CN207450213U - The rack construction and aircraft of aircraft - Google Patents

Abstract

The utility model discloses the rack construction and aircraft of a kind of aircraft, which includes：Frame main body and the upper cover being connected with the frame main body；Wherein, the both sides of the upper cover are provided with side thermal component.The utility model can provide better heat dissipation performance for aircraft, enhance the performance and flight safety of aircraft, and simple in structure, easily fabricated, and manufacture is at low cost.

Description

Aircraft frame structure and aircraft

Technical Field

The utility model relates to an aircraft technical field, concretely relates to heat dissipation technology of aircraft.

Background

In recent years, with the increasing development and maturity of aircraft technology, the application field of the aircraft is wider and wider, and along with this, users put forward higher and higher requirements on the performance of the aircraft.

For example, in the field of aircraft design, heat dissipation is closely related to the performance of the aircraft, and if heat generating components are not dissipated in time, the performance and flight safety of the aircraft are affected, for example, chips or parts are burned out.

However, the existing aircraft still has the problem that a good heat dissipation function cannot be realized, so that the components generate heat, and therefore, the flying of the aircraft has potential safety hazards. Moreover, heat dissipation is performed by a heat dissipation device such as a fan, so that on one hand, the whole weight is increased and electric energy is consumed, and on the other hand, the heat dissipation effect is not ideal.

In view of the above, there is a strong need in the art for a rack design with better heat dissipation.

Disclosure of Invention

The utility model aims at providing an aircraft rack construction and aircraft can provide better heat dispersion for the aircraft, strengthens the performance and the flight safety of aircraft, and simple structure, easily makes low in manufacturing cost.

In a first aspect of the present invention, there is provided a frame structure of an aircraft, comprising: the rack comprises a rack main body and an upper cover connected with the rack main body; wherein,

and side radiating parts are arranged on two sides of the upper cover, so that air flow can enter the side radiating parts when the aircraft flies.

Preferably, the side heat-radiating member includes: the side air inlet is matched with the shape of the side surface of the upper cover, and the blade plate is arranged inside the side air inlet.

Preferably, the blades are arranged obliquely, so that airflow can enter the side air inlet when the aircraft flies.

Preferably, the blades are arranged along the advancing direction of the aircraft in an inclined direction with a high front and a low rear, wherein the head of the aircraft body is arranged at the front, and the tail of the aircraft body is arranged at the rear.

Preferably, the angle of the blades with respect to the horizontal decreases gradually from the head of the fuselage to the tail.

Preferably, the side heat dissipation part further comprises a reinforcing rod, two ends of the reinforcing rod are fixedly arranged in the side air inlet, and the reinforcing rod is connected with each blade.

Preferably, the heat sink further comprises an intermediate heat sink member disposed at an intermediate position of the upper cover.

Preferably, the middle heat sink member includes a middle air inlet and a louver, wherein,

the blade plate is obliquely arranged in the middle air inlet, so that air flow can enter the middle air inlet when the aircraft flies.

Preferably, the middle heat dissipation part further comprises a reinforcing rod, two ends of the reinforcing rod are fixedly arranged in the middle air inlet, and the reinforcing rod is connected with each blade.

Preferably, the frame structure of the aircraft further comprises a front air inlet arranged on the upper cover, and the front air inlet is positioned at the front end of the middle heat dissipation component.

In a second aspect of the invention, an aircraft is provided, comprising a frame structure as described above.

Compared with the prior art, the embodiment of the utility model, following difference and effect have at least:

1. providing better heat dissipation performance.

2. The performance and flight safety of the aircraft are enhanced.

3. Simple structure, easy manufacture and low manufacturing cost.

It is understood that within the scope of the present invention, the above-mentioned technical features of the present invention and those specifically described below (e.g. in the examples) can be combined with each other to constitute new or preferred technical solutions. Not to be reiterated herein, but to the extent of space.

