Method and device for determining satellite orbit and satellite attitude after satellite-rocket separation

Abstract

The application relates to a satellite orbit and satellite attitude determination method and device after satellite-rocket separation. The method comprises the following steps: calculating a first position vector and a first velocity vector of a satellite in an epoch equatorial coordinate system according to orbit elements of a rocket in a second earth-centered equatorial coordinate system and a satellite-rocket separation time; determining a second position vector and a second velocity vector of the satellite in a J2000 earth-centered inertial coordinate system based on the first position vector and the first velocity vector; taking the second position vector and the second velocity vector as orbit initial values, and determining orbit parameters of the satellite before injection into an orbit by using an orbit recursion algorithm; determining an attitude direction cosine matrix of the satellite in the J2000 earth-centered inertial coordinate system at the satellite-rocket separation time based on a launch time; and taking the attitude direction cosine matrix as attitude initial values, and determining attitude parameters of the satellite before injection into the orbit by using a star gyroscope combination filtering method. The scheme can quickly determine the orbit and the attitude of the satellite, and improves the autonomous detection capability of the satellite.

Description

Technical Field

[0001] This invention relates to the field of spacecraft control, and in particular to a method and apparatus for determining the satellite orbit and attitude after separation from the launch vehicle. Background Technology

[0002] In the initial stage after separation from the launch vehicle and entering orbit, the satellite is unable to complete autonomous navigation due to inaccurate orbital data and the unavailability of star sensors. During this initial orbital phase, the satellite has diverse needs for rapid alignment with the sun, tracking and control, and pointing to various different targets.

[0003] Therefore, how to quickly determine the satellite's orbit and attitude after separation from the launch vehicle is a technical problem that urgently needs to be solved. Summary of the Invention

[0004] This invention provides a method and apparatus for determining a satellite's orbit and attitude after separation from the launch vehicle, which can quickly determine the satellite's orbit and attitude and improve the satellite's autonomous detection capabilities.

[0005] In a first aspect, embodiments of the present invention provide a method for determining a satellite's orbit and attitude after separation from the launch vehicle, comprising:

[0006] The orbital elements of the rocket in the true equatorial coordinate system of the epoch are determined based on the rocket's orbital elements in the second geocentric equatorial coordinate system and the time of separation between the rocket and the satellite.

[0007] Based on the orbital elements of the rocket in the true equatorial coordinate system of the specified epoch, calculate the first position vector and the first velocity vector of the satellite in the true equatorial coordinate system of the specified epoch;

[0008] Based on the moment of separation of the satellite from the launch vehicle, the first position vector, and the first velocity vector, the second position vector and the second velocity vector of the satellite in the J2000 geocentric inertial coordinate system are determined.

[0009] Using the second position vector and the second velocity vector as initial orbit values, the orbital parameters of the satellite before injection into the orbit are determined by an orbital recursion algorithm.

[0010] Calculate the first direction cosine matrix between the J2000 geocentric inertial coordinate system and the geocentric coordinate system, the second direction cosine matrix between the geocentric coordinate system and the launch coordinate system, the third direction cosine matrix between the launch inertial coordinate system and the rocket body coordinate system, and the fourth direction cosine matrix between the rocket body coordinate system and the celestial coordinate system in sequence.

[0011] Based on the launch time, the first direction cosine array, the second direction cosine array, the third direction cosine array, and the fourth direction cosine array, the attitude direction cosine array of the satellite in the J2000 geocentric inertial coordinate system at the moment of separation of the satellite and the rocket is determined.

[0012] Using the attitude direction cosine array as the initial attitude value, the attitude parameters of the satellite before injection orbit are determined by the star-sensitive gyroscope combined filtering method.

[0013] Secondly, embodiments of the present invention also provide a device for determining the satellite orbit and attitude after separation of the satellite from the launch vehicle, comprising:

[0014] The first determining unit is used to determine the rocket's orbital elements in the true equatorial coordinate system based on the rocket's orbital elements in the second geocentric equatorial coordinate system and the time of separation between the rocket and the satellite.

[0015] The first calculation unit is used to calculate the first position vector and the first velocity vector of the satellite in the true equatorial coordinate system based on the orbital elements of the rocket in the true equatorial coordinate system of the epoch.

