and sunlight. However, blackbodies have inherent temperature control errors, and the attenuation of sunlight diffuse reflectors can adversely affect on-board calibration. Calibration using typical ground features as reference sources is affected by the Earth's atmosphere. Besides traditional calibration methods, the Moon is increasingly valued as an on-orbit calibration source, primarily because lunar calibration offers unique advantages over other methods. First, the Moon has an extremely stable reflectivity, with annual irradiance variation of less than 10. -8 Furthermore, the lunar reflectance spectrum is continuous and smooth, which can better reflect the structure of the solar spectrum; secondly, the lunar spectral radiance value is within the dynamic range of most satellite remote sensing instruments, so there is no need to introduce other components into the optical path, which can simplify the calibration process; in addition, the moon can be observed by any Earth orbit satellite, which provides a cross-calibration method and ensures the consistency and stability of the calibration.

[0005] During normal Earth observation by the satellite payload, the Moon enters the payload's field of view once a month. However, due to the satellite's orbital motion, the payload's lunar scanning observation time is less than 10 seconds, resulting in only a limited number of effective images. To increase the payload's lunar observation time, it is necessary to either perform satellite pitch attitude maneuvers or add a one-dimensional drive function to the payload in addition to Earth scanning observations.

[0006] Currently, research on lunar calibration observation and control technology for remote sensing satellites mainly focuses on the lunar imaging calibration of linear array time-delay integral CCD sensors. The FY-3 satellite has carried out research on lunar calibration control of similar payloads. The FY-3 satellite adopts an offset momentum control scheme, and its attitude maneuver controller design is quite different from that of the zero momentum control scheme.

[0007] Patent document CN105486315A discloses a method for adjusting the attitude of a remote sensing satellite for absolute calibration of the moon. Starting from the lunar imaging calibration of a linear array time-delay integral CCD sensor, it proposes to correct the satellite attitude in two ways: firstly, by adjusting the satellite's attitude to rotate around its own Y-axis to control the observation velocity-to-height ratio centered on the moon's field of view; secondly, by adjusting the satellite's attitude to rotate around its own Z-axis to eliminate image shift of the lunar image in the remote sensor's field of view, thereby obtaining a clear lunar image. The payload sensor in this invention is a TeGeHg sensor, primarily used to detect the brightness temperature of the Earth's and Moon's surfaces, which differs significantly from the imaging capabilities of a linear array time-delay integral CCD sensor. Therefore, the control objectives and methods are quite different.

[0008] Patent document CN105446346A discloses a method for adjusting the attitude of a remote sensing satellite for relative lunar calibration. It determines the relative calibration time period through simulation, determines the initial attitude angle for lunar relative calibration through dual-vector attitude determination, determines the attitude control angular velocity during the lunar relative calibration process based on lunar imaging parameter analysis, eliminates the remote sensor image shift problem during the absolute lunar calibration process, and designs the satellite attitude trajectory through simulation. The payload sensor in patent document CN105446346A is a linear array time-delay integral CCD sensor, which requires addressing the remote sensor image shift problem. In this invention, the pitch attitude reference for lunar tracking maneuver control is the satellite pitch attitude angle - astro-lunar vector azimuth angle - offset angle, which differs significantly from the other two methods.

[0009] The paper "Attitude Maneuver Compensation Method for Lunar Calibration of Optical Remote Sensing Satellites" (Spacecraft Engineering, Vol. 25, No. 4, August 2016, pp. 5-12) addresses the mismatch between integration time and pushbroom velocity during lunar imaging by a spaceborne CCD camera. It proposes a method to compensate for the CCD camera's pushbroom velocity through attitude maneuvering, thus solving the oversampling problem and achieving matching between pushbroom velocity and integration time. The payload sensor in this invention is a TeGeHg sensor, and its imaging method differs significantly from the CCD camera imaging method described in that paper. Furthermore, it does not involve the high-reliability lunar calibration observation achieved through pitch attitude offset as described in this patent.

[0010] The paper "A Method for Lunar Calibration of Low-Earth Orbit High-Resolution Remote Sensing Satellites by Attitude Maneuver" (Optics and Precision Engineering, Vol. 28, No. 9, September 2020, pp. 1913-1923) proposes a method for lunar calibration by attitude maneuver, including techniques such as selecting the timing of lunar calibration, planning the attitude of the lunar calibration satellite, and selecting payload imaging parameters. However, it does not involve the "realizing reliable lunar calibration observation through pitch attitude offset" or "achieving non-zero maneuver control of attitude angle or angular velocity before and after maneuver through attitude planning and tracking, thereby improving observation efficiency" proposed in this invention.

[0011] The paper "Research on Active Lunar Calibration Control Technology of FY-3(05) Satellite" (Shanghai Aerospace, Vol. 38, No. 2, 2021, pp. 37-44) proposes an active lunar calibration control scheme for bias momentum satellites, including a lunar vector interpolation calculation method and a bias momentum satellite attitude maneuver control scheme, enabling the satellite to perform a 360° pitch maneuver to complete the lunar calibration observation control task. A key feature of the scheme provided in this paper is that the satellite needs to perform one maneuver (101 minutes) to achieve lunar observation. The control method proposed in this invention is mainly aimed at zero momentum satellites, offering flexible maneuverability and enabling lunar observation at any time within 5 minutes to 101 minutes. Summary of the Invention

[0012] To address the shortcomings of existing technologies, the purpose of this invention is to provide a method and system for zero-momentum satellite lunar calibration observation and control.

