A mission-free design method for launch vehicle and satellite separation timing

Abstract

The present application relates to a kind of launch vehicle star rocket separation timing taskless design method.The present application is automatically designed rocket separation timing according to star rocket separation mode, the position and speed of rocket before first star rocket separation, separation speed, the mass of satellite and the total mass of rocket, satellite installation position in rocket, satellite quantity, can avoid using STK software to call rocket orbit UTC time and orbit six root numbers, rocket navigation coordinate system attitude angle such as trajectory parameter for far-field safety analysis, can effectively reduce operation amount and artificial design process, improve work efficiency.Meanwhile, the present method can design the timing of star rocket separation section for different launch tasks, so as to realize the taskless design of star rocket separation scheme.

Description

Technical Field

[0001] This invention relates to the field of rocket ballistics technology, specifically to a mission-independent design method for the separation sequence of a launch vehicle and its satellite. Background Technology

[0002] After reaching its predetermined orbit, the launch vehicle releases the satellite, implementing satellite-rocket separation. Because the relative velocity between the rocket and satellite is relatively low after separation, the rocket needs to glide at its current speed for a period of time. Subsequent actions can only proceed once the distance between the rocket and satellite, and the distance between satellites (considering the separation of multiple satellites), reaches a safe level. Therefore, a far-field safety analysis of the rocket-satellite and satellite-satellite distances is required to design the satellite-rocket separation sequence after the rocket enters orbit.

[0003] For the far-field safety of the satellite-rocket system, STK software is generally used for analysis. This involves reading ballistic parameters such as the rocket's UTC time of orbit insertion, orbital six-pointer, and the rocket's attitude angles in the navigation coordinate system at separation. The software outputs the distances between the rocket and the satellite, and between the satellites themselves. Through manual judgment and iterative analysis, appropriate satellite-rocket separation schemes are derived to ensure that the far-field safety requirements are met for both the rocket and satellites, and between the satellites. However, this method has drawbacks: for different launch missions, the number of satellites and the separation method vary, requiring designers to use STK software for judgment and iterative analysis to develop appropriate separation schemes. This process is time-consuming, labor-intensive, and has a long design cycle. Summary of the Invention

[0004] This invention proposes a mission-independent design method for satellite-rocket separation timing, which can automatically design the satellite-rocket separation timing based on parameters such as the satellite-rocket separation method, the rocket's position and velocity just before the first satellite-rocket separation, the separation velocity, the satellite's mass and the rocket's total mass, the satellite's installation position inside the rocket, and the number of satellites, thereby meeting the mission-independent requirements for satellite-rocket separation.

[0005] A mission-independent design method for the separation sequence of a launch vehicle and its satellite, comprising:

[0006] An orbital coordinate system is established, and three attitude angles A, B, and C of the rocket body are set within the orbital coordinate system. The attitude angles A, B, and C at the moment of rocket separation are determined according to the rocket separation method.

[0007] Transform the attitude angles A, B, and C in the orbital coordinate system to the navigation coordinate system, and calculate the attitude angles from the rocket body coordinate system to the navigation coordinate system based on the components of A, B, and C in the navigation coordinate system.

[0008] The velocities of the satellite and rocket in the navigation coordinate system at the moment of separation are calculated based on the attitude angle from the rocket body coordinate system to the navigation coordinate system at the time of separation of each satellite, the satellite separation velocity, the rocket velocity just before each satellite-rocket separation, the mass of the satellite and the total mass of the rocket, and the installation position of the satellite inside the rocket.

[0009] The separation sequence is designed based on the number of satellites and the positions and velocities of the rocket and satellites at each separation time.

[0010] Furthermore, the orbital coordinate system is defined as follows: the origin is the rocket's center of mass, the X-axis points to the direction of the rocket's velocity vector at separation; the Y-axis is in the instantaneous orbital plane, points to the outside of the orbit, and is perpendicular to the X-axis; the Z-axis is perpendicular to both the X-axis and the Y-axis, satisfying the right-hand rule.

[0011] Furthermore, the attitude angles A, B, and C are set as follows:

[0012] Angle A is the angle between the projection of the rocket body's X-axis onto the orbital plane and the X-axis of the orbital coordinate system;

[0013] Angle B is the angle between the arrow body's X-axis and the orbital plane;

[0014] Angle C is the rotation angle of the rocket body around the X-axis. When C = 0, the Y-axis of the rocket body is in the orbital plane.

