CN111399531B - Hypersonic aircraft glide section guidance and attitude control integrated design method - Google Patents
Hypersonic aircraft glide section guidance and attitude control integrated design method Download PDFInfo
- Publication number
- CN111399531B CN111399531B CN202010327172.5A CN202010327172A CN111399531B CN 111399531 B CN111399531 B CN 111399531B CN 202010327172 A CN202010327172 A CN 202010327172A CN 111399531 B CN111399531 B CN 111399531B
- Authority
- CN
- China
- Prior art keywords
- control
- hypersonic
- aircraft
- guidance
- angle
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 238000000034 method Methods 0.000 title claims abstract description 44
- 238000013461 design Methods 0.000 title claims description 36
- 230000008859 change Effects 0.000 claims abstract description 10
- 230000014509 gene expression Effects 0.000 claims description 16
- 230000006870 function Effects 0.000 claims description 8
- 238000004590 computer program Methods 0.000 claims description 7
- 239000011159 matrix material Substances 0.000 claims description 4
- 238000005096 rolling process Methods 0.000 claims description 4
- 230000010354 integration Effects 0.000 claims 1
- 239000000126 substance Substances 0.000 claims 1
- 238000011160 research Methods 0.000 description 4
- 238000004088 simulation Methods 0.000 description 4
- 230000008901 benefit Effects 0.000 description 3
- 230000008878 coupling Effects 0.000 description 3
- 238000010168 coupling process Methods 0.000 description 3
- 238000005859 coupling reaction Methods 0.000 description 3
- 230000007423 decrease Effects 0.000 description 3
- 230000008569 process Effects 0.000 description 3
- 238000004458 analytical method Methods 0.000 description 2
- 238000009795 derivation Methods 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 238000005516 engineering process Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000000926 separation method Methods 0.000 description 2
- 238000013459 approach Methods 0.000 description 1
- 238000004364 calculation method Methods 0.000 description 1
- 230000003247 decreasing effect Effects 0.000 description 1
- 230000009189 diving Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000005457 optimization Methods 0.000 description 1
- 238000001228 spectrum Methods 0.000 description 1
- 238000012795 verification Methods 0.000 description 1
Images
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
- G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
- G05D1/0825—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using mathematical models
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Physics & Mathematics (AREA)
- Engineering & Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Pure & Applied Mathematics (AREA)
- Mathematical Optimization (AREA)
- Mathematical Physics (AREA)
- Mathematical Analysis (AREA)
- Algebra (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Automation & Control Theory (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The method comprises the steps of taking the geocentric distance of a current moving target and the latitude and longitude of the moving target as the input of a hypersonic aircraft gliding section guidance and attitude control integrated control model, generating a control rudder deflection angle according to a control scheme of the hypersonic aircraft gliding section guidance and attitude control integrated control model, inputting the control rudder deflection angle to a six-degree-of-freedom movement model of a hypersonic aircraft, and enabling the hypersonic aircraft to fly to the moving target to complete a flight task. The control scheme of the hypersonic aircraft glide section guidance and attitude control integrated model is designed based on the block dynamic surface method, the attitude control task requirement of the aircraft glide section is met, the guidance and control precision is high, the whole-course output change of the control quantity is smooth, the task adaptability of the aircraft is improved, and the problem of difficult coordination between the guidance and the control of the hypersonic aircraft is effectively solved.
Description
Technical Field
The invention relates to the field of aircraft control, in particular to a hypersonic aircraft glide phase guidance control method.
Background
The basic process of the hypersonic aerocraft gliding flight is to control the aerocraft to stably fly to reach a preset target end point under the condition of meeting all process constraint conditions. For the 'boosting-gliding' type hypersonic aircraft, unpowered gliding flight when the hypersonic aircraft is in the near space is the greatest advantage of the hypersonic aircraft compared with the traditional ballistic missile, and the range and the attack capability of the hypersonic aircraft are effectively improved.
However, the hypersonic aerocraft has extremely high flying speed, and the mass center movement around the mass center have the characteristics of fast time variation, nonlinearity, strong coupling, uncertainty and the like. The traditional aircraft guidance and control system is mainly based on a frequency spectrum separation theory to separately design control and guidance subsystems, coupling information among the subsystems is not utilized, and the integrated guidance and control system can fully utilize the coupling information among the subsystems to improve the performance of the whole control system.
At present, research aiming at the hypersonic aircraft mainly aims at the problems of a separation design method of a guidance control system of the hypersonic aircraft and an integrated design method of a guidance and attitude control system of a glide section, the research results of the integrated research of the guidance and attitude control of the hypersonic aircraft in the glide section are few, and related technologies are not disclosed so far.
Disclosure of Invention
Aiming at the problem of integrated design of guidance and attitude control of the glide phase of the hypersonic aerocraft in the prior art, the invention aims to provide an integrated design method of guidance and attitude control of the glide phase of the hypersonic aerocraft. Specifically, the invention relates to an integrated design method of a guidance attitude control system of an air-breathing general hypersonic aircraft model (GHV) in a glide section, and the hypersonic aircraft can well meet the task requirement of flight in the glide section by the integrated design of guidance and attitude control of the air-breathing general hypersonic aircraft model (GHV) in the glide section.
In order to achieve the technical purpose, the technical scheme adopted by the invention is as follows:
an integrated design method for guiding and attitude control of a hypersonic aircraft in gliding section is used for obtaining the geocentric distance r of a current moving target T Latitude of moving object
Longitude λ T The geocentric distance r of the current moving target is calculated T Latitude of moving object
Longitude λ T And as the input of the hypersonic aircraft glide section guidance and attitude control integrated control model, generating a control rudder deflection angle according to the control scheme of the hypersonic aircraft glide section guidance and attitude control integrated control model, inputting the control rudder deflection angle into a six-degree-of-freedom motion model of the hypersonic aircraft, and enabling the hypersonic aircraft to fly to a motion target to complete a flight task. Specifically, the hypersonic aircraft is an air-breathing general hypersonic aircraft.