Drawings

FIG. 1 is a schematic view of an aircraft flight attitude and configuration according to an embodiment of the present invention;

fig. 2 is a schematic view of a takeoff and landing static attitude according to an embodiment of the present invention;

fig. 3 is a schematic view of the flight attitude of an aircraft according to an embodiment of the invention;

fig. 4 is a schematic illustration of a take-off and landing static attitude of an aircraft according to an embodiment of the invention;

fig. 5 is a schematic view of the main components of an aircraft according to an embodiment of the invention;

fig. 6 is a schematic view of the main components of an aircraft according to an embodiment of the invention;

fig. 7 is an assembly schematic of an aircraft according to an embodiment of the invention;

FIG. 8 is a schematic view of a prior art vertical heat dissipation hole structure;

fig. 9 is an angle of incidence fairing of an aircraft according to an embodiment of the invention;

FIG. 10 is a schematic view of a prior art air flow and vertical louvers;

fig. 11 is a schematic view of an aircraft windward angle streamline louvers and air flow direction according to an embodiment of the present invention;

fig. 12 is an exploded schematic view of a tail antenna of an aircraft according to an embodiment of the invention;

fig. 13 is a schematic view of a tail antenna of an aircraft in a folded state according to an embodiment of the invention;

fig. 14 is a schematic illustration of a tail antenna reposition for an aircraft according to an embodiment of the present invention;

fig. 15 is a schematic view of a tail antenna structure of an aircraft according to an embodiment of the invention;

fig. 16 is a schematic view of a tail antenna component of an aircraft according to an embodiment of the invention;

fig. 17 is a schematic view of a folded condition of a tail antenna of an aircraft according to an embodiment of the invention;

fig. 18 is a schematic view of a tail antenna folded state of an aircraft according to an embodiment of the invention;

fig. 19 is a schematic view of a tail antenna folded state of an aircraft according to an embodiment of the invention;

fig. 20 is a schematic view of a tail antenna of an aircraft in a pre-impact state according to an embodiment of the invention;

fig. 21 is a schematic view of a rear wing antenna of an aircraft after being impacted according to an embodiment of the invention;

figure 22 is a schematic view of the main components of an aircraft according to one embodiment of the invention;

fig. 23 is a schematic view of the installation of a lens of an aircraft according to an embodiment of the invention;

fig. 24 is a schematic position diagram of a lens of an aircraft according to an embodiment of the invention;

fig. 25 is a schematic view of the main components of a lens of an aircraft according to an embodiment of the invention.

The same reference numbers will be used throughout the drawings to refer to the same or like elements or structures, wherein:

1: fuselage body

2: machine arm

3: propeller assembly

4: propeller protection assembly

5: blade

6: motor cover

7: connecting arm

8: connecting frame

9: a bent part

101: intermediate heat dissipation member

102: side heat dissipation component

103: front air inlet

104: reinforcing bar

105: leaf plate

106: side air inlet

107: middle air inlet

11: sheet antenna

12: antenna support

13: antenna slot

14: multi-angle clamping groove

15: first pin

16: second pin

17: rotating support shaft

18: folding angle limiting groove

19: mounting groove for rotating support shaft

21: main body clamping and fixing groove

22: groove assembly

51: image plate

52: main control board

53: battery with a battery cell

54: video camera

61: rack

611: frame main body

62: upper cover

63: rotary positioning shaft

64: support piece

Detailed Description

In the following description, numerous technical details are set forth in order to provide a better understanding of the present application. However, it will be understood by those skilled in the art that the technical solutions claimed in the present application can be implemented without these technical details and with various changes and modifications based on the following embodiments. Furthermore, the technical features mentioned in the embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

Term(s) for

As used herein, the term "aircraft" refers to flying equipment such as drones.

As used herein, the term "motor cover" refers to a component for housing the drive arrangement of an aircraft propeller.

As used herein, the term "bow" refers to a shape having a bent portion in the middle that is concave inward, and two ends that are bent toward the same side of the bent portion, in a "bow" shape.

As used herein, the term "air intake" refers to an opening in the fuselage of an aircraft for the introduction of an airflow.

As used herein, the term "louvers" refers to plate-like tabs disposed at the air intake opening in the direction of air flow entry.

As used herein, the term "multi-angle snap groove" refers to a groove structure capable of being snapped with a mounting groove and/or a groove for fixing at least two or more different angles.

As used herein, the term "rotating support shaft" refers to a support shaft for being rotatably fixed.

