Landing gear and aircraft

Abstract

This invention discloses a landing gear and an aircraft, relating to the field of aircraft structural components. The landing gear includes a nose landing gear and two main landing gears. The nose landing gear is located at the front of the fuselage and includes a nose wheel and a nose shock strut. One end of the nose shock strut is connected to the fuselage and located inside the fuselage, while the other end extends outside the fuselage and is elastically connected to the nose wheel. The two main landing gears are respectively located on opposite sides of the fuselage. Each main landing gear includes a rocker arm, a main shock absorber, and a main wheel. The rocker arm has a hinged end and a free end, respectively. The hinged end is hinged to the inside of the fuselage so that the rocker arm can only swing along the aircraft's heading. The free end is connected to the main wheel. This invention incorporates part of the nose landing gear structure inside the fuselage and directly hinges the rocker arms of the main landing gear to the fuselage, reducing the overall frontal area of ​​the landing gear, decreasing the aerodynamic drag and overall weight of the aircraft, thereby improving the overall performance and range of the aircraft.

Description

Technical Field

[0001] This invention relates to the field of aircraft structural components, and particularly to a landing gear and an aircraft. Background Technology

[0002] With the rapid development of the low-altitude economy, UAM (Urban Air Mobility) and eVTOL (Electric Vertical Takeoff and Landing) aircraft have become the two most favored concepts in this field. The landing gear system is a very important component of an aircraft, serving as the key carrier for the aircraft to achieve functions such as ground support, takeoff, landing, and taxiing.

[0003] In the existing technology, the wheeled landing gear used in eVTOL is usually non-retractable and accounts for 4% to 6% of the total weight of the aircraft. However, for eVTOL, 4% to 6% of the total weight is too large, and the frontal area of ​​the non-retractable landing gear is too large, resulting in excessive aerodynamic drag. Both factors will affect the overall performance and range of the aircraft. Summary of the Invention

[0004] The main objective of this invention is to propose a landing gear and an aircraft that addresses the problem of excessive landing gear weight and aerodynamic drag in the prior art, which affects the overall performance and range of the aircraft.

[0005] To achieve the above objectives, the present invention proposes a landing gear, wherein the landing gear is disposed at the bottom of the aircraft fuselage, and the landing gear comprises:

[0006] The front landing gear is located on the front side of the bottom of the fuselage. The front landing gear includes a front wheel and a front shock strut. One end of the front shock strut is connected to the fuselage and located inside the fuselage, and the other end extends out of the fuselage and is elastically connected to the front wheel.

[0007] Two main landing gears are respectively disposed on opposite sides of the bottom of the fuselage. Each main landing gear includes a rocker arm, a main shock absorber, and a main landing gear wheel. The two ends of the rocker arm are an articulated end and a free end, respectively. The articulated end is hinged to the inside of the fuselage so that the rocker arm can only swing along the heading of the aircraft. The free end is connected to the main landing gear wheel. The position on the rocker arm between the articulated end and the free end is elastically connected to the fuselage through the main shock absorber.

[0008] In one embodiment, the main buffer includes a main sleeve and a first buffer rod. The main sleeve is hinged to the inside of the machine body. A first end of the first buffer rod is sleeved inside the main sleeve. A second end of the first buffer rod extends out of the main sleeve and is hinged to the rocker arm near the hinged end. The first buffer rod can extend and retract along the extension direction of the main sleeve.

[0009] In one embodiment, the two main landing gears are symmetrically arranged about the plane of symmetry of the fuselage, and each rocker arm is inclined from the hinge end to the free end toward the tail side of the fuselage, and from the hinge end to the free end toward the direction away from the other rocker arm. The hinge end is hinged to the inside of the fuselage by a hinge rod, and the hinge rod is perpendicular to the plane of symmetry of the fuselage.

[0010] In one embodiment, the rocker arm includes an elastic damping rod and a rocker arm joint connected to each other. The end of the rocker arm joint away from the elastic damping rod forms the hinge end. The end of the rocker arm joint near the elastic damping rod is provided with a hinge seat that is hinged to the second end of the first buffer rod. The hinge end is provided with two spaced-apart hinge rings. The two ends of the hinge rod are rotatably passed through the two hinge rings respectively. The end of the elastic damping rod away from the rocker arm joint forms the free end.

[0011] In one embodiment, the rocker arm has an elliptical cross-section, and one long axis side of the rocker arm's cross-section faces the nose side of the fuselage.

[0012] In one embodiment, each of the main drive wheels is inclined from top to bottom toward the other main drive wheel, and the angle between the axis of each main drive wheel and the horizontal plane is 1° to 1.5°; or, the front buffer support is inclined from top to bottom toward the nose side of the body, and the angle between the axis of the front buffer support and the vertical line is 7° to 7.5°.