In the invention, based on the geometric model of the hypersonic aircraft (as shown in fig. 2), a six-degree-of-freedom motion model of the hypersonic aircraft is constructed as follows:
in the formula: v is the flight speed of the hypersonic aerocraft, theta is the local speed inclination angle, and sigma is the track yaw angle; omega x ，ω y ，ω z Respectively, the three-axis rotation angular velocity of the hypersonic aerocraft, m is the mass of the hypersonic aerocraft, gamma V To a roll angle, I x ，I y ，I z Respectively, three-axis rotational inertia, L, D and N are aerodynamic force borne by the hypersonic aerocraft, respectively, lift force, resistance force and lateral force, M x ，M y ，M z The aerodynamic moment, the rolling moment, the yawing moment and the pitching moment, which are respectively borne by the aircraft, and the expressions of the aerodynamic force and the aerodynamic moment are respectively as follows:
the above formula is an aerodynamic expression of the hypersonic deformable aircraft, wherein,
zero angle of attack lift coefficient;
Is the rate of change of coefficient of lift with respect to angle of attack, C D Is a coefficient of resistance, C N Are the lateral force coefficients, which are known quantities; b and c are transverse lateral reference length and longitudinal reference length of the hypersonic aircraft respectively, and m is xβ Is the partial derivative of roll torque with respect to side slip angle beta, m yβ For the partial derivative of the yaw moment with respect to the sideslip angle β, <' >>
Is a partial derivative matrix of the pitch moment with respect to the slip angle beta>
For roll moment coefficient versus roll rudder deflection angle delta x Partial derivative of (a), based on the partial derivative of (b)>
Respectively yaw moment coefficient vs. rudder deflection angle delta y The partial derivative of (a) is,
respectively pitch moment coefficient for elevator delta z The above quantities are known quantities. The form of α is: α is the angle of attack value, ζ L For lift uncertainty term, α = [1 α 3 ] T 。
Respectively, triaxial moment uncertainty.
q＝0.5ρV 2 Is dynamic pressure, ρ is the atmospheric density, a known quantity, S 0 The reference area is a known quantity for the hypersonic flight vehicle. The expression of the allowance in the six-degree-of-freedom motion model of the hypersonic aircraft is as follows:
mu is the gravitational constant, omega e Is the rotational angular velocity of the earth, J =1.5J 2 Is a harmonic coefficient of e The semimajor axis of the earth is a known quantity. r is the distance between the centers of the earth.
In the invention, a guidance and attitude control integrated control model of the hypersonic aerocraft in the glide section is constructed based on a guidance equation for controlling the course and the azimuth error of the hypersonic aerocraft facing control, a kinematic equation around the center of mass of the hypersonic aerocraft facing control and a kinematic equation around the center of mass of the hypersonic aerocraft facing control.
The control-oriented guidance equation for the control of the course and the azimuth error of the hypersonic aerocraft is as follows:
in the formula:
target command x of six-degree-of-freedom motion model of hypersonic flight vehicle with local speed inclination angle theta and track yaw angle sigma as state quantities 0C Is shown as follows
x 0c ＝[θ c σ c ] T
Derivation of x 0C Medium speed tilt angle command value theta c And a track yaw angle command value sigma c Is described in (1).
The speed dip angle instruction of the hypersonic aircraft for flying and supporting the moving target is as follows
Track yaw angle command value sigma c Is expressed as
The output x of the guidance equation for controlling the range and azimuth errors of the control-oriented hypersonic aerocraft 0 For local speed inclination theta and track yaw angle sigma, input
Is coefficient of lift C L Two components of (a).
In the invention, the control-oriented hypersonic deformable aircraft has the following kinematic equation around the mass center:
in the formula,
wherein β is a slip angle, ζ 01 、ζ 02 Is ζ 0 Two components of (a).
Solving x by control-oriented hypersonic deformable aircraft around centroid kinematic equation 1 ＝[α β γ V ] T Substituting into the aerodynamic expression of hypersonic deformable aircraft to calculate the lift coefficient C L And then the input of the motion equation of the hypersonic deformable aircraft relative to the moving target facing the control can be obtained
The input of the control-oriented hypersonic aerocraft around the center of mass kinematics equation is the three-axis rotation angular velocity vector x of the hypersonic aerocraft 2 ＝[ω x ω y ω z ] T 。
In the invention, the control-oriented hypersonic aircraft dynamic equation around the centroid is as follows:
in the formula
The output of the control-oriented hypersonic aircraft mechanical equation of motion around the center of mass is x 2 The input is control rudder deflection angle u = [ delta ] of the hypersonic aircraft x δ y δ z ] T 。
The control method comprises the following steps of simultaneously establishing a control equation for controlling the course and the azimuth error of the hypersonic aerocraft facing the control, a kinematic equation around the center of mass of the hypersonic aerocraft facing the control and a kinematic equation around the center of mass of the hypersonic aerocraft facing the control, namely establishing a control model integrating the guidance and the attitude control of the glide section of the hypersonic aerocraft as follows:
in the invention: based on a block dynamic surface method, a guidance and attitude control integrated control scheme of the hypersonic aircraft in a glide section is designed. Specifically, the control scheme of the hypersonic aircraft glide section guidance and attitude control integrated control model is as follows:
in the control scheme, the method comprises the following steps:
s 0 for the first dynamic plane defined, x 1 With its instruction value x 1d Difference of (a), x 1d For the first virtual control input, k 0 ＝diag(k 01 k 02 ) Given a positive gain constant; epsilon 01 And ε 02 A saturation function term gain to be given; sat (s, d) is a saturation function defined as:
γ Vd for the calculated roll angle command value, alpha v For the calculated roll angle command values, all pass x 1d Solving to obtain;
s 1 for the second dynamic plane of definition, x 1 With its instruction value x 1d Difference of (a), x 2d A virtual control input for a second dynamic surface; k is a radical of formula 1 ＝diag(k 11 ,k 12 ,k 13 ) Given a positive gain constant; tau is 1 ＝diag(τ 11 ,τ 12 ,τ 13 ) Is the time constant of the filter. According to the virtual control quantity x 2d Is obtained such that x 1 Reach the expected instruction value x 1d Three-channel angular rate virtual input expected value omega xd ,ω yd ,ω zd ；
Definition s 2 Is a third dynamic plane, is x 2 With its instruction value x 2d U is the control input for the third dynamic surface. k is a radical of formula 2 ＝diag(k 21 k 22 k 23 ) Given a positive gain constant, τ 2 ＝diag(τ 21 ,τ 22 ,τ 23 ) Virtually inputting a desired value x according to the three-channel angular rate as a time constant of a filter 2d And obtaining a design rudder deflection angle input u to complete the stable control and the guidance control of the aircraft attitude control system.
