CN113320717A - Guidance system reconstruction method for dealing with one-time ignition fault - Google Patents
Guidance system reconstruction method for dealing with one-time ignition fault Download PDFInfo
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Abstract
A guidance system reconstruction method for dealing with a primary ignition fault is characterized in that for an upper stage, a working mode of satellite orbit entering is ensured through two times of active section orbit changing, when primary ignition fails, secondary ignition can be started immediately, the upper stage of a sub-orbit changing section can reach a sufficient height, and an engine extrusion working mode is adopted by a second active section to push an upper stage/satellite assembly to an orbit entering point. The first active segment adopts an iterative guidance method, the second active segment adopts a guidance method and binds data, and the precision of the track entry is ensured. The guidance system reconstruction method for dealing with the primary ignition fault can fully utilize the extrusion working mode of the engine when the primary ignition fault occurs, autonomously judge and reconstruct on line, ensure that the satellite orbit-entering precision meets the requirement under the fault condition, and improve the reliability of the system.
Description
Technical Field
The invention relates to a guidance system reconstruction method for dealing with a one-time ignition fault, and belongs to the technical field of guidance control.
Background
When the upper stage launches a Low Earth Orbit (LEO) satellite task, a main engine is generally adopted for twice ignition and orbital transfer, and an iterative guidance method is adopted for twice orbital transfer. However, the upper stage orbital transfer thrust is provided by a single main engine, and the ignition of two active sections of the main engine is triggered by a single point, so that the risk of ignition failure exists. Therefore, the emission is lost due to one-time ignition failure, and the reliability of the system can be improved to a great extent by carrying out the guidance system reconstruction strategy according to the fault condition, so that the emission success rate is improved.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the method overcomes the defects of the prior art, provides a guidance system reconstruction method for dealing with primary ignition faults, ensures the working mode of satellite orbit entering for the upper stage through twice active section orbit changing, can immediately start secondary ignition when primary ignition fails, ensures that the upper stage can reach enough height on a sub-orbit changing section, and adopts an engine extrusion working mode for the second active section to push the upper stage/satellite assembly to the orbit entering point.
When the upper stage does not have faults, the two active section guidance methods both adopt iterative guidance, the iterative guidance has good adaptability to faults of a power system of the carrier, the iterative guidance is explicit guidance, and a control angle is generated according to an instantaneous state and estimation of the polarity of the future thrust condition, so that the description of the future thrust curve influences the accurate solution of the instantaneous control attitude angle. When the first ignition fails, the first active section can still adopt an iterative guidance method, but the thrust of the second active section adopting an engine extrusion working mode is very small, and the generated apparent acceleration is smaller than the gravity acceleration and is not enough to overcome the average gravity model deviation in iterative guidance, so that the second active section cannot be used, the guidance method and the binding element need to be replaced, and the tracking precision is ensured.
The purpose of the invention is realized by the following technical scheme:
a guidance system reconstruction method for dealing with one-time ignition faults comprises the following steps:
dividing an autonomous flight arc segment of the spacecraft into a first segment, a second segment and a tail segment according to a time sequence; normally, the engine of the spacecraft works in the local time of the first section and the second section;
on the ground, establishing a one-time firing fault flight sequence: after a primary ignition instruction in a first period is sent out, the state of the engine is identified within a preset time, when the identification result is that primary ignition fails, the engine is ignited again, the engine is in an extrusion working mode within the preset time, and the engine is in a normal working mode after the preset time; the spacecraft in the first section adopts an iterative guidance mode; in the second section, the engine is in an extrusion working mode, and the spacecraft adopts a positive thrust vector guidance mode; finally generating flight data under the fault state;
in the actual flight process, when a primary ignition fault occurs, guidance control is carried out by adopting a primary ignition fault flight time sequence and flight data under a fault state.
In the guidance system reconstruction method, the method for identifying the engine state within the preset time comprises the following steps: and when the axial speed increment of the spacecraft is smaller than the corresponding threshold value and the pressure after the engine pump and the rotating speed of the turbine of the spacecraft are respectively lower than the corresponding threshold values, the identification result is that the primary ignition fails.