[0013] In one embodiment, the front wheel and / or each of the main wheels are provided with a fairing.

[0014] In one embodiment, wheel-mounted switches are provided on the nose landing gear and / or the main landing gear, the wheel-mounted switches being configured to transmit air-to-ground signals.

[0015] The present invention also proposes a landing gear design method, applied to the aforementioned landing gear, the landing gear design method comprising the following steps:

[0016] The main landing gear is configured to absorb energy based on the landing kinetic energy and potential energy of the aircraft, wherein the energy absorption of the main landing gear includes energy absorption by the main landing gear wheels, energy absorption by the rocker arms, and energy absorption by the main shock absorbers.

[0017] The theoretical overload value of the aircraft is determined based on the vertical landing energy and / or the taxiing landing energy of the aircraft.

[0018] The buffer stroke of the main buffer is calculated based on the theoretical overload value and the law of conservation of energy.

[0019] Based on the buffer stroke of the main buffer, drop simulation is performed to obtain at least the overload simulation value and stroke simulation value of the main landing gear;

[0020] At least the theoretical overload value is adjusted, and the simulated overload value and simulated travel value are repeatedly calculated iteratively until the ratio of the energy absorbed by the main landing gear to the sum of the landing kinetic energy and potential energy of the aircraft is greater than or equal to the ratio threshold, and / or the difference between the sum of the landing kinetic energy and potential energy of the aircraft and the energy absorbed by the main landing gear is less than or equal to the difference threshold.

[0021] The present invention also proposes an aircraft having a fuselage, the bottom of which is provided with the aforementioned landing gear.

[0022] The technical solution of this invention involves placing a portion of the nose landing gear's front buffer strut inside the fuselage, with only one end of the strut extending outside the fuselage and elastically connected to the nose wheel. This allows most of the nose landing gear structure to be housed within the fuselage, reducing its frontal area and aerodynamic drag. Simultaneously, the main landing gear is directly hinged to the outside of the fuselage via a rocker arm, eliminating the need for an intermediate connecting structure. This reduces the number of structural components in the main landing gear, further lowering its frontal area and overall weight. By incorporating a portion of the nose landing gear structure within the fuselage and directly hinged the main landing gear's rocker arm to the fuselage, this invention reduces the overall frontal area of ​​the landing gear, decreases aerodynamic drag and overall weight, thereby improving the aircraft's overall performance and range. Attached Figure Description

[0023] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the structures shown in these drawings without creative effort.

[0024] Figure 1 This is a schematic diagram illustrating the assembly relationship between the landing gear and the fuselage provided by the present invention.

[0025] Figure 2 A schematic diagram of the landing gear structure provided by the present invention;

[0026] Figure 3A schematic diagram of the front landing gear provided by the present invention;

[0027] Figure 4 This is a schematic diagram of the main landing gear provided by the present invention;

[0028] Figure 5 A schematic diagram of the angle of the main landing gear provided by the present invention;

[0029] Figure 6 An assembly structure diagram of the landing gear and fairing provided for this invention;

[0030] Figure 7 A flowchart illustrating the design method for the main landing gear provided by this invention.

[0031] Explanation of icon numbers:

[0032] 100. Landing Gear; 1. Nose Landing Gear; 11. Nose Wheel; 12. Nose Buffer Strut; 121. Nose Sleeve; 122. Second Buffer Bar; 13. Anti-Sway Assembly; 131. Anti-Sway Body; 132. Torque Arm; 133. Drive Ring; 134. First Lever Arm; 135. Second Lever Arm; 14. Nose Wheel Onboard Switch; 15. Wheel Fork; 16. Nose Support Rod; 17. Baffle; 2. Main Landing Gear; 21. Rocker Arm; 211. Articulated End; 212. Free End; 213. Elastic Shock Absorber Bar; 214. Rocker Arm Joint; 215. Articulated Seat; 216. Articulated Ring; 22. Main Buffer; 221. Main Sleeve; 222. First Buffer Bar; 23. Main Wheel; 24. Articulated Rod; 25. Main Wheel Onboard Switch; 3. Fuselage; 4. Fairing.

[0033] The realization of the objective, functional features and advantages of the present invention will be further explained in conjunction with the embodiments and with reference to the accompanying drawings. Detailed Implementation

[0034] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present invention.

[0035] It should be noted that if the embodiments of the present invention involve directional indicators (such as up, down, left, right, front, back, etc.), the directional indicators are only used to explain the relative positional relationship and movement of the components in a specific posture. If the specific posture changes, the directional indicators will also change accordingly.

