Far-field safety analysis method and system after satellite-rocket separation

Abstract

The invention discloses a far-field safety analysis method and system after satellite-rocket separation, and the method comprises the steps: estimating the time-varying satellite-rocket relative distance between a satellite and a carrier rocket after the satellite-rocket separation of the same batch; calculating an included angle between a satellite guiding direction and a thrust direction of the carrier rocket; wherein the satellite guiding direction is the direction in which the carrier rocket points to the satellite; and according to preset far-field safety judgment conditions corresponding to N stages after the same batch of satellites and rockets are separated, performing far-field safety analysis on one or more of the relative distance of the satellites and rockets and the included angle (engine plume influence safety).

Description

Technical Field

[0001] The present application relates to the technical field of safety analysis after satellite separation, and in particular to a far-field safety analysis method and system after satellite-rocket separation. Background Art

[0002] Far-field safety after satellite-rocket separation means that after the satellite and the launch vehicle separate in the predetermined orbit, the satellite should avoid collision with other satellites and launch vehicles, and avoid being affected by the plume of the launch vehicle engine. The satellite-rocket separation can divide the period that requires far-field safety analysis into several flight segments, and formulate different far-field safety conditions for different flight segments to make far-field safety judgments.

[0003] The current far-field safety analysis after satellite-rocket separation can be performed with the help of STK software, or a dynamic model can be established for integral solution.

[0004] However, although the STK software can analyze whether the safety requirements are met by combining the attitude relationship of the rocket after reading in the ballistic data and orbital parameters, the far-field safety will change for different launch missions due to differences in launch sites, target orbital inclinations, flight attitude angles, etc. Therefore, the STK software still needs to rely on manual judgment of far-field safety, which is time-consuming and labor-intensive. When the number of satellites is large and the far-field analysis time is long, the dynamic model will have the problem of taking a long time to solve.

[0005] Therefore, there is an urgent need for a far-field safety analysis technology that saves time and effort and does not require the use of STK software. Summary of the invention

[0006] In order to solve or partially solve the technical problem that the far-field safety analysis after the separation of satellites and rockets in the prior art needs to rely on manual labor, which is time-consuming and labor-intensive, the present invention provides a far-field safety analysis method and system after the separation of satellites and rockets. By designing preset far-field safety judgment conditions corresponding to N stages after the separation of satellites and rockets in the same batch, after the separation of satellites and rockets in the same batch, by estimating the relative distance between the satellite and the rocket that changes with time, and calculating the angle between the satellite guidance direction and the thrust direction of the launch vehicle, there is no need to use STK software or rely on manual judgment, but the preset far-field safety judgment conditions can be used to directly perform far-field safety analysis, which saves time and labor. Under the same working conditions and inputs, the calculation results of this scheme are slightly different from those of the STK software. In addition, this scheme is less affected by the number of satellites and is suitable for far-field safety analysis after launching any number of satellites.

[0007] In order to solve the above technical problems, the first aspect of the present invention discloses a far-field safety analysis method after satellite-rocket separation, characterized in that the method comprises:

[0008] After the separation of satellites and rockets from the same batch, estimate the relative distance between the satellite and the launch vehicle over time;

[0009] Calculating the angle between the satellite guidance direction and the thrust direction of the carrier rocket; wherein the satellite guidance direction is the direction in which the carrier rocket points to the satellite;

[0010] According to the preset far-field safety judgment conditions corresponding to the N stages after the separation of the same batch of satellites and rockets, a far-field safety analysis is performed on one or more of the satellite-rocket relative distance and the angle, where N≥3 and is a positive integer.

[0011] Optionally, if there are more than two satellites in the same batch; before estimating the satellite-rocket relative distance between the satellite and the launch vehicle over time, specifically including:

[0012] Estimate the relative distances between two satellites in the same batch over time;

[0013] The performing of a far-field safety analysis on one or more of the satellite-rocket relative distance and the included angle according to the preset judgment conditions corresponding to the N stages specifically includes:

[0014] According to the preset judgment conditions corresponding to the N stages, a far-field safety analysis is performed on one or more of the satellite relative distance, the satellite-rocket relative distance, and the angle.

