CN113955113B - Miniature aircraft suitable for high-speed operation and control method - Google Patents
Miniature aircraft suitable for high-speed operation and control method Download PDFInfo
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- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 claims abstract description 13
- 238000011217 control strategy Methods 0.000 claims description 24
- 238000005096 rolling process Methods 0.000 claims description 6
- 230000009466 transformation Effects 0.000 claims description 6
- 238000006073 displacement reaction Methods 0.000 claims description 3
- 239000011159 matrix material Substances 0.000 claims description 3
- 238000005381 potential energy Methods 0.000 claims description 3
- 230000004044 response Effects 0.000 claims description 3
- 230000000087 stabilizing effect Effects 0.000 claims description 3
- 238000013016 damping Methods 0.000 claims 1
- 238000012423 maintenance Methods 0.000 abstract description 5
- 239000000463 material Substances 0.000 description 4
- 229920000271 Kevlar® Polymers 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 239000000835 fiber Substances 0.000 description 2
- 239000011152 fibreglass Substances 0.000 description 2
- 239000004761 kevlar Substances 0.000 description 2
- 239000002131 composite material Substances 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C39/00—Aircraft not otherwise provided for
- B64C39/02—Aircraft not otherwise provided for characterised by special use
- B64C39/028—Micro-sized aircraft
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/04—Helicopters
- B64C27/08—Helicopters with two or more rotors
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/04—Helicopters
- B64C27/08—Helicopters with two or more rotors
- B64C27/10—Helicopters with two or more rotors arranged coaxially
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/04—Helicopters
- B64C27/12—Rotor drives
- B64C27/14—Direct drive between power plant and rotor hub
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/02—Aircraft characterised by the type or position of power plants
- B64D27/24—Aircraft characterised by the type or position of power plants using steam or spring force
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
- B64U10/00—Type of UAV
- B64U10/10—Rotorcrafts
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
- B64U50/00—Propulsion; Power supply
- B64U50/10—Propulsion
- B64U50/19—Propulsion using electrically powered motors
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- 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
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- Aviation & Aerospace Engineering (AREA)
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- Combustion & Propulsion (AREA)
- Radar, Positioning & Navigation (AREA)
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Abstract
The invention provides a micro aircraft suitable for high-speed operation and a control method thereof, wherein the aircraft comprises a flight control module and a wing-shaped fuselage capable of generating vertical lift force during flight; a vertical duct is arranged in the middle of the machine body in the overlooking direction; a first rotor wing group capable of providing vertical lift force is arranged in the duct; a second rotor wing group capable of providing horizontal thrust is arranged at the rear edge of the tail part of the wing-shaped fuselage; the invention has the advantages of long service life, good flight control performance, low maintenance cost and the like.
Description
Technical Field
The invention relates to the technical field of micro aircrafts, in particular to a micro aircraft suitable for high-speed operation and a control method.
Background
Currently, multi-rotor micro-aircraft are proposed for many important fields, and the key infrastructure of the existing micro-aircraft is reluctant to cope with the complex environment, but the structure of the existing micro-aircraft still can be out of order and the maintenance cost is high. The stability of operation and control is greatly reduced when the high-speed operation is carried out. Designers must therefore provide a miniature aircraft with a long life, high operating speed, safer flight and low maintenance costs, so as to enable its use in important fields.
Disclosure of Invention
The invention provides a micro aircraft suitable for high-speed operation and a control method thereof, and the micro aircraft has the advantages of long service life, good flight control performance, low maintenance cost and the like.
The invention adopts the following technical scheme.
A micro-aircraft suitable for high-speed operation, the aircraft comprising a flight control module and a wing-shaped fuselage capable of generating vertical lift during flight; a vertical duct is arranged in the middle of the machine body in the downward viewing direction; a first rotor wing group capable of providing vertical lift force is arranged in the duct; and a second rotor group capable of providing horizontal thrust is arranged at the rear edge of the tail part of the wing-shaped fuselage.
The first rotor wing group comprises a first rotor wing (1) and a second rotor wing (2) which are vertically and coaxially arranged; the second rotor group is horizontal to setting up and third rotor (3), fourth rotor (4) for fuselage axis symmetric distribution.
The second rotor group is a rotor group capable of controlling the horizontal flight direction of the aircraft.
The first rotor wing, the second rotor wing, the third rotor wing and the fourth rotor wing are all driven by motors; the rotating directions of the first rotor wing and the second rotor wing are opposite; the rotating directions of the third rotor and the fourth rotor are opposite; the aircraft further comprises an electronic speed controller and a battery; the lower part of the aircraft body is provided with a lander.
