CN111692919A - Precise guidance control method for aircraft with ultra-close range - Google Patents
Precise guidance control method for aircraft with ultra-close range Download PDFInfo
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- CN111692919A CN111692919A CN202010048525.8A CN202010048525A CN111692919A CN 111692919 A CN111692919 A CN 111692919A CN 202010048525 A CN202010048525 A CN 202010048525A CN 111692919 A CN111692919 A CN 111692919A
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F41—WEAPONS
- F41G—WEAPON SIGHTS; AIMING
- F41G3/00—Aiming or laying means
- F41G3/22—Aiming or laying means for vehicle-borne armament, e.g. on aircraft
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
- F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
- F42B15/01—Arrangements thereon for guidance or control
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Abstract
The invention discloses an ultra-near range aircraft precise guidance control method, which comprises the steps of calculating the actual range of an aircraft, determining an initial firing angle according to the actual range, further selecting a nominal reference track, adjusting the pitching track of an inertial guidance section through the nominal reference track to enable the aircraft to have precise hitting capacity covering all the near range, limiting the initial amplitude of inclination correction through constraining the initial nominal inclination, and performing a series of processing such as smoothing and the like on a nominal inclination instruction to finally enable the aircraft to accurately hit a target in the ultra-near range.
Description
Technical Field
The invention relates to an accurate guidance and control method of a low-speed rotary aircraft at an ultra-close range, belonging to the field of low-cost guidance weapons.
Background
People have higher and higher requirements on the control precision and the like of the aircraft, have higher and higher requirements on other performances of the aircraft, generally have a relatively clear and applicable range, and have few relevant data on how to accurately guide the aircraft with the ultra-close range (greater than the minimum range).
The range is an important index of a rocket projectile, and the aircraft already has the range before leaving a factory, and the range refers to the maximum range and the minimum range which can be reached. These two data are also important tactical indicators of the aircraft, namely the effective range of use.
For a guidance control system under an aircraft ultra-close range condition, the following technical difficulties mainly exist:
1) if the ultra-low firing angle is adopted by the aircraft, the flight time is short, and the trajectory is too low to easily hit the ground;
2) the large firing angle is adopted, the limitation of a transmitting device is realized, the ultra-large firing angle cannot be realized, the flight elevation and the speed change are large, the peak speed is approximate to zero, on one hand, the control is difficult to ensure the instruction response precision, on the other hand, the large rudder deflection easily occurs at low speed, the rudder effect is reduced, and the control failure problem is caused;
3) when the satellite works normally, the aircraft flies for a period of time, and the initial velocity launching deviation needs to be corrected quickly.
In addition, due to the working characteristics of an engine and external interference such as wind, the aircraft generates a ballistic inclination angle jumping phenomenon in an ascending section; after the engine works, the mass of the projectile body is rapidly reduced, so that the lift-drag ratio is influenced, and the attack angle is changed violently. These uncertainties also present significant challenges to the design of guidance control systems for rocket projectiles beyond near range.
For the reasons, the inventor of the invention makes an intensive study on the precise guidance control of the existing ultra-near range aircraft, and aims to design a precise guidance control method of the ultra-near range aircraft, which can solve the problems.
Disclosure of Invention
In order to overcome the problems, the inventor of the invention carries out intensive research and designs an ultra-near-range aircraft precise guidance control method, the method comprises the steps of calculating the actual range of the aircraft, determining an initial firing angle according to the actual range, further selecting a nominal reference track, adjusting the pitch track of an inertial guidance section through the nominal reference track so that the aircraft has precise hitting capacity covering all the near range, then constraining the initial nominal inclination angle, carrying out amplitude limiting processing on the initial amplitude of inclination angle correction, carrying out a series of processing such as smoothing processing on a nominal inclination angle instruction and the like, and finally enabling the aircraft to accurately hit a target in the ultra-near range, thereby completing the invention.
Specifically, the invention aims to provide an accurate guidance control method for an aircraft with an ultra-close range, in the method, the actual range of the aircraft is calculated, and when the actual range is more than 2.88km and less than 8km, the initial firing angle of the aircraft is determined to be 70-80 degrees;
after the initial firing angle of the aircraft is determined, filling flight control parameters, geomagnetic parameters and nominal reference tracks into the aircraft;
after the aircraft launches and takes off according to the determined initial firing angle, the longitudinal overload and the lateral overload of the aircraft are respectively solved, the overload required is transmitted to the steering engine, and the steering engine steers to complete the control industry.
