CN109115035A - A kind of carrier rocket grade return phase precise guidance method based on trajectory forming - Google Patents
A kind of carrier rocket grade return phase precise guidance method based on trajectory forming Download PDFInfo
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
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Abstract
The invention proposes a kind of carrier rocket grade return phase precise guidance methods based on trajectory forming, belong to guidance and control technology field.The virtual inertia line of sight rate model that bookbinding target information forms before this method devises sub- grade state by executing agency of grid rudder, carrying output of navigating by arrow, penetrates, the trajectory forming optimal guidance expression formula under trajectory system has been derived, and then has obtained sub- grade return phase real-time flight overload instruction.The present invention can effectively improve sub- grade and settle in an area and sub- grade reuse return guidance precision, and by making full use of grid rudder in the high control efficiency of big dynamic pressure area, effective guarantee grade is landed with it is expected that the angle of fall is realized in the high-precision of target landing point.
Description
Technical field
The present invention relates to it is a kind of based on trajectory forming carrier rocket grade return phase precise guidance method, belong to guidance with
Control technology field.
Background technique
Currently, track commonly passes through densely populated areas without control free falling after the sub- level work of China's rocket, land
Speed is larger, and impact dispersion range is wide, remains weighing several tons from the air fall ground when have biggish destructive power, have it is potential
Security risk.Enormous impact force during chorista remains are born makes rocket body be likely to occur explosion, to surrounding objects
It causes serious harm, the personnel near drop point is caused with very big physiology and psychological impact, threaten ground staff's safety；Propellant
Remaining residue propellant in storage tank, due to generalling use conventional toxic propellant at present, it is born after propellant leakage also over the ground
Face personnel safety constitutes grave danger；Further, since the limitation of China's location of launching site, in order to ensure what chorista was settled in an area
Safety, it usually needs launching trajectory is adjusted, affects the carrying capacity of rocket to a certain extent.
It should be noted which results in single ignition due to that can not be reused using the rocket grade returned without control
The production of arrow and test period are longer, and rocket production process need to separate progress with launch mission, and should shift to an earlier date 1-2 and formulate transmitting
Plan and customized rocket, emission density are difficult to break through, is unable to satisfy quick response mission need.Therefore, it is delivered to break through tradition
The limitation of rocket, and launch cost is reduced, the concept of Control System for Reusable Launch Vehicle gradually generates and by industry weight in recent years
Depending on.Relative to the disposable carrier rocket of tradition, Control System for Reusable Launch Vehicle refers to that vehicle takes off from ground and completes predetermined transmitting
It is all or part of to return to simultaneously safe landing after task.In return course, not only need accurately to control drop point, simultaneously
The trajectory tilt angle to landing end is also needed to constrain, traditional method of guidance is difficult to realize biggish end trajectory tilt angle,
New return final guidance method urgently proposes.
Summary of the invention
The present invention is in order to solve the problems, such as that existing traditional guidance method is difficult to realize biggish terminal end trajectory tilt angle, together
When settle in an area towards rocket grade and be precisely controlled and reuse recycling demand, propose the sub- grade return phase system based on trajectory forming
Guiding method.This method is a kind of guidance method derived based on the theory of optimal control, its advantage is that biggish expectation may be implemented
End trajectory tilt angle compensates for the defect that conventional lead method is difficult to realize big end trajectory tilt angle, settles in an area for rocket grade
Control and all types of Reusable Launch Vehicles recycling control including vertically returning to vehicle are applicable in.Technical solution is detailed
Steps are as follows:
A kind of carrier rocket grade return phase precise guidance method based on trajectory forming, which comprises
Step 1: the real-time position that navigation system obtains carrier rocket grade under inertial system is carried using the arrow based on GPS/INS
It sets and real-time speed；
Step 2: the carrier rocket grade target longitudinal direction angle of sight is obtained using the drop point site and drop point speed of expectation target
With component of the line of sight angular speed of the lateral angle of sight and carrier rocket grade under the system of ground；
Step 3: according to ground system to the coordinate conversion matrix of trajectory system, the line of sight angle of carrier rocket grade is obtained
Component of the speed under trajectory system；
Step 4: trajectory is obtained using component of the line of sight angular speed under trajectory system and shapes optimal guidance law, and root
Trajectory forming optimal guidance overload instruction is obtained according to trajectory forming optimal Guidance；
Step 5: it is swung by grid rudder and generates pneumatic control torque to control the rotation of carrier rocket grade posture, realized
To the physical responses of trajectory forming optimal guidance overload instruction.