As used herein, the term "folding angle limiting groove" refers to a groove that is matched with the multi-angle catching groove and can be fixed with the multi-angle catching groove at least two or more different angles.

As used herein, the term "rotational positioning shaft" refers to a positioning shaft that is rotatably fixed.

As used herein, the term "damping gear" refers to a gear that achieves a speed reduction effect by increasing frictional resistance.

1. Aircraft frame structure

The present inventors have made extensive and intensive studies and have proposed a new frame structure. Referring to the drawings, the frame structure mainly comprises a frame main body, a horn, a propeller assembly and a propeller protection assembly. Wherein, the frame main body is connected with one end of the machine arm; the propeller assembly is connected with the other end of the horn and the propeller protection assembly respectively. Also, the propeller protection assembly is used not only to protect the propeller assembly, but also to serve as a landing gear.

More specifically, the propeller assembly further includes a motor cover for accommodating the motor, the motor cover being connected to the other end of the horn and the propeller protection assembly, respectively. The height of the motor cover is set so that the motor cover acts as an undercarriage to support the fuselage when the aircraft is landing on the reverse side.

More specifically, the propeller protection assembly includes a plurality of link arms and a plurality of link brackets, wherein one end of each link arm is connected to the motor cover and the other end is connected to the link brackets. The propeller assembly comprises a blade, and one end of each connecting arm, which is connected with the connecting frame, extends downwards and is connected with the connecting frame, so that an accommodating space capable of protecting the blade is formed.

According to the rack structure, not only a protection space for the propeller assembly, especially the propeller blades, is formed, but also a supporting structure convenient for lifting of the front side and the back side is formed, and furthermore, the structure is simple and convenient to manufacture.

The following detailed description will be made with further reference to the accompanying drawings.

As shown in fig. 1, 6 and 12, the airframe structure of the aircraft of the present embodiment includes: the device comprises a machine body 1, a machine arm 2, a propeller component 3 with blades 5 arranged at the bottom and a propeller protection component 4, wherein one end of the machine arm 2 is connected with the machine body 1, and the other end of the machine arm is connected with the propeller component 3; and, the propeller protection assembly 4 is used to protect the blades 5; and also for use as landing gear, as shown in figure 2.

More specifically, as shown in the figure, the propeller protection assembly 4 forms a three-dimensional protection space on the outer side of the blade 5, and can block objects from impacting the blade from the outer side of the blade 5, and meanwhile, the lower part of the propeller protection assembly 4 can stably support the aircraft, so that the landing gear function is realized. Therefore, the structure is beneficial to reducing the whole volume and the whole weight, and has simple structure, easy manufacture and lower whole manufacturing cost.

In this embodiment, the frame structure includes 4 propeller assemblies 3, which are respectively disposed on both sides of the head and both sides of the tail of the body 1, and are symmetrically disposed, in other words, the 4 propeller assemblies 3 are symmetrically disposed on the front end and the rear end of both sides of the body 1. The embodiment is not limited thereto, however, and in other embodiments the frame structure may comprise a different number of propeller assemblies 3, for example, 3, or 4 propeller assemblies may be provided on each side of the fuselage, etc.

In this embodiment, the propeller assembly 3 further comprises a motor cover 6 disposed on the upper portion of the blades 5, and the height of the motor cover 6 is greater than that of the fuselage 1, so that when the aircraft is lifted and lowered on the reverse side, the motor cover 6 serves as a landing gear capable of supporting the fuselage (as shown in fig. 3 and 4). Further, when the aircraft flies, the attitude of the aircraft cannot be adjusted in time before landing on the ground due to various reasons such as descent speed, wind speed, interference and the like, and in this case, the motor cover 6 of the aircraft in this embodiment can provide a reverse landing gear for the aircraft, can stably support the aircraft in an upside-down state to land on the ground, and can also support the aircraft in this state to take off. The structure supporting the forward and reverse double-sided landing enables the aircraft to have more flexible flight capability, can provide better protection for the landing of the aircraft, and is simple in structure, easy to manufacture and low in overall manufacturing cost.