According to the method for controlling the precise guidance of the ultra-near range aircraft, the aircraft can be controlled to emit at a large firing angle in the ultra-near range and finally hit a target.
Drawings
FIG. 1 is a graph showing a roll rate variation obtained in a simulation experimental example according to the present invention;
fig. 2 shows a longitudinal ballistic curve obtained in a simulation experimental example according to the present invention;
fig. 3 shows a lateral ballistic curve obtained in a simulation experimental example according to the present invention;
fig. 4 shows a graph of variation in ballistic inclination angle obtained in a simulation experimental example according to the present invention;
fig. 5 shows a pitch angle variation graph obtained in a simulation experimental example according to the present invention.
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 precise guidance control method for the ultra-close range aircraft, provided by the invention, only a geomagnetic sensor, an angular velocity gyroscope and satellite navigation are adopted, and a method of launching additional trajectory fast maneuver at a medium-high firing angle is adopted, so that the rocket projectile is launched at an approximate ultra-large firing angle, and the target is precisely hit at the ultra-close range. Specifically, the method comprises the following steps:
firstly, calculating an actual range of the aircraft, and when the actual range is more than 2.88km and less than 8km, determining that the initial firing angle of the aircraft is 70-80 degrees.
If the actual range is less than 2.88km, the aircraft cannot hit due to too close distance, and if the actual range is greater than 8km, guidance control can be performed by using a conventional aircraft control method in the prior art, which is not specifically described in the application.
The actual range is the relative distance between the launching point of the aircraft and the target point, the coordinate position of the launching point of the aircraft is known, and the actual range can be determined after the longitude coordinate, the latitude coordinate and the elevation information of the target point are determined. Preferably, the launching point of the aircraft and the target are both at the same altitude position in the present application, because the distance between the launching point and the target is small, and the height difference between the launching point and the target is ignored.
The initial shooting angle refers to an included angle between an initial speed direction and a horizontal plane when the aircraft launches, and the initial shooting angle is preferably 75-78 degrees, and more preferably 78 degrees.
According to the method, after the initial firing angle of the aircraft is determined, the flight control parameters, the geomagnetic parameters and the nominal reference track can be filled into the aircraft;
the flight control parameters comprise: launching point latitude, launching point longitude, launching point elevation, target point latitude, target point longitude, target point elevation, initial shooting angle, control channel starting control time, control command amplitude limiting re-compensation coefficient, pitching guidance command amplitude limiting and yawing guidance command amplitude limiting.
The geomagnetic parameters include: the flight time and the trajectory inclination angle are adopted to meet the working requirements of the missile-borne geomagnetic device. The flight time is the time required by the aircraft from launching to landing, and is an estimated value estimated according to the target position, and the ballistic inclination angle is the azimuth angle between the aircraft speed vector and the horizontal plane and is an estimated value estimated according to the simulated flight track.
The nominal reference trajectory includes a plurality of sets of data ordered in time, wherein uncontrolled longitudinal position, altitude, and ballistic dip angle information are included in each set of data. The uncontrolled longitudinal position is the position in the X-axis direction of the launching coordinate system, the height is the position in the Y-axis direction of the launching coordinate system, the origin of the launching coordinate system is the launching point of the aircraft, the X-axis of the launching coordinate system points to the target position as the elevation of the target point is approximately equal to the elevation of the launching point, the X-axis is a horizontal straight line, and the Y-axis is a coordinate axis perpendicular to the horizontal plane.
The nominal reference track is selected from uncontrolled ejection tables under the ejection angles of 40-84 degrees according to actual ejection ranges, the uncontrolled ejection tables under the ejection angles of 40-84 degrees are existing reference tables in the field, the nominal reference track is not particularly limited in the application, and the nominal reference track is given for different ejection ranges in the uncontrolled ejection tables under the ejection angles of 40-84 degrees.
