CN111336871B - Vertical attack guidance method based on circuitous flight - Google Patents
Vertical attack guidance method based on circuitous flight Download PDFInfo
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
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Abstract
The invention provides a vertical attack guidance method based on circuitous flight, which is beneficial to reducing the requirement of the guidance ammunition on overload capacity in the terminal guidance stage, can bypass barriers and prevent air-raid fire and plays a role in tactical deception. Setting a virtual target point of an arc raising section by applying a drop point prediction technology, and sharing maneuvering pressure of the guided munition to the front section and the middle section of a trajectory; and then, combining the arc-dropping section virtual target point, lifting the terminal point of the proportional guidance law constrained by the falling angle to be right above the actual target point, so that the guided ammunition can continuously complete the vertical attack task, and a certain terminal guidance allowance is reserved. On one hand, the maneuvering pressure of the final guidance section is reduced, the improvement of the final guidance precision and the improvement of the reliability of engineering are facilitated, on the other hand, the influence of the fall of a launching point and a target point on a guidance scheme in the actual battlefield application can be effectively reduced, and the improvement of the robustness of the scheme is facilitated.
Description
Technical Field
The invention relates to the technical field of ammunition guidance control, in particular to a vertical attack guidance method based on circuitous flight, which is not limited to guidance rocket projectiles and can be used for guidance cannonballs, guidance aeronautical projectiles, missiles and the like.
Background
With the gradual development of radar detection technology and defense system in modern battlefield, the battle mission puts higher requirements on the penetration capability and survival capability of guided munitions on the battlefield. Taking axe cruise missiles as an example, the missiles fly roundly above the land, bypass barriers and prevent air fire, even further utilize the barriers to shield the other party to detect own equipment such as radars and the like, and can protrude into and attack targets from the optimal direction. Meanwhile, the adoption of circuitous flight is also beneficial to confusing the prejudgment of enemy on own tactical intention, and plays a role in tactical deception.
Meanwhile, for some warheads, the falling angle is one of the great factors determining the warhead power. Taking the example of killing and blasting the warhead, because the fragment flying direction of the warhead is mainly concentrated near the normal plane of the bullet axis, most fragments of the warhead fly to the near ground and the sky when the falling angle is small, and the exertion of the power of the warhead is greatly influenced. And taking the penetration warhead as an example, the large falling angle is beneficial to the penetration effect of the warhead. Under the condition of a small falling angle, the penetration warhead is easy to slip on the surface of a hard-spot target to cause slippage, so that the penetration into the target body is impossible, and the attack efficiency is completely lost.
A related research field of guidance law with a falling angle constraint and the like is also carried out, and a thesis of optimal polynomial guidance law with an impact angle constraint (Zhao, Yang book wave, Mo jin He, et. optimal polynomial guidance law with an impact angle constraint [ J ]. the university of Beijing Physician university, 2018, v.38; No.280 (06): 91-94 +100.) aims at the problem of hitting of an air-to-ground missile on a static target on the ground, designs a polynomial guidance law, meets the constraints of miss distance and the impact angle, and provides an analytic solution of the state of a guidance system. However, the designed optimal polynomial trajectory needs larger maneuvering capacity at the end of the trajectory, and once the maneuvering capacity is saturated due to disturbance in practical engineering application, the final miss distance is larger, and the optimal polynomial trajectory can only fly roundly in a small range at the end of the trajectory, so that the optimal polynomial trajectory cannot play a role in tactical fraud to a great extent. According to the 'limited time convergence sliding mode guidance law with falling angle constraint' (Lepengcheng, limited time convergence sliding mode guidance law with falling angle constraint [ J ]. modern defense technology, 2017 (06): 70-77 +162.), the limited time convergence sliding mode guidance law with falling angle constraint is improved and designed, and the condition that the guidance system state converges to the sliding mode surface, the missile eye line angle rate converges to zero and the trajectory angle meets the requirement of the terminal falling angle in the limited time is proved. However, in an actual guidance control system, the sliding mode variable structure controls the switching characteristic which is discontinuous in nature, and the system shakes. Buffeting can affect the stability of the control system, resulting in a reduction in the accuracy of the missile hit.