In this embodiment, the propeller protection assembly 4 includes 8 connecting arms 7 and 2 connecting frames 8, the connecting frames 8 are respectively located at two sides of the body 1, one end of each connecting arm 7 is connected to one motor cover 6, and the other end is connected to the connecting frame 8. More specifically, as shown in the figures, each motor cover 6 is connected to 2 respective connecting arms 7, and each connecting bracket 8 connects all the connecting arms 7 located on the same side of the body 1, so as to constitute a protective structure for all the blades 5 on one side with all the connecting arms 7 on one side of the body 1. It can be seen that each connecting arm 7 is connected at one end to the motor cover 6 and at the other end to the connecting bracket 8, as viewed from either side of the fuselage 1, and the connecting arms 7 extend downwardly at the end adjacent to the connecting bracket 8, whereby all the connecting arms 7 on each side of the fuselage 1 and the connecting brackets 8 on that side constitute a protective structure for all the blades 5, which has a high degree of fastness and impact resistance. Under the condition, the blades of the aircraft can be well protected in the flying process, so that the impact of external objects on the blades is prevented, and the flying safety is improved. Meanwhile, the protection of the blades during the carrying process of the aircraft is also ensured. This is advantageous in that it provides the necessary blade protection while reducing the volume and weight, and has a simple structure and a low manufacturing cost.

As described above, in the present embodiment, the end of each connecting arm 7 connected to the connecting frame 8 extends downward to be connected to the connecting frame 8, thereby constituting an accommodating space capable of protecting the blade 5. Specifically speaking, every linking arm 7 is the shape of buckling, can be the right angle bending shape, but the utility model discloses be not limited to this, in other embodiments, every linking arm 7 also can be 120 and buckle, perhaps 130 and buckle, and the angle of buckling does not specifically restrict as long as every linking arm 7 can be after being connected with motor cover 6 and link 8, and the formation can protect the accommodation space of paddle 5 can.

In the present embodiment, the connecting frame 8 has a bent shape as shown in the figure, and has an arch structure. Specifically, the two ends of the connecting frame 8 are bent toward the fuselage 1, which is beneficial for protecting the blade 5, so that the front of the blade 5 is within the protection range of the connecting frame 8, and the blade 5 is prevented from being impacted by external objects from the front. Still further, as shown in the figure, the middle part of the connecting frame 8 is provided with a bent part 9, and the bent part 9 is recessed towards the side of the body 1, so that, when viewed from any side of the body 1, all the connecting arms 7 connected with 1 motor cover 6 are located at one side of the bent part 9, and all the connecting arms 7 connected with another motor cover 6 are located at the other side of the bent part 9, which is beneficial to enhancing the strength and impact resistance of the whole propeller protection assembly 4. The protection capability of the blade 5 is improved. However, the present invention is not limited thereto, and in other embodiments, the connecting frame 8 may have other shapes, for example, the connecting frame 8 may have a W shape, an S shape, or a linear shape, or the connecting frame may have two ends bent toward the body, and so on. The connecting frame 8 may have a plurality of bent portions 9, and the connecting frame 8 may be divided into a plurality of stages corresponding to the case where the body 1 has a plurality of motor covers 6 on each side. In this case, the connecting arms connected to the same motor cover 6 may be connected to the same segment of the connecting bracket 8.

In the present exemplary embodiment, each motor cover 6 is connected to 2 respective connecting arms 7. The angle between the 2 connecting arms connected to each motor casing 6 is anywhere between 30 deg. and 150 deg., for example 75-90 deg., most preferably 75 deg., in which case the force resolution is best. However, the present invention is not limited thereto, and in other embodiments, the motor cover 6 may be connected with 1, 3 or more connecting arms 7, respectively, and the included angle between adjacent connecting arms 7 may be any angle between 15-45 °, preferably 30-35 °, and most preferably 30 °, in which case the force resolution is the best. The advantage of connecting 1 linking arm 7 per motor casing 6 is that simple structure reduces weight. The advantage of having 3 or more connecting arms 7 per motor cover 6 is that it is stronger and more impact resistant.

