CN116661495A - Near-range deceleration control method for aircraft - Google Patents
Near-range deceleration control method for aircraft Download PDFInfo
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
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- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Abstract
The application discloses a near-range deceleration control method of an aircraft, wherein the control of the aircraft is divided into a longitudinal control part and a lateral control part, the lateral control part is a main deceleration means, and the longitudinal control part plays roles of guidance and auxiliary lateral deceleration; and further obtaining a final pitching overload instruction and a final yawing overload instruction, transmitting the final pitching overload instruction and the final yawing overload instruction to the steering engine, and controlling the aircraft to fly to the target through the steering engine.
Description
Technical Field
The application relates to a control method of an aircraft, in particular to a near-range deceleration control method of the aircraft.
Background
During the flight of an aircraft, the speed and maneuverability are often contradictory, and during the end guidance of a near range, the aircraft is required to have better turning maneuverability due to the shortened distance; in order to adapt to various ranges, the power of the aircraft is usually designed in a redundant way, which can lead to the too high flying speed of the aircraft in a near-range task; in conventional guidance flight, the proportional guidance method is a common guidance law, and determines that overload of the aircraft is in direct proportion to the speed and the target line-of-sight angular speed; thus, excessive speeds may cause the overload acceleration required for guidance to be large. The maximum overload acceleration which can be generated by the aircraft is usually fixed, and overload required by guidance is not suitable to exceed the overload capacity of the aircraft for a long time, otherwise, the guidance precision and the flight stability are greatly affected; to reduce the overload required for maneuver requires a significant reduction in speed.
The speed of an aircraft is related to thrust and drag. Because the engine has shorter working time, the aircraft enters a passive section, namely an unpowered flight section, soon after being launched. At this time, the resistance and the height variation are the main factors determining the speed. Typically the height variation depends on the trajectory, the regularity of which is relatively fixed. The drag of an aircraft is related to its aerodynamic profile and attitude, the latter being directly controllable.
In practical work, in order to promote universality of the aircraft, the aircraft is ensured to be applied to strike a short-range target under special conditions, but no control scheme for meeting the short-range target by performing deceleration control on the aircraft exists at present.
For the above reasons, the present inventors have made intensive studies on a near-range deceleration control method of an aircraft, in hopes of designing a near-range deceleration control method of an aircraft capable of solving the above-mentioned problems.
Disclosure of Invention
In order to overcome the problems, the inventor has conducted intensive researches and designs a near-range deceleration control method of an aircraft, wherein the control of the aircraft is divided into a longitudinal control part and a lateral control part, the lateral control part is a main deceleration means, and the longitudinal control part plays roles of guidance and auxiliary lateral deceleration; and a final pitching overload instruction and a final yawing overload instruction are obtained and transmitted to the steering engine, and the steering engine is used for executing control to control the aircraft to fly to the target, so that the application is completed.
Specifically, the application aims to provide a near-range deceleration control method of an aircraft, which comprises the following steps of:
step 1, loading program parameters into an aircraft before the aircraft is launched;
step 2, acquiring a final pitching overload instruction of the aircraft in real time after the aircraft is launched;
after the engine of the aircraft is in a low thrust stage or the engine of the aircraft is shut down, starting to acquire a final yaw overload instruction in real time;
and step 3, transmitting the final pitching overload instruction or the final pitching overload instruction and the final yawing overload instruction to the steering engine in real time, controlling steering operation of the steering engine according to the final pitching overload instruction or the final yawing overload instruction, and controlling the aircraft to fly to a target.
In the step 2, a pitch overload instruction is obtained first, and then the pitch overload instruction is limited in a specific interval, so as to obtain a final pitch overload instruction.
Wherein the pitch overload command is obtained in real time by the following formula (one):
wherein a is yc Representing a pitch overload command,
N p and N q All of which represent the guidance coefficient,
V r indicating the relative speed of the aircraft to the target,
q y representing the vertical line of sight angle of the aircraft to the target,
a first time derivative representing the vertical line of sight angle of the aircraft to the target,
q f indicating that the corrected end-of-line angle is falling,
t go representing the estimated time of flight remaining in the vehicle,
g represents the acceleration of gravity and,
θ represents the ballistic tilt angle.
Wherein the guidance factor N p And N q Obtained by the following formula (II):
wherein n represents a guidance parameter;
preferably, the guidance parameter n=n when the engine of the aircraft is in operation, i.e. when the aircraft is in the active section 0 ，
When the aircraft is in the passive section after the engine of the aircraft is shut down, i.e. after the active section is finished, the guidance parameter n is obtained by the following formula (III):
wherein r represents the distance of the aircraft from the target,
r 0 indicating the distance of the aircraft from the target at the end of the active segment,
r 1 representing the closest critical distance between the aircraft and the target,
n 1 indicating the final time guidance factor,
n 0 indicating the initial time derivative.