[0013] A zero-momentum satellite lunar calibration observation control method provided by the present invention includes:

[0014] Step S1: Determine the start time TO of the lunar calibration observation;

[0015] Step S2: Calculate the elevation angle α and azimuth angle β of the star and moon vectors at time TO;

[0016] Step S3: Determine the periodic offset, i.e., the pitch attitude offset angle γ:

[0017] γ=εsin(2πt / τ)

[0018] Where ε is the peak value of the bias, t is the time variable, and τ is the bias period; the bias period is selected as 1 / N of the entire monthly observation time or the monthly observation duration, where N is a positive integer, and the bias parameters ε and τ are noted above.

[0019] Step S4: Based on the azimuth angle β of the lunar and celestial vectors at time TO, generate and add the delay execution note: allow lunar calibration observations; and determine the pitch attitude maneuver;

[0020] The elevation attitude reference during lunar observation is (θ-β-γ), where θ is the satellite elevation attitude angle, which is given in real time by the satellite attitude determination; β is the azimuth angle of the lunar-satellite vector, which is given in real time by the onboard computer; and γ is the elevation attitude offset angle, which is given in real time by calculation.

[0021] Step S5: After the lunar calibration observation mission is completed, perform pitch maneuver to return. Note: Lunar calibration observation is prohibited; do not enter lunar observation during Earth observation.

[0022] Preferably, in step S1, the ground determines the start time TO of the lunar calibration observation based on the lunar phase angle, the interval between the illuminated and shadowed areas where the satellite is located, and the conditions of the satellite's telemetry and control arc.

[0023] The start time of the lunar calibration observation TO is chosen around the full moon to avoid the lunar phase angle of 0°. The lunar phase angle is the angle formed by the line connecting the payload and the moon and the line connecting the moon and the sun.

[0024] The start time (TO) and end time (Tend) of the lunar calibration observation are chosen when the satellite is in shadow. Before and after the time (TO), there are visible satellite telemetry and control arcs with annotation commands and numbers. During the lunar calibration maneuver observation, the star sensor is not affected by stray light.

[0025] Preferably, step S2 includes:

[0026] Step S2.1: Calculate the position vector of the Moon in the instantaneous geocentric mean equatorial coordinate system, denoted as the Moon vector; the origin of the instantaneous geocentric mean equatorial coordinate system is located at the Earth's center of mass, the X-axis points to the mean vernal equinox, the Z-axis points to the mean north celestial pole, and the Y-axis is perpendicular to the X-axis and Z-axis, forming a right-hand rectangular coordinate system;

[0027] Step S2.2: Transform the lunar vector to the satellite body coordinate system through coordinate transformation, and normalize the lunar vector to (Xb, Yb, Zb), where Xb, Yb, and Zb represent the normalized X-axis coordinate, Y-axis coordinate, and Z-axis coordinate in the satellite body coordinate system, respectively.

[0028] Step S2.3: Calculate the elevation angle α and azimuth angle β of the star and moon vectors at time TO;

[0029] In step S2.3:

[0030] At time TO, the elevation angle α of the celestial-lunar vector pointing from the satellite's center of mass to the center of the moon is:

[0031] α = acos(Yb)

[0032] The lunar vector is on the +Yb side, and the elevation angle α is defined as positive, and vice versa;

[0033] The azimuth angle β of the star-moon vector at time TO is:

[0034] β = atan2(Xb, Zb)

[0035] When the projection of the lunar vector onto the XbObZb plane of the satellite body coordinate system is in the +Xb half-plane, the azimuth angle β is positive, and vice versa; where Ob represents the origin of the satellite body coordinate system.

[0036] Preferably, in step S4: if the azimuth angle β of the lunar vector at time TO is 0, then a delayed execution note is generated and uploaded: allowing lunar calibration observation; the note package contains the time when the lunar calibration observation maneuver control begins, i.e., time TO; if the azimuth angle β at time TO is not 0, then the time Tj for the pitch attitude maneuver to β = 0 is calculated based on the magnitude of β, thereby determining TO-Tj as the delayed note execution time for the pitch attitude maneuver, and a delayed execution note is generated and uploaded: allowing lunar calibration observation, including pitch maneuver; the note package contains the start execution time of the pitch attitude maneuver, TO-Tj, and the lunar calibration observation is automatically executed after the maneuver is completed;

[0037] The pitch attitude maneuver to β=0 is performed using a pre-planning and post-tracking approach, ensuring that the pitch angular velocity relative to the orbital coordinate system is near the orbital angular velocity after the maneuver is completed. The initial three-axis attitude angles are (Φ, θ, Ψ), and the three-axis attitude angular velocities are... Φ, θ, Ψ All are near zero; the three-axis attitude angles after the maneuver are (Φ′, θ′, Ψ′), and the three-axis attitude angular velocities are... Φ′、Ψ′、 Both are near zero, and θ′ is near β at time TO. Near the orbital angular velocity; "near" means the difference is less than a preset numerical range.

[0038] According to the present invention, a zero-momentum satellite lunar calibration observation control method is provided, which adds a periodic attitude offset function during satellite pitch attitude maneuvers and adds a sinusoidal offset amount to the pitch attitude reference for lunar calibration observation. By measuring the satellite orbit, attitude, and the star-moon vector pointing from the satellite's center of mass to the center of the moon, the azimuth of the moon in the satellite's body coordinate system is calculated and determined, which is used as the reference for lunar calibration observation. The star-moon vector is always located in the YbObZb plane of the satellite's body coordinate system through attitude control.

[0039] A zero-momentum satellite lunar calibration observation and control system according to the present invention includes:

[0040] Module M1: Determines the start time TO of the lunar calibration observation;

[0041] Module M2: Calculates the elevation angle α and azimuth angle β of the star and moon vectors at time TO;

[0042] Module M3: Determines the periodic offset, i.e., the pitch attitude offset angle γ.

[0043] γ=εsin(2πt / τ)

[0044] Where ε is the peak value of the bias, t is the time variable, and τ is the bias period; the bias period is selected as 1 / N of the entire monthly observation time or the monthly observation duration, where N is a positive integer, and the bias parameters ε and τ are noted above.