When the nominal reference track is selected according to the actual range, the correction is carried out by the following method:
when the actual range is below 4.5km, a numerical value obtained by subtracting 200 meters from the actual range value is used as a range selection nominal reference track;
when the actual range is more than 4.5km and less than 6km, a numerical value obtained by adding 200 meters to the actual range value is taken as a range selection nominal reference track;
when the actual range is more than 6km and less than 7.5km, taking a numerical value obtained by adding 400 meters to the actual range value as a range selection nominal reference track;
and when the actual range is greater than 7.5km, a value obtained by adding 600 meters to the actual range value is used as the range to select the nominal reference track.
In the method, the optimization processing is carried out on the range parameters required by table look-up, and the purpose of range optimization is only to obtain a better nominal reference trajectory, so that the obtained nominal reference trajectory is more accurate, but the actual range of the aircraft is still the actual range. Through multiple times of simulation and engineering practice of the inventor, the flight path of the aircraft after being corrected is more reasonable and is more suitable for hitting a target in a close range.
In a preferred embodiment, after the aircraft launches and takes off according to the determined initial firing angle, a longitudinal guidance strategy and a lateral guidance strategy of the aircraft are respectively provided, namely, the longitudinal overload and the lateral overload which are needed by the aircraft are solved, the overload which is needed is transmitted to a steering engine, and the steering engine performs steering operation to complete control and guidance control operation.
In the pitching direction, namely the longitudinal direction, the longitudinal flight trajectory, namely the trajectory of the aircraft can be divided into an initial stable section, a climbing section and a final guide section.
And various interferences in the flying process are eliminated by controlling the damping loop on the basis of the measured projectile body angular velocity in the initial stable section, so that the aircraft can fly stably.
The initial stable section is a flight track in a period from the time when the aircraft is launched from the barrel until the geomagnetic sensor and the inertial measurement element on the aircraft start to work and output information.
The projectile body angular velocity is measured through a satellite navigation system, and the control damping loop can be a control damping loop applied to an aircraft in the field, and is not particularly limited in the application; the geomagnetic sensor and the inertial measurement unit may be any devices known in the art, and are not particularly limited in this application.
The climbing section is a flight track within a period from the time when the satellite navigation system finishes positioning and can accurately output the current position and speed information of the aircraft to the time when the satellite navigation system enters the final guidance section, and preferably, the aircraft is controlled to enter the climbing section at the 4 th second after being transmitted by setting related parameters of the satellite navigation system.
In the climbing section, the aircraft corrects the flight path according to the filled nominal reference path, so that the aircraft has smaller position deviation when entering the final guide section.
In particular, the aircraft obtains the longitudinal overload demand in the climbing section through the following formula (I),
ayc＝kc·θf+Gc(A)
Wherein, aycIndicating longitudinal required overload, thetaf＝θc-θ，θcRepresenting the desired ballistic inclination, taken from a nominal reference trajectory, theta representing the own ballistic inclination measured by the aircraft, in particular measured in real time by a gyroscope on the aircraft, kcThe guidance gain coefficient is represented as a fixed parameter preassembled on an aircraft, and the value of the fixed parameter is generally a number between 0 and 1, GcIndicating gravity overload compensation, preferably, GcGcos θ is a compensation value that varies in real time according to the trajectory inclination angle, θfThe tilt correction initial amplitude representing the nominal reference trajectory. By selecting a reasonable nominal reference track, the actual track of the climbing section can be more suitable for hitting a target at a close distance.
The desired ballistic inclination angle θcThe deployment is performed every 0.5 seconds and the longitudinal overload is resolved from the expected ballistic dip for each deployment.
Preferably, in the inertial guidance section, when the flight path is corrected according to the filled nominal reference path, in order to prevent the aircraft from generating the singular phenomenon of the pitch attitude angle in the climbing section, the following treatment is further performed:
adjusting a desired ballistic inclination angle theta from a nominal reference trajectorycThen, first, θ is determinedcIf theta is taken as the valuecWhen the value of (A) is 76.8 or more, then theta iscAdjusting the value to 76.8, and then utilizing the theta with the value of 76.8cSolving for guidance gain factor and longitudinalOverload is required.
