A method for autonomous orbit control of satellites using detachable independent propulsion modules

Abstract

A method for autonomous satellite orbit control using a separable independent propulsion module is proposed. This method does not rely on a ground-based telemetry and control system. By autonomously assessing the remaining propellant, it improves the efficiency of chemical propellant use in the satellite propulsion module while ensuring the safety of the satellite's chemical propellant. Furthermore, by calculating the satellite's orbit control parameters and formulating an orbit control ignition strategy based on orbit change constraints and all injection parameters, autonomous satellite orbit control can be achieved without relying on a ground-based telemetry and control system.

Description

Technical Field

[0001] This invention relates to a satellite autonomous orbit control method using a detachable independent propulsion module, belonging to the field of spacecraft orbit control. Background Technology

[0002] AsiaSat 6E is a geostationary broadband communications satellite equipped with a detachable independent propulsion module. The satellite uses a hybrid chemical-electric propulsion system for its orbit insertion. A separate chemical propulsion system (SPS) is connected in series below it. After separation from the launch vehicle, the SPS performs the chemical propulsion orbit change first, then the SPS is jettisoned, and the electric propulsion system completes the remaining orbit transfer and operational orbit position maintenance tasks.

[0003] Traditional satellites using chemical propulsion systems consume a large amount of propellant after launch for orbital maneuvers. Once in GEO orbit, sufficient chemical propellant is retained for position maintenance during on-orbit operation. Satellites using independent propulsion modules differ from traditional satellites in that their initial orbit after separation is a LEO orbit with a perigee of 200 km and an apogee of 500 km. The independent propulsion module is equipped with a bicomponent chemical propulsion system, carrying approximately 1845 kg of propellant. The satellite design utilizes all the chemical propellant from the independent propulsion module for lifting the satellite's semi-major axis, reserving approximately 50 kg of fuel for deorbit control after separation.

[0004] The existing orbital control methods using chemical propulsion systems are as follows:

[0005] 1. The satellite was launched into GTO orbit;

[0006] 2. Ground-based or spaceborne GNSS determines the satellite's orbit. The ground-based system uses the satellite's current orbit and ground-based telemetry and control capabilities to calculate the number of orbit control maneuvers and orbit control ignition cycles.

[0007] 3. Once the satellite enters the tracking and control arc, the ground sends remote control commands to control its orbit.

[0008] 4. After the satellite completes its orbital maneuver, the remaining propellant is used to maintain its working orbital position.

[0009] Using an independent propulsion module to lift a satellite into orbit has the following advantages compared to traditional satellites:

[0010] 1. The satellite initially operates in LEO orbit, with few tracking and control segments and short durations for each segment, making it difficult to support remote orbit changes.

[0011] 2. After completing the orbital maneuver, the propulsion module separates from the satellite and becomes deorbited. Aside from the propellant used for deorbiting, the remaining propellant is used for satellite orbit control.

[0012] Based on the above constraints, traditional ground-based remote-controlled satellite orbit control methods no longer meet the requirements for orbit lifting using independent propulsion modules. Summary of the Invention

[0013] The technical problem solved by this invention is that the traditional ground-based remote-controlled satellite orbit control method cannot meet the orbit lifting requirements of the independent propulsion module in the existing technology. Therefore, this invention proposes an autonomous orbit control method for satellites with a detachable independent propulsion module.

[0014] The present invention solves the above-mentioned technical problem through the following technical solution:

[0015] A method for autonomous orbit control of satellites using detachable independent propulsion modules includes:

[0016] Determine the constraints for satellite orbit change and the corresponding injection parameters required for satellite orbit change;

[0017] The satellite receives GNSS orbit determination data or injects orbit parameters from the ground, and together with the injected parameters required for satellite orbit change, performs an availability assessment.

[0018] Calculate satellite orbit control parameters based on the orbit parameters received from the satellite through availability assessment;

[0019] The orbit control ignition strategy is formulated based on the calculated satellite orbit control parameters.

[0020] The satellite orbit change constraints include the longest single working time and number of times the satellite propulsion system's orbit control thruster is activated, the feasibility of ignition during the Earth's shadow period of the satellite energy system, the orbit control strategy under the risk of low-orbit collision, and the remaining amount of propellant in the independent propulsion module after deorbiting. The injection parameters required for satellite orbit change are designed and input according to various satellite orbit change constraints to complete the satellite orbit change.

[0021] When the GNSS orbit determination data or the orbit parameters injected on the ground and the parameters required for satellite orbit change are both available, all data are used for satellite autonomous orbit control, and satellite orbit control parameters are calculated, including the next orbit control ignition time tj and ignition duration Tj.

[0022] If both the GNSS orbit determination data or the ground-based orbit parameters and the parameters required for satellite orbit change are unavailable, then the parameters required for satellite orbit change should be re-determined and the GNSS orbit determination data or ground-based orbit parameters should be received again until all data are available.

[0023] The specific method for calculating satellite orbit control parameters is as follows:

[0024] Determine the current satellite mass and dry weight, and calculate the propellant consumption based on the required velocity increment for the orbit change mission;

[0025] Judge whether the propellant consumption meets the consumption for the next ignition duration, and calculate the current ignition duration according to the judgment result;

[0026] Determine the next orbit control ignition time according to the current ignition duration.

[0027] The method for judging whether the propellant consumption meets the consumption for the next ignition duration is as follows:

[0028] Calculate the theoretical propellant consumption of the satellite M - M0 - Md according to the current satellite mass M, the satellite dry mass M0, and the remaining propellant amount Md required for the independent propulsion module to de-orbit;

[0029] Compare the propellant consumption dm calculated from the velocity increment required for the orbit transfer mission with the theoretical propellant consumption M - M0 - Md of the satellite. If dm < M - M0 - Md, calculate the current ignition duration Tj using dm; otherwise, calculate the current ignition duration Tj using M - M0 - Md.

[0030] The specific orbit control ignition strategy is as follows:

[0031] When the orbit control ignition time is tj - t n1 the satellite switches to the ignition attitude;

[0032] When the orbit control ignition time is tj - t n2 the satellite performs the bottoming operation;

[0033] When the orbit control ignition time is tj - t n3 the high-pressure self-locking valve of the satellite propulsion system is opened;

[0034] When the orbit control ignition time is tj, perform the orbit control ignition operation.

[0035] The remaining propellant amount Md is a preset value, and the method for evaluating and determining the actual value of the remaining propellant amount is as follows:

[0036] Measure the orbit transfer velocity increment according to the accelerometer installed in the independent propulsion module, calculate the satellite mass after the ignition ends according to the orbit transfer velocity increment and the preset specific impulse, and evaluate and obtain the remaining propellant amount using the satellite mass after the ignition ends.

[0037] The real-time value of the actual remaining propellant amount is calculated according to the comprehensive thrust and specific impulse parameters during the ignition; both the real-time value calculation method and the remaining propellant amount Md obtained by the evaluation and determination method can be used in the calculation of the current ignition duration Tj, and the priority of the real-time value calculation method and the evaluation and determination method is set according to the orbit transfer requirements.

[0038] When the satellite enters the final orbit, it uses the propulsion system tank hot spot method to jointly determine whether the propellant has reached the stop orbit control ignition threshold. If orbit control ignition is in progress and the remaining propellant triggers the hot spot method alarm, the orbit control engine is automatically shut down according to the orbit control ignition strategy, the orbit control is stopped, and the satellite switches to sun orientation.

[0039] Otherwise, it will continue to perform orbit changes according to the orbit control ignition strategy, and after completing all orbit changes, it will autonomously switch to the sun orientation mode and wait for the satellite to perform subsequent actions.

[0040] In the orbital control ignition strategy, t n1 t n2 t n3 Based on the requirements of the orbital change mission, the ignition duration Tj is determined, where j = 1 to n, and n is the total number of orbital changes.

[0041] The advantages of this invention compared to the prior art are:

[0042] This invention provides a satellite autonomous orbit control method using a separable independent propulsion module. By clearly defining the orbit change constraints and all injection parameters, the method calculates the satellite orbit control parameters and formulates the orbit control strategy. By autonomously evaluating the remaining propellant, the method improves the utilization efficiency of the satellite propulsion module's chemical propellant while ensuring the safety of the satellite's chemical propellant use. This enables autonomous satellite orbit control without relying on a ground-based telemetry and control system. Attached Figure Description

[0043] Figure 1 A schematic diagram of an independent propulsion module satellite autonomous orbit control method provided for the invention; Detailed Implementation

[0044] A method for autonomous satellite orbit control using a separable independent propulsion module is proposed. This method does not rely on a ground-based telemetry and control system. By autonomously assessing the remaining propellant, it improves the efficiency of chemical propellant use in the satellite propulsion module while ensuring the safety of the satellite's chemical propellant. Furthermore, by clearly defining the orbit change constraints and all injection parameters, it calculates the satellite's orbit control parameters and formulates an orbit control ignition strategy. This method achieves autonomous satellite orbit control without relying on a ground-based telemetry and control system.

[0045] The specific steps for the independent propulsion module satellite autonomous orbit control method are as follows:

[0046] Determine the constraints for satellite orbit change and the corresponding injection parameters required for satellite orbit change;

[0047] The satellite receives GNSS orbit determination data or injects orbit parameters from the ground, and together with the injected parameters required for satellite orbit change, performs an availability assessment.

[0048] Calculate satellite orbit control parameters based on the orbit parameters received from the satellite through availability assessment;

[0049] Formulate the orbit control ignition strategy based on the calculated satellite orbit control parameters.

[0050] The satellite orbit transfer constraint conditions include the maximum single working time and start-up times of the orbit control thrusters of the satellite propulsion system, the feasibility of ignition during the satellite's eclipse period in the energy system, the orbit control strategy under the risk of low-orbit collision of the satellite, and the remaining amount of deorbiting propellant in the independent propulsion module; the injection parameters required for satellite orbit transfer are designed and input according to various satellite orbit transfer constraint conditions to complete the satellite orbit transfer.

[0051] When both the GNSS orbit determination data or the ground uploaded orbit parameters and the injection parameters required for satellite orbit transfer are in an available state, all data are used for the satellite's autonomous orbit control, and the satellite orbit control parameters are calculated, including the next orbit control ignition time tj and the ignition duration Tj;

[0052] When both the GNSS orbit determination data or the ground uploaded orbit parameters and the injection parameters required for satellite orbit transfer are in an unavailable state, re-determine the injection parameters required for satellite orbit transfer and re-receive the GNSS orbit determination data or the ground uploaded orbit parameters until all data are in an available state.

[0053] The method for calculating the satellite orbit control parameters is specifically as follows:

[0054] Determine the current satellite mass and the satellite dry mass, and calculate the propellant consumption according to the required velocity increment for the orbit transfer mission;

[0055] Judge whether the propellant consumption meets the consumption for the next ignition duration, and calculate the current ignition duration according to the judgment result;

[0056] Determine the next orbit control ignition time according to the current ignition duration.

[0057] The method for judging whether the propellant consumption meets the consumption for the next ignition duration is as follows:

[0058] According to the current satellite mass M, the satellite dry mass M0, and the remaining amount of propellant Md required for the deorbiting of the independent propulsion module, calculate the theoretical propellant consumption of the satellite M - M0 - Md;

[0059] Compare the calculated propellant consumption dm obtained from the required velocity increment for the orbit transfer mission with the theoretical propellant consumption of the satellite M - M0 - Md. If dm < M - M0 - Md, calculate the current ignition duration Tj using dm; otherwise, calculate the current ignition duration Tj using M - M0 - Md.

[0060] The orbit control ignition strategy is specifically as follows:

[0061] When the orbit control ignition time is tj - t n1 the satellite switches to the ignition attitude;

[0062] When the rail control ignition time is tj-t n2 At that time, the satellite performed a bottom-diving operation;

[0063] When the rail control ignition time is tj-t n3 At that time, the high-pressure self-locking valve of the satellite propulsion system is opened;

[0064] When the track control ignition time is tj, the track control ignition operation is executed.

[0065] The remaining propellant quantity Md is a preset value. The method for assessing and determining the actual remaining propellant quantity is as follows:

[0066] The satellite mass after ignition is calculated based on the orbit change velocity increment measured by the accelerometer installed in the independent propulsion module and the preset specific impulse. The remaining propellant amount is obtained by assessing the satellite mass after ignition.

[0067] The remaining propellant quantity Md is a preset value. The method for assessing and determining the actual remaining propellant quantity is as follows:

[0068] The satellite mass after ignition is calculated based on the orbit change velocity increment measured by the accelerometer installed in the independent propulsion module and the preset specific impulse. The remaining propellant amount is obtained by assessing the satellite mass after ignition.

[0069] When the satellite enters the final orbit, it uses the propulsion system tank hot spot method to jointly determine whether the propellant has reached the stop orbit control ignition threshold. If orbit control ignition is in progress and the remaining propellant triggers the hot spot method alarm, the orbit control engine is automatically shut down according to the orbit control ignition strategy, the orbit control is stopped, and the satellite switches to sun orientation.

[0070] Otherwise, it will continue to perform orbit changes according to the orbit control ignition strategy, and after completing all orbit changes, it will autonomously switch to the sun orientation mode and wait for the satellite to perform subsequent actions.

[0071] In the orbital control ignition strategy, t n1 t n2 t n3 Based on the requirements of the orbital change mission, the ignition duration Tj is determined, where j = 1 to n, and n is the total number of orbital changes.

[0072] The following description, in conjunction with the accompanying drawings and preferred embodiments, provides further details:

[0073] In the current embodiment, a satellite autonomous orbit control method with detachable independent propulsion modules is adopted, such as... Figure 1 As shown, the specific steps are as follows:

[0074] (1) Define the orbit change constraints: including the longest working time of the propulsion system orbit control thruster and the number of times it can be started; whether the satellite energy system supports ignition during the satellite's shadow period; the requirements of the orbit control strategy for the risk of low-orbit collisions; the remaining propellant Md required for the independent propulsion module to deorbit, etc. The above constraints serve as input parameters for the satellite's autonomous orbit control design.

[0075] The initial orbit for satellite separation from the launch vehicle is a LEO orbit with a perigee of 200 km and an apogee of 500 km. The independent propulsion module is equipped with a bicomponent chemical propulsion system, using a 490N engine as the orbital control thruster. It carries approximately 1845 kg of propellant. The satellite design utilizes all the chemical propellant from the independent propulsion module for lifting the satellite's semi-major axis, reserving approximately 50 kg of fuel for deorbit control after separation. Considering factors such as collision risk and propulsion module fuel load, the orbital control design for the independent propulsion module is as follows:

[0076] (a) At the perigee of the initial orbit, the first orbital change is performed, increasing the apogee altitude from 500 km to 1300 km. The ignition direction is along the instantaneous inertial velocity direction.

[0077] (b) At the apogee of 1300 km, the second orbital change circularizes the orbit, transforming it into a near-circular orbit at an altitude of 1300 km. The ignition direction is along the instantaneous inertial velocity direction.

[0078] (c) Perform six (3rd to 8th) orbital changes at appropriate locations on a near-circular track at an altitude of 1300km to maximize the semi-major axis of the track.

[0079] (2) If the satellite receives the correct GNSS orbit determination data, or if new orbit parameters are injected from the ground, the autonomous orbit control function can be enabled.

[0080] Before satellites can enable autonomous orbit control, the ground needs to assess the health of the control system. GNSS receivers can provide satellite orbit information, and attitude sensors such as star sensors can output satellite attitude data.

[0081] (3) The satellite autonomously calculates its orbit control parameters based on the orbit data from step (2). These parameters include the next orbit control ignition time tj and ignition duration Tj (j = 1 to 8).

[0082] (31) Current satellite mass M, satellite dry weight M0 = 2500 kg;

[0083] (32) Based on the velocity increment dv required for orbit change and the specific impulse Isp of the orbit control thruster (unit: m / s), calculate the propellant consumption dm, where Ms is the current satellite mass;

[0084]

[0085] (33) If dm < M - 2550, the ignition duration of this time is calculated using dm; otherwise, the ignition duration Tj of this time is calculated using M - 2550.

[0086]

[0087] (4) According to the ignition strategy calculated in step (3), each time of orbit control ignition, the following orbit change process is executed on the satellite:

[0088] (41) tj - t n1 (t n1 = 600 s), the satellite transfers to the ignition attitude;

[0089] (42) tj - t n2 (t n2 = 120 s), the satellite executes the sinking operation;

[0090] (43) tj - t n3 (t n3 = 10 s), open the high - pressure self - locking valve of the propulsion system;

[0091] (44) At tj moment, execute the orbit control ignition operation.

[0092] (441) During the orbit change process, two methods are used on the satellite to estimate the remaining propellant of the satellite. One is to calculate the real - time remaining propellant Mt through the comprehensive thrust and specific impulse during the 490N ignition, and the expression is as follows:

[0093]

[0094] The other is to obtain the velocity increment of the orbit change through the accelerometer installed in the independent propulsion module, and calculate the mass of the satellite after the ignition ends according to the preset specific impulse. The expression is as follows:

[0095]

[0096] The priority can be set for the two types of evaluation methods;

[0097] (442) In the last few orbits of the orbit change, use the hot spot method of the propulsion system storage tank to jointly judge whether the propellant reaches the threshold of stopping the orbit control ignition. If the orbit control ignition is in progress and the remaining propellant touches the hot spot method alarm, the satellite autonomously closes the orbit control engine according to the strategy, stops the orbit control, and transfers to the sun - pointing orientation to ensure the energy and propellant safety of the satellite. Otherwise, after the satellite completes all orbits of the orbit change according to the orbit change strategy, it autonomously transfers to the sun - pointing orientation mode, completes the autonomous orbit control of the independent propulsion module, and waits for the satellite to execute subsequent operations.

[0098] Although the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solutions of the present invention by utilizing the methods and techniques disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and alterations made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the protection scope of the technical solutions of the present invention.

[0099] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for autonomous orbit control of a satellite using a detachable independent propulsion module, characterized in that... It includes: Determine the satellite orbit transfer constraint conditions and the corresponding injection parameters required for satellite orbit transfer; The satellite receives GNSS orbit determination data or ground-uplinked orbit parameters, and conducts an availability judgment together with the injection parameters required for satellite orbit transfer; Calculate the satellite orbit control parameters based on the orbit parameters received by the satellite that pass the availability judgment; Formulate an orbit control ignition strategy based on the calculated satellite orbit control parameters; The satellite orbit transfer constraint conditions include the single longest working time and the number of starts of the orbit control thruster of the satellite propulsion system, the feasibility of ignition during the satellite's eclipse period by the satellite energy system, the orbit control strategy under the risk of low-orbit collision of the satellite, and the remaining amount of deorbit propellant in the independent propulsion module; the injection parameters required for satellite orbit transfer are designed and input according to various satellite orbit transfer constraint conditions to complete satellite orbit transfer; The method for calculating the satellite orbit control parameters is specifically as follows: Determine the current satellite mass and the satellite dry mass, and calculate the propellant consumption according to the required velocity increment for the orbit transfer mission; Judge whether the propellant consumption meets the consumption during the next ignition duration, and calculate the current ignition duration according to the judgment result; Determine the next orbit control ignition time according to the current ignition duration; The method for judging whether the propellant consumption meets the consumption during the next ignition duration is as follows: According to the current satellite mass M, the satellite dry mass M0, and the remaining amount of propellant Md required for the deorbit of the independent propulsion module, calculate the theoretical propellant consumption of the satellite M - M0 - Md; Compare the calculated propellant consumption dm obtained from the required velocity increment for the orbit transfer mission with the theoretical propellant consumption of the satellite M - M0 - Md. If dm < M - M0 - Md, then the current ignition duration Tj is calculated using dm; otherwise, the current ignition duration Tj is calculated using M - M0 - Md.