In the guidance system reconstruction method, during the identification process, the fault identification words are set, the axial speed increment of the spacecraft, the pressure after the engine pump and the turbine speed are judged for multiple times, and in each judgment, when the axial speed increment of the spacecraft is smaller than the corresponding threshold value, and the pressure after the engine pump and the turbine speed of the spacecraft are respectively lower than the corresponding threshold values, the fault identification words are set to be 1, and when the fault identification words for multiple times are all set to be 1, the identification result is that one-time ignition fails.
In the guidance system reconstruction method, the positive thrust vector guidance mode is as follows: and adjusting the thrust direction of the engine to be consistent with the speed direction for accelerating, so as to improve the height of the track at the highest speed.
The guidance system reconstruction method has the axial speed increment threshold of
K is a coefficient of a threshold value,
a threshold thrust is determined for the engine,
and taking the theoretical takeoff mass.
A guidance system reconstruction device for dealing with one-time ignition faults divides an autonomous flight arc section of a spacecraft into a first section, a second section and a tail section according to a time sequence; normally, the engine of the spacecraft works in the local time of the first section and the second section; the guidance system reconfiguration device includes:
the primary ignition fault flight time sequence module is internally provided with a primary ignition fault flight time sequence: after an ignition instruction in the first period is sent out, the state of the engine is identified within a preset time, when the identification result is that the ignition fails in sequence, the engine is ignited again, the engine is in an extrusion working mode within the preset time, and the engine is in a normal working mode after the preset time; the spacecraft in the first section adopts an iterative guidance mode; in the second section, the engine is in an extrusion working mode, and the spacecraft adopts a positive thrust vector guidance mode; finally generating flight data under the fault state;
and the data switching module is used for controlling guidance control by adopting a one-time ignition fault flight time sequence and flight data in a fault state when one-time ignition fault occurs in the actual flight process.
A detection and reconstruction method for a guidance system under the condition of spacecraft ignition fault comprises the following steps:
dividing an autonomous flight arc segment of the spacecraft into a first segment, a second segment and a tail segment according to a time sequence; wherein the engines of the spacecraft operate during the first and second segments of local time;
after the engine is ignited once in the first section, the ignition state is identified, if the ignition fails once, the engine is in a squeezing working mode, otherwise, the engine is in a normal working mode; when the ignition is failed once, the ignition is restarted, and the engine is in a normal working mode; the spacecraft in the first section adopts an iterative guidance mode;
after entering the second section, the engine is in an extrusion working mode, and the spacecraft adopts a positive thrust vector guidance mode.
A detection and reconstruction device of a guidance system under the condition of ignition failure of a spacecraft divides an autonomous flight arc section of the spacecraft into a first section, a second section and a tail section according to a time sequence; wherein the engines of the spacecraft operate during the first and second segments of local time; the guidance system detection reconstruction device includes:
the first section detection guidance module is used for identifying the primary ignition state after the engine is ignited for one time in the first section, if the primary ignition fails, the engine is in an extrusion working mode, otherwise, the engine is in a normal working mode; when the ignition is failed once, the ignition is restarted, and the engine is in a normal working mode; the spacecraft in the first section adopts an iterative guidance mode;
and after the second section of the guidance module enters the second section, the engine is in an extrusion working mode, and the spacecraft adopts a positive thrust vector guidance mode.
Compared with the prior art, the invention has the following beneficial effects:
(1) the guidance system reconstruction method for dealing with the primary ignition fault can fully utilize the extrusion working mode of the engine when the primary ignition fault occurs, autonomously judge and reconstruct on line, ensure that the satellite orbit-entering precision meets the requirement under the fault condition, and improve the reliability of the system;
(2) redesigning the task track under the fault condition according to the primary ignition fault timing to obtain more optimized data, designing a guidance system based on the track parameters under the fault condition, and improving the track entering precision under the fault condition;
(3) according to the difference of the thrust of the normal working mode and the extrusion working mode of the engine, a proper guidance method is selected when the guidance system is reconstructed, and the rail entering precision under the comprehensive deviation is improved;
(4) according to the actual flight characteristics of the upper stage, a one-time ignition fault on-line judgment method is designed, the axial speed increment of the upper stage is taken as the main part, two parameters of the pressure after the pump of the engine and the turbine rotating speed of the engine are taken as the auxiliary parts, and continuous judgment is carried out for many times, so that the situations of misjudgment, missed judgment and the like are avoided.