The invention also provides a hypersonic aircraft glide section guidance and attitude control integrated system, which comprises:
a target information acquisition module for acquiring the geocentric distance r of the current moving target T Latitude of moving object
Longitude lambda T ；
The guidance module receives the target information acquired by the target information acquisition module, inputs the target information into a hypersonic aircraft glide section guidance and attitude control integrated control model which is pre-loaded on the guidance module, and generates a control rudder deflection angle according to a pre-designed control scheme of the hypersonic aircraft glide section guidance and attitude control integrated control model;
and the attitude control-control module receives the control rudder deflection angle generated by the guidance module and inputs the control rudder deflection angle into a six-degree-of-freedom motion model of the supersonic aircraft pre-loaded on the attitude control-control module to complete the stable motion of the hypersonic aircraft and realize the tracking control of the guidance instruction.
The invention also provides a hypersonic aircraft, which adopts the technical scheme that: a hypersonic aerocraft comprises an aerocraft body and an airborne circuit board arranged in the aerocraft body, wherein a processor and a memory are arranged on the airborne circuit board, a computer program is stored in the memory, and the processor executes the computer program and then realizes the step of the hypersonic aerocraft gliding section guidance and attitude control integrated design method.
The invention also provides a computer readable storage medium, wherein a computer program is stored on the computer readable storage medium, and the computer program is characterized in that when being executed by a processor, the computer program realizes the steps of the hypersonic aircraft glide section guidance and attitude control integrated design method.
Compared with the prior art, the invention has the following advantages:
the invention designs a guidance and attitude control integrated system aiming at the flight of a hypersonic aerocraft in a gliding section. The research of the integrated design of the guidance attitude control system of the glide phase is developed aiming at an air-breathing general hypersonic aerocraft model (GHV). Firstly, a glide section guidance attitude control integrated model based on range and azimuth error control is established, and the design of a guidance attitude control integrated method of a hypersonic aircraft in a glide section is completed by using a self-adaptive block dynamic surface design idea. The method is suitable for completing the flight task of the glide section of the hypersonic aircraft, has great significance in engineering application, effectively solves the problem of coordination and stability of guidance and attitude control of the hypersonic aircraft in the glide section, simultaneously ensures the robustness of a design method of a guidance control system, meets the flight task requirement of the glide section, and is suitable for integrated design of guidance and attitude control of the glide section of the hypersonic aircraft.
Drawings
FIG. 1 is a schematic flow chart of example 1
FIG. 2 is a geometric model diagram of a hypersonic aircraft
FIG. 3 is a three-dimensional variation graph of the latitude and longitude and the altitude of the track in the flight of the glide flight
FIG. 4 is a graph showing the variation of track altitude, latitude and longitude in the flight of glide flight
FIG. 5 is a graph of track angle, local velocity dip and velocity change in flight during the glide phase
FIG. 6 is a graph showing the Mach number and dynamic pressure changes in the flight in the glide phase
FIG. 7 is a graph of angle of attack, sideslip angle, and roll angle in glide flight
FIG. 8 is a graph of commanded angle tracking error in glide flight
FIG. 9 is a graph showing changes in a roll rudder, a yaw rudder, and an elevator in flight in the glide phase
Fig. 10 is a graph showing the change in triaxial angular velocity in flight in the glide section.
Detailed Description
To further clarify the objects, technical solutions and advantages of the embodiments of the present invention, the spirit of the present disclosure will be clearly described in the following drawings and detailed description, and any person skilled in the art who knows the embodiments of the present disclosure can make changes and modifications to the technology taught by the present disclosure without departing from the spirit and scope of the present disclosure. The exemplary embodiments of the present invention and the description thereof are provided to explain the present invention and not to limit the present invention.
FIG. 1 is a schematic control flow diagram of the present embodiment, which is used for collecting the distance r between the current moving target and the center of mass of the hypersonic aircraft T Latitude of moving object
Longitude λ T Will >>
And inputting a guidance module, wherein an outer ring loop is a guidance loop, and the trajectory planning and control is carried out on the hypersonic aircraft, so that the guidance precision is required to be high. The guidance module receives the information collected by the target information collection moduleAnd target information is used for generating a rate instruction for controlling the rudder deflection angle and the hypersonic aircraft. The inner loop is an attitude control loop, stable motion of the hypersonic aircraft and tracking control of a guidance instruction output by the guidance module are required to be realized, and high precision and robustness are required, so that the hypersonic aircraft can move to an input motion target.
In the embodiment, the integrated design method for guidance and attitude control of the glide section of the hypersonic aircraft comprises the following steps:
s1, constructing a six-degree-of-freedom motion model based on a geometric model of the hypersonic aircraft;
the geometric model of the hypersonic aircraft is shown in fig. 2, and the present embodiment is based on an air-breathing general hypersonic aircraft model (GHV).
The six-degree-of-freedom motion model of the hypersonic aircraft is as follows:
in the formula: v is the flight speed of the hypersonic aerocraft, theta is the local speed inclination angle, and sigma is the track yaw angle; omega x ，ω y ，ω z Respectively, the three-axis rotation angular velocity of the hypersonic aerocraft, m is the mass of the hypersonic aerocraft, gamma V To a roll angle, I x ，I y ，I z Respectively, three-axis moment of inertia;
l, D and N are aerodynamic forces borne by the hypersonic aerocraft, namely lift force, resistance force and lateral force, M x ，M y ，M z The aerodynamic moment, the rolling moment, the yawing moment and the pitching moment, which are respectively borne by the aircraft, and the expressions of the aerodynamic force and the aerodynamic moment are respectively as follows:
the above expression is an aerodynamic expression of the hypersonic deformable aircraft, wherein,
zero angle of attack lift coefficient;
Is the rate of change of coefficient of lift with respect to angle of attack, C D Is a coefficient of resistance, C N Are the lateral force coefficients, which are known quantities; b and c are transverse lateral reference length and longitudinal reference length of the hypersonic aircraft respectively, and m is xβ Is the partial derivative of roll torque with respect to the slip angle beta, m yβ For the partial derivative of the yaw moment with respect to the sideslip angle β, <' >>
Is a partial derivative matrix of the pitch moment with respect to the slip angle beta>
For roll moment coefficient versus roll rudder deflection angle delta x Is based on the partial derivative of (4)>
Respectively yaw moment coefficient to rudder deflection angle delta y The partial derivative of (a) of (b),
respectively pitch moment coefficient for elevator delta z The above quantities are known quantities. The form of α is: α is the angle of attack value, ζ L For lift uncertainty term, α = [1 α = 3 ] T 。
Respectively, triaxial moment uncertainty terms.
