Spacecraft attitude orbit navigation initial value on-orbit establishment method and device

Abstract

The application provides a spacecraft attitude orbit navigation initial value on-orbit establishment method and device, and belongs to the technical field of spacecraft navigation, which comprises the following steps: designing a navigation system single machine product configuration scheme and a single machine installation layout scheme; judging whether the spacecraft has been successfully separated from a mother platform; damping the angular velocity of the spacecraft; performing slow rotation search around two axes of the spacecraft body to preliminarily establish an inertial system attitude initial value; performing slow rotation search around two axes of the spacecraft body to preliminarily establish an earth-fixed system position and velocity initial value; performing attitude maneuvering based on the inertial system attitude initial value and the earth-fixed system position and velocity initial value, and adjusting the attitude to a final stable pointing direction which can simultaneously capture the attitude and the position and velocity; and calculating the position and velocity and the attitude in an expected reference system according to the attitude and the position and velocity, and taking the position and velocity and the attitude as the final inertial navigation initial value. The application can make the spacecraft and the released mother platform not need to be equipped with information transmission hardware and software interfaces, and can realize the capability of establishing the inertial navigation attitude and position and velocity initial value on the spacecraft.

Description

Technical Field

[0001] This invention belongs to the field of spacecraft navigation technology, specifically relating to an on-orbit method and apparatus for establishing initial values ​​for spacecraft attitude and orbit navigation. Background Technology

[0002] Obtaining spacecraft attitude and orbital navigation information through navigation technologies and systems is a fundamental prerequisite for spacecraft flight control. Inertial navigation is currently the most widely used navigation technology in engineering, and almost all spacecraft incorporate it as a basic configuration. Inertial navigation has the significant advantage of not relying on external artificial information sources. However, since it navigates by measuring angle and velocity increments and then integrating and accumulating them, obtaining initial values ​​of angle, position, and velocity is a necessary condition for completing navigation calculations. How to obtain the initial navigation state is called initial alignment in the field of inertial navigation technology, and it is one of the important research topics in inertial navigation.

[0003] There are two main methods for establishing initial values ​​for spacecraft inertial navigation: One method involves the ground-based telemetry, tracking, and command (TT&C) system calculating initial attitude values ​​before launch, either through third-party measurement methods or by the spacecraft's own sensors based on its stationary state on the ground. These initial values ​​are then combined with precise geographical location and time information from the launch site to calculate initial position and velocity values. These values ​​are then encoded into the spacecraft's navigation system. The spacecraft remains powered and continuously navigates throughout the launch phase, and its attitude and position / velocity are recorded at launch separation as initial values ​​for on-orbit navigation. The other method uses a navigation reference transfer mechanism. At launch separation, the launch vehicle's navigation system transmits initial attitude and position / velocity values ​​to the spacecraft's navigation system. Upon receiving these initial values, the spacecraft uses them as initial inertial navigation values ​​and initiates on-orbit navigation. The first method requires the spacecraft to be powered before launch and continuously powered and performing navigation calculations throughout the launch phase. The second method requires data interaction between the spacecraft and launch vehicle, necessitating hardware and software interfaces, which increases the weight, cost, and complexity of the entire system and reduces its reliability.

[0004] For spacecraft that only have a mechanical interface with the launch vehicle or release platform (hereinafter referred to as the mother platform), power is applied only after separation from the mother platform. At the time of power-up, the spacecraft has no prior information, which falls under the category of cold start. Under such circumstances, existing technical solutions cannot establish initial attitude and orbit values. One potential approach is to equip the spacecraft with products such as astronomical navigation equipment and GNSS navigation receivers to perform self-searching for initial navigation values. However, due to the influence of separation interference, the spacecraft is often in an unstable and uncertain state of tumbling at a large angular velocity after release. In order for each navigation product to work normally and output navigation information, it requires a stable working environment and certain pointing requirements. How to design a reasonably configured navigation system and design a logically clear and highly reliable initial navigation value establishment algorithm is the key problem to be solved by this invention. The application of this technology has important practical significance for improving the autonomy and reliability of space missions. Summary of the Invention

[0005] To address the aforementioned technical problems, this invention provides a method and apparatus for establishing initial values ​​of attitude, orbit, and navigation of a low-Earth orbit spacecraft in orbit. Using this apparatus and method, the spacecraft can establish its initial values ​​of inertial navigation attitude, position, and velocity in orbit without the need for a hardware or software interface for information transmission between the spacecraft and the release platform.