Further, the acquisition of component of the line of sight angular speed of carrier rocket grade under the system of ground described in step 2
Process are as follows:
Step 1: utilizing the position vector [x, y, z] and velocity vector of carrier rocket gradeObtain carrier rocket
Relative position velocity vector [x between grade and expectation target drop pointr, yr, zr] and [Vxr, Vyr, Vzr]:
Wherein, xr, yr, zrFor component of the Relative position vector of sub- grade and target under inertial system, xt, yt, ztFor target
Position vector under inertial system, x, y, z are position vector of the sub- grade under inertial system, vrx, vry, vrzExist for sub- grade and target
Relative velocity vector under inertial system,For velocity vector of the target under inertial system,It is sub- grade under inertial system
Velocity vector；
Step 2: determining longitudinal angle of sight and the lateral angle of sight using Relative position vector:
Wherein, qalcFor longitudinal angle of sight, qelcFor the lateral angle of sight；
Step 3: determining the rotation angle speed of sight relative inertness system under earth axes using relative position velocity vector
It spends, then line of sight rate under earth axes are as follows:
Wherein, Ωx、ΩyAnd ΩzIt is expressed as line of sight rate three axis components under earth axes；For square of sub- grade and target drop point relative distance.
Further, carrier rocket grade is obtained to the coordinate conversion matrix of trajectory system according to ground system described in step 3
Line of sight angular speed component form under trajectory system process are as follows:
That is:
Ωy2=-Ωxsinθ+Ωycosθ
Ωz2=Ωxsinψvcosθ+Ωzcosψv+Ωysinψvsinθ
Wherein, ω2Indicate the line of sight angular velocity vector of carrier rocket grade under trajectory system；θ is trajectory tilt angle, ψvFor
Trajectory deflection angle, Ωx2Indicate that line of sight rate vector is in the component of x-axis under trajectory system；Ωz2、Ωy2Respectively longitudinally, laterally sight
Angular speed.
Further, the derivation process of the forming of trajectory described in step 4 optimal guidance law are as follows:
Step 1: establishing carrier rocket grade in the equation of motion of fore-and-aft plane and lateral plane；
Step 2: obtaining system state equation using the equation of motion；
Step 3: obtaining the trajectory using the system state equation and shape optimal guidance law.
Further, the acquisition process of the equation of motion described in step 1 are as follows:
Step 1: it is closed using rocket grade movement relation is related to relative kinematic principle acquisition line-of-sight distance rate of change
It is formula:
Wherein: γDFor velocity vector fore-and-aft plane deflection,For velocity vector, λDFor longitudinal angle of sight,It is vertical
To line-of-sight rate by line, ηDFor the angle in fore-and-aft plane between velocity vector and sight, ρ is line-of-sight distance,For line-of-sight distance variation
Rate, ηTFor the angle of velocity vector in lateral plane and fore-and-aft plane, γTThe deflection for being speed in lateral plane, λTFor
The lateral angle of sight,For lateral line-of-sight rate by line；
Step 2: to formula in step 1 (5) about time t progress derivation, it is concluded that the fortune of fore-and-aft plane and lateral plane
Dynamic equation are as follows:
Wherein,Indicate acceleration.
Further, the acquisition process of system state equation described in step 2 are as follows:
Step 1: taking end conswtraint condition is that the angle of sight is equal with speed inclination angle, and LOS guidance is zero, it may be assumed that
Step 2: setting state variable:
System state equation is then obtained at this time are as follows:
End conswtraint condition are as follows:
Wherein, γDFIndicate desired speed inclination angle, x1Indicate the sum of the real-time angle of sight and trajectory tilt angle, x2Indicate that sight turns
Rate；
Step 3: providing assumed condition are as follows:
Wherein, TgIndicate residual non-uniformity；
Step 4: assumed condition in third step is substituted into formula (9), then state equation can simplify are as follows:
Note:
Then obtain the vector expression of state equation are as follows:
Further, the acquisition process of the forming of trajectory described in step 3 optimal guidance law are as follows:
Step 1: determining that trajectory constrains terminal condition using Linear Quadratic type performance indicator, wherein linear system two
Secondary type performance indicator are as follows:
Wherein xT(tf)Fx(tf) for constraining terminal condition,For reducing trajectory amplitude of fluctuation；
Based on modern control theory, the form of optimal control solution are as follows: u*=-R-1BTPx, wherein R=1,
Step 2: carrying out p solution using form of the inverse Riccati equation to optimal control solution, wherein optimal control solution
Form are as follows: u*=-R-1BTPx, wherein R=1,
Inverse Riccati equation are as follows:
It enables: E=p-1, then formula (15) converts are as follows:
Then:
Wherein E11、E12、E21、E22For four elements in E gusts.