[0015] Optionally, estimating the relative distances between satellites in the same batch that vary with time may include:

[0016] At the time t0 of satellite-rocket separation, the position velocities of satellites i and j in the launch system are converted into the position velocities of satellites i and j in the geocentric system; wherein time t0 is the time of satellite-rocket separation; satellites i and j are any two satellites from the same batch;

[0017] According to the position speed of the satellite i and the satellite j in the geocentric system, the instantaneous root of the satellite i and the satellite j at time t0 is calculated: Among them, the parameters of the instantaneous root include: semi-major axis a m , eccentricity e m , orbital inclination i m , right ascension of ascending node Ω m 、Arg of perigee ω m 、Mean anomaly angle M m ;

[0018] The instantaneous root of the satellite i and the satellite j at time t0 is: Converted to the square root of the satellite i and the satellite j at time t0:

[0019] According to the square root of the satellite i and the satellite j at time t0: Estimate the satellite i and the satellite j at t i The flat root of the moment: in, Δt represents t i The time difference between time t0 and time t1 is Δt = t2. i -t0;

[0020] The satellite i and the satellite j are placed at t i The flat root of the moment: Convert to the satellite i and the satellite j at t i The instant root of time:

[0021] According to the satellite i and the satellite j at t i The instant root of time: Determine the satellite i and the satellite j at t i The location at the moment;

[0022] According to the satellite i and the satellite j at t i The relative distance S between the satellite i and the satellite j changes with time. ij .

[0023] Optionally, the square root of the satellite i and the satellite j at time t0 is: Estimate the satellite i and the satellite j at t i The flat root of the moment: Specifically include:

[0024] According to the square root of the satellite i at time t0: Determine the following parameters of the satellite i: the rate of change of the right ascension of the ascending node W Ω , the rate of change of the perigee argument W ω , the rate of change of the mean perigee angle W M :

[0025]

[0026]

[0027] Where ω0 represents the angular velocity of the earth's rotation; J2 represents the second-order term in the earth's shape dynamics coefficient, which is a constant; A e represents the semi-major axis of the Earth's rotation ellipsoid, which is a constant;

[0028] Combined with i The time difference Δt between time t0 and time t1 determines the time difference between the satellite i and time t2. i Pinggen of the moment The parameters included in:

[0029] a m (Δt) = a m

[0030] e m (Δt) = e m

[0031] i m (Δt)=i m

[0032] Ω m (Δt)=Ω m +W Ω Δt

[0033] ω m (Δt)=ω m +W ω Δt

[0034] M m (Δt) = M m +W M Δt; where W Ω Δt represents the change in right ascension of the ascending node, W ω Δt represents the change in the argument of perigee, W M Δt represents the change in the mean perigee angle;

[0035] The satellite j is at t i Pinggen of the moment The estimation method and the satellite i at t i Pinggen of the moment The estimation method is the same.

[0036] Optionally, the estimating the relative distance between the satellite and the launch vehicle as it changes over time specifically includes:

[0037] By integrating the earth's gravity, engine thrust, aerodynamic force and the mass of the launch vehicle, the launch vehicle t is estimated. i The velocity of the launch system at the moment, and based on the launch vehicle t i The velocity integral of the launch system at the time is obtained as i The position under the launch system at all times;

[0038] The launch vehicle t i The position of the launch system at the moment, converting the launch vehicle t iThe position of the earth at all times;

[0039] According to the satellite i at t i The position of the launch vehicle at time t i The position at the moment, determines the relative distance S between the satellite i and the launch vehicle that changes with time i ; Wherein, the satellite i is any satellite from the same batch.

[0040] Optionally, the calculating the angle between the satellite guidance direction and the thrust direction of the launch vehicle specifically includes:

[0041] According to the satellite i at t i The position of the launch vehicle at time t i The position at the moment, find the vector of the satellite guidance direction in the geocentric system

[0042] The vector of the satellite guidance direction in the geocentric system Converted into the vector of the satellite guidance direction in the arrow system

[0043] Determine the vector of the launch vehicle thrust in the arrow system

[0044] According to the formula Determine the angle; where β i It represents the angle between the satellite guidance direction formed by satellite i and the launch vehicle and the thrust direction of the launch vehicle, acos() represents the inverse cosine formula, and || represents the modulus.

[0045] Optionally, the N stages include: a glide stage from the separation of the satellite and the rocket to the startup of the carrier rocket, an active stage from the startup to the shutdown of the carrier rocket, and an unpowered glide stage after the shutdown of the carrier rocket;

[0046] In the glide section after the separation of the satellite and the rocket and before the launch of the carrier rocket, the corresponding preset far-field safety judgment condition is: the relative distance S between the satellite and the rocket i >L1; or: the relative distance between the star and the arrow S i >L1 and the relative distance between satellites is S ij >L1; wherein L1 represents the first preset safety distance;

[0047] In the active section between the start-up and shutdown of the launch vehicle, the corresponding preset far-field safety judgment condition is one of the following:

[0048] Wherein, L2 represents the second preset safety distance, L3 represents the third preset safety distance, L4 represents the fourth preset safety distance, αsafe The cone angle in the plume influence range; wherein the plume influence range has an arc length of L2 and a half angle of α safe of cone;

[0049] During the unpowered gliding phase after the launch vehicle is shut down, the corresponding preset far-field safety judgment condition is: S i >L5; or: S i >L5, S ij >L6; wherein, L5 represents the fifth preset safety distance; L6 represents the sixth preset safety distance.