Wherein the corrected tip landing angle q when the aircraft is in the active section f Is given by q f0 ；
After the end of the active section of the aircraft, before the aircraft reaches the highest point, the corrected end falling angle q f Obtained by the following formula (IV):
q f ＝q f0 +Δq f (IV)
Wherein q f0 Indicating the desired tip landing angle at hit;
Δq f a correction deviation indicating the end drop angle;
after the aircraft passes the highest point, the corrected tip landing angle q f Is given by q f0 ；
Preferably, the corrected deviation Δq of the end falling angle f Obtained by the following substeps:
sub-step 1, obtaining a preliminary correction deviation by the following formula (five):
wherein Δq' f Representing the preliminary correction deviation;
V max representing the actual maximum speed at the end of the active segment;
V max0 indicating the theoretical maximum speed after the end of the active segment;
substep 2, incorporating said preliminary correction deviation Δq' f Limiting to a predetermined interval to obtain a corrected deviation deltaq of the end falling angle f ；
Preferably, in sub-step 2, the predetermined interval is [ -6,6].
Wherein the specific interval is [ -a y max ,a y max ]，
Wherein a is y max Indicating a maximum allowable overload;
preferably, the maximum allowable overload a y max Obtained by the following formula (six):
a y max ＝QS ref C nb max (six)
Wherein Q represents the dynamic pressure of air,
S ref indicating the reference area of the machine body,
C nb max representing the maximum normal force coefficient.
Wherein in said step 2, said final yaw overload command is obtained by the following formula (seven):
ψ c ＝ψ c0 K 2 +ψ cold (1-K 2 ) (seven)
Wherein, psi is c Representing a final yaw overload command;
ψ c0 representing a yaw angle command;
k represents a smoothing coefficient;
ψ cold indicating a yaw angle command to shift forward.
Wherein when the aircraft is in a left-offset shapeIn the state, the yaw angle command psi c0 ＝ψ v +β c ；
The yaw angle command ψ when the aircraft is in a right yaw state c0 ＝ψ v -β c ；
When the aircraft is in a left-offset state, if the lateral displacement z of the aircraft meets z < -delta z max The aircraft is switched to a right-hand state;
when the aircraft is in the right-offset state, if the lateral displacement z of the aircraft meets z > delta z max The aircraft is switched to a left-hand state;
wherein, psi is v Indicating the current ballistic deflection of the aircraft,
β c representing a sideslip command;
Δz max representing a maximum value of a deceleration single sideslip displacement of the aircraft;
preferably, the sideslip instruction is obtained by the sub-steps of:
a, obtaining a preliminary sideslip instruction through the following formula (eight):
wherein beta' c Representing a preliminary sideslip command;
t represents the flight time of the aircraft;
substep b, executing the preliminary sideslip command beta c ' restricted to a predetermined section, thereby obtaining side slip command beta c ；
Preferably, in sub-step b, the predetermined interval is [12, 16].
Wherein the smoothing coefficient K is obtained by the following formula (nine):
wherein e represents the base of the natural logarithm,
T S the time constant is represented by a time constant,
t represents the current time of flight and,
t 0 indicating the time at which the last slip commutation was performed.
In the step 2, after a final yaw overload instruction is obtained through a formula (seventh), whether the aircraft meets a lateral deceleration closing condition is judged in real time, and when the lateral deceleration closing condition is met, the final yaw overload instruction is obtained through a proportional guidance law;
preferably, when any one of the following three conditions is satisfied, that is, the lateral deceleration closing condition is satisfied;
condition one: the distance r between the aircraft and the target satisfies r < r min ；
Condition II: the aircraft has crossed the highest point and the current real-time estimated final velocity V f Satisfy V f ＜V f0 ；
And (3) a third condition: the absolute value of the lateral displacement of the aircraft |z| satisfies |z| > z max ；
Wherein r is min A critical target distance representing a closing lateral deceleration;
V f0 indicating a predetermined end speed;
z max the maximum threshold value of the lateral deviation of the flight trajectory.