Drawings
FIG. 1 is a schematic diagram of the reconstruction scheme of the primary ignition fault guidance system of the upper stage.
Fig. 2 is a timing diagram of the upper stage primary ignition fault rail design.
FIG. 3 is a flow chart of the reconstruction of the primary ignition fault guidance system of the upper stage.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, embodiments of the present invention will be described in detail with reference to the accompanying drawings.
The upper stage is a single engine which sends the satellite into the orbit through two times of active section orbital transfer, and the current carrier rocket can ensure stable posture to complete tasks by changing an operation model under the condition of redundant thrust, but the fault-tolerant control method of the two-time active working orbital transfer working mode of the main engine of the upper stage is not provided. The invention aims to reconstruct a guidance system when the first ignition fails when the LEO task is executed at the upper level, so as to realize fault-tolerant design.
The invention discloses a fault-tolerant control method for one-time ignition fault handling of an upper-level LEO task, which adopts the technical scheme that: after a primary ignition instruction is sent out, judging the primary ignition state, and if the ignition is successful, performing track entry according to a preset scheme; if the ignition fails, the ignition is immediately re-ignited in the first active section, and the second active section adopts an engine extrusion working mode to realize fault-tolerant control. The method comprises the following specific steps:
1) one-time ignition fault track design
The basic dynamics model of the upper level is established as follows:
in the formula: the origin of the launching coordinate system is fixedly connected with a rocket launching point o, oxgThe axis is in the horizontal plane of the emission point and points to the emission aiming direction; oygThe axis is vertical to the horizontal plane of the emission point and points upwards; ozgAxis and xgoygThe planes are perpendicular and form a right-hand coordinate system. And the launching inertial coordinate system O-xyz is completely overlapped with the ground launching coordinate system at the launching moment of the carrier rocket, and then is fixed in the inertial space. The relation position of the coordinate system and the earth center inertia coordinate system is fixed and invariable. The emission inertia coordinate system is a main coordinate system of guidance calculation, and navigation calculation and attitude angle calculation are carried out in the coordinate system, which is also called as a guidance calculation coordinate system. Arrow coordinate system o1-x1y1z1The origin of the coordinate system is taken at the rocket centroid; x is the number of1The axis points in the direction of the head along the rocket longitudinal axis; y is1The axis being in the longitudinal plane of the arrow and x1The axis is vertical and points upwards; z is a radical of1Axis and x1、y1The shaft constitutes a right-hand system.
The components of the apparent acceleration in the x, y and z axes of the inertia system, gxa、gya、gzaAre respectively the components of the gravitational acceleration on the x, y and z axes of the inertia generating system,
psi and gamma are respectively the pitch angle, yaw angle, roll angle, Vxa、Vya、VzaThe components of the upper stage speed on the x, y and z axes of the inertia generating system are respectively, m is the mass of the upper stage, and F is the main engine thrust of the upper stage. And aiming at the basic dynamic model, determining the track under the condition of the primary ignition fault according to the time sequence and the engine thrust under the condition of the primary ignition fault, thereby providing more optimized track parameters for subsequent guidance reconstruction.
2) And (4) automatically judging the primary ignition fault. And starting to judge the ignition fault condition after primary ignition, wherein the parameters participating in the judgment comprise the pressure delta P after the engine pump, the turbine rotating speed delta n and the axial speed increment delta W, the pressure after the engine pump and the turbine rotating speed are directly obtained through bus data, and the axial speed increment is obtained through measurement. Axial velocity increment threshold of
In the formula (I), the compound is shown in the specification,
the unit of (a) is meter/second; k is a threshold coefficient, generally 1/3-1/10 is taken, and in the invention, 1/4 is taken;
judging threshold thrust for the engine, wherein the unit is N; delta t is the time zone of axial velocity increment threshold detectionTime, in seconds; and m is the theoretical takeoff mass in kg. Firstly, judging the axial speed increment delta W of the upper stage, judging the pressure delta P behind the engine pump and the turbine speed delta n when the axial speed increment delta W is smaller than a threshold value, setting a fault identification word to be 1 when the two are lower than the respective threshold values, setting the fault identification word to be 0 under other conditions, and proving that one-time ignition fails when the fault identification words are 1 for 5 continuous times.