The rotors in the first rotor group are molded by glass fiber reinforced plastics; the rotors in the second rotor group are formed by Kevlar fibers; the rear edge of the wing-shaped fuselage tail is a wing-shaped thin end.
The overall dimensions of the aircraft do not exceed 15 centimetres.
A control method of a miniature aircraft suitable for high-speed operation is used for the miniature aircraft, the miniature aircraft is a three-rotor helicopter, a dynamic model used by a flight control module is based on a three-rotor helicopter control model using an Euler-Lagrange method, and generalized coordinates describing the position and the orientation of a rotorcraft in the control model are as follows:
q T = (x, y, z, ψ, θ, Φ) formula one;
wherein x, y and z represent the position of the center of mass of the three-rotor helicopter relative to an inertia system I; psi, theta, phi are the three euler yaw angles, pitch angles, and roll angles, representing the direction of the rotorcraft;
the control model is divided into translation coordinates and rotation coordinates, and is expressed by a formula
The translational kinetic energy of a rotorcraft is formulated as
Wherein m represents the mass of the rotorcraft; the rotary kinetic energy of a rotorcraft is formulated as
Wherein J represents moment of inertia; the gravitational potential energy of the rotorcraft is as follows:
u = mgz formula six;
from the above, the Lagrangian function of the rotorcraft control model is formulated as
L＝T tra +T rot -U formula seven;
the rotorcraft dynamics model is derived from the Euler-Lagrange equation and the external generalized force F, as follows:
wherein τ is a generalized moment, F ξ Translational thrust exerted on the rotorcraft due to flight control module inputs; the force acting on the motor rotor of the frame of the three-rotor helicopter body is expressed by a formula as follows:
wherein u is defined as
u＝f 1 +f 2 +f 3 cos α formula eleven;
wherein f is 1 ，f 2 ，f 3 Thrust generated by the three rotors respectively; alpha is an included angle between the thrust and the horizontal plane;
in the formula (f) i Thrust, k, generated for the ith motor i >0 is a constant, ω i Is the angular velocity of motor i; translational force F ξ And
have the following relations
Wherein R is a transformation matrix representing the direction of the rotorcraft, and R is expressed as
C and s in the formula respectively represent cos and sin;
the generalized moment η is expressed as:
τ＝[τ φ τ θ τ ψ ] T a formula fourteen;
wherein
τ φ ＝(f 2 -f 1 )l 1 A formula fifteen;
τ θ ＝-f 3 l 2 cosα+m 3 gl 2 +(f 2 +f 1 )l 3 -(m 1 +m 2 )gl 3 sixthly, a formula is formed;
τ ψ ＝f 3 l 2 sin α formula seventeen;
l 1 、l 2 、l 3 the force arms of three rotor motor rotors in the model are respectively;
combining xi and eta, decomposing the Euler Lagrange equation into a kinetic equation under a translation xi coordinate system and a kinetic equation under a rotation eta coordinate system, and expressing the kinetic equations as follows by formulas:
by combining the above formulas
The coriolis term, gyro term, and centrifuge term are defined by the formula:
the dynamic model of three rotor motor rotors is expressed by formula
When the three-rotor helicopter is in a hovering state, the control strategy for stabilizing the aircraft by the three-rotor helicopter control model is as follows:
the input variables of the control model are adjusted to
Transformation of kinetic model
Where x and y are coordinates in the horizontal plane, z is the vertical position, ψ is the yaw angle about the z-axis, θ is the pitch angle about the y-axis, and φ is the roll angle about the x-axis;
the control strategy controls the total thrust represented by u, an
Roll, pitch and yaw moments, respectively, to achieve control of the aircraft;
control of the vertical position z may be achieved by using the following control inputs:
controlling fly in control strategyHeight and yaw of the vehicle, in which a z1 、a z2 Is a normal number, z d The height required to be controlled; the yaw angle position can be controlled by an equation of
Derived by
The control strategy adjusts a controller parameter a z1 And a ψ1 To obtain good damped stable response of altitude and yaw angular displacement, respectively; controller parameter a ψ1 And a z2 Can be adjusted to improve tracking performance;
the time margin of the above equation may be such that r is the amount of time r that the control strategy controls the roll of the aircraft 1 →0；ψ→ψ d (ii) a Is further simplified to obtain
Tan phi is approximately equal to phi since phi is sufficiently small; then there is
The control strategy has a nested saturation control law expressed by formula
Wherein σ i(s) Is a saturation function defined as:
the control law can guarantee
Converge to zero;
when the control strategy is used for pitch control of an aircraft,
the control strategy can be simplified to
The same method as the previous method for controlling the rolling and rolling angle is adopted, and the formula is
The control law can guarantee
Converging to zero.