[0016] The second calculation unit is used to determine the second position vector and the second velocity vector of the satellite in the J2000 geocentric inertial coordinate system based on the satellite-rocket separation time, the first position vector, and the first velocity vector.

[0017] The orbit parameter determination unit is used to determine the orbit parameters of the satellite before injection into the orbit by using the second position vector and the second velocity vector as initial orbit values ​​and employing an orbit recursion algorithm.

[0018] The third calculation unit is used to sequentially calculate the first direction cosine matrix between the J2000 geocentric inertial coordinate system and the geocentric coordinate system, the second direction cosine matrix between the geocentric coordinate system and the launch coordinate system, the third direction cosine matrix between the launch inertial coordinate system and the rocket body coordinate system, and the fourth direction cosine matrix between the rocket body coordinate system and the celestial coordinate system.

[0019] The second determining unit is used to determine the attitude direction cosine array of the satellite in the J2000 geocentric inertial coordinate system at the time of satellite-rocket separation based on the launch time, the first direction cosine array, the second direction cosine array, the third direction cosine array and the fourth direction cosine array.

[0020] The attitude parameter determination unit is used to determine the attitude parameters of the satellite before injection into orbit by using the attitude direction cosine array as the initial attitude value and the star-sensitive gyroscope combined filtering method.

[0021] Thirdly, embodiments of the present invention also provide an electronic device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, it implements the method described in any embodiment of this specification.

[0022] Fourthly, embodiments of the present invention also provide a computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the methods described in any embodiment of this specification.

[0023] This invention provides a method and apparatus for determining a satellite's orbit and attitude after separation from the launch vehicle. By injecting the rocket launch time and satellite-launch separation-related parameters into the satellite, and performing multi-coordinate system transformations and other related processing on the rocket and satellite parameters, the initial orbit and initial attitude of the satellite can be determined. Based on the initial orbit and initial attitude, a high-precision orbit recursion algorithm and a star-sensitive gyroscope combined filtering method are used to quickly calculate the satellite's orbital and attitude parameters before orbit injection, enabling autonomous navigation and meeting the needs of solar search, telemetry, and orientation after separation. Therefore, this solution can quickly determine the satellite's orbit and attitude, improving the satellite's autonomous detection capabilities. Attached Figure Description

[0024] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0025] Figure 1 This is a schematic diagram of a method for determining the satellite orbit and attitude after separation from the launch vehicle, provided in an embodiment of the present invention.

[0026] Figure 2 This is a hardware architecture diagram of an electronic device provided in an embodiment of the present invention;

[0027] Figure 3 This is a structural diagram of a device for determining the satellite orbit and attitude after separation from the launch vehicle, provided in an embodiment of the present invention. Detailed Implementation

[0028] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0029] Please refer to Figure 1 This invention provides a method for determining a satellite's orbit and attitude after separation from the launch vehicle. The method includes:

[0030] Step 100: Determine the rocket's orbital elements in the true equatorial coordinate system based on the rocket's orbital elements in the second geocentric equatorial coordinate system and the time of rocket-satellite separation.

[0031] Step 102: Calculate the first position vector and the first velocity vector of the satellite in the true equatorial coordinate system based on the orbital elements of the rocket in the true equatorial coordinate system of the epoch.

[0032] Step 104: Based on the satellite-rocket separation time, the first position vector, and the first velocity vector, determine the satellite's second position vector and second velocity vector in the J2000 geocentric inertial coordinate system;

[0033] Step 106: Using the second position vector and the second velocity vector as initial orbit values, the orbit parameters of the satellite before injection into the orbit are determined using an orbit recursion algorithm;

[0034] Step 108: Calculate the first direction cosine matrix between the J2000 geocentric inertial coordinate system and the geocentric coordinate system, the second direction cosine matrix between the geocentric coordinate system and the launch coordinate system, the third direction cosine matrix between the launch inertial coordinate system and the rocket body coordinate system, and the fourth direction cosine matrix between the rocket body coordinate system and the celestial coordinate system in sequence.

