Undercarriage and aircraft

Abstract

The invention discloses an undercarriage and an aircraft, and relates to the technical field of aircraft structural parts, the undercarriage comprises a front undercarriage and two main undercarriages, the front undercarriage is arranged on the front side of an aircraft body, the front undercarriage comprises a front aircraft wheel and a front buffer supporting column, one end of the front buffer supporting column is connected with the aircraft body and arranged in the aircraft body, and the other end of the front buffer supporting column is connected with the aircraft body; the other end penetrates out of the machine body and is elastically connected with the front machine wheel; the two main undercarriages are arranged on the two opposite sides of the aircraft body respectively, each main undercarriage comprises a rocker arm, a main buffer and a main wheel, the two ends of each rocker arm are a hinged end and a free end respectively, the hinged ends are hinged to the inner side of the aircraft body so that the rocker arms can only swing in the course of the aircraft, and the free ends are connected with the main wheels. A part of the structure of the front undercarriage is arranged in the aircraft body, and the rocker arm of the main undercarriage is directly hinged to the aircraft body, so that the overall windward area of the undercarriage is reduced, the aerodynamic resistance and the overall weight of the aircraft are reduced, and the overall performance and endurance of the aircraft are improved.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft structural parts, and in particular to a landing gear and an aircraft. Background Art

[0002] As the low-altitude economy enters a period of rapid development, UAM (Urban Air Mobility) and eVTOL (Electric Vertical Takeoff and Landing) aircraft have become the two most popular concepts in the field of low-altitude economy. The landing gear system is a very important component in the aircraft and is the key carrier for the aircraft to achieve ground support, take-off, landing and taxiing.

[0003] The wheeled landing gear used in eVTOL in the prior art is usually a non-retractable landing gear, 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 heavy, and the windward area of ​​the non-retractable landing gear is too large, resulting in excessive aerodynamic resistance. Both factors will affect the overall performance and endurance of the aircraft. Summary of the invention

[0004] The main purpose of the present invention is to provide a landing gear and an aircraft, aiming to solve the problem in the prior art that the landing gear is too heavy and the aerodynamic resistance is too large, thereby affecting the overall performance and endurance of the aircraft.

[0005] In order to achieve the above-mentioned object, the present invention provides a landing gear, which is arranged at the bottom of the aircraft body, and comprises:

[0006] A front landing gear, the front landing gear is arranged at the front side of the bottom of the fuselage, the front landing gear comprises a front wheel and a front buffer strut, one end of the front buffer strut is connected to the fuselage and arranged inside the fuselage, and the other end passes through the outside of the fuselage and is elastically connected to the front wheel;

[0007] Two main landing gears, the two main landing gears are respectively arranged on opposite sides of the bottom of the fuselage, each of the main landing gears includes a rocker arm, a main buffer and a main wheel, the two ends of the rocker arm are respectively a hinged end and a free end, the hinged end is hinged to the inner side 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, and the position on the rocker arm located between the hinged end and the free end is elastically connected to the fuselage through the main buffer.

[0008] In one embodiment, the main buffer includes a main sleeve and a first buffer rod, the main sleeve is hinged to the inner side of the body, the first end of the first buffer rod is sleeved in the main sleeve, the second end of the first buffer rod extends out of the main sleeve and is hinged to a position on the rocker arm close to the hinged end, and the first buffer rod can be extended and retracted along the extension direction of the main sleeve.

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

[0010] In one embodiment, the rocker arm includes an elastic shock-absorbing rod and a rocker arm joint which are connected to each other, and the end of the rocker arm joint away from the elastic shock-absorbing rod forms the hinged end, and the end of the rocker arm joint close to the elastic shock-absorbing rod is provided with a hinged seat hinged to the second end of the first buffer rod; the hinged end is provided with two hinged rings arranged at intervals, and the two ends of the hinged rod are respectively rotatably inserted into the two hinged rings, and the end of the elastic shock-absorbing rod away from the rocker arm joint forms the free end.

[0011] In one embodiment, the cross section of the rocker arm is elliptical, and a long axis side of the cross section of the rocker arm faces the nose side of the machine body.

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

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

[0014] In one embodiment, a wheel-mounted switch is disposed on the front landing gear and / or the main landing gear, and the wheel-mounted switch is configured to send an air-to-ground signal.