q＝0.5ρV 2 Is the dynamic pressure, ρ is the atmospheric density, is a known quantity, S 0 The reference area is a known quantity for the hypersonic flight vehicle. The expression of the margin in the centroid motion model is as follows
Mu is the gravitational constant, omega e Is rotational angular velocity of the earth, J =1.5J 2 Is a coefficient of harmonics, a e Are known quantities for the earth's semi-major axis. r is the distance between the centers of the earth.
S2, constructing a guidance and attitude control integrated control model of the hypersonic aerocraft in the glide section based on a guidance equation for controlling the course and the azimuth error of the control-oriented hypersonic aerocraft, a kinematics equation around the center of mass of the control-oriented hypersonic aerocraft and a kinematics equation around the center of mass of the control-oriented hypersonic aerocraft.
The guidance equation for the control-oriented hypersonic aircraft range and azimuth error control is as follows:
in the formula:
a target command x of a six-degree-of-freedom kinematic model with a local speed inclination angle theta and a track yaw angle sigma as state quantities is assumed 0C Is represented as follows:
x 0c ＝[θ c σ c ] T
derivation of x 0C Medium speed tilt angle command value theta c And a track yaw angle command value sigma c The expression (c).
The speed dip angle instruction of the hypersonic aircraft to fly to the target point is as follows
Track yaw angle command value sigma c Is expressed as
The output x of the guidance equation for controlling the range and azimuth errors of the control-oriented hypersonic aerocraft 0 For local speed inclination angle theta and track yaw angle sigma, input
Is coefficient of lift C L Two components of (a).
The control-oriented hypersonic deformable aircraft has the following kinematic equation around the center of mass:
in the formula,
wherein β is a slip angle, ζ 01 、ζ 02 Is ζ 0 Two components of (a).
Solving x through control-oriented hypersonic deformable aircraft around centroid kinematic equation 1 ＝[α β γ V ] T Substituting into the aerodynamic expression (known expression) of hypersonic deformable aircraft to obtain lift coefficient C L And then the input of the motion equation of the hypersonic deformable aircraft relative to the target for control can be obtained
The input of the control-oriented hypersonic aerocraft around the center of mass kinematics equation is the three-axis rotation angular velocity vector x of the hypersonic aerocraft 2 ＝[ω x ω y ω z ] T 。
In the invention, the control-oriented hypersonic aircraft dynamic equation around the center of mass is as follows:
in the formula
The output of the control-oriented hypersonic aircraft mechanical equation of motion around the center of mass is x 2 The input is the control rudder deflection angle u = [ delta ] of the hypersonic aerocraft x δ y δ z ] T 。
The control method comprises the following steps of simultaneously establishing a guidance equation for controlling the range and the azimuth error of the hypersonic aerocraft facing the control, a kinematics equation around the center of mass of the hypersonic aerocraft facing the control and a kinematics equation around the center of mass of the hypersonic aerocraft facing the control, namely establishing an integrated control model of the glide section guidance and the attitude control of the hypersonic aerocraft as follows:
s3, designing a guidance and attitude control integrated control scheme of the hypersonic aircraft in a glide section based on a block dynamic surface method;
the control scheme of the hypersonic aircraft glide section guidance and attitude control integrated control model is as follows:
in the above control scheme:
s 0 for the first dynamic plane defined, x 1 With its instruction value x 1d Difference of (a), x 1d For the first virtual control input, k 0 ＝diag(k 01 k 02 ) Given a positive gain constant; epsilon 01 And epsilon 02 A gain of a saturation function term to be given; sat (s, d) is a saturation function defined as:
γ Vd for the calculated roll angle command value, alpha v For the calculated roll angle command values, all pass x 1d And (6) solving to obtain.
s 1 Is a second dynamic plane of definition, is x 1 With its instruction value x 1d Difference of (a), x 2d A virtual control input for a second dynamic surface; k is a radical of 1 ＝diag(k 11 ,k 12 ,k 13 ) Given a positive gain constant; tau is 1 ＝diag(τ 11 ,τ 12 ,τ 13 ) Is the time constant of the filter. According to the virtual control quantity x 2d Is obtained such that x 1 Reach the expected instruction value x 1d Three-channel angular rate virtual input desired value omega xd ,ω yd ,ω zd 。
Definition s 2 Is a third dynamic plane, is x 2 With its instruction value x 2d U is the control input for the third dynamic surface. k is a radical of formula 2 ＝diag(k 21 k 22 k 23 ) Given a positive gain constant, τ 2 ＝diag(τ 21 ,τ 22 ,τ 23 ) Virtually inputting a desired value x according to the three-channel angular rate as a time constant of a filter 2d And obtaining a design rudder deflection angle input u to complete the stable control and the guidance control of the aircraft attitude control system.
S4, the distance r of the current moving target relative to the mass center of the hypersonic aircraft T Latitude of moving object
Longitude λ T And as the input of the hypersonic aircraft glide section guidance and attitude control integrated model, generating a control rudder deflection angle instruction according to a control method of the hypersonic aircraft glide section guidance and attitude control integrated control model, inputting the control rudder deflection angle instruction into a six-degree-of-freedom motion model of the hypersonic aircraft, and enabling the hypersonic aircraft to track a motion target to complete a flight task.
Simulation verification is carried out based on the hypersonic aircraft glide section guidance and attitude control integrated design method provided by the following steps:
simulation calculation example:
in order to verify the effectiveness of the hypersonic aircraft glide phase guidance and attitude control integrated design method, numerical simulation is carried out on the model. The effect, initial state and integrated model parameter table of the hypersonic aircraft glide section guidance and attitude control integrated design method are shown in the following tables 1 and 2.
TABLE 1 initial diving state of aircraft and longitude and latitude of target point
TABLE 2 Integrated design method parameter optimization values
2. Analysis of results
The simulation results are shown in fig. 3-10.