[0035] Step 110: Based on the launch time, the first direction cosine array, the second direction cosine array, the third direction cosine array, and the fourth direction cosine array, determine the attitude direction cosine array of the satellite in the J2000 geocentric inertial coordinate system at the time of satellite-rocket separation.

[0036] Step 112: Using the attitude direction cosine array as the initial attitude value, the attitude parameters of the satellite before injection orbit are determined by the star-sensitive gyroscope combined filtering method.

[0037] In this embodiment, by injecting the rocket launch time and satellite-rocket separation parameters into the satellite, and performing multi-coordinate system transformations and other related processing on the rocket and satellite parameters, the satellite's initial orbit and initial attitude can be determined. Based on the initial orbit and initial attitude, a high-precision orbit recursion algorithm and a star-sensitive gyroscope combined filtering method are used to quickly calculate the satellite's orbital and attitude parameters before orbit injection, enabling autonomous navigation and meeting the needs for solar search, telemetry, and orientation after satellite-rocket separation. Therefore, this scheme can quickly determine the satellite's orbit and attitude, improving the satellite's autonomous detection capabilities.

[0038] The following description Figure 1 The execution method of each step is shown.

[0039] First, for step 100, the orbital elements of the rocket in the epochal true equatorial coordinate system are determined based on the rocket's orbital elements in the second geocentric equatorial coordinate system and the time of separation between the rocket and the satellite.

[0040] In this step, the geocentric second equatorial coordinate system is defined as follows: origin OE For the Earth's core, O E Z DX2 The axis points to the North Pole, O E X DX2 The axis points in the equatorial plane to the prime meridian at the moment of separation of the satellite and the rocket, O E Y DX2 As determined by the right-hand rule, all parameters related to rocket orbits are defined in the geocentric second equatorial coordinate system. Typically, rocket orbital parameters are defined in the geocentric second equatorial coordinate system.

[0041] The true equatorial coordinate system of the epoch is defined as: O E For the Earth's core, X TOD The line of intersection between the ecliptic plane and the true equatorial plane at an epoch is called the true vernal equinox. E X TOD From the Earth's center to the true vernal equinox, O E Z I Perpendicular to the true equatorial plane, coinciding with the Earth's axis of rotation, pointing towards the North Pole, O E Y TOD Determined by the right-hand rule.

[0042] Once the rocket launch time is known, the separation time is determined accordingly. Then, based on the rocket's orbital elements in the second geocentric equatorial coordinate system and the separation time, the rocket's orbital elements in the true equatorial coordinate system can be determined using the following formula:

[0043] a TOD =a DX2

[0044] e TOD =e DX2

[0045] incl TOD =incl DX2

[0046] Ω TOD =Ω eDX2 +λ G (T0)

[0047] ω TOD =ω DX2

[0048] M TOD =M DX2

[0049] In the formula, a is the semi-major axis of the orbit, e is the eccentricity of the orbit, incl is the orbital inclination, Ω is the right ascension of the ascending node, ω is the argument of perigee, M is the mean perigee angle, the subscript TOD indicates the true equatorial coordinate system, and the subscript DX2 indicates the geocentric second equatorial coordinate system; λ G(T0) is the Greenwich sidereal hour angle corresponding to the time T0 when the star separates from the rocket.

[0050] Regarding step 104, the J2000 geocentric inertial coordinate system is defined as: O E The intersection of the mean ecliptic plane and the mean equatorial plane at the J2000 epoch, centered on the Earth, is called the J2000 mean vernal equinox. E X I From the Earth's center to the vernal equinox, O E Z I Perpendicular to the equatorial plane, coinciding with the Earth's axis of rotation, pointing towards the North Pole, O E Y I As determined by the right-hand rule, all parameters related to the satellite orbit are defined in the J2000 geocentric inertial coordinate system.

[0051] The satellite's second position vector and second velocity vector in the J2000 geocentric inertial coordinate system are determined by the following formula:

[0052]

[0053]

[0054] In the formula, C P Let C be the precession matrix. N The precession matrix and the nutation matrix are determined based on the star-launch separation time T0; r TOD v is the first position vector; TOD r is the first velocity vector; I v is the second position vector; I This is the second velocity vector.

