CN109484675B - Spacecraft orbit-in control method by utilizing space vector matching - Google Patents
Spacecraft orbit-in control method by utilizing space vector matching Download PDFInfo
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Abstract
The invention discloses a spacecraft orbit entering control method by utilizing space vector matching, which comprises the following steps of: initializing spacecraft state parameters; calculating the number of the current orbits of the spacecraft, the geocentric radial diameter and the rotation angle of the current orbit relative to the standard orbit; obtaining a standard orbit velocity vector which is suitable for the current orbit state according to the geocentric radial and the rotation angle of the current orbit relative to the standard orbit; calculating the deviation amount of the speed vector, the residual flight time DT, the pitching program angle and the yawing program angle according to the standard orbit speed vector and the current actual orbit speed vector; and performing attitude control and shutdown control by using the calculated residual flight time DT, the pitch program angle and the yaw program angle, and relating to the technical field of track control. The invention has the advantages of strong real-time performance, high guidance precision, strong adjustment capability on the orbit, simple flight software on the spacecraft and low requirement on ground preparation and calculation work.
Description
Technical Field
The invention relates to the technical field of orbit control, in particular to a spacecraft orbit in-orbit control method by utilizing space vector matching.
Background
For a spacecraft adopting a solid engine at the base level, the engine generally adopts an exhausted shutdown mode, and the engine cannot adjust the shutdown time of the engine according to the current actual condition, so that a larger position and speed deviation may be generated relative to the theoretical condition, and the change amplitude of the program angle calculated by the last stage when the orbit planning is carried out again relative to the standard program angle may be larger. On the other hand, the spacecraft generally adopts an iterative guidance control mode in order to improve the final stage orbit-entering precision, but the traditional iterative guidance mode generally takes a terminal orbit as a control target, and has poor adaptability to the large attitude adjustment condition in the small overload state in the flight process.
Disclosure of Invention
The invention aims to overcome the defects of the background technology and provide a spacecraft orbit control method which has the advantages of simple and easily realized thought, small calculation amount, improvement on spacecraft final stage orbit accuracy and realization of large-range attitude control by utilizing space vector matching.
The invention provides a spacecraft orbit entering control method by utilizing space vector matching, which comprises the following steps of:
1) initializing spacecraft state parameters;
2) calculating the number of the current orbits of the spacecraft, the geocentric radius and the rotation angle of the current orbit relative to the standard orbit;
3) obtaining a standard orbit velocity vector which is suitable for the current orbit state according to the geocentric radius and the rotation angle of the current orbit relative to the standard orbit;
4) calculating the deviation amount of the speed vector according to the standard track speed vector and the current actual track speed vector, and calculating the residual flight time DT, the pitching program angle and the yawing program angle by using the deviation amount of the speed vector;
5) and performing attitude control and shutdown control by using the calculated remaining flight time DT, the pitch program angle and the yaw program angle.
On the basis of the technical scheme, the spacecraft state parameters in the step 1) comprise:
transmitting latitude, longitude, elevation, direction, current time of flight, position, velocity, geocentric radius, and an interpolation table of orbit parameters for the current orbit with respect to a standard orbit rotation angle.
On the basis of the technical scheme, the control quantity of the number of the tracks in the step 2) mainly comprises the following steps:
semi-major axis of the track, inclination angle of the track, and eccentricity of the track.
On the basis of the technical scheme, the standard orbit in the step 2) refers to a flight orbit of the spacecraft under the theoretical condition without any deviation condition, and the rotation angle refers to an included angle between an actual orbit and the standard orbit.
On the basis of the above technical solution, the obtaining of the standard orbit velocity vector that should be adapted to the current orbit state according to the geodesic radius and the rotation angle of the current orbit relative to the standard orbit in step 3) specifically means:
and searching a standard track parameter interpolation value suitable for the current track state according to a pre-bound track parameter interpolation table which changes along with the geocentric radius and the rotation angle of the current track relative to the standard track, and calculating to obtain a standard track speed vector.