[0015] The present invention also provides a landing gear design method, which is applied to the above-mentioned landing gear. The landing gear design method comprises the following steps:

[0016] configuring the energy absorption of the main landing gear based on the landing kinetic energy and potential energy of the aircraft, wherein the energy absorption of the main landing gear includes the energy absorption of the main wheels, the energy absorption of the rocker arms and the energy absorption of the main buffers;

[0017] determining a theoretical overload value of the aircraft based on the vertical landing energy and / or the rolling landing energy of the aircraft;

[0018] Calculating the buffering stroke of the main buffer based on the overload theoretical value and the law of conservation of energy;

[0019] Performing a drop shock simulation based on the buffer stroke of the main buffer to at least obtain an overload simulation value and a stroke simulation value of the main landing gear;

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

[0021] The present invention further provides an aircraft, which comprises a fuselage, and the above-mentioned landing gear is arranged at the bottom of the fuselage.

[0022] The technical solution of the present invention is to arrange a part of the front buffer strut of the front landing gear inside the fuselage, and only extend one end of the front buffer strut outside the fuselage and elastically connect it to the front wheel, so that most of the structure of the front landing gear is accommodated inside the fuselage, thereby reducing the frontal area of ​​the front landing gear and reducing the aerodynamic drag. At the same time, the main landing gear is directly hinged to the outside of the fuselage through a rocker arm, without the need to be indirectly connected to the fuselage by setting an intermediate transfer structure, thereby reducing the number of structural parts of the main landing gear, thereby reducing the frontal area of ​​the main landing gear and reducing the overall weight of the main landing gear. The present invention arranges part of the structure of the front landing gear in the fuselage and directly hinges the rocker arm of the main landing gear to the fuselage, thereby reducing the overall frontal area of ​​the landing gear, reducing the aerodynamic drag and overall weight of the aircraft, and improving the overall performance and endurance of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying creative work.

[0024] Figure 1 A schematic diagram of the assembly relationship between the landing gear and the fuselage provided by the present invention;

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

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

[0027] Figure 4 A schematic structural diagram of the main landing gear provided by the present invention;

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

[0029] Figure 6 The assembly structure diagram of the landing gear and fairing provided by the present invention;

[0030] Figure 7 A flowchart of a main landing gear design method provided by the present invention.

[0031] Description of Figure Numbers:

[0032] 100. Landing gear; 1. Front landing gear; 11. Front wheel; 12. Front buffer strut; 121. Front sleeve; 122. Second buffer rod; 13. Swing reduction assembly; 131. Swing reduction body; 132. Torque arm; 133. Transmission ring; 134. First lever arm; 135. Second lever arm; 14. Front wheel load switch; 15. Wheel fork; 16. Front support rod; 17. Baffle; 2. Main landing gear; 21. Rocker arm; 211. Articulated end; 212. Free end; 213. Elastic shock-absorbing rod; 214. Rocker arm joint; 215. Articulated seat; 216. Articulated ring; 22. Main buffer; 221. Main sleeve; 222. First buffer rod; 23. Main wheel; 24. Articulated rod; 25. Main wheel load switch; 3. Airframe; 4. Fairing.

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

[0034] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work 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 indications (such as up, down, left, right, front, back, etc.), the directional indications are only used to explain the relative position relationship, movement status, etc. between the components in a certain specific posture. If the specific posture changes, the directional indication will also change accordingly.

[0036] In addition, if there are descriptions involving "first", "second", etc. in the embodiments of the present invention, the descriptions of "first", "second", etc. are only used for descriptive purposes and cannot be understood as indicating or implying their relative importance or implicitly indicating the number of technical features indicated. Therefore, the features limited to "first" and "second" may explicitly or implicitly include at least one of the features. In addition, if "and / or" or "and / or" appears in the full text, its meaning includes three parallel solutions. Taking "A and / or B" as an example, it includes solution A, solution B, or solutions that satisfy both A and B. In addition, the technical solutions between the various embodiments can be combined with each other, but it must be based on the ability of ordinary technicians in this field to implement. When the combination of technical solutions is contradictory or cannot be implemented, it should be deemed that such combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0037] The wheeled landing gear used in eVTOL in the prior art is usually a non-retractable landing gear, 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 heavy, and the windward area of ​​the non-retractable landing gear is too large, resulting in excessive aerodynamic resistance. Both factors will affect the overall performance and endurance of the aircraft.