In the embodiment, the body 1, the horn 2 and the propeller protection assembly 4 are integrally formed, so that the manufacturing is simple and the manufacturing cost is low. However, the present invention is not limited thereto, and in other embodiments, each component may be connected by means of a snap, a screw hole, a screw, or the like.

The main advantages of the above-described airframe structure of an aircraft include:

1. can adapt to different flight attitudes, provide multiple rising and falling support attitudes, and realize rising and falling of the front and back sides.

2. The propeller assembly, especially the blade structure, is effectively protected, and the anti-impact performance and the safety are high.

3. Small volume and light weight.

4. Simple structure, easy manufacture and low manufacturing cost.

In conclusion, the frame structure of the aircraft has a very wide application prospect in the field of aircraft.

2. High efficiency louver design

The present inventors have made extensive and intensive studies and as a result, have proposed a novel heat dissipation hole structure in which side heat dissipation members having a shape matching with each other are provided on the upper surface of a body, that is, both sides of an upper cover, a middle heat dissipation member is provided in the middle, and louvers are sequentially arranged in an air inlet of the heat dissipation member, and the louvers are formed such that the louver direction is high in the front and low in the rear, and an angle with respect to the horizontal direction gradually decreases from the head of the body toward the tail wing, and are fixed by a reinforcing bar. Furthermore, the rack structure also comprises an air inlet arranged on the upper surface and positioned at the front end of the middle heat dissipation part. According to the structure, the angle formed by the improved air inlet and the blade plate structure effectively utilizes the air flow to fully take out the heat of the air flow, so that the service life of the aircraft is prolonged, and the risk that the chip or parts are burnt out due to overhigh temperature of the aircraft body is avoided. Through above-mentioned heat dissipation part, the even effectual heat dissipation of aircraft has effectively been realized.

This is explained in further detail below.

As shown in fig. 11 and 12, in the present embodiment, the airframe structure of the aircraft includes: fuselage 1, horn 2, screw assembly 3, the one end of horn 2 is connected with fuselage 1, and the other end is connected with screw assembly 3 to the upper surface of fuselage 1, that is, the both sides of upper cover 62 are provided with side heat dissipation part 102. In the present embodiment, the surface of the main body 1 is provided with a side heat-radiating member 102, which contributes to effective heat radiation of heat-generating components on both sides in the main body.

In the present embodiment, the side heat sink 102 includes a long strip-shaped side air inlet 106 matching with the shape of the side body of the fuselage 1, and a louver 105, wherein, as shown in fig. 9 and 11, the louver 105 is obliquely disposed in the side air inlet 106 in sequence, and the oblique direction of the louver 105 is set so that the airflow can enter the side air inlet 106 when the aircraft is in flight. In this case, the side air inlets 106 can be matched with the shapes of the two sides of the fuselage 1 to the greatest extent, so as to provide more effective air inlets, and the blades 105 are arranged in sequence and inclined, so that when the aircraft flies, the direction of the air entering is consistent with the inclined direction of the blades 105, thereby not only reducing the air resistance encountered by the aircraft in the flying process, but also leading more wind in the windward direction, and improving the heat dissipation efficiency. In contrast, as shown in fig. 8 and 10, the conventional heat dissipation holes in the prior art are vertical heat dissipation holes, which cannot effectively release body heat, and the aircraft flies at high temperature for a long time, which results in a short service life, is easy to burn out chips or parts, has a potential safety hazard, and is high in maintenance cost.

Further, in the present embodiment, the blades 105 are disposed in an inclined direction with a high front and a low rear along the forward direction of the aircraft, wherein the head of the fuselage 1 is located at the front and the tail of the fuselage 1 is located at the rear.

In the present embodiment, the angle of the louver 105 with respect to the horizontal direction decreases gradually from the head of the fuselage 1 toward the tail. In other words, the louvers 105 are relatively "vertical" near the head of the fuselage 1, and gradually "lie flat" in a rearward direction of the fuselage 1, which facilitates airflow entering the side intake vents 106 during flight of the aircraft and provides a better reduction in air drag. However, the present invention is not limited thereto, and in other embodiments, the included angle of the louver 105 relative to the horizontal direction may also be a constant value, that is, each louver 105 is parallel to each other, or the louver 105 may also adopt different included angles at different positions of the side air inlet 106.