Disclosure of Invention
In view of the above, the invention provides a vertical attack guidance method based on circuitous flight, which is beneficial to reducing the requirement of the guidance ammunition on overload capacity in the terminal guidance stage, can bypass obstacles and prevent air-fire power, and plays a role in tactical deception.
The invention is realized by the following technical scheme:
in the stage from the stop of a guided ammunition engine to the top point of a trajectory, taking a virtual target point of an arc raising section as a predicted drop point, and adjusting the trajectory by adopting a zero-effect miss distance control strategy; the arc-rising section virtual target point is arranged behind the task target point and is 0.1 to 1 time of the range from the task target point; the zero-effect miss distance is the horizontal coordinate difference value of the uncontrolled flight landing point and the arc raising section virtual target point;
guiding the guided ammunition to the virtual target point of the arc-falling section by using a proportional guidance law with a falling angle constraint at the stage from the trajectory vertex to the virtual target point of the arc-falling section, wherein the constraint falling angle is selected to form a vertical trajectory when the projectile flies to the virtual target point of the arc-falling section; entering a final guidance stage when the flight height meets the set threshold height; the virtual target point of the arc-reducing section is arranged above the plumb of the task target point, and the horizontal projections of the virtual target point and the task target point are overlapped.
The threshold height is the sum of the altitude of the virtual target point of the arc-dropping section and the distance reserved for avoiding the divergence of the proportional guidance instruction.
The elevation of the virtual target point of the arc-dropping section can be selected within the range of 0.1-0.3 times of the height of the full trajectory.
In the second stage, according to ammunition speed and position information provided by a guided ammunition navigation system, calculating the horizontal distance between the current point and the predicted landing point of the uncontrolled flight; obtaining the horizontal distance between the current point and the virtual target point of the arc raising section; and obtaining the horizontal coordinate difference value of the uncontrolled flight landing point and the arc raising section virtual target point according to the difference value of the two horizontal distances.
The method comprises the following steps of obtaining an uncontrolled predicted drop point by carrying out numerical integration on a rigid six-degree-of-freedom ballistic kinetic equation of a guided ammunition through a missile-borne computer.
The method comprises the steps of taking a residual estimated trajectory numerical value integral initial value as a table head in a flight envelope range, taking an estimated falling speed output by an integral terminal point as a table value, generating a data table in advance through offline calculation, writing the data table into an missile-borne computer, and obtaining an uncontrolled predicted falling point through interpolation operation in a real-time flight process.
Wherein, the integration initial value meshing density in the flight envelope reaches simultaneously: the length grid width is less than 400m, the velocity grid width is less than 10m/s, and the angle grid width is less than 1 deg.
The method comprises the steps of adopting a proxy model to carry out regression and fitting on a data table generated offline to generate an algebraic formula, and obtaining an estimated falling speed through algebraic formula operation in the real-time flight process.
And in the final guidance stage, the attack target is guided according to the proportion to complete guidance.
Wherein the guided munition comprises a guided rocket projectile, a guided cannonball, a guided aerobomb or a guided missile.
Has the advantages that:
1. the invention sets the virtual target point of the arc raising section by applying the drop point prediction technology, shares the maneuvering pressure of the guided ammunition to the front section and the middle section of the trajectory, can bypass barriers and air defense fire, and can intrude and attack the target from the optimal direction. Meanwhile, the zero-effect predicted landing point of the trajectory of the arc raising section is far away from the real target, so that the prejudgment of the enemy on own tactical intention can be effectively confused, and the tactical deception effect is played; and then, combining the arc-dropping section virtual target point, lifting the terminal point of the proportional guidance law of the falling angle constraint right above the actual target point, so that the projection of the center of mass of the projectile body in the horizontal plane is close to the target at the end of the adjustment section, and the trajectory is nearly vertical, so that the guided ammunition can continuously complete the vertical attack task, and a certain final guidance allowance is reserved. On one hand, the maneuvering pressure of the final guidance section is reduced, the improvement of the final guidance precision and the improvement of the reliability of engineering are facilitated, on the other hand, the influence of the fall of a launching point and a target point on a guidance scheme in the actual battlefield application can be effectively reduced, and the improvement of the robustness of the scheme is facilitated.
