CN115688390A - Carrier rocket satellite-rocket separation time sequence task-removing design method - Google Patents
Carrier rocket satellite-rocket separation time sequence task-removing design method Download PDFInfo
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Abstract
The invention relates to a carrier rocket satellite-rocket separation time sequence task-removing design method. According to the method, the rocket separation time sequence is automatically designed according to the satellite and rocket separation mode, the position and speed of the rocket at the moment before the first satellite and rocket separation, the separation speed, the satellite quality and the total rocket quality, the installation position of the satellite in the rocket and the number of the satellites, the STK software is prevented from being used for calling trajectory parameters such as the rocket in-orbit UTC time, six orbital parameters, the rocket navigation coordinate system attitude angle during the satellite and rocket separation for far-field safety analysis, the calculation amount and the manual design process can be effectively reduced, and the working efficiency is improved. Meanwhile, the method can design the time sequence of the satellite and arrow separation section for different launching tasks, thereby realizing the task-free design of the satellite and arrow separation scheme.
Description
Technical Field
The invention relates to the technical field of rocket trajectories, in particular to a carrier rocket satellite-rocket separation time sequence task-removing design method.
Background
And releasing the satellite after the carrier rocket reaches the preset orbit, and implementing satellite-rocket separation. Because the relative speed of the rocket and the satellite is low after separation, the rocket needs to slide for a period of time according to the current speed, and when the distance between the satellite and the rocket and the distance between the satellites (the separation condition of a plurality of satellites needs to be considered) reach a safe distance, subsequent actions can be carried out. Therefore, the far-field safety between the rocket and the satellite and between the satellite needs to be analyzed so as to design a satellite-rocket separation time sequence after the rocket enters the orbit.
For the far-field safety of the satellite and the rocket, STK software is generally needed to be used for analyzing, the distance between the rocket and the satellite and the distance between the satellites are output by reading trajectory parameters such as the rocket entry UTC time, six orbits and attitude angles of a rocket navigation coordinate system when the satellite and the rocket are separated, and a corresponding satellite and rocket separation scheme is given through manual judgment and repeated iteration, so that the far-field safety requirements between the rocket and the satellite and between the satellite and the satellite can be met. But the method has the following defects: for different launching tasks, due to the fact that the number of satellites and the satellite-arrow separation mode and the like are changed, a designer needs to perform judgment by combining with STK software and iterate repeatedly to give a corresponding satellite-arrow separation scheme, time and labor are consumed, and the design period is long.
Disclosure of Invention
The invention provides a carrier rocket satellite-rocket separation timing sequence task-removing design method, which can automatically design a satellite-rocket separation timing sequence according to parameters such as a satellite-rocket separation mode, the position and speed of a rocket immediately before the separation of the satellite and the rocket for the first time, separation speed, the mass of a satellite and the total mass of the rocket, the installation position of the satellite in the rocket, the number of satellites and the like, thereby meeting the satellite-rocket separation task-removing requirement.
A carrier rocket satellite-rocket separation time sequence task-removing design method comprises the following steps:
setting three attitude angles A, B, C of arrow bodies in the established orbit coordinate system, and determining an attitude angle A, B, C at the satellite-arrow separation time according to a satellite-arrow separation mode;
converting the attitude angle A, B, C in the orbit coordinate system into a navigation coordinate system, and calculating the attitude angle from the arrow coordinate system to the navigation coordinate system according to the component of A, B, C in the navigation coordinate system;
calculating the speeds of the satellite and the rocket in the navigation coordinate system at the satellite and rocket separation moment according to the attitude angle from the rocket body coordinate system to the navigation coordinate system when each satellite is separated, the satellite separation speed, the speed of the rocket at the moment before each satellite and rocket separation, the mass of the satellite and the total mass of the rocket, and the installation position of the satellite in the rocket;
and designing a satellite-rocket separation time sequence according to the number of satellites, the rocket at each satellite-rocket separation time and the position and speed of the satellites.