[0006] To achieve the above objectives, the present invention adopts the following technical solution:

[0007] A method for establishing initial values ​​for spacecraft attitude and orbit navigation in orbit on-orbit includes the following steps:

[0008] Step S110: Design a single-unit product configuration scheme for the navigation system;

[0009] Step S120: Design a single-unit installation layout scheme for the navigation system;

[0010] Step S130: After completing the single-unit configuration and layout, the spacecraft determines whether it has successfully separated from the mother platform;

[0011] Step S140: After confirming separation, the spacecraft dampens its own angular rate;

[0012] Step S150: The spacecraft performs a slow rotation search around its two axes to initially establish the initial attitude values ​​of the inertial frame;

[0013] Step S160: The spacecraft performs a slow rotation search around its two axes to initially establish the initial values ​​of the Earth-fixed position and velocity.

[0014] Step S170: Based on the inertial frame attitude navigation calculation results at the current moment and the initial value of the Earth-fixed position velocity, perform attitude maneuvers to adjust the attitude to a final stable direction that can simultaneously capture attitude and position velocity.

[0015] Step S180: Based on the attitude and position velocity results from step S170, calculate the position velocity and attitude results in the desired reference frame as the initial values ​​for the final inertial navigation.

[0016] An on-orbit device for establishing initial values ​​for spacecraft attitude and orbit navigation includes the following modules:

[0017] The inertial navigation module is used to measure the three-axis angle increments and perform inertial attitude navigation calculations.

[0018] The starlight attitude capture and measurement module is used to capture and identify stars, and to calculate the spacecraft's attitude relative to the inertial frame.

[0019] The satellite navigation module is used to receive satellite navigation signals and calculate the spacecraft's position and velocity.

[0020] The control module is used to control the spacecraft to achieve angular rate damping, slow rotation of the body around a specified axis, or pointing to the desired attitude direction based on the execution results of the calculation module.

[0021] The calculation module is used to execute the on-orbit establishment method for initial values ​​of spacecraft attitude, orbit, and navigation.

[0022] An electronic device includes a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the on-orbit establishment method for initial values ​​of spacecraft attitude and orbit navigation.

[0023] A non-transitory computer-readable storage medium storing a computer program that, when executed by a processor, implements the aforementioned method for establishing initial values ​​for spacecraft attitude, orbit, and navigation in orbit.

[0024] Beneficial effects:

[0025] This invention discloses a method and apparatus for establishing initial values ​​of spacecraft attitude, orbit, and navigation in orbit. Using this apparatus and method, the spacecraft can establish initial values ​​of inertial navigation attitude, position, and velocity in orbit without the need for a hardware or software interface for information transmission between the spacecraft and the release platform. Attached Figure Description

[0026] The specific embodiments of the present invention will be further described in detail below with reference to the accompanying drawings, wherein:

[0027] Figure 1 This is a flowchart of the on-orbit establishment method for initial values ​​of spacecraft attitude and orbit navigation according to an embodiment of the present invention;

[0028] Figure 2 This is a schematic diagram of the single-unit installation layout;

[0029] Figure 3The curve shows the change in attitude angular velocity under rate damping.

[0030] Figure 4 The curves showing the changes in the optical axis direction of the star sensor and the visible sky region of stars;

[0031] Figure 5 The curve representing the angle between the slow-rotation axis and the negative normal of the theoretical initial orbital plane in a star-sensitive search.

[0032] Figure 6 The curves showing the changes in the visible sky area of ​​the GNSS antenna beam center direction and navigation signal during the attitude adjustment phase are shown.

[0033] Figure 7 The curves showing the changes in the visible sky area of ​​the navigation signal along the beam center direction of the GNSS antenna in the slow-rotation phase;

[0034] Figure 8 This is the attitude angle variation curve for the three-axis stable segment relative to the ground;

[0035] Figure 9 The curve shows the attitude angular velocity variation during the three-axis ground-based stable segment.

[0036] Figure 10 The curves show the changes in the optical axis of the star sensor and the beam center pointing of the GNSS antenna during the three-axis stable phase of the ground. Detailed Implementation

[0037] Exemplary embodiments of the present disclosure will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art.

[0038] like Figure 1 As shown, the on-orbit establishment method for initial values ​​of spacecraft attitude and orbit navigation according to the present invention includes the following steps:

[0039] Step S110: Design the configuration scheme for a standalone navigation system product, as follows:

[0040] Step S110-1: To realize the inertial navigation function, configure an inertial measurement device composed of gyroscopes with three-axis angle increment measurement capability;

[0041] Step S110-2: Configure the star sensor to realize the attitude search function;

[0042] Step S110-3: To realize the position and velocity search function, configure a GNSS navigation receiver, and the configured navigation receiver should have a receiving sensitivity corresponding to the navigation signal strength at the orbital altitude at the time of release;

[0043] Step S110-4: To realize the attitude and position velocity signal search function, a low-speed attitude control actuator is configured. Specifically, a jet thruster system, flywheel assembly or control torque gyroscope group or other angular momentum exchange devices can be selected. The configured actuator should have three-axis attitude maneuver control capability.