The application has the beneficial effects that:
(1) According to the near-range deceleration control method of the aircraft, which is provided by the application, the near-range deceleration control method is mainly applied to the aircraft with long-range flight capability, so that the aircraft can actively decelerate to improve terminal guidance precision when the aircraft executes a short-range task, and finally can accurately hit a near-range target;
(2) According to the near-range deceleration control method for the aircraft, which is provided by the application, the thrust control of the engine is not required, the method is suitable for the aircraft pushed by the solid engine, and the aircraft does not need to be additionally provided with an additional deceleration actuating mechanism, so that the universality is strong;
(3) According to the near-range deceleration control method of the aircraft, provided by the application, the longitudinal direction and the transverse direction can be simultaneously controlled, and the longitudinal channel adjusts the trajectory height and the falling angle by adjusting the guidance law coefficient so as to influence the speed; the transverse channel adopts a mode of laterally sliding alternately, so that not only can the resistance be increased and a remarkable deceleration effect be provided, but also the trajectory can be basically stabilized, and the terminal guidance is not influenced;
(4) According to the near-range deceleration control method of the aircraft, which is provided by the application, a perfect condition mechanism is provided, the time for starting and ending deceleration can be accurately mastered, and the influence of the deceleration process on the flight stability is avoided;
(5) According to the near-range deceleration control method for the aircraft, provided by the application, the protection measures are arranged on the instructions generated in each link, so that the size of the instructions can be ensured to be within the executable range of the aircraft, and smooth transition measures are arranged when the instructions are switched, so that the system is prevented from losing stability due to oversized instructions or abrupt changes.
Drawings
FIG. 1 is a schematic diagram of the overall logic process of the near-range deceleration control method of the present application for an aircraft;
FIG. 2 shows a graph of the flight trajectories of two aircraft in an experimental example of the application;
FIG. 3 shows a graph of the velocity of two aircraft over time in an experimental example of the application;
FIG. 4 shows a graph of the longitudinal acceleration of two aircraft over time in an experimental example of the application;
FIG. 5 shows a graph of the lateral acceleration of two aircraft over time in an experimental example of the application;
FIG. 6 shows a graph of the change over time of the pitch rudder deflection angle of two aircraft in an experimental example of the application;
FIG. 7 shows a graph of yaw rudder deflection angle versus time for two aircraft in an experimental example of the application;
fig. 8 shows a graph of the sideslip angle of two aircraft over time in an experimental example of the application.
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 controlling near-range deceleration of an aircraft, as shown in fig. 1, the method comprises the following steps:
step 1, loading program parameters into an aircraft before the aircraft is launched;
preferably, the program parameters bound in the present application include:
closest critical distance r of aircraft and target 1 The specific value of the device needs to be selected and set according to the target distance;
initial time guidance coefficient n 0 The value of the catalyst can be n 0 ＝-0.8；
Terminal time guidance coefficient n 1 The value of the catalyst can be n 1 ＝-0.05；
Theoretical maximum speed V after the end of the active segment max0 The specific value of the device is required to be selected and set according to parameters such as the target distance, the model of the aircraft and the like;
desired tip landing angle q at hit f0 The specific value of the method can be selected and set according to the performance parameters and actual requirements of the aircraft;
maximum value deltaz of single sideslip displacement of aircraft deceleration max The specific value of the method can be selected and set according to the performance parameters and actual requirements of the aircraft;
time constant T S The time required for the smoothing coefficient K to increase from 0 to 0.6321 is determined to describe the rate of change of the smoothing coefficient. In order to ensure the smooth effect, the specific value of the device is at least less than one third of the time interval of each reversing when the aircraft decelerates and sideslips, and the device can be adjusted according to the specific experimental effect;
critical target distance r for closing lateral deceleration min The specific value is formulated according to the actual range, and the distance between the targets is ensured to be smaller than r min After flyingThe walker has enough time allowance for adjusting and stabilizing the gesture and enters the terminal guidance stage;
predetermined final speed V of aircraft f0 The specific values of the method are as follows: with V-shape aircraft f0 When terminal guidance is carried out at the speed of (2), the overload required by the guidance is smaller than the maximum available overload;
maximum critical value z of lateral deviation of flight trajectory max The specific value of the method is generally a lateral displacement value generated within seconds of sideslip of the aircraft, and the lateral displacement value is not excessively large, so that the overall trajectory of the aircraft is ensured not to generate larger deviation.
Step 2, acquiring a final pitching overload instruction of the aircraft in real time after the aircraft is launched;
after the engine of the aircraft is in a low thrust stage or the engine of the aircraft is shut down, starting to acquire a final yaw overload instruction in real time;
in the step 2, a pitch overload instruction is obtained first, and then the pitch overload instruction is limited in a specific interval, so that a final pitch overload instruction is obtained.