[0015] 4. The launch vehicle satellite-rocket separation timing de-tasking design method according to claim 1, characterized in that, the transformation of attitude angles A, B, and C in the orbital coordinate system to the navigation coordinate system includes:

[0016] Using A, B, and C as Euler angles, the transformation matrix between the rocket body coordinate system and the orbital coordinate system is obtained, and the coordinate basis of the satellite in the rocket body coordinate system is established. Transform to the orbital coordinate system using the transformation matrix between the rocket body coordinate system and the orbital coordinate system; base The components in the orbital coordinate system are denoted as:

[0017] Will Transformation from orbital coordinate system to geocentric coordinate system, base The components in the geocentric coordinate system are denoted as:

[0018] Will Transform from geocentric coordinate system to launch coordinate system, base The components in the launch coordinate system are denoted as:

[0019] Will Transform from launch coordinate system to navigation coordinate system, base The components in the navigation coordinate system are denoted as:

[0020] According to the base The attitude angles from the rocket body coordinate system to the navigation coordinate system are obtained by calculating the components within the navigation coordinate system:

[0021]

[0022]

[0023] γ=γ 0

[0024] Where, γ 0 The roll angle of the navigation coordinate system at the moment of satellite-rocket separation is determined in advance by the rocket's telemetry and control requirements and the satellite-rocket separation method.

[0025] Each coordinate system is defined as follows:

[0026] Geocentric coordinate system: The origin of the coordinate system is at the Earth's center Oe. OeXe lies in the equatorial plane and points to the prime meridian. The OeZe axis is perpendicular to the equatorial plane and coincides with the Earth's rotation axis. OeYe can be obtained by the right-hand rule.

[0027] Launch coordinate system: The origin is located at the launch origin. The OY axis is taken as a vertical line passing through the launch point, with upward as positive. The OX axis is perpendicular to the OY axis and points to the theoretical launch direction. The OZ axis, together with the OX and OY axes, forms a right-handed rectangular coordinate system.

[0028] Rocket body coordinate system: The origin of the coordinate system is located at the center of mass of the rocket. The OX1 axis is consistent with the longitudinal axis of symmetry of the rocket body and points towards the head. The OY1 axis is perpendicular to the OX1 axis, located in the longitudinal plane of symmetry of the rocket, and points upward. The OZ1 axis, together with the OX1 axis and the OY1 axis, forms a right-handed rectangular coordinate system.

[0029] Navigation coordinate system: The navigation coordinate system coincides with the launch coordinate system at the moment of rocket ignition. After ignition, the position of the origin of the coordinate system moves at the speed of the launch point at the moment of launch, while the directions of the coordinate axes OXd, OYd, and OZd remain unchanged.

[0030] Furthermore, the calculation of the velocities of the satellite and rocket in the navigation coordinate system immediately after separation, based on the attitude angle of each satellite at separation time, satellite separation velocity, rocket velocity instant before the first separation, satellite mass and rocket total mass, and satellite installation position within the rocket, includes:

[0031] In order to obtain ψ and γ are Euler angles. The coordinate basis of the satellite's installation position inside the rocket, sat_HJ{satx,saty,satz}, is transformed to the navigation coordinate system through the transformation matrix between the rocket body coordinate system and the navigation coordinate system. The components of the satellite's installation position basis in the navigation coordinate system are denoted as: sat1_HJ{satx1,saty1,satz1}.

[0032] Based on the satellite's mass and the rocket's total mass m_sat, m_HJ, and the satellite's separation velocity SEP_V, the momentum theorem is used to calculate the separation velocities Δv_sat and Δv_HJ obtained by the rocket and satellite immediately after separation.

[0033] Using sat1_HJ{satx1,saty1,satz1}, Δv_sat, Δv_HJ, and the rocket velocity HJV(HJVX, HJVY, HJVZ) instant before the first separation, calculate the velocities SATV1(SATVX1, SATVY1, SATVZ1) and HJV1(HJVX1, HJVY1, HJVZ1) instant after separation.