I.e. when thetacWhen the value of (A) is above 76.8,
θf＝76.8-θ，ayc＝kc(76.8-θ)+Gc；
if thetacWhen the value of (A) is 76.8 or less, the value of (A) is used as it is without any special treatmentcAnd solving the guidance gain coefficient and longitudinal overload.
Preferably, θ if solvedfWhen the value is more than 8, then thetafAdjusting the value to 8, and then utilizing the theta with the value of 8fCalculating longitudinal required overload;
theta if solvedfWhen the value is less than-8, then thetafAdjusting the value to-8, and then using the theta with the value of-8fCalculating longitudinal required overload;
theta if solvedfThe value is between-8 and 8, and the theta is directly utilized without special treatmentfThe calculation needs overload in the longitudinal direction.
I.e. when thetafWhen the value of (A) is 8 or more, ayc＝kc·8+Gc；
When theta isfWhen the value of (A) is more than-8, ayc＝kc·(-8)+Gc；
When theta isfHas a value of-8 to 8, ayc＝kc·θf+Gc。
More preferably, the longitudinal overload is smoothed by the following method;
if k is0If > 0, then k is0Is set to 0 and the subsequent processing is carried out, i.e. if k0If > 0, then k is0Set the value of (2) to 0 and perform the following steps (2) and (3); if k is0K is not more than 0, k is not corrected0By special treatment, directly adding k0Substituted into the following step 2 and stepProcessing in step 3;
i.e. thetad＝θf·k0，θf＝θc-θ；
I.e. thetaRepair the＝θc+θd(ii) a Through thetaRepair theAlternative thetacCalculating longitudinal required overload;
that is, when it is necessary to consider smoothing the longitudinal overload demand, the formula of the longitudinal overload demand becomes ayc＝kc·(θRepair the-θ)+GcI.e. ayc＝kc·(θc+θf·k0-θ)+Gc。
Wherein k is0Represents a smoothing coefficient, θdDenotes the smoothed nominal inclination correction value, in equation (two), t1The value of the starting control moment of the climbing section is 0, t represents flight time, and dt represents climbing correction duration, namely time for entering the climbing section.
In the application, the aircraft enters the climbing section at the 4 th second after being launched, and k is the time lapse0Will get closer to 0, when the aircraft flies for 10 seconds, t is 10, dt is 6, k is0Is-0.6667, when the aircraft flies for 20 seconds, t is 20, dt is 16, k is0The value of (B) is-0.25.
In a preferred embodiment, when the aircraft reaches the highest point position, the aircraft is switched to a final guide section, the angular velocity of the line of sight of the missile is obtained through the current position of the aircraft and the target position of filling, and a pitching guide instruction is formed by adopting a proportional guide law to control the aircraft to fly to the target. And the current position information of the aircraft is obtained by receiving satellite signals through a satellite navigation system on the aircraft. When the aircraft reaches the highest point, the trajectory inclination angle of the aircraft is zero, and the aircraft judges whether the aircraft should enter the final guide section or not by monitoring the numerical change of the trajectory inclination angle.
In the final guide section, the aircraft obtains longitudinal overload demand through the following formula (III),
aycindicating that an overload is required in the longitudinal direction,
for longitudinal line-of-sight angular velocity, GcIndicating gravity overload compensation, preferably, GcGcos θ is a compensation value that varies in real time according to the trajectory inclination, VmFor the flight speed of the aircraft, obtained by a satellite navigation system, NyThe value of the longitudinal guidance coefficient is 2-4, and the value of the longitudinal guidance coefficient is preferably 4 in the application.
Preferably, the maximum available overload of the aircraft is adopted to carry out amplitude limiting processing on the acceleration instruction in the whole process, namely when the calculated required overload is larger than the maximum available overload, the steering engine is controlled only through the maximum available overload.
The maximum available overload is solved by the following equation:
where ac _ M represents the maximum available overload, cn represents the normal force coefficient, s represents the aircraft surface area, and is a known quantity pre-installed on the aircraft, M represents the total mass of the aircraft, and is a known quantity pre-installed on the aircraft, and q represents the dynamic pressure.
The normal force coefficient is preset in the aircraft and can be read from the aircraft.
The dynamic pressure is obtained by the following formula,
q＝0.5ρv2
ρ represents the atmospheric density at the location of the aircraft, with the value being a fixed value pre-installed in the aircraft, and v represents the speed of the aircraft, provided by the satellite navigation system solution.