[0044] Step S120: Design a standalone installation layout scheme for the navigation system, as follows:

[0045] Step S120-1: Inertial Measurement Unit Installation Layout Design. The three-axis gyroscope is installed and laid out according to conventional methods, ensuring that each axis of the gyroscope corresponds to and is parallel to the axes of the spacecraft system.

[0046] Step S120-2: Star Sensor Installation Layout Design. The main constraints on star sensor installation are that its optical axis should point towards the starry sky and should not be obstructed by Earth or other near-field spacecraft such as the mother platform. Furthermore, the star sensor's optical axis should be far from the engine plume interference zone. Based on these constraints, considering that the spacecraft's main flight attitude is generally that the Z-axis points towards the Earth's center and the X-axis points forward within the orbital plane, and that the main source of plume interference risk is the high-thrust orbital control engine, whose possible installation direction is generally along the positive and negative directions of the spacecraft's axes, the design direction for the star sensor is to have its optical axis at a 45° angle to the XOY plane of the spacecraft system and biased towards -Z. A recommended specific direction is to have the optical axis along the bisectors of the +X and -Z angles.

[0047] Step S120-3: GNSS navigation receiver installation layout design. The main constraint for GNSS navigation receiver installation is that the beam center of its signal receiving antenna should point in the direction where navigation signals exist. For low-Earth orbit spacecraft, the center of the receivable signal source is the local zenith direction. Therefore, when only one antenna is configured, it can be designed so that the antenna beam center points in the -Z direction of the spacecraft system. When multiple antennas are configured, the beam center of one of the antennas should point in the -Z direction of the spacecraft system.

[0048] Step S120-4: Installation layout design of low-speed attitude control actuator. Since attitude control is generally an essential function and system for spacecraft to complete their main missions, this invention does not specifically design the installation layout of its actuator; it only requires three-axis attitude control capability.

[0049] Step S130: After completing the individual unit configuration and layout, the spacecraft determines whether it has successfully separated from the mother platform, as follows:

[0050] Successful separation is a necessary prerequisite for navigation initial value search. The separation judgment method can be designed according to the specific configuration of individual spacecraft units and components. When the spacecraft is equipped with limit switches, a reliable separation signal can be directly provided by the limit switches; otherwise, if an attitude control engine is equipped, the separation judgment can be made by starting the attitude control engine for a test spray and measuring the angular velocity increment generated by the excitation; finally, when no attitude control engine is equipped, other attitude control actuators such as flywheel groups or control torque gyroscope groups must be equipped. These actuators can be used to excite the angular velocity, and the separation judgment can be completed based on the measured actual angular velocity increment.

[0051] Step S140: After confirming separation, the spacecraft dampens its own angular rate, as follows:

[0052] Using gyroscopes to perform three-axis angle increment measurements on spacecraft And calculate the corresponding triaxial angular velocities. :

[0053] ;

[0054] in, The sampling period for gyroscope measurements. These are the angular increments of the spacecraft around its X, Y, and Z axes. These are the angular velocities of the spacecraft around its X, Y, and Z axes, respectively. The command torque is calculated using the following proportional control law based on the angular velocity. calculate:

[0055] ;

[0056] in, , , This is the proportional control coefficient; finally, the low-speed attitude control actuator configured by the system completes the actual control torque output and ultimately completes the angular velocity rate damping task.

[0057] Step S150: The spacecraft performs a slow rotation search around its two axes to initially establish the initial attitude values ​​of the inertial frame, as follows:

[0058] Step S150-1: Rotation Axis Design. The inertial frame attitude is established by measuring the starry sky region using a star sensor. When the optical axis is blocked by the Earth or other spacecraft, the star sensor cannot complete the measurement and output the attitude; therefore, rotation is required to search the sky region. In the worst-case scenario, a full-sky search is required, necessitating full rotations around two mutually perpendicular axes. Furthermore, the rotation axis should not coincide with the star sensor's optical axis; otherwise, the pointed-to sky region will remain unchanged, and the search cannot be completed. Considering these factors, the unit direction vector of the star sensor's optical axis is denoted as... The +X axis vector of the spacecraft system is Design of rotating shaft , as follows:

[0059] ;

[0060] Step S150-2: Slow Rotation Angular Velocity Design. Due to the limited dynamic response capability of star sensors, sky searches can only be conducted when the rotation angular velocity does not exceed the dynamic capability of the star sensor; on the other hand, the search angular velocity should not be too slow, ensuring that the visible star area during two orbital searches does not change excessively due to orbital motion. Therefore, considering that the dynamic capability of current star sensors is generally not less than... The spacecraft's orbital period is no less than 90 minutes, and the rotational angular velocity for the designed sky search is... .