It will can be obtained after formula (17) expansion:
From E symmetry: E12=E21, then formula (18) changes are as follows:
Consider terminal condition are as follows: E11(tf)=E12(tf)=E32(tf)=0 then can obtain formula (19) integral:
Then Riccati non trivial solution are as follows:
Then p gusts are as follows:
Step 3: p gusts of substitution optimal control solutions can then be obtained:
The sub- grade of rocket one can be obtained in the optimal motor-driven guidance law of fore-and-aft plane.
Step 4: by longitudinal angle of sight λDIt is substituted for lateral angle of sight λT, longitudinal line-of-sight rate by lineIt is substituted forEnd
Desired speed tilt angle gammaDFReplace with end desired speed drift angle γTF, ηDIt is substituted for ηT, by the process weight of the first step to third step
The new lateral plane trajectory that derives shapes optimal guidance law, and the lateral plane trajectory shapes optimal guidance law are as follows:
Step 5: can be obtained by formula (24) and (23), longitudinally and laterally overload instruction are as follows:
Wherein, λDAnd λTAs q in formula (2)alcAnd qelc,WithAs Ω in formula (5)z2And Ωy2。
The invention has the advantages that:
The invention proposes a kind of carrier rocket grade return phase precise guidance methods based on trajectory forming.This method is set
Bookbinding target information forms virtual before having counted sub- grade state by executing agency of grid rudder, carrying output of navigating by arrow, having penetrated
Inertia line of sight rate model has derived the trajectory forming optimal guidance expression formula under trajectory system, and then has obtained sub- grade return phase
Real-time flight overload instruction.
A kind of carrier rocket grade return phase precise guidance method based on trajectory forming proposed by the present invention can be mentioned effectively
High sub- grade, which settles in an area to reuse with sub- grade, returns to guidance precision, by making full use of grid rudder in the high control effect of big dynamic pressure area
Rate, effective guarantee grade are landed with it is expected that the angle of fall is realized in the high-precision of target landing point.Flight is returned without control compared to tradition
Mode, the present invention can significantly improve impact accuracy to ten meter levels, and can realize it is vertical return to landing point, will be returned in carrier rocket
It falls area's control after rise and the reusable field of VTOL in future plays a significant role.
Detailed description of the invention
Fig. 1 is the carrier rocket grade schematic diagram of the present invention in head installation grid rudder.
Fig. 2 is the flow chart of the carrier rocket grade return phase precise guidance method of the present invention based on trajectory forming.
Carrier rocket grade movement relation figure of the present invention when Fig. 3.
Specific embodiment
The present invention will be further described combined with specific embodiments below, but the present invention should not be limited by the examples.
Embodiment 1:
A kind of carrier rocket grade return phase precise guidance method based on virtual proportional guidance, as depicted in figs. 1 and 2,
Described method includes following steps:
Step 1: navigation system is carried based on the arrow comprising devices such as GPS/INS and obtains carrier rocket grade under inertial system
Real time position and real-time speed；
Step 2: drop point site and drop point speed based on expectation target obtain carrier rocket grade --- and target longitudinally regards
Line angle and the lateral angle of sight and carrier rocket grade --- component of the line of sight angular speed under the system of ground；
Step 3: based on ground system to the coordinate conversion matrix of trajectory system, carrier rocket grade --- line of sight is obtained
Component of the angular speed under trajectory system；
Step 4: deriving trajectory forming optimal guidance law, obtains trajectory forming optimal guidance overload instruction；
Step 5: it is swung by grid rudder and generates pneumatic control torque to control the rotation of carrier rocket grade posture, realized
To the physical responses of trajectory forming optimal guidance overload instruction.