[0038] In order to solve the above problems, the present invention provides a landing gear 100 .

[0039] Please combine Figures 1 to 4 The landing gear 100 of the present embodiment is arranged at the bottom of the fuselage 3 of the aircraft. The landing gear 100 includes a front landing gear 1 and two main landing gears 2. The front landing gear 1 is arranged at the front side of the bottom of the fuselage 3. The front landing gear 1 includes a front wheel 11 and a front buffer strut 12. One end of the front buffer strut 12 is connected to the fuselage 3 and arranged inside the fuselage 3, and the other end passes through the fuselage 3 and is elastically connected to the front wheel 11; the two main landing gears 2 are respectively arranged on opposite sides of the bottom of the fuselage 3. Each main landing gear 2 includes a rocker arm 21, a main buffer 22 and a main wheel 23. The two ends of the rocker arm 21 are respectively a hinged end 211 and a free end 212. The hinged end 211 is hinged to the inner side 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 between the hinged end 211 and the free end 212 on the rocker arm 21 is elastically connected to the fuselage 3 through the main buffer 22.

[0040] It can be understood that the direction of travel of the aircraft is the front-to-back direction, and the two main landing gears 2 are respectively arranged on the left and right sides of the fuselage 3, and usually, the two main landing gears 2 are symmetrically arranged. When landing, the front wheel 11 contacts the ground, and the deformation of the front wheel 11 itself can absorb part of the impact force, and the front buffer strut 12 elastically connected to the front wheel 11 is elastically compressed under the impact force generated by the weight of the fuselage, thereby absorbing another part of the impact force; and when the main wheel 23 contacts the ground, the deformation of the main wheel 23 itself can absorb part of the impact force, the rocker arm 21 rotates relative to the fuselage 3 under the impact force, absorbing another part of the impact force, and the main buffer 22 elastically connected to the rocker arm 21 is elastically compressed under the action of the impact force, thereby absorbing another part of the impact force, so that the aircraft can land smoothly.

[0041] The technical solution of the present invention is to dispose a part of the front buffer strut 12 of the front landing gear 1 inside the fuselage 3, and only extend one end of the front buffer strut 12 outside the fuselage 3 and elastically connected to the front wheel 11, so that most of the structure of the front landing gear 1 is accommodated inside the fuselage 3, thereby reducing the frontal area of ​​the front landing gear 1 and reducing aerodynamic resistance. At the same time, the main landing gear 2 is directly hinged to the inner side of the fuselage 3 through the rocker arm 21, without the need to be indirectly connected to the fuselage 3 through the intermediate transfer structure, thereby reducing the number of structural parts of the main landing gear 2, thereby reducing the frontal area of ​​the main landing gear 2 and reducing the overall weight of the main landing gear 2. The main landing gear 2 adopts a rocker arm landing gear, and the rocker arm 21 can be set at any position to meet the parameter requirements of the aircraft, thereby improving the flexibility of the layout and reducing the overall volume of the landing gear 100. The present invention reduces the overall frontal area of ​​the landing gear 100 by arranging part of the structure of the front landing gear 1 in the fuselage 3 and directly hinges the rocker arm 21 of the main landing gear 2 to the fuselage 3, reduces the aerodynamic drag and the overall weight of the aircraft, and thus improves the overall performance and endurance of the aircraft. It should be noted that the rocker arm 21 can only swing along the heading of the aircraft, which avoids the rocker arm 21 from swinging left and right when the friction between the tire and the ground is too large during vertical landing, resulting in the rocker arm 21 not being able to fully swing along the heading, avoids the main buffer 22 from being stuck and reducing the efficiency of the main buffer 22 due to the left and right swing of the rocker arm 21, and can also avoid the problem of increasing the overload of the landing gear 100 and reducing the life of the landing gear 100, and can prevent the tire from being deformed and causing air leakage or tire blowout.

[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 inner side of the body 3. The first end of the first buffer rod 222 is sleeved in the main sleeve 221. The second end of the first buffer rod 222 extends out of the main sleeve 221 and is hinged to a position on the rocker arm 21 near the hinge end 211. The first buffer rod 222 can be extended and retracted along the extension direction of the main sleeve 221.