In a preferred embodiment, the lateral flight trajectory/trajectory of the aircraft can be divided into an initial stable segment and a final guided segment in the yaw direction, i.e. laterally.
In the initial stable section, the same as the longitudinal guidance scheme, based on the measured angular velocity of the projectile body, various interferences in the flight process are eliminated by controlling a damping loop, so that the aircraft flies stably;
when the aircraft reaches the highest point, the lateral guidance is directly transferred into the final guidance section, namely, the lateral final guidance section is entered. And (3) flying into a lateral final control section, and after control, forming a guidance instruction by adopting a proportional guidance law, and controlling the aircraft to fly to a target, specifically, obtaining lateral overload required by the following formula (IV) in the lateral final control section:
azcindicating that an overload is required laterally,
for lateral line-of-sight angular velocity, VmFor the flight speed of the aircraft, obtained by a satellite navigation system, NzThe value of the lateral guidance coefficient is 2-4, and the value of the lateral guidance coefficient is preferably 4 in the application.
In a preferred embodiment, in the lateral direction, upon entering the lateral terminal section, the magnitude of the lateral deviation angle is subjected to a calculation comparison,
when the slip angle is below 10 degrees, the start control time of the lateral channel is delayed by 3s, namely the lateral channel is started and controlled after the pitching direction enters the final guide section and outputs pitching overload for 3 s;
when the slip angle is larger than 10 degrees and smaller than 20 degrees, the start control time of the lateral channel is delayed by 1.5 s; namely, after the pitching direction enters the final guide section and outputs pitching overload for 1.5s, the lateral channel is started to control;
when the slip angle is more than 20 degrees, the lateral channel and the longitudinal channel are synchronously controlled.
The control starting refers to the moment when the control command is started to be input;
the slip angle refers to an included angle between an aircraft speed vector and an aircraft longitudinal symmetric plane, and the value is calculated and provided by a satellite navigation system.
By the adoption of the control method for delaying the lateral control starting time, the lateral direction of the aircraft can be more stable, the overall control track is more stable, and the final hit rate is higher.
In a preferred embodiment, when the initial stable section, the pitching direction and the lateral direction are only controlled by the damping loop, the two control channels only introduce the angular velocity of the aircraft to realize damping stability augmentation, and through angular velocity feedback, the elimination of the swing of the projectile body is accelerated, so that the flight stability of the aircraft is maintained.
In the application, the aircraft only controls the pitching channel and the yawing channel, the rolling channel is not controlled, and the steering engine for executing the control command in the application is an existing steering engine in the field, and the application is not particularly limited to this.
Experimental example:
the point with the relative distance of 2.88km from the launching point of the aircraft is set as a target through computer simulation, the guidance control is carried out through the precise guidance control method of the aircraft with the ultra-close range provided by the invention,
the specific control process comprises the following steps: according to the position information of the transmitting point and the position information of the target point, the actual range of the aircraft is calculated to be 2.88km, the initial firing angle of the aircraft is determined to be 78 degrees, and flight control parameters, geomagnetic parameters and nominal reference tracks are filled into the aircraft. And a value obtained by subtracting 200 meters from the actual range value is used as the range to select the nominal reference track.
Flight control parameters include: the latitude of the launching point is 38.75734 degrees, the longitude of the launching point is 105.60569 degrees, the elevation of the launching point is 1380m, the latitude of the target point is 38.971994 degrees, the longitude of the target point is 105.30979 degrees, the elevation of the target point is 1239m, the initial firing angle is 78 degrees, the control channel starting control time is 20.1s, the control instruction limiting and pitching guiding instruction limiting and limiting range is 80, the yawing guiding instruction limiting range is 80, the re-compensation coefficient is 0.2, the pitching guiding instruction limiting range is 0.2, and the yawing guiding instruction limiting range is 0.2.