Wherein, carrier rocket grade in step 2 --- line of sight angular speed component form derivation process under the system of ground
Are as follows:
Step 1: by the position vector [x, y, z] and velocity vector of carrier rocket gradeObtain carrier rocket grade
Relative position velocity vector [x between expectation target drop pointr, yr, zr] and [Vxr, Vyr, Vzr]:
Wherein, xr, yr, zrFor component of the Relative position vector of sub- grade and target under inertial system, xt, yt, ztFor target
Position vector under inertial system, x, y, z are position vector of the sub- grade under inertial system, vrx, vry, vrzExist for sub- grade and target
Relative velocity vector under inertial system,For velocity vector of the target under inertial system,It is sub- grade under inertial system
Velocity vector；
Step 2: determining longitudinal angle of sight and the lateral angle of sight using Relative position vector:
Wherein, qalcFor longitudinal angle of sight, qelcFor the lateral angle of sight；
Step 3: determining the rotation angle speed of sight relative inertness system under earth axes using relative position, velocity vector
It spends, then line of sight rate under earth axes are as follows:
Wherein,For square of sub- grade and target drop point relative distance.
Line of sight rate is obtained to the coordinate conversion matrix of trajectory system under trajectory system point based on ground system in step 3
The process of amount form are as follows:
That is:
Ωy2=-Ωxsinθ+Ωycosθ
Ωz2=Ωxsinψvcosθ+Ωzcosψv+Ωysinψvsinθ
Wherein θ is trajectory tilt angle, ψvFor trajectory deflection angle, Ωz2、Ωy2Respectively longitudinally, laterally line-of-sight rate by line.
The forming of trajectory described in step 4 optimal guidance law derivation process is described as follows:
Firstly, establishing carrier rocket grade in the equation of motion of fore-and-aft plane and lateral plane, it is shown that specific step is as follows.
Step 1: being derived from based on the sub grade movement relation figure of rocket one and relative kinematic principle:
Wherein: ηD=λD+γD, ηT=λT-γT, γDFor velocity vector fore-and-aft plane deflection,For velocity vector,
λDFor longitudinal angle of sight,For longitudinal line-of-sight rate by line, ηDFor the angle in fore-and-aft plane between velocity vector and sight, ρ is sight
Distance,For line-of-sight distance rate of change, ηTFor the angle of velocity vector in lateral plane and fore-and-aft plane, γTIt is speed in side
Deflection into plane, λTFor the lateral angle of sight,For lateral line-of-sight rate by line；
Step 2: to formula in step 1 (5) about time t carrying out derivation, it is concluded that fore-and-aft plane and lateral plane
The equation of motion are as follows:
Wherein,Indicate acceleration,
Secondly, the optimum guidance law to fore-and-aft plane derives, detailed step are as follows:
Step 1: taking end conswtraint condition is that the angle of sight is equal with speed inclination angle, and LOS guidance is zero, it may be assumed that
Step 2: setting state variable:
System state equation is then obtained at this time are as follows:
End conswtraint condition are as follows:
Step 3: providing assumed condition are as follows: Wherein TgIndicate residual non-uniformity；
Step 4: assumed condition in step 3 is substituted into formula (9), then state equation can simplify are as follows:
Note:
Then obtain the vector expression of state equation are as follows:
Finally, carrying out optimum guidance law derivation, detailed step based on the state space system completed is established are as follows:
Step 1: selected Linear Quadratic type performance indicator are as follows:
Wherein xT(tf)Fx(tf) for constraining terminal condition,For reducing trajectory amplitude of fluctuation；
Based on modern control theory, the form of optimal control solution are as follows: u*=-R-1BTPx, wherein R=1,
Step 2: carrying out p solution based on inverse Riccati equation, known by modern control theory, inverse Riccati equation are as follows:
It enables: E=p-1, then formula (15) converts are as follows:
Then:
Wherein E11、E12、E21、E22For four elements in E gusts.
It will can be obtained after formula (17) expansion:
From E symmetry: E12=E21, then formula (18) changes are as follows:
Consider terminal condition are as follows: E11(tf)=E12(tf)=E32(tf)=0 then can obtain formula (19) integral:
Then Riccati non trivial solution are as follows:
Then p gusts are as follows:
Step 3: p gusts of substitution optimal control solutions can then be obtained:
The sub- grade of rocket one can be obtained in the optimal motor-driven guidance law of fore-and-aft plane.
Step 4: by longitudinal angle of sight λDIt is substituted for lateral angle of sight λT, longitudinal line-of-sight rate by lineIt is substituted forEnd
Desired speed tilt angle gammaDFReplace with end desired speed drift angle γTF, ηDIt is substituted for ηT, lateral plane is derived again by first three step
Trajectory shapes optimal guidance law are as follows:
Step 5: can be obtained by formula (24) and (23), longitudinally and laterally overload instruction are as follows:
Wherein, λDAnd λTAs q in formula (2)alcAnd qelc,WithAs Ω in formula (5)z2And Ωy2。
The present invention settles in an area for existing carrier rocket grade, and range is excessive and precision deficiency is recycled in reuse and proposes, draws
Enter arrow and carry the sub- grade state of navigation output, bookbinding target information, derives virtual inertia line of sight rate computation model, and pass through bullet
Road shapes optimum guidance law, obtains sub- grade return phase zero-miss guidance instruction.Effectively improve carrier rocket grade settle in an area and
Reuse recycling guidance precision.
Although the present invention has been disclosed in the preferred embodiment as above, it is not intended to limit the invention, any to be familiar with this
The people of technology can do various changes and modification, therefore protection of the invention without departing from the spirit and scope of the present invention
Range should subject to the definition of the claims.
Claims (7)
1. a kind of carrier rocket grade return phase precise guidance method based on trajectory forming, which is characterized in that the method packet
It includes:
Step 1: using based on GPS/INS arrow carry navigation system obtain inertial system under carrier rocket grade real time position and
Real-time speed；
Step 2: the carrier rocket grade target longitudinal direction angle of sight and side are obtained using the drop point site and drop point speed of expectation target
To component of the line of sight angular speed of the angle of sight and carrier rocket grade under the system of ground；
Step 3: according to ground system to the coordinate conversion matrix of trajectory system, the line of sight angular speed of carrier rocket grade is obtained
Component under trajectory system；
Step 4: trajectory is obtained using component of the line of sight angular speed under trajectory system and shapes optimal guidance law, and according to bullet
Road shapes optimal Guidance and obtains trajectory forming optimal guidance overload instruction；
Step 5: it is swung by grid rudder and generates pneumatic control torque to control the rotation of carrier rocket grade posture, realized to bullet
Road shapes the physical responses of optimal guidance overload instruction.
2. precise guidance method according to claim 1, which is characterized in that the target of carrier rocket grade described in step 2 regards
The acquisition process of component of the angular velocity under the system of ground are as follows:
Step 1: utilizing the position vector [x, y, z] and velocity vector of carrier rocket gradeObtain carrier rocket grade with
Relative position velocity vector [x between expectation target drop pointr,yr,zr] and [Vxr,Vyr,Vzr]:
Wherein, xr,yr,zrFor component of the Relative position vector of sub- grade and target under inertial system, xt,yt,ztIt is target used
Property system under position vector, x, y, z be position vector of the sub- grade under inertial system, vrx,vry,vrzIt is sub- grade and target in inertia
Relative velocity vector under system,For velocity vector of the target under inertial system,For speed of the sub- grade under inertial system
Spend vector；
Step 2: determining longitudinal angle of sight and the lateral angle of sight using Relative position vector:
Wherein, qalcFor longitudinal angle of sight, qelcFor the lateral angle of sight；
Step 3: determining the angular velocity of rotation of sight relative inertness system under earth axes using relative position velocity vector, then
Line of sight rate under earth axes are as follows:
Wherein, Ωx、ΩyAnd ΩzIt is expressed as line of sight rate three axis components under earth axes；
For square of sub- grade and target drop point relative distance.
3. precise guidance method according to claim 1, which is characterized in that according to ground system to trajectory system described in step 3
Coordinate conversion matrix is to obtain the process of line of sight angular speed component form under trajectory system of carrier rocket grade are as follows:
That is:
Ωy2=-Ωxsinθ+Ωycosθ
Ωz2=Ωxsinψvcosθ+Ωzcosψv+Ωysinψvsinθ
Wherein, ω2Indicate the line of sight angular velocity vector of carrier rocket grade under trajectory system；θ is trajectory tilt angle, ψvFor trajectory
Drift angle, Ωx2Indicate that line of sight rate vector is in the component of x-axis under trajectory system；Ωz2、Ωy2Respectively longitudinally, laterally the angle of sight is fast
Rate.
4. precise guidance method according to claim 1, which is characterized in that trajectory described in step 4 shapes optimal guidance law
Derivation process are as follows:
Step 1: establishing carrier rocket grade in the equation of motion of fore-and-aft plane and lateral plane；
Step 2: obtaining system state equation using the equation of motion；
Step 3: obtaining the trajectory using the system state equation and shape optimal guidance law.
