CN116880526A - Guidance method of composite guidance aircraft under satellite rejection condition - Google Patents
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Abstract
The application discloses a guidance method of a composite guidance aircraft under a satellite rejection condition, which comprises the steps that in a middle guidance section of the aircraft, the aircraft is controlled to fly to a target based on satellite signals and an attitude sensing system, in the process, if the satellite rejection is encountered, the aircraft is controlled to obtain the expected acceleration based on the satellite signals applied at the last moment until the real-time satellite signals are obtained again, and after the aircraft emits for a preset time, a laser guide head is started; after a target is captured by the laser seeker, the expected acceleration of the aircraft is obtained in real time through the novel line-of-sight angle constraint guidance law, a rudder command is generated based on the expected acceleration of the aircraft, the rudder is controlled to work, the aircraft is controlled to fly towards the target, the target is impacted by the expected line-of-sight angle, the deviation caused by satellite rejection is compensated and corrected through the novel line-of-sight angle constraint guidance law, and finally the aircraft hits the target.
Description
Technical Field
The application relates to an aircraft guidance control method, in particular to a guidance method of a composite guidance aircraft under a satellite rejection condition.
Background
A composite guidance aircraft refers to an aircraft whose flight trajectory is guided in different ways at the same stage or at different stages (e.g., mid-section or end-section). The satellite and laser composite guidance aircraft utilizes a satellite navigation system to conduct middle guidance, and controls the aircraft to complete self-control flight of the middle section and approach a target area; the terminal guidance section adopts laser guidance to track and hit the hit target. Compared with the traditional single-mold guided vehicle, the satellite and laser composite guided vehicle greatly improves the hit precision of the middle and remote guided vehicle.
However, due to the vulnerability of the satellite signals, when the aircraft is in a harsh environment, the satellite signals may be blocked and altered at any time, i.e. satellite rejection occurs. Under the satellite rejection condition, the composite guidance aircraft adopting the satellite navigation system in the middle guidance section can generate huge deviation. In order to make the aircraft hit the target accurately, the error generated in the middle guide section is required to be corrected in the final guide section, however, in the prior art, an effective correction scheme is lacking for the aircraft entering the final guide section under the condition of larger error, so that a deviation generated due to satellite rejection and the like in the final guide section correction is needed greatly, and finally, the aircraft can still have higher hit precision, and further, the method has important significance in improving the anti-interference capability of the aircraft.
Based on the above problems, the present inventors have conducted intensive studies on a composite guidance method of terminal guidance by using laser guidance, so as to expect to design an aircraft control method having a strong correction capability, which can still ensure hit accuracy in the presence of satellite rejection and the like.
Disclosure of Invention
In order to overcome the problems, the inventor has conducted intensive researches and designs a guidance method of a composite guidance aircraft under the condition of satellite rejection, in the method, the aircraft obtains the expected acceleration of the aircraft based on satellite signals and an attitude sensing system in a middle guidance section, and accordingly the aircraft is controlled to fly to a target, in the process, if the satellite rejection is encountered, the expected acceleration of the aircraft is obtained based on the satellite signals applied at the last moment until the real-time satellite signals are obtained again, and after the aircraft emits for a preset time, a laser guide head is started; after a laser seeker captures a target, the expected acceleration of the aircraft is obtained in real time through a novel line-of-sight angle constraint guidance law, a rudder command is generated based on the expected acceleration of the aircraft, the rudder is controlled to work, the aircraft is controlled to fly towards the target, the target is impacted by the expected line-of-sight angle, the deviation caused by satellite rejection is compensated and corrected through the novel line-of-sight angle constraint guidance law, and finally the aircraft hits the target, so that the application is completed.