2. In the third stage, the proportional guidance law with the falling angle constraint is adopted, so that the requirement on the maneuvering capability of the guided ammunition is greatly reduced, the requirement on the on-ammunition hardware is not high, the additional cost is low, and the implementation is easy.
Drawings
FIG. 1 is a flow chart of a vertical attack guidance method based on circuitous flight.
FIG. 2 is a diagram illustrating second-stage null-effect prediction according to the present invention.
FIG. 3 is a schematic diagram of a virtual target point of a third-stage arc-dropping segment according to the present invention.
FIG. 4 is a graph of simulation results of the present invention. FIG. 4(a) is a simulation (r) -trajectory diagram; FIG. 4 (b) is a diagram of ballistic inclination angles under simulation (c).
Detailed Description
The invention is described in detail below by way of example with reference to the accompanying drawings.
The invention provides a vertical attack guidance method based on circuitous flight, which is divided into four stages according to the flight trajectory of guided munitions for convenient expression, wherein the first stage is from the launching time of the guided munition to t1Time of day t1＝tmax+ δ t, where tmaxThe maximum working time of the ammunition engine under different temperature conditions, wherein deltat is a design allowance taken for ensuring the working time of the ammunition engine covered at the first stage, and is generally taken for more than 1 s; the second stage is from t1Moment to ballistic vertex; the third stage is that the altitude from the trajectory vertex to the center of mass of the guided munition is reduced to the altitude close to the virtual target point of the arc-dropping section, and the altitude is mathematically described as H being less than or equal to Hd+ δ H. Wherein Hd+ δ H is called the threshold height, H is the current flight altitude of the guided munition, HdThe elevation of the virtual target point of the arc-dropping section designed according to the tactical use requirement is selected within the range of about 0.1-0.3 times of the total trajectory height, and the delta H leaves a distance for avoiding the divergence of the proportional guidance instruction and can be 500m +/-200 m; and the fourth stage is a final guidance stage, and the final guidance is carried out by adopting classical proportion guidance or a proportion guidance law with falling angle constraint, so that the guidance error is further reduced under the condition of ensuring vertical attack.
The flow chart of the invention is shown in fig. 1, and specifically comprises the following steps:
the first stage, the flight track control is carried out;
in the second stage, an arc-rising section virtual target point is arranged at a distance of 0.1 to 1 times of the range from the target point behind the task target point; obtaining the coordinates of the predicted uncontrolled flight landing point through a rigid six-degree-of-freedom trajectory kinetic equation of the guided munition; obtaining a deflection instruction of a steering engine of the guided ammunition according to a horizontal coordinate difference value between the uncontrolled flight falling point and the virtual target point of the arc raising section, deflecting the steering engine under the action of the instruction to enable the guided ammunition to maneuver by taking the target point of the arc raising section as a predicted falling point, and entering a third stage when the trajectory inclination angle is less than 0 degrees;
in order to realize circuitous flight, an arc raising section virtual target point is set according to a tactical mission in the stage; and adjusting the trajectory according to the relation between the virtual target point of the arc raising section and the predicted uncontrolled flight landing point (the difference value of the two is the zero-effect miss distance).
The arc-lifting section virtual target point can be arranged behind the task target point by 0.1 to 1 time of range from the target point according to requirements. If the virtual target point of the arc-lifting section is in front of or behind the task target point but the distance is less than 0.1 time of the range, the circuitous flying effect is not obvious by adopting the invention. Meanwhile, the roundabout flight turning distance is larger as the distance of the arc raising section virtual target point behind the task target point is larger. However, when the virtual target point of the arc-lifting section is behind the task target point and the distance is greater than 1 time of the range, on one hand, the effect of further increasing the turn-back distance of roundabout flight is not obvious, on the other hand, a more rigorous requirement is provided for the maneuvering capability of the guided rocket, so that the waste of missile-borne resources can be caused, and the cost is greatly increased.