[0034] Furthermore, the design of the satellite-rocket separation sequence based on the number of satellites, the position and velocity of the rocket and satellite at the moment of the first satellite-rocket separation includes:

[0035] Numerical integration or orbit prediction is performed on the position and velocity of the rocket and each separated satellite immediately after each separation. After each satellite separation, the distances between all separated satellites and the distance between the rocket and the separated satellites are calculated. When the distance between the rocket and the separated satellites is greater than the safety distance L_HJTOWX, and the distance between the separated satellites is greater than the safety distance L_WXTOWX, the numerical integration or prediction duration is the interval ΔT between the current satellite separation and the next satellite separation. The above steps are repeated until the last satellite is separated.

[0036] Furthermore, the design of the satellite-rocket separation sequence based on the number of satellites, the position and velocity of the rocket and satellite at the moment of the first satellite-rocket separation includes designing the rocket deorbit passivation sequence: after the last satellite separates from the rocket, the interval ΔTLG between satellite-rocket separation and deorbit passivation is designed, and the rocket mass is updated to m_HJ-m_satall, where m_satall is the total mass of all satellites;

[0037] Numerical integration or orbit prediction is performed on the position and velocity of the rocket and each satellite immediately after separation. The distance between the rocket and each satellite, and the distance between all satellites are calculated in real time. When the distance between the rocket and each satellite reaches the safe distance L_HJTOWX and the distance between all satellites reaches the safe distance L_WXTOWX, the integration or prediction time is the interval △TLG from separation to deorbiting.

[0038] The advantages of this invention compared to the prior art are:

[0039] This invention automatically designs the rocket separation sequence based on the satellite-rocket separation method, the rocket's position and velocity just before the first separation, the separation velocity, the satellite's mass and the rocket's total mass, the satellite's installation position within the rocket, and the number of satellites. This avoids the need to use STK software to retrieve ballistic parameters such as the rocket's UTC insertion time, orbital six-point numbers, and rocket attitude angles at separation for far-field safety analysis. This effectively reduces computational load and manual design processes, improving work efficiency. Furthermore, this method can design the satellite-rocket separation sequence for different launch missions, thus achieving mission-independent design of the satellite-rocket separation scheme. Attached Figure Description

[0040] Figure 1 This is a flowchart illustrating the task-independent design process for the separation sequence of the launch vehicle and satellite in this invention. Detailed Implementation

[0041] A mission-independent design method for launch vehicle satellite-rocket separation timing, the design process is as follows: Figure 1 As shown, the specific steps include:

[0042] Step 1: Calculate the attitude angles of the rocket's navigation coordinate system at each separation based on the separation method and the rocket's position just before separation.

[0043] Due to constraints imposed by satellite requirements, satellite-rocket adapter design, and space-based telemetry and control, the separation attitude angles for each mission may differ, ultimately requiring the provision of separation attitude angles within the navigation coordinate system. Since it is difficult to directly design attitude angles based on constraints within the navigation coordinate system, a coordinate system that intuitively describes the rocket's attitude at separation is first needed, called the "orbit coordinate system."

[0044] Establish an orbital coordinate system: the origin is the rocket's center of mass, the X-axis points in the direction of the rocket's velocity vector at separation; the Y-axis is in the instantaneous orbital plane, points out of the orbit, and is perpendicular to the X-axis; the Z-axis is perpendicular to the X-axis and Y-axis, satisfying the right-hand rule.

[0045] Define three attitude angles A, B, and C within the orbital coordinate system. The three angles are set as follows:

[0046] Angle A is the angle between the projection of the rocket body's X-axis onto the orbital plane and the X-axis of the orbital coordinate system;

[0047] Angle B is the angle between the arrow body's X-axis and the orbital plane;

[0048] Angle C is the rotation angle of the rocket body around the X-axis. When C = 0, the Y-axis of the rocket body is in the orbital plane.