3) And selecting a guidance method. The invention relates to two guidance methods:
a) iterative guidance: determining a program angle of a thrust vector during an iterative guided computation
And psiζCan be approximated as a linear function of time (in an orbital coordinate system), and shows from the law of control of the rocket, the control angle formed to ensure that the rocket reaches a predetermined velocity vector at a target point
Occupying the whole control angle
ψζThe control angle for satisfying the position constraint is only a small amount of the main part, and for solving, the control attitude angle can be simplified into the following form:
in the formula k2t-k1、e2t-e1To satisfy the control amount of the terminal position constraint.
b) Positive thrust vector guidance
The positive thrust vector guidance method adjusts the thrust direction to be consistent with the speed direction, accelerates the whole force and improves the height of the track at the highest speed.
Vx, Vy, Vx are navigation speeds,
ψcxand the command program angle is used for providing the attitude control system, and the attitude control system realizes a guidance command through the attitude adjusting tracking program angle.
Example (b):
step one, establishing primary ignition fault flight data on the ground. A one-time fired fault flight sequence is established based on engine operating characteristics and time constraints under one-time fired fault conditions, as shown in fig. 1-3. Compared with the flight time sequence under the normal condition, firstly, after the ignition instruction of the first active section is sent out, the fault self-identification section is added, the time of the section is set to be 15s according to the working characteristic constraint of the engine, when the fault track is designed, the main engine of the section is set to be in an extrusion working mode, the working state of the first active section is the same as that of the standard track, and when the fault track is designed, the thrust of the engine is iterated according to the extrusion working mode to generate new flight data. After the design of the primary ignition fault track of the upper level is finished, the guidance system is designed by adopting a reconstruction strategy of iterative guidance of a first active section and positive thrust vector guidance of a second active section, the requirement of satellite orbit entering precision is met under the comprehensive deviation through Monte Carlo targeting simulation, and the verification of flight data under the fault state is finished.
And step two, automatically identifying the primary ignition fault condition on line. After the ignition instruction of the first active section is sent out, three parameters of overload, engine pump back pressure and turbine speed in the bus are judged and identified on line. Setting the judgment condition as
(first, the axial speed increment of the upper stage is judged, and when the axial speed increment is smaller than the threshold valueAnd judging the back pressure of the engine pump and the turbine speed, setting the fault identifier word to be 1 when the back pressure of the engine pump and the turbine speed are both lower than the threshold value, setting the fault identifier word to be 0 under other conditions, and proving that one-time ignition fails when the fault identifier words are all 1 for 5 continuous times. )
And step three, executing a reconstruction strategy and performing data switching. According to the online identification result, when primary ignition fails, secondary ignition is immediately used 15s after the ignition moment of the first active section, and the success of sub-rail track change is ensured. The second active section adopts an engine extrusion working mode to provide power, the thrust provided by the engine extrusion working mode is smaller than that provided by a normal working mode, and the assumed condition of iterative guidance is not met, so that the second active section needs to switch a guidance strategy, adopts a positive thrust vector guidance strategy, and raises the upper stage to the height of the required orbit of the satellite in the second active section as much as possible, thereby reducing errors, and simultaneously, switching fault data, thereby improving the orbit entering precision.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.
Claims (8)
1. A guidance system reconstruction method for dealing with one-time ignition fault is characterized by comprising the following steps:
dividing an autonomous flight arc segment of the spacecraft into a first segment, a second segment and a tail segment according to a time sequence; normally, the engine of the spacecraft works in the local time of the first section and the second section;
on the ground, establishing a one-time firing fault flight sequence: after a primary ignition instruction in a first period is sent out, the state of the engine is identified within a preset time, when the identification result is that primary ignition fails, the engine is ignited again, the engine is in an extrusion working mode within the preset time, and the engine is in a normal working mode after the preset time; the spacecraft in the first section adopts an iterative guidance mode; in the second section, the engine is in an extrusion working mode, and the spacecraft adopts a positive thrust vector guidance mode; finally generating flight data under the fault state;
in the actual flight process, when a primary ignition fault occurs, guidance control is carried out by adopting a primary ignition fault flight time sequence and flight data under a fault state.