[0061] Step S150-3: Perform slow rotation search. Starting from the current attitude, under the action of the attitude maneuver control law, rotate sequentially according to the designed angular velocity. , Slowly rotate one revolution. During the rotation, if the star sensor outputs a valid attitude, the current step is completed and immediately exited; if it rotates around... , If the star sensor still does not output a valid attitude after one slow rotation, the slow rotation search continues with the same rotation direction and angular velocity until a valid attitude is successfully output.

[0062] Around , The attitude control law for the axis slow rotation maneuver is designed as follows:

[0063] ;

[0064] in, The angular velocity of the spacecraft's attitude, measured by the gyroscope. , , This is the proportional control coefficient.

[0065] Step S150-4: Calculate the spacecraft's attitude relative to the inertial frame based on the attitude output by the star sensor and the star sensor's installation matrix relative to the spacecraft, as follows:

[0066] ;

[0067] in, To represent quaternion multiplication, The attitude output by the star sensor. The quaternion corresponding to the star sensor's installation matrix relative to the spacecraft. Let it be its conjugate quaternion. The attitude of the spacecraft relative to the inertial frame.

[0068] Step S150-5: Calculate the spacecraft's attitude relative to the inertial frame based on the star sensor output. As initial values, inertial attitude navigation calculations are performed.

[0069] Step S160: The spacecraft performs a slow rotation search around its two axes to initially establish the initial values ​​of its position and velocity relative to Earth, as detailed below:

[0070] Step S160-1: Rotation Axis Design. Similar to attitude search, position and velocity search should involve full rotation along two axes perpendicular to the GNSS navigation receiver antenna beam center. Since the antenna beam center is the spacecraft's Z-axis, a rotation axis design is required. , These are the +Y and +X axes of the spacecraft system, respectively:

[0071] ;

[0072] in, The +X axis direction vector of the spacecraft system. It is the +Y axis direction vector of the spacecraft system.

[0073] Step S160-2: Slow Rotation Angular Velocity Design. Since a GNSS navigation receiver requires a certain signal acquisition and tracking time from receiving the navigation signal to outputting the navigation result, the rotation angular velocity cannot be too large; on the other hand, too small a rotation angular velocity reduces the responsiveness of initial navigation value establishment. Taking into account that the cold start time of current GNSS navigation receivers is generally no more than 120 seconds and the effective beamwidth for low Earth orbit is generally no less than... The rotational angular velocity for the navigation signal search is designed to be... .

[0074] Step S160-3: Perform a slow spin search.

[0075] To improve the position and velocity search speed, the negative normal vector of the spacecraft's theoretical orbital plane is first obtained. Perform attitude maneuvers to make the rotation axis point to Then, starting from the current attitude, under the action of the attitude maneuver control law, it rotates sequentially around the designed angular velocity. , Slowly rotate one revolution. During the rotation, if the GNSS navigation receiver outputs a valid position and velocity, the current step is completed and immediately exited; if it rotates around... , If the receiver still does not output a valid position velocity after one slow rotation, the slow rotation search continues with the same rotation direction and angular velocity until valid data is successfully output.

[0076] Rotation axis point to Inertial frame command direction cosine array for:

[0077] ;

[0078] in, This is the coordinate transformation matrix of the spacecraft's intrinsic system relative to the inertial frame, calculated based on the current attitude navigation values. It is a rotating axis The direction vector in an inertial frame. It refers to the spacecraft's position and velocity in an inertial frame of reference. It is a cosine matrix for command direction. The three-axis unit vectors of the coordinate system are calculated using the process described above. This is based on the direction cosine matrix. The corresponding inertial frame command attitude quaternion can be obtained by conversion. Used for attitude maneuver control. Maneuvers the current attitude to the commanded attitude. The control law is designed as follows:

[0079] ;

[0080] ;

[0081] in, To represent quaternion multiplication, express conjugate, The current attitude quaternion, It is the error attitude quaternion. for The vector part, The angular velocity of the spacecraft's attitude, calculated from measurements by a gyroscope. , , , , , These are the proportional and derivative control coefficients, respectively. It is the command control torque.

[0082] Around , The attitude control law for the axis slow rotation maneuver is designed as follows:

[0083] ;

[0084] in, The angular velocity of the spacecraft's attitude, calculated from measurements by a gyroscope. , , This is the proportional control coefficient. It is the command control torque.

[0085] Step S170: Based on the inertial frame attitude navigation calculation results at the current moment and the initial value of the Earth-fixed position and velocity, perform attitude maneuvers to adjust the orientation to a final stable direction that can simultaneously capture attitude and position and velocity, as follows:

[0086] Step S170-1: Calculation of the position and velocity of the inertial frame spacecraft. The GNSS navigation receiver outputs the spacecraft's fixed position and velocity on Earth, denoted as follows: , Let the current time be... A certain reference time and its corresponding Greenwich right ascension are respectively , Then the Greenwich right ascension at the current moment for:

[0087] ;

[0088] in, rad / s is the angular velocity of Earth's rotation. This is the current time. From this, the spacecraft's position in the inertial frame can be obtained. ,speed for:

[0089] ;

[0090] ;

[0091] in, It is the coordinate transformation matrix from the inertial frame of reference to the Earth-fixed frame of reference. , They are respectively The first and second components.