[0036] Furthermore, if the embodiments of this invention involve descriptions such as "first" or "second," these descriptions are for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined with "first" or "second" may explicitly or implicitly include at least one of those features. Additionally, the use of "and / or" or "and / or" throughout the text includes three parallel solutions. For example, "A and / or B" includes solution A, solution B, or a solution where both A and B are satisfied simultaneously. Furthermore, the technical solutions of the various embodiments can be combined with each other, but this must be based on the ability of those skilled in the art to implement them. When the combination of technical solutions is contradictory or impossible to implement, it should be considered that such a combination of technical solutions does not exist and is not within the scope of protection claimed by this invention.

[0037] In the existing technology, the wheeled landing gear used in eVTOL is usually non-retractable and accounts for 4% to 6% of the total weight of the aircraft. However, for eVTOL, 4% to 6% of the total weight is too large, and the frontal area of ​​the non-retractable landing gear is too large, resulting in excessive aerodynamic drag. Both factors will affect the overall performance and range of the aircraft.

[0038] To address the above problems, this invention proposes a landing gear 100.

[0039] Please combine Figures 1 to 4 In this embodiment, the landing gear 100 is disposed at the bottom of the fuselage 3 of the aircraft. The landing gear 100 includes a nose landing gear 1 and two main landing gears 2. The nose landing gear 1 is disposed at the front side of the bottom of the fuselage 3. The nose landing gear 1 includes a nose wheel 11 and a nose shock strut 12. One end of the nose shock strut 12 is connected to the fuselage 3 and disposed inside the fuselage 3, and the other end extends out of the fuselage 3 and is elastically connected to the nose wheel 11. The two main landing gears 2 are respectively disposed on opposite sides of the bottom of the fuselage 3. Each main landing gear 2 includes a rocker arm 21, a main shock absorber 22 and a main wheel 23. The two ends of the rocker arm 21 are a hinged end 211 and a free end 212, respectively. The hinged end 211 is hinged to the inside of the fuselage 3 so that the rocker arm 21 can only swing along the heading of the aircraft. The free end 212 is connected to the main wheel 23. The position on the rocker arm 21 between the hinged end 211 and the free end 212 is elastically connected to the fuselage 3 through the main shock absorber 22.

[0040] Understandably, the aircraft travels in a forward-backward direction, with the two main landing gears 2 positioned on the left and right sides of the fuselage 3, and typically arranged symmetrically. During landing, the nose wheel 11 contacts the ground, and its deformation absorbs some of the impact force. The front buffer strut 12, elastically connected to the nose wheel 11, compresses elastically under the impact force generated by the weight of the fuselage, thus absorbing another portion of the impact force. When the main landing gear 23 contacts the ground, its deformation absorbs some of the impact force. The rocker arm 21 rotates relative to the fuselage 3 under the impact force, absorbing another portion of the impact force. The main buffer 22, elastically connected to the rocker arm 21, compresses elastically under the impact force, absorbing yet another portion of the impact force, allowing the aircraft to land smoothly.

[0041] The technical solution of the present invention places a portion of the front buffer strut 12 of the front landing gear 1 inside the fuselage 3, with only one end of the front buffer strut 12 extending outside the fuselage 3 and elastically connected to the front wheel 11. This allows most of the structure of the front landing gear 1 to be housed inside the fuselage 3, thereby reducing the frontal area of ​​the front landing gear 1 and reducing aerodynamic drag. At the same time, the main landing gear 2 is directly hinged to the inside of the fuselage 3 via the rocker arm 21, without the need for an intermediate transition structure to indirectly connect it to the fuselage 3. This reduces the number of structural components of the main landing gear 2, thereby reducing the frontal area of ​​the main landing gear 2 and the overall weight of the main landing gear 2. The main landing gear 2 adopts a rocker arm type landing gear, which allows the rocker arm 21 to be set in any position to meet the parameter requirements of the aircraft, improving the flexibility of the layout and reducing the overall volume of the landing gear 100. This invention reduces the overall frontal area of ​​the landing gear 100 by incorporating part of the front landing gear 1 within the fuselage 3 and directly hinges the rocker arm 21 of the main landing gear 2 to the fuselage 3. This reduces the aerodynamic drag and overall weight of the aircraft, thereby improving the overall performance and range of the aircraft. It should be noted that the rocker arm 21 can only swing along the flight path of the aircraft. This avoids the rocker arm 21 not being able to swing fully along the flight path due to excessive friction between the tires and the ground during vertical landing. It also avoids the main shock absorber 22 becoming stuck due to the left-right swing of the rocker arm 21, reducing its efficiency, increasing the overload of the landing gear 100, and shortening its lifespan. Furthermore, it prevents tire deformation, which could lead to air leaks or tire blowouts.