[0055] For step 106, the second position vector and the second velocity vector are used as initial orbit values. A high-precision orbit recursion algorithm is used for numerical integration to determine the orbit parameters of the satellite before injection into the orbit, which can be used for autonomous positioning and navigation.

[0056] For step 108, the geocentric coordinate system is defined as: O E For the Earth's core, X E The axis points to the intersection of the Prime Meridian and the Equator, Z. E The axis is parallel to the Earth's axis and points to the North Pole. E Y E Determined by the right-hand rule.

[0057] The launch coordinate system is defined as follows: the origin O is at the launch point, OX is in the horizontal plane at the launch point pointing in the launch aiming direction, the OY axis is along the vertical line of the launch point upwards, and the OZ axis is determined by the right-hand rule.

[0058] The launch inertial coordinate system is defined as follows: at the instant of rocket liftoff, O... A Point O coincides with launch point O, and its coordinate axes also coincide with the axes of the launch coordinate system. After the rocket lifts off, O... A The point and the coordinate system remain stationary in inertial space.

[0059] Based on the above definition of coordinate systems and the phase position relationship between each coordinate system, the first direction cosine matrix between the J2000 geocentric inertial coordinate system and the geocentric coordinate system, the second direction cosine matrix between the geocentric coordinate system and the launch coordinate system, the third direction cosine matrix between the launch inertial coordinate system and the rocket body coordinate system, and the fourth direction cosine matrix between the rocket body coordinate system and the celestial coordinate system can be determined in sequence.

[0060] In some implementations, the specific calculation process is as follows:

[0061] (1) The first direction cosine matrix C between the J2000 geocentric inertial coordinate system and the geocentric coordinate system EI It is determined by the following formula:

[0062] C EI =C G *C N *C P

[0063] In the formula, C G Let C be the Greenwich stellar time angle matrix. P Let C be the precession matrix. N The nutation matrix is ​​defined as follows, and the precession matrix and the nutation matrix are determined based on the star-launch separation time.

[0064] (2) The second direction cosine matrix C between the geocentric coordinate system and the launch coordinate system GE It is determined by the following formula:

[0065]

[0066] In the formula, R X R Y R Z These represent the direction cosine matrices rotating around the X, Y, and Z axes, respectively. α0 is the angle between the OX axis and the north tangent of the meridian passing through the launch point O, i.e., the geocentric azimuth. The OX axis is the direction pointing to the launch aiming direction in the horizontal plane of the launch point under the launch coordinate system. Let λ0 be the geocentric latitude of launch point O on the Earth's surface, and λ0 be the available longitude of launch point O on the Earth's surface.

[0067] (3) The third-direction cosine matrix C between the launch inertial coordinate system and the rocket body coordinate system BG (T0) are determined by the following formulas:

[0068]

[0069] In the formula, R X R Y R Z Let represent the direction cosine matrices of rotation about the X-axis, Y-axis, and Z-axis, respectively. T0 is the moment of separation between the satellite and the launch vehicle. γ(T0), ψ(T0) These are the rocket's roll angle, yaw angle, and pitch angle at the moment of separation from the satellite.

[0070] (4) The fourth-direction cosine matrix C between the rocket body coordinate system and the star coordinate system SaB It was determined in the following way:

[0071] Make the satellite's Z-axis coincide with the rocket's X-axis, the satellite's X-axis coincide with the rocket's Y-axis, and the satellite's Y-axis coincide with the rocket's Z-axis.

[0072] The fourth direction cosine matrix between the arrow body coordinate system and the star body coordinate system is determined as follows:

[0073]

[0074] Regarding step 110, the attitude direction cosine array C SaI It was determined in the following way:

[0075] Based on the fact that the launch inertial frame coincides with the launch system at the satellite launch time, the first direction cosine array C at the satellite launch time is determined. EI =C EI (T F );

[0076] Based on the fact that the launch inertial frame remains unchanged after satellite launch, the attitude orientation cosine matrix C of the satellite in the J2000 geocentric inertial coordinate system at the moment of satellite-rocket separation is determined. SaI for:

[0077] C SaI =C SaB *C BG (T0)*C GE *C EI (T F )

[0078] In the formula, T0 is the moment of separation between the satellite and the rocket, and T F The time of satellite launch is specified; the attitude direction cosine array is in quaternion form.

[0079] Finally, for step 112, the attitude direction cosine array is used as the initial attitude value, and the attitude parameters of the satellite before injection into the orbit are determined by the star-sensitive gyroscope combined filtering method for autonomous positioning and navigation.

[0080] like Figure 2 , Figure 3 As shown, this invention provides a device for determining a satellite's orbit and attitude after separation from the launch vehicle. The device can be implemented in software, hardware, or a combination of both. From a hardware perspective, as... Figure 2 The diagram shown is a hardware architecture diagram of an electronic device for determining the satellite orbit and attitude after separation from the launch vehicle, provided in an embodiment of the present invention. Except for... Figure 2 In addition to the processor, memory, network interface, and non-volatile memory shown, the electronic device in the embodiment may also include other hardware, such as a forwarding chip responsible for processing packets. Taking software implementation as an example, such as... Figure 3 As shown, a device in a logical sense is formed by the CPU of the electronic device in which it is located reading the corresponding computer program from the non-volatile memory into the memory for execution.

[0081] This embodiment provides a device for determining a satellite's orbit and attitude after separation from the launch vehicle, comprising:

[0082] The first determining unit 300 is used to determine the orbital elements of the rocket in the true equatorial coordinate system based on the rocket's orbital elements in the second geocentric equatorial coordinate system and the time of separation between the rocket and the satellite.

[0083] The first calculation unit 302 is used to calculate the first position vector and the first velocity vector of the satellite in the true equatorial coordinate system based on the orbital elements of the rocket in the true equatorial coordinate system of the epoch.

[0084] The second calculation unit 304 is used to determine the second position vector and the second velocity vector of the satellite in the J2000 geocentric inertial coordinate system based on the satellite-rocket separation time, the first position vector and the first velocity vector;

[0085] The orbit parameter determination unit 306 is used to use the second position vector and the second velocity vector as initial orbit values ​​and to determine the orbit parameters of the satellite before injection into the orbit using an orbit recursion algorithm.

[0086] The third calculation unit 308 is used to sequentially calculate the first direction cosine matrix between the J2000 geocentric inertial coordinate system and the geocentric coordinate system, the second direction cosine matrix between the geocentric coordinate system and the launch coordinate system, the third direction cosine matrix between the launch inertial coordinate system and the rocket body coordinate system, and the fourth direction cosine matrix between the rocket body coordinate system and the star coordinate system.

[0087] The second determining unit 310 is used to determine the attitude direction cosine array of the satellite in the J2000 geocentric inertial coordinate system at the time of satellite-rocket separation based on the launch time, the first direction cosine array, the second direction cosine array, the third direction cosine array and the fourth direction cosine array.

[0088] The attitude parameter determination unit 312 is used to determine the attitude parameters of the satellite before injection into orbit by using the attitude direction cosine array as the initial attitude value and the star-sensitive gyroscope combined filtering method.

[0089] In this embodiment of the invention, the first determining unit 300 can be used to execute step 100 in the above method embodiment, the first calculating unit 302 can be used to execute step 102 in the above method embodiment, the second calculating unit 304 can be used to execute step 104 in the above method embodiment, the orbital parameter determining unit 306 can be used to execute step 106 in the above method embodiment, the third calculating unit 308 can be used to execute step 108 in the above method embodiment, the second determining unit 310 can be used to execute step 110 in the above method embodiment, and the attitude parameter determining unit 312 can be used to execute step 312 in the above method embodiment.

[0090] In some implementations, the orbital elements in the true equatorial coordinate system of the epoch are determined by the following formula:

[0091] a TOD =a DX2

[0092] e TOD =e DX2

[0093] incl TOD =incl DX2

[0094] Ω TOD =Ω eDX2 +λ G (T0)

[0095] ω TOD =ω DX2

[0096] M TOD =M DX2

[0097] In the formula, a is the semi-major axis of the orbit, e is the eccentricity of the orbit, incl is the orbital inclination, Ω is the right ascension of the ascending node, ω is the argument of perigee, M is the mean perigee angle, the subscript TOD indicates the true equatorial coordinate system, and the subscript DX2 indicates the geocentric second equatorial coordinate system; λ G(T0) is the Greenwich Mean Time angle corresponding to the separation time T0 of the spacecraft and rocket. The separation time T0 is determined based on the time of rocket launch.

[0098] In some implementations, the satellite's second position vector and second velocity vector in the J2000 geocentric inertial coordinate system are determined by the following formula:

[0099]

[0100]

[0101] In the formula, C P Let C be the precession matrix. N The precession matrix and the nutation matrix are determined based on the star-launch separation time; r TOD v is the first position vector; TOD r is the first velocity vector; I v is the second position vector; I This is the second velocity vector.

[0102] In some implementations, the first direction cosine matrix C between the J2000 geocentric inertial coordinate system and the geocentric coordinate system... EI It is determined by the following formula:

[0103] C EI =C G *C N *C P

[0104] In the formula, C G Let C be the Greenwich stellar time angle matrix. P Let C be the precession matrix. N The nutation matrix is ​​defined as follows, and the precession matrix and the nutation matrix are determined based on the star-launch separation time.

[0105] In some implementations, the second directional cosine matrix C between the geocentric coordinate system and the launch coordinate system GE And the third-direction cosine matrix C between the launch inertial coordinate system and the rocket body coordinate system. BG (T0) are determined by the following formulas:

[0106]

[0107]

[0108] In the formula, R X R Y R ZThese represent the direction cosine matrices rotating around the X, Y, and Z axes, respectively. α0 is the angle between the OX axis and the north tangent of the meridian passing through the launch point O, i.e., the geocentric azimuth. The OX axis is the direction pointing to the launch aiming direction in the horizontal plane of the launch point under the launch coordinate system. Let λO be the geocentric latitude of launch point O on the Earth's surface, and λ0 be the available longitude of launch point O on the Earth's surface; T0 be the time of separation between the satellite and the launch vehicle, and γ(T0), ψ(T0). These are the rocket's roll angle, yaw angle, and pitch angle at the moment of separation from the satellite.

[0109] In some implementations, the fourth-direction cosine matrix C between the rocket body coordinate system and the celestial coordinate system... SaB It was determined in the following way:

[0110] Make the satellite's Z-axis coincide with the rocket's X-axis, the satellite's X-axis coincide with the rocket's Y-axis, and the satellite's Y-axis coincide with the rocket's Z-axis.

[0111] The fourth direction cosine matrix between the arrow body coordinate system and the star body coordinate system is determined as follows:

[0112]

[0113] In some embodiments, the attitude direction cosine array C SaI It was determined in the following way:

[0114] Based on the fact that the launch inertial frame coincides with the launch system at the satellite launch time, the first direction cosine array C at the satellite launch time is determined. EI =C EI (T F );

[0115] Based on the fact that the launch inertial frame remains unchanged after satellite launch, the attitude orientation cosine matrix C of the satellite in the J2000 geocentric inertial coordinate system at the moment of satellite-rocket separation is determined. SaI for:

[0116] C SaI =C SaB *C BG (T0)*C GE *C EI (T F )

[0117] In the formula, T0 is the moment of separation between the satellite and the rocket, and T F The time of satellite launch is specified; the attitude direction cosine array is in quaternion form.

[0118] It is understood that the structures illustrated in the embodiments of the present invention do not constitute a specific limitation on a device for determining a satellite's orbit and attitude after separation from the launch vehicle. In other embodiments of the present invention, a device for determining a satellite's orbit and attitude after separation from the launch vehicle may include more or fewer components than illustrated, or combine some components, or split some components, or arrange the components differently. The illustrated components may be implemented in hardware, software, or a combination of software and hardware.

[0119] The information interaction and execution process between the modules in the above-mentioned device are based on the same concept as the method embodiment of the present invention, and the specific details can be found in the description of the method embodiment of the present invention, and will not be repeated here.

[0120] This invention also provides an electronic device, including a memory and a processor. The memory stores a computer program, and when the processor executes the computer program, it implements a method for determining the satellite orbit and satellite attitude after separation from the launch vehicle, according to any embodiment of this invention.

[0121] This invention also provides a computer-readable storage medium storing a computer program. When executed by a processor, the computer program causes the processor to perform a method for determining the satellite orbit and attitude after separation from the launch vehicle, according to any embodiment of this invention.

[0122] Specifically, a system or apparatus equipped with a storage medium may be provided, on which software program code implementing the functions of any of the embodiments described above is stored, and the computer (or CPU or MPU) of the system or apparatus may read and execute the program code stored in the storage medium.

[0123] In this case, the program code read from the storage medium can itself implement the function of any of the above embodiments, and therefore the program code and the storage medium storing the program code constitute part of the present invention.

[0124] Examples of storage media used to provide program code include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Alternatively, program code can be downloaded from a server computer via a communication network.

[0125] Furthermore, it should be clear that not only can the program code read by the computer be executed, but also the operating system or other components operating on the computer can be instructed based on the program code to perform some or all of the actual operations, thereby realizing the function of any of the embodiments described above.

[0126] Furthermore, it is understood that the program code read from the storage medium is written to the memory set in the expansion board inserted into the computer or to the memory set in the expansion module connected to the computer. Then, based on the instructions of the program code, the CPU or other components installed on the expansion board or expansion module execute some and all of the actual operations, thereby realizing the function of any of the above embodiments.

[0127] It should be noted that, in this document, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0128] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for determining a satellite's orbit and attitude after separation from the launch vehicle, characterized in that, include: The orbital elements of the rocket in the true equatorial coordinate system of the epoch are determined based on the rocket's orbital elements in the second geocentric equatorial coordinate system and the time of separation between the rocket and the satellite. Calculate the first position vector and the first velocity vector of the satellite in the true equatorial coordinate system based on the orbital elements of the rocket in the true equatorial coordinate system. Based on the satellite-rocket separation time, the first position vector, and the first velocity vector, the second position vector and the second velocity vector of the satellite in the J2000 geocentric inertial coordinate system are determined. Using the second position vector and the second velocity vector as initial orbit values, the orbit parameters of the satellite before injection into the orbit are determined by an orbit recursion algorithm. Calculate the first direction cosine matrix between the J2000 geocentric inertial coordinate system and the geocentric coordinate system, the second direction cosine matrix between the geocentric coordinate system and the launch coordinate system, the third direction cosine matrix between the launch inertial coordinate system and the rocket body coordinate system, and the fourth direction cosine matrix between the rocket body coordinate system and the celestial coordinate system in sequence. Based on the launch time, the first direction cosine array, the second direction cosine array, the third direction cosine array, and the fourth direction cosine array, the attitude direction cosine array of the satellite in the J2000 geocentric inertial coordinate system at the moment of separation of the satellite and the rocket is determined. Using the attitude direction cosine array as the initial attitude value, the attitude parameters of the satellite before injection orbit are determined by the star-sensitive gyroscope combined filtering method.

2. The method according to claim 1, characterized in that, The orbital elements in the true equatorial coordinate system of the epoch are determined by the following formula: a TOD =a DX2 And TOD =and DX2 incl TOD =incl DX2 Oh TOD =Oh eDX2 +λ G (T0) oh TOD =ω DX2 M TOD =M DX2 In the formula, a is the semi-major axis of the orbit, e is the eccentricity of the orbit, incl is the orbital inclination, Ω is the right ascension of the ascending node, ω is the argument of perigee, M is the mean perigee angle, the subscript TOD indicates the true equatorial coordinate system, and the subscript DX2 indicates the geocentric second equatorial coordinate system; λ G (T0) is the Greenwich Mean Time angle corresponding to the separation time T0 of the spacecraft and rocket. The separation time T0 is determined based on the time of rocket launch.

3. The method according to claim 2, characterized in that, The satellite's second position vector and second velocity vector in the J2000 geocentric inertial coordinate system are determined by the following formula: In the formula, C P Let C be the precession matrix. N The precession matrix and the nutation matrix are determined based on the star-launch separation time; r TOD v is the first position vector; TOD r is the first velocity vector; I v is the second position vector; I This is the second velocity vector.

4. The method according to claim 2, characterized in that, The first direction cosine matrix C between the J2000 geocentric inertial coordinate system and the geocentric coordinate system EI It is determined by the following formula: C EI =C G *C N *C P In the formula, C G Let C be the Greenwich stellar time angle matrix. P Let C be the precession matrix. N The nutation matrix is ​​defined as follows, and the precession matrix and the nutation matrix are determined based on the star-launch separation time.

5. The method according to claim 4, characterized in that, The second direction cosine matrix C between the geocentric coordinate system and the launch coordinate system GE And the third-direction cosine matrix C between the launch inertial coordinate system and the rocket body coordinate system. BG (T0) are determined by the following formulas: In the formula, R X R Y R Z These represent the direction cosine matrices rotating around the X, Y, and Z axes, respectively. α0 is the angle between the OX axis and the north tangent of the meridian passing through the launch point O, i.e., the geocentric azimuth. The OX axis is the direction pointing to the launch aiming direction in the horizontal plane of the launch point under the launch coordinate system. Let λO be the geocentric latitude of launch point O on the Earth's surface, and λ0 be the available longitude of launch point O on the Earth's surface; T0 be the time of separation between the satellite and the launch vehicle, and γ(T0), ψ(T0) These are the rocket's roll angle, yaw angle, and pitch angle at the moment of separation from the satellite.

6. The method according to claim 5, characterized in that, The fourth direction cosine matrix C between the rocket coordinate system and the celestial coordinate system SaB It was determined in the following way: Make the satellite's Z-axis coincide with the rocket's X-axis, the satellite's X-axis coincide with the rocket's Y-axis, and the satellite's Y-axis coincide with the rocket's Z-axis. The fourth direction cosine matrix between the arrow body coordinate system and the star body coordinate system is determined as follows:

7. The method according to claim 6, characterized in that, The attitude direction cosine array C SaI It was determined in the following way: Based on the fact that the launch inertial frame coincides with the launch system at the satellite launch time, the first direction cosine array C at the satellite launch time is determined. EI =C EI (T F ); Based on the fact that the launch inertial frame remains unchanged after satellite launch, the attitude orientation cosine matrix C of the satellite in the J2000 geocentric inertial coordinate system at the moment of satellite-rocket separation is determined. SaI for: C SaI =C SaB *C BG (T0)*C GE *C EI (T F ) In the formula, T0 is the moment of separation between the satellite and the rocket, and T F The time of satellite launch is specified; the attitude direction cosine array is in quaternion form.

8. A device for determining a satellite's orbit and attitude after separation from the launch vehicle, characterized in that, include: The first determining unit is used to determine the rocket's orbital elements in the true equatorial coordinate system based on the rocket's orbital elements in the second geocentric equatorial coordinate system and the time of separation between the rocket and the satellite. The first calculation unit is used to calculate the first position vector and the first velocity vector of the satellite in the true equatorial coordinate system based on the orbital elements of the rocket in the true equatorial coordinate system of the epoch. The second calculation unit is used to determine the second position vector and the second velocity vector of the satellite in the J2000 geocentric inertial coordinate system based on the satellite-rocket separation time, the first position vector, and the first velocity vector. The orbit parameter determination unit is used to determine the orbit parameters of the satellite before injection into the orbit by using the second position vector and the second velocity vector as initial orbit values ​​and employing an orbit recursion algorithm. The third calculation unit is used to sequentially calculate the first direction cosine matrix between the J2000 geocentric inertial coordinate system and the geocentric coordinate system, the second direction cosine matrix between the geocentric coordinate system and the launch coordinate system, the third direction cosine matrix between the launch inertial coordinate system and the rocket body coordinate system, and the fourth direction cosine matrix between the rocket body coordinate system and the celestial coordinate system. The second determining unit is used to determine the attitude direction cosine array of the satellite in the J2000 geocentric inertial coordinate system at the time of satellite-rocket separation based on the launch time, the first direction cosine array, the second direction cosine array, the third direction cosine array and the fourth direction cosine array. The attitude parameter determination unit is used to determine the attitude parameters of the satellite before injection into orbit by using the attitude direction cosine array as the initial attitude value and the star-sensitive gyroscope combined filtering method.

9. A computing device comprising a memory and a processor, wherein the memory stores a computer program, and the processor, when executing the computer program, implements the method as described in any one of claims 1-7.

10. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the method of any one of claims 1-7.

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