A method for implementing orbit control by attitude bias using out-of-plane thrusters

Abstract

The embodiment of the present disclosure relates to a method for implementing orbit control by using out-of-plane thrusters with attitude bias. The method comprises the following steps: determining a satellite orbit control time interval by using a star sensor; determining a minimum available residual fuel amount of a satellite and calculating a satellite orbit control available speed increment by using characteristic parameters of the satellite; calculating a total semi-major axis control amount and a total jet time length and determining a satellite orbit control strategy according to the available speed increment, a semi-major axis and a satellite attitude bias angle; calculating a thruster efficiency factor and adjusting control parameters in the satellite orbit control strategy according to the thruster efficiency factor. The embodiment of the present disclosure can realize that the control time interval during the satellite orbit control of the GEO satellite with abnormal tangential thruster is not constrained, and the orbit keeping requirement after the satellite orbit control is met.

Description

TECHNICAL FIELD

[0001] The embodiment of the present disclosure relates to the technical field of on-orbit management of a spacecraft, and particularly relates to a method for implementing de-orbiting control by using an out-of-plane thruster through attitude bias. BACKGROUND

[0002] Due to the functional degradation of on-orbit GEO (Geosynchronous Earth Orbit) satellite on-board equipment or insufficient on-board fuel, there is a risk of satellite disconnection and loss of control, and the satellite needs to be de-orbited and passivated to avoid adverse effects on other satellites in the GEO orbit and to no longer occupy valuable orbital resources in the GEO orbit. In the related art, the perturbation characteristics of the GEO tomb orbit are utilized by using the second-order lunar term, and the direction of the perigee is set to 90 degrees or 270 degrees to effectively suppress the descent of the perigee of the large eccentricity tomb orbit. The first ignition time needs the satellite to be located at 90 degrees or 270 degrees in right ascension. However, the above method has certain requirements for the control time, and cannot be directly applied to the de-orbiting control of GEO satellites with abnormal tangential thrusters.

[0003] Therefore, it is necessary to improve one or more problems in the related technical solutions.

[0004] It should be noted that this section aims to provide background or context to the technical solutions of the present disclosure stated in the claims. The description herein is not admitted to be prior art merely because it is included in this section. SUMMARY

[0005] The purpose of the embodiment of the present disclosure is to provide a method for implementing de-orbiting control by using an out-of-plane thruster through attitude bias, thereby at least overcoming one or more problems caused by the limitations and defects of the related art.

[0006] According to the embodiment of the present disclosure, a method for implementing de-orbiting control by using an out-of-plane thruster through attitude bias is provided, and the method comprises:

[0007] determining a satellite de-orbiting control time interval by using a star sensor;

[0008] determining a minimum available residual fuel amount of a satellite by using characteristic parameters of the satellite, and calculating a speed increment available for de-orbiting control of the satellite;

[0009] calculating a total semi-major axis control amount and a total jet time length according to the available speed increment, a semi-major axis and an attitude bias angle of the satellite, and determining a satellite de-orbiting control strategy;

[0010] calculating a thruster efficiency factor, and adjusting control parameters in the satellite de-orbiting control strategy according to the thruster efficiency factor.

[0011] In an embodiment of the present disclosure, the process of determining the satellite de-orbit control time interval by using the star sensor includes:

[0012] According to the parameters of the star sensor, the position of the sun vector is calculated, and the relative angle between the sun vector position and the star sensitive light axis is obtained;

[0013] According to the relative angle between the sun vector position and the star sensitive light axis, the time period when the star sensor is disturbed by stray light is determined, thereby determining the satellite de-orbit control time interval.

[0014] In an embodiment of the present disclosure, the relative angle between the sun vector position and the star sensitive light axis includes:

[0015] The sun out-of-plane angle and the sun in-plane angle.

[0016] In an embodiment of the present disclosure, when the sun out-of-plane angle α < 110° and the sun in-plane angle -30° < β < 130°, the corresponding time is the time period when the star sensor is disturbed by stray light.