Further, the orbital coordinate system is defined as: the origin is the rocket center of mass, and the X axis points to the velocity vector direction of the rocket during separation; the Y axis is positioned in the instantaneous track plane, points to the outer side of the track and is vertical to the X axis; the Z axis is vertical to the X axis and the Y axis and meets the right hand rule.
Further, the attitude angle A, B, C is set to:
the angle A is an included angle between the projection of the X axis of the arrow body in the track plane and the X axis of the track coordinate system;
the angle B is an included angle of the X axis of the arrow body deviating from the plane of the track;
the angle C is the rotation angle of the arrow body around the X axis, and the Y axis of the arrow body is in the track plane when C = 0.
4. The launch vehicle satellite-rocket separation timing sequence tasking design method of claim 1, wherein said converting attitude angle A, B, C in an orbital coordinate system to a navigational coordinate system comprises:
obtaining a conversion matrix of the rocket body coordinate system and the orbit coordinate system by taking A, B, C as an Euler angle, and enabling a coordinate base of a satellite in the rocket body coordinate system
Converting the conversion matrix into a track coordinate system through an arrow coordinate system and the track coordinate system; substrate
The components in the orbital coordinate system are noted as:
will be provided with
From orbital to geocentric coordinate systems, base
The components in the geocentric coordinate system are noted as:
will be provided with
From the geocentric coordinate system to the emission coordinate system, the base
The components in the emission coordinate system are noted as:
will be provided with
From the emission coordinate system to the navigation coordinate system, base
The components in the navigation coordinate system are noted as:
according to the substrate
And calculating the components in the navigation coordinate system to obtain the attitude angle from the rocket coordinate system to the navigation coordinate system:
γ＝γ 0
wherein, γ 0 The roll angle of the navigation coordinate system at the satellite-rocket separation time is determined in advance by the measurement and control requirements of the rocket and the satellite-rocket separation mode.
Each coordinate system is defined as:
geocentric coordinate system: the origin of coordinates is at the earth's center Oe, oeXe points in the equatorial plane to the original meridian, the OeZe axis is perpendicular to the equatorial plane and coincides with the earth's rotation axis, oeYe is available from the right hand rule.
Emission coordinate system: the origin of coordinates is located at the emission origin, the OY axis is a plumb line passing through the emission point, the direction is positive, the OX axis is perpendicular to the OY axis and points to the theoretical direction, and the OZ axis, the OX axis and the OY axis form a right-hand rectangular coordinate system;
arrow coordinate system: the origin of coordinates is located in the center of mass of the rocket, the axis OX1 is consistent with the longitudinal symmetry axis of the rocket body and points to the direction of the head, the axis OY1 is perpendicular to the axis OX1 and located in the longitudinal symmetry plane of the rocket and points to the upper part, and the axis OZ1, the axis OX1 and the axis OY1 form a right-hand rectangular coordinate system;
a navigation coordinate system: the navigation coordinate system is coincided with the launching coordinate system at the rocket launching moment, after the rocket is ignited, the position of the origin of coordinates moves at the linking speed of the launching point at the launching moment, and the directions of the coordinate axes OXd, OYd and OZD are kept unchanged.
Further, the calculating the speed of the satellite and the rocket in the instantaneous navigation coordinate system after the separation of the satellite and the rocket according to the attitude angle of the navigation coordinate system when each satellite is separated, the satellite separation speed, the speed of the instantaneous rocket before the separation of the satellite and the rocket for the first time, the mass of the satellite and the total mass of the rocket and the installation position of the satellite in the rocket comprises the following steps:
to obtain
Psi and gamma are Euler angles, a coordinate base sat _ HJ { satx, party, satz } of the satellite installation position in the rocket is converted into a navigation coordinate system through a conversion matrix of the rocket body coordinate system and the navigation coordinate system, and the component of the base of the satellite installation position in the navigation coordinate system is recorded as: sat1_ HJ { satx1, saty1, satz1};
according to the mass of the satellite, the total mass m _ sat and m _ HJ of the rocket and the separation speed SEP _ V of the satellite, calculating the separation speeds delta V _ sat and delta V _ HJ obtained by the rocket and the satellite at the moment after the satellite and the rocket are separated by adopting a momentum theorem;
the velocities of the satellite and rocket at the moment after the separation of the satellite and the rocket (SATVX 1, SATVY1, SATVZ 1) and HJV1 (HJVX 1, HJVY1 and HJVZ 1) are calculated from sat1_ HJ { satx1, saty1, satz1}, delta v _ sat, delta v _ HJ and the velocity of the rocket at the moment before the separation of the satellite and the rocket for the first time (HJVX, HJVY and HJVZ 1).
Further, the designing of the satellite-rocket separation time sequence according to the number of satellites, the rocket at the first satellite-rocket separation time, and the positions and the speeds of the satellites comprises:
performing numerical integration or orbit prediction on the position and the speed of the instantaneous rocket and each separated satellite after each satellite and rocket separation; when the distance between the rocket and the separated satellite is greater than the safe distance L _ HJTOWX and the distance between the separated satellites is greater than the safe distance L _ WXTOWX, the numerical integration or forecast time length is the interval time length Delta T between the separation of the satellite and the rocket at the time and the separation of the satellite and the satellite at the next time; and repeating the steps until the last satellite is separated.
Further, the satellite and rocket separation time sequence is designed according to the number of satellites, the rocket at the first satellite and rocket separation time, and the position and the speed of the satellites, and the method comprises the following steps: after the last satellite and the satellite are separated, designing the interval time delta TLG from the separation of the satellite and the satellite to the off-orbit passivation, and updating the rocket mass to m _ HJ-m _ satall, wherein m _ satall is the total mass of all satellites;
and performing numerical integration or orbit prediction on the position and the speed of the instant rocket and each satellite after the satellite and the rocket are separated, calculating the distance between the rocket and each satellite and the distance between all satellites in real time, and 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, wherein the integration or prediction time length is the interval time length delta TLG from the separation of the satellite and the rocket to the off-orbit passivation.
Compared with the prior art, the invention has the beneficial effects that:
according to the method, the rocket separation time sequence is automatically designed according to the satellite-rocket separation mode, the position and speed of the rocket at the moment before the first satellite-rocket separation, the separation speed, the satellite mass and the total rocket mass, the installation position of the satellite in the rocket and the number of the satellites, and the far field safety analysis by calling ballistic parameters such as the rocket in-orbit UTC time, six orbital parameters, the attitude angle of a rocket navigation coordinate system during satellite-rocket separation and the like by using STK software can be avoided, the calculation amount and the manual design process can be effectively reduced, and the working efficiency is improved. Meanwhile, the method can design the time sequence of the satellite and rocket separation section for different launching tasks, thereby realizing the off-duty design of the satellite and rocket separation scheme.
Drawings
FIG. 1 is a flow chart of a carrier rocket satellite-rocket separation timing sequence task-based design.
Detailed Description
A carrier rocket satellite-rocket separation time sequence task-removing design method is shown in a figure 1, and specifically comprises the following steps:
the method comprises the following steps: and calculating the attitude angle of the rocket navigation coordinate system during the separation of the satellite and the rocket each time according to the satellite and rocket separation mode and the position of the rocket at the moment before the separation of the satellite and the rocket.
Due to the restriction of factors such as the requirement of a satellite party, the design of a satellite-rocket adapter, space-based measurement and control and the like, the attitude angle design of satellite-rocket separation of each task may be different, and finally, the separation attitude angle in the navigation coordinate system needs to be provided. Since it is difficult to design the attitude angle directly in the navigation coordinate system according to the constraint condition, a coordinate system for visually describing the attitude of the arrow body during separation is first required, which is called an "orbital coordinate system":
establishing a track coordinate system: the origin is the rocket center of mass, and the X axis points to the velocity vector direction of the rocket during separation; the Y axis is positioned in the instantaneous track plane, points to the outer side of the track and is vertical to the X axis; the Z axis is vertical to the X axis and the Y axis and meets the right hand rule.
Three attitude angles A, B, C are defined within the orbital coordinate system. The three angles are set as:
the angle A is an included angle between the projection of the X axis of the arrow body in the track plane and the X axis of the track coordinate system;
the angle B is an included angle of the X axis of the arrow body deviating from the plane of the track;
the angle C is the rotation angle of the arrow body around the X axis, and the Y axis of the arrow body is in the track plane when C = 0.
Three angles A, B, C are given according to the satellite and arrow separation mode, and the satellite and arrow separation mode in the embodiment is as follows: after the rocket is in orbit, the rocket rotates 90 degrees around the Z axis of the orbit coordinate system, namely the rocket is raised by 90 degrees. At this time, A, B, C is 90 °,0 °, respectively. After determining the attitude angles of the three orbital coordinate systems, converting the attitude angles into a navigation coordinate system, and comprising the following steps:
obtaining a conversion matrix of the rocket coordinate system and the orbit coordinate system by taking A, B, C as an Euler angle, and enabling a coordinate base of a satellite in the rocket coordinate system
And converting the coordinate system of the rocket body into the track coordinate system through a conversion matrix of the coordinate system of the rocket body and the track coordinate system. Substrate
The components in the orbital coordinate system are denoted as:
substrate
Depending on the installation location of the satellite in the rocket;
the conversion matrix of the rocket body coordinate system and the track coordinate system is as follows:
will be provided with
Converting the orbit coordinate system into a geocentric coordinate system (the transformation matrix from the orbit coordinate system to the geocentric coordinate system needs to use the speed and the position of the rocket at the satellite-rocket separation time, namely the position (X _ HJ, Y _ HJ, Z _ HJ) and the speed (Vx _ HJ, vy _ HJ, vz _ HJ)) of the rocket at the moment before the satellite-rocket separation at the moment, and inputting the position (X _ HJ, Y _ HJ, Z _ HJ) and the speed (Vx _ HJ, vy _ HJ, vz _ HJ)) of the rocket at the moment before the satellite-rocket separation at the moment into a base
In the geocentric systemThe component of (c) is noted as:
obtaining six orbit numbers (an orbit inclination angle i, a rising intersection point right ascension omega and a near place argument w) of the geocentric system at the moment by the position and the speed of the geocentric coordinate system at the moment of satellite and rocket separation, and obtaining a conversion matrix from the orbit coordinate system to the geocentric coordinate system by the six orbit numbers:
will be provided with
Converting the geocentric coordinate system into an emission coordinate system, and recording the components of the substrate in the emission coordinate system as:
the transformation matrix of the geocentric coordinate system and the emission coordinate system is as follows:
in the formula:
A 0 -theoretical direction;
B 0 -the geodetic latitude of the launch site;
L 0 -the geodetic longitude of the emission point.
Will be provided with
Converting the transmitting coordinate system into a navigation coordinate system, and recording the components of the substrate in the navigation coordinate system as:
the transformation matrix of the emission coordinate system and the navigation coordinate system is as follows:
M FS2DH ＝A -1 BA
in the formula:
A 0 -theoretical firing;
B 0 -the geodetic latitude of the launch site;
ω e -angular velocity of rotation of the earth;
t-the time elapsed since the fire was emitted to the current moment.
The coordinate systems in the above conversion process are defined as:
geocentric coordinate system: the origin of coordinates is at the earth's center Oe, oeXe points in the equatorial plane to the original meridian, the OeZe axis is perpendicular to the equatorial plane and coincides with the earth's rotation axis, oeYe is available from the right hand rule.
Emission coordinate system: the origin of coordinates is located at the emission origin, the OY axis is a plumb line passing through the emission point, the direction is positive, the OX axis is perpendicular to the OY axis and points to the theoretical direction, and the OZ axis, the OX axis and the OY axis form a right-hand rectangular coordinate system;
arrow coordinate system: the origin of coordinates is located at the center of mass of the rocket, the axis OX1 is consistent with the longitudinal symmetry axis of the rocket body and points to the head direction, the axis OY1 is perpendicular to the axis OX1, is located in the longitudinal symmetry plane of the rocket and points to the upper direction, and the axis OZ1, the axis OX1 and the axis OY1 form a right-hand rectangular coordinate system;
navigation coordinate system (navigation system for short in fig. 1): the navigation coordinate system is coincided with the launching coordinate system at the rocket launching moment, after the rocket is ignited, the position of the origin of coordinates moves at the linking speed of the launching point at the launching moment, and the directions of the coordinate axes OXd, OYd and OZD are kept unchanged.
And calculating the attitude angle of the navigation coordinate system according to the components of the substrate in the navigation coordinate system:
γ＝γ 0
wherein, γ 0 The roll angle of the navigation coordinate system at the satellite-rocket separation time is determined in advance by the measurement and control requirements of the rocket and the satellite-rocket separation mode, and is taken as-165 degrees in the example.
Step two: and calculating the speeds of the satellite and the rocket in the instantaneous navigation coordinate system after the satellite and the rocket are separated according to the attitude of the navigation coordinate system when each satellite is separated, the separation speed, the speed of the instantaneous rocket before the satellite and the rocket are separated for the first time, the mass of the satellite and the total mass of the rocket, and the installation position of the satellite in the rocket.
Obtained in step one
Psi and gamma are euler angles, and the coordinate base sat _ HJ { satx, party, satz } of the satellite installation position in the rocket is determined (sat _ HJ in the embodiment is the same as the above-mentioned one
) The satellite installation direction is located on the Y axis of the rocket body coordinate system in the embodiment, the coordinate base is sat _ HJ {0,1,0}, and the component of the base sat _ HJ { satx, sar, satz } in the navigation coordinate system is recorded as: sat1_ HJ { satx1, saty1, satz1}.
The conversion matrix of the rocket body coordinate system and the navigation coordinate system is as follows:
the conversion matrix from the arrow coordinate system to the navigation coordinate system is the same as the conversion matrix from the arrow coordinate system to the orbit coordinate system, and is different from euler angles.
The subsequent calculation in this embodiment is performed in the navigation coordinate system, and in other embodiments, the calculation may be performed in other coordinate systems, and the calculation is convenient mainly according to which coordinate system the provided separation speed is relative to.
And (3) according to the mass of the satellite, the total mass m _ sat and m _ HJ of the rocket and the separation speed SEP _ V (relative to the rocket) of the satellite, calculating the separation speeds delta V _ sat and delta V _ HJ obtained by the rocket and the satellite at the moment after the satellite and the rocket are separated by using a momentum theorem.
Calculating the velocities SATV1 (SATVX 1, SATVY1, SATVZ 1) and HJV1 (HJVX 1, HJVY1, HJVZ 1) of the satellite and the rocket at the moment after the separation of the satellite and the rocket, the velocity HJV (HJVX, HJVY, HJVZ) of the rocket at the moment before the separation of the satellite and the rocket for the first time, and the velocities SATV1 (SATVX 1, SATVY1, SATVZ 1) and HJV1 (HJVX 1, HJVY1, HJVZ 1) of the satellite and the rocket at the moment after the separation of the satellite and the rocket (a navigation coordinate system):
SATVX1＝HJVX-Δv_sat*satx1
SATVY1＝HJVY-Δv_sat*saty1
SATVZ1＝HJVZ-Δv_sat*satz1
HJVX1＝HJVX-Δv_HJ*satx1
HJVY1＝HJVY-Δv_HJ*saty1
HJVZ1＝HJVZ-Δv_HJ*satz1
step three: and designing a satellite-rocket separation time sequence according to the number of satellites, the rocket at the first satellite-rocket separation time and the position and speed of the satellites.
Designing the flight time sequence according to the number of satellites, as shown in table 1;
TABLE 1
The steps for calculating the rocket sliding time such as delta T2, delta T3 … and delta TLG are as follows:
performing numerical integration or orbit prediction on the position and the speed of the instantaneous rocket and each separated satellite after each satellite and rocket separation, and judging whether the real-time satellite and rocket distance and the inter-satellite distance meet the requirements or not; when the distance between the rocket and the separated satellite is greater than the safe distance L _ HJTOWX and the distance between the separated satellites is greater than the safe distance L _ WXTOWX, the numerical integration or forecast time length is the interval time length Delta T between the separation of the satellite and the rocket at the time and the separation of the satellite and the satellite at the next time; and repeating the steps until the last satellite is separated.
Specifically, numerical integration or orbit prediction is carried out on the positions and the speeds of the rocket and the first satellite at the moment after the first satellite and the rocket are separated, the distance between the rocket and the first satellite is calculated in real time (only one separated satellite exists at the moment, and the distance between the separated satellites does not need to be calculated), and when the distance between the rocket and the first satellite reaches the safe distance L _ HJTOWX, the integration or prediction time length is the interval time length Delta T2 between the first satellite and rocket separation and the second satellite and rocket separation. In this example, L _ HJTOWX is 200m, and L _WXTOWX is 70m. And after the satellite and the rocket are separated for the second time, updating the rocket mass to m _ HJ-m _ sata-m _ satb, wherein m _ satb and m _ satb are the satellite mass for the first separation of the satellite and the rocket and the second separation of the satellite and the rocket, repeating the first step and the second step, and calculating to obtain the speed of the satellite and the rocket at the moment after the second separation of the satellite and the rocket. And performing numerical integration or orbit prediction on the position and the speed of the instantaneous rocket and each separated satellite after the second satellite-rocket separation, calculating the distance between the rocket and each separated satellite and the distance between the first satellite and the second satellite in real time, wherein when the distances between the rocket and each separated satellite reach the safe distance L _ HJTOWX and the distances between the first satellite and the second satellite reach the safe distance L _ WXTOWX, the integration or prediction duration is the interval duration DeltaT 3 of the second satellite-rocket separation and the third satellite-rocket separation. And repeating the steps, and subtracting 1 from the number of satellites after each separation until the number of the satellites is not more than 0, namely completing the separation of the last satellite.
Designing a rocket off-orbit passivation time sequence delta TLG: after the last satellite is separated, the interval time delta TLG from the separation of the satellite and the arrow to the off-orbit passivation is designed. And updating the rocket mass to m _ HJ-m _ satall, wherein m _ satall is the total mass of all satellites. And performing numerical integration or orbit prediction on the positions and the speeds of the instant rocket and each satellite after the satellite and the rocket are separated, calculating the distances L1, L2 and L3 between the rocket and all the satellites in real time, and calculating the distances L _1to2 and L _1to3 between all the satellites, wherein when the distances between the rocket and all the satellites reach the safe distance L _ HJTOWX and the distances between all the satellites reach the safe distance L _ WXTOWX, the integration or prediction time length is the interval time delta TLG from the separation of the satellite and the passivation of the satellite from the orbit.
When only one satellite exists, after the rocket is separated from the satellite for the first time, the rocket is deactivated after sliding delta TLG; when a plurality of satellites are separated, the rocket sequentially separates the satellites and the rocket according to the designed time sequence and separation posture, and after the last satellite is separated, the rocket is subjected to orbit departure passivation after sliding delta TLG.
The present invention is not limited to the above-described embodiments, and it will be apparent to those skilled in the art that various modifications and improvements can be made without departing from the principle of the present invention, and such modifications and improvements are also considered to be within the scope of the present invention. Those not described in detail in this specification are within the skill of the art.
Claims (9)
1. A carrier rocket satellite-rocket separation time sequence task-removing design method is characterized by comprising the following steps:
establishing an orbit coordinate system, setting three attitude angles A, B, C of arrow bodies in the orbit coordinate system, and determining an attitude angle A, B, C at the satellite-arrow separation time according to a satellite-arrow separation mode;
converting the attitude angle A, B, C in the orbit coordinate system into a navigation coordinate system, and calculating the attitude angle from the arrow coordinate system to the navigation coordinate system according to the component of A, B, C in the navigation coordinate system;
calculating the speeds of the satellite and the rocket in the navigation coordinate system at the satellite and rocket separation moment according to the attitude angle from the rocket body coordinate system to the navigation coordinate system when each satellite is separated, the satellite separation speed, the speed of the rocket at the moment before each satellite and rocket separation, the mass of the satellite and the total mass of the rocket, and the installation position of the satellite in the rocket;
and designing a satellite-rocket separation time sequence according to the number of satellites, the rocket at each satellite-rocket separation time and the position and speed of the satellites.
2. The launch vehicle satellite-rocket separation timing sequence tasking design method of claim 1, wherein the orbit coordinate system is defined as: the origin is the rocket center of mass, and the X axis points to the velocity vector direction of the rocket during separation; the Y axis is positioned in the instantaneous track plane, points to the outer side of the track and is vertical to the X axis; the Z axis is vertical to the X axis and the Y axis and meets the right hand rule.
3. The launch vehicle satellite-rocket separation timing sequence tasking design method of claim 2, wherein the attitude angle A, B, C is set as:
the angle A is an included angle between the projection of the X axis of the arrow body in the track plane and the X axis of the track coordinate system;
the angle B is an included angle of the X axis of the arrow body deviating from the plane of the track;
the angle C is the rotation angle of the arrow body around the X axis, and the Y axis of the arrow body is in the track plane when C = 0.
4. The launch vehicle satellite-rocket separation timing sequence tasking design method of claim 1, wherein said converting attitude angle A, B, C in an orbital coordinate system to a navigational coordinate system comprises:
obtaining a conversion matrix of the rocket coordinate system and the orbit coordinate system by taking A, B, C as an Euler angle, and enabling a coordinate base of a satellite in the rocket coordinate system
Converting the conversion matrix into a track coordinate system through an arrow coordinate system and the track coordinate system; substrate
The components in the orbital coordinate system are noted as:
will be provided with
From orbital to geocentric coordinate systems, bases
The components in the geocentric coordinate system are denoted as:
will be provided with
From the geocentric coordinate system to the emission coordinate system, the base
The components in the emission coordinate system are noted as:
will be provided with
From the emission coordinate system to the navigation coordinate system, the base
The components in the navigation coordinate system are noted as:
according to the substrate
And calculating the components in the navigation coordinate system to obtain the attitude angle from the rocket coordinate system to the navigation coordinate system:
γ＝yO
wherein gamma 0 is the rolling angle of the navigation coordinate system at the satellite-rocket separation time, and is determined in advance by the measurement and control requirements of the rocket and the satellite-rocket separation mode.
5. The carrier rocket satellite-rocket separation time sequence task-removing design method as claimed in claim 4, wherein the calculating of the satellite and rocket speeds in the instantaneous navigation coordinate system after the separation of the satellite and the rocket according to the navigation coordinate system attitude angle when each satellite is separated, the satellite separation speed, the speed of the instantaneous rocket before the separation of the satellite and the rocket for the first time, the satellite mass and the rocket total mass and the installation position of the satellite in the rocket comprises:
to obtain
Psi and gamma are Euler angles, a coordinate base sat _ HJ { satx, party, satz } of the satellite installation position in the rocket is converted into a navigation coordinate system through a conversion matrix of the rocket body coordinate system and the navigation coordinate system, and the component of the base of the satellite installation position in the navigation coordinate system is recorded as: sat1_ HJ { satx1, saty1, satz1};
according to the mass of the satellite, the total mass m _ sat and m _ HJ of the rocket and the separation speed SEP _ V of the satellite, calculating the separation speeds delta V _ sat and delta V _ HJ obtained by the rocket and the satellite at the moment after the satellite and the rocket are separated by adopting a momentum theorem;
the velocities SATV1 (SATVX 1, SATVY1, SATVZ 1) and HJV1 (HJVX 1, HJVY1, HJVZ 1) of the satellite and the rocket at the moment after the separation of the satellite and the rocket are calculated from sat1_ HJ { satx1, satz1}, delta v _ sat, delta v _ HJ and the velocity HJV (HJVX, HJVY, HJVZ 1) of the rocket at the moment before the separation of the satellite and the rocket for the first time.
6. The method for mission-free design of a carrier rocket satellite-rocket separation time sequence according to claim 1, wherein the design of the satellite-rocket separation time sequence according to the number of satellites, the rocket at the first satellite-rocket separation time, and the positions and speeds of the satellites comprises the following steps:
performing numerical integration or orbit prediction on the position and the speed of the instant rocket and each separated satellite after each satellite and rocket separation; when the distance between the rocket and the separated satellite is greater than the safe distance L _ HJTOWX and the distance between the separated satellites is greater than the safe distance L _ WXTOWX, the numerical integration or forecast time length is the interval time length Delta T between the separation of the satellite and the rocket at the time and the separation of the satellite and the satellite at the next time; and repeating the steps until the last satellite is separated.
7. The method for mission-free design of a carrier rocket satellite-rocket separation time sequence according to claim 6, wherein the design of the satellite-rocket separation time sequence according to the number of satellites, the first satellite-rocket separation time, and the positions and speeds of the satellites comprises the following steps: after the last satellite and the satellite are separated, designing the interval time delta TLG from the separation of the satellite and the satellite to the off-orbit passivation, and updating the rocket mass to m _ HJ-m _ satall, wherein m _ satall is the total mass of all satellites;
and performing numerical integration or orbit prediction on the position and the speed of the instant rocket and each satellite after the satellite and the rocket are separated, calculating the distance between the rocket and each satellite and the distance between all satellites in real time, and 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, wherein the integration or prediction time length is the interval time length delta TLG from the separation of the satellite and the rocket to the off-orbit passivation.
8. A carrier rocket satellite-rocket separation time sequence task-removing design system is characterized by comprising:
the coordinate system establishing module is used for establishing an orbit coordinate system, setting three attitude angles A, B, C of an arrow body in the orbit coordinate system, and determining an attitude angle A, B, C according to a satellite-arrow separation mode;
the attitude angle calculation module is used for converting an attitude angle A, B, C in the orbit coordinate system into a navigation coordinate system and calculating the attitude angle of the navigation coordinate system according to the component of A, B, C in the navigation coordinate system;
the speed calculation module is used for calculating the speeds of the satellite and the rocket in the instantaneous navigation coordinate system after the separation of the satellite and the rocket according to the attitude angle of the navigation coordinate system when each satellite is separated, the separation speed of the satellite, the speed of the rocket in the moment before the separation of the satellite and the rocket for the first time, the mass of the satellite and the total mass of the rocket and the installation position of the satellite in the rocket;
and the time sequence design module is used for designing a satellite and rocket separation time sequence according to the number of the satellites, the rocket at the first satellite and rocket separation time and the position and speed of the satellites.
9. The system for mission-free design of separation timing sequence of launch vehicle satellite and rocket according to claim 8, wherein the design method according to any one of claims 1-7 is adopted to design the separation timing sequence of launch vehicle satellite and rocket.
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| CN116501077B (en) * | 2023-06-27 | 2023-09-15 | 航天科工火箭技术有限公司 | Rocket attitude angle automatic optimization method constrained by space-based measurement and control |
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