[0042] In one embodiment, the main buffer 22 includes a main sleeve 221 and a first buffer rod 222. The main sleeve 221 is hinged to the inside of the machine body 3. The first end of the first buffer rod 222 is sleeved inside the main sleeve 221. The second end of the first buffer rod 222 extends out of the main sleeve 221 and is hinged to the rocker arm 21 near the hinge end 211. The first buffer rod 222 can extend and retract along the extension direction of the main sleeve 221.

[0043] The first buffer rod 222 is hinged to the rocker arm 21, allowing the first buffer rod 222 to adaptively adjust its connection angle with the rocker arm 21 when the rocker arm 21 rotates relative to the body 3, thereby better buffering and absorbing impact forces. The swing of the rocker arm 21 along the flight direction can prevent the rocker arm 21 from bearing bending moments caused by lateral and transverse forces, and prevent the main buffer 22 from jamming or being damaged due to bending moments. At the same time, the hinge point between the rocker arm 21 and the first buffer rod 222 is set close to the hinge end 211 of the first buffer rod 222 to avoid excessive torque at the hinge end 211, thus improving safety. The telescopic design of the first buffer rod 222 reduces the overall volume of the main buffer 22, thereby reducing the frontal area and aerodynamic drag.

[0044] It should be noted that the main sleeve 221 and the first buffer rod 222 can be buffered by hydraulic means in the prior art. That is, the main sleeve 221 is the outer cylinder and the first buffer rod 222 is the inner cylinder, also called the piston rod. The piston rod is hollow and filled with oil, while the outer cylinder is filled with air, thereby achieving the effect of compression buffering. The relevant applications in the prior art are relatively mature, and will not be elaborated on here.

[0045] In one embodiment, the two main landing gears 2 are symmetrical about the plane of symmetry of the fuselage 3 (see...). Figure 2 The dashed lines P in the diagram are symmetrically arranged. Each rocker arm 21 is inclined towards the tail side of the body 3 from the hinge end 211 to the free end 212, and from the hinge end 211 to the free end 212, it is inclined away from the other rocker arm 21. The hinge end 211 is hinged to the inside of the body 3 through a hinge rod 24, and the hinge rod 24 is perpendicular to the symmetry plane of the body 3.

[0046] The symmetrically arranged main landing gear 2 provides uniform support during takeoff, landing, and taxiing, balancing the aircraft's center of gravity and preventing tilting and rolling due to uneven loading, thus improving the overall stability of the landing gear 100. Simultaneously, the symmetrical arrangement ensures that the impact force during landing is evenly transmitted to the fuselage 3, reducing localized stress concentration, minimizing fatigue and damage caused by uneven stress, and extending the lifespan of the aircraft and landing gear 100. Furthermore, the rocker arm 21 tilts from its hinged end 211 to its free end 212 towards the tail side of the fuselage 3, and from its hinged end 211 to its free end 212 away from the other rocker arm 21, forming an outward-tilting V-shaped structure towards the tail. This increases the main wheel track to meet anti-rollover and anti-backward-rollover requirements. Additionally, the rocker arm 21 tilts downwards and rearwards, lowering the overall height of the main landing gear 2, making it suitable for aircraft with shorter fuselages, lowering the overall center of gravity, and improving stability.

[0047] Furthermore, the rocker arm 21 includes an elastic damping rod 213 and a rocker arm joint 214 connected to each other. The end of the rocker arm joint 214 away from the elastic damping rod 213 forms a hinge end 211. The end of the rocker arm joint 214 near the elastic damping rod 213 is provided with a hinge seat 215 that is hinged to the second end of the first buffer rod 222. The hinge end 211 is provided with a hinge ring 216. The hinge rod 24 is rotatably passed through the hinge ring 216. The end of the elastic damping rod 213 away from the rocker arm joint 214 forms a free end 212.

[0048] The shock absorber 213 is made of elastic material. If the impact force is too large during landing, or if the buffer travel of the landing gear 100 is limited, the elastic shock absorber 213 can absorb the impact force through its own deformation, thereby further improving the buffering performance of the nose landing gear 1. At the same time, the rocker arm joint 214 is connected to the first buffer bar 222 through the hinge seat 215, which improves the hinge stability between the first buffer bar 222 and the rocker arm joint 214. The rocker arm 21 is made of rigid material, which makes the rocker arm 21 less prone to deformation. This can prevent the deformation of the rocker arm 21 from causing the main buffer 22 to bear the bending moment generated by the lateral and transverse forces, and prevent the main buffer 22 from getting stuck or damaged due to the bending moment. The first buffer bar 222 is not connected to the elastic shock absorber 213, so that the deformation of the elastic shock absorber 213 does not affect the stability of the hinge. This can also prevent the main buffer 22 from bearing the bending moment generated by the lateral and transverse forces, thus extending the service life of the main buffer 22.