[0045] Module M4: Based on the azimuth angle β of the lunar and celestial vectors at time TO, it generates and adds the delay execution note number, allowing lunar calibration observations and determining pitch attitude maneuvers;

[0046] The elevation attitude reference during lunar observation is (θ-β-γ), where θ is the satellite elevation attitude angle, which is given in real time by the satellite attitude determination; β is the azimuth angle of the lunar-satellite vector, which is given in real time by the onboard computer; and γ is the elevation attitude offset angle, which is given in real time by calculation.

[0047] Module M5: After completing the lunar calibration observation mission, perform pitch maneuver to return. Note: Lunar calibration observation is prohibited; do not enter lunar observation during Earth observation.

[0048] Preferably, in module M1, the ground determines the start time TO of the lunar calibration observation based on the lunar phase angle, the interval between the illuminated and shadowed areas where the satellite is located, and the conditions of the satellite tracking arc.

[0049] The start time of the lunar calibration observation TO is chosen around the full moon to avoid the lunar phase angle of 0°. The lunar phase angle is the angle formed by the line connecting the payload and the moon and the line connecting the moon and the sun.

[0050] The start time (TO) and end time (Tend) of the lunar calibration observation are chosen when the satellite is in shadow. Before and after the time (TO), there are visible satellite telemetry and control arcs with annotation commands and numbers. During the lunar calibration maneuver observation, the star sensor is not affected by stray light.

[0051] Preferably, the module M2 includes:

[0052] Module M2.1: Calculates the position vector of the Moon in the instantaneous geocentric mean equatorial coordinate system, denoted as the Moon vector; the origin of the instantaneous geocentric mean equatorial coordinate system is located at the Earth's center of mass, the X-axis points to the mean vernal equinox, the Z-axis points to the mean north celestial pole, and the Y-axis is perpendicular to the X-axis and Z-axis, forming a right-handed rectangular coordinate system;

[0053] Module M2.2: Transforms the lunar vector to the satellite body coordinate system through coordinate transformation and normalizes the lunar vector to (Xb, Yb, Zb), where Xb, Yb, and Zb represent the normalized X-axis coordinate, Y-axis coordinate, and Z-axis coordinate in the satellite body coordinate system, respectively.

[0054] Module M2.3: Calculate the elevation angle α and azimuth angle β of the star and moon vectors at time TO;

[0055] In module M2.3:

[0056] At time TO, the elevation angle α of the celestial-lunar vector pointing from the satellite's center of mass to the center of the moon is:

[0057] α = acos(Yb)

[0058] The lunar vector is on the +Yb side, and the elevation angle α is defined as positive, and vice versa;

[0059] The azimuth angle β of the star-moon vector at time TO is:

[0060] β = atan2(Xb, Zb)

[0061] When the projection of the lunar vector onto the XbObZb plane of the satellite body coordinate system is in the +Xb half-plane, the azimuth angle β is positive, and vice versa; where Ob represents the origin of the satellite body coordinate system.

[0062] Preferably, in module M4: if the azimuth angle β = 0 at time TO, then a delayed execution note is generated and uploaded: allowing lunar calibration observation; the note package contains the time when the lunar calibration observation maneuver control begins, i.e., time TO; if the azimuth angle β ≠ 0 at time TO, then the time Tj for the pitch attitude maneuver to β = 0 is calculated based on the magnitude of β, thereby determining TO-Tj as the delayed note execution time for the pitch attitude maneuver, and a delayed execution note is generated and uploaded: allowing lunar calibration observation, including pitch maneuver; the note package contains the start execution time of the pitch attitude maneuver, TO-Tj, and the lunar calibration observation is automatically executed after the maneuver is completed;

[0063] The pitch attitude maneuver to β=0 is performed using a pre-planning and post-tracking approach, ensuring that the pitch angular velocity relative to the orbital coordinate system is near the orbital angular velocity after the maneuver is completed. The initial three-axis attitude angles are (Φ, θ, Ψ), and the three-axis attitude angular velocities are... Φ, θ, Ψ All are near zero; the three-axis attitude angles after the maneuver are (Φ′, θ′, Ψ′), and the three-axis attitude angular velocities are... Φ′、Ψ′、 Both are near zero, and θ′ is near β at time TO. Near the orbital angular velocity; "near" means the difference is less than a preset numerical range.

[0064] According to the present invention, a zero-momentum satellite lunar calibration observation and control system is provided, which adds a periodic attitude offset function during satellite pitch attitude maneuvers and adds a sinusoidal offset amount to the pitch attitude reference for lunar calibration observation. By measuring the satellite orbit, attitude, and the star-moon vector pointing from the satellite's center of mass to the center of the moon, the azimuth of the moon in the satellite's body coordinate system is calculated and determined, which is used as the reference for lunar calibration observation. The star-moon vector is always located in the YbObZb plane of the satellite's body coordinate system through attitude control.

[0065] Compared with the prior art, the present invention has the following beneficial effects:

[0066] 1. This invention proposes to add a sinusoidal offset to the pitch attitude reference for lunar calibration observation, which can effectively avoid the risk caused by a large deviation between the satellite attitude reference and the optical axis of the payload, and ensure that the payload can achieve lunar calibration observation and that the observation time meets the requirements. That is, the lunar observation time is ensured to meet the requirements by adjusting the pitch attitude offset.

[0067] 2. The attitude maneuver in this invention adopts a planning-then-tracking approach, which is suitable for working conditions where the attitude angular velocity is not zero before and after the maneuver, thus shortening the stabilization time after the maneuver and improving the efficiency of load observation.

