CN111290002B - Aircraft lateral deviation correction system applied to satellite signal unstable area - Google Patents
Aircraft lateral deviation correction system applied to satellite signal unstable area Download PDFInfo
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- CN111290002B CN111290002B CN201811533718.1A CN201811533718A CN111290002B CN 111290002 B CN111290002 B CN 111290002B CN 201811533718 A CN201811533718 A CN 201811533718A CN 111290002 B CN111290002 B CN 111290002B
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/13—Receivers
- G01S19/35—Constructional details or hardware or software details of the signal processing chain
- G01S19/37—Hardware or software details of the signal processing chain
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/13—Receivers
- G01S19/23—Testing, monitoring, correcting or calibrating of receiver elements
- G01S19/235—Calibration of receiver components
Abstract
The invention discloses an aircraft sideslip correction system applied to a satellite signal unstable area, which comprises a quasi-satellite guidance resolving module, a microprocessor module and a navigation ratio output module, wherein the quasi-satellite guidance resolving module is used for providing the microprocessor module with aircraft position and speed information at the current moment required by calculating the overload required by the sideslip during satellite loss; the navigation ratio output module is used for providing a real-time variable navigation ratio for the microprocessor module, and the navigation ratio output module obtains the real-time variable navigation ratio according to information such as a total range, a real-time lateral deviation distance and the like when the aircraft starts and controls, so that the guidance performance of the aircraft is improved, a target is ensured to enter a field range of a seeker when the aircraft enters terminal guidance, in addition, the navigation ratio is continuously changed in a small amplitude, large amplitude vibration of a flight track is not caused, the flight process is ensured to be stable, and the final hit precision is high.
Description
Technical Field
The invention relates to the field of guidance control of aircrafts, in particular to an aircraft sideslip correction system applied to an unstable satellite signal area.
Background
For a remote guidance aircraft, in order to improve the range of the aircraft, various measures are mostly adopted in the climbing section of a flight trajectory to enable the climbing height of the aircraft to be higher, such as rocket range extension, bottom row technology or high-power gunpowder, and the like, but the measures usually prolong the flight time of the climbing section of the aircraft, so that the starting and controlling time of the aircraft is generally set to be 50s after launching. The long flight time before starting control makes the aircraft unable to control the aircraft to fly to the target along the expected trajectory in the time, and the influence of the lateral wind, the magnus force generated by self rotation and the interference of the launching end often forces the aircraft to have a large lateral deviation distance during starting control, while even the general lateral guidance method and system can control the aircraft to fly to the target, when the aircraft enters the last guidance section, the general lateral guidance method and system often have difficulty in controlling the aircraft to make the target enter the field of view of the guidance head, and the evaluation standard of entering the field of view is as follows: and when the distance is 3km from the target, the lateral deviation is less than 600 m.
In addition, in the process of traveling of the aircraft, the aircraft is highly likely to be interfered by signals on part of road sections, satellite signals cannot be received clearly, the whole guidance control process is based on basic information provided by the satellite signals, the phenomenon of satellite loss of the high-overload remote guidance aircraft is more serious, and if the satellite signals are lost before a final guidance section, guidance cannot be performed naturally, and the lateral deviation of the aircraft is likely to be larger;
if the aircraft cannot enable the target to enter the field of view of the guide head when entering the final guide section, the aircraft cannot capture the target in the final guide section, and the target is probably missed finally; in the guidance control process of the aircraft, if guidance laws with large differences are adopted for different stages, the flight trajectory of the aircraft is inevitably vibrated greatly, and the stability of the aircraft is reduced;
for the above reasons, the present inventors have conducted intensive studies on existing aircraft control systems in order to design a new aircraft yaw correction system applied to an unstable satellite signal region, which can solve the above problems.
Disclosure of Invention
In order to overcome the problems, the inventor of the invention makes a keen study and designs an aircraft sideslip correction system applied to an unstable satellite signal area, wherein the system comprises a quasi-satellite guidance resolving module, a microprocessor module and a navigation ratio output module, and the quasi-satellite guidance resolving module is used for providing the microprocessor module with the aircraft position and speed information at the current moment required by overload for calculating the sideslip when a satellite is lost; providing real-time variable navigation ratio for the microprocessor module through the navigation ratio output module; the navigation ratio output module brings the total range, the real-time lateral deviation distance and the projection length of the connecting line between the aircraft point and the target point on the connecting line of the transmitting point and the target point into consideration of a guidance algorithm during starting and controlling of the aircraft to obtain a scientific and reasonable navigation ratio which changes in real time, so that the guidance performance of the navigation ratio is improved, a target is ensured to enter a field range of a seeker during entering the final guidance, in addition, the navigation ratio is continuously changed in a small amplitude, the large amplitude vibration of a flight track is not caused, the stable flight process is ensured, and the final hit precision is high, thereby completing the invention.