As can be seen from the figure 3, the hypersonic flight vehicle can well realize the flight task of the glide phase by the hypersonic flight vehicle glide phase guidance and attitude control integrated design method provided by the invention. As can be seen from fig. 4, when the aircraft flies to the terminal point, the longitude and latitude of the aircraft are 35.0001 ° and 30.0006 °, the errors of the longitude and latitude are 0.0002 ° and 0.0006 °, the height of the glide starting point is 55000m, the height of the terminal point is 49999.99m, the guidance precision of the aircraft is high, and the flight time is 1450s. As can be seen from FIG. 5, the high-precision robust controller can well realize the control of the local speed inclination angle and the track yaw angle value under the condition of guidance, and the flight end point speed is 4605m/s. As can be seen from fig. 6, as the speed gradually decreases during flight, the mach number gradually decreases. And the atmospheric density gradually increases as the altitude decreases. As can be seen from FIG. 7, the tracking control of the aircraft angle is good, the angle of attack gradually increases during flight, and the angle of attack at the tail end is 8.95 degrees; the sideslip angle is almost kept unchanged at 0 degrees, and a BTT flight mode is maintained; the roll angle gradually decreased and sharply increased toward the end point, at which the roll angle value was 28.8 °. As can be seen from fig. 8, the tracking conditions of the attack angle, the sideslip angle, and the roll angle are good, and except for the initial adjustment section and the approach to the target, the tracking errors of the attack angle, the sideslip angle, and the roll angle are large due to the rapid command change, and the tracking error in the whole process is extremely small. As can be seen from fig. 9, the aircraft rudder deflection angle changes more drastically during the initial adjustment phase, and after the glide state is stabilized, the elevator deflection angle increases slowly while the rudder deflection angle is maintained substantially around 0 °. As can be seen from fig. 10, the angular velocity variation of the aircraft is smoother and the global angular velocity variation does not exceed the maximum value.
The hypersonic aircraft glide section guidance and attitude control integrated design method based on the hypersonic aircraft glide section guidance and attitude control integrated design method can well meet the task requirements of guidance and attitude control of the aircraft glide section, the attitude control task requirements of the aircraft glide section are met, guidance and control accuracy is high, and the whole-course output change of the controlled variable is smooth.
The analysis shows that the hypersonic aircraft glide section guidance and attitude control integrated design method provided by the invention can well meet the task requirements of guidance and attitude control of the aircraft glide section and realize the task requirements of attitude control of the aircraft glide section, the guidance and control precision is high, and the whole-course output change of the control quantity is smooth. In order to better realize the flight task of the aircraft in the glide phase, the control scheme of the hypersonic aircraft glide phase guidance and attitude control integrated model is designed based on the block dynamic surface method, the task adaptability of the aircraft is improved, the effectiveness of the method in the glide phase is verified, and the problem of difficult coordination between the guidance and the control of the hypersonic aircraft is effectively solved.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention, and all modifications and equivalents of the present invention, which are made by the contents of the present specification and the accompanying drawings, or directly/indirectly applied to other related technical fields, are included in the scope of the present invention.
Claims (6)
1. A hypersonic aircraft glide phase guidance and attitude control integrated design method is characterized by comprising the following steps: obtaining the geocentric distance r of the current moving target T Latitude of moving object
Longitude lambda T The geocentric distance r of the current moving target is calculated T Latitude of moving target>
Longitude λ T As the input of the hypersonic aircraft glide section guidance and attitude control integrated control model, generating a control rudder deflection angle according to the control scheme of the hypersonic aircraft glide section guidance and attitude control integrated control model, inputting the control rudder deflection angle into a six-degree-of-freedom motion model of the hypersonic aircraft, and enabling the hypersonic aircraft to fly to a motion target to complete a flight task; the method comprises the following steps of constructing a six-degree-of-freedom motion model of the hypersonic aerocraft based on a geometric model of the hypersonic aerocraft as follows:
in the formula: v is the flight speed of the hypersonic aerocraft, theta is the local speed inclination angle, and sigma is the track yaw angle; omega x ，ω y ，ω z Are respectively hypersonic flightThree axes rotation angular velocity of the aircraft, m being the mass of the hypersonic aircraft, gamma V To a roll angle, I x ，I y ，I z Respectively, three-axis rotational inertia, L, D and N are aerodynamic force borne by the hypersonic aerocraft, respectively, lift force, resistance force and lateral force, M x ，M y ，M z The aerodynamic moment, the rolling moment, the yawing moment and the pitching moment, which are respectively borne by the aircraft, and the expressions of the aerodynamic force and the aerodynamic moment are respectively as follows:
wherein,
zero angle of attack lift coefficient;
Is the rate of change of lift coefficient with respect to angle of attack, C D Is a coefficient of resistance, C N Are the lateral force coefficients, which are known quantities; b and c are transverse lateral reference length and longitudinal reference length of the hypersonic aircraft respectively, and m is xβ Is the partial derivative of roll torque with respect to side slip angle beta, m yβ For the partial derivative of the yaw moment with respect to the sideslip angle β, <' >>
Is a partial derivative matrix of the pitch moment with respect to the slip angle beta>
For roll moment coefficient versus roll rudder deflection angle delta x Is based on the partial derivative of (4)>
Respectively yaw moment coefficient vs. rudder deflection angle delta y Partial derivative of (a), based on the partial derivative of (b)>
Are respectively a bowCoefficient of pitching moment on elevator delta z A partial derivative of (a), each of the aforementioned quantities being known quantities; the form of α is: α is the angle of attack value, ζ L For lift uncertainty term, α = [1 α 3 ] T ；