In this embodiment, the side heat sink member 102 further includes a reinforcing rod 104 fixedly disposed in the side air inlet 106 at both ends, and the reinforcing rod 104 penetrates each louver 105. The louver 105 can be more firmly disposed in the side air intake opening 106 by the reinforcing rod 104, and can withstand a large airflow pressure. In this embodiment, 1 reinforcing rod 104 is provided in the side air inlet 106, but the present invention is not limited thereto, and in other embodiments, 2 or more reinforcing rods 104 may be provided.

In this embodiment, the frame structure of the aircraft further includes an intermediate heat sink 101 disposed on the upper surface of the aircraft, i.e., in the middle of the upper cover 62, the intermediate heat sink 101 includes an intermediate air inlet 107 and louvers 105, the louvers 105 are sequentially arranged and obliquely disposed in the intermediate air inlet 107, and the oblique direction of the louvers 105 is set to enable the air flow to enter the intermediate air inlet 107 when the aircraft is in flight. Therefore, the air flow encountered in the flight of the aircraft can be effectively utilized to dissipate heat of the heat generating components arranged in the middle of the interior of the aircraft. The heat dissipation effect of the heat generating components at various positions of the whole aircraft is more perfect by matching with the side heat dissipation components 102 at the two sides.

In this embodiment, the middle heat sink 101 further includes a stiffener 104 fixed at both ends thereof in the middle air inlet 107, and the stiffener 104 penetrates each of the louvers 105. The louver 105 can be more firmly disposed in the middle intake opening 107 by the reinforcing rod 104, and can withstand a large airflow pressure. In this embodiment, 1 reinforcing rod 104 is disposed in the middle air inlet 107, but the present invention is not limited thereto, and in other embodiments, 2 or more reinforcing rods 104 may be disposed.

In this embodiment, the airframe structure of the aircraft further includes a forward air inlet 103 provided on the upper surface of the fuselage 1, and the forward air inlet 103 is located at the front end of the intermediate heat sink 101, i.e., near the head of the fuselage 1. The opening direction of the forward air inlet 103 is set so that the air flow can enter the forward air inlet 103 when the aircraft is flying. More specifically, the partial covering panel in the upper surface of the fuselage 1, in the middle of the fuselage 1 near the head, extends slightly downwards, so that it creates an opening with the upper surface, which opening is directed exactly in line with the direction of advance of the aircraft, so as to allow the entry of the air flow. The provision of the forward air inlet 103, together with the side heat-dissipating component 102 and the middle heat-dissipating component 101, provides integrated, more efficient heat dissipation, so that heat-generating components in the aircraft fuselage 1 are uniformly dissipated.

The main advantages of the heat dissipation hole structure include:

1. providing better heat dissipation performance.

2. The performance and flight safety of the aircraft are enhanced.

3. Simple structure, easy manufacture and low manufacturing cost.

To sum up, the utility model provides a louvre structure has very wide application prospect in the aircraft field.

3. Angle-adjustable antenna

The present inventors have made extensive and intensive studies and as a result, have proposed a novel angle-adjustable antenna structure. The antenna support and the groove component which are matched with each other are arranged at the tail wing part of the aircraft, the pin with the plurality of bending parts at the inner side edge is arranged on the antenna support, the pin can be inserted into and fixed in the groove component at the tail wing part at different angles, furthermore, another fixed pivot is provided between the antenna support and the groove component through the rotating support shaft and the rotating support shaft mounting groove, the antenna component can be flexibly adjusted to different postures according to requirements, namely, the antenna can be adjusted to different angles, the angle can be stably adjusted, the change condition of communication signals of the aircraft in the flight process can be coped with, and the performance of transmitting and receiving the signals of the antenna is obviously improved.

As described in further detail below.

As shown in fig. 12, 13, 14, 15, and 16, the antenna structure in this embodiment includes: an antenna 11, an antenna mount 12, and a notch assembly 22, wherein the antenna 11 is fixed to the antenna mount 12, and the antenna mount 12 is angularly adjustably disposed at a tail position of the aircraft by the notch assembly 22, in other words, the orientation of the antenna 11 is also angularly adjustable, as shown in fig. 17, 18, and 19.