[0049] It should be noted that the number of hinge rings 216 can be one or two. One hinge ring 216 provides higher rigidity, while two hinge rings 216 can increase the contact area between the hinge rod 24 and the hinge end 211, thereby improving the load-bearing capacity and the uniformity of force distribution, and preventing the hinge end 211 from falling off or being damaged. The number of hinge rings 216 can be selected according to actual design requirements. The hinge rod 24 is perpendicular to the plane of symmetry, which limits the rotation direction of the rocker arm 21, ensuring that the rocker arm 21 can only rotate around the hinge rod 24. Directional rotation, or swinging along the heading, can prevent the rocker arm 21 from not swinging fully due to excessive friction between the tires and the ground during vertical landing. It can also prevent the main buffer 22 from bearing bending moments caused by lateral and transverse forces when the rocker arm 21 swings. This prevents the main buffer 22 from getting stuck or damaged due to bending moments caused by lateral and transverse forces, thereby reducing the overload of the landing gear 100, extending the life of the landing gear 100, and preventing the main wheel 23 from being deformed, causing air leakage or tire blowout, thus improving the overall stability of the aircraft.

[0050] It should be noted that the elastic damping rod 213 can be made of metal materials with low elastic modulus, such as tube springs, leaf springs, and elastic steel. Specifically, it can be made of elastic steel with an elastic modulus of 210 Gpa, titanium alloy with an elastic modulus of 110 Gpa, or aluminum alloy with an elastic modulus of 70 Gpa. The elastic modulus of the material of the elastic damping rod 213 must be lower than 210 Gpa to meet its deformation requirements.

[0051] In one embodiment, the rocker arm 21 has an elliptical cross-section, with one major axis of the cross-section facing the nose of the fuselage 3. The elliptical cross-section of the rocker arm 21 can effectively reduce air resistance, and when the major axis is aligned with the flight direction (nose direction) of the aircraft, the frontal area can be further reduced, thereby reducing the impact of airflow on the rocker arm 21 during flight and optimizing the aerodynamic performance of the aircraft.

[0052] In one embodiment, each main wheel 23 is tilted from top to bottom toward the other main wheel 23, and the angle between the axis of each main wheel 23 and the horizontal plane is 1° to 1.5°. Understandably, since the rocker arm 21 is tilted outwards, the impact force received during descent will cause the main wheel 23 to generate lateral friction outwards. To avoid excessive lateral friction causing deformation, air leakage, or damage to the main wheel 23, each main wheel 23 is tilted inwards at a certain angle from top to bottom toward the other main wheel 23. This inward tilt angle causes the outer side of the main wheel 23 to slightly lift when it contacts the ground, reducing lateral friction between the main wheel 23 and the ground. This protects the main wheel 23 while reducing the stress on the hub and bearings, thus improving the safety and stability of the main wheel 23. The tilt angle of the main wheel 23 (see...) Figure 2 The effect is best when the angle A is 1°~1.5°. The tilt angle of the main wheel 23 is the same as the angle between the axis of the main wheel 23 and the horizontal plane. Therefore, the angle between the axis of each main wheel 23 and the horizontal plane is 1°~1.5°.

[0053] Please see Figure 5 In one embodiment, the front buffer strut 12 is inclined from top to bottom toward the nose side of the fuselage 3, and the angle between the axis of the front buffer strut 12 and the vertical line is 7° to 7.5°. The front buffer strut 12 is inclined forward at a certain angle ( Figure 5 (Angle B in the diagram), and the strut axis is at a certain distance from the wheel axis, which increases the mechanical stabilizer distance L1 and the geometric stabilizer distance L2 of the nose landing gear 1 (see...). Figure 5 The L1 and L2 in the model allow the front wheel to generate a centering torque during takeoff and make it less prone to swaying.

[0054] Please see Figure 6In one embodiment, the front wheel 11 is covered with a fairing, and in another embodiment, each main wheel 23 is covered with a fairing 4. Understandably, the fairing 4 is used to reduce wind resistance, and the fairing 4 is teardrop-shaped to further reduce the aerodynamic drag at the main wheels 23 and the front wheel 11 of the aircraft.

[0055] In one embodiment, the front buffer support 12 includes a front sleeve 121 and a second buffer rod 122. The front sleeve 121 is connected to the machine body 3 and is at least partially disposed inside the machine body 3. The first end of the second buffer rod 122 is sleeved inside the front sleeve 121, and the second end of the second buffer rod 122 passes through the machine body 3 and extends out of the front sleeve 121. The second buffer rod 122 can extend and retract along the extension direction of the front sleeve 121. The front wheel 11 is connected to the second end of the second buffer rod 122 through a wheel fork 15.