Preferably, the pitch overload command is obtained in real time by the following formula (one):
wherein a is yc Representing a pitch overload command,
N p and N q All of which represent the guidance coefficient,
V r when the target is a static target, acquiring speed information of the aircraft by a satellite receiver carried on the aircraft, or acquiring speed information of the aircraft by an inertia element, and further directly acquiring the relative speed; when the target is a moving target, capturing the target by a guide head on the aircraft and performing distance measurement (for example, the guide head has a distance measurement function), and acquiring the target speed by a radar or receiving instruction data and other modes, so as to acquire the relative speed by combining the speed calculation of the aircraft;
q y the vertical direction line of sight angle of the connecting line of the aircraft and the target is shown, and the vertical direction line of sight angle is measured after the target is captured by the guide head;
first time derivative representing vertical line of sight angle of aircraft to target, measured after capture of target by leader or via q y Performing differential (or difference) operation to obtain;
q f indicating a corrected tip fall angle;
t go representing the estimated remaining time of flight, the available estimation formula is t go ＝r/V r ；
g represents gravity acceleration, and the value of g is 9.8;
θ represents the ballistic tilt angle, obtained in real time by navigation systems (such as satellite navigation and inertial navigation).
Preferably, the guidance factor N p And N q Obtained by the following formula (II):
wherein n represents a guidance parameter;
preferably, the guidance parameter n=n when the engine of the aircraft is in operation, i.e. when the aircraft is in the active section 0 ，
When the aircraft is in the passive section after the engine of the aircraft is shut down, i.e. after the active section is finished, the guidance parameter n is obtained by the following formula (III):
wherein r represents the distance of the aircraft from the target,
r 0 when the active section is finished, the distance between the aircraft and the target is represented, and the value of the distance is calculated according to the position of the aircraft and the known target position;
r 1 representing the nearest critical distance between the aircraft and the target, generally taking 5km, which can be adjusted according to practical conditions. After the actual distance is smaller than the value, all guidance parameters are kept unchanged, so that the terminal flight of the aircraft is more stable.
n 1 Indicating the final time guidance factor,
n 0 indicating the initial time derivative.
Preferably, the corrected tip landing angle q when the aircraft is in the active section f Is given by q f0 ；
After the end of the active section of the aircraft, before the aircraft reaches the highest point, the corrected end falling angle q f Obtained by the following formula (IV):
q f ＝q f0 +Δq f (IV)
Wherein q f0 Indicating the desired tip landing angle at hit;
Δq f a correction deviation indicating the end drop angle;
after the aircraft passes the highest point, the corrected tip landing angle q f Is given by q f0 ；
Preferably, the corrected deviation Δq of the end falling angle f Obtained by the following substeps:
sub-step 1, obtaining a preliminary correction deviation by the following formula (five):
wherein Δq' f Representing the preliminary correction deviation;
V max representing the actual maximum speed at the end of the active segment; the value is obtained by an inertia element on the aircraft or a satellite receiver carried on the aircraft based on satellite signals;
V max0 indicating the theoretical maximum speed after the end of the active segment;
substep 2, incorporating said preliminary correction deviation Δq' f Limited to pre-treatmentDetermining the interval to obtain the correction deviation delta q of the end falling angle f ；
Preferably, in sub-step 2, the predetermined interval is [ -6,6].
When Deltaq' f If the value of (a) exceeds the interval, Δq f The value of (a) is more than one interval boundary, otherwise, deltaq' f ＝Δq f For example, if Δq' f With a value of-8, Δq f Has a value of-6 if Δq' f With a value of 3, Δq f If Δq 'has a value of 3' f Is 9, Δq f Has a value of 9.
In a preferred embodiment, the specific interval is [ -a y max ,a y max ]That is, if the value of the pitch overload instruction exceeds this section, the value of the pitch overload instruction eventually becomes beyond the section boundary on one side.
Wherein a is y max Indicating a maximum allowable overload;
preferably, the maximum allowable overload a y max Obtained by the following formula (six):
a y max ＝QS ref C nb max (six)
Wherein Q represents the dynamic pressure of air,
S ref representing a reference area of the machine body;
C nb max representing the maximum normal force coefficient.
According to the aerodynamic data of the aircraft, a normal force coefficient table is obtained, and the data of the table have the following meanings: at a mach number, the aircraft balances the normal force coefficient by maintaining a rudder deflection angle at an angle of attack. The data coordinates in the table are mach number and angle of attack, respectively.