Specifically, the invention aims to provide an aircraft sideslip correction system applied to an unstable satellite signal area, which comprises a pseudo-satellite guidance resolving module 1 and a microprocessor module 2;
the quasi-satellite guidance resolving module 1 is used for providing the position and speed information of the aircraft at the current moment required by calculating the overload required by the sidesway for the microprocessor module 2 when the satellite is lost.
In the microprocessor module 2, the required sideslip overload is obtained by multiplying the navigation ratio, the flight speed of the aircraft and the angular rate of the line of sight of the missile in the sideslip direction;
preferably, the lateral bias is obtained in real time by the following formula (one):
wherein, aM sideIndicating that the yaw requires overload, N indicating the navigational ratio, V indicating the flight speed of the aircraft,
representing the angular rate of the aircraft's yaw direction line of sight.
The system further comprises a storage module 3, wherein the storage module 3 is used for storing the position and speed information of 3 continuous moments on the aircraft;
preferably, when a satellite is lost, the quasi-satellite guidance resolving module 1 retrieves the position and speed information of the continuous 3 moments from the storage module 3, and reconstructs and fits the position and speed information of the current moment according to the retrieved information;
more preferably, the position and speed information at the current time is transmitted to the microprocessor module 2 and stored in the storage module 3. 4. The system of claim 1,
the system further comprises:
an antenna 4 for receiving satellite signals,
an anti-interference module 5 connected to the antenna 4 for filtering the satellite signal,
the receiver 6 is used for receiving the satellite signals subjected to filtering processing, converting the satellite signals into navigation messages and transmitting the navigation messages to the storage module 3;
and the satellite guidance calculation module 7 is used for calling the navigation message in the storage module 3 and calculating the position and speed information at the current moment.
Preferably, the receiver 6 comprises one or more of a GPS receiver, a beidou receiver and a GLONASS receiver;
and the receivers respectively receive corresponding satellite signals.
The receiver 6 is further configured to obtain a star number corresponding to each satellite signal;
when the number of the satellites of each satellite signal is lower than a set value, the satellite signals are considered to be in a satellite loss state, and the quasi-satellite guidance resolving module 1 is controlled to start to work;
when at least one of the satellite numbers of each satellite signal is not lower than a set value, transmitting the satellite signal type information with the highest satellite number to a satellite guidance resolving module 7, and the satellite guidance resolving module 7 retrieves a navigation message corresponding to the satellite signal from the storage module 3 and resolves the position and speed information of the current moment according to the navigation message;
preferably, the position and speed information at the current time is also stored in the storage module 3 while being transferred to the microprocessor module 2.
Wherein, the system is also provided with a navigation ratio output module 8 for calculating the navigation ratio;
the navigation ratio output module 8 outputs the lateral deviation distance z of the aircraft according to the control starting timemSelects the corresponding navigation ratio N and transmits the navigation ratio N to the microprocessor module 1 in real time.
Wherein the offset distance z of the aircraft during the controlmWhen the lateral deviation is large,
when in use
When the temperature of the water is higher than the set temperature,
when in use
And xmWhen the speed is higher than 3km,
when x ismWhen the length is less than or equal to 3km, N is 4
Wherein x ismRepresenting the length, x, of the projection of the line between the point of the aircraft and the target point on the line between the emission point and the target pointmThe value of (A) is a value obtained by real-time measurement and calculation, and changes along with the position change of the aircraft; x is the number of*Representing the length, x, of the projection of the line between the aircraft point and the target point on the line between the launch point and the target point at the time of the take-off*Take a constant value during the calculation.
Wherein the offset distance z of the aircraft during the controlmIn the case of a medium lateral offset,
when x ismWhen the speed is higher than 3km,
when x ismWhen the length is less than or equal to 3km, N is 4.
Wherein the offset distance z of the aircraft during the controlmWhen the lateral deviation is small, the device can be used,
N＝4。
wherein the offset distance z of the aircraft during the takeoff controlmWhen the value is more than 1800m, the offset distance zmIs large lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is between 600m and 1800m, the lateral offset distance zmIs a medium lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is below 600m, the offset distance zmIs a small lateral deviation.