The invention adopts the structure of the miniature duct three rotors, is matched with a proper composite material and a unique control method, so that the aircraft has the characteristics of long service life, high-speed operation, safer flight and low maintenance cost. The structure is more reliable and the control is more stable during high-speed flight. So as to enable its application in particular fields.
The invention provides a novel multi-rotor aircraft structure which is good in flight control performance; the design consists of three groups of rotors, wherein one group of coaxial rotors is used for vertical take-off and landing, and the other two rotors are used for generating forward speed and controlling the horizontal direction; the fuselage is designed as a wing profile and also provides a certain vertical lift. The design makes the structure more compact, and the device can be more suitable for the environment of high-speed flight. Meanwhile, the invention improves the structure of the traditional aircraft, so that the micro aircraft has better controllability in high-speed work, reduces the quality of the whole aircraft and improves the efficiency
According to the scheme, the material suitable for the high-speed operation of the micro aircraft can be determined through fluid-solid coupling simulation, namely the most important index concerned when the material is optimized aiming at the deformation of the micro aircraft is ensured, so that the deformation of the whole body and the deformation of the rotor wing of the micro aircraft are kept in a controllable range under the high-speed operation; mainly the selection of the material of the rotor wing under high-speed flight; the materials of the rotors of each group should not be the same, since the functions of the rotors of each group are different.
The invention provides a nonlinear controller based on a nested saturation technology by establishing a dynamic model of a system through an Euler-Lagrange method and adopting a nonlinear control strategy. The controller enables the closed loop system to be globally stable. The three-rotor helicopter can safely and autonomously operate. The proposed nonlinear controller performs better than a classical state feedback linear controller.
Drawings
The invention is described in further detail below with reference to the following figures and detailed description:
FIG. 1 is a schematic of the present invention;
FIG. 2 is a schematic top view of the present invention;
FIG. 3 is a schematic diagram of a three-rotor helicopter control model of the present invention;
FIG. 4 is a schematic diagram of the forces applied to a three-rotor helicopter control model of the present invention;
in the figure: 1-a first rotor; 2-a second rotor; 3-a third rotor; 4-a fourth rotor; 5, a motor; 6-wing fuselage; 7-a flight control module; 8-a lander; 9-a duct; 10-a first rotor set; 11-second rotor set.
Detailed Description
As shown in the figure, a micro-aircraft suitable for high-speed operation comprises a flight control module 7 and a wing-shaped fuselage 6 capable of generating vertical lift force during flight; a vertical duct 9 is arranged in the middle of the machine body in the downward view direction; a first rotor wing group 10 capable of providing vertical lift force is arranged in the duct; and a second rotor group 11 capable of providing horizontal thrust is arranged at the rear edge of the tail part of the wing-shaped fuselage.
The first rotor wing group comprises a first rotor wing 1 and a second rotor wing 2 which are vertically and coaxially arranged; and the second rotor wing group comprises a third rotor wing 3 and a fourth rotor wing 4 which are horizontally arranged and symmetrically distributed relative to the axis of the fuselage.
The second rotor group is a rotor group capable of controlling the horizontal flight direction of the aircraft.
The first rotor wing, the second rotor wing, the third rotor wing and the fourth rotor wing are all driven by a motor 5; the rotating directions of the first rotor wing and the second rotor wing are opposite; the rotating directions of the third rotor and the fourth rotor are opposite; the aircraft further comprises an electronic speed controller and a battery; the lower part of the aircraft fuselage is provided with a lander 8.
The rotors in the first rotor group are molded by glass fiber reinforced plastics; the rotors in the second rotor group are formed by Kevlar fibers; the rear edge of the wing-shaped fuselage tail is a wing-shaped thin end.
The overall dimensions of the aircraft do not exceed 15 centimetres.