For a rocket projectile, the coordinates of the predicted uncontrolled flight landing point P can be obtained by numerical integration of the missile-borne computer, and the kinetic equation is as follows:
the formula (2) is a rigid six-degree-of-freedom ballistic kinetic equation of a conventional rocket projectile, and can accurately describe the mass center motion and the attitude motion of the rocket projectile. For the rocket projectile in medium and long distance, inertial navigation is mounted on the general projectile, the initial value of the state variable in the formula (2) can be obtained through the current value output by the inertial navigation, and the kinetic equation is processedIntegrating the values, wherein the integral termination condition is that y is less than or equal to ytAnd obtaining the uncontrolled predicted drop point P of the rocket projectile.
Numerical integration is carried out on the formula (2), and the obtained predicted drop point precision can meet the guidance requirement, but the time-consuming problem caused by the numerical integration cannot be completely avoided. Therefore, a residual estimated trajectory numerical value integral initial value is used as a table head in a flight envelope range, an estimated falling speed output by an integral terminal point is used as a table value, a data table is generated in advance through offline calculation and written into an missile-borne computer, and a predicted falling point is obtained through interpolation operation in the real-time flight process. Trial calculation shows that when the grid division density in the flight envelope simultaneously reaches the length grid width below 400m, the speed grid width below 10m/s and the angle grid width below 1 degree, the requirement of the guidance method can be met.
Further, in the case where there is mesh correspondence data for estimating the amount of miss, the speed of obtaining the predicted amount of miss by interpolation is related to the size of the data amount of the stored mesh. If the flight envelope range is large or the grid division is dense, the speed of interpolation calculation is very significantly slow, even slower than the speed of direct online prediction by adopting rigid six-degree-of-freedom model integration. Therefore, a polynomial function or a proxy model such as a neural network can be adopted to carry out regression and fitting on the off-line generated prior data to generate an algebraic formula, and algebraic calculation is adopted to replace interpolation calculation, so that the calculation amount is further reduced.
As shown in FIG. 2, the task target point T has the coordinate of (x)T,yT) The set virtual target point T' of the arc rising section has the coordinates of
The predicted uncontrolled flight landing point P coordinate is (x)P,yP) Wherein y isP＝yTZero effective miss distance
The angle of attack α is the control amount output at that timezExpressed as:
wherein K1Is a proportionality coefficient, is selected by numerical simulation according to specific guided ammunition products, the y coordinates of a virtual target point T' of an arc rising section and a task target point T are equal, the x coordinate is different by a target offset L,
is an important variable for realizing circuitous flight.
The third stage, firstly, selecting a virtual target point of the arc descending section above the plumb of the task target point, superposing the horizontal projections of the virtual target point and the task target point, then guiding the guided ammunition to the virtual target point of the arc descending section by using a proportional guidance law with a falling angle constraint, wherein the constraint falling angle is selected to enable the missile to be close to forming a vertical trajectory when flying to the vicinity of the virtual target point of the arc descending section, and when the missile approaches the virtual target point and the flying height of the missile meets the set threshold height, switching to the fourth stage;
in order to lift the terminal point of the falling angle constraint proportional guidance law to be right above the actual target point, the projection of the center of mass of the projectile body in the horizontal plane is close to the target at the final stage of the adjustment section, and meanwhile, the terminal guidance allowance is reserved, after the flying posture of the missile is judged to enter the third stage, the virtual target point T' of the arc-reducing section is selected to be vertically above the task target point T, and the horizontal projections of the two are overlapped, as shown in fig. 3. And guiding the guided munition to the virtual target point of the arc-descending segment by using a proportional guidance law with a falling angle constraint, wherein the constraint falling angle is selected to be close to forming a vertical trajectory when the projectile body flies to the vicinity of the virtual target point of the arc-descending segment, and the projection of the center of mass of the guided munition on the horizontal plane is close to and coincident with the mission target point.
And fourthly, guiding the attack target according to the proportion to finish guidance.
The terminal guidance can be carried out by adopting classical proportion guidance or proportion guidance law with falling angle constraint, the guidance error is further reduced under the condition of ensuring vertical attack, and the terminal guidance is not detailed because the prior art is adopted.
Furthermore, in the second stage, according to ammunition speed and position information provided by a guided ammunition navigation system, the horizontal distance between a task target point and an uncontrolled flight predicted landing point is calculated by means of numerical integration and the like, the horizontal distance between the current task target point and a virtual target point is solved by a space distance solving method, a deflection instruction of a guided ammunition steering engine is obtained by multiplying a certain scale factor according to the difference value between the task target point and the virtual target point, and the steering engine deflects under the action of the instruction, so that the guided ammunition is maneuvered by taking the virtual target point as the predicted landing point.
The navigation system related in the invention is a navigation system which adopts means such as inertial navigation, satellite navigation, ground radio positioning navigation and the like, or adopts other technical measures to obtain motion state information such as speed, position and the like in flight.
The invention relates to a vertical attack guidance method of a guided rocket projectile based on miss distance estimation, which comprises the following specific implementation steps:
wherein the content of the first and second substances,
is the zero lift drag coefficient of the projectile,
the lifting coefficient and the lateral force coefficient of the projectile body are dimensionless proportionality coefficients, and particularly, for the consideration of simplifying the calculation process,
may all take zero.
The aerodynamic moment and the aerodynamic force are similar in calculation method, and the expression is as follows:
wherein M isx、My、MzThe rolling force, the yawing moment and the pitching moment which are respectively applied to the projectile body; m isx、my、 mzThe rolling moment coefficient, the yawing moment coefficient and the pitching moment coefficient of the projectile body are obtained; q is dynamic pressure, srefIs the maximum cross-sectional area. The formula for q is:
wherein rho is the atmospheric density, the atmospheric density rho is the function of the flight height y, the example adopts the international standard atmospheric condition, and the atmospheric density is fitted into a 6 th-order polynomial function of the flight height y, and the fitting expression is as follows:
ρ＝λ1y6+λ2y5+λ3y4+λ4y3+λ5y2+λ6y+λ7 (7)
values of fitting coefficients of various items are shown in table 1:
TABLE 1 atmospheric Density fitting coefficients
C in formula (4)d0The zero-lift drag coefficient is related to the Mach number Ma of the guided rocket, is generally obtained through a wind tunnel force measurement test and is given in the form of a table function, different results are obtained for the missile types, for example, the table 2 is a table function of the zero-lift drag coefficient of a certain missile, and the calculation is carried out through linear interpolation.
TABLE 2 coefficient of zero lift drag Cd0Tabular function
| Ma | 0.4 | 0.9 | 1.15 | 1.5 | 2 | 3 | 4 |
| Cd0 | 0.4095 | 0.4843 | 0.6720 | 0.5846 | 0.4894 | 0.3707 | 0.3057 |
Wherein the Mach number Ma of the missile is the speed and the local sound velocity vsThe ratio of (A) to (B):
in the invention, the sound velocity is a function of the flying height y, and the sound velocity is fitted into a 6 th-order polynomial function of the flying height y according to the international standard atmospheric condition, wherein the fitting expression is as follows:
vs＝η1y6+η2y5+η3y4+η4y3+η5y2+η6y+η7 (9)
values of fitting coefficients of the terms are shown in table 3:
TABLE 3 Sound velocity fitting coefficient Table
Step 3, current time t0Current position v of guided rocket received from navigation system0、θ0、
x0、y0、z0、
ψ0、γ0And (3) performing numerical integration solution on the formula (2) by adopting a fixed step length numerical integration algorithm, and taking an initial value of the integral as the current time t of the missile motion information measured by the navigation system0Of motion information values, i.e.