[0050] Optionally, the performing of a far-field safety analysis on one or more of the satellite-rocket relative distance and the angle according to preset far-field safety judgment conditions corresponding to N stages after the satellite-rocket separation of the same batch specifically includes:

[0051] With t i The time change Δt between time t and time t0 is a time unit. Each time unit is combined with the preset far-field safety judgment conditions corresponding to the stage to perform a far-field safety analysis on one or more of the satellite-rocket relative distance and the angle.

[0052] Optionally, the method further includes:

[0053] In the unpowered gliding section after the launch vehicle is shut down, the relative distance S between the rocket and the satellite is used. i and / or the relative distance S of the satellites ij Conduct far-field safety analysis on the changing trend of

[0054] If the relative distance between the star and the arrow is S i is monotonically increasing, and / or the satellite relative distance S ij If the changing trend of is periodically increasing, it means that the safety requirements are met.

[0055] In a second aspect of the present invention, a far-field safety analysis system after satellite-rocket separation is disclosed, the system comprising:

[0056] An estimation module, used for estimating the relative distance between the satellite and the launch vehicle over time after the satellite and the rocket of the same batch are separated;

[0057] A calculation module, used to calculate the angle between the satellite guidance direction and the thrust direction of the carrier rocket; wherein the satellite guidance direction is the direction in which the carrier rocket points to the satellite;

[0058] The analysis module is used to perform far-field safety analysis on one or more of the relative distance between the satellite and the rocket and the angle according to the preset far-field safety judgment conditions corresponding to the N stages after the separation of the satellite and the rocket from the same batch, where N≥3 and is a positive integer.

[0059] Through one or more technical solutions of the present invention, the present invention has the following beneficial effects or advantages:

[0060] The present invention provides a far-field safety analysis method and system after satellite-rocket separation. By designing preset far-field safety judgment conditions corresponding to N stages after the same batch of satellite-rocket separation, after the same batch of satellite-rocket separation, by estimating the satellite-rocket relative distance between the satellite and the carrier rocket that changes with time, and calculating the angle between the satellite guidance direction and the thrust direction of the carrier rocket, there is no need to use STK software or rely on manual judgment, but the preset far-field safety judgment conditions can be used to directly perform far-field safety analysis, saving time and labor. Under the same working conditions and inputs, the calculation results of this scheme are slightly different from those of the STK software. In addition, this scheme is less affected by the number of satellites and is suitable for far-field safety analysis after launching any number of satellites.

[0061] The above description is only an overview of the technical solution of the present invention. In order to more clearly understand the technical means of the present invention, it can be implemented according to the contents of the specification. In order to make the above and other purposes, features and advantages of the present invention more obvious and easy to understand, the specific implementation methods of the present invention are listed below. BRIEF DESCRIPTION OF THE DRAWINGS

[0062] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present invention. Moreover, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:

[0063] Figure 1 A flow chart of a far-field safety analysis method after satellite-rocket separation according to an embodiment of the present invention is shown;

[0064] Figure 2 A schematic diagram showing a variation trend of the relative distance between a satellite and an arrow according to an embodiment of the present invention;

[0065] Figure 3 A schematic diagram showing a satellite relative distance variation trend according to an embodiment of the present invention;

[0066] Figure 4 A schematic diagram of a far-field safety analysis system after satellite-rocket separation according to an embodiment of the present invention is shown. DETAILED DESCRIPTION

[0067] The exemplary embodiments of the present invention will be described in more detail below with reference to the accompanying drawings. Although the exemplary embodiments of the present invention are shown in the accompanying drawings, it should be understood that the present invention can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided in order to enable a more thorough understanding of the present invention and to enable the scope of the present invention to be fully communicated to those skilled in the art.

[0068] In a first aspect, an embodiment of the present invention provides a far-field safety analysis method after satellite-rocket separation. The method is applicable to the far-field safety analysis of one or more satellites of the same batch after satellite-rocket separation.

[0069] See also Figure 1 , the method comprises at least the following steps:

[0070] S101, after the separation of satellites and rockets in the same batch, estimate the relative distance between the satellite and the launch vehicle as it changes over time.

[0071] S102, calculating the angle between the satellite guidance direction and the thrust direction of the carrier rocket, wherein the satellite guidance direction is the direction in which the carrier rocket points to the satellite.

