CN116661495A - Near-range deceleration control method for aircraft - Google Patents

Abstract

The invention discloses a short-range deceleration control method for an aircraft. In the method, the control of the aircraft is divided into two parts: longitudinal control and lateral control. The function of deceleration; and then get the final pitch overload command and the final yaw overload command, which are passed to the steering gear, and the control is performed through the steering gear to control the aircraft to fly to the target.

Description

A kind of aircraft short-range deceleration control method

technical field

The invention relates to a control method of an aircraft, in particular to an aircraft short-range deceleration control method.

Background technique

During the flight of the aircraft, the speed and maneuverability are usually contradictory. In the short-range terminal guidance, due to the shortening of the distance, the aircraft is required to have better turning maneuverability; and in order to adapt to various ranges, the power of the aircraft It is usually designed for redundancy, which will cause its flying speed to be too fast in short-range missions; in conventional guided flight, the proportional guidance method is a commonly used guidance law, and the proportional guidance method determines the overload and speed of the aircraft and Target line-of-sight angular velocity is directly proportional; therefore, excessive velocity will result in greater g-g acceleration required for guidance. The maximum overload acceleration that an aircraft can produce is usually fixed, and the overload required for guidance should not exceed the overload capability of the aircraft for a long time, otherwise it will greatly affect the guidance accuracy and flight stability; in order to reduce the overload required for maneuvering, it is necessary to significantly reduce speed.

The speed of the aircraft is related to thrust and drag. Due to the short working time of the engine, the aircraft will soon enter the passive section after launch, that is, the unpowered flight section. At this point, drag and altitude change are the main factors that determine speed. Usually the height change depends on the trajectory, and its law is relatively fixed. The drag of an aircraft is related to its aerodynamic shape and flight attitude, which can be directly controlled.

In actual work, in order to improve the universality of the aircraft and ensure that the aircraft can be used to strike short-range targets under special circumstances, there is currently no control scheme for slowing down the aircraft to meet short-range targets.

Due to the above reasons, the present inventor has done in-depth research on the short-range deceleration control method of the aircraft, in order to design a short-range deceleration control method of the aircraft that can solve the above-mentioned problems.

Contents of the invention

In order to overcome the above-mentioned problems, the inventor has carried out intensive research and designed a method for controlling the deceleration of an aircraft at short range. In this method, the control of the aircraft is divided into two parts: longitudinal control and lateral control. Lateral control is the main deceleration. means, the longitudinal control plays the role of guidance and auxiliary lateral deceleration; and then obtains the final pitch overload command and the final yaw overload command, which are all transmitted to the steering gear, and the steering gear is used to perform control, control the aircraft to fly to the target, and complete the present invention .

Specifically, the object of the present invention is to provide a kind of aircraft short-range deceleration control method, and this method comprises the following steps:

Step 1, before the launch of the aircraft, the program parameters are bound in the aircraft;

Step 2, starting to obtain the final pitch overload command of the aircraft in real time after the aircraft is launched;

When the engine of the aircraft is in the low-thrust stage or the engine of the aircraft is shut down, the final yaw overload command is obtained in real time;

In step 3, the final pitch overload command, or the final pitch overload command and the final yaw overload command are transmitted to the steering gear in real time, and the steering gear is controlled accordingly to control the aircraft to fly to the target.

Wherein, in the step 2, the pitch overload command is obtained first, and then the pitch overload command is limited to a specific interval, so as to obtain the final pitch overload command.

Wherein, the pitching overload instruction is obtained in real time through the following formula (1):

Among them, a _yc represents the pitch overload command,

Both N _p and N _q represent the guidance coefficient,

V _r represents the relative speed of the aircraft and the target,

q _y represents the vertical line of sight angle between the aircraft and the target,

Indicates the first-order time derivative of the vertical line-of-sight angle between the aircraft and the target,

q _f represents the correction end fall angle,

t _go represents the estimated remaining flight time,

g is the acceleration due to gravity,

θ represents the ballistic inclination.