[0043] The first buffer rod 222 is hinged to the rocker arm 21, so that when the rocker arm 21 rotates relative to the body 3, the first buffer rod 222 can adaptively adjust its direct connection angle with the rocker arm 21, so as to better buffer and absorb the impact force. The swing of the rocker arm 21 along the heading can prevent the rocker arm 21 from being subjected to the bending moment caused by the lateral force and the transverse force, and prevent the main buffer 22 from being stuck or damaged due to the bending moment. At the same time, the hinge between the rocker arm 21 and the first buffer rod 222 is arranged close to the hinge end 211 of the first buffer rod 222, so as to avoid excessive torque at the hinge end 211, thereby improving safety. The retractable setting of the first buffer rod 222 reduces the overall volume of the main buffer 22, thereby reducing the windward area and aerodynamic resistance.

[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 an outer cylinder, and the first buffer rod 222 is an inner cylinder also called a piston rod. The interior of the piston rod is hollow and filled with oil, while the outer cylinder is filled with air, thereby achieving a compression buffering effect. The relevant applications in the prior art are relatively mature and will not be elaborated 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 dotted line P) in the figure is symmetrically arranged, each rocker arm 21 is inclined toward the tail side of the body 3 from the hinge end 211 to the free end 212, and is inclined from the hinge end 211 to the free end 212 in a direction away from the other rocker arm 21, and the hinge end 211 is hinged to the inner side 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 can provide uniform support during take-off, landing and taxiing, balance the center of gravity of the aircraft, avoid tilting and rolling due to eccentric loading, and improve the overall stability of the landing gear 100. At the same time, the symmetrical arrangement allows the impact force of the aircraft during landing to be evenly transmitted to the fuselage 3, reducing local stress concentration, reducing fatigue and damage caused by uneven force, and extending the life of the aircraft and the landing gear 100. At the same time, the rocker arm 21 is inclined from the hinge end 211 to the free end 212 toward the tail side of the fuselage 3, and is inclined from the hinge end 211 to the free end 212 in a direction away from the other rocker arm 21, forming an outward eight-shaped structure inclined toward the tail side. On the one hand, the main wheelbase can be increased to meet the requirements of the anti-rollover angle and the anti-backward rollover angle. On the other hand, the rocker arm 21 is inclined toward the rear and downward, reducing the overall height of the main landing gear 2, which is suitable for aircraft with a shorter fuselage, reducing the overall center of gravity of the aircraft, and improving stability.

[0047] Furthermore, the rocker arm 21 includes an elastic shock-absorbing rod 213 and a rocker arm joint 214 which are connected to each other. The end of the rocker arm joint 214 away from the elastic shock-absorbing rod 213 forms a hinged end 211, and the end of the rocker arm joint 214 close to the elastic shock-absorbing rod 213 is provided with a hinged seat 215 hinged to the second end of the first buffer rod 222; the hinged end 211 is provided with a hinge ring 216, and the hinge rod 24 is rotatably inserted into the hinge ring 216, and the end of the elastic shock-absorbing rod 213 away from the rocker arm joint 214 forms a free end 212.

[0048] The shock-absorbing rod is an elastic shock-absorbing rod 213 made of elastic material. If the impact force is too large during landing, or the buffering stroke of the landing gear 100 is limited, the elastic shock-absorbing rod 213 can absorb the impact force through its own deformation, thereby further improving the buffering performance of the front landing gear 1. At the same time, the rocker arm joint 214 is connected to the first buffer rod 222 through the hinge seat 215, which improves the hinge stability of the first buffer rod 222 and the rocker arm joint 214. The rocker arm 21 is made of rigid material, so that the rocker arm 21 is not easy to deform, which can prevent the deformation of the rocker arm 21 from causing the main buffer 22 to bear the bending moment caused by the lateral force and the transverse force, and prevent the main buffer 22 from being stuck or damaged due to the bending moment. The first buffer rod 222 is not connected to the elastic shock-absorbing rod 213, which prevents the deformation of the elastic shock-absorbing rod 213 from affecting the stability of the hinge, and can also prevent the main buffer 22 from bearing the bending moment caused by the lateral force and the transverse force, thereby 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 has higher rigidity, while two hinge rings 216 can increase the contact area between the hinge rod 24 and the hinge end 211, improve the pressure bearing capacity, and also improve the force uniformity, so as to avoid 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 symmetry plane, so that it can limit the rotation direction of the rocker arm 21, so that the rocker arm 21 can only rotate with the hinge rod 24 as the rotation axis. Directional rotation, that is, swinging along the heading, can prevent the rocker arm 21 from being unable to swing fully due to excessive friction between the tire and the ground when the aircraft lands vertically, and can also prevent the main buffer 22 from being subjected to the bending moment caused by the lateral force and the transverse force when the rocker arm 21 swings, so that the main buffer 22 will not be stuck or damaged due to the bending moment caused by the lateral force and the transverse force, thereby reducing the increase in the overload of the landing gear 100, extending the life of the landing gear 100, and can prevent the main wheel 23 from being deformed to cause air leakage or tire blowout, thereby improving the overall stability of the aircraft.