[0092] Step S170-2: Execute the command attitude calculation to simultaneously capture attitude and position / velocity. Specifically, this attitude is a three-axis stable attitude relative to the Earth, with the -Z axis pointing to the local zenith and the +Z axis pointing to the Earth's center. To be compatible with conventional three-axis stable states relative to the Earth, the +Y axis of the spacecraft system is taken as pointing to the negative normal to the orbital plane. The coordinate transformation matrix from the inertial frame to the orbital frame is then... for:

[0093] ;

[0094] ;

[0095] Based on the coordinate transformation matrix The corresponding attitude quaternion can be obtained by conversion. Three-axis stabilization relative to the Earth is generally controlled by attitude control relative to the orbital system. The attitude quaternion relative to the orbital system in the current state is:

[0096] ;

[0097] The command attitude at this time is:

[0098] ;

[0099] Inertial frame command attitude angular velocity for:

[0100] ;

[0101] in, is the Earth's gravitational constant.

[0102] Make the current posture and attitude angular velocity Maneuver and stabilize to the commanded attitude and command attitude angular velocity The control law is designed as follows:

[0103] ;

[0104] ;

[0105] in, It is the attitude error quaternion. for The vector part, , , , , , These are the proportional and differential control coefficients, respectively.

[0106] Step S180: Based on the attitude and position velocity results from step S170, calculate the position velocity and attitude results in the desired reference frame as the initial values ​​for the final inertial navigation.

[0107] Specifically, the desired reference frame navigation values ​​generally include: inertial frame attitude, orbital frame attitude, inertial frame position and velocity, and Earth-fixed frame position and velocity. Among these, depending on the designed sensor installation method, the spacecraft's inertial frame attitude can be calculated using the star sensor output value, and the GNSS receiver can output the Earth-fixed frame position and velocity. , Step S180 aims to calculate the attitude result in the desired reference frame, which refers to the attitude of the orbital frame. The attitude of the orbital frame can be calculated based on the attitude of the inertial frame, as follows:

[0108] ;

[0109] in, This is the coordinate transformation matrix of the spacecraft's intrinsic system relative to the inertial frame, calculated based on the attitude navigation values ​​output by the current star sensor. Let be the coordinate transformation matrix of the orbital system relative to the inertial system, derived from , The coordinate transformation matrix of the spacecraft system relative to the orbital system is obtained from steps S170-1 and S170-2. The corresponding orbital attitude quaternion or Euler attitude angle can be obtained from this matrix.

[0110] Step S180 aims to calculate the desired position and velocity in the reference frame, which refers to the position and velocity in the inertial frame. , , which can be determined by , The calculation is shown in step S170-1.

[0111] The above method will be described in more detail below with a specific example. This example describes a method for calculating the line-of-sight measurement of spatial targets on a strapdown platform, which includes the following steps:

[0112] Step S110: Design the system's standalone product configuration scheme.

[0113] The system's single-unit configuration consists of a "thrust attitude control system composed of a three-axis gyroscope, a star sensor, a GNSS navigation receiver, and four jet thrusters".

[0114] Step S120: Design the system stand-alone installation layout scheme.

[0115] See Figure 2 The installation layout of each unit is as follows: the axes of the three-axis gyroscope are parallel to the axes of the spacecraft's body coordinate system; the optical axis of the star sensor is along the direction of the angle bisectors of the +X and -Z angles of the spacecraft system; the beam center of the GNSS navigation receiver antenna points to the -Z direction of the spacecraft system; and all four jet thrusters are installed in the XOZ plane and are at an angle of 45° to both the X and Z axes.

[0116] Step S130: Separation and determination after release, as detailed below:

[0117] The spacecraft is equipped with limit switches, which provide a reliable separation signal directly.

[0118] Step S140: Separate the angular rate damping.

[0119] The spacecraft entered a 500km sun-synchronous orbit, and the initial orbital elements after release and separation were: , These are the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, argument of perigee, and true anomaly angle, respectively; the initial three-axis attitude angles after separation are... (3-1-2 reversal sequence), the initial triaxial angular velocity relative to the inertial frame is ° / s; Spacecraft inertia matrix is Using a gyroscope to measure the angular velocity of the three axes Measurements were performed, and the following proportional control law regarding angular velocity was used to determine the command torque. The calculated command torque is output by the jet thruster attitude control system:

[0120] ;

[0121] Refer to the rate-damping curve based on this control law. Figure 3 The criterion for the completion of rate damping was that the resultant angular velocity was no greater than 0.1° / s. The rate damping took a total of about 2 seconds, and the resultant angular velocity of the three axes was about 0.096° / s when it was completed. Figure 3 In the middle, the ordinate of each subplot , , These are the angular velocities along the X, Y, and Z axes, respectively.