Among them, the guidance coefficients N _p and N _q are obtained by the following formula (2):

Among them, n represents the guidance parameter;

Preferably, when the engine of the aircraft is working, that is, when the aircraft is in the active phase, the guidance parameter n=n ₀ ,

After the engine of the aircraft is shut down, that is, after the active segment is over, when the aircraft is in the passive segment, the guidance parameter n is obtained by the following formula (3):

Among them, r represents the distance between the aircraft and the target,

r ₀ represents the distance between the aircraft and the target at the end of the active segment,

r ₁ represents the shortest critical distance between the aircraft and the target,

n ₁ represents the guidance coefficient at the final moment,

n ₀ represents the guidance coefficient at the initial moment.

Wherein, when the aircraft is in the active segment, the value of the corrected end fall angle _qf is _qf0 ;

After the active segment of the aircraft ends, before the aircraft reaches the highest point, the corrected end fall angle _qf is obtained by the following formula (4):

q _f ＝q _f0 +Δq _f (4)

Among them, q _f0 represents the expected end drop angle at the time of hit;

Δq _f represents the correction deviation of the end fall angle;

After the aircraft passes the highest point, the value of the corrected end fall angle _qf is _qf0 ;

Preferably, the corrected deviation Δq _f of the end fall angle is obtained through the following sub-steps:

Sub-step 1, obtain the preliminary correction deviation through the following formula (5):

Among them, Δq′ _f represents the preliminary correction deviation;

V _max represents the actual maximum speed at the end of the active segment;

V _max0 represents the theoretical maximum speed after the end of the active segment;

Sub-step 2, limiting the preliminary correction deviation Δq' _f to a predetermined interval, thereby obtaining the correction deviation Δq _f of the end fall angle;

Preferably, in sub-step 2, the predetermined interval is [-6, 6].

Wherein, the specific interval is [-a _{y max} , a _{y max} ],

Among them, a _{y max} represents the maximum allowable overload;

Preferably, the maximum allowable overload a _{y max} is obtained by the following formula (6):

a _{y max} = QS _ref C _{nb max} (6)

Among them, Q represents the air dynamic pressure,

S _ref represents the body reference area,

C _{nb max} represents the maximum normal force coefficient.

Wherein, in the step 2, the final yaw overload command is obtained by the following formula (7):

ψ _c ＝ψ _c0 K ² +ψ _cold (1-K ² ) (7)

Among them, ψ _c represents the final yaw overload command;

ψ _c0 represents the yaw angle command;

K represents the smoothing coefficient;

ψ _cold represents the yaw angle command before commutation.

Wherein, when the aircraft is in the state of left yaw, the yaw angle command ψ _c0 =ψ _v +β _c ;

When the aircraft is in a state of right yaw, the yaw angle command ψ _c0 =ψ _v -β _c ;

When the aircraft is in the state of left deviation, if the lateral displacement z of the aircraft satisfies z<-Δz _max , the aircraft will switch to the state of right deviation;

When the aircraft is in the right-biased state, if the lateral displacement z of the aircraft satisfies z>Δz _max , the aircraft will switch to the left-biased state;

Among them, ψ _v represents the current trajectory deflection angle of the aircraft,

β _c represents the sideslip command;

Δz _max represents the maximum value of a single sideslip displacement when the aircraft decelerates;

Preferably, the sideslip instruction is obtained through the following sub-steps:

In sub-step a, the preliminary sideslip instruction is obtained through the following formula (8):

Among them, β′ _c represents the preliminary sideslip command;

t represents the flight time of the aircraft;

Sub-step b, limiting the preliminary sideslip instruction β _c ′ to a predetermined interval, thereby obtaining the sideslip instruction β _c ;

Preferably, in sub-step b, the predetermined interval is [12, 16].

Wherein, described smoothing coefficient K obtains by following formula (9):

Among them, e represents the base of the natural logarithm,

T _S represents the time constant,

t represents the current flight time,

t ₀ represents the time of the last skidding changeover.

Wherein, in the step 2, after obtaining the final yaw overload command through formula (7), it is judged in real time whether the aircraft meets the lateral deceleration closing condition, and when the lateral deceleration closing condition is met, the final yaw The overload command is obtained through the proportional guidance guidance law;

Preferably, when any one of the following three conditions is satisfied, the lateral deceleration closing condition is satisfied;

Condition 1: The distance r between the aircraft and the target satisfies r<r _min ;

Condition 2: The aircraft has passed the highest point, and the current real-time estimated terminal velocity V _f satisfies V _f < V _f0 ;

Condition 3: The absolute value of the lateral displacement |z| of the aircraft satisfies |z|＞z _max ;

Among them, r _min represents the critical target distance for turning off the lateral deceleration;

V _f0 represents the predetermined final velocity;

z _max The maximum critical value of the lateral deviation of the flight trajectory.

The beneficial effects that the present invention has include:

(1) According to the aircraft short-range deceleration control method provided by the present invention, it is mainly used in aircraft with long-range flight capabilities, so that the aircraft can actively decelerate to improve the accuracy of the terminal guidance when performing short-range missions, and finally can accurately hits short-range targets;

(2) According to the aircraft short-range deceleration control method provided by the present invention, there is no requirement for the thrust control of the engine, and it is suitable for aircraft driven by solid motors, and the aircraft does not need to add additional deceleration actuators, so it has strong versatility;

(3) According to the aircraft short-range deceleration control method provided by the present invention, the longitudinal and lateral directions can be controlled simultaneously. The longitudinal channel adjusts the ballistic height and the drop angle by adjusting the guidance law coefficient, thereby affecting the speed; the lateral channel adopts left and right sideslip alternately The method can not only increase the resistance, provide a significant deceleration effect, but also basically stabilize the trajectory without affecting the terminal guidance;

(4) According to the aircraft short-range deceleration control method provided by the present invention, there is a perfect condition mechanism, which can accurately grasp the timing of the start and end of deceleration, and avoid the impact of the deceleration process on flight stability;

(5) According to the aircraft short-range deceleration control method provided by the present invention, the instructions generated in each link are provided with protective measures, which can ensure that the size of the instruction is within the executable range of the aircraft, and a smooth transition is provided when the instruction is switched. Measures to avoid system destabilization due to excessive command size or sudden change.

Description of drawings

FIG. 1 shows a schematic diagram of the overall logic process of the aircraft short-range deceleration control method of the present application;

Fig. 2 shows the flight trajectory curve diagram of two aircrafts in the experimental example of the present application;

Fig. 3 shows two aircraft speed curves with time in the experimental example of the application;

Fig. 4 shows two aircraft longitudinal acceleration curves with time in the experimental example of the application;

Fig. 5 shows two aircraft lateral acceleration curves with time in the experimental example of the application;

Fig. 6 shows two aircraft pitch rudder deflection angles curve graphs with time in the experimental example of the application;

Fig. 7 shows two aircraft yaw rudder deflection curves with time in the experimental example of the application;

Fig. 8 shows curves of sideslip angle of two aircrafts changing with time in the experimental example of the present application.

Detailed ways

The present invention will be further described in detail through the drawings and examples below. Through these descriptions, the features and advantages of the present invention will become more apparent.

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 superior or better than other embodiments. While various aspects of the embodiments are shown in drawings, the drawings are not necessarily drawn to scale unless specifically indicated.

According to a kind of aircraft short-range deceleration control method provided by the present invention, as shown in Figure 1, the method comprises the following steps:

Step 1, before the launch of the aircraft, the program parameters are bound in the aircraft;

Preferably, the program parameters of binding in the present application include:

The shortest critical distance r ₁ between the aircraft and the target, its specific value needs to be selected and set according to the target distance;

Guidance coefficient n ₀ at the initial moment, its value can be n ₀ =-0.8;

Guidance coefficient n ₁ at the last moment, its value can be n ₁ =-0.05;

The theoretical maximum speed V _max0 after the active segment is over, and its specific value needs to be selected and set according to the target distance and aircraft model;

The expected end fall angle q _f0 when hitting, its specific value can be selected and set according to the performance parameters and actual needs of the aircraft;

The maximum value of the side-slip displacement Δz _max for a single side-slip of the aircraft deceleration, its specific value can be selected and set according to the performance parameters and actual needs of the aircraft;

The time constant T _S determines the time required for the smoothing coefficient K to increase from 0 to 0.6321, and is used to describe the speed of smoothing coefficient changes. In order to ensure the smoothing effect, its specific value should be at least less than one-third of the time interval of each direction change when the aircraft decelerates and skids, which can be adjusted according to the specific experimental results;

The critical target distance r _min for turning off the lateral deceleration, its specific value is determined according to the actual range, and it should be ensured that after the target distance is less than r _min , the aircraft has enough time margin to adjust, stabilize the attitude and enter the final guidance stage;

The specific value of the predetermined final velocity V _f0 of the aircraft should satisfy: when the aircraft performs terminal guidance at the speed of V _f0 , the overload required for guidance is less than the maximum available overload;

The maximum critical value of the flight trajectory lateral deviation z _max , its specific value is generally the lateral displacement value generated within a few seconds of the aircraft skidding, and should not be too large to ensure that the overall trajectory of the aircraft does not produce a large deviation.

Step 2, starting to obtain the final pitch overload command of the aircraft in real time after the aircraft is launched;

When the engine of the aircraft is in the low-thrust stage or the engine of the aircraft is shut down, the final yaw overload command is obtained in real time;

In the step 2, the pitch overload command is obtained first, and then the pitch overload command is limited to a specific interval, so as to obtain the final pitch overload command.

Preferably, the pitching overload command is obtained in real time through the following formula (1):

Among them, a _yc represents the pitch overload command,

Both N _p and N _q represent the guidance coefficient,

V _r represents the relative speed between the aircraft and the target. When the target is a static target, the speed information of the aircraft itself can be obtained through the satellite receiver on the aircraft, or the speed information of the aircraft itself can be obtained through the inertial components, and then the relative speed can be obtained directly Size; when the target is a moving target, the seeker on the aircraft can capture the target and measure the distance (for example, the seeker has the function of ranging), and the target speed can also be obtained by means of radar or receiving command data, and then combined with The relative speed is obtained by calculating the speed of the aircraft itself;

q _y represents the vertical line of sight angle between the aircraft and the target, which is measured after the seeker captures the target;

Indicates the first-order time derivative of the vertical line-of-sight angle between the aircraft and the target, which is measured by the seeker after capturing the target or obtained by differential (or differential) calculations via q _y ;

q _f represents the correction end fall angle;

t _go represents the estimated remaining flight time, and the available estimation formula is t _go =r/V _r ;

g represents the acceleration of gravity, and its value is 9.8;

θ represents the ballistic inclination, which is obtained in real time by navigation systems (such as satellite navigation and inertial navigation).

Preferably, the guidance coefficients _Np and _Nq are obtained by the following formula (2):

Among them, n represents the guidance parameter;

Preferably, when the engine of the aircraft is working, that is, when the aircraft is in the active phase, the guidance parameter n=n ₀ ,

After the engine of the aircraft is shut down, that is, after the active segment is over, when the aircraft is in the passive segment, the guidance parameter n is obtained by the following formula (3):

Among them, r represents the distance between the aircraft and the target,

r ₀ represents the distance between the aircraft and the target at the end of the active segment, and its value is calculated based on the position of the aircraft itself and the known target position;

r ₁ represents the shortest critical distance between the aircraft and the target, generally 5km, which can be adjusted according to the actual situation. When the actual distance is less than this value, all guidance parameters remain unchanged to make the final flight of the aircraft more stable.

n ₁ represents the guidance coefficient at the final moment,

n ₀ represents the guidance coefficient at the initial moment.

Preferably, when the aircraft is in the active segment, the value of the corrected end fall angle _qf is _qf0 ;

After the active segment of the aircraft ends, before the aircraft reaches the highest point, the corrected end fall angle _qf is obtained by the following formula (4):

q _f ＝q _f0 +Δq _f (4)

Among them, q _f0 represents the expected end drop angle at the time of hit;

Δq _f represents the correction deviation of the end fall angle;

After the aircraft passes the highest point, the value of the corrected end fall angle _qf is _qf0 ;

Preferably, the corrected deviation Δq _f of the end fall angle is obtained through the following sub-steps:

Sub-step 1, obtain the preliminary correction deviation through the following formula (5):

Among them, Δq′ _f represents the preliminary correction deviation;

V _max represents the actual maximum speed at the end of the active segment; its value is obtained by the inertial components on the aircraft, or by the satellite receiver on the aircraft based on satellite signals;

V _max0 represents the theoretical maximum speed after the end of the active segment;

Sub-step 2, limiting the preliminary correction deviation Δq' _f to a predetermined interval, thereby obtaining the correction deviation Δq _f of the end fall angle;

Preferably, in sub-step 2, the predetermined interval is [-6, 6].

When the value of Δq' _f exceeds this interval, the value of Δq _f exceeds the interval boundary on one side, otherwise, Δq' _f = Δq _f , for example, if the value of Δq' _f is -8, then the value of Δq _f is -6, if the value of Δq' _f is 3, then the value of Δq _f is 3, and if the value of Δq' _f is 9, then the value of Δq _f is 9.

In a preferred embodiment, the specific interval is [ _{-ay max} , a _{y max} ], that is, if the value of the pitch overload command exceeds this interval, the value of the final pitch overload command becomes an interval exceeding one side boundary.

Among them, a _{y max} represents the maximum allowable overload;

Preferably, the maximum allowable overload a _{y max} is obtained by the following formula (6):

a _{y max} = QS _ref C _{nb max} (6)

Among them, Q represents the air dynamic pressure,

S _ref represents the body reference area;

C _{nb max} represents the maximum normal force coefficient.

In this application, a table of normal force coefficients is obtained according to the aerodynamic data of the aircraft. The meaning of the data in this table is: at a certain Mach number, the normal force of the aircraft when it reaches equilibrium at a certain angle of attack by maintaining a certain rudder deflection angle coefficient. The data coordinates in the table are Mach number and angle of attack respectively.

The current Mach number Ma and aerodynamic pressure Q are obtained by reading the aircraft navigation data and the preset atmospheric model; two-dimensional linear interpolation is performed according to the normal force coefficient table, and the interpolation coordinates are Ma and the maximum allowable angle of attack of the aircraft α _max ; get the maximum normal force coefficient C _{nb max} through interpolation.

In a preferred embodiment, the final yaw overload command is obtained by the following formula (7):

ψ _c ＝ψ _c0 K ² +ψ _{c old} (1-K ² ) (7)

Among them, ψ _c represents the final yaw overload command;

ψ _c0 represents the yaw angle command;

K represents the smoothing coefficient;

ψ _{c old} represents the yaw angle command before commutation.

In this application, the yaw overload means to control the reciprocating swing of the aircraft in the yaw direction, that is, to control the flight direction of the aircraft to switch from left to right, so as to reduce the speed of the aircraft to the greatest extent, and to avoid system instability caused by command mutations when switching from left to right , to perform command smoothing when switching.

Preferably, when the aircraft is in a state of left yaw, the yaw angle command ψ _c0 =ψ _v +β _c ;

When the aircraft is in a state of right yaw, the yaw angle command ψ _c0 =ψ _v -β _c ;

Said left deflection refers to: when viewed from top to bottom on a horizontal plane, the aircraft performs a sideslip maneuver towards the left side of the horizontal velocity direction, the sideslip angle is positive, and the resulting sideslip aerodynamic force points to the left side of the horizontal velocity, making the aircraft The horizontal speed is shifted to the left, and vice versa when it is shifted to the right.

In this application, when the aircraft enters the deceleration phase and starts to obtain the final yaw overload command, the initial state of the aircraft can be arbitrarily set to be left or right.

When the aircraft is in the state of left deviation, if the lateral displacement z of the aircraft satisfies z<-Δz _max , the aircraft will switch to the state of right deviation;

When the aircraft is in the right-biased state, if the lateral displacement z of the aircraft satisfies z>Δz _max , the aircraft will switch to the left-biased state;

The lateral displacement z of the aircraft is obtained by calculating the position coordinates obtained by inertial navigation or satellite navigation;

Among them, _ψv represents the current ballistic deflection angle of the aircraft, and its acquisition method is: measured by inertial navigation or satellite navigation;

β _c represents the sideslip command;

Δz _max represents the maximum value of a single sideslip displacement when the aircraft decelerates;

Preferably, the sideslip command β _c is obtained through the following sub-steps:

In sub-step a, the preliminary sideslip instruction is obtained through the following formula (8):

Among them, β′ _c represents the preliminary sideslip instruction;

t represents the flight time of the aircraft; in this application, the timing starts after the aircraft is launched, that is, the time when the aircraft is launched is t=0.

Sub-step b, limiting the preliminary sideslip instruction β _c ′ to a predetermined interval, thereby obtaining the sideslip instruction β _c ;

Preferably, in sub-step b, the predetermined interval is [12, 16]. That is, if the value of β' _c exceeds this interval, the value of β _c becomes beyond the boundary of the interval on one side.

In the present application, the sideslip command and the final yaw overload command in formula (7) are calculated and obtained after the aircraft is launched for a predetermined time, and then control the left and right deflection of the aircraft to decelerate. The predetermined time can be selected according to specific needs Setting, but it must be after the high-thrust stage of the aircraft, it can be a low-thrust stage, or it can be a passive stage. When the working time of the engine of the aircraft is 20s, the first 10s is a high-thrust stage, and the rest is a low-thrust stage, and the engine starts at the beginning of the flight (flight time t=0). The sideslip instruction and the final yaw overload instruction in formula (7) can be obtained at t=12s after the launch of the aircraft.

Preferably, the smoothing coefficient K is obtained by the following formula (9):

Among them, e represents the base of the natural logarithm,

T _S represents the time constant,

t represents the current flight time,

t ₀ represents the time when the last side-slip reversing, in the present application, the side-slip reversing refers to that the aircraft is switched from the left-biased state to the right-biased state, or the aircraft is switched from the right-biased state to the left-biased state;

At each side-slip changeover, record the time t at that moment and overwrite it as t ₀ , if t=23s for a certain side-slip changeover, then modify _t0 to _t0 =23s until the next time The value of t ₀ is updated again when the sideslip changes direction.

In this application, by setting the smoothing coefficient, the final yaw overload command obtained by formula (7) can be smoothly transitioned, and the command mutation will not cause system instability when switching between left and right directions.

In a preferred embodiment, the final yaw overload command obtained by formula (7) in this application is the main deceleration command. When the aircraft meets certain conditions, the deceleration command can be canceled, and the common guidance command can be used to control the aircraft. Fly to the target.

Specifically, in the step 2, after the final yaw overload command is obtained through formula (7), it is judged in real time whether the aircraft satisfies the lateral deceleration closing condition, and when the lateral deceleration closing condition is met, the final The yaw overload command is obtained through the proportional guidance guidance law; here, other guidance laws can also be selected for guidance, and the settings can be selected according to specific needs;

Preferably, when any one of the following three conditions is satisfied, the lateral deceleration closing condition is satisfied;

Condition 1: The distance r between the aircraft and the target satisfies r<r _min ;

Condition 2: The aircraft has passed the highest point, and the current real-time estimated terminal velocity V _f satisfies V _f < V _f0 ;

Condition 3: The absolute value of the lateral displacement |z| of the aircraft satisfies |z|＞z _max ;

Among them, r _min represents the critical target distance for turning off the lateral deceleration;

V _f0 represents the predetermined final velocity;

z _max is the maximum critical value of the lateral deviation of the flight trajectory;

The distance r between the aircraft and the target is measured by inertial navigation or satellite navigation;

The real-time estimated final velocity V _f is obtained through simulation estimation based on inertial navigation or satellite navigation measurement data.

In step 3, the final pitch overload command, or the final pitch overload command and the final yaw overload command are transmitted to the steering gear in real time, and the steering gear is controlled accordingly to control the aircraft to fly to the target.

Experimental example

Select two aircraft of the same model with the same engine, and launch the two aircraft from the same location to the same near target in the same environment; specifically, the engines on both aircraft work for 22 seconds; the launch location The distance between the target and the target is 14km; this experiment example is carried out in the way of mathematical simulation.

The first aircraft is guided and controlled by the short-range deceleration control method described in the present invention, wherein:

Step 1, before the launch of the aircraft, the parameters of the stapling program in the aircraft include:

The shortest critical distance between the aircraft and the target is r ₁ =2.5km; the initial guidance coefficient n ₀ =-0.8; the last guidance coefficient n ₁ =0.05; the theoretical maximum speed V _max0 after the active segment ends =426.5m/s; Desired terminal drop angle q _f0 =80°; aircraft deceleration single sideslip displacement maximum value Δz _max =100m; time constant T _S =0.25s; critical target distance r _min for closing lateral deceleration =2.5km; The velocity V _f0 =161.7m/s; the maximum critical value of the flight trajectory lateral deviation z _max =100m.

Step 2, starting to obtain the final pitch overload command of the aircraft in real time after the aircraft is launched;

Obtain the pitch overload command in real time through the following formula (1):

The guidance coefficients N _p and N _q are obtained by the following formula (2):

When the engine of the aircraft is working, that is, when the aircraft is in the active phase, the guidance parameter n=n ₀ ,

After the engine of the aircraft is shut down, that is, after the active segment is over, when the aircraft is in the passive segment, the guidance parameter n is obtained by the following formula (3):

Preferably, when the aircraft is in the active segment, the value of the corrected end fall angle _qf is _qf0 ;

After the active segment of the aircraft ends, before the aircraft reaches the highest point, the corrected end fall angle _qf is obtained by the following formula (4):

q _f ＝q _f0 +Δq _f (four)

The preliminary correction deviation is obtained by the following formula (5):

Limiting the preliminary correction deviation Δq' _f to a predetermined interval [-6, 6], thereby obtaining the correction deviation Δq _f of the end fall angle;

Limit the pitch overload command to a specific interval [-a _{y max} , a _{y max} ], so as to obtain the final pitch overload command;

The maximum allowable overload a _{y max} is obtained by the following formula (6):

a _{y max} = QS _ref C _{nb max} (6)

At 12 seconds after the launch of the aircraft, the final yaw overload command is obtained through the following formula (7):

ψ _c ＝ψ _c0 K ² +ψ _{c old} (1-K ² ) (7)

When the aircraft is in a state of left yaw, the yaw angle command ψ _c0 =ψ _v +β _c ;

When the aircraft is in a state of right yaw, the yaw angle command ψ _c0 =ψ _v -β _c ;

When the aircraft is in the state of left deviation, if the lateral displacement z of the aircraft satisfies z<-Δz _max , the aircraft will switch to the state of right deviation;

When the aircraft is in the right-biased state, if the lateral displacement z of the aircraft satisfies z>Δz _max , the aircraft will switch to the left-biased state;

Obtain the preliminary sideslip instruction through the following formula (8):

Limiting the preliminary sideslip instruction β′ _c to a predetermined interval [12, 16], thereby obtaining the sideslip instruction β _c ;

The smoothing coefficient K is obtained by the following formula (9):

In step 3, the final pitch overload command, or the final pitch overload command and the final yaw overload command are transmitted to the steering gear in real time, and the steering gear is controlled accordingly to control the aircraft to fly to the target.

In the 45th second of the flight of the aircraft, if the distance between the aircraft and the target is detected to be r<r _min , then the method of obtaining the final yaw overload command is switched to Among them, the value of N is 4, V represents the speed of the aircraft, /> Indicates the line-of-sight angular velocity 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 (8), change it to: directly set β _c ′=0, that is, do not take any sideslip deceleration measures.

(2) In formula (3), directly set n=n ₁ =0.05, that is, the guidance parameter is constant.

The remaining steps of the second aircraft scheme remain the same as the first aircraft, so it is equivalent to not taking the deceleration method of the present invention, but only doing the scheme of conventional guidance flight.

Finally, the flight trajectory curves of the first aircraft and the second aircraft are shown in Figure 2, the speed-versus-time curves are shown in Figure 3, and the longitudinal and lateral acceleration curves of the aircraft are shown in Figures 4 and 5, respectively. , the curves of aircraft pitch and yaw rudder deflection with time are shown in Figure 6 and Figure 7 respectively, and the curve of aircraft sideslip angle with time is shown in Figure 8. Show.

It can be shown from the above simulation results that, compared with the second aircraft, after the first aircraft is decelerated by adopting the scheme of the present invention, it has a significant lateral maneuver after reaching the highest speed point, and the speed is significantly reduced. Therefore its terminal velocity is lower. Although both can hit the target, because of the lower speed of the first aircraft, the overload (acceleration) during the final maneuver is significantly smaller, so the final fall angle is more accurate.

This experimental example shows that the method of the present invention can effectively reduce the speed of the aircraft, reduce the required overload of the aircraft, and improve the guidance accuracy of the terminal of the short-range aircraft.

Table I)

The present invention has been described above in conjunction with preferred embodiments, but these embodiments are only exemplary and serve as illustrations only. On this basis, various replacements and improvements can be made to the present invention, all of which fall within the protection scope of the present invention.

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.