[0049] Based on the satellite-rocket separation method, angles A, B, and C are given. In this embodiment, the satellite-rocket separation method is as follows: After the rocket enters orbit, it rotates 90° around the Z-axis of the orbital coordinate system, that is, the rocket "raises its head" by 90°. At this time, A, B, and C are 90°, 0°, and 0° respectively. After determining the attitude angles of the three orbital coordinate systems, they need to be transformed to the navigation coordinate system. The steps are as follows:

[0050] Using A, B, and C as Euler angles, the transformation matrix between the rocket body coordinate system and the orbital coordinate system is obtained, and the coordinate basis of the satellite in the rocket body coordinate system is established. Transform to the orbital coordinate system using a transformation matrix between the rocket body coordinate system and the orbital coordinate system. Base The components in the orbital coordinate system are denoted as: base It depends on the satellite's installation location within the rocket;

[0051] The transformation matrix between the rocket body coordinate system and the orbital coordinate system is:

[0052]

[0053] Will Transformation from orbital coordinates to geocentric coordinates (the transformation matrix from orbital to geocentric coordinates requires the rocket's velocity and position at the moment of separation, i.e., the input at this point is the rocket's position (X_HJ, Y_HJ, Z_HJ) and velocity (Vx_HJ, Vy_HJ, Vz_HJ) just before separation), base The component within the geocentric system is denoted as:

[0054] From the position and velocity in the geocentric coordinate system at the moment of star-launch separation, calculate the six orbital roots of the geocentric system at that moment (requiring the orbital inclination i, right ascension of the ascending node Ω, and argument of perigee w). From the six orbital roots, obtain the transformation matrix from the orbital coordinate system to the geocentric coordinate system.

[0055]

[0056] Will Transforming from the geocentric coordinate system to the launch coordinate system, the components of the base in the launch coordinate system are denoted as:

[0057] The transformation matrix between the geocentric coordinate system and the launch coordinate system is:

[0058]

[0059] In the formula:

[0060] A0—Theoretical firing direction;

[0061] B0—Latitude of the launch point;

[0062] L0 – Longitude of the launch point.

[0063] Will The components of the base in the navigation coordinate system are denoted as follows:

[0064] The transformation matrix between the launch coordinate system and the navigation coordinate system is:

[0065] M FS2DH =A -1 BA

[0066]

[0067]

[0068] In the formula:

[0069] A0—Theoretical firing direction;

[0070] B0—Latitude of the launch point;

[0071] ω e —Earth's rotational angular velocity;

[0072] t — the time elapsed from ignition and launch to the current moment.

[0073] The coordinate systems in the above transformation process are defined as follows:

[0074] Geocentric coordinate system: The origin of the coordinate system is at the Earth's center Oe. OeXe lies in the equatorial plane and points to the prime meridian. The OeZe axis is perpendicular to the equatorial plane and coincides with the Earth's rotation axis. OeYe can be obtained by the right-hand rule.

[0075] Launch coordinate system: The origin is located at the launch origin. The OY axis is taken as a vertical line passing through the launch point, with upward as positive. The OX axis is perpendicular to the OY axis and points to the theoretical launch direction. The OZ axis, together with the OX and OY axes, forms a right-handed rectangular coordinate system.

[0076] Rocket body coordinate system: The origin of the coordinate system is located at the center of mass of the rocket. The OX1 axis is consistent with the longitudinal axis of symmetry of the rocket body and points towards the head. The OY1 axis is perpendicular to the OX1 axis, located in the longitudinal plane of symmetry of the rocket, and points upward. The OZ1 axis, together with the OX1 axis and the OY1 axis, forms a right-handed rectangular coordinate system.

[0077] Navigation coordinate system ( Figure 1 The navigation coordinate system (abbreviated as navigation system) coincides with the launch coordinate system at the moment of rocket ignition. After ignition, the position of the origin of the coordinate system moves at the speed of the launch point at the moment of launch, while the directions of the coordinate axes OXd, OYd, and OZd remain unchanged.

[0078] Calculate the attitude angles of the navigation coordinate system based on the components of the base in the navigation coordinate system:

[0079]

[0080]

[0081] γ=γ 0

[0082] Where, γ 0 The roll angle of the navigation coordinate system at the moment of satellite-rocket separation is determined in advance by the rocket's telemetry and control requirements and the satellite-rocket separation method; in this example, it is taken as -165°.