[0122] Step S150: Perform slow rotation search around the two axes of the body to initially establish the initial attitude values ​​of the inertial frame.

[0123] Based on the star-aspect mounting method, the component of its optical axis in the spacecraft system is: The two rotation axes perpendicular to the optical axis are , Furthermore, the magnitude of the rotational angular velocity for the sky area search is... =1° / s. Two seconds after the rate damping ends, the star-sensitive slow-rotation search begins. The initial inertial frame position and velocity are... , The attitude of the inertial frame is (3-1-2 sequence) , , These are the roll angle, pitch angle, and yaw angle, respectively, and the attitude angular velocity relative to the inertial frame is... ° / s. Under the following attitude maneuver control law, the spacecraft orbits sequentially at the designed angular velocity. , Slow spin:

[0124] .

[0125] See results Figure 4 As can be seen from the image, the visible star region consistently remains within an angle greater than 90° with the Earth's center. Initially, the star sensor's optical axis was 85.33° below the boundary of the visible star region, within the region where stars were not visible. With the orbital... The star sensor's optical axis gradually rises during slow rotation, reaching the visible sky boundary after 84.14 seconds. It remains stable for a short period, then rises to approximately 11.63° above the visible sky boundary after 95 seconds. At this point, it is confirmed that the star sensor has outputted a good quality attitude result, and the slow rotation search is terminated. Subsequently, the gyroscope-based attitude inertial navigation system uses the star sensor output at this time as the initial value to continuously calculate the inertial frame attitude.

[0126] Step S160: Perform a slow rotation search around the two axes of the body to initially establish the initial values ​​of the Earth's fixed-axis position and velocity.

[0127] In this embodiment, after the star-sensitive search is completed, the position and velocity search continues. The initial position and velocity of the inertial frame are: , The attitude of the inertial frame is (3-1-2 turn sequence), the relative inertial frame attitude angular velocity is ° / s. In the initial state, the GNSS navigation receiver antenna beam center is approximately 22.49° below the visible sky area, and the GNSS receiver is in a non-positioning state where it cannot output a valid position and velocity. Therefore, a slow rotation search is required.

[0128] The two rotation axes perpendicular to each other are the direction of the GNSS antenna beam center. , Furthermore, the rotational angular velocity for the navigation signal search is... =0.5° / s.

[0129] In the slow spin search, the components of the negative normal vector of the spacecraft's theoretical initial orbital plane in the inertial frame are first obtained through the binding parameters. Design of rotating shaft point to Inertial frame command attitude quaternion for To enable attitude maneuvering to the commanded attitude. The control law is: , .

[0130] See results Figure 5 As can be seen from the figure, after approximately 25 seconds, the rotating axis... The trajectory basically coincides with the negative normal of the theoretical initial orbital plane. A short period of stability is maintained until 40 seconds to complete this phase of attitude maneuvering. During this period, the relationship between the angle between the GNSS antenna beam center and the boundary of the visible sky area for navigation signals is described in [reference needed]. Figure 6 As can be seen from the figure, the visible sky area for navigation signals remained within a range with an angle greater than 80° to the direction of the Earth's center. After about 2.242 seconds of maneuvering, the spacecraft was in the visible sky area for navigation signals. By the end of this phase, GNSS had performed about 38 seconds of navigation signal acquisition and tracking processing.

[0131] Around , The attitude control law for the axis slow rotation maneuver is designed as follows:

[0132] ;

[0133] See results Figure 7 As can be seen from the figure, during the slow rotation, the GNSS antenna beam direction is always in the visible sky area of ​​the navigation signal. Since the navigation signal has been acquired and tracked for a period of time during the attitude maneuver, and considering the receiver cold start time as 60s, the receiver has successfully located and output a valid position and velocity when the slow rotation reaches about 22s. It continues to stabilize for a short period of time, and then exits the position and velocity search at 30s. At this time, it can be ensured that the GNSS navigation receiver has output a position and velocity result of good quality.

[0134] Step S170: Perform attitude maneuvers to adjust the orientation to a final stable direction that can simultaneously capture attitude and position velocity.

[0135] In this embodiment, the spacecraft's Earth-fixed frame position and velocity output from the GNSS navigation receiver are transformed into a coordinate system to obtain the inertial frame position and velocity. , Based on this, the current attitude quaternion, commanded attitude quaternion, and commanded attitude angular velocity relative to the orbital system can be calculated. The initial attitude of the orbital system is... (3-1-2 turn sequence), the relative inertial frame attitude angular velocity is ° / s. The three-axis ground-based stable attitude control law design for the orbital system is as follows:

[0136] , .