Respectively are three-axis moment uncertainty items; />
q＝0.5ρV 2 Is the dynamic pressure, ρ is the atmospheric density, is a known quantity, S 0 A reference area of the hypersonic aerocraft is a known quantity; the expression of the allowance in the six-degree-of-freedom motion model of the hypersonic aircraft is as follows:
mu is the gravitational constant, omega e Is the rotational angular velocity of the earth, J =1.5J 2 Is a harmonic coefficient of e The semimajor axis of the earth is a known quantity; r is the geocentric distance;
output x of guidance equation for control-oriented hypersonic aircraft range and azimuth error control 0 For local speed inclination theta and track yaw angle sigma, input
Is coefficient of lift C L The control-oriented guidance equation for the control of the course and the azimuth error of the hypersonic aerocraft is as follows:
in the formula:
assuming six-degree-of-freedom motion model of hypersonic flight vehicle with local velocity inclination angle theta and track yaw angle sigma as state quantitiesType of target instruction x 0C Is represented as follows:
x 0c ＝[θ c σ c ] T
wherein theta is c Representing the speed inclination angle command, sigma, of the hypersonic aircraft flight-support moving target c Representing a track yaw angle instruction value;
the control-oriented hypersonic deformable aircraft has the following kinematic equation around the center of mass:
in the formula,
wherein β is a slip angle, ζ 01 、ζ 02 Is ζ 0 Two components of (a);
the control-oriented hypersonic aircraft dynamic equation around the center of mass is as follows:
in the formula
The output of the control-oriented hypersonic aircraft mechanical equation of motion around the center of mass is x 2 The input is control rudder deflection angle u = [ delta ] of the hypersonic aircraft x δ y δ z ] T ；
The control method comprises the following steps of simultaneously establishing a control equation for controlling the course and the azimuth error of the hypersonic aerocraft facing the control, a kinematic equation around the center of mass of the hypersonic aerocraft facing the control and a kinematic equation around the center of mass of the hypersonic aerocraft facing the control, namely establishing a control model integrating the guidance and the attitude control of the glide section of the hypersonic aerocraft as follows:
the control scheme of the hypersonic aircraft glide section guidance and attitude control integrated control model is as follows:
in the above control scheme:
s 0 for the first dynamic plane defined, x 1 With its instruction value x 1d Difference of (a), x 1d For the first virtual control input, k 0 ＝diag(k 01 k 02 ) Given a positive gain constant; epsilon 01 And ε 02 A saturation function term gain to be given; sat (s, d) is a saturation function defined as:
s 1 for the second dynamic plane of definition, x 1 With its instruction value x 1d Difference of (a), x 2d A virtual control input for a second dynamic surface; k is a radical of 1 ＝diag(k 11 ,k 12 ,k 13 ) Given a positive gain constant; tau is 1 ＝diag(τ 11 ,τ 12 ,τ 13 ) For filteringThe time constant of the device; according to the virtual control quantity x 2d Is obtained such that x 1 Reach the expected instruction value x 1d Three-channel angular rate virtual input desired value omega xd ,ω yd ,ω zd ；
Definition s 2 Is a third dynamic plane, is x 2 With its instruction value x 2d U is the control input of the third dynamic surface; k is a radical of 2 ＝diag(k 21 k 22 k 23 ) Given a positive gain constant, τ 2 ＝diag(τ 21 ,τ 22 ,τ 23 ) Virtually inputting a desired value x according to the three-channel angular rate as a time constant of a filter 2d And obtaining a design rudder deflection angle input u to complete the stable control and guidance control of the aircraft attitude control system.
2. The hypersonic aircraft gliding section guidance and attitude control integrated design method as claimed in claim 1, characterized in that: the hypersonic aerocraft is an air-breathing general hypersonic aerocraft.
3. The hypersonic aircraft glide phase guidance and attitude control integrated design method according to claim 1 or 2, characterized in that: the speed dip angle instruction of the hypersonic aircraft to fly to the moving target is as follows:
track yaw angle command value sigma c Is expressed as
4. The hypersonic aircraft gliding section guidance and attitude control integrated design method as claimed in claim 3, characterized in that: by the hypersonic deformation aircraft moving around the mass center facing to the controlSolving x by chemical equation 1 ＝[α β γ V ] T Substituting into the aerodynamic expression of hypersonic deformable aircraft to calculate the lift coefficient C L And then the input of the motion equation of the hypersonic deformable aircraft relative to the moving target facing the control can be obtained
The input of the control-oriented hypersonic aerocraft around the center of mass kinematics equation is the three-axis rotation angular velocity vector x of the hypersonic aerocraft 2 ＝[ω x ω y ω z ] T 。
5. The utility model provides a hypersonic aircraft glide phase guidance and attitude control integration system which characterized in that includes:
a target information acquisition module for acquiring the geocentric distance r of the current moving target relative to the hypersonic aircraft T Latitude phi of moving object T Longitude λ, longitude T ；
The guidance module receives the target information acquired by the target information acquisition module, inputs the target information into a hypersonic aircraft glide section guidance and attitude control integrated control model which is pre-loaded on the guidance module, and generates a control rudder deflection angle according to a pre-designed control scheme of the hypersonic aircraft glide section guidance and attitude control integrated control model;
the attitude control-control module receives the control rudder deflection angle generated by the guidance module and inputs the control rudder deflection angle into a six-degree-of-freedom motion model of the supersonic aircraft pre-loaded on the attitude control-control module to complete the stable motion of the hypersonic aircraft and realize the tracking control of the guidance instruction;
the method comprises the following steps of constructing a six-degree-of-freedom motion model of the hypersonic aerocraft based on a geometric model of the hypersonic aerocraft as follows:
in the formula: v is the flight speed of the hypersonic aerocraft, theta is the local speed inclination angle, and sigma is the track yaw angle; omega x ，ω y ，ω z Respectively, the three-axis rotation angular velocity of the hypersonic aerocraft, m is the mass of the hypersonic aerocraft, gamma V To a roll angle, I x ，I y ，I z Respectively, three-axis rotational inertia, L, D and N are aerodynamic force borne by the hypersonic aerocraft, respectively, lift force, resistance force and lateral force, M x ，M y ，M z The aerodynamic moment, the rolling moment, the yawing moment and the pitching moment, which are respectively borne by the aircraft, and the expressions of the aerodynamic force and the aerodynamic moment are respectively as follows:
wherein,
zero angle of attack lift coefficient;
Is the rate of change of coefficient of lift with respect to angle of attack, C D Is a coefficient of resistance, C N Are the lateral force coefficients, which are known quantities; b and c are transverse lateral reference length and longitudinal reference length of the hypersonic aircraft respectively, and m is xβ Is the partial derivative of roll torque with respect to the slip angle beta, m yβ For the partial derivative of the yaw moment with respect to the sideslip angle β, <' >>
Is a partial derivative matrix of the pitch moment with respect to the slip angle beta>
For roll moment coefficient versus roll rudder deflection angle delta x Partial derivative of (a), based on the partial derivative of (b)>
Are respectively yaw moment coefficient pairAt rudder deflection angle delta y Partial derivative of (a), based on the partial derivative of (b)>
Respectively pitch moment coefficient for elevator delta z A partial derivative of (a), each of the above quantities being known quantities; the form of α is: α is the angle of attack value, ζ L For lift uncertainty term, α = [1 α = 3 ] T ；
Respectively are three-axis moment uncertainty items;
q＝0.5ρV 2 is dynamic pressure, ρ is the atmospheric density, a known quantity, S 0 A reference area of the hypersonic aerocraft is a known quantity; the expression of the allowance in the six-degree-of-freedom motion model of the hypersonic aerocraft is as follows:
mu is the gravitational constant, omega e Is rotational angular velocity of the earth, J =1.5J 2 Is a coefficient of harmonics, a e The semimajor axis of the earth is a known quantity; r is the geocentric distance;
output x of guidance equation for control-oriented hypersonic aircraft range and azimuth error control 0 For local speed inclination angle theta and track yaw angle sigma, input
Is coefficient of lift C L The control-oriented guidance equation for the control of the course and the azimuth error of the hypersonic aerocraft is as follows:
in the formula:
assuming a target command x of a six-degree-of-freedom motion model of the hypersonic aerocraft by taking a local velocity inclination angle theta and a track yaw angle sigma as state quantities 0C Is represented as follows:
x 0c ＝[θ c σ c ] T
wherein theta is c Speed dip angle instruction, sigma, representing hypersonic flight vehicle flight-offset moving target c Representing a track yaw angle instruction value;
the control-oriented hypersonic deformable aircraft has the following kinematic equation around the center of mass:
in the formula,
wherein β is a slip angle, ζ 01 、ζ 02 Is ζ 0 Two components of (a);
the control-oriented hypersonic aircraft dynamic equation around the center of mass is as follows:
in the formula
The output of the control-oriented hypersonic aircraft mechanical equation of motion around the center of mass is x 2 The input is control rudder deflection angle u = [ delta ] of the hypersonic aircraft x δ y δ z ] T ；
The control method comprises the following steps of simultaneously establishing a guidance equation for controlling the range and the azimuth error of the hypersonic aerocraft facing the control, a kinematics equation around the center of mass of the hypersonic aerocraft facing the control and a kinematics equation around the center of mass of the hypersonic aerocraft facing the control, namely establishing an integrated control model of the glide section guidance and the attitude control of the hypersonic aerocraft as follows:
the control scheme of the hypersonic aircraft glide section guidance and attitude control integrated control model is as follows:
in the above control scheme:
s 0 for the first dynamic plane defined, x 1 With its instruction value x 1d Difference of (a), x 1d For the first virtual control input, k 0 ＝diag(k 01 k 02 ) Given a positive gain constant; epsilon 01 And ε 02 A gain of a saturation function term to be given; sat (s, d) is a saturation function defined as:
γ Vd for the calculated roll angle command value, alpha v For the calculated roll angle command values, all pass x 1d Solving to obtain;
s 1 for the second dynamic plane of definition, x 1 With its instruction value x 1d Difference of (a), x 2d A virtual control input for a second dynamic surface; k is a radical of 1 ＝diag(k 11 ,k 12 ,k 13 ) Given a positive gain constant; tau. 1 ＝diag(τ 11 ,τ 12 ,τ 13 ) Is the time constant of the filter; according to the virtual control quantity x 2d Is obtained such that x 1 Reach the expected instruction value x 1d Three-channel angular rate virtual input desired value omega xd ,ω yd ,ω zd ；
Definition s 2 Is a third dynamic surface, is x 2 With its instruction value x 2d U is the control input of the third dynamic surface; k is a radical of 2 ＝diag(k 21 k 22 k 23 ) Given a positive gain constant, τ 2 ＝diag(τ 21 ,τ 22 ,τ 23 ) Virtually inputting a desired value x according to the three-channel angular rate as a time constant of a filter 2d And obtaining a design rudder deflection angle input u to complete the stable control and the guidance control of the aircraft attitude control system.
6. The utility model provides a hypersonic aircraft, includes the organism and establishes the airborne circuit board in the organism, be equipped with treater and memory on the airborne circuit board, its characterized in that: the memory stores a computer program, and the processor realizes the steps of the hypersonic flight vehicle glide slope guidance and attitude control integrated design method as claimed in any one of claims 1 to 4 when executing the computer program.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010327172.5A CN111399531B (en) | 2020-04-23 | 2020-04-23 | Hypersonic aircraft glide section guidance and attitude control integrated design method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010327172.5A CN111399531B (en) | 2020-04-23 | 2020-04-23 | Hypersonic aircraft glide section guidance and attitude control integrated design method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN111399531A CN111399531A (en) | 2020-07-10 |
| CN111399531B true CN111399531B (en) | 2023-04-07 |
Family
ID=71435324
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202010327172.5A Active CN111399531B (en) | 2020-04-23 | 2020-04-23 | Hypersonic aircraft glide section guidance and attitude control integrated design method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN111399531B (en) |