[0072] S103, according to the preset far-field safety judgment conditions corresponding to the N stages after the separation of the same batch of satellites and rockets, one or more of the relative distance between the satellites and rockets and the angle (the impact of engine plume on safety) are analyzed for far-field safety. N ≥ 3 and is a positive integer

[0073] Among them, when there is only one satellite, it is only necessary to calculate the two parameters of satellite-rocket relative distance and angle according to steps S101 to S103 to perform far-field safety analysis. When there are more than two satellites, before S101, estimate the relative distance of satellites in the same batch that changes over time. On this basis, according to the preset judgment conditions corresponding to the N stages, perform far-field safety analysis on one or more of the satellite relative distance, satellite-rocket relative distance, and angle.

[0074] Assuming that the total number of satellites in the same batch is M, the following is the implementation logic of estimating the relative distance between two satellites in the same batch that changes over time for the scenario with more than two satellites. If there is only one satellite, the far-field safety analysis between satellites is not considered.

[0075] As an optional embodiment, the implementation logic of estimating the relative distances between two satellites in the same batch that change over time specifically includes the following steps:

[0076] S201, at the time t0 of separation of the rocket and the satellite, convert the position velocity of satellite i and satellite j in the launch system into the position velocity of satellite i and satellite j in the geocentric system.

[0077] Among them, time t0 represents the moment of satellite-rocket separation; satellite i and satellite j are any two satellites from the same batch, i∈N, j∈N, and i≠j.

[0078] The launch system is a coordinate system centered on the satellite and established according to the specific design or functional requirements of the satellite. The launch system is a non-inertial system and is affected by multiple forces such as the earth's gravity and atmospheric resistance.

[0079] The geocentric system is a right-handed coordinate system with the center of mass of the earth as the origin, the axis of rotation of the earth as the Z axis (usually pointing to the North Pole), and the X and Y axes in the equatorial plane and perpendicular to each other. The geocentric system is an inertial system that does not consider the influence of the rotation and revolution of the earth on the coordinate system. It is used to describe the absolute position and speed of objects around the earth (such as satellites, rockets, etc.).

[0080] During the conversion process, the transformation matrix of the launch system relative to the geocentric system is first determined, and then the transformation matrix is ​​used to convert the position and velocity of satellite i and satellite j in the launch system into the position and velocity of satellite i and satellite j in the geocentric system.

[0081] S202, according to the position and velocity of satellite i and satellite j in the geocentric system, find the instantaneous root of satellite i and satellite j at time t0:

[0082] The instantaneous root is the orbital root of the satellite at a certain moment (for example, at t0). The parameters included in the instantaneous root are: semi-major axis a m , eccentricity e m , orbital inclination i m , right ascension of ascending node Ω m 、Arg of perigee ω m 、Mean anomaly angle M m It is worth noting that the calculation method for the mutual conversion between position velocity, instantaneous root and flat root is a conventional conversion method, which will not be repeated here or in the subsequent parts. In addition, the mutual conversion between position velocity, instantaneous root and flat root implemented by satellite i and satellite j is independent of each other.

[0083] S203, the instantaneous root of satellite i and satellite j at time t0: Converted to the square root of satellite i and satellite j at time t0:

[0084] S204, based on the square root of satellite i and satellite j at time t0: Estimate satellite i and satellite j at t i The flat root of the moment:

[0085] in, Δt represents ti The time difference between time t0 and time t1 is Δt = t2. i -t0. Therefore, satellite i and satellite j are at t i Pinggen of the moment The parameters included can be replaced by Δt.

[0086] In this embodiment, the J2 perturbation theory can be used to calculate the satellite i and satellite j at t i Pinggen of the moment Make an estimate.

[0087] In the estimated satellite i at t i Pinggen of the moment In the process, according to the square root of satellite i at time t0: Determine the following parameters of satellite i: the rate of change of right ascension of ascending node W Ω , the rate of change of the perigee argument W ω , the rate of change of the mean perigee angle W M :

[0088]

[0089]

[0090]

[0091] Where ω0 represents the angular velocity of the earth's rotation; J2 represents the second-order term in the earth's shape dynamics coefficient, which is a constant; A e Represents the semi-major axis of the Earth's rotation ellipsoid, which is a constant.

[0092] After obtaining the right ascension change rate W of the ascending node of satellite i Ω , the rate of change of the perigee argument W ω , the rate of change of the mean perigee angle W M Afterwards, combined with t i The time difference Δt between time t and time t0 determines the time difference between satellite i and time t i Pinggen of the moment The parameters included in:

[0093] a m (Δt) = a m

[0094] e m (Δt) = e m

[0095] i m (Δt)=i m

[0096] Ωm (Δt)=Ω m +W Ω Δt

[0097] ω m (Δt)=ω m +W ω Δt

[0098] M m (Δt) = M m +W M Δt

[0099] Among them, W Ω Δt represents the change in right ascension of the ascending node, W ω Δt represents the change in the argument of perigee, W M Δt represents the change in the mean perigee angle.

[0100] According to the above technical solution, the satellite i at t i Pinggen of the moment It is worth noting that satellite j is i Pinggen of the moment The estimation method and satellite i at t i Pinggen of the moment The estimation method is the same as that of , so I will not repeat it here.

[0101] S205, satellite i and satellite j are placed at t i The flat root of the moment: Convert to satellite i and satellite j at t i The root of the moment: The conversion method is conventional and will not be described in detail here.

[0102] S206, according to satellite i and satellite j at t i The instant root of time: Determine the satellite i and satellite j at t i The position at the moment.

[0103] S207, according to satellite i and satellite j at t i The position at the moment, determine the relative distance S between satellite i and satellite j over time ij .

[0104] The above is the estimated relative distance S of satellite i and satellite j over time. ij The specific implementation process of, in practical applications, the relative distance between any two satellites can be estimated according to the above method.

[0105] As an optional implementation, the implementation logic of estimating the satellite-rocket relative distance between the satellite and the launch vehicle that changes over time in S101 is described below, which specifically includes the following steps:

[0106] S301, integrate the Earth's gravity, engine thrust, aerodynamic force and the mass of the launch vehicle to estimate the launch vehicle t i The velocity of the launch system at the moment, and based on the launch vehicle t i The velocity integral of the launch system at the time is obtained as i The position under the launch system at all times.

[0107] In this embodiment, since the launch vehicle is divided into N stages after the separation of the satellite and the rocket, there is an active stage between the start-up and shutdown of the launch vehicle in the N stages, the relative distance between the satellite and the rocket cannot be estimated by referring to the J2 perturbation theory. Based on this, this embodiment estimates the launch vehicle t i The position under the launch system at all times.

[0108] Specifically, referring to the earth's gravity, engine thrust, aerodynamic force and the mass of the launch vehicle, the velocity formula is integrated and solved to estimate the launch vehicle t i The velocity of the launch system at the moment. Further, according to the position formula, further integration is performed to obtain the launch vehicle t i The position under the launch system at all times.

[0109] In this embodiment, the speed formula is:

[0110]

[0111] Where v represents the launch vehicle t i The velocity in the launch frame at that moment, v x 、v y 、v z Indicates launch vehicle t i The velocity components in the launch system at each moment, represents the first-order derivative of each velocity component, G represents the earth's gravity, G x , G y , G z represents the components of the earth's gravity in the launch system, F represents the engine thrust, and F x 、F y 、F z represents the components of the engine thrust in the launch system, Q represents the aerodynamic force, Q x , Q y , Q z represents the components of aerodynamic force in the launch system, and m represents the mass of the launch vehicle.

[0112] In this embodiment, the position formula is:

[0113]

[0114] Where R represents the launch vehicle t i The position in the launch system at the moment, R x , R y , R z Indicates launch vehicle t i The position components of the emission system at each moment, Represents the first-order derivative of each position component.

[0115] S302, will launch the rocket t i The position of the launch system at the moment, converting the launch vehicle t i The position under the geocentric system at all times.

[0116] In the conversion process, the conversion matrix of the launch system relative to the geocentric system is first determined, and then the conversion matrix is ​​used to convert the launch vehicle t i The position of the launch system at the moment is converted into the launch vehicle t i The position under the geocentric system at all times.

[0117] S303, according to the satellite i at t i The position of the launch vehicle at time t i The position at the moment determines the relative distance S between satellite i and the launch vehicle that changes with time. i .

[0118] Wherein, satellite i is any satellite in the same batch. Any satellite at t i The position at any moment can be obtained through the technical solutions of the aforementioned steps: S201 to S206.

[0119] By using the technical solutions of S301 to S303, the relative distance between any satellite and the launch vehicle over time can be calculated. It is worth noting that if there are more than two satellites, it is necessary to calculate the relative distance between each satellite and the launch vehicle over time and perform far-field safety.

[0120] As an optional implementation, the implementation logic of calculating the angle between the satellite guidance direction and the launch vehicle thrust direction in S102 is described below, which specifically includes the following steps:

[0121] S401, according to satellite i at t i The position of the launch vehicle at time t i The position at the moment, find the vector of the satellite guidance direction in the geocentric system The satellite guidance direction is the direction of the line between the satellite and the launch vehicle. Different satellites have different corresponding satellite guidance directions. i The position of the launch vehicle at time t i The position at the moment can be used to obtain the vector of the satellite guidance direction in the geocentric system.