Obtaining the current Mach number Ma and the current aerodynamic pressure Q by reading the navigation data of the aircraft and a preset atmospheric model; performing two-dimensional linear interpolation operation according to the normal force coefficient table, wherein the interpolation coordinates are Ma and the maximum allowable attack angle alpha of the aircraft max The method comprises the steps of carrying out a first treatment on the surface of the Obtaining the maximum normal force coefficient C through interpolation operation nb max 。
In a preferred embodiment, the final yaw overload command is obtained by the following formula (seven):
ψ c ＝ψ c0 K 2 +ψ c old (1-K 2 ) (seven)
Wherein, psi is c Representing a final yaw overload command;
ψ c0 representing a yaw angle command;
k represents a smoothing coefficient;
ψ c old indicating a yaw angle command to shift forward.
According to the application, the yaw overload instruction is used for controlling the aircraft to swing back and forth in the yaw direction, namely controlling the flying direction of the aircraft to switch left and right, so that the speed of the aircraft is reduced to the greatest extent, and in order to avoid unstable system caused by abrupt change of instructions in the switching of the left and right directions, the instruction smoothing process is performed in the switching process.
Preferably, the yaw angle command ψ is when the aircraft is in a left-hand state c0 ＝ψ v +β c ；
The yaw angle command ψ when the aircraft is in a right yaw state c0 ＝ψ v -β c ；
The left offset means: when the horizontal plane is viewed from top to bottom, the aircraft performs sideslip maneuver towards the left side of the horizontal velocity direction, the sideslip angle is positive, the generated sideslip aerodynamic force points to the left side of the horizontal velocity, so that the horizontal velocity of the aircraft is offset leftwards, and the right offset is opposite.
In the application, when the aircraft enters a deceleration stage and a final yaw overload instruction is obtained, the initial state of the aircraft can be set to be left-biased or right-biased at will.
When the aircraft is in a left-offset state, if the lateral displacement z of the aircraft meets z < -delta z max The aircraft is switched to a right-hand state;
when the aircraft is in the right-offset state, if the lateral displacement z of the aircraft meets z > delta z max The aircraft is switched to a left-hand state;
the lateral displacement z of the aircraft is obtained in the following way: acquiring position coordinates by inertial navigation or satellite navigation, and then calculating;
wherein, psi is v Representing the current ballistic deflection of the aircraft, the method of obtaining is as follows: obtained by inertial navigation or satellite navigation measurement;
β c representing a sideslip command;
Δz max representing a maximum value of a deceleration single sideslip displacement of the aircraft;
preferably, the sideslip command beta c Obtained by the following substeps:
a, obtaining a preliminary sideslip instruction through the following formula (eight):
wherein beta' c Representing a preliminary sideslip command;
t represents the flight time of the aircraft; in the present application, the timing is started after the aircraft is launched, i.e., the time at which the aircraft is launched is t=0.
Substep b, executing the preliminary sideslip command beta c ' restricted to a predetermined section, thereby obtaining side slip command beta c ；
Preferably, in sub-step b, the predetermined interval is [12, 16]]. I.e. if beta' c If the value of (2) exceeds the interval, beta c The value of (2) becomes beyond the interval boundary on one side.
In the application, the sideslip instruction and the final yaw overload instruction in the formula (seventh) are obtained by starting to calculate after the aircraft emits for a preset time, and further control the left and right deflection of the aircraft to reduce the speed, wherein the preset time can be selected and set according to specific requirements, but the preset time can be a small thrust stage or a passive stage after the high thrust stage of the aircraft. When the engine operating time of the aircraft is 20s, the first 10s is a high thrust phase, the rest is a low thrust phase, and the engine is started at the beginning of the flight (flight time t=0). The side slip command and the final yaw overload command in equation (seven) may be obtained beginning at t=12 s after the aircraft is launched.
Preferably, the smoothing coefficient K is obtained by the following formula (nine):
wherein e represents the base of the natural logarithm,
T S the time constant is represented by a time constant,
t represents the current time of flight and,
t 0 the time of the last sideslip reversing is represented, and in the application, the sideslip reversing refers to that the aircraft is switched from a left-biased state to a right-biased state or the aircraft is switched from the right-biased state to the left-biased state;
at each sideslip commutation, the time t of the moment is recorded and covered as t 0 If t=23s during a certain sideslip commutation, then t will be 0 Modifying the coverage to t 0 =23 s, until the next sideslip commutation t is updated again 0 Is a value of (2).
By setting the smoothing coefficient, the final yaw overload instruction obtained in the formula (seven) can be smoothly transited, and the system instability caused by abrupt instruction change during the left-right direction switching can be avoided.