[0083] Step 2: Calculate the velocities of the satellite and rocket in the navigation coordinate system immediately after separation based on the attitude of each satellite in the navigation coordinate system at the time of separation, the separation velocity, the velocity of the rocket just before the first separation, the mass of the satellite and the total mass of the rocket, and the installation position of the satellite inside the rocket.

[0084] The result obtained in step one ψ and γ are Euler angles, and the coordinate basis of the satellite's installation position inside the rocket is sat_HJ{satx, saty, satz} (in this embodiment, sat_HJ is equivalent to the above). The satellite is converted to the navigation coordinate system using a transformation matrix between the rocket body coordinate system and the navigation coordinate system. In this embodiment, the satellite installation direction is located on the Y-axis of the rocket body coordinate system, and the coordinate basis is sat_HJ{0,1,0}. The components of the basis sat_HJ{satx,saty,satz} in the navigation coordinate system are denoted as sat1_HJ{satx1,saty1,satz1}.

[0085] The transformation matrix between the rocket body coordinate system and the navigation coordinate system is:

[0086]

[0087] The transformation matrix from the rocket body coordinate system to the navigation coordinate system is the same as the transformation matrix from the rocket body coordinate system to the orbit coordinate system; the difference lies in the Euler angles.

[0088] The subsequent calculations in this embodiment are all performed in the navigation coordinate system. In other embodiments, calculations can also be performed in other coordinate systems, depending on which coordinate system the provided separation velocity is relative to, to facilitate the calculation.

[0089] Based on the satellite's mass and the rocket's total mass m_sat, m_HJ, and the separation velocity SEP_V (satellite relative to rocket), the momentum theorem is used to calculate the separation velocities Δv_sat and Δv_HJ obtained by the rocket and satellite immediately after separation.

[0090] Using the components of the base in the navigation coordinate system, sat1_HJ{satx1,saty1,satz1}, the separation velocities of the rocket and satellite after separation, and the rocket's velocity HJV(HJVX, HJVY, HJVZ) instant before the first separation, calculate the velocities of the satellite and rocket (SATV1(SATVX1, SATVY1, SATVZ1) and HJV1(HJVX1, HJVY1, HJVZ1)) instant after separation (navigation coordinate system):

[0091] SATVX1=HJVX-Δv_sat*satx1

[0092] SATVY1=HJVY-Δv_sat*saty1

[0093] SATVZ1=HJVZ-Δv_sat*satz1

[0094] HJVX1=HJVX-Δv_HJ*satx1

[0095] HJVY1=HJVY-Δv_HJ*saty1

[0096] HJVZ1=HJVZ-Δv_HJ*satz1

[0097] Step 3: Design the separation sequence of the satellite and rocket based on the number of satellites, the position and velocity of the rocket and satellite at the moment of the first separation.

[0098] The flight sequence is designed based on the number of satellites, as shown in Table 1;

[0099] Table 1

[0100]

[0101] The calculation steps for rocket glide times such as △T2, △T3, ..., △TLG are as follows:

[0102] Numerical integration or orbit prediction is performed on the position and velocity of the rocket and each separated satellite immediately after each separation to determine whether the real-time distance between the rocket and the satellite and the distance between satellites meet the requirements. After each satellite separation, the distance between all separated satellites and the distance between the rocket and the separated satellites are calculated. When the distance between the rocket and the separated satellites is greater than the safe distance L_HJTOWX, and the distance between separated satellites is greater than the safe distance L_WXTOWX, the numerical integration or prediction duration is the interval ΔT between the current rocket separation and the next rocket separation. The above steps are repeated until the last satellite is separated.

[0103] Specifically, numerical integration or orbit prediction is performed on the positions and velocities of the rocket and the first satellite immediately after the initial separation. The distance between the rocket and the first satellite is calculated in real time (at this point, only one separated satellite exists, so the distance between separated satellites does not need to be calculated). When the distance between the rocket and the first satellite reaches a safe distance L_HJTOWX, the integration or prediction time is the interval ΔT2 between the first and second separations. In this example, L_HJTOWX is taken as 200m, and L_WXTOWX as 70m. After the second separation, the rocket mass is updated to m_HJ - m_sata - m_satb, where m_sata and m_satb are the satellite masses after the first and second separations, respectively. Steps one and two are repeated to calculate the velocities of the satellite and rocket immediately after the second separation. Numerical integration or orbit prediction is performed on the position and velocity of the rocket and each separated satellite immediately after the second separation. The distances between the rocket and each separated satellite, and between the first and second satellites, are calculated in real time. When the distances between the rocket and each separated satellite reach a safe distance L_HJTOWX, and the distances between the first and second satellites reach a safe distance L_WXTOWX, the integration or prediction time at this point is the interval ΔT3 between the second and third separations. This process is repeated, decreasing the number of satellites by 1 after each separation, until the number of satellites is no greater than 0, indicating that the last satellite has separated.