[0068] 3. This invention determines the Moon's position in the satellite's coordinate system by measuring satellite orbit, attitude, and calculating the star-moon vector (the vector pointing from the satellite's center of mass to the Moon's center). This position is used as the benchmark for lunar calibration observation. Attitude control ensures that the star-moon vector is always located within the YbObZb plane of the satellite's coordinate system. Attached Figure Description

[0069] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0070] Figure 1 This is a schematic diagram of the lunar observation process provided by the present invention.

[0071] Figure 2 This is a schematic diagram illustrating the principle of the present invention. Detailed Implementation

[0072] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0073] Although precise measurements of the relative pointing relationship between the payload and the star sensor were performed on the ground, significant deviations exist between the actual and calculated relative pointing angles due to errors in satellite attitude measurement, control, and payload optical axis pointing. These errors are compounded by the influence of forces in the vibration environment during the launch vehicle's active phase and thermal factors during the on-orbit flight phase. Furthermore, if control is achieved using the attitude reference output from the star sensor, the lunar apparent angle is only 0.5°, and the payload's instantaneous field of view is only 0.6°. This could result in the payload failing to scan the Moon, leading to failed lunar calibration observations. Therefore, achieving long-term lunar calibration observations through satellite pitch attitude maneuvers in orbit carries considerable risk.

[0074] To address the pointing deviation issue during lunar calibration observations, this invention proposes adding a periodic attitude offset function during satellite pitch maneuvers. This adds a sinusoidally varying offset to the pitch attitude reference for lunar calibration observations, ensuring the payload achieves lunar calibration observations while meeting time requirements and mitigating the risks associated with large pointing deviations. Specifically, considering the satellite lacks a lunar sensor and cannot directly measure its attitude relative to the moon, the moon's position in the satellite's body coordinate system is determined through satellite orbit measurement, attitude measurement, and calculation of the lunar-satellite vector (the vector pointing from the satellite's center of mass to the moon's center). This position serves as the reference for lunar calibration observations, and attitude control ensures the lunar-satellite vector remains within the YbObZb plane of the satellite's body coordinate system.

[0075] The following is in conjunction with the appendix Figure 1 The steps of this invention will be described in detail.

[0076] A zero-momentum satellite lunar calibration observation control method provided by the present invention includes:

[0077] 1) The ground determines the start time TO for lunar calibration observation based on factors such as the lunar phase angle, the satellite's location in the illuminated or shadowed area, and the satellite's telemetry and control arc. Specifically, before lunar observation, it is confirmed that the satellite has completed all remaining on-orbit tests and is in good working order. The start time TO for lunar calibration observation is chosen around the full moon (avoiding a lunar phase angle of 0°, which is the angle formed by the line connecting the payload and the moon, and the line connecting the moon and the sun). The start time TO and end time Tend are preferably chosen when the satellite is in the shadowed area. There are visible telemetry and control arcs with annotation commands and numbers before and after the TO time. During the lunar calibration maneuver observation, the star sensor is not affected by stray light such as ground atmospheric light.

[0078] 2) Calculate the position vector (Xe, Ye, Ze) of the Moon in the instantaneous geocentric equatorial coordinate system, denoted as the Moon vector;

[0079] Where Xe, Ye, and Ze represent the X-axis coordinate, Y-axis coordinate, and Z-axis coordinate in the instantaneous geocentric mean equatorial coordinate system, respectively;

[0080] The origin of the instantaneous geocentric mean equatorial coordinate system is located at the Earth's center of mass, the X-axis points to the mean vernal equinox, the Z-axis points to the mean north celestial pole, and the Y-axis is perpendicular to the X-axis and Z-axis, forming a right-handed rectangular coordinate system;

[0081] 3) Transform the lunar vector to the satellite's body coordinate system using coordinate transformation, and normalize the lunar vector to (Xb, Yb, Zb), i.e., Xb 2 +Yb 2 +Zb 2 =1;

[0082] Where Xb, Yb, and Zb represent the normalized X-axis coordinates, Y-axis coordinates, and Z-axis coordinates in the satellite body coordinate system, respectively;

[0083] The origin of the satellite's coordinate system is the satellite's geometric center. The X-axis is the direction of the satellite's flight, the Z-axis is the square from the Earth's center of mass to the satellite's center of mass, and the Y-axis is perpendicular to the X-axis and Z-axis, forming a right-handed rectangular coordinate system with coordinates Xo, Yo, and Zo.

[0084] 4) Ground-based calculation of the lunar-satellite vector elevation angle α at time TO; specifically, the calculation of the lunar-satellite vector elevation angle α from the satellite's center of mass to the moon's center at time TO:

[0085] α = acos(Yb)

[0086] The lunar vector is on the +Yb side, with the elevation angle α defined as positive and the opposite as negative.

[0087] 5) Ground calculation of the azimuth angle β of the star and moon vector at time TO; Calculation of the azimuth angle β of the star and moon vector at time TO:

[0088] β = atan2(Xb, Zb)

[0089] When the projection of the lunar vector onto the XbObZb plane of the satellite body coordinate system lies in the +Xb half-plane, the azimuth angle β is positive, and vice versa; where Ob represents the origin of the satellite body coordinate system.

[0090] 6) Determine the periodic offset, i.e., the pitch attitude offset angle γ:

[0091] γ=εsin(2πt / τ)

[0092] Where ε is the peak value of the bias, t is the time variable, and τ is the bias period. The bias period is selected as 1 / N of the entire lunar observation time or the lunar observation duration (N is a positive integer), and the bias parameters ε and τ are noted above.

[0093] For example, if the instantaneous field of view of the payload is 0.6°, the peak offset ε = 0.3°, and the offset period τ = 10 min, then the pitch attitude offset angle γ is:

[0094] γ = 0.0052sin(2πt / 600)

[0095] Generate pitch attitude offset measurement data and upload it;

[0096] 7) If the azimuth angle β of the lunar vector is 0 at time TO, then generate and add the delay execution note "Allow lunar calibration observation". The note package contains the time when the lunar calibration observation maneuver control starts, i.e., time TO; the system starts the lunar calibration observation maneuver control at time TO.

[0097] If the azimuth angle β≠0 at time TO, the ground calculates the time Tj for the pitch attitude maneuver to β=0 based on the magnitude of β. Thus, TO-Tj is determined as the execution time of the "pitch attitude maneuver" delay note. The delay execution note "Allow lunar calibration observation (including pitch maneuver)" is generated and added. The note package contains the start time of the pitch attitude maneuver, TO-Tj. After the maneuver is completed, the "lunar calibration observation" is automatically executed.

[0098] 8) Satellite pitch and attitude maneuvers include two scenarios: attitude maneuver control during the switch from Earth observation mode to lunar observation mode, and lunar tracking maneuver control during lunar calibration observation mode. Both are initiated through delayed execution of annotations, allowing for flexible maneuver timing and independence from the constraint of the lunar-satellite vector azimuth angle β = 0. When β ≠ 0, the satellite first performs an attitude maneuver to β = 0 before autonomously entering lunar tracking control; when β = 0, it directly enters lunar tracking control. Since the satellite's pitch angular velocity is not zero (close to its orbital angular velocity) during lunar tracking maneuver control, a pre-planning and post-tracking approach is used to complete both scenarios, shortening the attitude convergence and stabilization time after the maneuver and improving observation efficiency.

[0099] The pitch attitude maneuver to β=0 is performed using a pre-planning and then tracking approach, ensuring that the pitch angular velocity relative to the orbital coordinate system is near the orbital angular velocity after the maneuver is completed. This allows for immediate lunar observation after the maneuver. Specifically, the pitch attitude maneuver to β=0 is performed using a pre-planning and then tracking approach. The initial three-axis attitude angles are (Φ, θ, Ψ), and the three-axis attitude angular velocities are... Φ, θ, Ψ All are near zero; the three-axis attitude angles after the maneuver are (Φ′, θ′, Ψ′), and the three-axis attitude angular velocities are... Φ′、Ψ′、 Both are near zero, and θ′ is near β at time TO. Near the orbital angular velocity; "near" means that the difference is less than a preset numerical range, for example, Φ near zero means that the difference between Φ and zero falls within the preset numerical range;

[0100] Φ, θ, and Ψ represent the satellite roll attitude angle, satellite pitch attitude angle, and satellite yaw attitude angle at the start of the pitch attitude maneuver, respectively.

[0101] These represent the satellite roll rate, satellite pitch rate, and satellite yaw rate at the start of the pitch attitude maneuver, respectively.

[0102] Φ′, θ′, and Ψ′ represent the satellite roll attitude angle, satellite pitch attitude angle, and satellite yaw attitude angle after the pitch attitude maneuver is completed, respectively.

[0103] These represent the satellite roll rate, satellite pitch rate, and satellite yaw rate after the pitch attitude maneuver is completed, respectively.

[0104] 9) The pitch attitude reference during lunar observation is (θ-β-γ), where θ is the satellite pitch attitude angle, which is given in real time by the satellite attitude determination; β is the azimuth angle of the lunar-satellite vector, which is given in real time by the onboard computer according to steps 2), 3), and 5); and γ is the pitch attitude offset angle, which is given in real time according to step 6).

[0105] 10) After lunar observation is completed, a pitch maneuver is performed for return. The return process adopts a planning-then-tracking approach, ensuring that the pitch angular velocity relative to the orbital coordinate system is near 0 upon return. Earth observation resumes immediately upon return. Specifically, after lunar observation, the same planning-then-tracking maneuver is used to resume Earth observation. The initial three-axis attitude angles are (Φ1, θ1, Ψ1), and the three-axis attitude angular velocities are... Φ1、Ψ1、 Both are near zero, θ1 is not zero, and is given by the star sensor measurement. The angular velocity is not zero and is near the orbital angular velocity; the three-axis attitude angles after the maneuver are (Φ1′, θ1′, Ψ1′), and the three-axis attitude angular velocities are... Φ1′、θ1′、Ψ1′、 All are near zero;

[0106] Φ1, θ1, and Ψ1 represent the satellite roll attitude angle, satellite pitch attitude angle, and satellite yaw attitude angle at the start of the Earth observation recovery maneuver, respectively.

[0107] These represent the satellite roll rate, satellite pitch rate, and satellite yaw rate at the start of the pitch attitude recovery maneuver.

[0108] Φ1′, θ1′, and Ψ1′ represent the satellite roll attitude angle, satellite pitch attitude angle, and satellite yaw attitude angle, respectively, after the completion of the Earth observation maneuver.

[0109] These represent the satellite roll rate, satellite pitch rate, and satellite yaw rate at the start of the pitch attitude maneuver after the Earth observation recovery maneuver is completed;

[0110] 11) After the lunar calibration observation task is completed, note "Lunar calibration observation prohibited" to ensure that the control subsystem does not enter the lunar observation mode when in Earth observation mode.

[0111] The present invention will now be described in more detail.

[0112] In this invention, the satellite payload employs a cross-orbit circular scanning mode, with the rotation axis along the satellite's flight direction +X. The field of view along the trajectory is only 0.6°, and all channels are equipped with cold-space signal observation capabilities. The satellite control subsystem adopts a three-axis zero-momentum control scheme, enabling pitch attitude maneuvers at any angle within 0–360°. Reliable lunar calibration observations are achieved through on-orbit offset of the pitch attitude reference. During normal Earth observation, the Earth signal acquisition range is ±55.1° on both sides of the nadir point. On the side of the satellite facing away from the sun, with +Yb as the starting position of 0°, the cold-space signal acquisition range, i.e., the lunar calibration observation range, is 19°–22°. The scanning start position can be changed in orbit via calibration, thus altering the Earth signal acquisition range to the cold-space signal acquisition range.

[0113] The pitch attitude reference for lunar tracking maneuver control is (θ-β-γ), where θ is the satellite pitch attitude angle, β is the lunar vector azimuth angle, and γ is the pitch attitude offset angle. The lunar vector azimuth angle is determined using a combination of analytical and fitting methods. When the satellite requires long-term lunar calibration observations, a high-precision initial value of the lunar vector is uploaded, and the analytical method is used. When the satellite only needs lunar calibration observations within one day, the position of the lunar vector within one day is given through JPL / DE ephemeris simulation, and the fitting method is used. The pitch attitude offset angle γ is a small angle with sinusoidal variation. The angle is 0° at the start of lunar observation, and the peak value is selected between 0.1° and 0.3° based on actual conditions. The period is the lunar observation period; if the lunar observation time is long, the period can be 1 / N of the lunar observation duration (N is a positive integer).

[0114] This invention also provides a zero-momentum satellite lunar calibration observation control system. This system can be implemented by executing the process steps of the zero-momentum satellite lunar calibration observation control method. Therefore, the zero-momentum satellite lunar calibration observation control method can be understood as a preferred embodiment of the zero-momentum satellite lunar calibration observation control system. Specifically, according to the zero-momentum satellite lunar calibration observation control system provided by this invention, it includes:

[0115] Module M1: Determines the start time TO of the lunar calibration observation;

[0116] Module M2: Calculates the elevation angle α and azimuth angle β of the star and moon vectors at time TO;

[0117] Module M3: Determines the periodic offset, i.e., the pitch attitude offset angle γ.

[0118] γ=εsin(2πt / τ)

[0119] Where ε is the peak value of the bias, t is the time variable, and τ is the bias period; the bias period is selected as 1 / N of the entire monthly observation time or the monthly observation duration, where N is a positive integer, and the bias parameters ε and τ are noted above.

[0120] Module M4: Based on the azimuth angle β of the lunar and celestial vectors at time TO, it generates and adds the delay execution note number, allowing lunar calibration observations and determining pitch attitude maneuvers;

[0121] The elevation attitude reference during lunar observation is (θ-β-γ), where θ is the satellite elevation attitude angle, which is given in real time by the satellite attitude determination; β is the azimuth angle of the lunar-satellite vector, which is given in real time by the onboard computer; and γ is the elevation attitude offset angle, which is given in real time by calculation.

[0122] Module M5: After completing the lunar calibration observation mission, perform pitch maneuver to return. Note: Lunar calibration observation is prohibited; do not enter lunar observation during Earth observation.

[0123] In module M1, the ground determines the start time TO of the lunar calibration observation based on the lunar phase angle, the interval between the illuminated and shadowed areas where the satellite is located, and the conditions of the satellite's telemetry and control arc.

[0124] The start time of the lunar calibration observation TO is chosen around the full moon to avoid the lunar phase angle of 0°. The lunar phase angle is the angle formed by the line connecting the payload and the moon and the line connecting the moon and the sun.

[0125] The start time (TO) and end time (Tend) of the lunar calibration observation are chosen when the satellite is in shadow. Before and after the time (TO), there are visible satellite telemetry and control arcs with annotation commands and numbers. During the lunar calibration maneuver observation, the star sensor is not affected by stray light.

[0126] The module M2 includes:

[0127] Module M2.1: Calculates the position vector of the Moon in the instantaneous geocentric mean equatorial coordinate system, denoted as the Moon vector; the origin of the instantaneous geocentric mean equatorial coordinate system is located at the Earth's center of mass, the X-axis points to the mean vernal equinox, the Z-axis points to the mean north celestial pole, and the Y-axis is perpendicular to the X-axis and Z-axis, forming a right-handed rectangular coordinate system;

[0128] Module M2.2: Transforms the lunar vector to the satellite body coordinate system through coordinate transformation and normalizes the lunar vector to (Xb, Yb, Zb), where Xb, Yb, and Zb represent the normalized X-axis coordinate, Y-axis coordinate, and Z-axis coordinate in the satellite body coordinate system, respectively.

[0129] Module M2.3: Calculate the elevation angle α and azimuth angle β of the star and moon vectors at time TO;

[0130] In module M2.3:

[0131] At time TO, the elevation angle α of the celestial-lunar vector pointing from the satellite's center of mass to the center of the moon is:

[0132] α = acos(Yb)

[0133] The lunar vector is on the +Yb side, and the elevation angle α is defined as positive, and vice versa;

[0134] The azimuth angle β of the star-moon vector at time TO is:

[0135] β = atan2(Xb, Zb)

[0136] When the projection of the lunar vector onto the XbObZb plane of the satellite body coordinate system is in the +Xb half-plane, the azimuth angle β is positive, and vice versa; where Ob represents the origin of the satellite body coordinate system.

[0137] Those skilled in the art will understand that, in addition to implementing the system, apparatus, and their modules provided by this invention in purely computer-readable program code, the same program can be implemented 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, apparatus, and their modules provided by this invention can be considered a hardware component, and the modules included therein for implementing various programs can also be considered structures within the hardware component; alternatively, modules for implementing various functions can be considered both software programs implementing the method and structures within the hardware component.

[0138] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A method for lunar calibration observation and control of a zero-momentum satellite, characterized in that, include: Step S1: Determine the start time TO of the lunar calibration observation; Step S2: Calculate the elevation angle α and azimuth angle β of the star and moon vectors at time TO; Step S3: Determine the periodic offset, i.e., the pitch attitude offset angle γ: γ=εsin(2πt / τ) Where ε is the peak value of the bias, t is the time variable, and τ is the bias period; the bias period is selected as 1 / N of the entire monthly observation time or the monthly observation duration, where N is a positive integer, and the bias parameters ε and τ are noted above. Step S4: Based on the azimuth angle β of the lunar and celestial vectors at time TO, generate and add the delay execution note: allow lunar calibration observations; and determine the pitch attitude maneuver; The elevation attitude reference during lunar observation is θ-β-γ, where θ is the satellite elevation attitude angle, which is given in real time by the satellite attitude determination; β is the azimuth angle of the lunar-satellite vector, which is given in real time by the onboard computer; and γ is the elevation attitude offset angle, which is given in real time by calculation. Step S5: After the lunar calibration observation mission is completed, perform pitch maneuver to return. Note: Lunar calibration observation is prohibited; do not enter lunar observation mode during Earth observation. The peak bias ε is selected between 0.1° and 0.3°.

2. The zero-momentum satellite lunar calibration observation and control method according to claim 1, characterized in that, In step S1, the ground determines the start time TO of the lunar calibration observation based on the lunar phase angle, the interval between the illuminated and shadowed areas where the satellite is located, and the conditions of the satellite's tracking and control arc. The start time of the lunar calibration observation TO is chosen around the full moon to avoid the lunar phase angle of 0°. The lunar phase angle is the angle formed by the line connecting the payload and the moon and the line connecting the moon and the sun. The start time (TO) and end time (Tend) of the lunar calibration observation are chosen when the satellite is in shadow. Before and after the time (TO), there are visible satellite telemetry and control arcs with annotation commands and numbers. During the lunar calibration maneuver observation, the star sensor is not affected by stray light.

3. The zero-momentum satellite lunar calibration observation and control method according to claim 1, characterized in that, Step S2 includes: Step S2.1: Calculate the position vector of the Moon in the instantaneous geocentric mean equatorial coordinate system, denoted as the Moon vector; the origin of the instantaneous geocentric mean equatorial coordinate system is located at the Earth's center of mass, the X-axis points to the mean vernal equinox, the Z-axis points to the mean north celestial pole, and the Y-axis is perpendicular to the X-axis and Z-axis, forming a right-hand rectangular coordinate system; Step S2.2: Transform the lunar vector to the satellite body coordinate system through coordinate transformation, and normalize the lunar vector to (Xb, Yb, Zb), where Xb, Yb, and Zb represent the normalized X-axis coordinate, Y-axis coordinate, and Z-axis coordinate in the satellite body coordinate system, respectively. Step S2.3: Calculate the elevation angle α and azimuth angle β of the star and moon vectors at time TO; In step S2.3: At time TO, the elevation angle α of the celestial-lunar vector pointing from the satellite's center of mass to the center of the moon is: α = acos(Yb) The lunar vector is on the +Yb side, and the elevation angle α is defined as positive, and vice versa; The azimuth angle β of the star-moon vector at time TO is: β = atan2(Xb, Zb) When the projection of the lunar vector onto the XbObZb plane of the satellite body coordinate system is in the +Xb half-plane, the azimuth angle β is positive, and vice versa; where Ob represents the origin of the satellite body coordinate system.

4. The zero-momentum satellite lunar calibration observation and control method according to claim 1, characterized in that, In step S4: if the azimuth angle β of the lunar vector at time TO is 0, then the delayed execution note is generated and uploaded: allowing lunar calibration observation; the note package contains the time when the lunar calibration observation maneuver control begins, i.e., time TO; if the azimuth angle β≠0 at time TO, then the time Tj for the pitch attitude maneuver to β=0 is calculated based on the magnitude of β, thereby determining TO-Tj as the delayed note execution time for the pitch attitude maneuver, and the delayed execution note is generated and uploaded: allowing lunar calibration observation, including pitch maneuver; the note package contains the start execution time of the pitch attitude maneuver, TO-Tj, and the lunar calibration observation is automatically executed after the maneuver is in place; The pitch attitude maneuver to β=0 is performed using a pre-planning and post-tracking approach, ensuring that the pitch angular velocity relative to the orbital coordinate system is near the orbital angular velocity after the maneuver is completed. The initial three-axis attitude angles are (Φ, θ, Ψ), and the three-axis attitude angular velocities are (Φ, θ, Ψ). , , ), Φ, θ, Ψ, , , All are near zero; the three-axis attitude angles after the maneuver are (Φ′, θ′, Ψ′), and the three-axis attitude angular velocities are ( , , ), Φ′、Ψ′、 , Both are near zero, and θ′ is near β at time TO. Near the orbital angular velocity; "near" means the difference is less than a preset numerical range.

5. A method for lunar calibration observation and control of a zero-momentum satellite, characterized in that, Add a periodic attitude offset function during satellite pitch attitude maneuvers, and add a sinusoidal offset to the pitch attitude reference for lunar calibration observations; calculate and determine the Moon's azimuth in the satellite's body coordinate system through satellite orbit measurement, attitude measurement, and the star-moon vector pointing from the satellite's center of mass to the Moon's center, and use this as the reference for lunar calibration observations; ensure that the star-moon vector is always located within the YbObZb plane of the satellite's body coordinate system through attitude control. Determine the periodic offset, i.e., the pitch attitude offset angle γ: γ=εsin(2πt / τ); where ε is the peak value of the offset, t is the time variable, and τ is the offset period; the offset period is selected as 1 / N of the entire lunar observation time or the lunar observation duration, where N is a positive integer, and the offset parameters ε and τ are noted above. The peak value of the bias is selected between 0.1° and 0.3°.

6. A zero-momentum satellite lunar calibration and observation control system, characterized in that, include: Module M1: Determines the start time TO of the lunar calibration observation; Module M2: Calculates the elevation angle α and azimuth angle β of the star and moon vectors at time TO; Module M3: Determines the periodic offset, i.e., the pitch attitude offset angle γ. γ=εsin(2πt / τ) Where ε is the peak value of the bias, t is the time variable, and τ is the bias period; the bias period is selected as 1 / N of the entire monthly observation time or the monthly observation duration, where N is a positive integer, and the bias parameters ε and τ are noted above. Module M4: Based on the azimuth angle β of the lunar and celestial vectors at time TO, it generates and adds the delay execution note number, allowing lunar calibration observations and determining pitch attitude maneuvers; The elevation attitude reference during lunar observation is θ-β-γ, where θ is the satellite elevation attitude angle, which is given in real time by the satellite attitude determination; β is the azimuth angle of the lunar-satellite vector, which is given in real time by the onboard computer; and γ is the elevation attitude offset angle, which is given in real time by calculation. Module M5: After completing the lunar calibration observation mission, perform pitch maneuver to return. Note: Lunar calibration observation is prohibited; do not enter lunar observation during Earth observation. The bias parameter ε is selected between 0.1° and 0.3°.

7. The zero-momentum satellite lunar calibration and observation control system according to claim 6, characterized in that, In module M1, the ground determines the start time TO of the lunar calibration observation based on the lunar phase angle, the interval between the illuminated and shadowed areas where the satellite is located, and the conditions of the satellite's telemetry and control arc. The start time of the lunar calibration observation TO is chosen around the full moon to avoid the lunar phase angle of 0°. The lunar phase angle is the angle formed by the line connecting the payload and the moon and the line connecting the moon and the sun. The start time (TO) and end time (Tend) of the lunar calibration observation are chosen when the satellite is in shadow. Before and after the time (TO), there are visible satellite telemetry and control arcs with annotation commands and numbers. During the lunar calibration maneuver observation, the star sensor is not affected by stray light.

8. The zero-momentum satellite lunar calibration and observation control system according to claim 6, characterized in that, The module M2 includes: Module M2.1: Calculates the position vector of the Moon in the instantaneous geocentric mean equatorial coordinate system, denoted as the Moon vector; the origin of the instantaneous geocentric mean equatorial coordinate system is located at the Earth's center of mass, the X-axis points to the mean vernal equinox, the Z-axis points to the mean north celestial pole, and the Y-axis is perpendicular to the X-axis and Z-axis, forming a right-handed rectangular coordinate system; Module M2.2: Transforms the lunar vector to the satellite body coordinate system through coordinate transformation and normalizes the lunar vector to (Xb, Yb, Zb), where Xb, Yb, and Zb represent the normalized X-axis coordinate, Y-axis coordinate, and Z-axis coordinate in the satellite body coordinate system, respectively. Module M2.3: Calculate the elevation angle α and azimuth angle β of the star and moon vectors at time TO; In module M2.3: At time TO, the elevation angle α of the celestial-lunar vector pointing from the satellite's center of mass to the center of the moon is: α = acos(Yb) The lunar vector is on the +Yb side, and the elevation angle α is defined as positive, and vice versa; The azimuth angle β of the star-moon vector at time TO is: β = atan2(Xb, Zb) When the projection of the lunar vector onto the XbObZb plane of the satellite body coordinate system is in the +Xb half-plane, the azimuth angle β is positive, and vice versa; where Ob represents the origin of the satellite body coordinate system.

9. The zero-momentum satellite lunar calibration and observation control system according to claim 6, characterized in that, In module M4: if the azimuth angle β=0 at time TO, then the delayed execution note is generated and uploaded: allowing lunar calibration observation; the note package contains the time when the lunar calibration observation maneuver control begins, i.e., time TO; if the azimuth angle β≠0 at time TO, then the time Tj for the pitch attitude maneuver to β=0 is calculated based on the magnitude of β, thereby determining TO-Tj as the delayed note execution time for the pitch attitude maneuver, and the delayed execution note is generated and uploaded: allowing lunar calibration observation, including pitch maneuver; the note package contains the start execution time of the pitch attitude maneuver, TO-Tj, and the lunar calibration observation is automatically executed after the maneuver is in place; The pitch attitude maneuver to β=0 is performed using a pre-planning and post-tracking approach, ensuring that the pitch angular velocity relative to the orbital coordinate system is near the orbital angular velocity after the maneuver is completed. The initial three-axis attitude angles are (Φ, θ, Ψ), and the three-axis attitude angular velocities are (Φ, θ, Ψ). , , ), Φ, θ, Ψ, , , All are near zero; the three-axis attitude angles after the maneuver are (Φ′, θ′, Ψ′), and the three-axis attitude angular velocities are ( , , ), Φ′、Ψ′、 , Both are near zero, and θ′ is near β at time TO. Near the orbital angular velocity; "near" means the difference is less than a preset numerical range.

10. A zero-momentum satellite lunar calibration and observation control system, characterized in that, Add a periodic attitude offset function during satellite pitch attitude maneuvers, and add a sinusoidal offset to the pitch attitude reference for lunar calibration observations; calculate and determine the Moon's azimuth in the satellite's body coordinate system through satellite orbit measurement, attitude measurement, and the star-moon vector pointing from the satellite's center of mass to the Moon's center, and use this as the reference for lunar calibration observations; ensure that the star-moon vector is always located within the YbObZb plane of the satellite's body coordinate system through attitude control. Determine the periodic offset, i.e., the pitch attitude offset angle γ: γ=εsin(2πt / τ); where ε is the peak value of the offset, t is the time variable, and τ is the offset period; the offset period is selected as 1 / N of the entire lunar observation time or the lunar observation duration, where N is a positive integer, and the offset parameters ε and τ are noted above. The peak value of the bias is selected between 0.1° and 0.3°.

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