Specifically, the application aims to provide a guidance method of a composite guidance aircraft under a satellite rejection condition, which comprises the following steps:
step 1, an aircraft obtains expected acceleration of the aircraft in a middle guidance section based on satellite signals and an attitude sensing system, the aircraft is controlled to fly to a target according to the expected acceleration, and after the aircraft emits for a preset time, a laser guide head is started;
and 2, after the laser seeker captures the target, acquiring the expected acceleration of the aircraft in real time through the novel sight angle constraint guidance law, generating a rudder instruction based on the expected acceleration of the aircraft, controlling the rudder operation of a steering engine, controlling the aircraft to fly to the target, and colliding the target with the expected sight angle.
Wherein, in step 1, when the satellite is rejected, the desired acceleration of the aircraft is obtained based on the satellite signal applied at the last moment until the real-time satellite signal is retrieved.
In step 2, the novel sight angle constraint guidance law obtains the expected acceleration of the aircraft in real time through the following formula (one):
wherein a is M Indicating the desired acceleration of the aircraft,
r represents the distance between the aircraft and the target, and is calculated by ephemeris data of the aircraft and the target;
representing the component in line of sight of the relative velocity between the aircraft and the target,
the angular velocity of the view line of the bullet mesh is represented,
θ represents the view angle of the bullet, obtained by the laser guidance head,
θ d indicating that the desired angle of view is to be presented,
representing the angle between the aircraft velocity vector and the line of sight,
m represents a constant greater than 0,
s represents the surface of the sliding die,
beta represents a constant greater than 0.
The application has the beneficial effects that:
(1) The guidance method of the composite guidance aircraft under the satellite rejection condition has extremely high universality, and can correct the flight track under the condition that the middle guidance section has deviation, so that the aircraft finally hits the target, and adverse effects of external factors such as satellite signal interference and the like on hit precision are eliminated from the result;
(2) According to the method for guiding the composite guided vehicle under the satellite rejection condition, the expected acceleration of the vehicle is obtained in real time through the novel sight angle constraint guidance law, so that the vehicle can hit a target at an expected sight angle, and the vehicle can hit a weak link of the target at a specific angle at the expected sight angle, so that the optimal hit effect is obtained;
(3) According to the guidance method of the composite guidance aircraft under the satellite rejection condition, which is provided by the application, the problem that the aircraft has larger deviation at the beginning of terminal guidance can be solved pertinently, and the problem that the aircraft is unstable and is not aligned due to overload exceeding the actual working range of a steering engine in short-term use is avoided; the novel sight angle constraint guidance law provides continuous smooth overload for use through the sliding die surface; correspondingly, the traditional proportional guidance law is a guidance method for making the rotation angular velocity of the velocity vector of the aircraft proportional to the rotation angular velocity of the sight, if the deviation of the initial stage of terminal guidance is large, the situation that the required normal overload is smaller than the available normal overload is necessarily caused in order to achieve the constraint effect, and finally the aircraft is unstable and is misaligned.
Drawings
FIG. 1 shows a schematic overall logic diagram of a method of guidance of a composite guided vehicle under satellite rejection conditions of the present application;
FIG. 2 shows a schematic representation of a flight trajectory of an aircraft in an embodiment;
fig. 3 shows a schematic view of the line of sight angle of an aircraft over time in an embodiment.
Detailed Description
The application is further described in detail below by means of the figures and examples. The features and advantages of the present application will become more apparent from the description.
The word "exemplary" is used 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. Although various aspects of the embodiments are illustrated in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
According to the method for guiding the composite guided vehicle under the satellite refusal condition provided by the application, as shown in fig. 1, the method comprises the following steps:
step 1, an aircraft obtains expected acceleration of the aircraft in a middle guidance section based on satellite signals and an attitude sensing system, the aircraft is controlled to fly to a target according to the expected acceleration, and after the aircraft emits for a preset time, a laser guide head is started; before launching the aircraft, estimating a time of flight of the aircraft based on the target location information, the aircraft launch location, and the aircraft flight speed, and setting the aircraft to start the laser guide head into the terminal guidance phase about 4 km from the target.