Taking integral termination condition as y<yt；
using the formula Δ R-R' xT-xP+ L obtaining the zero-effect miss distance delta R to the virtual target point of the arc raising section, and then controlling the formula alpha according to the arc raising sectionz＝K1Calculating an attack angle instruction by multiplying the X delta R to obtain a deflection instruction of the steering engine of the guided rocket projectile, wherein the steering engine acts on the deflection instructionAnd (4) deflecting downwards to enable the missile to maneuver by taking the virtual target point of the arc raising section as a predicted drop point. Wherein x isPIs the X-axis coordinate of the integral endpoint of equation (2).
And 5, entering a third stage of the guidance scheme when the trajectory inclination angle is less than 0 DEG, and setting a virtual target point of the arc-dropping section. And selecting the virtual target point T 'of the arc-reducing section to be vertically above the actual target point T, and superposing the horizontal projections of the virtual target point T' and the actual target point T. Guiding the rocket projectile to a virtual target point of a descending arc section by using a proportional guidance law with a falling angle constraint, and expressing an instruction formula by using an attack angle as follows:
wherein, N, K2The specific values can be calculated by solving the optimal control problem off line according to specific rocket projectile products, and can also refer to a value range N belonging to [2,20 ] to obtain gain of a proportional guidance coefficient and a falling angle constraint term respectively]，K2∈[-50,-2]And (4) selecting according to trajectory simulation data. V is the relative flight speed of the rocket projectile and the virtual target of the descending arc section,
angular velocity of the projectile in line of sight, theta, for the rocket to reach the virtual targetdThe trajectory inclination angle of the rocket projectile at the current moment and the trajectory inclination angle required by the battle mission are theta in the vertical attack guidance missiondTake 90 degrees. t is tgoThe remaining flight time of the projectile to the virtual target may be estimated by dividing the distance of the projectile from the virtual target by the relative flight speed of the projectile.
When the missile approaches the virtual target point of the arc-dropping section, the trajectory inclination angle of the missile is very close to the trajectory inclination angle required by the battle mission according to the characteristic of the falling angle constraint proportion guidance law. When the flying altitude is lower than the set threshold height HdAfter + δ H, the fourth stage is entered. Wherein HdThe altitude of the arc-dropping section virtual target designed according to tactical use requirements is selected in the range of about 0.1-0.3 times of the full trajectory height, and delta H is reserved for avoiding the divergence of the proportional guidance instructionThe distance can be 500m +/-200 m.
And 6, after the fourth stage is carried out, the guided rocket projectile flies to the target according to the control instruction calculated by the proportional guidance law, and because the guided rocket projectile forms a vertical trajectory inclination angle and the projection of the center of mass of the guided rocket projectile in the horizontal plane of the target point almost coincides with the task target, a vertical falling angle is naturally formed when the guided rocket projectile hits the target.
The method is characterized in that different radio angles are uniformly adopted for emission, different offsets L, target positions (20, 0, 0) (unit: km) and threshold heights (5 km) are set, and numerical simulation is carried out on different target point positions shown in a table 4 by adopting a guidance method introduced by the patent. The simulation results are shown in fig. 4.
TABLE 4 simulation scenarios and Main index parameters
| Simulation scenario | Angle of emanation/° | L/m | Miss amount/m | Falling angle/° |
| ① | 30 | 5000 | 0.1289 | －89.4625 |
| ② | 25 | 2000 | 0.2624 | －90.5693 |
| ③ | 40 | 2000 | 0.0396 | －90.5165 |
As can be seen from the graph, the guidance method of the invention realizes the circuitous flight of the guided rocket, ensures the vertical attack on the task target and the miss distance within 0.3 m, and has higher precision. The simulation conditions and the results can be compared to obtain the theta required by the first stage of the schemetrackMay be given according to tactical technical requirements. The magnitude of the offset L determines the extent of the circuitous flight of the rocket projectile. In the simulation scene II, L is too small, circuitous flight is not obvious, and in the simulation scene I, L is relatively large, and obvious circuitous flight appears.
In summary, the above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (10)
1. A vertical attack guidance method based on circuitous flight is characterized in that in the stage from the stop of a guided ammunition engine to the top of a trajectory, a zero-effect miss control strategy is adopted to adjust the trajectory by taking a virtual target point of an arc raising section as a predicted drop point; the arc-rising section virtual target point is arranged behind the task target point and is 0.1 to 1 time of the range from the task target point; the zero-effect miss distance is a horizontal coordinate difference value of an uncontrolled prediction falling point and a virtual target point of an arc rising section;
guiding the guided ammunition to the virtual target point of the arc-falling section by using a proportional guidance law with a falling angle constraint at the stage from the trajectory vertex to the virtual target point of the arc-falling section, wherein the constraint falling angle is selected to form a vertical trajectory when the projectile flies to the virtual target point of the arc-falling section; entering a final guidance stage when the flight height meets the set threshold height; the virtual target point of the arc-reducing section is arranged above the plumb of the task target point, and the horizontal projections of the virtual target point and the task target point are overlapped.
2. The vertical attack guidance method based on detour flight according to claim 1, wherein the threshold height is the sum of the altitude of the virtual target point of the arc-dropping segment and the distance set aside for avoiding the divergence of the proportional guidance command.
3. The vertical attack guidance method based on roundabout flight according to claim 1, wherein the elevation of the arc-dropping virtual target point is selectable within the range of 0.1-0.3 times of the full ballistic height.
4. The vertical attack guidance method based on roundabout flight according to claim 1, characterized in that in the second phase, the horizontal distance between the current point and the uncontrolled predicted landing point is calculated according to ammunition speed and position information provided by a guided ammunition navigation system; obtaining the horizontal distance between the current point and the virtual target point of the arc raising section; obtaining a horizontal coordinate difference value of the uncontrolled prediction falling point and the arc raising section virtual target point according to the difference value of the two horizontal distances; wherein the second stage is from t1From time to ballistic vertex, where t1＝tmax+ δ t, where tmaxIs the maximum working time of the ammunition engine under different temperature conditions, deltat is the design margin taken for ensuring the first stage to cover the working time of the engine, and the first stage is from the launching time of the guided ammunition to t1The time of day.
5. The vertical attack guidance method based on circuitous flight according to claim 1, characterized in that the uncontrolled predicted drop point is obtained by numerical integration of rigid six-degree-of-freedom ballistic kinetic equations of guided munitions by an onboard computer.
6. The roundabout-type-flight-based vertical attack guidance method according to claim 5, wherein a residual estimated ballistic numerical value integral initial value is used as a header in a flight envelope range, an estimated falling speed output by an integral end point is used as a table value, a data table is generated in advance through offline calculation and written into a missile-borne computer, and an uncontrolled predicted falling point is obtained through interpolation operation in a real-time flight process.
7. The vertical attack guidance method based on roundabout flight according to claim 6, wherein the integration initial value meshing density in the flight envelope simultaneously achieves: the length grid width is less than 400m, the velocity grid width is less than 10m/s, and the angle grid width is less than 1 deg.
8. The vertical attack guidance method based on detour flight as claimed in claim 6, wherein the proxy model is adopted to carry out regression and fitting on the data table generated off-line to generate an algebraic formula, and the estimated falling speed is obtained through algebraic formula operation in the real-time flight process.
9. The vertical attack guidance method based on roundabout flight according to claim 1, characterized in that the terminal guidance phase guides the attack targets according to the proportion to complete guidance.
10. The vertical attack guidance method based on roundabout flight according to claim 1, wherein the guided munition comprises a guided rocket projectile, a guided cannonball, a guided missile or a guided missile.
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| CN111856924B (en) * | 2020-08-06 | 2022-12-13 | 西安睿高测控技术有限公司 | Control method of guided ammunition using relay type steering engine |
| CN114754628B (en) * | 2022-03-31 | 2023-08-04 | 南京理工大学 | Flying body trajectory control method based on drop point prediction and virtual tracking |
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