In a preferred embodiment, the final yaw overload command obtained by the formula (seventh) in the application is a main deceleration command, and when the aircraft meets a certain condition, the deceleration command can be canceled, and the aircraft is controlled to fly to the target by using a common guidance command.
Specifically, in the step 2, after the final yaw overload instruction is obtained by the formula (seventh), whether the aircraft meets a lateral deceleration closing condition is determined in real time, and when the lateral deceleration closing condition is met, the final yaw overload instruction is obtained by a proportional guidance law; other guidance laws can be selected for guidance, and the guidance can be selected and set according to specific needs;
preferably, when any one of the following three conditions is satisfied, that is, the lateral deceleration closing condition is satisfied;
condition one: the distance r between the aircraft and the target satisfies r < r min ；
Condition II: the aircraft has crossed the highest point and the current real-time estimated final velocity V f Satisfy V f ＜V f0 ；
And (3) a third condition: the absolute value of the lateral displacement of the aircraft |z| satisfies |z| > z max ；
Wherein r is min A critical target distance representing a closing lateral deceleration;
V f0 indicating a predetermined end speed;
z max a maximum critical value of lateral deviation of the flight trajectory;
the distance r between the aircraft and the target is obtained by inertial navigation or satellite navigation measurement;
the real-time estimated final velocity V f Based on inertial navigation or satellite navigation measurement data, is obtained through simulation estimation.
And step 3, transmitting the final pitching overload instruction or the final pitching overload instruction and the final yawing overload instruction to the steering engine in real time, controlling steering operation of the steering engine according to the final pitching overload instruction or the final yawing overload instruction, and controlling the aircraft to fly to a target.
Experimental example
Selecting two aircraft with the same engine model, and transmitting the two aircraft from the same place to the same near target under the same environment; specifically, the engine operating time on both aircraft is 22s; the distance between the transmitting place and the target is 14km; the experimental example is performed in a mathematical simulation mode.
The first aircraft performs guidance control by the near-range deceleration control method, wherein:
step 1, loading the aircraft with program parameters prior to the aircraft launch, including:
closest critical distance r of aircraft and target 1 =2.5 km; initial time guidance coefficient n 0 -0.8; terminal time guidance coefficient n 1 =0.05; theoretical maximum speed V after the end of the active segment max0 ＝426.5m/s；Desired tip landing angle q at hit f0 =80°; maximum value deltaz of single sideslip displacement of aircraft deceleration max =100m; time constant T S =0.25 s; critical target distance r for closing lateral deceleration min =2.5 km; predetermined final speed V of aircraft f0 = 161.7m/s; maximum critical value z of lateral deviation of flight trajectory max ＝100m。
Step 2, acquiring a final pitching overload instruction of the aircraft in real time after the aircraft is launched;
obtaining a pitch overload instruction in real time by the following formula (I):
guidance factor N p And N q Obtained by the following formula (II):
guidance parameter n=n when the engine of the aircraft is in operation, i.e. when the aircraft is in the active section 0 ，
When the aircraft is in the passive section after the engine of the aircraft is shut down, i.e. after the active section is finished, the guidance parameter n is obtained by the following formula (III):
preferably, the corrected tip landing angle q when the aircraft is in the active section f Is given by q f0 ；
After the end of the active section of the aircraft, before the aircraft reaches the highest point, the corrected end falling angle q f Obtained by the following formula (IV):
q f ＝q f0 +Δq f (IV)
The preliminary correction deviation is obtained by the following formula (five):
the preliminary correction deviation deltaq 'is corrected' f Limited to a predetermined interval [ -6,6]Thereby obtaining the correction deviation delta q of the end falling angle f ；
Limiting the pitch overload command to a specific interval [ -a [ - ] y max ,a y max ]Obtaining a final pitch overload command;
said maximum allowable overload a y max Obtained by the following formula (six):
a y max ＝QS ref C nb max (six)
At 12 seconds after the aircraft launch, a final yaw overload command is obtained by the following equation (seven):
ψ c ＝ψ c0 K 2 +ψ c old (1-K 2 ) (seven)
The yaw angle command ψ when the aircraft is in a left-hand state c0 ＝ψ v +β c ；
The yaw angle command ψ when the aircraft is in a right yaw state c0 ＝ψ v -β c ；
When the aircraft is in a left-offset state, if the lateral displacement z of the aircraft meets z < -delta z max The aircraft is switched to a right-hand state;
when the aircraft is in the right-offset state, if the lateral displacement z of the aircraft meets z > delta z max The aircraft is switched to a left-hand state;
obtaining a preliminary sideslip command by the following formula (eight):
directing the preliminary sideslip command beta' c Is limited to a predetermined interval [12, 16]]Thereby obtaining sideslip command beta c ；
The smoothing coefficient K is obtained by the following formula (nine):
and step 3, transmitting the final pitching overload instruction or the final pitching overload instruction and the final yawing overload instruction to the steering engine in real time, controlling steering operation of the steering engine according to the final pitching overload instruction or the final yawing overload instruction, and controlling the aircraft to fly to a target.
At 45s of the flight of the aircraft, the distance r between the aircraft and the target is monitored to be less than r min The obtaining mode of the final yaw overload instruction is switched toWherein N takes on a value of 4, V represents the speed of the aircraft, < >>Indicating the angular velocity of the bullet eye line of sight in the horizontal direction.
The second aircraft also adopts the above scheme, but makes the following changes in the core step of deceleration:
(1) In formula (eight), the following is used: direct beta c ' =0, i.e. no sideslip deceleration measures are taken.
(2) In formula (iii), let n=n directly 1 =0.05, i.e. the guidance parameter is constant.
The remaining steps of the second aircraft solution remain the same as those of the first aircraft, and are therefore equivalent to a solution that does not take the deceleration method of the application, but only makes a conventional guided flight.
Finally, the flight track curves of the first aircraft and the second aircraft are shown in fig. 2, the speed change curve with time is shown in fig. 3, the longitudinal acceleration change curve with time and the transverse acceleration change curve with time are shown in fig. 4 and fig. 5 respectively, the pitch angle change curve with time and the yaw rudder deflection angle change curve with time are shown in fig. 6 and fig. 7 respectively, the sideslip angle change curve with time is shown in fig. 8, and the final section trajectory part parameters of the two are shown in the following table (one).
The simulation results can show that compared with the second aircraft, after the first aircraft adopts the scheme of the application to decelerate, the first aircraft has obvious lateral maneuver after reaching the highest speed point, the speed is obviously reduced, and therefore, the tail end speed is lower. Although both hit the target, the first aircraft is significantly less overloaded (accelerated) when the end maneuver occurs due to the lower speed, and therefore the end landing angle is more accurate.
The experimental example shows that the method can effectively reduce the speed of the aircraft, reduce the overload required by the aircraft and improve the terminal guidance precision of the near-range aircraft.
Watch 1
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 (10)
1. A near-range deceleration control method of an aircraft is characterized by comprising the following steps of:
the method comprises the following steps:
step 1, loading program parameters into an aircraft before the aircraft is launched;
step 2, acquiring a final pitching overload instruction of the aircraft in real time after the aircraft is launched;
after the engine of the aircraft is in a low thrust stage or the engine of the aircraft is shut down, starting to acquire a final yaw overload instruction in real time;
and step 3, transmitting the final pitching overload instruction or the final pitching overload instruction and the final yawing overload instruction to the steering engine in real time, controlling steering operation of the steering engine according to the final pitching overload instruction or the final yawing overload instruction, and controlling the aircraft to fly to a target.
2. The aircraft near-range deceleration control method according to claim 1, wherein:
in the step 2, a pitch overload instruction is obtained first, and then the pitch overload instruction is limited in a specific interval, so that a final pitch overload instruction is obtained.
3. The aircraft near-range deceleration control method according to claim 2, wherein:
the pitch overload instruction is obtained in real time by the following formula (I):
wherein a is yc Representing a pitch overload command,
N p and N q All of which represent the guidance coefficient,
V r indicating the relative speed of the aircraft to the target,
q y representing the vertical line of sight angle of the aircraft to the target,
a first time derivative representing the vertical line of sight angle of the aircraft to the target,
q f indicating that the corrected end-of-line angle is falling,
t go representing the estimated time of flight remaining in the vehicle,
g represents the acceleration of gravity and,
θ represents the ballistic tilt angle.
4. A near-range deceleration control method for an aircraft according to claim 3, wherein:
guidance factor N p And N q Obtained by the following formula (II):
wherein n represents a guidance parameter;
preferably, the guidance parameter n=n when the engine of the aircraft is in operation, i.e. when the aircraft is in the active section 0 ，
When the aircraft is in the passive section after the engine of the aircraft is shut down, i.e. after the active section is finished, the guidance parameter n is obtained by the following formula (III):
wherein r represents the distance of the aircraft from the target,
r 0 indicating the distance of the aircraft from the target at the end of the active segment,
r 1 representing the closest critical distance between the aircraft and the target,
n 1 indicating the final time guidance factor,
n 0 indicating the initial time derivative.