And (3) selecting a nominal reference track by taking 2.68km obtained by subtracting 200 meters from the actual range value of 2.88km as the range, and obtaining the nominal reference track corresponding to the 2.68km range as shown in the following table one:
In the climbing section in the pitch direction, the aircraft obtains the longitudinal overload demand through the following formula (five), ayc＝kc·(θc+θf·k0-θ)+Gc(V)
Wherein, thetacRepresents the expected ballistic inclination angle, is taken from the table I, theta represents the measured ballistic inclination angle of the aircraft, and kcRepresenting a guidance gain coefficient, with a value of 0.8, GcIndicating gravity overload compensation, Gc＝gcosθ；
Starting control at the 4 th second after the aircraft is launched, entering a climbing section:
retrieve θ from Table onecIs 82.3 deg., since it is greater than 76.8 deg., it is adjusted to 76.8 deg., at which the measured trajectory inclination of the aircraft itself is 78 deg., then theta isf76.8-78-1.2, between-8 and 8, without additional adjustment,
t is 4, t1A value of 0, dt a value of 0, k0Can not be usedResolving, so the point is not smoothed,
the longitudinal overload requirement at this point is thus:
ayc＝1.058；
at 4.5 seconds after aircraft launch, θ is retrieved from Table IcIs 82.2 deg., since it is greater than 76.8 deg., it is adjusted to 76.8 deg., at which the measured trajectory inclination of the aircraft itself is 78 deg., then theta isf76.8-78-1.2, between-8 and 8, without additional adjustment,
the value of t is: 4.5, t1The value of dt is 0.5 to obtain k0The value of (a) is-8,
the longitudinal overload requirement at this point is thus:
ayc＝8.757；
by analogy, one theta is called every 0.5 secondcObtaining longitudinal required overload by combining the theta obtained by current detection;
the aircraft enters the terminal pilot segment at 34 seconds,
at the final stage of the pitching direction, the aircraft obtains the longitudinal overload demand through the following formula (III),
Gcvalue is GcGcos θ, longitudinal guidance coefficient NyIs 4;
when entering the lateral final section, the lateral overload demand is obtained by the following formula (four):
lateral guidance coefficient NzIs 4;
in the simulation process, the trajectory simulation period is 5ms, a rudder control instruction is generated, the instruction is transmitted to the simulator through the simulation injection serial port to start trajectory simulation, and then the projectile attitude angle and angular velocity information, the position information and the roll angle and angular velocity information output by the simulator are fed back to the missile-borne computer so as to solve the rudder control instruction of the next simulation period. And (4) according to the steps, circularly iterating until the aircraft lands, and finishing the simulation.
The simulation curves of the 2.88km distance aircraft beyond the near range are finally obtained and are shown in figures 1 to 5.
From the roll rate profile of the aircraft shown in FIG. 1, the aircraft rotational speed is finally stabilized, indicating that the aircraft is flying smoothly.
As can be seen from fig. 2, the target and the launch point are both located at a point where the altitude is zero, i.e., on the abscissa axis of fig. 2, and the aircraft takes off from a position where the rightmost abscissa axis is 0, and lands at the target point position after flying 2.88km in the direction of the abscissa axis, so that the aircraft accurately hits the target as viewed in the pitch direction.
As can be seen from FIG. 3, the aircraft takes off from the launching points with the horizontal and vertical coordinates of 0, a certain lateral deviation exists in the flight process, and the lateral deviation is basically corrected to 0 when the aircraft lands after flying for 2.88km, so that the lateral deviation of the aircraft in the landing process is extremely small when viewed from the lateral deviation direction, and the aircraft can accurately hit the target.
As can be seen in fig. 4, the initial firing angle of the aircraft during launching is 78 degrees, the angle gradually decreases, and the terminal landing angle reaches about-80 degrees during final landing, which belongs to large landing angle.
As can be seen in FIG. 5, the pitch angle of the aircraft before landing, i.e. at the 60 th to 75 th seconds, is stable, and the flight of the tail end of the aircraft is stable.
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 (9)
1. An accurate guidance control method for an aircraft with an ultra-close range is characterized in that in the method,
calculating the actual range of the aircraft, and determining that the initial firing angle of the aircraft is 70-80 degrees when the actual range is more than 2.88km and less than 8 km;
after the initial firing angle of the aircraft is determined, filling flight control parameters and a nominal reference track into the aircraft;
after the aircraft launches and takes off according to the determined initial firing angle, the longitudinal overload and the lateral overload of the aircraft are respectively solved, the overload required is transmitted to the steering engine, and the steering engine steers to complete the control industry.