The invention has the advantages that:
according to the aircraft sidesway correction system applied to the satellite signal unstable area, provided by the invention, the radial range from a target when an aircraft starts controlling, the real-time sidesway distance and the projection length of a connecting line between the aircraft located point and a target point on the connecting line of a transmitting point and the target point are taken into consideration of a guidance algorithm, so that the navigation ratio can be adaptively adjusted according to the self sidesway condition and the flight condition of the aircraft, namely, the navigation ratio is increased when the sidesway is large, and the navigation ratio is decreased when the sidesway is small;
in addition, the invention provides an aircraft sideslip correction system applied to an unstable satellite signal area, the change of the navigation ratio is smooth and continuous, and the deflection failure of an actuating mechanism caused by the discontinuity of control quantity is avoided;
by arranging the pseudo-satellite guidance resolving module, the reconstructed satellite signal can be fitted in time when the satellite is lost, the position and speed information of the aircraft can be continuously output, the aircraft can still be controlled to stably fly under the condition that the satellite signal is lost, and the problem of uncontrollable aircraft caused by satellite loss in the navigation process is solved; .
Drawings
FIG. 1 is a logic diagram of the overall structure of a high overload resistant aircraft sideslip correction system according to a preferred embodiment of the present invention;
FIG. 2 is a schematic structural diagram of a high overload resistant antenna in a high overload resistant aircraft sideslip correction system according to a preferred embodiment of the present invention;
FIG. 3 illustrates a schematic diagram of the location of the target point, the launch point and the aircraft in accordance with a preferred embodiment of the present invention;
FIG. 4 shows a trajectory graph related to lateral deviation and a shooting distance after control activation, namely a lateral trajectory graph after control activation, in a simulation experiment of the invention;
FIG. 5 shows the trajectory profile of the present invention after the start of control and before the final guide segment, which is related to the lateral deviation and the shooting distance, i.e. the lateral trajectory profile before entering the final guide segment;
fig. 6 shows a diagram of a lateral ballistic trajectory of an aircraft before the aircraft enters a terminal section in a simulation experiment of the present invention.
Description of the reference numerals
1-quasi satellite guidance resolving module
2-microprocessor module
3-memory module
4-aerial
5-anti-interference module
6-receiver
7-satellite guidance resolving module
8-navigation ratio output module
9-holding tank
10-protective baffle
Detailed Description
The invention is explained in more detail below with reference to the figures and examples. The features and advantages of the present invention will become more apparent from the description.
The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. While the various aspects of the embodiments are presented in drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
According to the invention, an aircraft sideslip correction system applied to satellite signal instability areas is provided, as shown in fig. 1, and comprises: the system comprises a quasi-satellite guidance resolving module 1, a microprocessor module 2 and a navigation ratio output module 8; wherein the content of the first and second substances,
the microprocessor module 2 is used for calculating the required sideslip overload required by the aircraft sideslip correction;
the quasi-satellite guidance resolving module 1 is used for providing the position and speed information of the aircraft at the current moment required by calculating the overload required by the sidesway for the microprocessor module 2 when the satellite is lost.
The overload needing to be used is index data used for controlling the workload of a steering engine on the aircraft, and the steering engine on the aircraft performs steering operation according to the calculated overload needing to be used. The lateral bias requiring overload is the lateral overload that the steering engine needs to provide in order to eliminate the lateral bias.
In a preferred embodiment, in the microprocessor module 2, the sidesway overload demand is obtained by multiplying the navigation ratio, the flight speed of the aircraft and the side way direction line-of-sight angular rate;
the microprocessor module 2 is a core part of the whole satellite guidance system, in the application, the microprocessor module 2 can select a high-performance 32-bit floating point DSP chip TMS320C6713 of TI company, 8 parallel processing units are arranged in the chip, the external clock input is selected to be 50MHz, and the PLL in the processor multiplies the frequency to 200 MHz.
Preferably, the lateral bias is obtained in real time by the following formula (one):
wherein, aM sideIndicating that the yaw requires overload, N indicating the navigational ratio, V indicating the flight speed of the aircraft,
representing the angular rate of the aircraft's yaw direction line of sight. Since the aim of the application is to study the correction of the lateral deviation, the angular rate of the visual line of the bullet eyes in the lateral deviation direction is abbreviated as the angular rate of the visual line of the bullet eyes, and the lateral deviation requiring overload can also be abbreviated as overload requiring.
The flight speed of the aircraft is measured in real time by a navigation module 2 on the aircraft, the line-of-sight angular rate of the missile can be measured in real time by a sensing element or can be obtained by calculation, and generally, the normal line-of-sight angular rate of the missile and the line-of-sight angular rate of the missile in the lateral deviation direction can be obtained by aircraft position information and target point position information which are solved by satellite signals in a middle guidance section; and directly measuring by a platform laser guide head during final guide section to obtain the normal line-of-sight angular rate of the bullet eyes and the lateral deviation direction line-of-sight angular rate of the bullet eyes, wherein the normal line-of-sight angular rate and the lateral deviation direction line-of-sight angular rate are not particularly limited in the application.
The overload needing to be used is a special term in the field, and in the guidance control process of the guidance aircraft, the overload needing to be used must be firstly solved and converted into an overload instruction, and then the steering engine is controlled to steer;
in a preferred embodiment, the system further comprises a storage module 3, said storage module 3 being adapted to store position and speed information for 3 consecutive moments on the aircraft;
when receiving new position and speed information in the storage module 3, automatically covering the earliest position and speed information, so that only 3 groups of information are reserved in the storage module 3 for standby calling; a set of position and velocity information is resolved each time a satellite signal is received, referred to as a time instant, preferably 50ms apart.
Preferably, when a satellite is lost, the quasi-satellite guidance resolving module 1 retrieves the position and speed information of the continuous 3 moments from the storage module 3, and reconstructs and fits the position and speed information of the current moment according to the retrieved information;
more preferably, the position and speed information at the current moment is transmitted to the microprocessor module 2 and simultaneously stored in the storage module 3, the position and speed information is transmitted to the microprocessor module 2 so that the microprocessor module 2 can calculate overload, guidance control is provided for the aircraft, the position and speed information in the storage module 3 is updated in real time after being transmitted to the storage module 3, and the position and speed information at the next moment can be calculated conveniently by calling the information at any time.
In a preferred embodiment, as shown in fig. 1 and 2, the system further comprises:
an antenna 4 for receiving satellite signals,
the anti-interference module 5 is connected with the antenna 4 and used for filtering the satellite signals and eliminating noise interference in the satellite signals;
the receiver 6 is used for receiving the satellite signals subjected to filtering processing, converting the satellite signals into navigation messages and transmitting the navigation messages to the storage module 3; the navigation message is a message which is broadcasted to a user by a navigation satellite and used for describing the operation state parameters of the navigation satellite, and comprises system time, ephemeris, almanac, correction parameters of a satellite clock, health conditions of the navigation satellite, ionospheric delay model parameters and the like; the parameters of the navigation message provide time information for the user, and the position coordinate and the speed of the user can be calculated by utilizing the parameters of the navigation message;
and the satellite guidance calculation module 7 is used for calling the navigation message in the storage module 3 and calculating the position and the speed information of the aircraft at the current moment according to the navigation message.
Wherein, preferably, the receiver 6 comprises one or more of a GPS receiver, a beidou receiver and a GLONASS receiver; more preferably, the receiver 6 comprises a GPS receiver, a beidou receiver and a GLONASS receiver;
the receivers receive corresponding satellite signals respectively, namely the GPS receiver receives GPS satellite signals, the Beidou receiver receives Beidou satellite signals, and the GLONASS receiver receives GLONASS satellite signals.
Further preferably, the receiver 6 is further configured to obtain a star number corresponding to each satellite signal; the GPS receiver is used for acquiring the number of stars corresponding to the GPS satellite signals, the Beidou receiver is used for acquiring the number of stars corresponding to the Beidou satellite signals, and the GLONASS receiver is used for acquiring the number of stars corresponding to the GLONASS satellite signals;
when the number of the satellites of each satellite signal is lower than a set value, the satellite signals are considered to be in a satellite loss state, and the quasi-satellite guidance resolving module 1 is controlled to start to work; the set value can be set according to the actual working condition and can be 4-5, and the set value is preferably set to be 4 in the invention; the specific judgment process can be carried out in the receiver, and the star number information can also be gathered to the microprocessor module, and the microprocessor module uniformly judges and sends out a control instruction;
when at least one of the satellite numbers of the satellite signals is not lower than a set value, determining that no satellite is lost at the moment, transmitting the satellite signal type information with the highest satellite number to a satellite guidance resolving module 7, and the satellite guidance resolving module 7 retrieves a navigation message corresponding to the satellite signal from the storage module 3 and resolves the position and speed information at the current moment according to the navigation message; if the number of the stars of the Beidou satellite signals is the largest, the navigation message corresponding to the GPS satellite signals is called, and the position and speed information at the current moment is calculated according to the navigation message, and if the number of the stars of the Beidou satellite signals is the largest, the navigation message corresponding to the Beidou satellite signals is called, and the position and speed information at the current moment is calculated according to the navigation message.
Preferably, the position and speed information at the current moment is transmitted to the microprocessor module 2 and simultaneously stored in the storage module 3, the position and speed information is transmitted to the microprocessor module 2 so that the microprocessor module 2 can calculate overload, guidance control is provided for the aircraft, the position and speed information in the storage module 3 is updated in real time after being transmitted to the storage module 3, and the position and speed information at the next moment can be calculated conveniently by calling the information at any time.
In a preferred embodiment, as shown in fig. 2, the antenna 4 is in the shape of a sheet, for receiving satellite signals in case of high overload,
preferably, the antenna 4 is arranged on the outer wall of the aircraft,
more preferably, be provided with recessed holding tank 9 on the outer wall of aircraft, antenna 4 is installed in holding tank 9, holding tank 9's degree of depth size is greater than the thickness size of antenna, and is provided with guard flap 10 in antenna 4 outside.
Antenna 4 is fixed in the bottom of holding tank 9, preferably, the holding tank just can hold antenna 4, and the lateral wall of holding tank can provide the side direction spacing for antenna 4, prevents that antenna 4 from moving in a twinkling of an eye, protective baffle 10 is fixed at the top of holding tank, and inside the holding tank was arranged in completely to its self, can make the aircraft surface level and smooth basically, protective baffle external shape suits with the appearance profile of aircraft, can be the arc, also can be dull and stereotyped, protective baffle inboard and antenna 4 looks butt for fixed antenna 4 can not remove and destroy at acceleration process antenna 4.
The protective baffle 10 is used for protecting the antenna 4 on the inner side of the aircraft in the acceleration stage of the aircraft, and the antenna 4 is prevented from being damaged in the acceleration process, when the aircraft enters the guidance stage, the protective baffle 10 is separated from the aircraft, so that the antenna 4 is exposed outside, satellite signals can be conveniently received by the protective baffle 10 and the protective baffle 10 is prevented from shielding/interfering the satellite signals. Preferably, the antenna 4 is similar to a steering engine on an aircraft and needs to be started in a guidance stage, so that the protective baffle 10 and a baffle outside the steering engine of the aircraft can be synchronously controlled and synchronously separated.
The antenna 4 is in a sheet shape, that is, the antenna 4 is a sheet antenna or a sheet antenna, and the antenna may be a rectangular flat plate or an arc plate with a radian, and may be arranged according to the outline of the aircraft, in this application, the arc plate with the radian is preferred to match with the outline of the aircraft, and in the rolling process of the aircraft, the time for receiving the satellite signal by the arc plate antenna with the radian is longer, the signal strength is better,
preferably, the antenna 4 is provided with a plurality of pieces which are uniformly distributed around the aircraft, preferably, the antenna 4 is provided with 4 pieces, and in the application, the antenna 4 is preferably arranged along the circumferential direction of the rolling of the aircraft so as to ensure that the satellite signal receiving capability of the aircraft is not weakened when the aircraft rolls at a high speed.
Compared with a traditional conical antenna or annular antenna, the flaky antenna 4 in the application has the advantages that the occupied space area is small, the influence of external noise or interference is not easily caused, the integration level of the flaky antenna is higher, and the satellite signal receiving capacity is stronger.
Preferably, the sheet-shaped antenna 4 may be made of the same material as that of a conventional loop antenna or a cone antenna, and the thickness of the antenna 4 may be reduced as much as possible on the basis of ensuring stability and physical strength, so as to reduce cost;
preferably, the length of the antenna 4 is preferably 120-200 mm, the width of the antenna 4 is preferably 50-70 mm, and the thickness of the antenna is 4-8 mm.
The above equation (a) is also an overload demand calculation equation which is the most widely applied proportion guidance law in the field, but the guidance law in the prior art generally takes a fixed value, and the navigation ratio in the guidance law is adjusted by the navigation ratio output module 8 to give different overload demands.
The navigation ratio output module 8 outputs the lateral deviation distance z of the aircraft according to the control starting timemSelects the corresponding navigation ratio N and transmits the navigation ratio N to the microprocessor module 1 in real time.
In the invention, the position of the aircraft, the target position and the launching position are all regarded as one point, namely the position of the aircraft, the target point and the launching point are obtained;
the offset distance zmAs shown in fig. 3, the target point and the launching point are connected by a straight line, and the distance between the point where the aircraft is located and the straight line is the offset distance; to refer to the extent to which the aircraft is sailing off in the lateral direction.
The starting control point is a time node in the flight process of the aircraft, the aircraft flies in an uncontrolled inertia mode before the starting control point, and when the aircraft passes through the time node, a guidance control system on the aircraft starts to work, so that the flight direction of the aircraft is adjusted, the flight deviation is corrected, and the aircraft can finally hit a target.
In a preferred embodiment, the yaw distance z of the aircraft is determined as a function of the departure controlmSelects the corresponding navigation ratio N to calculate the yaw overload.