5. A near-range deceleration control method for an aircraft according to claim 3, wherein:
the corrected end falling angle q when the aircraft is in the active section f Is given by q f0 ；
After the end of the active section of the aircraft, before the aircraft reaches the highest point, the corrected end falling angle q f Obtained by the following formula (IV):
q f ＝q f0 +Δq f (IV)
Wherein q f0 Indicating the desired tip landing angle at hit;
Δq f a correction deviation indicating the end drop angle;
after the aircraft passes the highest point, the corrected tip landing angle q f Is given by q f0 ；
Preferably, the end falls at an angleCorrection deviation deltaq f Obtained by the following substeps:
sub-step 1, obtaining a preliminary correction deviation by the following formula (five):
wherein Δq' f Representing the preliminary correction deviation;
V max representing the actual maximum speed at the end of the active segment;
V max0 indicating the theoretical maximum speed after the end of the active segment;
substep 2, incorporating said preliminary correction deviation Δq' f Limiting to a predetermined interval to obtain a corrected deviation deltaq of the end falling angle f ；
Preferably, in sub-step 2, the predetermined interval is [ -6,6].
6. The aircraft near-range deceleration control method according to claim 2, wherein:
the specific interval is [ -a y max ,a y max ]，
Wherein a is y max Indicating a maximum allowable overload;
preferably, the maximum allowable overload a y max Obtained by the following formula (six):
a y max ＝QS ref C nb max (six)
Wherein Q represents the dynamic pressure of air,
S ref indicating the reference area of the machine body,
C nb max representing the maximum normal force coefficient.
7. The aircraft near-range deceleration control method according to claim 1, wherein:
in said step 2, said final yaw overload command is obtained by the following formula (seven):
ψ c ＝ψ c0 K 2 +ψ c old (1-K 2 ) (seven)
Wherein, psi is c Representing a final yaw overload command;
ψ c0 representing a yaw angle command;
k represents a smoothing coefficient;
ψ c old indicating a yaw angle command to shift forward.
8. The aircraft near-range deceleration control method according to claim 7, wherein:
the yaw angle command ψ when the aircraft is in a left-hand state c0 ＝ψ v +β c ；
The yaw angle command ψ when the aircraft is in a right yaw state c0 ＝ψ v -β c ；
When the aircraft is in a left-offset state, if the lateral displacement z of the aircraft meets z < -delta z max The aircraft is switched to a right-hand state;
when the aircraft is in the right-offset state, if the lateral displacement z of the aircraft meets z > delta z max The aircraft is switched to a left-hand state;
wherein, psi is v Indicating the current ballistic deflection of the aircraft,
β c representing a sideslip command;
Δz max representing a maximum value of a deceleration single sideslip displacement of the aircraft;
preferably, the sideslip instruction is obtained by the sub-steps of:
a, obtaining a preliminary sideslip instruction through the following formula (eight):
wherein beta' c Representing a preliminary sideslip command;
t represents the flight time of the aircraft;
substep b, combining said preliminary stepsSideslip command beta' c Is limited to a predetermined interval, thereby obtaining a sideslip command beta c ；
Preferably, in sub-step b, the predetermined interval is [12, 16].
9. The aircraft near-range deceleration control method according to claim 7, wherein:
the smoothing coefficient K is obtained by the following formula (nine):
wherein e represents the base of the natural logarithm,
T S the time constant is represented by a time constant,
t represents the current time of flight and,
t 0 indicating the time at which the last slip commutation was performed.
10. The aircraft near-range deceleration control method according to claim 7, wherein:
in the step 2, after a final yaw overload instruction is obtained through the step (seventh), judging whether the aircraft meets a lateral deceleration closing condition in real time, and when the lateral deceleration closing condition is met, obtaining the final yaw overload instruction through a proportional guidance law;
preferably, when any one of the following three conditions is satisfied, that is, the lateral deceleration closing condition is satisfied;
condition one: the distance r between the aircraft and the target satisfies r < r min ；
Condition II: the aircraft has crossed the highest point and the current real-time estimated final velocity V f Satisfy V f ＜V f0 ；
And (3) a third condition: the absolute value z of the lateral displacement of the aircraft is such that z > z max ；
Wherein r is min A critical target distance representing a closing lateral deceleration;
V f0 indicating a predetermined end speed;
z max the maximum threshold value of the lateral deviation of the flight trajectory.
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