[0137] See results Figure 8 and Figure 9 , Figure 8 The y-axis of each subplot , , These are roll angle, pitch angle, and yaw angle. Figure 9 The y-axis of each subplot , , These represent the angular velocities of the X, Y, and Z axes, respectively. As can be seen from the figure, after approximately 3 seconds, the system enters a stable triaxial state relative to the ground and maintains this state thereafter. (See reference...) Figure 10 Under the three-axis stable state, the optical axis of the star sensor and the beam center of the GNSS antenna always maintain a stable pointing, and the star sensor and GNSS navigation receiver can continuously output high-quality attitude and position velocity information.

[0138] Step S180: Under the three-axis stable attitude relative to the ground, the star sensor continuously outputs the inertial frame attitude quaternion. The GNSS receiver continuously outputs the Earth's fixed position and velocity. , In this embodiment, a typical set of output results using the UTC time of 00:09:00 on October 8, 2025 is taken as follows:

[0139] , , ;

[0140] When the corresponding orbital attitude is required and position and velocity of the inertial frame , The results obtained through calculation are as follows:

[0141] , , .

[0142] The present invention also provides an on-orbit device for establishing initial values ​​for spacecraft attitude and orbit navigation, comprising the following modules:

[0143] The inertial navigation module is used to measure the three-axis angle increments and perform inertial attitude navigation calculations.

[0144] The starlight attitude capture and measurement module is used to capture and identify stars, and to calculate the spacecraft's attitude relative to the inertial frame.

[0145] The satellite navigation module is used to receive satellite navigation signals and calculate the spacecraft's position and velocity.

[0146] The control module is used to control the spacecraft to achieve angular rate damping, slow rotation of the body around a specified axis, or pointing to the desired attitude direction based on the execution results of the calculation module.

[0147] The calculation module is used to execute the above method for establishing initial values ​​for spacecraft attitude, orbit, and navigation in orbit.

[0148] The present invention also provides an electronic device, including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the steps of the above-described method for establishing initial values ​​for spacecraft attitude and orbit navigation in orbit.

[0149] The present invention also provides a non-transitory computer-readable storage medium storing a computer program thereon, which, when executed by a processor, implements the steps of the above-described method for establishing initial values ​​for spacecraft attitude and orbit navigation in orbit.

[0150] Those skilled in the art will understand that embodiments of the present invention can be provided as methods, systems, or computer program products. Therefore, the present invention can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, the present invention can take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code. The solutions in the embodiments of the present invention can be implemented using various computer languages, such as the object-oriented programming language Java and the interpreted scripting language JavaScript.

[0151] This invention is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the invention. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, produce implementations of the flowchart illustrations and / or block diagrams. Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0152] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0153] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0154] Although preferred embodiments of the invention have been described, those skilled in the art, upon learning the basic inventive concept, can make other changes and modifications to these embodiments. If such modifications and variations of the invention fall within the scope of the claims and their equivalents, the invention is also intended to include such modifications and variations.

Claims

1. A method for establishing initial values ​​for spacecraft attitude and orbit navigation in orbit, characterized in that, Includes the following steps: Step S110: Design a single-unit product configuration scheme for the navigation system; Step S120: Design a single-unit installation layout scheme for the navigation system; Step S130: After completing the single-unit configuration and layout, the spacecraft determines whether it has successfully separated from the mother platform; Step S140: After confirming separation, the spacecraft dampens its own angular rate; Step S150: The spacecraft performs a slow rotation search around its two axes to initially establish the initial attitude values ​​of the inertial frame; Step S160: The spacecraft performs a slow rotation search around its two axes to initially establish the initial values ​​of the Earth-fixed position and velocity. Step S170: Perform attitude maneuvers based on the current inertial frame attitude navigation calculation results and the initial value of the Earth-fixed position and velocity, and adjust the attitude to a final stable direction that can simultaneously capture attitude and position and velocity. Step S180: Based on the attitude and position-velocity results from step S170, calculate the position-velocity and attitude results in the desired reference frame, which will be used as the initial values ​​for the final inertial navigation. Step S110 includes: Step S110-1: To realize the inertial attitude navigation function, configure an inertial measurement device composed of gyroscopes with three-axis angle increment measurement capability; Step S110-2: Configure the star sensor to realize the attitude search function; Step S110-3: To realize the position and velocity search function, a GNSS navigation receiver is configured, and the configured navigation receiver has a receiving sensitivity corresponding to the navigation signal strength at the orbital altitude at the time of release; Step S110-4: To realize the attitude and position velocity signal search function, a low-speed attitude control actuator is configured. The configured actuator has three-axis attitude maneuver control capability. Step S150 includes: Step S150-1: Rotation Axis Design: Let the unit direction vector of the star sensor optical axis be... The +X axis vector of the spacecraft system is Design of rotating shaft , as follows: ； Step S150-2: Slow Rotation Angular Velocity Design: The magnitude of the rotation angular velocity for the sky search is designed as follows: ; Step S150-3: Perform slow rotation search: Starting from the current attitude, under the action of the attitude maneuver control law, rotate sequentially according to the designed angular velocity. , Slowly rotate one revolution; during the rotation, if the star sensor outputs a valid attitude, the current step is completed and immediately exited; when circling... , If the star sensor still does not output a valid attitude after one slow rotation is completed, the slow rotation search will continue with the same rotation direction and angular velocity until a valid attitude is successfully output. Around , The attitude control law for the axis slow rotation maneuver is designed as follows: ； in, The angular velocity of the spacecraft's attitude, calculated from gyroscope measurements. , , This is the proportional control coefficient; Step S150-4: Calculate the spacecraft's attitude relative to the inertial frame based on the attitude output by the star sensor and the star sensor's installation matrix relative to the spacecraft. Step S150-5: Using the spacecraft's attitude relative to the inertial frame calculated from the star sensor output as the initial value, perform inertial attitude navigation calculations; Step S160 includes: Step S160-1: Rotation Axis Design: During position and velocity search, the receiver should rotate a full circle along two axes perpendicular to the center of the GNSS navigation receiver antenna beam. The rotation axis design is as follows. , These are the +Y and +X axes of the spacecraft system, respectively: ； in, The +X axis direction vector of the spacecraft system. It is the +Y axis direction vector of the spacecraft system; Step S160-2: Slow Rotation Angular Velocity Design: Design the rotation angular velocity for navigation signal search to be... ; Step S160-3: Perform a slow spin search.