[0017] In an embodiment of the present disclosure, the process of determining the minimum available residual fuel amount of the satellite by using the characteristic parameters of the satellite includes:

[0018] According to the characteristic parameters of the satellite, the residual amount of propellant is calculated by using the PVT method, and the characteristic parameters include the pressure and temperature data of the propellant tank;

[0019] According to the residual amount of propellant, the minimum available residual fuel amount of the satellite is calculated.

[0020] In an embodiment of the present disclosure, the process of calculating the total semi-major axis control amount and the total jet time length according to the available speed increment, the semi-major axis and the satellite attitude offset angle includes:

[0021] The tangential speed increment is obtained according to the tangential semi-major axis target control amount;

[0022] The de-orbit control total speed increment and the normal speed increment are obtained according to the tangential speed increment, and the normal semi-major axis target control amount is obtained;

[0023] The total semi-major axis control amount is obtained according to the tangential semi-major axis target control amount and the normal semi-major axis target control amount, and the total jet time length is obtained according to the de-orbit control total speed increment.

[0024] In an embodiment of the present disclosure, the process of determining the satellite de-orbit control strategy includes:

[0025] According to the total semi-major axis control amount and the total jet time length, a control method combining multiple average amount maneuver control and correction amount maneuver control is adopted.

[0026] In an embodiment of the present disclosure, the control is performed at time intervals spaced 12 hours apart at far and near points, a batch of test spraying is performed first, and the remaining target control amount is distributed according to 2*N times of control.

[0027] In an embodiment of the present disclosure, the process of calculating the thruster efficiency factor and adjusting the control parameters in the post-control orbit determination, calculating the actual semi-major axis change amount after control, and obtaining the thruster efficiency factor comprises:

[0028] In an embodiment of the present disclosure, the process of calculating the thruster efficiency factor and adjusting the control parameters in the post-control orbit determination, calculating the actual semi-major axis change amount after control, and obtaining the thruster efficiency factor comprises:

[0029] If the efficiency factor is in the range of (0.85, 0.97] or [1.03, 1.05), the parameters including the thruster calibration coefficient are adjusted.

[0030] The technical scheme provided by the embodiment of the present disclosure can include the following beneficial effects:

[0031] In an embodiment of the present disclosure, by using the above method of realizing orbit control by attitude bias using out-of-plane thrusters, on the one hand, the orbit height of the GEO satellite after orbit control is above 280 kilometers above the synchronous belt, and the eccentricity is less than 0.003, avoiding the problem of the satellite frequently crossing the synchronous belt due to excessive satellite eccentricity. On the other hand, the problem of limited control time due to the influence of stray light on star sensitivity when using large-angle attitude bias for out-of-plane thrust orbit control due to abnormal tangential thrusters is solved. BRIEF DESCRIPTION OF DRAWINGS

[0032] The accompanying drawings, which are incorporated into and form part of the specification, illustrate embodiments consistent with the present disclosure and, together with the specification, serve to explain the principles of the present disclosure. Obviously, the drawings in the following description are only some embodiments of the present disclosure, and other drawings can be obtained from these drawings without creative labor for those skilled in the art.

[0033] Figure 1 FIG. 1 shows a step diagram of a method of realizing orbit control by attitude bias using out-of-plane thrusters in an exemplary embodiment of the present disclosure;

[0034] Figure 2 FIG. 2 shows a relationship diagram between tangential and normal velocities after attitude bias of a satellite in an exemplary embodiment of the present disclosure. DETAILED DESCRIPTION

[0035] Example implementations are now described with reference to the drawings. Example implementations can, however, be implemented in many different forms and should not be construed as limited to the examples set forth herein; rather, these examples are provided so that this disclosure will be thorough and complete, and will fully convey the scope of example implementations to those skilled in the art. The described features, structures, or characteristics can be combined in one or more implementations.

[0036] In addition, the drawings are merely schematic and are not drawn to scale. Identical reference numerals in the figures designate identical or similar parts throughout the figures and text. Some of the blocks in the drawings are functional blocks that represent functions implemented by a processor, software, or other implementation, module, or unit that is not necessarily implemented as described.