A control method of a miniature aircraft suitable for high-speed operation is used for the miniature aircraft, the miniature aircraft is a three-rotor helicopter, a dynamic model used by a flight control module is based on a three-rotor helicopter control model using an Euler-Lagrange method, and generalized coordinates describing the position and the orientation of a rotorcraft in the control model are as follows:
q T = (x, y, z, ψ, θ, Φ) formula one;
wherein x, y and z represent the position of the three-rotor helicopter mass center relative to an inertia system I; psi, theta, phi are the three euler yaw angles, pitch angles, and roll angles, representing the direction of the rotorcraft;
the control model is divided into translation coordinates and rotation coordinates and is expressed by a formula
The translational kinetic energy of a rotorcraft is formulated as
Wherein m represents the mass of the rotorcraft; the rotary kinetic energy of a rotorcraft is formulated as
Wherein J represents moment of inertia; the gravitational potential energy of the rotorcraft is as follows:
u = mgz formula six;
from the above, the Lagrangian function of the rotorcraft control model is formulated as
L＝T tra +T rot -U formula seven;
the rotorcraft dynamics model is derived from the Euler-Lagrange equation and the external generalized force F, as follows:
wherein τ is a generalized moment, F ξ Translational thrust exerted on the rotorcraft due to flight control module inputs;
the force acting on the motor rotor of the frame of the three-rotor helicopter body is expressed by a formula as follows:
wherein u is defined as
u＝f 1 +f 2 +f 3 cos α formula eleven;
wherein f is 1 ，f 2 ，f 3 Thrust generated by the three rotors respectively; alpha is an included angle between the thrust and the horizontal plane;
in the formula, f i Thrust generated for the i-th motor, k i >0 is a constant, ω i Is the angular velocity of motor i; translational force F ξ And with
Has the following relationship
Wherein R is a transformation matrix representing the direction of the rotorcraft, and R is expressed by
C and s in the formula respectively represent cos and sin;
generalized moment η is expressed as:
τ＝[τ φ τ θ τ ψ ] T a formula fourteen;
wherein
τ φ ＝(f 2 -f 1 )l 1 A formula fifteen;
τ θ ＝-f 3 l 2 cos α+m 3 gl 2 +(f 2 +f 1 )l 3 -(m 1 +m 2 )gl 3 sixthly, a formula is formed;
τ ψ ＝f 3 l 2 sin α formula seventeen;
l 1 、l 2 、l 3 the force arms of three rotor motor rotors in the model are respectively;
combining xi and eta, decomposing the Euler Lagrange equation into a kinetic equation under a translation xi coordinate system and a kinetic equation under a rotation eta coordinate system, and expressing the equations as follows:
by combining the above formulas
The coriolis term, gyro term, and centrifuge term are defined by the formula:
the dynamic model of three rotor motor rotors is expressed by formula
When the three-rotor helicopter is in a hovering state, the control strategy for stabilizing the aircraft by the three-rotor helicopter control model is as follows:
the input variables of the control model are adjusted to
Transformation of kinetic model
Where x and y are coordinates in the horizontal plane, z is the vertical position, ψ is the yaw angle about the z-axis, θ is the pitch angle about the y-axis, and φ is the roll angle about the x-axis;
the control strategy controls the total thrust represented by u, and
roll, pitch and yaw moments, respectively, to achieve control of the aircraft;
control of the vertical position z may be achieved by using the following control inputs:
in the control strategy controlling the altitude and yaw of the aircraft, where a z1 、a z2 Is a normal number, z d The height required to be controlled; the yaw angle position can be controlled by an equation of
Derived by
The control strategy adjusts a controller parameter a z1 And a ψ1 To obtain good damped stable responses for altitude and yaw angular displacements, respectively; controller parameter a ψ1 And a z2 Can be adjusted to improve tracking performance;
the time margin of the above equation may be such that r is the amount of time r that the control strategy controls the roll of the aircraft 1 →0；ψ→ψ d (ii) a Is further simplified to obtain
Tan phi is approximately equal to phi since phi is sufficiently small; then there is
The control strategy has a nested saturation control law expressed by formula
Wherein σ i(s) Is a saturation function defined as:
the control law may guarantee that y,
φ，
converge to zero;
when the control strategy is used for pitch control of an aircraft,
the control strategy can be simplified to
The same method as the previous method for controlling the rolling and rolling angles is adopted, and the formula is
The control law may guarantee that x,
θ，
converging to zero.