In the application, before the aircraft is launched, ephemeris information of the aircraft and a target, preset expected flight time and expected line of sight angle information are set in the aircraft, and after the aircraft is launched, the aircraft starts program setting before the highest point and enters a middle guidance section.
In the middle guidance section, when the satellite is rejected, the desired acceleration of the aircraft is obtained based on the satellite signal applied at the last moment until the real-time satellite signal is recovered.
In the application, the satellite rejection means that the aircraft cannot receive signals from satellites and can not update the state information of the aircraft based on real-time satellite signals;
preferably, in the middle guidance section, the aircraft is continuously under the satellite rejection state within 16 seconds, the generated direction error is relatively small, the laser signal diffusely reflected at the target can be captured in time after the laser guidance head starts working, and the aircraft can finally hit the target by correcting the novel sight angle constraint guidance law in the subsequent terminal guidance section.
In the application, the attitude sensing system adopts the geomagnetic sensor which is responsible for measuring the rolling angle and the rotating speed of the aircraft, and compared with the traditional space orientation gyro, the geomagnetic sensor is not affected by the constraint of the frame angle, so that the attitude sensing system can work normally under the condition of high overload.
And 2, after the laser seeker captures the target, acquiring the expected acceleration of the aircraft in real time through the novel sight angle constraint guidance law, generating a rudder instruction based on the expected acceleration of the aircraft, controlling the rudder operation of a steering engine, controlling the aircraft to fly to the target, and colliding the target with the expected sight angle.
According to the application, the steering engine servo system is arranged on the aircraft, and generates the steering deflection command according to the expected acceleration of the aircraft, so as to control steering engine deflection, and meanwhile, the steering deflection state can be fed back to the microprocessor for resolving the expected acceleration, so that the steering deflection command is subjected to feedforward compensation, and the steering precision is improved.
In the application, the strapdown laser guide head is selected as the laser guide head and is responsible for measuring the sight angle, and compared with the traditional platform guide head, the strapdown guide head does not need to be arranged on a platform and can be directly and fixedly connected on an aircraft, thereby saving the loading space of the aircraft body and having good working performance under the condition of high overload.
After the laser guide head captures the laser echo reflected by the target, the target can be considered as captured, and the target enters the terminal guidance section.
In step 2, the novel view angle constraint guidance law obtains the expected acceleration of the aircraft in real time through the following formula (one):
wherein a is M Indicating the desired acceleration of the aircraft,
r represents the distance between the aircraft and the target, and is calculated by ephemeris data of the aircraft and the target;
representing an aircraft and a targetA component of the relative velocity between the two in line of sight, which can be obtained from the satellite signal;
in the present application, the r and r are normally obtained by satellite signals still during the terminal guidance phaseWhen satellite rejection is encountered, then r and +.>In the present application, the laser signal irradiating the target is an intermittent signal, and can be emitted by a third party, namely, at intervals of a preset time t 1 Irradiating once, each irradiation duration being constant, the duration and interval time t of the irradiation being pre-stored in the aircraft 1 The time from the ith emission of laser light to the reception of the diffuse reflected laser light is t ki The speed of light is c, the distance between the laser irradiator and the target is s, and the speed of the aircraft is V M ，r i Representing the distance between the aircraft and the target measured the i-th time, then
r i ＝ct ki -s
r i+1 ＝ct k(i+1) -s
Because of the small time interval, the component in the line of sight direction of the relative velocity of the aircraft and the target measured at the ith time is considered as:
preferably, an active laser irradiator can also be mounted on the laser guidance head of the aircraft, by means of which the target is actively irradiated, so that the corresponding r and r are obtained
Indicating the eye movementThe angular velocity of the line of sight is obtained by: />
θ represents the view angle of the bullet, obtained by the laser guidance head,
θ d indicating that the desired angle of view is to be presented,
represents the included angle between the velocity vector of the aircraft and the line of sight, simply referred to as the leading angle of the velocity vector of the aircraft,
m represents a constant greater than 0, preferably of value 3000,
s represents the surface of the sliding die,
beta represents a constant greater than 0, preferably taking a value of 1.5.
In a preferred embodiment, the slip-form surface S is obtained by the following formula (two):
wherein t represents the time elapsed after the start of terminal guidance, and the time when the laser guide head captures the laser signal diffusely reflected from the target is taken as the timing starting point.
Preferably, the method comprises the steps of,by subjecting +.>Obtaining a derivative:
wherein a' M Indicating the desired acceleration of the aircraft at the previous moment, when the aircraft has just entered the terminal guidance zone, the new line of sight angle is first activatedThe a 'is when the beam guidance law obtains the desired acceleration of the aircraft' M The value of (2) is the expected acceleration at the last moment of the middle guidance section. In the aircraft of the application, the middle guidance section carries out guidance control on the aircraft through the traditional proportional guidance law.
V M Indicating the speed of the aircraft and,
indicate +.>When the new line of sight angle constraint guidance law is enabled for the first time to obtain the desired acceleration of the aircraft just after entering the terminal guidance zone, the +.>The value of (2) is 0.
Examples
Three identical aircrafts are arranged, the initial speed of each aircrafts is 800m/s, the target is 30km away from the launching position of each aircrafts, and the environmental disturbance suffered by the three aircrafts is completely consistent. The three aircrafts all cast the laser fairing 110 seconds after the emission, the target is captured by the laser guidance head, namely, the simulation system provides simulated laser signals for the three aircrafts 110 seconds after the emission.
The first aircraft belongs to the normal condition, namely the satellite signals are given to the first aircraft in real time in the middle guidance section and the final guidance section of the aircraft, and the situation that the satellite refuses exists is avoided, and the control process of the aircraft comprises the following steps:
step 1, the aircraft obtains the expected acceleration of the aircraft in the middle guidance section based on satellite signals and an attitude sensing system, the aircraft is controlled to fly to a target according to the expected acceleration, and after the aircraft is transmitted 110, a laser guide head is started to receive a simulated laser signal; and 2, after the laser seeker captures the target, namely after receiving the simulated laser signal, acquiring the expected acceleration of the aircraft in real time through the novel sight angle constraint guidance law, generating a rudder instruction based on the expected acceleration of the aircraft, controlling the steering engine to rudder, controlling the aircraft to fly to the target, and colliding the target with the expected sight angle.
The novel sight angle constraint guidance law obtains the expected acceleration of the aircraft in real time through the following formula (I):
wherein a is M Indicating the desired acceleration of the aircraft,
r represents the distance between the aircraft and the target;
representing the component in line of sight of the relative velocity between the aircraft and the target,
the angular velocity of the view line of the bullet mesh is represented,
θ represents the view angle of the bullet eye,
θ d indicating the expected line of sight angle, taking the value of 70 degrees,
representing the angle between the aircraft velocity vector and the line of sight,
m represents a constant, takes a value of 3000,
s represents the surface of the sliding die,
beta represents a constant and has a value of 1.5.
The slip-form surface S is obtained by the following formula (two):
wherein t represents the time counted from 110 seconds after the aircraft is launched;
by subjecting +.>Obtaining a derivative:
wherein a' M Indicating a desired acceleration of the aircraft at a previous moment;
V M indicating the speed of the aircraft and,
indicate +.>The initial value is 0.
Eventually the first aircraft steadily hits the target at the desired line of sight angle.
The second aircraft is in a modified condition, i.e. satellite rejection is present in both the mid-guided and end-guided segments of the aircraft, satellite rejection is present in the 90-100 and 105-110 seconds, and the simulated satellite signal is not provided to the aircraft during this period, the control process of the second aircraft comprising the steps of:
step 1, the aircraft obtains the expected acceleration of the aircraft in the middle guidance section based on satellite signals and an attitude sensing system, the aircraft is controlled to fly to a target according to the expected acceleration, and after the aircraft is transmitted 110, a laser guide head is started to receive a simulated laser signal; in the process, when the satellite is rejected, the expected acceleration of the aircraft is obtained based on the satellite signal applied at the last moment until the real-time satellite signal is obtained again;
and 2, after the laser seeker captures the target, namely after receiving the simulated laser signal, acquiring the expected acceleration of the aircraft in real time through the novel sight angle constraint guidance law, generating a rudder instruction based on the expected acceleration of the aircraft, controlling the steering engine to rudder, controlling the aircraft to fly to the target, and colliding the target with the expected sight angle.
The novel sight angle constraint guidance law obtains the expected acceleration of the aircraft in real time through the following formula (I):
wherein a is M Indicating the desired acceleration of the aircraft,
r represents the distance between the aircraft and the target;
representing the component in line of sight of the relative velocity between the aircraft and the target,
the angular velocity of the view line of the bullet mesh is represented,
θ represents the view angle of the bullet eye,
θ d indicating the expected line of sight angle, taking the value of 70 degrees,
representing the angle between the aircraft velocity vector and the line of sight,
m represents a constant, takes a value of 3000,
s represents the surface of the sliding die,
beta represents a constant and has a value of 1.5.
When the satellite is rejected,
r i ＝ct ki -s
r i+1 ＝ct k(i+1) -s
r i representing the distance between the aircraft and the target measured the i-th time,
representing the component in the direction of line of sight of the i-th measured relative velocity of the aircraft to the target:
r i+1 representing the distance between the aircraft and the target measured at the i+1st time,
c represents the speed of light at which the light is emitted,
t ki indicating the time elapsed from the ith laser light emission to the reception of the diffusely reflected laser light of that time,
s denotes the distance between the laser illuminator and the target.
The slip-form surface S is obtained by the following formula (two):
wherein t represents the time counted from 110 seconds after the aircraft is launched;
by subjecting +.>Obtaining a derivative:
wherein a' M Indicating a desired acceleration of the aircraft at a previous moment;
V M indicating the speed of the aircraft and,
indicate +.>The initial value is 0.
Eventually the second aircraft can still hit the target at the desired line of sight angle.
A third aircraft belonging to an abnormal situation, i.e. in a middle guidance section of the aircraft, a satellite rejection situation, i.e. in the 90 th to 100 th seconds, is present, which period does not provide a simulated satellite signal into the aircraft, the control procedure of the third aircraft comprising the steps of:
step 1, the aircraft obtains the expected acceleration of the aircraft in the middle guidance section based on satellite signals and an attitude sensing system, the aircraft is controlled to fly to a target according to the expected acceleration, and after the aircraft is transmitted 110, a laser guide head is started to receive a simulated laser signal; in the process, when the satellite is rejected, the expected acceleration of the aircraft is obtained based on the satellite signal applied at the last moment until the real-time satellite signal is obtained again;
step 2, after the laser guide head captures the target, namely after receiving the simulated laser signal, the target is guided by the proportion to obtain the guiding rateTo obtain in real time a desired acceleration of the aircraft; n is 4, V and +.>Are obtained by satellite signals;
eventually the third aircraft fails to hit the target.
The flight track of the three aircrafts is shown in fig. 2, the line-of-sight angle change curve chart of the three aircrafts with time is shown in fig. 3, and according to the result shown in fig. 2, the guidance method of the composite guidance aircrafts under the satellite rejection condition can enable the aircrafts to hit targets normally; under abnormal conditions, namely under satellite rejection conditions, because the aircraft generates deviation in the middle guidance section, the aircraft is controlled by adopting the proportional guidance rate, so that the aircraft finally cannot accurately hit the target; under the correction condition, namely, the guidance method of the composite guidance aircraft under the satellite rejection condition is adopted to correct the deviation generated by the aircraft under the satellite rejection condition, and finally, the target is accurately hit. As can be seen from fig. 3, the guidance control of the aircraft by the guidance method of the composite guidance aircraft under the satellite rejection condition can enable the aircraft to hit the target at the desired line of sight angle, regardless of whether the satellite rejection condition exists.