[0056] At least a portion of the front sleeve is located inside the fuselage, with only a portion of the front sleeve and the second buffer rod 122 extending outside the fuselage 3. This reduces the volume of the front buffer support 12 exposed outside the fuselage, thereby reducing the frontal area and aerodynamic drag.

[0057] It should be noted that the front sleeve 121 and the second buffer rod 122 can be buffered by hydraulic means in the prior art. That is, the front sleeve 121 is the outer cylinder and the second buffer rod 122 is the inner cylinder, also called the piston rod. The piston rod is hollow and filled with oil, while the outer cylinder is filled with air, thereby achieving the effect of compression buffering. The relevant applications in the prior art are relatively mature, and will not be elaborated on here.

[0058] In one embodiment, the front sleeve 121 is hinged inside the fuselage 3, and the front landing gear 1 also includes a front strut 16. One end of the front strut 16 is connected to the fuselage 3, and the other end is connected to the front sleeve 121. The front strut 16 is used to stabilize and limit the front landing gear 1 to prevent the front landing gear 1 from rotating too much.

[0059] In one embodiment, the front landing gear 1 further includes a sway damping assembly 13, which includes a sway damping body 131, a torque arm 132, and a transmission ring 133. The sway damping body 131 is sleeved on the outside of the front sleeve 121, and the transmission ring 133 is sleeved on the outside of the second end of the second buffer rod 122. The torque arm 132 includes two hinged first arms 134 and second arms 135. The end of the first arm 134 away from the second arm 135 is hinged to the sway damping body 131, and the end of the second arm 135 away from the first arm 134 is hinged to the transmission ring 133. The sway damping body 131 is a sway damper or a steering servo.

[0060] The rotation of the wheel fork 15 is transmitted to the torque arm 132 via the transmission ring 133, and then to the anti-sway body 131 via the torque arm 132. During landing or high-speed taxiing, the nose wheel 11 may experience high-frequency swaying due to uneven ground or other factors. The anti-sway body 131 provides damping to reduce the rapid left-right swaying of the nose wheel, preventing swaying and enabling the aircraft to maintain a stable course during ground taxiing, reducing the risk of yaw and runway deviation caused by tire swaying. The hinged relationship between the first lever arm 134 and the second lever arm 135 and the anti-sway body 131 and the transmission ring 133 allows the torque arm 132 to adjust the angle between the first lever arm 134 and the second lever arm 135 during the extension and retraction of the second buffer rod 122, preventing damage to the torque arm 132 during the extension and retraction of the second buffer rod 122.

[0061] The sway damper can be any existing sway damper. When the sway damper body 131 is a turning servo, it can give the front wheel 11 the ability to actively turn. At the same time, the turning servo itself also has a sway damping function, thereby improving the stability of the aircraft and its functionality.

[0062] In one embodiment, wheel-mounted switches are provided on the nose landing gear 1 and / or the main landing gear 2, and the wheel-mounted switches are configured to send air-to-ground signals.

[0063] Specifically, in one embodiment, a main wheel-mounted switch 25 is provided at the hinge end 211 of the rocker arm joint 214. The main wheel-mounted switch 25 is used to send a ground-to-air signal to the flight control computer when the rocker arm 21 swings, thereby determining whether the aircraft is in the take-off state or the landing state. Understandably, the main wheel-mounted switch 25 can be a mature wheel-mounted switch in the prior art, which will not be described here.

[0064] In another embodiment, a baffle 17 and a front wheel switch 14 are provided at the hinge of the first lever arm 134 and the second lever arm 135. The baffle 17 is used to cooperate with the front wheel switch 14 to send a ground-to-air signal to the flight control computer when the second buffer rod 122 is in motion, so as to determine whether the aircraft is in the take-off state or the landing state. Understandably, the front wheel switch 14 can be a mature wheel switch in the prior art, which will not be described here.

[0065] The main landing gear is a non-traditional strutless rocker arm landing gear, with the rocker arm extending to the side and rear, primarily to meet the requirements for anti-rollover angle and anti-rollover angle, as well as some other design parameters. According to landing gear design specifications, the landing gear arrangement must ensure that the anti-rollover angle is greater than the rollover angle and the anti-rollover angle is ≤63°. The anti-rollover angle in this application is 62.54°, meeting the requirements.

[0066] Please see Figure 7The present invention also provides a landing gear design method, applied to the aforementioned landing gear, the landing gear design method comprising the following steps:

[0067] S100: Based on the landing kinetic energy and potential energy of the aircraft, configure the energy absorption of the main landing gear, wherein the energy absorption of the main landing gear includes the energy absorption of the main landing gear wheel, the energy absorption of the rocker arm and the energy absorption of the main shock absorber.

[0068] First, determine the landing kinetic energy of the aircraft as... Potential energy is (GL)* ,

[0069] S200: Determine the theoretical overload value of the aircraft based on its vertical landing energy and / or taxiing landing energy;

[0070] S300: Calculate the buffer stroke of the main buffer based on the theoretical overload value and the law of conservation of energy;

[0071] Specifically, according to the law of conservation of energy, the sum of the landing kinetic energy and potential energy of the aircraft is equal to the energy absorbed by the main landing gear 2's main wheels 23, rocker arms 21, and main shock absorber 22. Therefore, the buffer stroke of the main shock absorber 22 can be calculated using the following formula:

[0072] +(GL)* = NG+ NG+ NG

[0073] in This indicates the equivalent mass of main landing gear 2. G represents the landing descent speed of the aircraft, and G represents the landing equivalent weight of the main landing gear 2 (i.e., *g), L represents the lift during the aircraft's landing process, which is a coefficient not greater than 1 multiplied by G (usually taken as 0.67 or 1), and N is the landing gear 2 landing overload. NG represents the energy absorbed by the main buffer 22 ( Main buffer travel, (main buffer efficiency coefficient) NG indicates the energy absorbed by the elastic damping rod 213 of the rocker arm 21. For the travel of the rocker arm's elastic damping rod, (Efficiency coefficient of the elastic damping rod of the rocker arm). NG indicates the energy absorbed by the main wheel 23 ( For the compression stroke of the main engine wheel, (Efficiency coefficient of the main wheel).

[0074] In the above formulas, the stroke of the elastic damping rod of rocker arm 21 changes linearly, satisfying Hooke's Law F=kx, which can be verified through finite element simulation, with an efficiency coefficient of 0.45; the stroke of the main wheel 23 can be obtained from the tire static pressure curve, with an efficiency coefficient of 0.47; the efficiency coefficient of the main buffer 22 ranges from 0.65 to 0.80. Considering the structural form and design experience of the main buffer 22, the initial value can be taken as 0.75, and iteratively applied in subsequent design processes. Therefore, the above formulas only contain the overload N and the stroke of the main buffer. The value is not yet determined, while N is the theoretical overload value determined in step S200. The main buffer stroke is calculated using the above formula. ;

[0075] S400: Based on the buffer stroke of the main buffer, perform drop simulation to obtain at least the overload simulation value and stroke simulation value of the main landing gear;

[0076] Based on the main buffer 22 stroke Drop simulation was performed to obtain the overall overload simulation value of the main landing gear 2, the overall stroke simulation value of the main landing gear, and the overall simulation efficiency coefficient of the main landing gear.

[0077] S500: At least adjust the theoretical overload value, iteratively calculate the simulated overload value and the simulated travel value until the ratio of the energy absorbed by the main landing gear to the sum of the landing kinetic energy and potential energy of the aircraft is greater than or equal to the ratio threshold, and / or the difference between the sum of the landing kinetic energy and potential energy of the aircraft and the energy absorbed by the main landing gear is less than or equal to the difference threshold.

[0078] Specifically, taking a difference threshold of 30% of the aircraft's landing kinetic and potential energy, and a ratio threshold of 70%, as an example, the simulated energy absorbed by the main landing gear 2 is calculated using the overall overload simulation value, overall stroke simulation value, overall simulation efficiency coefficient, and equivalent mass of the main landing gear 2. This energy is then compared with the sum of the aircraft's landing kinetic and potential energy. If the difference between the sum of the aircraft's landing kinetic and potential energy and the simulated absorbed energy is greater than or equal to 30% of the sum of the aircraft's landing kinetic and potential energy... If the ratio of the simulated absorbed energy to the sum of the aircraft's landing kinetic energy and potential energy is less than or equal to 70% of the sum of the aircraft's landing kinetic energy and potential energy, then adjust the theoretical overload value and iterate the overload simulation value, stroke simulation value, and simulation efficiency coefficient until the difference between the calculated simulated absorbed energy of the main landing gear 2 as a whole and the sum of the aircraft's landing kinetic energy and potential energy is less than 30% of the sum of the aircraft's landing kinetic energy and potential energy, or the ratio of the simulated absorbed energy of the main landing gear 2 as a whole to the sum of the aircraft's landing kinetic energy and potential energy is greater than 70%, then the design requirements are met.

[0079] The main landing gear absorbs landing energy through a three-stage buffer system, increasing material utilization and reducing landing overload. The three-stage buffer structure consists of the main landing gear wheels, the rocker arm's elastic damping rods, and the main shock absorber. During the design process, the kinetic energy of the aircraft landing was first determined to be... Potential energy is (GL)* Therefore, according to the law of conservation of energy, we can obtain:

[0080] +(GL)* = NG+ NG+ NG

[0081] The present invention also proposes an aircraft having a fuselage 3, the bottom of which is provided with the aforementioned landing gear 100. The specific structure of the landing gear 100 is as described in the above embodiments. Since this aircraft adopts all the technical solutions of all the above embodiments, it has at least all the beneficial effects brought about by the technical solutions of the above embodiments, which will not be elaborated here.

[0082] The above description is merely an optional embodiment of the present invention and does not limit the patent scope of the present invention. Any equivalent structural transformations made based on the content of the present invention specification and drawings, or direct / indirect applications in other related technical fields, are included within the patent protection scope of the present invention.

Claims

1. A landing gear, characterized in that, The landing gear is disposed at the bottom of the aircraft fuselage, and the landing gear includes: The front landing gear is located on the front side of the bottom of the fuselage. The front landing gear includes a front wheel and a front shock strut. One end of the front shock strut is connected to the fuselage and located inside the fuselage, and the other end extends out of the fuselage and is elastically connected to the front wheel. Two main landing gears are respectively disposed on opposite sides of the bottom of the fuselage. Each main landing gear includes a rocker arm, a main shock absorber, and a main wheel. The main shock absorber includes a first shock rod. The two ends of the rocker arm are a hinged end and a free end, respectively. The hinged end is hinged to the inside of the fuselage so that the rocker arm can only swing along the heading of the aircraft. The free end is connected to the main wheel. The position on the rocker arm between the hinged end and the free end is elastically connected to the fuselage through the main shock absorber. The two main landing gears are symmetrically arranged about the plane of symmetry of the fuselage. Each rocker arm is inclined towards the tail side of the fuselage from the hinge end to the free end, and is inclined away from the other rocker arm from the hinge end to the free end. The hinge end is hinged to the inside of the fuselage through a hinge rod, and the hinge rod is perpendicular to the plane of symmetry of the fuselage. The rocker arm includes an elastic damping rod and a rocker arm joint connected to each other. The end of the rocker arm joint away from the elastic damping rod forms the hinge end. The end of the rocker arm joint near the elastic damping rod is provided with a hinge seat that is hinged to the first buffer rod. The hinge end is provided with a hinge ring. The hinge rod is rotatably passed through the hinge ring. The end of the elastic damping rod away from the rocker arm joint forms the free end. Wheel-mounted switches are provided on the nose landing gear and / or the main landing gear, and the wheel-mounted switches are configured to send air-to-ground signals.

2. The landing gear as claimed in claim 1, characterized in that, The main buffer also includes a main sleeve, which is hinged to the inside of the machine body. The first end of the first buffer rod is sleeved inside the main sleeve, and the second end of the first buffer rod extends out of the main sleeve and is hinged to the rocker arm near the hinge end. The first buffer rod can extend and retract along the extension direction of the main sleeve.

3. The landing gear as described in any one of claims 1 to 2, characterized in that, The rocker arm has an elliptical cross-section, and one long axis of the rocker arm's cross-section faces the nose side of the fuselage.

4. The landing gear as described in any one of claims 1 to 2, characterized in that, Each of the main drive wheels is inclined from top to bottom toward the other main drive wheel, and the angle between the axis of each main drive wheel and the horizontal plane is 1° to 1.5°; or, the front buffer support is inclined from top to bottom toward the nose side of the body, and the angle between the axis of the front buffer support and the vertical line is 7° to 7.5°.

5. The landing gear as described in any one of claims 1 to 2, characterized in that, The front wheel and / or each of the main wheels are provided with a fairing.

6. A landing gear design method, characterized in that, The landing gear design method, applicable to any one of claims 1 to 5, comprises the following steps: The main landing gear is configured to absorb energy based on the landing kinetic energy and potential energy of the aircraft, wherein the energy absorption of the main landing gear includes energy absorption by the main landing gear wheels, energy absorption by the rocker arms, and energy absorption by the main shock absorbers. The theoretical overload value of the aircraft is determined based on the vertical landing energy and / or the taxiing landing energy of the aircraft. The buffer stroke of the main buffer is calculated based on the theoretical overload value and the law of conservation of energy. Based on the buffer stroke of the main buffer, drop simulation is performed to obtain at least the overload simulation value and stroke simulation value of the main landing gear; At least the theoretical overload value is adjusted, and the simulated overload value and simulated travel value are repeatedly calculated iteratively until the ratio of the energy absorbed by the main landing gear to the sum of the landing kinetic energy and potential energy of the aircraft is greater than the ratio threshold, and / or the difference between the sum of the landing kinetic energy and potential energy of the aircraft and the energy absorbed by the main landing gear is less than the difference threshold.

7. An aircraft, characterized in that, The aircraft has a fuselage, and the bottom of the fuselage is provided with landing gear as described in any one of claims 1 to 5.

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