Wherein preferably the offset z of the aircraft at the time of takeoff controlmWhen the lateral deviation is large,
when in use
When the temperature of the water is higher than the set temperature,
when in use
And xmWhen the speed is higher than 3km,
when x ismWhen the length is less than or equal to 3km, N is 4
Wherein x ismRepresenting the length, x, of the projection of the line between the point of the aircraft and the target point on the line between the emission point and the target pointmThe value of (A) is a variation value obtained by real-time measurement and calculation; as the position of the aircraft changes; x is the number of*Representing the length, x, of the projection of the line between the aircraft point and the target point on the line between the launch point and the target point at the time of the take-off*Taking a constant value in the calculation process; x is the number ofm、x*And zmSee the schematic diagram shown in fig. 3;
according to the above calculation formula, when
During the process, the calculation formula of the navigation ratio N is changed, but the value of N is gradually changed along the curve all the time, no abrupt change point exists, the N is smooth and continuous, the aircraft can only provide continuous and stable overload, and larger instantaneous overload is not needed to be provided due to the abrupt change of the navigation ratio, so that the deflection failure of an actuating mechanism caused by the discontinuity of the control quantity is avoided.
In a preferred embodiment, the offset z of the aircraft is measured during the takeoff controlmIn the case of a medium lateral offset,
when x ismWhen the speed is higher than 3km,
when x ismWhen the length is less than or equal to 3km, N is 4.
At xmWhen the speed is less than or equal to 3km, the aircraft enters the terminal systemAnd the guide section and the lateral deviation are corrected to be within an allowable range, so that a guide head on the aircraft can capture a target, and the target is guided by adopting a proportional guide law, wherein the guide head can be a laser guide head and the like.
In a preferred embodiment, the offset z of the aircraft is measured during the takeoff controlmWhen the lateral deviation is small, the device can be used,
n is 4; namely, only fixed navigation ratio is needed to be used for guidance calculation when the vehicle is deflected to a small side.
In a preferred embodiment, the offset z of the aircraft is the distance of the aircraft during the takeoff controlmWhen the value is more than 1800m, the offset distance zmIs large lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is between 600m and 1800m, the lateral offset distance zmIs a medium lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is below 600m, the offset distance zmIs a small lateral deviation. And selecting corresponding navigation ratio calculation formulas according to different lateral deviation amounts, so that ammunition under different lateral deviation amounts can enable the target point to enter the field of view before the final guide section, namely the guide head captures the target.
In a preferred embodiment, said xmAnd zmAll are obtained by real-time solution, and the solution process comprises
Pre-stored longitude and latitude coordinates of the launching point and the longitude and latitude coordinates of the target point are called,
the longitude and latitude coordinates of the position of the aircraft are calculated in real time through a quasi-satellite guidance calculating module or a satellite guidance calculating module,
then x is calculated according to the real-time position relation among the position of the aircraft, the launching point and the target pointmAnd zmThe calculation relationship may be as shown in fig. 3, and a specific calculation method may be a method known in the art, which is not particularly limited in this application.
The aircraft adopts a proportion guidance law based on the gradual change of the navigation ratio of the satellite signals to guide before the final guidance segment, and can capture the laser signals during the final guidance segment, thereby switching to laser guidance at the final guidance segment and greatly improving the hit precision.
In a preferred embodiment, since the present invention is directed to a method and a system for correcting aircraft lateral deviation, during the research process, all points need to be projected onto the same plane for research, so all points involved in the present invention, such as an aircraft point, an emission point, a target point, a start control point, and the like, refer to the projected point of the point on the same horizontal plane.
In a preferred embodiment, the data transmission between the receiver 6 and the memory module, and between the memory module and the modules such as the microprocessor module 3, the pseudo-satellite guidance calculating module 1, the satellite guidance calculating module 7, etc., is performed through a data bus, and the data bus integrates an a/D converter, a D/a converter, an 422/485/232 interface, and an SPI/SCI interface, so that information can be transmitted more quickly and with less loss.
In a preferred embodiment, when a satellite is lost, the quasi-satellite guidance resolving module 1 obtains the aircraft position and speed information at the current moment through the following formula (two) and formula (three);
wherein the content of the first and second substances,
xi,yi,zirespectively representing coordinates of the aircraft in the directions of an x axis, a y axis and a z axis under a ground coordinate system at the ith moment;
respectively, the aircraft sits on the ground at the ith momentThe speed along the directions of the x axis, the y axis and the z axis is marked; by analogy, xi-1,yi-1,zi-1Respectively are coordinates of the aircraft in the directions of an x axis, a y axis and a z axis under a ground coordinate system at the moment i-1;
the speed of the aircraft in the directions of an x axis, a y axis and a z axis under a ground coordinate system at the moment i-1 respectively, namely xi-1,yi-1,zi-1Together represent the position information of the aircraft at time i-1,
collectively representing speed information of the aircraft at time i-1; x is the number ofi-2,yi-2,zi-2Respectively the coordinates of the aircraft in the directions of an x axis, a y axis and a z axis under the ground coordinate system at the moment i-2;
the speeds of the aircraft in the directions of an x axis, a y axis and a z axis under a ground coordinate system at the moment i-2 are respectively; and delta t is a satellite guidance period, and the general value of delta t is 50 ms.