On the basis of the above technical solution, the calculating the speed deviation amount, the remaining flight time DT, the pitch program angle, and the yaw program angle according to the standard orbit speed vector and the current actual orbit speed vector in step 4) specifically means:
and converting the standard track speed vector and the actual track speed vector into the same reference coordinate system, calculating the speed deviation amount in three coordinate directions, and calculating the residual flight time DT, the pitch program angle and the yaw program angle according to the speed deviation amount in the three coordinate directions.
On the basis of the above technical solution, the specifically performing attitude control and shutdown control by using the remaining flight time t in step 5) includes:
if the residual flight time DT is less than or equal to delta T, stopping the calculation of the guidance system, and closing the engine after the flight is finished according to the residual flight time DT calculated at the last time and the corresponding pitching program angle and yawing program angle;
and if the residual flight time DT & ltDELTA & gt is obtained, returning to the step 1) to perform guidance calculation of the next period.
On the basis of the technical scheme, spacecraft state parameters, interpolation calculation, calculation of residual flight time DT, pitching program angle and yawing program angle, attitude control and shutdown control are initialized in each guidance calculation period.
Compared with the prior art, the invention has the following advantages:
according to the spacecraft orbit in-orbit control method based on space vector matching, high-precision in-orbit is achieved through control over the intermediate flight speed vector of a spacecraft, the method is simple and easy to implement in engineering, and the spacecraft orbit in-orbit control method has a high engineering application value; compared with the traditional guidance scheme, the control method has the advantages of strong real-time performance, high guidance precision, strong adjustment capability on the orbit, simple flight software on the spacecraft, low requirement on ground preparation calculation work and suitability for the space spacecraft with high-precision guidance and large-attitude adjustment.
Drawings
FIG. 1 is a flowchart of a spacecraft orbit entry control method using space vector matching according to an embodiment of the present invention;
fig. 2 is a schematic diagram of a guidance-in section of a transition track in dynamic rotation according to an embodiment of the present invention.
Detailed Description
The invention is described in further detail below with reference to the figures and the embodiments.
Example one
Referring to fig. 1, an embodiment of the present invention provides a spacecraft orbit entry control method using space vector matching, including the following steps:
1) initializing spacecraft state parameters;
2) calculating the number of the current orbits of the spacecraft, the geocentric radius and the rotation angle of the current orbit relative to the standard orbit;
3) obtaining a standard orbit velocity vector which is suitable for the current orbit state according to the geocentric radius and the rotation angle of the current orbit relative to the standard orbit;
4) calculating the deviation amount of the speed vector according to the standard track speed vector and the current actual track speed vector, and calculating the residual flight time DT, the pitching program angle and the yawing program angle by using the deviation amount of the speed vector;
5) and performing attitude control and shutdown control by using the calculated remaining flight time DT, the pitch program angle and the yaw program angle.
According to the spacecraft orbit in-orbit control method based on space vector matching, high-precision in-orbit is achieved through control over the intermediate flight speed vector of a spacecraft, the method is simple and easy to implement in engineering, and the spacecraft orbit in-orbit control method has a high engineering application value; compared with the traditional guidance scheme, the control method has the advantages of strong real-time performance, high guidance precision, strong adjustment capability on the orbit, simple flight software on the spacecraft, low requirement on ground preparation calculation work and suitability for the space spacecraft with high-precision guidance and large-attitude adjustment.
Example two
Referring to fig. 1 to 2, an embodiment of the present invention provides a spacecraft orbit entry control method using space vector matching, including the following steps:
1) initializing spacecraft state parameters;
before iterative computation, current state parameters of a spacecraft need to be initialized, and the method mainly comprises the following steps:
transmitting latitude, longitude, elevation, direction, current time of flight, position, velocity, geocentric radius, and an interpolation table of orbit parameters for the current orbit with respect to a standard orbit rotation angle.