2. The method for establishing initial values ​​for spacecraft attitude and orbit navigation in orbit according to claim 1, characterized in that, Step S120 includes: Step S120-1: Inertial Measurement Unit Installation Layout Design: Install the three-axis gyroscope in a conventional manner, so that each axis of the gyroscope corresponds to and is parallel to each axis of the spacecraft system; Step S120-2: Star sensor installation layout design: The installation direction of the star sensor is designed so that the optical axis is at a 45° angle with the XOY plane of the spacecraft system and deviates towards -Z; Step S120-3: GNSS navigation receiver installation layout design: When only one antenna is configured, the antenna beam center of the GNSS navigation receiver is pointed to the -Z direction of the spacecraft system. When multiple antennas are configured, the beam center of one of the antennas is pointed to the -Z direction of the spacecraft system. Step S120-4: Installation layout design of low-speed attitude control actuator: Its installation layout should enable it to have three-axis attitude control capability.

3. The method for establishing initial values ​​for spacecraft attitude and orbit navigation in orbit according to claim 2, characterized in that, Step S130 includes: when the spacecraft is equipped with a limit switch, the limit switch directly provides a separation signal; otherwise, if an attitude control engine is equipped, the attitude control engine is started and tested, and the angular velocity increment generated by the excitation is measured for judgment; when no attitude control engine is equipped, the attitude control actuator is used to excite the angular velocity, and the separation judgment is completed based on the measured actual angular velocity increment.

4. The method for establishing initial values ​​for spacecraft attitude and orbit navigation in orbit according to claim 3, characterized in that, Step S140 includes: Using gyroscopes to perform three-axis angle increment measurements on spacecraft And calculate the corresponding triaxial angular velocities. : ； in, The sampling period for gyroscope measurements; These are the angular increments of the spacecraft around its X, Y, and Z axes. These are the angular velocities of the spacecraft around its X, Y, and Z axes, respectively; the command torque is calculated using the following proportional control law regarding the angular velocity. calculate: ； in, , , This is the proportional control coefficient; finally, the low-speed attitude control actuator configured by the system completes the actual control torque output and ultimately completes the angular velocity rate damping task.

5. A device for establishing initial values ​​for spacecraft attitude and orbit navigation in orbit, characterized in that, Includes the following modules: The inertial attitude navigation module is used to measure the three-axis angle increments and perform inertial attitude navigation calculations. The starlight attitude capture and measurement module is used to capture and identify stars, and to calculate the spacecraft's attitude relative to the inertial frame. The satellite navigation module is used to receive satellite navigation signals and calculate the spacecraft's position and velocity. The control module is used to control the spacecraft to achieve angular rate damping, slow rotation of the body around a specified axis, or pointing to the desired attitude direction based on the execution results of the calculation module. A calculation module is used to execute the on-orbit establishment method for initial values ​​of spacecraft attitude, orbit, and navigation as described in any one of claims 1-4.

6. An electronic device, characterized in that, It includes a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the on-orbit establishment method for initial values ​​of spacecraft attitude, orbit, and navigation according to any one of claims 1-4.

7. A non-transitory computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by a processor, the computer program implements the method for establishing initial values ​​for spacecraft attitude, orbit, and navigation in orbit according to any one of claims 1-4.

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