Claims (1)
1. A method for controlling a miniature aircraft suitable for high-speed operation is characterized by comprising the following steps: the miniature aircraft is a three-rotor helicopter, a dynamics model used by a flight control module is based on a three-rotor helicopter control model using an Euler-Lagrange method, and in the control model, generalized coordinates describing the position and the orientation of the rotor helicopter are as follows:
q T = (x, y, z, ψ, θ, Φ) formula one;
wherein x, y and z represent the position of the three-rotor helicopter mass center relative to an inertia system I; psi, theta, phi are the three euler yaw angles, pitch angles, and roll angles, representing the direction of the rotorcraft;
the control model is divided into translation coordinates and rotation coordinates, and is expressed by a formula
The translational kinetic energy of a rotorcraft is formulated as
Wherein m represents the mass of the rotorcraft; the rotary kinetic energy of a rotorcraft is formulated as
Wherein J represents moment of inertia; the gravitational potential energy of the rotorcraft is as follows:
u = mgz formula six;
from the above, the Lagrangian function of the rotorcraft control model is formulated as
L＝T tra +T rot -U formula seven;
the rotorcraft dynamics model is derived from the Euler-Lagrange equation and the external generalized force F, as follows:
wherein τ is the generalized moment, F ξ Translational thrust exerted on the rotorcraft due to flight control module inputs;
the force acting on the motor rotor of the frame of the three-rotor helicopter body is expressed by a formula as follows:
wherein u is defined as
u＝f 1 +f 2 +f 3 cos α formula eleven;
wherein f is 1 ，f 2 ，f 3 Thrust generated by the three rotors respectively; alpha is an included angle between the thrust and the horizontal plane;
in the formula (f) i Thrust generated for the i-th motor, k i > 0 is a constant, omega i Is the angular velocity of motor i; translation force F ξ And with
Have the following relations
Wherein R is a transformation matrix representing the direction of the rotorcraft, and R is expressed by
C and s in the formula respectively represent cos and sin;
the generalized moment η is expressed as:
τ＝[τ φ τ θ τ ψ ] T a fourteen formula;
wherein
τ φ ＝(f 2 -f 1 )l 1 A formula fifteen;
τ θ ＝-f 3 l 2 cosα+m 3 gl 2 +(f 2 +f 1 )l 3 -(m 1 +m 2 )gl 3 sixthly, a formula is formed;
τ ψ ＝f 3 l 2 sin α formula seventeen;
l 1 、l 2 、l 3 the force arms of the rotors of the three rotor wing motors in the model are respectively;
combining xi and eta, decomposing the Euler Lagrange equation into a kinetic equation under a translation xi coordinate system and a kinetic equation under a rotation eta coordinate system, and expressing the kinetic equations as follows by formulas:
by combining the above formulas
The coriolis term, gyro term, and centrifuge term are defined by the equations:
the dynamic model of three rotor motor rotors is expressed by formula
When the three-rotor helicopter is in a hovering state, the control strategy for stabilizing the aircraft by the three-rotor helicopter control model is as follows:
the input variable of the control model is adjusted to
Transformation of kinetic model
Where x and y are coordinates in the horizontal plane, z is the vertical position, ψ is the yaw angle about the z-axis, θ is the pitch angle about the y-axis, and φ is the roll angle about the x-axis;
the control strategy controls the total thrust represented by u, and
roll, pitch and yaw moments, respectively, to achieve control of the aircraft;
control of the vertical position z may be achieved by using the following control inputs:
in the control strategy controlling the altitude and yaw of the aircraft, where a z1 、a z2 Is a normal number, z d The height required to be controlled; the yaw angle position can be controlled by an equation of
Derived by
The control strategy adjusts a controller parameter a z1 And a ψ1 To respectively obtainObtaining good damping stable response of height and yaw angular displacement; controller parameter a ψ1 And a z2 Can be adjusted to improve tracking performance;
the time margin of the above equation may be such that r is the amount of time that the control strategy controls the roll of the aircraft 1 →0；ψ→ψ d (ii) a Is further simplified to obtain
Tan phi is approximately equal to phi since phi is sufficiently small; then there is
The control strategy has a nested saturation control law expressed by formula
Wherein σ i(s) Is a saturation function defined as:
the control law may guarantee that y,
φ，
converge to zero;
when the control strategy is used for pitch control of an aircraft,
the control strategy can be simplified to
The same method as the previous method for controlling the rolling and rolling angle is adopted, and the formula is
The control law may guarantee that x,
θ，
converge to zero.
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| EP1901153A1 (en) * | 2006-09-12 | 2008-03-19 | OFFIS e.V. | Control system for unmanned 4-rotor-helicopter |
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