The application has been described above in connection with preferred embodiments, which are, however, exemplary only and for illustrative purposes. On this basis, the application can be subjected to various substitutions and improvements, and all fall within the protection scope of the application.
Claims (7)
1. The method for guiding the composite guided aircraft under the satellite rejection condition is characterized by comprising the following steps of:
step 1, an aircraft obtains expected acceleration of the aircraft in a middle guidance section based on satellite signals and an attitude sensing system, the aircraft is controlled to fly to a target according to the expected acceleration, and after the aircraft emits for a preset time, a laser guide head is started;
and 2, after the laser seeker captures the target, acquiring the expected acceleration of the aircraft in real time through the novel sight angle constraint guidance law, generating a rudder instruction based on the expected acceleration of the aircraft, controlling the rudder operation of a steering engine, controlling the aircraft to fly to the target, and colliding the target with the expected sight angle.
2. The method of guidance of a composite guided vehicle under satellite rejection conditions according to claim 1, wherein,
in step 1, when the satellite is rejected, the desired acceleration of the aircraft is obtained based on the satellite signal applied at the last moment, until the real-time satellite signal is retrieved.
3. The method of guidance of a composite guided vehicle under satellite rejection conditions according to claim 1, wherein,
in step 2, the novel line-of-sight angle constraint guidance law obtains the desired acceleration of the aircraft in real time by the following equation (one):
wherein a is M Indicating the desired acceleration of the aircraft,
r denotes the distance between the aircraft and the target,
representing the component in line of sight of the relative velocity between the aircraft and the target,
the angular velocity of the view line of the bullet mesh is represented,
θ represents the view angle of the bullet eye,
θ d indicating that the desired angle of view is to be presented,
representing the angle between the aircraft velocity vector and the line of sight,
m represents a constant greater than 0,
s represents the surface of the sliding die,
beta represents a constant greater than 0.
4. A method of guidance of a composite guided vehicle under satellite rejection conditions according to claim 3,
the slip-form surface S is obtained by the following formula (two):
where t represents the time elapsed after the start of terminal guidance.
5. A method of guidance of a composite guided vehicle under satellite rejection conditions according to claim 3,
by subjecting +.>Obtaining a derivative:
wherein a' M Indicating the desired acceleration of the aircraft at the last moment,
V M indicating the speed of the aircraft and,
indicate +.>
6. The method of guidance of a composite guided vehicle under satellite rejection conditions according to claim 1, wherein,
in step 2, when the satellite is rejected, the distance r between the aircraft and the target is obtained by the following formula (four):
r i ＝ct ki s (fourth)
Wherein r is i Representing the distance between the aircraft and the target measured the i-th time,
c represents the speed of light at which the light is emitted,
t ki indicating from the ith laser emission to the th laser emissionThe time that it takes to receive this time of diffuse reflection of the laser light,
s denotes the distance between the laser illuminator and the target.
7. The method of guidance for a composite guided vehicle under satellite outage conditions of claim 6,
in step 2, when the satellite is rejected, the component in line of sight of the relative velocity between the aircraft and the target is obtained by the following formula (five)
Wherein r is i Representing the distance between the aircraft and the target measured the i-th time,
t 1 the time interval between two laser shots is indicated,
r i+1 ＝ct k(i+1) -s
the time from the (i+1) th time of emitting the laser light to the time of receiving the diffusely reflected laser light is t k(i+1) ，
r i+1 Representing the distance between the aircraft and the target measured at the i+1st time.
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