Examples of the experiments
In order to verify that the aircraft sideslip correction system applied to the satellite signal unstable area can normally work when the satellite signal is interrupted, and has better sideslip correction capability compared with a traditional guidance control system under the condition of large sideslip during starting and controlling, and the hit rate can be improved, two sets of simulation verifications are adopted to respectively carry out simulation;
experiment one:
three aircrafts with the same model are launched to the same target position at the same launching place, for each aircraft, the target point is within the range, the distance between the target point and the launching point is 2 kilometers, the rotating speed of the aircraft in the advancing process is controlled to be 6-10 revolutions per second, the overload on each aircraft is over 10000g, the flight track of each aircraft is mapped, and then a graph 4 is obtained;
in the simulation process, the position and speed information of the aircraft is calculated in real time through computer simulation, and is converted into satellite signals which are transmitted to a control system of the aircraft in the form of satellite signals.
In the aircraft sidesway correction system applied to the satellite signal unstable region, which is installed in all the three aircrafts, the satellite signal is received through the antenna shown in fig. 2, the satellite signal is subjected to filtering processing through the anti-interference module, the satellite signal subjected to filtering processing is received through the receiver, and the satellite signal is converted into a navigation message and is transmitted to the storage module; when the satellite is lost, fitting and reconstructing a satellite signal through a pseudo-satellite guidance resolving module to obtain the position and speed information of the aircraft at the current moment; when the satellite is not lost, the position and speed information at the current moment is calculated by the satellite guidance calculating module, the sidesway overload is calculated by the microprocessor module, and a guidance instruction is continuously provided for the aircraft in the guidance section.
Wherein, the first aircraft does not encounter the problem of losing stars during the flight process, and finally smoothly reaches the target point, which is represented by a trajectory curve of the lost stars in fig. 4;
the second aircraft loses the satellite signal within 5s from 36s to 41s after being transmitted, and finally still successfully reaches the target point, which is represented by a lost star 1 track curve in fig. 4;
the third aircraft loses satellite signals in an area 10000m-12000m away from the launching point, and finally still successfully reaches the target point, which is represented by a lost satellite 2 track curve in fig. 4.
The experiments show that the target can still be finally hit by the satellite signals which are lost in stages under the condition that the aircraft sideslip correction system applied to the satellite signal unstable area provided by the invention is installed.
Experiment two:
setting the shooting distance between the starting control time of the aircraft and the target to be 20km and the lateral deviation to be 3 km; need to be at a distance from the targetEnsuring that the lateral deviation is within 600m at the 3km position, namely enabling the guide head to capture a target when entering a final guide section, wherein the flying speed of the aircraft is 300m/s, and the flying direction is parallel to a connecting line from a launching point to the target point; for this example, the ballistic curves in fig. 5 and fig. 6 are obtained by ballistic simulation, wherein the first scheme (solid line) represents the ballistic curve obtained by using the high overload resistant aircraft lateral deviation correction system provided by the present application, the second scheme (dotted line) represents the ballistic curve obtained by using the conventional proportional guidance algorithm,
where N is 4, the resulting ballistic curve.
FIG. 5 shows a diagram of the lateral ballistic trajectories of the aircraft after takeoff and control; fig. 6 shows lateral ballistic trajectory diagrams before the aircraft enters the final section in both scenarios, i.e., fig. 5 and 6 are not complete lateral ballistic trajectory diagrams, but are partial phase lateral ballistic trajectory diagrams.
The shooting distance in the invention refers to: calculating from the starting control time of the aircraft, and projecting the flight distance of the aircraft on the connecting line of the emission point and the target point; in the experimental example, the shooting distance when starting control is 0, and the shooting distance when just hitting a target is 20 km;
as can be seen from fig. 5, the trajectory correction condition obtained by the aircraft yaw correction system for resisting high overload provided by the present application is obviously due to the trajectory correction condition obtained by the conventional proportional guidance algorithm, and under the same large yaw condition, that is, the yaw is 3km, the aircraft yaw correction system for resisting high overload provided by the present application can effectively control the aircraft to fly to the target, whereas the conventional proportional guidance algorithm finally has a miss distance of about 200m and cannot accurately hit the target.
As can be seen from FIG. 6, the high overload resistant aircraft lateral deviation correction system provided by the present application can be used as desired at xmCorrecting the lateral deviation to be within 600m when the lateral deviation is 3km, and accurately obtaining the lateral deviation to be about 400 m; the traditional proportional guidance algorithm can not complete the task index, and is in xmAbout 610 meters of lateral deviation is still left when the lateral deviation is 3 km;
therefore, the comparison can show that the aircraft sideslip correction system applied to the satellite signal unstable area can effectively correct the sideslip and reduce the miss distance.
Through two experimental examples, the aircraft sideslip correction system applied to the satellite signal unstable area provided by the invention can normally work in the satellite signal unstable area, so that the sideslip is corrected, and a target is hit.
The present invention has been described above in connection with preferred embodiments, but these embodiments are merely exemplary and merely illustrative. On the basis of the above, the invention can be subjected to various substitutions and modifications, and the substitutions and the modifications are all within the protection scope of the invention.
Claims (4)
1. An aircraft yaw correction system for use in areas of satellite signal instability, the system comprising: the system comprises a quasi-satellite guidance resolving module (1) and a microprocessor module (2);
the quasi-satellite guidance resolving module (1) is used for providing the position and speed information of the aircraft at the current moment required by calculating the required overload of the lateral deviation for the microprocessor module (2) when the satellite is lost;
in the microprocessor module (2), the required sideslip overload is obtained by multiplying the navigation ratio, the flight speed of the aircraft and the angular rate of the line of sight of the missile in the sideslip direction;
the lateral deviation is acquired in real time by the following formula (one):
wherein, aM sideIndicating that the yaw requires overload, N indicating the navigational ratio, V indicating the flight speed of the aircraft,
representing the angular rate of the lateral deviation direction line of sight of the aircraft;
the system is also provided with a navigation ratio output module (8) for calculating a navigation ratio;
the navigation ratio output module (8) outputs the lateral deviation distance z of the aircraft during control startingmThe corresponding navigation ratio N is selected according to the size of the navigation ratio, and the navigation ratio N is transmitted to the microprocessor module (2) in real time;
offset distance z of aircraft during takeoff and controlmWhen the lateral deviation is large,
when in use
When the temperature of the water is higher than the set temperature,
when in use
And xmWhen the speed is higher than 3km,
when x ismWhen the length is less than or equal to 3km, N is 4
Wherein x ismRepresenting the length, x, of the projection of the line between the point of the aircraft and the target point on the line between the emission point and the target pointmThe value of (A) is a value obtained by real-time measurement and calculation, and changes along with the position change of the aircraft; x is the number of*Representing the length of a connecting line between the aircraft location point and the target point projected on the connecting line between the emission point and the target point at the starting and controlling time;
offset distance z of aircraft during takeoff and controlmIn the case of a medium lateral offset,
when x ismWhen the speed is higher than 3km,
when x ismWhen the length is less than or equal to 3km, N is 4;
offset distance z of aircraft during takeoff and controlmWhen the lateral deviation is small, the device can be used,
N＝4；
when starting controlOffset z of the aircraftmWhen the value is more than 1800m, the offset distance zmIs large lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is between 600m and 1800m, the lateral offset distance zmIs a medium lateral deviation;
offset distance z of aircraft when taking off controlmWhen the value is below 600m, the offset distance zmIs a small lateral deviation.
2. The system of claim 1,
the system also comprises a storage module (3), wherein the storage module (3) is used for storing the position and speed information of 3 continuous moments on the aircraft;
when a satellite is lost, the quasi-satellite guidance resolving module (1) retrieves the position and speed information of the continuous 3 moments from the storage module (3), and reconstructs and fits the position and speed information of the current moment according to the retrieved information;
the position and speed information of the current moment is transmitted to the microprocessor module (2) and is also stored in the storage module (3).
3. The system of claim 1,
the system further comprises:
an antenna (4) for receiving satellite signals,
an anti-interference module (5) connected with the antenna (4) and used for filtering the satellite signals,
the receiver (6) is used for receiving the satellite signals subjected to filtering processing, converting the satellite signals into navigation messages and transmitting the navigation messages to the storage module (3);
the satellite guidance calculation module (7) is used for calling the navigation message in the storage module (3) and calculating the position and speed information at the current time;
the receiver (6) comprises one or more of a GPS receiver, a Beidou receiver and a GLONASS receiver;
and the receivers respectively receive corresponding satellite signals.
4. The system of claim 3,
the receiver (6) is also used for acquiring the corresponding star number of each satellite signal;
when the number of the satellites of each satellite signal is lower than a set value, the satellite signals are considered to be in a satellite loss state, and a quasi-satellite guidance resolving module (1) is controlled to start working;
when at least one of the satellite numbers of the satellite signals is not lower than a set value, the satellite signal type information with the highest satellite number is transmitted to the satellite guidance resolving module (7), and the satellite guidance resolving module (7) retrieves the navigation message corresponding to the satellite signal from the storage module (3) and resolves the position and speed information of the current moment according to the navigation message.
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