Families Citing this family (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113759956B (en) * | 2020-12-14 | 2024-05-28 | 北京天兵科技有限公司 | Flight trajectory design method for sub-orbit aircraft |
| CN113050682B (en) * | 2021-03-12 | 2022-04-22 | 中国人民解放军国防科技大学 | Missile boosting section correction guidance method and device and hypersonic aircraft |
| CN113418499B (en) * | 2021-05-13 | 2022-09-23 | 青岛杰瑞自动化有限公司 | Method and system for resolving roll angle of rotary aircraft |
| CN114265420B (en) * | 2021-12-09 | 2023-08-29 | 中国运载火箭技术研究院 | Guidance control integrated design method suitable for high dynamic and slow response control |
| CN114355777B (en) * | 2022-01-06 | 2023-10-10 | 浙江大学 | Dynamic gliding method and system based on distributed pressure sensor and sectional attitude control |
| CN114815888B (en) * | 2022-04-28 | 2023-05-19 | 哈尔滨工业大学 | Affine form guidance control integrated control method |
| CN115016291B (en) * | 2022-07-13 | 2023-11-10 | 西北工业大学 | Wind field information-based anti-interference attitude control system and method for aircraft |
| CN117518836B (en) * | 2024-01-04 | 2024-04-09 | 中南大学 | Robust deep reinforcement learning guidance control integrated method for variant aircraft |
Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5050086A (en) * | 1990-04-30 | 1991-09-17 | The Boeing Company | Aircraft lateral-directional control system |
| US5100084A (en) * | 1990-04-16 | 1992-03-31 | Space Systems/Loral, Inc. | Method and apparatus for inclined orbit attitude control for momentum bias spacecraft |
| US9639085B1 (en) * | 2015-08-05 | 2017-05-02 | The United States Of America As Represented By The Secretary Of The Air Force | Phugoid peaks trajectory for hypersonic glide vehicles |
| CN107102547A (en) * | 2017-05-10 | 2017-08-29 | 北京控制工程研究所 | A kind of RLV landing phase Guidance Law acquisition methods based on sliding mode control theory |
| CN107867387A (en) * | 2017-10-16 | 2018-04-03 | 中国航天空气动力技术研究院 | Outflow Waverider aircraft layout in a kind of |
| CN207758991U (en) * | 2017-12-25 | 2018-08-24 | 厦门大学 | It is a kind of can VTOL composite wing supersonic speed unmanned plane |
| CN110456810A (en) * | 2019-07-31 | 2019-11-15 | 湖北航天技术研究院总体设计所 | Hypersonic aircraft head body separation system and its design method, control method |
| CN110471456A (en) * | 2019-08-22 | 2019-11-19 | 中国人民解放军国防科技大学 | The guidance of hypersonic aircraft dive section, attitude control, deformation integrated control method |
| CN111003196A (en) * | 2019-12-31 | 2020-04-14 | 中国人民解放军国防科技大学 | Full-wave-rider aircraft and design method and system thereof |
-
2020
- 2020-04-23 CN CN202010327172.5A patent/CN111399531B/en active Active
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5100084A (en) * | 1990-04-16 | 1992-03-31 | Space Systems/Loral, Inc. | Method and apparatus for inclined orbit attitude control for momentum bias spacecraft |
| US5050086A (en) * | 1990-04-30 | 1991-09-17 | The Boeing Company | Aircraft lateral-directional control system |
| US9639085B1 (en) * | 2015-08-05 | 2017-05-02 | The United States Of America As Represented By The Secretary Of The Air Force | Phugoid peaks trajectory for hypersonic glide vehicles |
| CN107102547A (en) * | 2017-05-10 | 2017-08-29 | 北京控制工程研究所 | A kind of RLV landing phase Guidance Law acquisition methods based on sliding mode control theory |
| CN107867387A (en) * | 2017-10-16 | 2018-04-03 | 中国航天空气动力技术研究院 | Outflow Waverider aircraft layout in a kind of |
| CN207758991U (en) * | 2017-12-25 | 2018-08-24 | 厦门大学 | It is a kind of can VTOL composite wing supersonic speed unmanned plane |
| CN110456810A (en) * | 2019-07-31 | 2019-11-15 | 湖北航天技术研究院总体设计所 | Hypersonic aircraft head body separation system and its design method, control method |
| CN110471456A (en) * | 2019-08-22 | 2019-11-19 | 中国人民解放军国防科技大学 | The guidance of hypersonic aircraft dive section, attitude control, deformation integrated control method |
| CN111003196A (en) * | 2019-12-31 | 2020-04-14 | 中国人民解放军国防科技大学 | Full-wave-rider aircraft and design method and system thereof |
Non-Patent Citations (1)
| Title |
|---|
| 顾攀飞等.高超声速飞行器再入自适应容错制导控制一体化设计.南京航空航天大学学报.2018,第第50卷卷(第第50卷期),全文. * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN111399531A (en) | 2020-07-10 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN111399531B (en) | Hypersonic aircraft glide section guidance and attitude control integrated design method | |
| Lozano | Unmanned aerial vehicles: Embedded control | |
| CN110471456B (en) | Hypersonic aircraft diving section guidance, attitude control and deformation integrated control method | |
| Jung et al. | Development and application of controller for transition flight of tail-sitter UAV | |
| CN102073755B (en) | Motion control simulation method for near-space hypersonic aircraft | |
| CN105159305B (en) | A kind of quadrotor flight control method based on sliding moding structure | |
| CN103760905B (en) | Based on fuzzy feedforward list rotor unmanned helicopter attitude nonlinear robust control method | |
| CN108803639A (en) | A kind of quadrotor flight control method based on Backstepping | |
| CN105425812B (en) | Unmanned aerial vehicle automatic landing trajectory control method based on dual models | |
| CN106707749B (en) | A kind of control method for bionic flapping-wing flying robot | |
| CN107491083B (en) | Four-rotor-wing autonomous landing method based on saturation self-adaptive sliding mode control | |
| Jung et al. | A comprehensive flight control design and experiment of a tail-sitter UAV | |
| CN112198885B (en) | Unmanned aerial vehicle control method capable of meeting autonomous landing requirement of maneuvering platform | |
| CN107678442B (en) | Dual-model-based four-rotor autonomous landing control method | |
| Garcia et al. | Attitude stabilization with real-time experiments of a tail-sitter aircraft in horizontal flight | |
| Garcia et al. | Robust geometric navigation of a quadrotor UAV on SE (3) | |
| CN109582039A (en) | A kind of optimal reconfiguration method under the J2 perturbation using Relative Navigation information | |
| Hegde et al. | Transition flight modeling and robust control of a VTOL unmanned quad tilt-rotor aerial vehicle | |
| Zhang et al. | Wind field disturbance analysis and flight control system design for a novel tilt-rotor UAV | |
| Lin et al. | Standoff tracking of a ground target based on coordinated turning guidance law | |
| Cheng et al. | Hover-to-cruise transition control for high-speed level flight of ducted fan UAV | |
| Zou et al. | Optimal hovering control of a tail-sitter via model-free fast terminal slide mode controller and cuckoo search algorithm | |
| CN114265420B (en) | Guidance control integrated design method suitable for high dynamic and slow response control | |
| Yao et al. | Trajectory tracking controller based on PID-NLADRC | |
| Ashari et al. | Flight trajectory control system on fixed wing UAV using linear quadratic regulator |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |