A landing stage retractable four-rotor mars aircraft and rotor system

Abstract

The application relates to a four-rotor Mars flying vehicle with a retractable landing frame and a rotor system, and belongs to the technical field of Mars rotor flying vehicles. The application aims at the technical problems that a conventional four-rotor Mars flying vehicle cannot be retracted and automatically unfolded, and cannot independently perform an autonomous flight task on Mars. The application comprises a rotor module, a system module, a body module and a support arm module. The upper end of the body module is provided with the system module, the side surface of the body module is connected with one end of the support arm module, the other end of the support arm module is connected with the rotor module, the support arm module provides support for the body module, and the rotor module controls the flight state of the body module. The rotor blade, the rotor support arm and the landing frame can be retracted along joints to meet the size envelope requirement of a launch and on-orbit stage carrying spacecraft. After being deployed to the surface of Mars, the rotor support arm and the landing frame can be automatically unfolded, and the rotor blade can be unfolded by using centrifugal force, so that efficient and high-reliability deployment on Mars is realized.

Description

Technical Field

[0001] The present invention relates to a Martian aircraft and a system thereof, belonging to the technical field of Martian rotary-wing aircraft. Background Art

[0002] As Martian surface exploration missions progress, traditional exploration methods such as rovers and orbiters are gradually showing their limitations. Humanity needs more efficient and flexible exploration vehicles to carry out scientific missions. The thin atmosphere on the Martian surface makes it possible for rotorcraft to carry out flight exploration missions. Currently, the Ingenuity Mars helicopter has been successfully deployed on the Martian surface. The Ingenuity helicopter is a coaxial twin-rotor aircraft with limited payload capacity. Because there is no axial flow coupling interference between the rotors, quadrotors have higher payload capacity and hovering efficiency than coaxial rotor aircraft of the same diameter. However, the larger envelope of quadrotors makes them difficult to directly integrate into launch vehicles. Therefore, quadrotor Mars vehicles require the ability to retract and extend the rotor assembly and landing legs to meet space launch requirements. However, most current quadrotor Mars vehicles lack complete functional components and systems.

[0003] Therefore, there is an urgent need to propose a four-rotor Mars aircraft with a retractable landing pad and a rotor system to solve the above technical problems. Summary of the Invention

[0004] The present invention addresses the technical issues of conventional quadrotor Mars aircraft, which are unable to fold and automatically unfold, and are incapable of independently performing autonomous flight missions on Mars. The present invention provides a quadrotor Mars aircraft with a retractable landing pad and a rotor system. A brief overview of the present invention is provided below to provide a basic understanding of certain aspects of the present invention. It should be understood that this overview is not an exhaustive overview of the present invention. It is not intended to identify key or important aspects of the present invention, nor is it intended to limit the scope of the present invention.

[0005] The technical solution of the present invention:

[0006] A four-rotor Martian aircraft with a retractable landing frame includes a rotor module, a system module, a fuselage module and an arm module. The system module is installed on the upper end of the fuselage module, the side of the fuselage module is connected to one end of the arm module, and the other end of the arm module is connected to the rotor module. The arm module provides support for the fuselage module, and the rotor module controls the flight status of the fuselage module.

[0007] Preferably: the rotor module includes rotor blades, blade folding shafts, blade clamps, blade clamping nuts, gaskets, motor rotor housings, outer rotor propulsion motors, motor mounting interfaces and rotor module mounting interfaces, the stator of the outer rotor propulsion motor is connected to the rotor module mounting interface through the motor mounting interface, the rotor of the outer rotor propulsion motor is connected to the motor rotor housing, a blade folding shaft is provided on the motor rotor housing, a gasket, rotor blades, another gasket, a blade clamp, and a blade clamping nut are sequentially installed on the blade folding shaft, and the rotor module mounting interface is installed with the support arm module.

[0008] Preferably: the two blade folding shafts are embedded in the top of the motor rotor housing, the propeller clamp is mounted on the two blade folding shafts through the end through holes at both ends of the propeller clamp, the middle circular hole of the propeller clamp is connected to the shaft protruding from the middle of the motor rotor housing, and the two rotor blades are arranged in a circumferential array with the shaft protruding from the middle of the motor rotor housing as the center.

[0009] Preferably, the system module includes a communication antenna, a solar panel mounting interface and a solar panel, the solar panel is connected to the fuselage module via the solar panel mounting interface, and the communication antenna is mounted in the center of the solar panel.

[0010] Preferably: the fuselage module includes a fuselage upper frame, an arm mounting interface, a fuselage side panel, a fuselage column, a rear black and white camera mounting plate, a rear black and white camera, an electronic core module lower mounting frame, a fuselage lower frame, a front black and white camera, a front black and white camera mounting plate, an electronic core module upper mounting frame, a posture sensor module, a battery pack upper bracket, a lithium battery, a battery pack lower bracket, a circuit board connecting block, a laser altimeter, a color camera, a circuit board and a fuselage thermal insulation skin, a rear black and white camera mounting plate and a front black and white camera mounting plate are respectively arranged at the diagonal positions of the fuselage lower frame, the rear black and white camera is installed on the lower side of the rear black and white camera mounting plate, the front black and white camera is installed on the lower side of the front black and white camera mounting plate, and the lenses of the rear black and white camera and the front black and white camera are facing The lower arrangement is as follows: the upper side of the fuselage lower frame is connected to the electronic core module lower mounting frame, the upper side of the electronic core module lower mounting frame is installed with a battery pack lower bracket, the battery pack lower bracket is provided with a lithium battery, the lower side of the fuselage upper frame is connected to the electronic core module upper mounting frame, the lower side of the electronic core module upper mounting frame is connected to the battery pack upper bracket, the battery pack upper bracket is installed on the upper side of the lithium battery, four circuit boards are provided on the side of the electronic core module upper mounting frame, the bottom ends of adjacent circuit boards are connected by a circuit board connecting block, a laser altimeter is installed below the circuit board connecting block on the front side, the lens of the laser altimeter is downward, a color camera is installed on the front side of the laser altimeter, the color camera lens is tilted forward and downward, and the attitude sensor module is installed above the battery pack upper bracket;

[0011] The lower fuselage frame is connected to the upper fuselage frame through four evenly arranged fuselage columns. Fuselage side strips are set at the four corners of the lower fuselage frame and the upper fuselage frame. The upper fuselage frame, fuselage side strips, fuselage columns 3-4 and the lower fuselage frame are surface-mounted with fuselage insulation skin. Arm mounting interfaces are set at the midpoints of the four long sides of the upper fuselage frame, and the arm mounting interfaces are installed with the arm modules.

[0012] Preferably, the number of the arm modules is four, and the four arm modules are arranged in a circumferential array.

[0013] Preferably: the arm module includes an arm mounting interface, a retraction and expansion joint bracket, an arm locking spring, an arm locking pin, an arm rotating bearing, an arm rotating shaft, an arm expansion torsion spring and a rotor arm, one end of the rotor arm is connected to the arm mounting interface, the other end of the rotor arm is connected to the rotor module mounting interface, the arm mounting interface is connected to the retraction and expansion joint bracket, the middle part of the arm rotating shaft is provided with an arm expansion torsion spring and an arm mounting interface, the end of the arm rotating shaft is connected to the retraction and expansion joint bracket through the arm rotating bearing, the two ends of the arm expansion torsion spring are respectively connected to the arm mounting interface and the retraction and expansion joint bracket, a guide hole is processed on the arm mounting interface, an arm locking spring is installed in the guide hole, an arm locking pin is installed at the end of the arm locking spring, the retraction and expansion joint bracket is processed with a positioning hole, the arm locking pin cooperates with the retraction and expansion joint bracket and the positioning holes on the side of the retraction and expansion joint bracket.

[0014] Preferably: the support arm module also includes a landing gear deployment stop hook, a landing gear deployment torsion spring, a landing gear locking torsion spring, a landing gear locking wedge shaft, a landing gear locking wedge block, a landing gear retraction and extension shaft, a buffer flexible hinge, a landing gear vibration absorbing block, a landing leg rod and a foot pad, the landing gear locking wedge shaft and the landing gear retraction and extension shaft are all connected to the retraction and extension joint bracket, the landing gear locking wedge block is connected to the retraction and extension joint bracket through the landing gear locking wedge shaft, the landing gear locking wedge shaft is sleeved with a landing gear locking torsion spring, the landing gear deployment stop hook is connected to the retraction and extension joint bracket through the landing gear retraction and extension shaft, and the landing gear retraction and extension shaft is sleeved with a landing gear locking torsion spring. It is equipped with a landing gear deployment torsion spring, a landing gear locking torsion spring, and one end of the landing gear deployment torsion spring is connected to the retraction and deployment joint bracket, the other end of the landing gear locking torsion spring is connected to the landing gear locking wedge, the other end of the landing gear deployment torsion spring is connected to the landing gear deployment stop hook, the upper end of the landing gear deployment stop hook has a groove, the landing gear locking wedge cooperates with the groove of the landing gear deployment stop hook, the lower end of the landing gear deployment stop hook is connected to one end of the buffer flexible hinge, the side of the buffer flexible hinge is connected to the landing gear vibration absorbing block, the other end of the buffer flexible hinge is connected to the landing leg rod, and the end of the landing leg rod is bonded to the foot pad.

[0015] A retractable rotor system with a landing gear, wherein a power supply circuit board, an attitude sensor module, a laser altimeter, and a navigation circuit board are connected in parallel to a flight control circuit board. The navigation circuit board is electrically connected to a communication thermal control circuit board. A rear black-and-white camera, a front black-and-white camera, and a color camera are connected in parallel to the navigation circuit board.

[0016] Preferably: the solar panel and the lithium battery are electrically connected to the power circuit board, the power circuit board is electrically connected to the communication thermal control circuit board, and the communication antenna and the battery heating film are electrically connected to the communication thermal control circuit board.

[0017] The present invention has the following beneficial effects:

[0018] 1. The present invention carries a payload on the belly of the fuselage to perform scientific exploration and sample return missions on Mars.

[0019] 2. The present invention realizes autonomous flight control and independent exploration on Mars.

[0020] 3. The rotor blades, rotor arms and landing gear of the present invention can be folded along the joints to meet the size envelope requirements of the carrier spacecraft during launch and on-orbit phases.

[0021] 4. After being deployed on the surface of Mars, the present invention can automatically unfold the rotor arms and landing gear, and use centrifugal force to unfold the rotor blades, thereby achieving efficient and reliable deployment on Mars.

[0022] 5. The retraction and extension joints of the rotor arm and the landing gear in the present invention are integrated into the same modular structure, so that the four-rotor Mars aircraft has a compact structure and lighter weight. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] Figure 1 This is a three-dimensional diagram of a four-rotor Mars spacecraft with a retractable landing gear in its unfolded state;

[0024] Figure 2 It is a structural diagram of the rotor module;

[0025] Figure 3 This is a schematic diagram of the structure of a four-rotor Mars aircraft with a retractable landing frame in a folded state;

[0026] Figure 4 It is a three-dimensional image of the fuselage module;

[0027] Figure 5 It is a schematic diagram of the partial structure of the fuselage module;

[0028] Figure 6 It is an exploded view of the partial structure of the fuselage module;

[0029] Figure 7 is a cross-sectional view of the arm module;

[0030] Figure 8 is a partial structure diagram of the branch arm module;

[0031] Figure 9 is Figure 3 is an enlarged view of A in FIG. 4;

[0032] Figure 10 is a partial structure diagram of the landing leg retractable four-rotor Mars aircraft in the retracted state;

[0033] Figure 11 is a landing leg retractable rotor system diagram.

[0034] In the figure: 1-rotor module, 2-system module, 3-airframe module, 4-branch arm module, 1-1-rotor blade, 1-2-blade folding shaft, 1-3-blade clamp, 1-4-blade compression nut, 1-5-pad, 1-6-motor rotor housing, 1-7-outer rotor propulsion motor, 1-8-motor mounting interface, 1-9-rotor module mounting interface, 2-1-communication antenna, 2-2-solar panel mounting interface, 2-3-solar cell panel, 3-1-airframe upper frame, 3-2-branch arm mounting interface, 3-3-airframe side strip, 3-4-airframe stand, 3-5-rear-mounted black and white camera mounting plate, 3-6-rear-mounted black and white camera, 3-7-electronic core module lower mounting bracket, 3-8-airframe lower frame, 3-9-front-mounted black and white camera, 3-10-front-mounted black and white camera mounting plate, 3-11-electronic core module upper mounting bracket, 3-12-attitude sensor module, 3-13-battery pack upper support, 3-14-lithium battery, 3-15-battery pack lower support, 3-16-circuit board connecting block, 3-17-laser altimeter, 3-18-color camera, 3-19-circuit board, 3-20-airframe heat insulation skin, 4-1-branch arm mounting interface, 4-2-retractable joint support, 4-3-landing leg expansion stop hook, 4-4-landing leg expansion torsional spring, 4-5-landing leg locking torsional spring, 4-6-landing leg locking wedge shaft, 4-7-landing leg locking wedge block, 4-8-landing leg retraction shaft, 4-9-cushioning flexible hinge, 4-10-landing leg vibration absorbing block, 4-11-branch arm locking spring, 4-12-branch arm locking pin, 4-13-branch arm rotating bearing, 4-14-bearing pre-tightening block, 4-15-branch arm rotating shaft, 4-16-branch arm expansion torsional spring, 4-17-rotor branch arm, 4-18-landing leg rod, 4-19-foot pad. DETAILED DESCRIPTION

[0035] To make the objectives, technical solutions, and advantages of the present invention more clearly apparent, the present invention is described below using specific embodiments shown in the accompanying drawings. However, it should be understood that these descriptions are merely illustrative and are not intended to limit the scope of the present invention. In addition, in the following description, descriptions of well-known structures and technologies are omitted to avoid unnecessary confusion of the concepts of the present invention.

[0036] Specific implementation method 1: Combination Figure 1-10 The present embodiment is described. This embodiment is a four-rotor Mars aircraft with a retractable landing gear, comprising a rotor module 1, a system module 2, a fuselage module 3 and an arm module 4. The system module 2 is installed on the upper end of the fuselage module 3, the side of the fuselage module 3 is connected to one end of the arm module 4, and the other end of the arm module 4 is connected to the rotor module 1. The arm module 4 provides landing support for the fuselage module 3 and transmits the aerodynamic force generated by the rotor module 1 to the fuselage module 3. The four rotor modules 1 generate propulsion for the aircraft to take off and attitude control torque for controlling the flight state. The rotor module 1 and the arm module 4 of the present invention can be folded during the launch phase and automatically unfolded after being deployed on Mars.

[0037] Specific implementation method 2: Combination Figure 1-10 Describe this embodiment. This embodiment is a four-rotor Mars aircraft with a retractable landing gear. The rotor module 1 includes a rotor blade 1-1, a blade folding shaft 1-2, a blade clamp 1-3, a blade tightening nut 1-4, a pad 1-5, a motor rotor housing 1-6, an outer rotor propulsion motor 1-7, a motor mounting interface 1-8 and a rotor module mounting interface 1-9. The stator of the outer rotor propulsion motor 1-7 is fixedly connected to the rotor module mounting interface 1-9 through the motor mounting interface 1-8, and the rotor of the outer rotor propulsion motor 1-7 is fixedly connected to the motor rotor housing 1-6. The motor rotor housing 1-6 is provided with a blade folding shaft 1-2, and a pad is sequentially installed on the blade folding shaft 1-2. 1-5, rotor blade 1-1, another pad 1-5, blade clamp 1-3, blade clamping nut 1-4, rotate the blade clamping nut 1-4 to clamp the rotor blade 1-1, the motor mounting interface 1-8 and the rotor module mounting interface 1-9 both have cavities, and the rotor module mounting interface 1-9 is installed with the support arm module 4; the rotor blade 1-1 can rotate around the blade folding axis 1-2, but there is a certain rotation damping under the action of the blade clamping nut 1-4. In the retraction stage, the two rotor blades 1-1 are retracted along their respective blade folding axes 1-2 to be parallel to the rotor support arm 4-17. After the aircraft is deployed, the outer rotor propulsion motor 1-7 rotates slowly, relying on centrifugal force to throw the rotor blade 1-1 away. Figure 2 In the working position shown, in order to prevent Martian dust from contaminating the inside of the motor, the motor rotor housing 1-6 is a sealed structure integrally formed using carbon fiber composite materials.

[0038] Specific implementation method three: Combination Figure 1-10 The present embodiment is described as follows. This embodiment is a four-rotor Mars aircraft with a retractable landing gear. Two blade folding shafts 1-2 are symmetrically pre-embedded in the top of the motor rotor housing 1-6. The propeller clamps 1-3 are mounted on the two blade folding shafts 1-2 through the end through holes at both ends of the propeller clamps 1-3. The middle circular hole of the propeller clamps 1-3 is connected to the shaft protruding from the middle of the motor rotor housing 1-6. The two rotor blades 1-1 are arranged in a circumferential array with the shaft protruding from the middle of the motor rotor housing 1-6 as the center.

[0039] Specific implementation method four: Combination Figure 1-10 The present embodiment is described as follows. This embodiment is a four-rotor Mars aircraft with a retractable landing gear. The system module 2 includes a communication antenna 2-1, a solar panel mounting interface 2-2 and a solar panel 2-3. The solar panel 2-3 is connected to the fuselage module 3 via four solar panel mounting interfaces 2-2. The communication antenna 2-1 is mounted in the center of the solar panel 2-3 via four sets of bolts. The solar panel 2-3 is surface-mounted with thin-film solar cells to charge the power supply system of the present invention. The communication antenna 2-1 wirelessly communicates with the signal base station on the surface of Mars.

[0040] Specific implementation method five: Combination Figure 1-10This embodiment describes a quadrotor-type Mars aircraft with a retractable landing gear. The fuselage module 3 includes a fuselage upper frame 3-1, an arm mounting interface 3-2, a fuselage side slat 3-3, a fuselage column 3-4, a rear black and white camera mounting plate 3-5, a rear black and white camera 3-6, an electronic core module lower mounting frame 3-7, a fuselage lower frame 3-8, a front black and white camera 3-9, a front black and white camera mounting plate 3-10, an electronic core module upper mounting frame 3-11, an attitude sensor module 3-12, a battery pack upper bracket 3-13, a lithium battery 3-14, a battery pack lower bracket 3-15, and a circuit board connection block 3- 16, laser altimeter 3-17, color camera 3-18, circuit board 3-19 and fuselage thermal insulation skin 3-20, rear black and white camera mounting plate 3-5 and front black and white camera mounting plate 3-10 are respectively arranged at the diagonal positions of the fuselage lower frame 3-8, rear black and white camera 3-6 is installed on the lower side of the rear black and white camera mounting plate 3-5, and front black and white camera 3-9 is installed on the lower side of the front black and white camera mounting plate 3-10, and the lenses of the rear black and white camera 3-6 and the front black and white camera 3-9 are arranged downward, and the upper side of the fuselage lower frame 3-8 is connected to the electronic core module lower mounting frame 3-7 by bolts, and the upper side of the electronic core module lower mounting frame 3-7 is connected to the electronic core module lower mounting frame 3-7. The battery pack lower bracket 3-15 is installed on the side, and 15 lithium batteries are arranged on the battery pack lower bracket 3-15. A battery heating film 3-21 is arranged on the side of the lithium battery. The lower side of the upper frame 3-1 on the fuselage is bolted to the electronic core module upper mounting bracket 3-11. The lower side of the electronic core module upper mounting bracket 3-11 is bolted to the battery pack upper bracket 3-13. The battery pack upper bracket 3-13 is installed on the upper side of the 15 lithium batteries. The battery pack upper bracket 3-13 and the battery pack lower bracket 3-15 sandwich the 15 lithium batteries in the middle. The lithium battery and the battery pack upper bracket 3-13 and the battery pack lower bracket 3-15 are positioned by stoppers. The lithium battery is connected through wires and the copper sheet are connected in series and parallel, four circuit boards 3-19 are provided on the side of the side of the mounting frame 3-11 on the electronic core module, the bottom ends of adjacent circuit boards 3-19 are connected and fixed by circuit board connecting blocks 3-16, the front black and white camera 3-9 is located in front of the color camera 3-18, a laser altimeter 3-17 is installed below the circuit board connecting block 3-16 located in the front side, the lens of the laser altimeter 3-17 is facing downward, a color camera 3-18 is installed in front of the laser altimeter 3-17, the lens of the color camera 3-18 is tilted forward and downward, and the attitude sensor module 3-12 is installed in the upper center position of the battery pack upper bracket 3-13;

[0041] The lower fuselage frame 3-8 is connected to the upper fuselage frame 3-1 through four evenly arranged fuselage columns 3-4. The axis of the fuselage column 3-4 passes through the midpoint of the four long sides of the lower fuselage frame 3-8 and the upper fuselage frame 3-1. The lower fuselage frame 3-8 and the upper fuselage frame 3-1 are arranged in parallel. Fuselage side strips 3-3 are provided at the four corners of the lower fuselage frame 3-8 and the upper fuselage frame 3-1. Two fuselage side strips 3-3 are pasted at each corner. The upper fuselage frame 3-1, the fuselage side strips 3-3, the fuselage columns 3-4 and the lower fuselage frame 3-8 are surface-mounted with fuselage thermal insulation skin 3-20. The midpoints of the four long sides of the upper fuselage frame 3-1 are provided with support arm mounting interfaces 3-2. The support arm mounting interfaces 3-2 are installed with the support arm modules 4. The upper fuselage frame 3-1 and the lower fuselage frame 3-8 are made of carbon fiber composite materials, and the interior is prefabricated according to the mechanical installation requirements of other parts of the fuselage module. The threaded holes are buried. The middle of the upper frame 3-1 of the fuselage has a "cross" shaped carbon fiber tubular structure. The cross-section of the tube is rectangular, and the edge is an octagonal prism-shaped carbon fiber plate. The lower frame 3-8 of the fuselage is a "field" shaped carbon fiber tubular structure. The cross-section of the tube is square, and the four corners are connected by short sides. The fuselage column 3-4 is a carbon fiber tube with a square cross-section. The battery pack upper bracket 3-13, the battery pack lower bracket 3-15, the electronic core module upper mounting frame 3-11 and the electronic core module lower mounting frame 3-7 are all made of titanium alloy. The mass is concentrated. On the basis of ensuring the stability of the device and facilitating attitude adjustment, the lightweight of the device is achieved. The attitude sensor module 3-12 contains two inertial measurement units and a three-axis inclinometer. The four circuit boards are respectively a power board responsible for power management and power supply, a navigation board responsible for guidance and navigation calculations, a flight control board responsible for flight control, and a communication and thermal control board.

[0042] Specific implementation method six: combination Figure 1-10 The present embodiment will be described. This embodiment is a four-rotor Mars aircraft with a retractable landing frame. The number of the arm modules 4 is four, and the four arm modules 4 are arranged in a circumferential array.

[0043] Specific implementation method seven: combination Figure 1-10The present embodiment is described. The present embodiment is a four-rotor Mars aircraft with a retractable landing gear. The arm module 4 includes an arm mounting interface 4-1, a retractable joint bracket 4-2, an arm locking spring 4-11, an arm locking pin 4-12, an arm rotating bearing 4-13, a bearing preload block 4-14, an arm rotating shaft 4-15, an arm deployment torsion spring 4-16 and a rotor arm 4-17. The rotor arm 4-17 has a cavity therein. One end of the rotor arm 4-17 is fixedly connected to the arm mounting interface 4-1, and the other end of the rotor arm 4-17 is fixedly connected to the rotor module mounting interface 1-9. The rotor arm 4-17 and the rotor module mounting interface 1-9 both have a waist-shaped hole for passing the electronic cable of the rotor module 1 to the fuselage module 3. The arm mounting interface 3-2 is fixedly connected to the retractable joint bracket 4-2. The middle part of the arm shaft 4-15 is provided with an arm expansion torsion spring 4-16 and an arm mounting interface 4-1. The end of the arm shaft 4-15 is connected to the side of the expansion and retraction joint bracket 4-2 through the arm rotation bearing 4-13. The bearing preload block 4-14 presses against the outer ring of the arm rotation bearing 4-13 to realize the preload of the shaft system. The two ends of the arm expansion torsion spring 4-16 are respectively fixedly connected to the arm mounting interface 4-1 and the expansion and retraction joint bracket 4-2. A guide hole parallel to the arm shaft 4-15 is machined on the arm mounting interface 4-1, and an arm locking spring 4-11 is installed in the guide hole. An arm locking pin 4-12 is installed at the end of the arm locking spring 4-11. A positioning hole is machined on the side of the expansion and retraction joint bracket 4-2. The arm locking pin 4-12 cooperates with the side of the expansion and retraction joint bracket 4-2 and the positioning hole on the side of the expansion and retraction joint bracket 4-2.

[0044] The inner ring of the support arm rotating bearing 4-13 is connected to the support arm rotating shaft 4-15, and the support arm rotating shaft 4-15 passes through the through hole of the support arm mounting interface 4-1. The support arm expansion torsion spring 4-16 is sleeved on the support arm rotating shaft 4-15. One end of the support arm expansion torsion spring 4-16 hooks the expansion and contraction joint bracket 4-2, and the other end hooks the hole of the support arm mounting interface 4-1. There are guide holes on both sides of the support arm mounting interface 4-1, and the support arm locking spring 4-11 is installed in the guide hole. The outer side of the support arm locking spring 4-11 presses the support arm locking pin 4-12. There are circular holes on both sides of the expansion and contraction joint bracket 4-2. The positioning hole of the conical hole has the same shape as the tip of the support arm lock pin 4-12. The landing gear locking wedge shaft 4-6 and the landing gear retraction and extension shaft 4-8 are fixed on the retraction and extension joint bracket 4-2. The landing gear locking torsion spring 4-5 is sleeved on the landing gear locking wedge shaft 4-6. The landing gear deployment torsion spring 4-4 ​​is sleeved on the landing gear retraction and extension shaft 4-8. One end of the landing gear locking torsion spring 4-5 is constrained by the retraction and extension joint bracket 4-2, and the other end is inserted into the hole of the landing gear locking wedge block 4-7. One end of the landing gear deployment torsion spring 4-4 ​​is constrained by the retraction and extension joint bracket 4-2. , and the other end is inserted into the hole of the landing gear deployment stop lock hook 4-3; when the arm lock pin 4-12 is pressed back into the guide hole by the side of the retraction and deployment joint bracket 4-2, the arm locking spring 4-11 is compressed, and the arm mounting interface 4-1 connected to the rotor arm 4-17 can rotate around the arm rotation axis 4-15. During this process, the arm deployment torsion spring 4-16 is deformed. When the rotor arm 4-17 is deployed to the horizontal position under the action of the arm deployment torsion spring 4-16, the arm locking spring 4-11 pops the arm lock pin 4-12 into the positioning hole, locking the rotor arm 4-17. 7. The landing gear deployment stop lock hook 4-3 can rotate around the landing gear retraction and extension shaft 4-8. During this process, the landing gear deployment torsion spring 4-4 ​​is deformed and generates a restoring force for deploying the landing gear. When the constraint is lost, the landing leg buffer flexible hinge 4-9 is deployed under the action of the landing gear deployment torsion spring 4-4 ​​until the landing gear deployment stop lock hook 4-3 is restricted from moving by the landing gear locking wedge shaft 4-6. At this time, the landing gear locking wedge block 4-7 is wedged into the groove of the landing gear deployment stop lock hook 4-3 under the action of the landing gear locking torsion spring 4-5, completing the locking of the landing gear.

[0045] Specific implementation method eight: combination Figure 1-10The present embodiment is described. The present embodiment is a four-rotor Mars aircraft with a retractable landing gear. The arm module 4 also includes a landing gear deployment stop hook 4-3, a landing gear deployment torsion spring 4-4, a landing gear locking torsion spring 4-5, a landing gear locking wedge shaft 4-6, a landing gear locking wedge block 4-7, a landing gear retraction and extension shaft 4-8, a buffering flexible hinge 4-9, a landing gear vibration absorbing block 4-10, a landing leg rod 4-18 and a foot pad 4-19. The landing gear locking wedge shaft 4-6 and the landing gear retraction and extension shaft 4-8 are both connected to the retraction and extension joint bracket 4-2. The landing gear locking wedge block 4-7 is connected to the retraction and extension joint bracket 4-2 through the landing gear locking wedge shaft 4-6. The landing gear locking torsion spring 4-5 is installed on the landing gear locking wedge shaft 4-6. The landing gear deployment stop hook 4-3 is connected to the retraction and extension joint bracket 4-8 through the landing gear retraction and extension shaft 4-8. The landing gear stowage and extension shaft 4-8 is provided with a landing gear stowage and extension torsion spring 4-4, a landing gear locking torsion spring 4-5 and one end of the landing gear stowage and extension torsion spring 4-4 ​​are connected to the stowage and extension joint bracket 4-2, the other end of the landing gear locking torsion spring 4-5 is connected to the landing gear locking wedge 4-7, the other end of the landing gear stowage and extension torsion spring 4-4 ​​is connected to the landing gear stowage and extension stop hook 4-3, the upper end of the landing gear stowage and extension stop hook 4-3 has a groove, the landing gear locking wedge 4-7 is matched with the groove of the landing gear stowage and extension stop hook 4-3, the lower end of the landing gear stowage and extension stop hook 4-3 is connected to one end of the buffer flexible hinge 4-9, the side of the buffer flexible hinge 4-9 is connected to the landing gear vibration absorbing block 4-10, the other end of the buffer flexible hinge 4-9 is connected to the landing leg rod 4-18, and the end of the landing leg rod 4-18 is bonded to the foot pad 4-19.

[0046] Working principle of the present invention:

[0047] Mars spacecraft folding: The quadrotor Mars spacecraft is folded to fit inside the launch vehicle, first by folding the landing gear, then the rotor module, and then the rotor arms;

[0048] Landing gear retraction: The landing gear locking wedge 4-7 is pulled out of the slot of the landing gear deployment stop hook 4-3, and the landing gear deployment stop hook 4-3 rotates around the landing gear retraction and deployment rotation axis 4-8 until the landing leg rod 4-18 is parallel to the central axis of the fuselage module 3. During this process, the landing gear deployment torsion spring 4-4 ​​is deformed and stores energy, and the landing gear locking wedge 4-7 maintains contact with the surface of the landing gear deployment stop hook 4-3 under the action of the landing gear locking torsion spring 4-5;

[0049] Folding of the rotor module 1: The two rotor blades 1-1 rotate along the blade folding axis 1-2 until they are parallel to the rotor support arm 4-17.

[0050] Retraction of the rotor arm 4-17: the arm locking pin 4-12 is pressed out from the hole of the retraction and extension joint bracket 4-2, and its top end contacts the outer surface of the retraction and extension joint bracket 4-2. The arm locking spring 4-11 is in a compressed state, and the arm mounting interface 4-1 connected to the rotor arm 4-17 rotates around the arm rotation axis 4-15. During this process, the arm deployment torsion spring 4-16 is deformed, and the rotor arm 4-17 is folded to be parallel to the central axis of the fuselage module 3.

[0051] Mars spacecraft deployment: After arriving on Mars, the quadcopter Mars spacecraft is freed from its surrounding constraints and deployed, first by deploying the rotor arms, then the landing legs, and finally the rotor modules.

[0052] Deployment and locking of the rotor arm 4-17: After the constraint is released, the arm mounting interface 4-1 rotates upward under the action of the arm deployment torsion spring 4-16. When the rotor arm 4-17 rotates to the horizontal position, the arm locking pin 4-12 aligns with the hole of the stowage joint bracket 4-2 and is inserted into the hole of the stowage joint bracket 4-2 under the push of the arm locking spring 4-11, thereby achieving the deployment and locking of the rotor arm 4-17.

[0053] Landing gear deployment and locking: After the restraint is released, the landing gear deployment stop hook 4-3 rotates around the landing gear retraction and deployment rotation axis 4-8 under the action of the landing gear deployment torsion spring 4-4 ​​until the landing gear deployment stop hook 4-3 is limited by the landing gear locking wedge rotation axis 4-6. The landing gear locking wedge 4-7 is wedged into the groove of the landing gear deployment stop lock hook 4-3 under the action of the landing gear locking torsion spring 4-5, thereby realizing the deployment and locking of the landing gear;

[0054] Deployment of the rotor module: The outer rotor propulsion motor 1-7 drives the rotor blades 1-1 to rotate, and the two rotor blades 1-1 are deployed around the blade folding axis 1-2 to the working position under the action of centrifugal force.

[0055] Specific implementation method nine: Combination Figure 1-11 The present embodiment is described. A rotor system with a retractable landing pad is provided in the present embodiment. A four-rotor Mars aircraft with a retractable landing pad is used. The power supply circuit board, attitude sensor module 3-12, laser altimeter 3-17, and navigation circuit board are connected in parallel to the flight control circuit board. The navigation circuit board is electrically connected to the communication thermal control circuit board. The rear black and white camera 3-6, the front black and white camera 3-9, and the color camera 3-18 are connected in parallel to the navigation circuit board.

[0056] Specific implementation method ten: Combination Figure 1-11In this embodiment, the rotors of the landing frame can be folded and unfolded, the solar panels 2-3 and lithium batteries 3-14 are electrically connected to the power circuit board, the power circuit board is electrically connected to the communication and thermal control circuit board, the communication antenna 2-1 and the battery heating film 3-21 are electrically connected to the communication and thermal control circuit board; the circuit board 3-19 includes four circuit boards, namely the power circuit board, the flight control circuit board, the navigation circuit board and the communication and thermal control circuit board, the power circuit board receives the power charged by the solar panels 2-3 and stores the power in the lithium batteries 3-14, manages the discharge program of the lithium batteries 3-14, supplies power to the flight control circuit board and the communication and thermal control circuit board, the flight control board supplies power to the navigation circuit board, the flight control board receives attitude measurement data from the attitude sensing module 3-12 and aircraft height data from the laser altimeter 3-17, the navigation circuit board receives image data from the front navigation camera 3-9, the rear navigation camera 3-6 and the color camera 3-18, performs fusion calculation by using an algorithm to obtain the position and attitude data of the aircraft, the flight control circuit board and the navigation circuit board exchange data, the navigation circuit board sends the fused position and attitude data to the flight control board, the flight control board sends thermal control instructions to the navigation circuit board, the navigation circuit board forwards the thermal control data to the communication and thermal control circuit board to control the battery heating film 3-21, the communication and thermal control board exchanges data with the communication antenna 2-1 to receive wireless instructions from the signal base station and send the detection data of the aircraft to the signal base station; the rotors generate propulsion force to overcome gravity and control torque to adjust the attitude, the blades of adjacent rotor modules 1 rotate in opposite directions, while the blades of rotor modules 1 at diagonal positions rotate in the same direction, the rotor arm 4-17 connects the rotor module 1 to the body module 3, the landing leg 4-18 supports the Mars aircraft, the foot pad 4-19 at the end has strong wear resistance and directly contacts the surface of Mars to increase the friction between the landing frame and the surface of Mars, the covered body thermal insulation skin 3-20 has a polyimide coating to reduce heat leakage of electronic components inside the body, the 15 lithium batteries 3-14 store power for the quadcopter Mars aircraft to perform detection tasks independently, the four circuit boards 3-19 realize power management, guidance, navigation, thermal control, communication, flight control and other functions of the Mars aircraft, the front and rear black and white cameras 3-9 and 3-6 perform binocular vision navigation function, the attitude sensor module 3-12 can measure the inclination, angular velocity and acceleration of the body around the three axes in space, the laser altimeter 3-17 can measure the height of the aircraft from the surface of Mars, the color camera 3-18 can take color pictures of the surface of Mars, the guidance, navigation and control modules of the Mars aircraft calculate the state of the aircraft using real-time data obtained by the cameras and sensors and realize autonomous flight control.

[0057] It should be noted that in the above embodiments, as long as the technical solutions are not contradictory, they can be permuted and combined. Those skilled in the art can exhaust all possibilities based on the mathematical knowledge of permutations and combinations. Therefore, the present invention will no longer describe the technical solutions after permutations and combinations one by one, but it should be understood that the technical solutions after permutations and combinations have been disclosed by the present invention.

[0058] The foregoing description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Those skilled in the art will readily appreciate that various modifications and variations of the present invention are possible. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention are intended to be within the scope of protection of the present invention.

Claims

1. A four-rotor Mars aircraft with a retractable landing gear, characterized by: The invention comprises a rotor module (1), a system module (2), a fuselage module (3) and an arm module (4), wherein the upper end of the fuselage module (3) is installed with the system module (2), the side of the fuselage module (3) is connected to one end of the arm module (4), and the other end of the arm module (4) is connected to the rotor module (1), the arm module (4) provides support for the fuselage module (3), and the rotor module (1) controls the flight state of the fuselage module (3); The number of the support arm modules (4) is four, and the four support arm modules (4) are arranged in a circumferential array; The arm module (4) comprises a second arm mounting interface (4-1), a retractable joint bracket (4-2), an arm locking spring (4-11), an arm locking pin (4-12), an arm rotating bearing (4-13), an arm rotating shaft (4-15), an arm deployment torsion spring (4-16) and a rotor arm (4-17), one end of the rotor arm (4-17) is connected to the second arm mounting interface (4-1), the other end of the rotor arm (4-17) is connected to the rotor module mounting interface (1-9), the first arm mounting interface (3-2) is connected to the retractable joint bracket (4-2), the middle part of the arm rotating shaft (4-15) is equipped with an arm deployment torsion spring (4-16), the second arm mounting interface (4-13), and the rotor module mounting interface (1-9). The mounting interface (4-1) is connected to the expansion joint bracket (4-2) at the end of the support arm rotating shaft (4-15) through the support arm rotating bearing (4-13), and the two ends of the support arm expansion torsion spring (4-16) are respectively connected to the second support arm mounting interface (4-1) and the expansion joint bracket (4-2). A guide hole is machined on the second support arm mounting interface (4-1), and a support arm locking spring (4-11) is installed in the guide hole. An arm locking pin (4-12) is installed at the end of the support arm locking spring (4-11). The expansion joint bracket (4-2) is machined with a positioning hole. The support arm locking pin (4-12) is matched with the expansion joint bracket (4-2) and the positioning hole on the side of the expansion joint bracket (4-2). The support arm module (4) further comprises a landing gear deployment stop hook (4-3), a landing gear deployment torsion spring (4-4), a landing gear locking torsion spring (4-5), a landing gear locking wedge rotating shaft (4-6), a landing gear locking wedge block (4-7), a landing gear retraction and deployment rotating shaft (4-8), a buffering flexible hinge (4-9), a landing gear vibration absorbing block (4-10), a landing leg rod (4-18) and a foot pad (4-19), a landing gear locking wedge rotating shaft (4-6) The landing gear retraction and extension shaft (4-8) is connected to the retraction and extension joint bracket (4-2), the landing gear locking wedge (4-7) is connected to the retraction and extension joint bracket (4-2) through the landing gear locking wedge shaft (4-6), the landing gear locking wedge shaft (4-6) is equipped with a landing gear locking torsion spring (4-5), the landing gear deployment stop hook (4-3) is connected to the retraction and extension joint bracket (4-2) through the landing gear retraction and extension shaft (4-8), and the landing gear retraction and extension shaft The upper part (4-8) is provided with a landing gear deployment torsion spring (4-4), one end of the landing gear locking torsion spring (4-5) and the landing gear deployment torsion spring (4-4) is connected to the retraction and expansion joint bracket (4-2), the other end of the landing gear locking torsion spring (4-5) is connected to the landing gear locking wedge (4-7), and the other end of the landing gear deployment torsion spring (4-4) is connected to the landing gear deployment stop hook (4-3), and the upper end of the landing gear deployment stop hook (4-3) has a groove. The landing gear locking wedge (4-7) is matched with the groove of the landing gear deployment stop hook (4-3), the lower end of the landing gear deployment stop hook (4-3) is connected to one end of the buffer flexible hinge (4-9), the side of the buffer flexible hinge (4-9) is connected to the landing gear vibration absorbing block (4-10), the other end of the buffer flexible hinge (4-9) is connected to the landing leg rod (4-18), and the end of the landing leg rod (4-18) is bonded to the foot pad (4-19).

2. The quadrotor-type Martian aircraft with a retractable landing platform according to claim 1, characterized in that: The rotor module (1) comprises a rotor blade (1-1), a blade folding shaft (1-2), a blade clamp (1-3), a blade pressing nut (1-4), a pad (1-5), a motor rotor housing (1-6), an outer rotor propulsion motor (1-7), a motor mounting interface (1-8) and a rotor module mounting interface (1-9). The stator of the outer rotor propulsion motor (1-7) is connected to the rotor module mounting interface (1-9) via the motor mounting interface (1-8), the rotor of the outer rotor propulsion motor (1-7) is connected to the motor rotor housing (1-6), a blade folding shaft (1-2) is provided on the motor rotor housing (1-6), a pad (1-5), a rotor blade (1-1), another pad (1-5), a blade clamp (1-3) and a blade pressing nut (1-4) are sequentially mounted on the blade folding shaft (1-2), and the rotor module mounting interface (1-9) is mounted on the support arm module (4).

3. The quadrotor-type Martian aircraft with a retractable landing platform according to claim 2, characterized in that: Two blade folding shafts (1-2) are pre-buried in the top of the motor rotor housing (1-6); a blade clamp (1-3) is sleeved on the two blade folding shafts (1-2) through end through holes at both ends of the blade clamp (1-3); a central circular hole of the blade clamp (1-3) is connected to a shaft protruding from the middle of the motor rotor housing (1-6); and the two rotor blades (1-1) are arranged in a circumferential array with the shaft protruding from the middle of the motor rotor housing (1-6) as the center.

4. The quadrotor-type Martian aircraft with a retractable landing platform according to claim 2, characterized in that: The system module (2) includes a communication antenna (2-1), a solar panel mounting interface (2-2) and a solar panel (2-3). The solar panel (2-3) is connected to the fuselage module (3) through the solar panel mounting interface (2-2). The communication antenna (2-1) is installed in the center of the solar panel (2-3).

5. The quadrotor-type Martian aircraft with a retractable landing platform according to claim 4, characterized in that: The fuselage module (3) includes a fuselage upper frame (3-1), a first arm mounting interface (3-2), a fuselage side panel (3-3), a fuselage column (3-4), a rear black and white camera mounting plate (3-5), a rear black and white camera (3-6), an electronic core module lower mounting frame (3-7), a fuselage lower frame (3-8), a front black and white camera (3-9), a front black and white camera mounting plate (3-10), an electronic core module upper mounting frame (3-11), a posture sensor module (3-12), a battery pack upper bracket (3-13), a lithium battery (3-14), a battery The lower bracket (3-15), the circuit board connecting block (3-16), the laser altimeter (3-17), the color camera (3-18), the circuit board (3-19) and the fuselage heat insulation skin (3-20) are respectively arranged at the diagonal positions of the fuselage lower frame (3-8), the rear black and white camera mounting plate (3-5) and the front black and white camera mounting plate (3-10) are respectively arranged, the rear black and white camera (3-6) is installed on the lower side of the rear black and white camera mounting plate (3-5), the front black and white camera (3-9) is installed on the lower side of the front black and white camera mounting plate (3-10), the rear black and white camera (3-6) and The lens of the front black and white camera (3-9) is arranged downward, the upper side of the fuselage lower frame (3-8) is connected to the electronic core module lower mounting frame (3-7), the upper side of the electronic core module lower mounting frame (3-7) is installed with a battery pack lower bracket (3-15), and a lithium battery is arranged on the battery pack lower bracket (3-15), the lower side of the fuselage upper frame (3-1) is connected to the electronic core module upper mounting frame (3-11), the lower side of the electronic core module upper mounting frame (3-11) is connected to the battery pack upper bracket (3-13), and the battery pack upper bracket (3-13) is installed on the upper side of the lithium battery, the electronic core module Four circuit boards (3-19) are arranged on the side of the core module upper mounting frame (3-11), the bottom ends of adjacent circuit boards (3-19) are connected by a circuit board connecting block (3-16), a laser altimeter (3-17) is installed below the circuit board connecting block (3-16) on the front side, the lens of the laser altimeter (3-17) is downward, a color camera (3-18) is installed on the front side of the laser altimeter (3-17), the lens of the color camera (3-18) is tilted forward and downward, and the attitude sensor module (3-12) is installed above the battery pack upper bracket (3-13); The lower fuselage frame (3-8) is connected to the upper fuselage frame (3-1) through four evenly arranged fuselage columns (3-4); fuselage side panels (3-3) are provided at the four corners of the lower fuselage frame (3-8) and the upper fuselage frame (3-1); the upper fuselage frame (3-1), the fuselage side panels (3-3), the fuselage columns 3-4 and the lower fuselage frame (3-8) are surface-mounted with a fuselage heat insulation skin (3-20); a first arm mounting interface (3-2) is provided at the midpoint of the four long sides of the upper fuselage frame (3-1); and the first arm mounting interface (3-2) is installed with the arm module (4).

6. A retractable landing gear rotor system, characterized in that: A four-rotor Mars aircraft with a retractable landing frame according to any one of claims 1 to 5 is adopted, wherein a power supply circuit board, an attitude sensor module (3-12), a laser altimeter (3-17), and a navigation circuit board are connected in parallel to a flight control circuit board, the navigation circuit board is electrically connected to a communication thermal control circuit board, and a rear black and white camera (3-6), a front black and white camera (3-9), and a color camera (3-18) are connected in parallel to the navigation circuit board.

7. The retractable landing gear rotor system according to claim 6, characterized in that: The solar cell panel (2-3) and the lithium battery (3-14) are electrically connected to the power circuit board, the power circuit board is electrically connected to the communication thermal control circuit board, and the communication antenna (2-1) and the battery heating film (3-21) are electrically connected to the communication thermal control circuit board.

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