[0037] In the present example implementation, a method for implementing deorbiting control by using out-of-plane thrusters with attitude bias is first provided. As shown in FIG. 1, the method can include steps S101-S104. Figure 1

[0038] Step S101: determining a satellite deorbiting control time interval by using a star sensor;

[0039] Step S102: determining a minimum available residual fuel amount of a satellite by using a characteristic parameter of the satellite, and calculating a speed increment available for satellite deorbiting control;

[0040] Step S103: calculating a total semi-major axis control amount and a total jet time length according to the available speed increment, a semi-major axis, and a satellite attitude bias angle, and determining a satellite deorbiting control strategy;

[0041] Step S104: calculating a thruster efficiency factor, and adjusting control parameters in the satellite deorbiting control strategy according to the thruster efficiency factor.

[0042] Specifically, in step S101, when a satellite adopts a large-angle attitude bias for out-of-plane thrust deorbiting control due to abnormal tangential thrusters, there is a problem that the control time is limited due to the influence of stray light on the star sensor. When a small-angle attitude bias (bias angle less than 10°) is adopted, there is a time interval during which control can be performed throughout the day. Selecting this time interval for deorbiting control can satisfy a control method that combines multiple average amount maneuvers and correction amount maneuvers, and the time interval for control at the apogee and perigee is 12 hours apart.

[0043] The controllable time interval calculation method considering the influence of stray light on the star sensor is as follows:

[0044] Calculate the position of the sun vector in the orbit system

[0045] ​According to the star sensor installation layout design, the vector of the sun vector in the star sensor measurement coordinate system is represented as: In the formula, A st is a conversion matrix.

[0046] According to the star sensor affected by stray light conditions, the time period in which the star sensor is affected by stray light is determined, so as to determine the controllable time interval.

[0047] The sun vector The components of the star sensor measurement coordinate system x, y, z are respectively Define the out-of-solar-plane angle The in-solar-plane angle When the sun vector in the star sensor measurement coordinate system satisfies the following geometric relationship: α < 110° and -30° < β < 130°, the star sensor is affected by the sun stray light and cannot normally output information, and the time period corresponding to the remaining cases is the satellite controllable time period.

[0048] In step S102, according to the pressure and temperature data of the satellite propellant tank, the PVT (gas law) method is used to calculate the remaining fuel of the satellite. For the remaining fuel calculation result, the remaining fuel calculation error and the propellant extrusion efficiency need to be considered. According to the actual situation of the satellite in orbit, the calculation is corrected: the PVT method is used to calculate the remaining amount of propellant, the theoretical calculation error is 1.9% of the tank, and the extrusion efficiency index of the tank is 99%, and the minimum usable remaining fuel of the satellite is obtained. The formula of the minimum usable remaining fuel of the satellite is:

[0049] Δm = Δm0- Δm0·0.019- Δm0·0.01 (1)

[0050] In the formula, Δm is the minimum usable remaining fuel of the satellite, and Δm0 is the propellant consumption.

[0051] Given the satellite constants such as mass and specific impulse, and the satellite remaining fuel calculated in step two, the available velocity increment of the satellite is calculated using the momentum conservation formula:

[0052] Δm·I sp = F·Δt = m·ΔV (2)

[0053]

[0054] In the formula, I sp The specific impulse of the thruster; F is the engine thrust, Δt is the working time of the thruster, m is the mass of the satellite, and ΔV is the available velocity increment.

[0055] In step S103, the key of deorbiting control by out-of-plane thrusters is to yaw the satellite by a small angle a, and to realize the raising of semi-major axis and the discharge of fuel simultaneously under the condition of step one calculation. The tangential semi-major axis target control amount Δa T :

[0056] Δa T = 2a s V T / V s (4)

[0057] It can be known that: Figure 2

[0058] V T = V sin a (5)

[0059] V N = V cos a (6)

[0060] Wherein, V T is the tangential velocity increment, V N is the normal velocity increment, V is the total velocity increment of deorbiting control, V s is the nominal orbit velocity, a s is the nominal orbit altitude.

[0061] From formula (5), it can be known that V:

[0062] V = V T / sin a = Δa T V s / (2a s · sin a) (7)

[0063] The normal velocity increment applied by north-south control changes the velocity direction and the velocity size at the same time, thus causing the change of semi-major axis, which has little effect on the semi-major axis of the satellite when the normal velocity increment is small, and is generally not considered, but in the process of deorbiting control, the amount needs to be considered, and the influence of normal velocity increment on semi-major axis is set as Δa NT :

[0064]

[0065] The total semi-major axis change amount Δa and the total jet time length ΔT are respectively:

[0066] Δa = Δa T + Δa NT (9)

[0067] ΔT = mV / F (10)

[0068] In the formula, F is the engine thrust, and m is the satellite mass.​

[0069] According to the total calculated semi-major axis control amount and total jet time length, a control method combining multiple average amount maneuver control and correction amount maneuver control is adopted, control is performed at the interval of 12 hours at the far and near points, in order to ensure the reliability of control, generally, a batch of test jet is performed first, the remaining target control amount is evenly distributed for 2*N times of control implementation, so as to ensure that the eccentricity is less than 0.003.

[0070] In step S104, the purpose of control effect evaluation and error correction after each batch of control is to improve the control accuracy and efficiency of the subsequent batch and to successfully complete the de-orbiting control task of the satellite. The parameter error generally examined in de-orbiting control is semi-major axis error, and the errors caused by angle offset error and other factors are considered in the thruster efficiency factor based on semi-major axis. Engineering practice has proved that this method is feasible, and the thruster efficiency factor calculation formula is:

[0071]

[0072] In the formula, Δa R is the actual semi-major axis change amount (unit: kilometer), and Δa G is the target semi-major axis control amount (unit: kilometer).

[0073] The control parameter adjustment process based on the thruster efficiency factor evaluation of the semi-major axis is as follows:

[0074] Determine the post-control orbit;

[0075] Calculate the actual change amount of the post-control semi-major axis, and calculate the thruster efficiency factor according to formula (11);

[0076] Evaluate the thruster efficiency of this control. Generally, the efficiency factor is in the range of (0.85, 0.97] or [1.03, 1.05] for general accuracy, (0.97, 0.99] or [1.01, 1.03] for higher accuracy, and (0.99, 1.01) for high accuracy;

[0077] When the control accuracy is general, the thruster calibration coefficient of the next control batch needs to be adjusted. The adjustment method is to adjust the target semi-major axis control amount, and the calculation formula is Δa G = Δa' G / k, in which Δa G is the target semi-major axis control amount (unit: kilometer), Δa' G is the target semi-major axis control amount before adjustment (unit: kilometer), and k is the thruster efficiency factor.

[0078] The present application is applicable to the case where the tangential thruster on the satellite is abnormal and needs to be de-orbited by using the out-of-plane thruster, and the remaining fuel on the satellite is sufficient.

[0079] In one embodiment, the process of determining the satellite off-orbit control time interval by using the star sensor includes: calculating the solar vector position according to the parameters of the star sensor, and obtaining the relative angle between the solar vector position and the star sensor light axis; determining the time period in which the star sensor is disturbed by stray light according to the relative angle between the solar vector position and the star sensor light axis, thereby determining the satellite off-orbit control time interval. The relative angle between the solar vector position and the star sensor light axis includes a solar out-of-plane angle and a solar in-plane angle. In one embodiment, when the solar out-of-plane angle a < 110° and the solar in-plane angle -30° < β < 130°, the corresponding time is the time period in which the star sensor is disturbed by stray light.

[0080] Specifically, before the eastward position is maintained, a 90° yawing maneuver is required. To ensure normal operation of the star sensor, the working time interval of the tangential thruster is as follows:

[0081] During the period from March 10 to October 20 each year, the working time is from 21:00 to 3:00 the next morning in Beijing time; and the shadow period needs to be avoided.

[0082] During the period from October 20 to March 10 of the next year, the working time is from 17:00 to 3:00 the next morning in Beijing time.

[0083] When normal control is performed without yawing attitude maneuver or small-angle maneuver, to ensure normal operation of the star sensor, the working time interval of the normal thruster is as follows:

[0084] During the period from March 10 to October 20 each year, the working time is from 21:00 to 11:00 the next morning in Beijing time; and the shadow period needs to be avoided.

[0085] During the period from October 20 to March 10 of the next year, the working time is all day.

[0086] In one embodiment, the process of determining the minimum usable residual fuel amount of the satellite by using the characteristic parameters of the satellite includes: calculating the residual propellant amount by using the PVT method according to the characteristic parameters of the satellite, the characteristic parameters including the pressure and temperature data of the propellant tank; and calculating the minimum usable residual fuel amount of the satellite according to the residual propellant amount.

[0087] Specifically, the residual propellant amount of 50 kg is calculated by using the PVT method according to the pressure and temperature data of the propellant tank of the satellite, the theoretical calculation error of which is 1.9% of the tank, and the extrusion efficiency index of the tank is 99%, and the minimum usable residual fuel amount of the satellite is 47.68 kg.

[0088] The satellite mass is 500 kg, the specific impulse of the thruster is 2087.3 N / (kg / s), the available residual fuel is 47.68 kg, the thruster thrust is 10 N, and the satellite can generate a total speed increment of about 199.05 m / s and a total jet time of about 4796.12 seconds during control.

[0089] In one embodiment, the process of calculating the total semi-major axis control amount and the total jet time according to the available speed increment, the semi-major axis, and the satellite attitude offset angle comprises: obtaining a tangential speed increment according to the tangential semi-major axis target control amount; obtaining an out-of-orbit control total speed increment and a normal speed increment according to the tangential speed increment, and obtaining a normal semi-major axis target control amount; obtaining a total semi-major axis control amount according to the tangential semi-major axis target control amount and the normal semi-major axis target control amount, and obtaining the total jet time according to the out-of-orbit control total speed increment.

[0090] In one embodiment, the process of determining the satellite out-of-orbit control strategy comprises: using a control method combining multiple average amount maneuvers and correction amount maneuvers according to the total semi-major axis control amount and the total jet time. The control is performed at time intervals of 12 hours apart at the apogee and the perigee, 1 batch of trial injection is performed first, and the remaining target control amount is then evenly distributed according to 2*N control implementations.

[0091] Specifically, according to the limitation of the total jet time and the target of lifting the semi-major axis by 300 km, the satellite attitude offset angle is set to 4 degrees, the normal speed increment is considered to change the actual semi-major axis by 323.64 km, the total speed increment is 156.79 m / s, the control strategy is set to 1 batch of trial injection, and the remaining target control amount is evenly distributed for 4 control implementations. The specific control strategy is shown in Table 1.

[0092] Table 1: Out-of-orbit control strategy of a satellite

[0093]

[0094] In one embodiment, the process of calculating the thruster efficiency factor and adjusting the control parameters in the satellite out-of-orbit control strategy according to the thruster efficiency factor comprises: determining the post-control orbit, calculating the actual semi-major axis change amount after control, and obtaining the thruster efficiency factor; if the efficiency factor is in the range of (0.85, 0.97] or [1.03, 1.05), parameter adjustment is performed, and the parameters include the thruster calibration coefficient. The thruster calibration coefficient is adjusted by adjusting the target semi-major axis control amount.

[0095] Specifically, the test injection control orbit determination result in the above shows that the efficiency factor of the thruster is 1.13, and the first batch of orbit control is not error corrected considering various possible accidental factors. The orbit determination result after the first batch of control shows that the efficiency factor of the thruster is 1.15. Considering the control efficiency of the previous two times, the thruster calibration coefficient of the second control is corrected to 1.14. The orbit determination result after the control shows that the efficiency factor of the thruster after correction is 1.008. The calibration coefficient after correction is used in the subsequent two batches of control, and the efficiency factors of the thrusters are 1.05 and 1.001 respectively. The specific parameters in the control process are shown in Table 2.

[0096] Table 2 Parameters in the orbit control process of a satellite

[0097]

[0098] Through the above method of realizing orbit control by using out-of-plane thrusters through attitude bias, on the one hand, the orbit height of the GEO satellite after the orbit control is ensured to be above 280 kilometers above the synchronous belt, and the eccentricity is less than 0.003, thereby avoiding the problem that the satellite frequently crosses the synchronous belt due to the excessive eccentricity of the satellite. On the other hand, the problem that the control time is limited due to the influence of stray light on the star sensor is solved when the out-of-plane thrust orbit control is performed by using a large-angle attitude bias due to the abnormal tangential thruster.

[0099] In the description of the present specification, the description of the terms "one embodiment", "some embodiments", "an example", "a specific example", or "some examples" and the like means that the specific features, structures, materials or characteristics described in connection with the embodiment or example are included in at least one embodiment or example of the present disclosure. In the present specification, the illustrative description of the above terms does not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art can combine and combine different embodiments or examples described in the specification.

[0100] Other embodiments of the present disclosure will be readily apparent to those skilled in the art upon considering the specification and practicing the invention disclosed herein. The present application is intended to cover any variations, uses or adaptive changes to the present disclosure following the general principles of the present disclosure and including common knowledge or conventional techniques in the art that are not disclosed by the present disclosure. The specification and examples are only considered as exemplary, and the true scope and spirit of the present disclosure are indicated by the appended claims.

Claims

1. A method for derailment control using an out-of-plane thruster via attitude offset, characterized in that, include: The satellite deorbit control time interval is determined using a star sensor; where: Based on the parameters of the star sensor, the position of the solar vector is calculated, and the relative angle between the solar vector position and the star sensor optical axis is obtained; Based on the relative angle between the solar vector position and the star sensor optical axis, the time period during which the star sensor is affected by stray light is determined, thereby determining the satellite deorbit control time interval; the relative angle between the solar vector position and the star sensor optical axis includes: the outward angle and the inward angle. By using the satellite's characteristic parameters, the minimum amount of remaining usable fuel for the satellite is determined, and the available velocity increment for satellite deorbit control is calculated. Based on the available velocity increment, semi-major axis, and satellite attitude offset angle, the total semi-major axis control quantity and total jet duration are calculated, and the satellite deorbit control strategy is determined; wherein, the method for calculating the total semi-major axis control quantity and total jet duration is as follows: The tangential velocity increment is obtained based on the target control quantity of the tangential semi-major axis; Based on the tangential velocity increment, the total velocity increment and normal velocity increment of the derailment control are obtained, and the target control quantity of the normal semi-major axis is obtained. The total semi-major axis control quantity is obtained based on the tangential semi-major axis target control quantity and the normal semi-major axis target control quantity, and the total jet duration is obtained based on the total speed increment of the derailment control. Calculate the thruster efficiency factor, and adjust the control parameters in the satellite deorbit control strategy according to the thruster efficiency factor.

2. The method for derailment control using an out-of-plane thruster via attitude offset as described in claim 1, characterized in that, When the outward angle α of the solar surface is less than 110° and the inward angle -30° < β < 130°, the corresponding time period is the time during which the star sensor is interfered with by stray light.

3. The method for derailment control using an out-of-plane thruster via attitude offset as described in claim 1, characterized in that, The process of determining the minimum usable remaining fuel of a satellite using its characteristic parameters includes: Based on the satellite's characteristic parameters, the remaining propellant was calculated using the PVT method. These characteristic parameters included the pressure and temperature data of the propellant tank. Based on the remaining propellant, the minimum amount of remaining fuel that the satellite can use is calculated.

4. The method for derailment control using an out-of-plane thruster via attitude offset as described in claim 1, characterized in that, The process of determining the satellite deorbit control strategy includes: Based on the total semi-major axis control quantity and the total jet duration, a control method combining multiple averaging quantity maneuver control and correction quantity maneuver control is adopted.

5. The method for derailment control using an out-of-plane thruster via attitude offset as described in claim 4, characterized in that, Control was implemented by setting a 12-hour interval between near and far locations. A test spray was conducted first, and then the remaining target control amount was evenly distributed based on 2*N control operations.

6. The method for derailment control using an out-of-plane thruster via attitude offset as described in claim 1, characterized in that, The process of calculating the thruster efficiency factor and adjusting the control parameters in the satellite deorbit control strategy based on the thruster efficiency factor includes: After the trajectory is determined, the actual change in the semi-major axis after control is calculated, and the thruster efficiency factor is obtained. If the efficiency factor is in the range of (0.85, 0.97) or [1.03, 1.05), then parameter adjustments are made, including thruster calibration coefficients.

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