[0050] It should be noted that the elastic shock-absorbing rod 213 can be a metal material with a low elastic modulus, such as a tube spring, a leaf spring, or an elastic steel. Specifically, elastic steel with an elastic modulus of 210 Gpa, a titanium alloy with an elastic modulus of 110 Gpa, or an aluminum alloy with an elastic modulus of 70 Gpa can be selected. The elastic modulus of the material of the elastic shock-absorbing rod 213 needs to be lower than 210 Gpa to meet its deformation requirements.

[0051] In one embodiment, the cross section of the rocker arm 21 is elliptical, and a long axis side of the cross section of the rocker arm 21 faces the nose side of the fuselage 3. The rocker arm 21 with an elliptical cross section can effectively reduce air resistance, and when the long axis side is consistent with the flight direction (nose direction) of the aircraft, the frontal area can be further reduced, the influence of airflow on the rocker arm 21 during flight is reduced, and the aerodynamic performance of the aircraft is optimized.

[0052] In one embodiment, each main wheel 23 is inclined from top to bottom toward another main wheel 23, and the angle between the axis of each main wheel 23 and the horizontal plane is 1° to 1.5°. It can be understood that since the rocker arm 21 is inclined toward the outside, the impact force received during the landing process will cause the main wheel 23 to generate lateral friction toward the outside. In order to prevent the main wheel 23 from being deformed, leaked or damaged due to excessive lateral friction, each main wheel 23 is inclined from top to bottom toward another main wheel 23 by a certain inclination angle. The inclination angle causes the outer side of the main wheel 23 to be slightly lifted when it contacts the ground, thereby reducing the lateral friction between the main wheel 23 and the ground, protecting the main wheel 23 while reducing the stress on the hub and the bearing, thereby improving the safety and stability of the main wheel 23. The inclination angle of the main wheel 23 (see Figure 2 The best effect is achieved when the angle A) in the figure is 1° to 1.5°, and the inclination angle of the main wheel 23 is the same as the angle between the axis of the main wheel 23 and the horizontal plane, so the angle between the axis of each main wheel 23 and the horizontal plane is 1° to 1.5°.

[0053] See also Figure 5 In one embodiment, the front buffer pillar 12 is inclined from top to bottom toward the nose side of the body 3, and the angle between the axis of the front buffer pillar 12 and the vertical line is 7° to 7.5°. The front buffer pillar 12 is inclined forward at a certain angle ( Figure 5 The angle B in the middle), and the strut axis is at a certain distance from the wheel axis, which increases the mechanical stability distance L1 and geometric stability distance L2 of the front landing gear 1 (see Figure 5 L1 and L2 in the figure), so that the front wheels of the aircraft can generate a centering torque during the taxiing process and are not prone to shimmy.

[0054] See also Figure 6In one embodiment, the front wheel 11 is provided with a fairing on its outer cover, and in another embodiment, each main wheel 23 is provided with a fairing 4 on its outer cover. It can be understood that the fairing 4 is used to reduce wind resistance, and the fairing 4 is provided in a teardrop shape to further reduce the aerodynamic resistance at the main wheel 23 and the front wheel 11 of the aircraft.

[0055] In one embodiment, the front buffer pillar 12 includes a front sleeve 121 and a second buffer rod 122. The front sleeve 121 is connected to the body 3, and the front sleeve 121 is at least partially arranged inside the body 3. The first end of the second buffer rod 122 is sleeved in the front sleeve 121, and the second end of the second buffer rod 122 is passed through the body 3 and extends out of the front sleeve 121. The second buffer rod 122 can be retracted and extended 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 fork 15.

[0056] At least part of the front sleeve is arranged inside the body, and only a part of the front sleeve and the second buffer rod 122 extend outside the body 3, so that the volume of the front buffer strut 12 exposed outside the body is reduced, thereby reducing the frontal area and aerodynamic resistance.

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

[0058] In one embodiment, the front sleeve 121 is hinged inside the body 3, and the front landing gear 1 also includes a front strut 16, one end of the front strut 16 is connected to the body 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 also includes a sway reduction assembly 13, which includes a sway reduction body 131, a torque arm 132 and a transmission ring 133. The sway reduction 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 mutually hinged first lever arms 134 and second lever arms 135. One end of the first lever arm 134 away from the second lever arm 135 is hinged to the sway reduction body 131, and one end of the second lever arm 135 away from the first lever arm 134 is hinged to the transmission ring 133, wherein the sway reduction body 131 is a sway reducer or a turning servo.

[0060] The rotation of the wheel fork 15 is transmitted to the torque arm 132 through the transmission ring 133, and then transmitted to the sway reduction body 131 through the torque arm 132. When landing or gliding at high speed, the front wheel 11 may have high-frequency oscillation due to uneven ground or other factors. The sway reduction body 131 reduces the rapid left and right shaking of the front wheel by providing a damping effect, avoids the occurrence of swaying phenomenon, enables the aircraft to maintain a stable driving direction when gliding on the ground, and reduces the risk of yaw and runway deviation caused by tire swinging. The first force arm 134 and the second force arm 135 are hinged to the sway reduction body 131 and the transmission ring 133, so that when the second buffer rod 122 is extended and retracted, the torque arm 132 can be extended and retracted by swinging to adjust the angle between the first force arm 134 and the second force arm 135, thereby avoiding damage to the torque arm 132 when the second buffer rod 122 is extended and retracted.

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

[0062] In one embodiment, a wheel-mounted switch is disposed on the front landing gear 1 and / or the main landing gear 2, and the wheel-mounted switch is configured to send an air-to-ground signal.

[0063] Specifically, in one embodiment, a main wheel-mounted switch 25 is provided at the hinged 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, so as to determine whether the aircraft is in a take-off state or a landing state. It can be understood that the main wheel-mounted switch 25 can adopt a mature wheel-mounted switch in the prior art, which will not be explained here.

[0064] In another embodiment, a baffle 17 and a front wheel load 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 load switch 14 to send a ground-to-air signal to the flight control computer when the second buffer rod 122 moves, so as to determine whether the aircraft is in a take-off state or a landing state. It can be understood that the front wheel load switch 14 can adopt a mature wheel load switch in the prior art, which will not be explained here.

[0065] The main landing gear is a non-traditional pillarless rocker-arm landing gear, with the rocker arm extending to the side and rear, mainly to meet the requirements of anti-rollback angle and anti-rollover angle and some other design parameters. According to the landing gear design specification, the layout of the landing gear must meet the requirements that the anti-rollback angle is greater than the rollback angle; the anti-rollover angle must be ≤63°. The anti-rollover angle of this application is 62.54°, which meets the index requirements.

[0066] See also Figure 7The present invention also provides a landing gear design method, which is applied to the above-mentioned landing gear. The landing gear design method comprises the following steps:

[0067] S100: configuring the energy absorption of the main landing gear based on the landing kinetic energy and potential energy of the aircraft, wherein the energy absorption of the main landing gear includes energy absorption of the main wheels, energy absorption of the rocker arms and energy absorption of the main buffers;

[0068] First, determine the landing kinetic energy of the aircraft as Potential energy is (GL)*(S s +S d +S t ),

[0069] S200: Determining a theoretical overload value of the aircraft based on the vertical landing energy and / or the rolling landing energy of the aircraft;

[0070] S300: Calculating the buffering stroke of the main buffer based on the overload theoretical 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 wheels 23, the rocker arm 21 and the main buffer 22 on the main landing gear 2, so the buffer stroke of the main buffer 22 is calculated by the following formula:

[0072]

[0073] Where m d l represents the equivalent mass of the main landing gear 2, V Z represents the landing sinking speed of the aircraft, G represents the landing equivalent weight of the main landing gear 2 (i.e. m dl *g), L represents the lift-to-weight ratio of the aircraft during landing, which is a coefficient not greater than 1 multiplied by G (usually 0.67 or 1), N is the landing overload of the main landing gear 2, S s n s NG represents the energy absorbed by the main buffer 22 (S s is the main buffer stroke, n s is the main buffer efficiency coefficient), S d n d NG represents the energy absorbed by the elastic damping rod 213 of the rocker arm 21 (S d is the elastic damping rod stroke of the rocker arm, n d is the efficiency coefficient of the elastic shock absorber rod of the rocker arm), S t n t NG represents the energy absorbed by the main wheel 23 (S t is the compression stroke of the main wheel, n t is the efficiency coefficient of the main engine wheel);

[0074] In the above formula, the stroke of the elastic shock-absorbing rod of the rocker arm 21 changes linearly, satisfying Hooke's law F=kx, which can be verified by finite element simulation, and the efficiency coefficient is 0.45; the stroke of the main wheel 23 can be obtained by querying the tire static pressure curve, and the efficiency coefficient is 0.47; the efficiency coefficient of the main buffer 22 is between 0.65 and 0.80. Combined with the structure of the main buffer 22 and design experience, the initial value can be 0.75, and then iterated in the subsequent design process. Therefore, in the above formula, there are only overload N and main buffer stroke S s Undetermined, and N is the theoretical overload value determined in step S200, and the main buffer stroke S is calculated by the above formula s ;

[0075] S400: performing a drop shock simulation based on the buffer stroke of the main buffer to at least obtain an overload simulation value and a stroke simulation value of the main landing gear;

[0076] Based on the stroke S of the main buffer 22 s Performing a drop shock simulation to obtain an overall overload simulation value of the main landing gear 2, an overall stroke simulation value of the main landing gear, and an overall simulation efficiency coefficient of the main landing gear;

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

[0078] Specifically, taking the difference threshold as 30% of the landing kinetic energy and potential energy of the aircraft and the ratio threshold as 70% as an example, the simulated absorption energy of the main landing gear 2 as a whole is calculated by the overload simulation value of the main landing gear 2 as a whole, the stroke simulation value of the main landing gear 2 as a whole, the simulated efficiency coefficient of the main landing gear 2 as a whole and the equivalent mass of the main landing gear 2, and compared with the sum of the landing kinetic energy and potential energy of the aircraft. If the difference between the sum of the landing kinetic energy and potential energy of the aircraft and the simulated absorption energy is greater than or equal to 30% of the sum of the landing kinetic energy and potential energy of the aircraft, the simulated absorption energy is calculated. Or if the ratio of the simulated absorbed energy to the sum of the landing kinetic energy and potential energy of the aircraft is less than or equal to 70% of the sum of the landing kinetic energy and potential energy of the aircraft, the overload theoretical value is adjusted and the overload simulation value, the stroke simulation value and the simulation efficiency coefficient are iterated until the difference between the calculated simulated absorbed energy of the main landing gear 2 as a whole and the sum of the landing kinetic energy and potential energy of the aircraft is less than 30% of the sum of the landing kinetic energy and potential energy of the aircraft or the ratio of the simulated absorbed energy of the main landing gear 2 as a whole to the sum of the landing kinetic energy and potential energy of the aircraft is greater than 70%, which meets the design requirements.

[0079] The main landing gear absorbs the landing energy through three-level buffering, which increases the utilization rate of materials and reduces the landing overload. Its three-level buffering structure is the main wheel, the elastic shock absorber rod of the rocker arm and the main buffer. In the design process, the kinetic energy of the aircraft landing is first determined to be Potential energy is (GL)*(S s +S d +S t ), then according to the law of conservation of energy we can get:

[0080]

[0081] The present invention further provides an aircraft, which has a fuselage 3, and the bottom of the fuselage 3 is provided with the above-mentioned landing gear 100. The specific structure of the landing gear 100 refers to the above-mentioned embodiment. Since the present aircraft adopts all the technical solutions of all the above-mentioned embodiments, it has at least all the beneficial effects brought by the technical solutions of the above-mentioned embodiments, which will not be described one by one here.

[0082] The above descriptions are only optional embodiments of the present invention, and are not intended to limit the patent scope of the present invention. Any equivalent structural transformations made using the contents of the present invention's specification and drawings, or direct / indirect applications in other related technical fields are included in the patent protection scope of the present invention.

Claims

1. A landing gear, characterized in that: The landing gear is arranged at the bottom of the aircraft body, and the landing gear comprises: A front landing gear, the front landing gear is arranged at the front side of the bottom of the fuselage, the front landing gear comprises a front wheel and a front buffer strut, one end of the front buffer strut is connected to the fuselage and arranged inside the fuselage, and the other end passes through the outside of the fuselage and is elastically connected to the front wheel; Two main landing gears, the two main landing gears are respectively arranged on opposite sides of the bottom of the fuselage, each of the main landing gears includes a rocker arm, a main buffer and a main wheel, the two ends of the rocker arm are respectively a hinged end and a free end, the hinged end is hinged to the inner side 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, and the position on the rocker arm located between the hinged end and the free end is elastically connected to the fuselage through the main buffer.

2. The landing gear according to claim 1, characterized in that The main buffer includes a main sleeve and a first buffer rod. The main sleeve is hinged to the inner side of the body. The first end of the first buffer rod is sleeved in the main sleeve. The second end of the first buffer rod extends out of the main sleeve and is hinged to a position on the rocker arm close to the hinged end. The first buffer rod can be extended and retracted along the extension direction of the main sleeve.

3. The landing gear according to claim 2, characterized in that: The two main landing gears are symmetrically arranged about the symmetry plane of the fuselage, and each rocker arm is inclined from the hinged end to the free end toward the tail side of the fuselage, and is inclined from the hinged end to the free end in a direction away from the other rocker arm, and the hinged end is hinged to the inner side of the fuselage through a hinged rod, and the hinged rod is perpendicular to the symmetry plane of the fuselage.

4. The landing gear according to claim 3, characterized in that The rocker arm includes an elastic shock-absorbing rod and a rocker arm joint that are connected to each other. The end of the rocker arm joint away from the elastic shock-absorbing rod forms the hinged end, and the end of the rocker arm joint close to the elastic shock-absorbing rod is provided with a hinged seat hinged to the second end of the first buffer rod; the hinged end is provided with a hinge ring, the hinge rod is rotatably inserted into the hinge ring, and the end of the elastic shock-absorbing rod away from the rocker arm joint forms the free end.

5. The landing gear according to any one of claims 1 to 4, characterized in that The cross section of the rocker arm is elliptical, and a long axis side of the cross section of the rocker arm faces the nose side of the machine body.

6. The landing gear according to any one of claims 1 to 4, characterized in that Each main wheel is inclined from top to bottom toward another main wheel, and the angle between the axis of each main wheel and the horizontal plane is 1° to 1.5°; or, the front buffer pillar is inclined from top to bottom toward the nose side of the machine body, and the angle between the axis of the front buffer pillar and the vertical line is 7° to 7.5°.

7. A landing gear according to any one of claims 1 to 4, characterized in that The front wheel and / or each main wheel outer cover is provided with a fairing.

8. The landing gear according to any one of claims 1 to 4, characterized in that The front landing gear and / or the main landing gear are provided with a wheel-mounted switch, and the wheel-mounted switch is configured to send an air-to-ground signal.

9. A method for designing a landing gear, characterized in that: Applied to the landing gear according to any one of claims 1 to 8, the design method of the landing gear comprises the following steps: configuring the energy absorption of the main landing gear based on the landing kinetic energy and potential energy of the aircraft, wherein the energy absorption of the main landing gear includes the energy absorption of the main wheels, the energy absorption of the rocker arms and the energy absorption of the main buffers; determining a theoretical overload value of the aircraft based on the vertical landing energy and / or the rolling landing energy of the aircraft; Calculating the buffering stroke of the main buffer based on the overload theoretical value and the law of conservation of energy; Performing a drop shock simulation based on the buffer stroke of the main buffer to at least obtain an overload simulation value and a stroke simulation value of the main landing gear; At least the overload theoretical value is adjusted, and the overload simulation value and the stroke simulation value are iteratively calculated repeatedly until the ratio of the absorbed energy of the main landing gear obtained by simulation calculation to the sum of the landing kinetic energy and potential energy of the aircraft is greater than a ratio threshold, and / or the difference between the sum of the landing kinetic energy and potential energy of the aircraft and the absorbed energy of the main landing gear is less than a difference threshold.

10. An aircraft, characterized in that: The aircraft comprises a fuselage, and a landing gear according to any one of claims 1 to 8 is arranged at the bottom of the fuselage.

Images