[0104] Design the rocket deorbit passivation timing △TLG: After the last satellite separates from the rocket, begin designing the interval △TLG between separation and deorbit passivation. Update the rocket mass to m_HJ - m_satall, where m_satall is the total mass of all satellites. Perform numerical integration or orbit prediction on the position and velocity of the rocket and each satellite immediately after separation. Calculate in real-time the distances L1, L2, L3... between the rocket and all satellites, and the distances L_1to2, L_1to3... between all satellites. When the distances between the rocket and all satellites reach the safe distance L_HJTOWX and the distances between all satellites reach the safe distance L_WXTOWX, the integrated or predicted time at this point is the interval △TLG between separation and deorbit passivation.

[0105] When there is only one satellite, after the first satellite-rocket separation, the rocket coasts in a ΔTLG and then deorbits. When multiple satellites are separating, the rocket separates the satellites sequentially according to the designed timing and separation attitude. After the last satellite is separated, the rocket coasts in a ΔTLG and then deorbits.

[0106] This invention is not limited to the embodiments described above. Those skilled in the art can make various improvements and modifications without departing from the principles of this invention, and these improvements and modifications are also considered within the scope of protection of this invention. Contents not described in detail in this specification are prior art known to those skilled in the art.

Claims

1. A method for de-tasking the design of launch vehicle satellite-rocket separation timing, characterized in that, include: Establish an orbital coordinate system, set three attitude angles A, B, and C of the rocket body within the orbital coordinate system, and determine the attitude angles A, B, and C at the moment of rocket separation according to the rocket separation method; Transform the attitude angles A, B, and C in the orbital coordinate system to the navigation coordinate system, and calculate the attitude angles from the rocket body coordinate system to the navigation coordinate system based on the components of A, B, and C in the navigation coordinate system. The velocities of the satellite and rocket in the navigation coordinate system at the moment of separation are calculated based on the attitude angle from the rocket body coordinate system to the navigation coordinate system at the time of separation of each satellite, the satellite separation velocity, the rocket velocity just before each satellite-rocket separation, the mass of the satellite and the total mass of the rocket, and the installation position of the satellite inside the rocket. The separation sequence of the satellite and rocket is designed based on the number of satellites and the positions and velocities of the rocket and satellites at each separation time. The orbital coordinate system is defined as follows: the origin is the rocket's center of mass; the X-axis points in the direction of the rocket's velocity vector at separation; the Y-axis lies in the instantaneous orbital plane, points outward from the orbit, and is perpendicular to the X-axis; the Z-axis is perpendicular to both the X-axis and Y-axis, satisfying the right-hand rule. The attitude angles A, B, and C are set as follows: Angle A is the angle between the projection of the rocket body's X-axis onto the orbital plane and the X-axis of the orbital coordinate system; Angle B is the angle between the arrow body's X-axis and the orbital plane; C is the rotation angle of the rocket body around the X-axis. When C=0, the Y-axis of the rocket body is in the orbital plane. The process of transforming the attitude angles A, B, and C in the orbital coordinate system to the navigation coordinate system includes: Using A, B, and C as Euler angles, the transformation matrix between the rocket body coordinate system and the orbital coordinate system is obtained, and the coordinate basis of the satellite in the rocket body coordinate system is established. Transform to the orbital coordinate system using the transformation matrix between the rocket body coordinate system and the orbital coordinate system; base The components in the orbital coordinate system are denoted as: ; Will Transformation from orbital coordinate system to geocentric coordinate system, base The components in the geocentric coordinate system are denoted as: ; Will Transform from geocentric coordinate system to launch coordinate system, base The components in the launch coordinate system are denoted as: ; Will Transform from launch coordinate system to navigation coordinate system, base The components in the navigation coordinate system are denoted as: ; According to the base The attitude angles from the rocket body coordinate system to the navigation coordinate system are obtained by calculating the components within the navigation coordinate system: ， in, The roll angle of the navigation coordinate system at the moment of satellite-rocket separation is determined in advance by the rocket's telemetry and control requirements and the satellite-rocket separation method; The calculation of the satellite and rocket velocities in the navigation coordinate system immediately after separation, based on the attitude angle of each satellite at separation time, satellite separation velocity, rocket velocity instantaneously before each separation, satellite mass and rocket total mass, and satellite installation position within the rocket, includes: In order to obtain Using Euler angles, the coordinate basis of the satellite's installation position inside the rocket, sat_HJ{satx,saty,satz}, is transformed to the navigation coordinate system through the transformation matrix between the rocket body coordinate system and the navigation coordinate system. The components of the satellite's installation position basis in the navigation coordinate system are denoted as: sat1_HJ{satx1,saty1,satz1}. Based on the satellite's mass, the rocket's total mass m_sat and m_HJ, and the satellite's separation velocity SEP_V, the momentum theorem is used to calculate the separation velocities of the rocket and satellite immediately after separation. ; From sat1_HJ{satx1,saty1,satz1}, Calculate the velocity of the rocket HJV (HJVX, HJVY, HJVZ) instant before each separation of the satellite and rocket, and calculate the velocities of the satellite and rocket SATV1 (SATVX1, SATVY1, SATVZ1) and HJV1 (HJVX1, HJVY1, HJVZ1) instant after separation. The design of the satellite-rocket separation sequence based on the number of satellites and the positions and velocities of the rocket and satellites at each separation time includes: Numerical integration or orbit prediction is performed on the position and velocity of the rocket and each separated satellite immediately after each separation. After each satellite separation, the distances between all separated satellites and the distance between the rocket and the separated satellites are calculated. When the distance between the rocket and the separated satellites is greater than the safety distance L_HJTOWX, and the distance between the separated satellites is greater than the safety distance L_WXTOWX, the numerical integration or prediction duration is the interval ΔT between the current satellite separation and the next satellite separation. The above steps are repeated until the last satellite is separated.