2. The close-range aircraft precision guidance control method of claim 1,
when the actual range is below 4.5km, a numerical value obtained by subtracting 200 meters from the actual range value is used as a range selection nominal reference track;
when the actual range is more than 4.5km and less than 6km, a numerical value obtained by adding 200 meters to the actual range value is taken as a range selection nominal reference track;
when the actual range is more than 6km and less than 7.5km, taking a numerical value obtained by adding 400 meters to the actual range value as a range selection nominal reference track;
and when the actual range is greater than 7.5km, a value obtained by adding 600 meters to the actual range value is used as the range to select the nominal reference track.
3. The close-range aircraft precision guidance control method of claim 1,
in the inertial guidance section in the pitching direction, the aircraft obtains longitudinal overload requirement through the following formula (I),
ayc＝kc·θf+Gc(A)
Wherein, aycIndicating longitudinal required overload, thetaf＝θc-θ，θcRepresenting a desired ballistic inclination, taken from a nominal reference trajectory; theta denotes the measured own projectile of the aircraftTrack inclination angle, kcRepresenting a representative guidance gain factor, GcIndicating gravity overload compensation, thetafThe tilt correction initial amplitude representing the nominal reference trajectory.
4. The close-range aircraft precision guidance control method of claim 3,
adjusting a desired ballistic inclination angle theta from a nominal reference trajectorycThen, first, θ is determinedcIf theta is taken as the valuecIf the value of (A) is 76.8 or more, then thetacThe value is adjusted to 76.8.
5. The close-range aircraft precision guidance control method of claim 4,
theta if solvedfWhen the value is more than 8, then thetafThe value is adjusted to 8;
theta if solvedfWhen the value is less than-8, then thetafThe value is adjusted to-8.
6. The close-range aircraft precision guidance control method of claim 4,
smoothing the longitudinal overload by the following method;
step 1, calculating a smoothing coefficient through the following formula (II),
step 2, calculating a nominal inclination angle correction value after smoothing treatment through a smoothing coefficient,
i.e. thetad＝θf·k0，θf＝θc-θ；
Step 3, correcting the expected trajectory inclination angle theta through the nominal inclination correction valuecObtaining a corrected ballistic inclination angle thetaRepair the，
I.e. thetaRepair the＝θc+θd；
Through thetaRepair theSubstitution of theta in the formula (one)cCalculating longitudinal required overload;
wherein k is0Represents a smoothing coefficient, θdIndicating the smoothed nominal inclination correction, t1T denotes the time of flight and dt denotes the climb correction duration, 0.
7. The close-range aircraft precision guidance control method of claim 3,
in the pitching direction, when the aircraft reaches the highest point, the aircraft is transferred to the final guide section, and in the final guide section in the pitching direction, the aircraft obtains longitudinal overload required by the following formula (III),
aycindicating that an overload is required in the longitudinal direction,
for longitudinal line-of-sight angular velocity, GcIndicating gravity overload compensation, VmIs the flight speed of the aircraft, NyThe value of the longitudinal guidance coefficient is 2-4, and the value of the longitudinal guidance coefficient is preferably 4.
8. The close-range aircraft precision guidance control method of claim 3,
when entering the lateral final guide section, the lateral required overload is obtained through the following formula (IV),
azcindicating that an overload is required laterally,
for lateral line-of-sight angular velocity, VmIs the flight speed of the aircraft, NzThe value of the lateral guidance coefficient is 2-4, and the value of the lateral guidance coefficient is preferably 4.
9. The close-range aircraft precision guidance control method of claim 4,
when the aircraft enters the lateral terminal section,
when the slip angle is below 10 degrees, the start control time of the lateral channel is delayed by 3s, namely the lateral channel is started and controlled after the pitching direction enters the final guide section for 3 s;
when the slip angle is larger than 10 degrees and smaller than 20 degrees, the start control time of the lateral channel is delayed by 1.5 s; namely, after the pitch direction enters the final guide section for 1.5s, the lateral channel is controlled;
when the slip angle is more than 20 degrees, the lateral channel and the longitudinal channel are synchronously controlled.
Priority Applications (1)
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