5. precise guidance method according to claim 4, which is characterized in that the acquisition process of the equation of motion described in step 1 are as follows:
Step 1: line-of-sight distance rate of change correlativity formula is obtained using rocket grade movement relation and relative kinematic principle:
Wherein: ηD=λD+γD, ηT=λT-γT, γDFor velocity vector fore-and-aft plane deflection,For velocity vector, λDFor
Longitudinal angle of sight,For longitudinal line-of-sight rate by line, ηDFor the angle in fore-and-aft plane between velocity vector and sight, ρ is line-of-sight distance
From,For line-of-sight distance rate of change, ηTFor the angle of velocity vector in lateral plane and fore-and-aft plane, γTIt is speed lateral
Deflection in plane, λTFor the lateral angle of sight,For lateral line-of-sight rate by line；
Step 2: to formula in step 1 (5) about time t progress derivation, it is concluded that the movement side of fore-and-aft plane and lateral plane
Journey are as follows:
Wherein,Indicate acceleration.
6. precise guidance method according to claim 4, which is characterized in that the acquisition of system state equation described in step 2
Journey are as follows:
Step 1: taking end conswtraint condition is that the angle of sight is equal with speed inclination angle, and LOS guidance is zero, it may be assumed that
Step 2: setting state variable:
System state equation is then obtained at this time are as follows:
End conswtraint condition are as follows:
Wherein, x1Indicate the sum of the real-time angle of sight and trajectory tilt angle, x2Indicate LOS guidance；
Step 3: providing assumed condition are as follows:
Wherein, TgIndicate residual non-uniformity；
Step 4: assumed condition in third step is substituted into formula (9), then state equation can simplify are as follows:
Note:
Then obtain the vector expression of state equation are as follows:
7. according to the precise guidance method of claim 4 or 6, which is characterized in that trajectory described in step 3 shapes optimal guidance law
Acquisition process are as follows:
Step 1: determining that trajectory constrains terminal condition using Linear Quadratic type performance indicator, wherein Linear Quadratic type
Performance indicator are as follows:
Wherein xT(tf)Fx(tf) for constraining terminal condition,For reducing trajectory amplitude of fluctuation；
Based on modern control theory, the form of optimal control solution are as follows: u*=-R-1BTPx, wherein R=1,
Step 2: carrying out p solution using form of the inverse Riccati equation to optimal control solution, wherein the form of optimal control solution
Are as follows: u*=-R-1BTPx, wherein R=1,
Inverse Riccati equation are as follows:
It enables: E=p-1, then formula (15) converts are as follows:
Then:
Wherein E11、E12、E21、E22For four elements in E gusts.
It will can be obtained after formula (17) expansion:
From E symmetry: E12=E21, then formula (18) changes are as follows:
Consider terminal condition are as follows: E11(tf)=E12(tf)=E32(tf)=0 then can obtain formula (19) integral:
Then Riccati non trivial solution are as follows:
Then p gusts are as follows:
Step 3: p gusts of substitution optimal control solutions can then be obtained:
The sub- grade of rocket one can be obtained in the optimal motor-driven guidance law of fore-and-aft plane.
Step 4: by longitudinal angle of sight λDIt is substituted for lateral angle of sight λT, longitudinal line-of-sight rate by lineIt is substituted forEnd expectation
Speed tilt angle gammaDFReplace with end desired speed drift angle γTF, ηDIt is substituted for ηT, pushed away again by the process of the first step to third step
Lateral plane trajectory forming optimal guidance law is led, the lateral plane trajectory shapes optimal guidance law are as follows:
Step 5: can be obtained by formula (24) and (23), longitudinally and laterally overload instruction are as follows:
Wherein, λDAnd λTAs q in formula (2)alcAnd qelc,WithAs Ω in formula (5)z2And Ωy2。
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| CN111473696A (en) * | 2020-03-24 | 2020-07-31 | 北京理工大学 | Guidance rocket vertical attack guidance method based on drop point estimation |
| CN111473696B (en) * | 2020-03-24 | 2021-01-26 | 北京理工大学 | Guidance rocket vertical attack guidance method based on drop point estimation |
| CN111591470A (en) * | 2020-04-28 | 2020-08-28 | 西北工业大学 | Aircraft accurate soft landing closed-loop guidance method adapting to thrust adjustable mode |
| CN112033235A (en) * | 2020-09-04 | 2020-12-04 | 清华大学 | Air combat three-body confrontation optimal prediction guidance method adopting golden section search |
| CN117288047A (en) * | 2023-10-10 | 2023-12-26 | 北京理工大学 | Two-dimensional correction fuze drop point prediction control method insensitive to model errors |
| CN117288047B (en) * | 2023-10-10 | 2024-04-12 | 北京理工大学 | Two-dimensional correction fuze drop point prediction control method insensitive to model errors |
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