[0122] S402, the vector that guides the satellite direction to the geocentric system Converted into the vector of the satellite guidance direction in the arrow system The arrow system is a coordinate system built with the arrow body (such as a rocket or missile) as the center. When converting, first determine the transformation matrix of the arrow system relative to the geocentric system, then multiply the vector of the arrow body in the geocentric system by the transformation matrix to obtain the representation of the vector in the arrow system.

[0123] S403, determine the vector of the launch vehicle thrust in the arrow system Among them, the vector of the launch vehicle thrust is Generally, it is along the axis of the rocket body and can be represented by a simple unit vector. For rockets using vector thrust, the corresponding vector needs to be calculated based on the angle of thrust deflection.

[0124] S404, according to the formula Determine the angle; where α i It represents the angle between the satellite guidance direction formed by satellite i and the launch vehicle and the thrust direction of the launch vehicle, scos() represents the inverse cosine formula, and || represents the modulus.

[0125] According to the technical solutions of S401 to S404, the angle between the satellite guidance direction formed by each satellite and the carrier rocket and the thrust direction of the carrier rocket can be determined. Different satellites have different corresponding angles.

[0126] As an optional embodiment, the same batch of satellites and rockets can be divided into N stages after separation. Exemplarily, the same batch of satellites and rockets can be divided into three stages after separation, namely: the glide stage from the separation of satellites and rockets to the start-up of the carrier rocket, the active stage from the start-up to the shutdown of the carrier rocket, and the unpowered glide stage after the shutdown of the carrier rocket. Different stages correspond to different preset far-field safety judgment conditions, and the preset far-field safety judgment conditions can be adaptively changed according to the number of satellites.

[0127] During the glide phase from the separation of the satellite and rocket to the start-up of the launch vehicle, the corresponding preset far-field safety judgment conditions are:

[0128] Relative distance between star and arrow S i>L1, this condition is applied to the safety analysis scenario with only one satellite; or,

[0129] Relative distance between star and arrow S i >L1 and the relative distance between satellites is S ij >L1, this condition is applied to the safety analysis scenario of more than two satellites; where L1 represents the first preset safety distance. It is worth noting that if there are more than two satellites, this condition needs to be used to determine the relative distance between the rocket and the satellite and the relative distance between the satellites. When all satellites meet the above conditions, it can be determined that the long-range safety requirements are met.

[0130] During the active phase between the launch vehicle startup and shutdown, the corresponding preset far-field safety judgment condition is one of the following:

[0131]

[0132] Among them, conditions (1) and (2) are applied to the safety analysis scenario of more than two satellites. Conditions (3) and (4) are applied to the safety analysis scenario of only one satellite. Meeting any of the above conditions can meet the far-field safety requirements in the active segment between startup and shutdown of the launch vehicle. It is worth noting that if there are more than two satellites, one of the above conditions (1) and (2) is required to determine the relative distance between the rocket and the satellite, the relative distance between the satellites and the angle. When all satellites meet the requirements, it can be determined that the remote safety requirements are met.

[0133] Wherein, L2 represents the second preset safety distance, L3 represents the third preset safety distance, L4 represents the fourth preset safety distance, α safe The cone angle in the plume influence range; where the plume influence range has an arc length of L2 and a half angle of α safe of cone;

[0134] During the unpowered glide phase after the launch vehicle is shut down, the corresponding preset far-field safety judgment conditions are:

[0135] S i >L5, this condition is applied to the safety analysis scenario with only one satellite; or

[0136] S i >L5, S ij >L6, this condition is applied to the safety analysis scenario of more than two satellites; among them, L5 represents the fifth preset safety distance; L6 represents the sixth preset safety distance. It is worth noting that if there are more than two satellites, this condition needs to be used to determine the relative distance between the rocket and the satellite and the relative distance between the satellites. When all satellites meet the above conditions, it can be determined that the long-range safety requirements are met.

[0137] In the process of safety analysis, ti The time change Δt between time t and time t0 is the time unit. Each time unit is combined with the preset far-field safety judgment conditions corresponding to the stage to conduct a far-field safety analysis on one or more of the relative distances and angles between the satellite and the rocket.

[0138] In the specific implementation process, for each stage, t i The time change Δt between time t and time t0 is a time unit, and a far-field safety analysis is performed for each time unit in combination with the preset far-field safety judgment condition corresponding to the stage until all stages are passed.

[0139] As an optional embodiment, during the unpowered gliding phase after the launch vehicle is shut down, the relative distance S between the rocket and the satellite can be used. i and / or satellite relative distance S ij The trend of changes in the far-field safety analysis is carried out; see Figure 2 This is a schematic diagram of the changing trend of the relative distance between the star and the rocket. Figure 3 , is a schematic diagram of the relative distance between satellites. i The trend of change is monotonically increasing (such as Figure 2 ), and / or the relative distance S of the satellites ij The trend of change is periodic increase (such as Figure 3 ), it means that the safety requirements are met.