2. A method for autonomous orbit control of a satellite with a separable independent propulsion module according to claim 1, characterized in that: When both the GNSS orbit determination data or the ground-uplinked orbit parameters and the injection parameters required for satellite orbit transfer are in an available state, all the data are used for the autonomous orbit control of the satellite, and the satellite orbit control parameters are calculated, including the next orbit control ignition time tj and the ignition duration Tj; When both the GNSS orbit determination data or the ground-uplinked orbit parameters and the injection parameters required for satellite orbit transfer are in an unavailable state, re-determine the injection parameters required for satellite orbit transfer and re-receive the GNSS orbit determination data or the ground-uplinked orbit parameters until all the data are in an available state.

3. A method for autonomous orbit control of a satellite with a separable independent propulsion module according to claim 2, characterized in that: The orbit control ignition strategy is specifically as follows: When the rail control ignition time is tj-t n1 At that time, the satellite entered the ignition attitude; When the rail control ignition time is tj-t n2 At that time, the satellite performed a bottom-diving operation; When the rail control ignition time is tj-t n3 At that time, the high-pressure self-locking valve of the satellite propulsion system is opened; When the orbit control ignition time is at the tj time, perform the orbit control ignition operation.

4. A method for autonomous orbit control of a satellite with a separable independent propulsion module according to claim 3, characterized in that: The remaining amount of propellant Md is a preset value, and the method for evaluating and determining the actual value of the remaining amount of propellant is: Measure the velocity increment for orbit transfer according to the accelerometer installed in the independent propulsion module, calculate the satellite mass after the ignition ends according to the velocity increment for orbit transfer and the preset specific impulse, and evaluate and obtain the remaining amount of propellant by using the satellite mass after the ignition ends.

5. A satellite autonomous orbit control method using a detachable independent propulsion module according to claim 4, characterized in that: The real-time value of the actual remaining propellant quantity is calculated based on the combined thrust and specific impulse parameters during ignition. The remaining propellant quantity Md obtained by both the real-time value calculation method and the evaluation and determination method can be used in the calculation of the ignition duration Tj. The priority of the real-time value calculation method and the evaluation and determination method is set according to the orbit change requirements.

6. A satellite autonomous orbit control method using a detachable independent propulsion module according to claim 5, characterized in that: When the satellite enters the final orbit, it uses the propulsion system tank hot spot method to jointly determine whether the propellant has reached the stop orbit control ignition threshold. If orbit control ignition is in progress and the remaining propellant triggers the hot spot method alarm, the orbit control engine is automatically shut down according to the orbit control ignition strategy, the orbit control is stopped, and the satellite switches to sun orientation. Otherwise, it will continue to perform orbit changes according to the orbit control ignition strategy, and after completing all orbit changes, it will autonomously switch to the sun orientation mode and wait for the satellite to perform subsequent actions.

7. A satellite autonomous orbit control method using a detachable independent propulsion module as described in claim 6, characterized in that: In the orbital control ignition strategy, t n1 t n2 t n3 Based on the requirements of the orbital change mission, the ignition duration Tj is determined, where j = 1~n, and n is the total number of orbital changes.

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