2. The launch vehicle satellite-rocket separation timing de-tasking design method according to claim 1, characterized in that... The design of the satellite-rocket separation sequence based on the number of satellites and the positions and velocities of the rocket and satellites at each satellite-rocket separation time includes designing the rocket deorbit passivation sequence: after the last satellite separates from the rocket, the interval ΔTLG between satellite-rocket separation and deorbit passivation is designed, and the rocket mass is updated to m_HJ-m_satall, where m_satall is the total mass of all satellites and m_HJ is the total mass of the rocket; Numerical integration or orbit prediction is performed on the position and velocity of the rocket and each satellite immediately after separation. The distance between the rocket and each satellite, and the distance between all satellites are calculated in real time. When the distance between the rocket and each satellite reaches the safe distance L_HJTOWX and the distance between all satellites reaches the safe distance L_WXTOWX, the integration or prediction time is the interval △TLG from separation to deorbiting.

3. A mission-independent design system for launch vehicle satellite-rocket separation timing, comprising designing the rocket-satellite separation timing using the design method described in any one of claims 1-2, characterized in that, include: The coordinate system establishment module is used to establish the orbital coordinate system, set the three attitude angles A, B, and C of the rocket body in the orbital coordinate system, and determine the attitude angles A, B, and C according to the star-rocket separation method. The attitude angle calculation module is used to transform the attitude angles A, B, and C in the orbit coordinate system to the navigation coordinate system, and calculate the attitude angles in the navigation coordinate system based on the components of A, B, and C in the navigation coordinate system. The velocity calculation module is used to calculate the velocities of the satellite and rocket in the navigation coordinate system immediately after separation, based on the attitude angle of each satellite at the time of separation, the satellite separation velocity, the rocket velocity just before each separation, the mass of the satellite and the total mass of the rocket, and the satellite's installation position inside the rocket. The timing design module is used to design the separation timing of the satellite and rocket based on the number of satellites and the positions and velocities of the rocket and satellites at each separation time.

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