[0140] In the second aspect, based on the same inventive concept as the far-field safety analysis method after satellite-rocket separation provided in the embodiment of the first aspect, the embodiment of the present invention further provides a far-field safety analysis system after satellite-rocket separation, see Figure 4 , the system comprising:

[0141] An estimation module 401 is used to estimate the relative distance between the satellite and the launch vehicle over time after the satellite and the rocket of the same batch are separated;

[0142] A calculation module 402 is used to calculate the angle between the satellite guidance direction and the thrust direction of the carrier rocket; wherein the satellite guidance direction is the direction in which the carrier rocket points to the satellite;

[0143] The analysis module 403 is used to perform far-field safety analysis on one or more of the satellite-rocket relative distance and the angle according to the preset far-field safety judgment conditions corresponding to the N stages after the separation of the same batch of satellites and rockets, where N≥3 and is a positive integer.

[0144] It should be noted that the far-field safety analysis system after satellite-rocket separation provided in the embodiment of the present invention, wherein the specific manner in which each module performs operations has been described in detail in the method embodiment provided in the above-mentioned first aspect. The specific implementation process can refer to the method embodiment provided in the above-mentioned first aspect, and will not be elaborated here.

[0145] Although the preferred embodiments of the present application have been described, those skilled in the art may make other changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications that fall within the scope of the present application.

[0146] Obviously, those skilled in the art can make various changes and modifications to the present application without departing from the spirit and scope of the present application. Thus, if these modifications and variations of the present application fall within the scope of the claims of the present application and their equivalents, the present application is also intended to include these modifications and variations.

Claims

1. A far-field safety analysis method after satellite-rocket separation, characterized in that: The method comprises: After the satellite-rocket separation of the same batch, estimate the relative distance between the satellite and the launch vehicle over time; Calculating the angle between the satellite guidance direction and the thrust direction of the carrier rocket; wherein the satellite guidance direction is the direction in which the carrier rocket points to the satellite; According to the preset far-field safety judgment conditions corresponding to the N stages after the separation of the same batch of satellites and rockets, a far-field safety analysis is performed on one or more of the satellite-rocket relative distance and the angle, where N≥3 and is a positive integer.

2. The method according to claim 1, characterized in that If there are more than two satellites in the same batch; before estimating the relative distance between the satellite and the launch vehicle over time, specifically including: Estimate the relative distances between two satellites in the same batch over time; The performing of a far-field safety analysis on one or more of the satellite-rocket relative distance and the included angle according to the preset judgment conditions corresponding to the N stages specifically includes: According to the preset judgment conditions corresponding to the N stages, a far-field safety analysis is performed on one or more of the satellite relative distance, the satellite-rocket relative distance, and the angle.

3. The method according to claim 2, characterized in that The estimating of the relative distances between two satellites in the same batch varying over time specifically includes: At the time t0 of satellite-rocket separation, the position velocities of satellites i and j in the launch system are converted into the position velocities of satellites i and j in the geocentric system; wherein time t0 is the time of satellite-rocket separation; satellites i and j are any two satellites from the same batch; According to the position speed of the satellite i and the satellite j in the geocentric system, the instantaneous root of the satellite i and the satellite j at time t0 is calculated: Among them, the parameters of the instantaneous root include: semi-major axis a m , eccentricity e m , orbital inclination i m , right ascension of ascending node Ω m 、Arg of perigee ω m 、Mean anomaly angle M m ; The instantaneous root of the satellite i and the satellite j at time t0 is: Converted to the square root of the satellite i and the satellite j at time t0: According to the square root of the satellite i and the satellite j at time t0: Estimate the satellite i and the satellite j at t i The flat root of the moment: in, Δt represents t i The time difference between time t0 and time t1 is Δt = t2. i -t0; The satellite i and the satellite j are placed at t i The flat root of the moment: Convert to the satellite i and the satellite j at t i The instant root of time: According to the satellite i and the satellite j at t i The instant root of time: Determine the satellite i and the satellite j at t i The location at the moment; According to the satellite i and the satellite j at t i The relative distance S between the satellite i and the satellite j changes with time. ij .

4. The method according to claim 3, characterized in that The square root of the satellite i and the satellite j at time t0 is: Estimate the satellite i and the satellite j at t i The flat root of the moment: Specifically include: According to the square root of the satellite i at time t0: Determine the following parameters of the satellite i: the rate of change of the right ascension of the ascending node W Ω , the rate of change of the perigee argument W ω , the rate of change of the mean perigee angle W M : Where ω0 represents the angular velocity of the earth's rotation; J2 represents the second-order term in the earth's shape dynamics coefficient, which is a constant; A e represents the semi-major axis of the Earth's rotation ellipsoid, which is a constant; Combined with i The time difference Δt between time t0 and time t1 determines the time difference between the satellite i and time t2. i Pinggen of the moment The parameters included in: Among them, W Ω Δt represents the change in the right ascension of the ascending node, W ω Δt represents the change in the perigee argument, W M Δt represents the change in the mean perigee angle; The satellite j is at t i Pinggen of the moment The estimation method and the satellite i at t i Pinggen of the moment The estimation method is the same.

5. The method according to claim 3, characterized in that The estimating the relative distance between the satellite and the launch vehicle over time specifically includes: By integrating the earth's gravity, engine thrust, aerodynamic force and the mass of the launch vehicle, the launch vehicle t is estimated. i The velocity of the launch system at the moment, and based on the launch vehicle t i The velocity integral of the launch system at the time is obtained as i The position under the launch system at all times; The launch vehicle t i The position of the launch system at the moment, converting the launch vehicle t i The position of the earth at all times; According to the satellite i at t i The position of the launch vehicle at time t i The position at the moment, determines the relative distance S between the satellite i and the launch vehicle that changes with time i ; Wherein, the satellite i is any satellite from the same batch.

6. The method according to claim 5, characterized in that The calculating the angle between the satellite guidance direction and the thrust direction of the carrier rocket specifically includes: According to the satellite i at t i The position of the launch vehicle at time t i The position at the moment, find the vector of the satellite guidance direction in the geocentric system The vector of the satellite guidance direction in the geocentric system Converted into the vector of the satellite guidance direction in the arrow system Determine the vector of the launch vehicle thrust in the arrow system According to the formula Determine the angle; where α i It represents the angle between the satellite guidance direction formed by satellite i and the launch vehicle and the thrust direction of the launch vehicle, acos() represents the inverse cosine formula, and || represents the modulus.

7. The method according to any one of claims 1 to 6, characterized in that: The N stages include: a glide stage from the separation of the satellite and the rocket to the start-up of the carrier rocket, an active stage from the start-up to the shutdown of the carrier rocket, and an unpowered glide stage after the shutdown of the carrier rocket; In the glide section after the separation of the satellite and the rocket and before the launch of the carrier rocket, the corresponding preset far-field safety judgment condition is: the relative distance S between the satellite and the rocket i >L1; or: the relative distance between the star and the arrow S i >L1 and the relative distance between satellites is S ij >L1; wherein L1 represents the first preset safety distance; In the active section between the start-up and shutdown of the launch vehicle, the corresponding preset far-field safety judgment condition is one of the following: Wherein, L2 represents the second preset safety distance, L3 represents the third preset safety distance, L4 represents the fourth preset safety distance, α safe The cone angle in the plume influence range; wherein the plume influence range has an arc length of L2 and a half angle of α safe of cone; During the unpowered gliding phase after the launch vehicle is shut down, the corresponding preset far-field safety judgment condition is: S i >L5; or: S i >L5, S ij >L6; wherein, L5 represents the fifth preset safety distance; L6 represents the sixth preset safety distance.

8. The method according to claim 7, characterized in that The far-field safety analysis is performed on one or more of the relative distance between the satellite and the rocket and the angle according to the preset far-field safety judgment conditions corresponding to the N stages after the separation of the satellite and the rocket from the same batch, specifically including: With t i The time change Δt between time t and time t0 is a time unit. Each time unit is combined with the preset far-field safety judgment conditions corresponding to the stage to perform a far-field safety analysis on one or more of the satellite-rocket relative distance and the angle.

9. The method according to claim 7, characterized in that The method further comprises: In the unpowered gliding section after the launch vehicle is shut down, the relative distance S between the rocket and the satellite is used. i and / or the relative distance S of the satellites ij Conduct far-field safety analysis on the changing trend of If the relative distance between the star and the arrow is S i is monotonically increasing, and / or the satellite relative distance S ij If the changing trend of is periodically increasing, it means that the safety requirements are met.

10. A far-field safety analysis system after satellite-rocket separation, characterized in that: The system comprises: An estimation module, used for estimating the relative distance between the satellite and the launch vehicle over time after the satellite and the rocket of the same batch are separated; A calculation module, used to calculate the angle between the satellite guidance direction and the thrust direction of the launch vehicle; Wherein, the satellite guidance direction is the direction in which the carrier rocket points to the satellite; The analysis module is used to perform far-field safety analysis on one or more of the relative distance between the satellite and the rocket and the angle according to the preset far-field safety judgment conditions corresponding to the N stages after the separation of the satellite and the rocket from the same batch, where N≥3 and is a positive integer.

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