2) Calculating the number of the current orbits of the spacecraft, the geocentric radius and the rotation angle of the current orbit relative to the standard orbit;
calculating the orbit number of a current point according to current spacecraft orbit state parameters, and solving the geocentric vector sum and the rotation angle relative to a standard orbit in real time according to the orbit number, wherein the orbit number control quantity mainly comprises: a semi-major axis of the track, a track inclination angle and a track eccentricity; the standard orbit refers to a flight orbit of the spacecraft under the theoretical condition without any deviation condition, and the actual flight orbit deviates from the standard orbit under the influence of various deviation factors in the actual condition; the rotation angle refers to an included angle between the actual track and the standard track.
3) Obtaining a standard orbit velocity vector which is suitable for the current orbit state according to the geocentric radius and the rotation angle of the current orbit relative to the standard orbit;
finding a standard speed vector corresponding to the current geocentric vector diameter and the track according to a two-dimensional number table which is bound in advance and is interpolated along with the geocentric vector and a rotating angle relative to the standard track, comparing the standard speed vector with the current actual flying speed, and calculating a speed deviation value; corresponding theoretical orbit velocity vectors correspond to different geocentric vector diameters and orbit rotation angles, and the spacecraft can accurately enter a preset transfer orbit only when the actual orbit velocity reaches or approaches the theoretical orbit velocity; the speed deviation amount is a vector state, and comprises the magnitude and the direction of the speed deviation amount.
4) Calculating a speed deviation amount, a residual flight time DT, a pitching program angle and a yawing program angle according to the standard orbit speed vector and the current actual orbit speed vector;
and converting the standard track speed vector and the actual track speed vector into the same reference coordinate system, calculating the speed deviation amount in three coordinate directions, and calculating the residual flight time DT, the pitch program angle and the yaw program angle according to the speed deviation amount in the three coordinate directions.
5) And carrying out attitude control and shutdown control by using the residual flight time DT.
Specifically, the method comprises the following steps: if the residual flight time DT is less than or equal to delta T, stopping the calculation of the guidance system, and closing the engine after finishing the flight according to the last calculated residual flight time DT and the corresponding pitching program angle and yawing program angle;
and if the residual flight time DT & ltDELTA & gt is obtained, returning to the step 1) to perform guidance calculation of the next period.
As shown in fig. 2, fig. 2 is a guidance-in interval schematic diagram of dynamic rotation of a transition orbit according to an embodiment of the present invention, in which a trajectory group formed by an elliptical orbit shape and rotation thereof is used as a core, an orbit parameter of a current point of a spacecraft is used as an initial value, a flight trajectory determined by considering J2 perturbation mainly includes orbit parameters such as a geocentric vector, an absolute velocity, a position-velocity vector included angle, a true paraxial point angle, and the like, and specifically, the transition orbit is a trajectory that is determined by considering a perturbation of J2The process is as follows: calculating the geocentric radial diameter of the current point and the rotation angle of the current orbit relative to the standard elliptical orbit according to the current velocity vector and the position vector of the spacecraft, and interpolating from a preset reference orbit to determine a corresponding standard orbit entering velocity vector (including the size and the direction), wherein an interpolation interval corresponds to that in FIG. 2
The remaining flight time, pitch program angle and yaw program angle are then adjusted based on the difference between the actual velocity vector and the standard orbit entry velocity vector. Updating the initial velocity vector and the position vector of the spacecraft again, repeating the process, updating the final-stage residual flight time, the pitching program angle and the yawing program angle, stopping the calculation of the guidance system when the time control requirement is met, namely if the residual flight time DT is less than or equal to delta T, and closing the engine after the flight is finished according to the last calculated residual flight time DT and the corresponding pitching program angle and yawing program angle; if the remaining flight time DT>And returning to delta T, and performing guidance calculation of the next period.