2. The guidance system reconstruction method according to claim 1, wherein the method of identifying the engine state within a preset time is: and when the axial speed increment of the spacecraft is smaller than the corresponding threshold value and the pressure after the engine pump and the rotating speed of the turbine of the spacecraft are respectively lower than the corresponding threshold values, the identification result is that the primary ignition fails.
3. The guidance system reconstruction method according to claim 1, characterized in that in the identification process, a fault identification word is set, and multiple judgments are performed on the axial velocity increment of the spacecraft, the post-engine-pump pressure and the turbine speed, wherein in each judgment, when the axial velocity increment of the spacecraft is smaller than a corresponding threshold value, and the post-engine-pump pressure and the turbine speed of the spacecraft are respectively lower than the corresponding threshold values, the fault identification word is set to 1, and when the fault identification words for a plurality of consecutive times are all set to 1, the identification result is that one-time ignition fails.
4. The guidance system reconstruction method according to claim 1, wherein the positive thrust vector guidance mode is: and adjusting the thrust direction of the engine to be consistent with the speed direction for accelerating, so as to improve the height of the track at the highest speed.
5. The guidance system reconstruction method of claim 2 or 3, wherein the axial velocity increment threshold is
K is a coefficient of a threshold value,
a threshold thrust is determined for the engine,
and taking the theoretical takeoff mass.
6. A guidance system reconstruction device for dealing with a primary ignition fault is characterized in that an autonomous flight arc segment of a spacecraft is divided into a first segment, a second segment and a tail segment according to a time sequence; normally, the engine of the spacecraft works in the local time of the first section and the second section; the guidance system reconfiguration device includes:
the primary ignition fault flight time sequence module is internally provided with a primary ignition fault flight time sequence: after an ignition instruction in the first period is sent out, the state of the engine is identified within a preset time, when the identification result is that the ignition fails in sequence, the engine is ignited again, the engine is in an extrusion working mode within the preset time, and the engine is in a normal working mode after the preset time; the spacecraft in the first section adopts an iterative guidance mode; in the second section, the engine is in an extrusion working mode, and the spacecraft adopts a positive thrust vector guidance mode; finally generating flight data under the fault state;
and the data switching module is used for controlling guidance control by adopting a one-time ignition fault flight time sequence and flight data in a fault state when one-time ignition fault occurs in the actual flight process.
7. A detection and reconstruction method for a guidance system under the condition of spacecraft ignition fault is characterized by comprising the following steps:
dividing an autonomous flight arc segment of the spacecraft into a first segment, a second segment and a tail segment according to a time sequence; wherein the engines of the spacecraft operate during the first and second segments of local time;
after the engine is ignited once in the first section, the ignition state is identified, if the ignition fails once, the engine is in a squeezing working mode, otherwise, the engine is in a normal working mode; when the ignition is failed once, the ignition is restarted, and the engine is in a normal working mode; the spacecraft in the first section adopts an iterative guidance mode;
after entering the second section, the engine is in an extrusion working mode, and the spacecraft adopts a positive thrust vector guidance mode.
8. A detection and reconstruction device of a guidance system under the condition of ignition failure of a spacecraft is characterized in that an autonomous flight arc section of the spacecraft is divided into a first section, a second section and a tail section according to a time sequence; wherein the engines of the spacecraft operate during the first and second segments of local time; the guidance system detection reconstruction device includes:
the first section detection guidance module is used for identifying the primary ignition state after the engine is ignited for one time in the first section, if the primary ignition fails, the engine is in an extrusion working mode, otherwise, the engine is in a normal working mode; when the ignition is failed once, the ignition is restarted, and the engine is in a normal working mode; the spacecraft in the first section adopts an iterative guidance mode;
and after the second section of the guidance module enters the second section, the engine is in an extrusion working mode, and the spacecraft adopts a positive thrust vector guidance mode.
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