CN210027899U - Four rotor crafts of biax slope - Google Patents
Four rotor crafts of biax slope Download PDFInfo
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- CN210027899U CN210027899U CN201920653997.9U CN201920653997U CN210027899U CN 210027899 U CN210027899 U CN 210027899U CN 201920653997 U CN201920653997 U CN 201920653997U CN 210027899 U CN210027899 U CN 210027899U
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Abstract
The utility model provides a double-shaft inclined four-rotor aircraft, the aircraft body comprises a plurality of power arms uniformly arranged at the peripheral edge of the aircraft body, and a sensor unit, a controller unit, a control distributor unit, a vehicle dynamics system module and a wireless communication module which are arranged at the aircraft body; the wireless communication module can establish wireless communication with the ground control station; the power arm comprises a rotor and a tilting mechanism; the rotor wing is driven by a driving motor at a motor fixing seat supported by the tilting mechanism; the tilting mechanism is connected with the controller unit; the controller unit adjusts the rotor position and rotor orientation via the tilt mechanism to adjust the flight of the aircraft; the utility model can automatically adjust the rotating speed and the rotor tilting angle of the four rotors according to the input instruction, and feed back the self flying state by utilizing the stability of the closed-loop control system, thereby automatically completing the self-flying regulation and control work and achieving the purposes of flying pitching, rolling and yawing; the high efficiency and the high reliability of the flight control are ensured.
Description
Technical Field
The utility model belongs to the technical field of small-size unmanned vehicles technique and specifically relates to a four rotor crafts of biax slope.
Background
The double-shaft inclined rotor wing configuration refers to the rotor wing configuration that a certain included angle is formed between the rotating plane of the rotor wing and the plane of the aircraft body through rotation of the supporting arm and the motor fixing seat. Most of the traditional four-rotor aircraft's support arm and motor fixing base are fixed mounting on the organism, rotary motion is all done to four rotors of aircraft in same plane, the rotor can only realize every single move, roll and yaw motion through the rotational speed that changes driving motor, need four variable speed driving motor and regulation control system, and the function of meeting an emergency is relatively poor and the in-process is steady inadequately, if the aircraft needs in time to accomplish above-mentioned 3 kinds of movements, first method is that need install driver additional in aircraft nose department or aircraft tail department installs driver and tail pipe additional and reach the flight requirement. However, as the number of devices increases, the weight of the aircraft body increases, more rotors or longer propellers are needed to maintain the hovering motion of the aircraft, and although the reliability of the aircraft is enhanced by the method of adding the rotors, the flexibility is difficult to guarantee. The second method is the method of using variable pitch propellers, which seeks to maintain the weight of the aircraft while maintaining high flexibility, but which requires the use of more complex systems of propellers and servomotors, is complex to implement, has a high failure rate, and may be accompanied by personnel injury in the event of failure. The method for realizing the flexible flying motion of the multi-rotor aircraft has the advantages of low lifting force and weight, complex structure and low efficiency, and greatly restricts the wide application of the multi-rotor aircraft in various fields.
Disclosure of Invention
The utility model provides a four rotor crafts of biax slope, it can be according to the rotational speed and the rotor angle of verting of input command automatically regulated four rotors, can carry out remote control, can utilize closed-loop control system's stability according to different flight environmental conditions again, feed back the flight state of self, accomplish the work of self-flying regulation and control automatically, it can reach the purpose of flight every single move, roll over and driftage through the rotor that verts; the high efficiency and the high reliability of the flight control are ensured, and the size of the aircraft can be reduced, so that the flexibility is improved.
The utility model adopts the following technical scheme.
A double-shaft tilting four-rotor aircraft is characterized in that a fuselage (1) of the aircraft comprises a plurality of power arms uniformly arranged on the periphery of the fuselage, a sensor unit (2), a controller unit (3), a control distributor unit (4), a vehicle dynamics system module (5) and a wireless communication module (6) which are arranged on the fuselage; the wireless communication module can establish wireless communication with a ground control station (7);
the power arm comprises a rotor (108) and a tilt mechanism (115); the rotor is driven by a drive motor (107) at a motor mount (105) supported by the tilt mechanism; the tilting mechanism is connected with the controller unit; the controller unit adjusts rotor position and rotor orientation via a tilt mechanism to adjust the flight of the aircraft.
The number of the power arms is four.
The tilting mechanism is connected with a servo motor (106) at the machine body by the starting end of a supporting arm (103), and the tail end of the supporting arm is hinged with the middle part of a motor fixing seat; the servo motor drives the supporting arm to rotate so as to enable the rotor wing to tilt; a fixing block (111) for fixing the steering engine (109) is arranged at the supporting arm; the steering wheel links to each other with the lower extreme of motor fixing base through link mechanism (104), the steering wheel drives the swing of motor fixing base through link mechanism makes the rotor produce yawing force.
The link mechanism is a parallelogram mechanism, the inclination angle α of the rotor wing can be adjusted between 0 degree and 90 degrees, and the inclination angle α can be 0 degree but can not be 90 degrees.
The fuselage further comprising a landing gear (101) and an electrical equipment bay (102); a gravity sensor (112), a gyroscope (113), a remote sensor (114) and a plurality of batteries (116) are arranged in the electrical equipment cabin; the output shaft of the servo motor and the contact position of the rotor wing supporting arm and the electrical equipment cabin (102) are connected through a bearing (110).
The sensor unit (2) is responsible for the connection work between the sensor unit and the controller unit (3), receives the measurement information of the sensor unit and converts the received measurement information into an electric signal according to a certain rule and outputs the electric signal to the controller unit (3); the controller unit (3) performs corresponding processing and sends out control command signals after finishing receiving the transmission information of the sensor unit (2) and the wireless communication module (5);
the control distributor unit (4) completes the distribution work of the control command signals of the controller unit (3), and distributes the received command information of the controller unit (3) to four servo motors (106), four driving motors (107) and four steering engines (109) of the aircraft in order to realize the adjustment of the flight state;
the vehicle dynamics system module (5) comprises an actuator unit (501) and a sensor unit (502), the sensor unit (502) comprises a gravity sensing module (503), a perturbation sensing module (504), a gravity sensor (111), a gyroscope (113) and a remote sensor (114), and executes an instruction for controlling the distributor unit (4), and the vehicle dynamics system module measures real-time gravity and perturbation external force through the sensor unit (502) and feeds back the measured real-time gravity and perturbation external force to the controller unit (3) to form a closed-loop system;
the wireless communication module (6) comprises a transmitting module (601) and a receiving module (602), the working modes of the transmitting module (601) and the receiving module (602) can be switched simultaneously and do not interfere with each other, the transmitting module (602) is connected with the controller unit (3), and the receiving module (602) is in communication with the ground control station (7) in a wireless mode so that the ground control station (7) can execute azimuth instruction operation on the aircraft.
The sensor unit (2) is an IMU comprised of an SBG system and IG-500, and has an embedded processor for outputting filtered attitude and position data; said controller unit (3) is based on a PD controller with 3 SISDs, which processes the chip set maximum criterion of Digilent of the board for reading the signal; the controller unit (2) reads serial information at MAX32 from output signals of the sensor unit (2), and commands the servo motor (106), the driving motor (107) and the steering engine (109) by sending out PWM signals;
and a receiving module (602) in the wireless communication module (6) is communicated with the ground control station (7) through a Spektrum satellite receiver and a DX6i RC transmitter.
The absolute rotating speed of the output shaft of the driving motor is the vector sum of the angular speed of the aircraft in the three directions of XYZ, the rotating speed of the steering engine and the rotating speed of the servo motor, the physical characteristics of four supporting arms of the aircraft are consistent, one supporting arm is taken for stress analysis, and the absolute rotating speed of the output shaft of the driving motor is the absolute rotating speed of the output shaft of the steering engine and the absolute rotating speed of the servo motor
Angular acceleration of
Vector in the formulaiij ijj ikjThe unit vector representing a reference system j is represented by a reference system i, p, q and r respectively represent the angular velocity of the aircraft in three directions of a reference coordinate system XYZ, η and gamma represent the tilting angle of a double shaft during servo motion, η 'gamma' represents the tilting angular velocity of the double shaft during servo motion, and omega represents the rotating speed of a driving motor;
using the Euler equation3M＝3I3α+3ω×3I3ω (formula 3),
calculating the torque of the output shaft of the drive motor3M, then using the rotation matrix R3to1Calculation of the aircraft's own moments in a reference system1MGγroThe calculation formula is as follows:
the thrust torque of the connecting rod structure is as follows:
wherein L is the length of the parallelogram mechanism; h is the height of the parallelogram mechanism;
the thrust and torque coefficients of the rotor are defined as:
T＝ρA(ΩR)2CT(6) and Q ═ ρ a (Ω R)2RCQ(equation 7);
in the formula CTAnd CQThe thrust and the torque coefficient of the connecting rod structure.
The total thrust moment of the aircraft is then:
the absolute stress condition of the aircraft in six degrees of freedom is as follows:
x, Y and Z are respectively expressed as the absolute acting force of the aircraft on XYZ axes; l, M and N are respectively expressed as the absolute acting moments of the aircraft on XYZ axes; u, v and w are respectively expressed as the angular velocities of the whole aircraft on XYZ axes; when the aircraft is under the comprehensive stress condition, absolute angular velocities of the aircraft in three directions of an XYZ coordinate axis are as follows:
in the formula, phi ', theta ' and psi ' are respectively expressed as absolute angular velocities of the aircraft in three directions of XYZ coordinate axes; phi is an included angle between the central axis of the aircraft body and the Z axis of the longitudinal axis; theta is the included angle between the projection line of the central axis of the aircraft body on the XY plane and the X axis.
The utility model provides a pair of novel four rotor crafts of biax slope can utilize the rotation of support arm and motor fixing base to make the rotor vert, reach the purpose that the flight is freely met an emergency through verting the rotor then, can be according to different flight environmental conditions again, utilize closed-loop control system's stability, feed back the flight state of self, accomplish the work that oneself flies the regulation and control automatically, because the optimization of the organism structure of aircraft has reduced the application of some driving system equipment, reduce the size of a scale of aircraft to a certain extent, become more nimble flight. The flight that has realized four rotor crafts is intelligent, and the person of controlling only needs input command just can obtain its requirement, also avoids the danger that the error operation of the person of controlling brought, the utility model provides a novel four rotor crafts of biax slope advantage that stands out of convenience, safety and intelligence has extensive application prospect for military use and civilian use.
Drawings
The invention will be described in further detail with reference to the following drawings and detailed description:
FIG. 1 is a schematic diagram of the operation control system of the present invention;
figure 2 is a schematic top view of the aircraft of the present invention;
FIG. 3 is a schematic view of the tilting mechanism of the present invention;
FIG. 4 is a schematic block diagram of a vehicle dynamics system of the present invention;
fig. 5 is a schematic diagram of information exchange between the controller unit and the ground control station according to the present invention;
fig. 6 is a schematic diagram of the communication between the controller unit and each of the actuator units and the sensor units according to the present invention;
fig. 7 is a schematic diagram of a reference coordinate system of the aircraft in flight (one of the rotors is in a tilting state).
In the drawings: 1-a fuselage; 2-a sensor unit; 3-a controller unit; 4-controlling the dispenser unit; 5-a vehicle dynamics system module; 6-a wireless communication module; 7-a ground control station;
101-a landing gear; 102-an electrical equipment compartment; 103-a support arm; 104-a linkage mechanism; 105-a motor mount; 106-servo motor; 107-drive motor; 108-rotor; 109-a steering engine; 110-a bearing; 111-fixing blocks; 112-gravity sensor; 113-a gyroscope; 114-remote sensing; 115-a tilt mechanism; 116-a battery;
501-an actuator unit; 502-a sensor unit; 503-gravity sensing module; 504-perturbation sensing module;
601-a transmitting module; 602-a receiving module.
Detailed Description
As shown in fig. 1-7, a biaxial oblique four-rotor aircraft, the fuselage 1 of which comprises a plurality of power arms uniformly arranged at the periphery of the fuselage, and a sensor unit 2, a controller unit 3, a control distributor unit 4, a vehicle dynamics system module 5 and a wireless communication module 6 which are arranged at the fuselage; the wireless communication module can establish wireless communication with the ground control station 7;
the power arm comprises a rotor 108 and a tilt mechanism 115; the rotor is driven by a drive motor 107 at a motor mount 105 supported by the tilt mechanism; the tilting mechanism is connected with the controller unit; the controller unit adjusts rotor position and rotor orientation via a tilt mechanism to adjust the flight of the aircraft.
The number of the power arms is four.
The tilting mechanism is connected with a servo motor 106 at the machine body by the starting end of a supporting arm 103, and the tail end of the supporting arm is hinged with the middle part of a motor fixing seat; the servo motor drives the supporting arm to rotate so as to enable the rotor wing to tilt; a fixing block 111 for fixing the steering engine 109 is arranged at the supporting arm; the steering wheel links to each other through link mechanism 104 and the lower extreme of motor fixing base, the steering wheel drives the swing of motor fixing base through link mechanism makes the rotor produce yawing force.
The link mechanism is a parallelogram mechanism, the inclination angle α of the rotor wing can be adjusted between 0 degree and 90 degrees, and the inclination angle α can be 0 degree but can not be 90 degrees.
The fuselage also comprises a landing gear 101 and an electrical equipment bay 102; a gravity sensor 112, a gyroscope 113, a remote sensor 114 and a plurality of batteries 116 are arranged in the electric equipment cabin; the output shaft of the servo motor and the contact position of the rotor support arm with the electrical equipment bay 102 are connected by a bearing 110.
The sensor unit 2 is responsible for the connection work between the sensor unit and the controller unit 3, receives the measurement information of the sensor unit, converts the received measurement information into an electric signal according to a certain rule and outputs the electric signal to the controller unit 3; the controller unit 3 performs corresponding processing and sends out control command signals after finishing receiving the information transmitted by the sensor unit 2 and the wireless communication module 5;
the control distributor unit 4 completes the distribution work of the control command signals of the controller unit 3, and distributes the received command information of the controller unit 3 to four servo motors 106, four driving motors 107 and four steering engines 109 of the aircraft in order to realize the adjustment of the flight state;
the vehicle dynamics system module 5 comprises an actuator unit 501 and a sensor unit 502, the sensor unit 502 comprises a gravity sensing module 503, a perturbation sensing module 504, a gravity sensor 111, a gyroscope 113 and a remote sensor 114, and executes instructions for controlling the distributor unit 4, and the vehicle dynamics system module measures real-time gravity and perturbation external force through the sensor unit 502 and feeds back the measured force to the controller unit 3 to form a closed-loop system;
the wireless communication module 6 comprises a transmitting module 601 and a receiving module 602, the operating modes of the transmitting module 601 and the receiving module 602 can be switched simultaneously and do not interfere with each other, the transmitting module 602 is connected with the controller unit 3, and the receiving module 602 is in wireless communication with the ground control station 7 so that the ground control station 7 executes the azimuth instruction operation on the aircraft.
The sensor unit 2 is an IMU comprised of an SBG system and IG-500, which has an embedded processor for outputting filtered attitude and position data; the controller unit 3 is based on a PD controller with 3 SISDs, which processes the chip set maximum criteria of Digilent of the board for reading the signal; the controller unit 2 reads serial information at MAX32 from the output signals of the sensor unit 2, and instructs the servo motor 106, the driving motor 107 and the steering engine 109 by sending out PWM signals;
the receiving module 602 in the wireless communication module 6 communicates with the ground control station 7 through a Spektrum satellite receiver and a DX6i RC transmitter.
The absolute rotating speed of the output shaft of the driving motor is the vector sum of the angular speed of the aircraft in the three directions of XYZ, the rotating speed of the steering engine and the rotating speed of the servo motor, the physical characteristics of four supporting arms of the aircraft are consistent, one supporting arm is taken for stress analysis, and the absolute rotating speed of the output shaft of the driving motor is the absolute rotating speed of the output shaft of the steering engine and the absolute rotating speed of the servo motor
Angular acceleration of
Vector in the formulaiij ijj ikjThe unit vector representing a reference system j is represented by a reference system i, p, q and r respectively represent the angular velocity of the aircraft in three directions of a reference coordinate system XYZ, η and gamma represent the tilting angle of a double shaft during servo motion, η 'gamma' represents the tilting angular velocity of the double shaft during servo motion, and omega represents the rotating speed of a driving motor;
using the Euler equation3M＝3I3α+3ω×3I3ω (formula 3),
calculating the torque of the output shaft of the drive motor3M, then using the rotation matrix R3to1Calculation of the aircraft's own moments in a reference system1MGγroThe calculation formula is as follows:
the thrust torque of the connecting rod structure is as follows:
wherein L is the length of the parallelogram mechanism; h is the height of the parallelogram mechanism;
the thrust and torque coefficients of the rotor are defined as:
T＝ρA(ΩR)2CT(6) and Q ═ ρ a (Ω R)2RCQ(equation 7);
in the formula CTAnd CQThe thrust and the torque coefficient of the connecting rod structure.
The total thrust moment of the aircraft is then:
the absolute stress condition of the aircraft in six degrees of freedom is as follows:
x, Y and Z are respectively expressed as the absolute acting force of the aircraft on XYZ axes; l, M and N are respectively expressed as the absolute acting moments of the aircraft on XYZ axes; u, v and w are respectively expressed as the angular velocities of the whole aircraft on XYZ axes; when the aircraft is under the comprehensive stress condition, absolute angular velocities of the aircraft in three directions of an XYZ coordinate axis are as follows:
in the formula, phi ', theta ' and psi ' are respectively expressed as absolute angular velocities of the aircraft in three directions of XYZ coordinate axes; phi is an included angle between the central axis of the aircraft body and the Z axis of the longitudinal axis; theta is the included angle between the projection line of the central axis of the aircraft body on the XY plane and the X axis.
Example (b):
when the flight attitude of the aircraft is changed, the controller unit controls the tilting mechanism, the servo motor drives the supporting arm to rotate so that the rotor wing vertically tilts in the direction perpendicular to the supporting arm, and the steering engine drives the motor fixing seat to swing through the connecting rod mechanism, so that the rotor wing tilts inwards or outwards at the plane where the aircraft body is located, the lift output direction of the aircraft is changed, and the flight attitude of the aircraft is changed.
Claims (7)
1. The utility model provides a four rotor crafts of biax slope which characterized in that: the aircraft body (1) of the aircraft comprises a plurality of power arms uniformly arranged on the periphery of the aircraft body, and a sensor unit (2), a controller unit (3), a control distributor unit (4), a vehicle dynamics system module (5) and a wireless communication module (6) which are arranged on the aircraft body; the wireless communication module can establish wireless communication with a ground control station (7);
the power arm comprises a rotor (108) and a tilt mechanism (115); the rotor is driven by a drive motor (107) at a motor mount (105) supported by the tilt mechanism; the tilting mechanism is connected with the controller unit; the controller unit adjusts rotor position and rotor orientation via a tilt mechanism to adjust the flight of the aircraft.
2. The dual-axis tiltrotor aircraft according to claim 1, wherein: the number of the power arms is four.
3. The dual-axis tiltrotor aircraft according to claim 1, wherein: the tilting mechanism is connected with a servo motor (106) at the machine body by the starting end of a supporting arm (103), and the tail end of the supporting arm is hinged with the middle part of a motor fixing seat; the servo motor drives the supporting arm to rotate so as to enable the rotor wing to tilt; a fixing block (111) for fixing the steering engine (109) is arranged at the supporting arm; the steering wheel links to each other with the lower extreme of motor fixing base through link mechanism (104), the steering wheel drives the swing of motor fixing base through link mechanism makes the rotor produce yawing force.
4. The dual-axis tilting quadrotor aircraft according to claim 3, wherein the linkage is a parallelogram linkage, the tilt angle α of the rotor is adjustable between 0 ° and 90 °, and the tilt angle α is 0 ° but not 90 °.
5. A dual-axis tiltrotor aircraft according to claim 3, wherein: the fuselage further comprising a landing gear (101) and an electrical equipment bay (102); a gravity sensor (112), a gyroscope (113), a remote sensor (114) and a plurality of batteries (116) are arranged in the electrical equipment cabin; the output shaft of the servo motor and the contact position of the rotor wing supporting arm and the electrical equipment cabin (102) are connected through a bearing (110).
6. The dual-axis tiltrotor aircraft according to claim 5, wherein: the sensor unit (2) is responsible for the connection work between the sensor unit and the controller unit (3), receives the measurement information of the sensor unit and converts the received measurement information into an electric signal according to a certain rule and outputs the electric signal to the controller unit (3); the controller unit (3) performs corresponding processing and sends out control command signals after finishing receiving the transmission information of the sensor unit (2) and the wireless communication module (6);
the control distributor unit (4) completes the distribution work of the control command signals of the controller unit (3), and distributes the received command information of the controller unit (3) to four servo motors (106), four driving motors (107) and four steering engines (109) of the aircraft in order to realize the adjustment of the flight state;
the vehicle dynamics system module (5) comprises an actuator unit (501) and a sensor unit (502), the sensor unit (502) comprises a gravity sensing module (503), a perturbation sensing module (504), a gravity sensor (112), a gyroscope (113) and a remote sensor (114), and executes an instruction for controlling the distributor unit (4), and the vehicle dynamics system module measures real-time gravity and perturbation external force through the sensor unit (502) and feeds back the measured real-time gravity and perturbation external force to the controller unit (3) to form a closed-loop system;
the wireless communication module (6) comprises a transmitting module (601) and a receiving module (602), the working modes of the transmitting module (601) and the receiving module (602) can be switched simultaneously and do not interfere with each other, the transmitting module (601) is connected with the controller unit (3), and the receiving module (602) is in communication with the ground control station (7) in a wireless mode so that the ground control station (7) can execute azimuth instruction operation on the aircraft.
7. The dual-axis tiltrotor aircraft according to claim 6, wherein: the sensor unit (2) is an IMU comprised of an SBG system and IG-500, and has an embedded processor for outputting filtered attitude and position data; said controller unit (3) is based on a PD controller with 3 SISDs, which processes the chip set maximum criterion of Digilent of the board for reading the signal; the controller unit (3) reads serial information at MAX32 from output signals of the sensor unit (2), and commands the servo motor (106), the driving motor (107) and the steering engine (109) by sending out PWM signals;
and a receiving module (602) in the wireless communication module (6) is communicated with the ground control station (7) through a Spektrum satellite receiver and a DX6i RC transmitter.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201920653997.9U CN210027899U (en) | 2019-05-09 | 2019-05-09 | Four rotor crafts of biax slope |
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| Application Number | Priority Date | Filing Date | Title |
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| CN201920653997.9U CN210027899U (en) | 2019-05-09 | 2019-05-09 | Four rotor crafts of biax slope |
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| CN210027899U true CN210027899U (en) | 2020-02-07 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110015415A (en) * | 2019-05-09 | 2019-07-16 | 福州大学 | A kind of bi-axial tilt quadrotor |
-
2019
- 2019-05-09 CN CN201920653997.9U patent/CN210027899U/en not_active Expired - Fee Related
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110015415A (en) * | 2019-05-09 | 2019-07-16 | 福州大学 | A kind of bi-axial tilt quadrotor |
| CN110015415B (en) * | 2019-05-09 | 2024-02-09 | 福州大学 | Double-shaft tilting four-rotor aircraft |
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