According to the spacecraft orbit in-orbit control method based on space vector matching, high-precision in-orbit is achieved through control over the intermediate flight speed vector of a spacecraft, the method is simple and easy to implement in engineering, and the spacecraft orbit in-orbit control method has a high engineering application value; compared with the traditional guidance scheme, the control method has the advantages of strong real-time performance, high guidance precision, strong adjustment capability on the orbit, simple flight software on the spacecraft, low requirement on ground preparation calculation work and suitability for the space spacecraft with high-precision guidance and large-attitude adjustment.
Various modifications and variations of the embodiments of the present invention may be made by those skilled in the art, and they are also within the scope of the present invention, provided they are within the scope of the claims of the present invention and their equivalents.
What is not described in detail in the specification is prior art that is well known to those skilled in the art.
Claims (6)
1. A spacecraft orbit control method utilizing space vector matching is characterized by comprising the following steps:
1) initializing spacecraft state parameters;
2) calculating the number of the current orbits of the spacecraft, the geocentric radius and the rotation angle of the current orbit relative to the standard orbit;
3) obtaining a standard orbit velocity vector which is suitable for the current orbit state according to the geocentric radius and the rotation angle of the current orbit relative to the standard orbit;
4) calculating the deviation amount of the speed vector according to the standard track speed vector and the current actual track speed vector, and calculating the residual flight time DT, the pitching program angle and the yawing program angle by using the deviation amount of the speed vector;
5) performing attitude control and shutdown control by using the calculated residual flight time DT, the pitch program angle and the yaw program angle;
the spacecraft state parameters in the step 1) comprise:
transmitting an orbit parameter interpolation table of latitude, longitude, elevation, direction, current flight time, position, speed, geocentric radius and current orbit rotation angle relative to a standard orbit;
the step 3) of obtaining the standard orbit velocity vector which is suitable for the current orbit state according to the geocentric radial and the rotation angle of the current orbit relative to the standard orbit specifically refers to:
and searching a standard track parameter interpolation value suitable for the current track state according to a pre-bound track parameter interpolation table which changes along with the geocentric radius and the rotation angle of the current track relative to the standard track, and calculating to obtain a standard track speed vector.
2. A spacecraft orbit input control method using space vector matching according to claim 1, wherein the orbit number control quantity in the step 2) mainly comprises:
semi-major axis of the track, inclination angle of the track, and eccentricity of the track.
3. A spacecraft orbit entry control method using space vector matching as claimed in claim 1, wherein: the standard orbit in the step 2) refers to a flight orbit of the spacecraft under the theoretical condition without any deviation condition, and the rotation angle refers to an included angle between an actual orbit and the standard orbit.
4. A spacecraft orbit entering control method using space vector matching according to claim 1, wherein the step 4) of calculating the velocity deviation amount, the remaining flight time DT, the pitch program angle and the yaw program angle according to the standard orbit velocity vector and the current actual orbit velocity vector specifically means:
and converting the standard track speed vector and the actual track speed vector into the same reference coordinate system, calculating the speed deviation amount in three coordinate directions, and calculating the residual flight time DT, the pitch program angle and the yaw program angle according to the speed deviation amount in the three coordinate directions.
5. The method for controlling orbit entering of spacecraft by using space vector matching according to claim 1, wherein said attitude control and shutdown control by using said calculated pitch program angle, yaw program angle and remaining time of flight DT in step 5) specifically includes:
if the residual flight time DT is less than or equal to delta T, stopping the calculation of the guidance system, and closing the engine after finishing the flight according to the last calculated residual flight time DT and the corresponding pitching program angle and yawing program angle;
and if the residual flight time DT & ltDELTA & gt is obtained, returning to the step 1) to perform guidance calculation of the next period.
6. A spacecraft orbit entry control method using space vector matching as claimed in claim 5, wherein: in each guidance calculation period, spacecraft state parameters, interpolation calculation, residual flight time DT, pitching program angle and yawing program angle calculation need to be initialized, and attitude control and shutdown control are carried out.
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