More specifically, in the present embodiment, two multi-angle engaging grooves 14 are disposed below the antenna bracket 12, a rotating bracket shaft 17 is disposed on the multi-angle engaging grooves 14, and the groove component 22 includes a rotating bracket shaft mounting groove 19 whose position and size are matched with the rotating bracket shaft 17.

Specifically, in the present embodiment, two multi-angle engaging grooves 14 are disposed below the antenna bracket 12, a rotating bracket shaft 17 is disposed on the multi-angle engaging grooves 14, and the groove component 22 includes a rotating bracket shaft mounting groove 19 whose position and size are matched with the rotating bracket shaft 17. The antenna support 12 is fixed but rotatable by its swivel support shaft 17 being inserted into a swivel support shaft mounting groove 19 on the fuselage, in which case, if an external force is applied, the antenna support 12 can be turned over a certain angle around the swivel support shaft 17 to avoid breaking directly by the action of the external force, while at the same time the swivel support shaft 17 can still remain embedded in the swivel support shaft mounting groove 19, in other words, the antenna support 12 is releasably disposed in the swivel support shaft mounting groove 19 at the tail position of the aircraft by the swivel support shaft 17.

In this embodiment, the multi-angle engaging groove 14 is provided with a first pin 15 and a second pin 16, and the fuselage of the aircraft is provided with a main body engaging and fixing groove 21 and a folding angle limiting groove 18, which are matched with the positions and sizes of the first pin 15 and the second pin 16. So that the first pin 15 and the second pin 16 can be inserted into the body-catching and fixing groove 21 and the folding-angle restricting groove 18, respectively.

In this embodiment, the first pin 15 and the second pin 16 form at least 2 grooves for engaging at different angles. Specifically, the inner edges of the first pin 15 and the second pin 16 are provided with 2 bending portions, so that the first pin 15 and the second pin 16 can be engaged with the main body engaging and fixing groove 21 and the folding angle limiting groove 18 at different angles. In other words, since the inner side edges of the first pin 15 and the second pin 16 are provided with 2 bending portions, when the first pin 15 and the second pin 16 are engaged with the main body engaging and fixing groove 21 and the folding angle limiting groove 18 at different angles, the postures of the antenna support 12 are different, and thus the direction of the antenna 11 is different, so that the antenna can flexibly adapt to flying states at different angles, and the performance of the antenna is better. Although the first pin 15 and the second pin 16 are provided with 2 bent portions at the inner edges thereof in the present embodiment, the present invention is not limited thereto, and in other embodiments, 3 or more bent portions may be provided at the inner edges of the first pin 15 and the second pin 16, so that the antenna 11 can be adjusted to more different angles.

In the present embodiment, the antenna 11 is a patch antenna, but the present invention is not limited thereto, and in other embodiments, the antenna 11 may be other types of antennas, for example, a wire antenna, a helical antenna, or the like.

In the present embodiment, the antenna slot 13 is disposed on the antenna bracket 12 for accommodating the antenna 11. The tail location is the position of the tail of the aircraft.

The main advantages of the above antenna structure include:

1. the performance of the antenna for transmitting and receiving signals is remarkably improved.

2. Simple structure, easy manufacture and low manufacturing cost.

To sum up, the utility model provides an antenna structure is showing the ability that has improved antenna emission and received signal to simple structure, low in manufacturing cost consequently has very wide application prospect in the aircraft field.

4. Anti-breaking antenna structure

The present inventors have made extensive and intensive studies and as a result, have proposed a novel break-resistant antenna structure. Specifically speaking, the antenna support is arranged on the aircraft in a partially releasable manner through a groove component arranged at the tail position of the aircraft, for example, through a rotating support shaft and a rotating support shaft mounting groove arranged at the tail position, when the antenna support encounters external force impact or impact, the antenna support can be turned over along the rotating support shaft, so that a first pin and a second pin inserted into a folding angle limiting groove and a main body clamping fixing groove can be released, namely, the multi-angle clamping groove and the groove component can be released, but the rotating support shaft is still rotatable and connected with the rotating support shaft mounting groove, thereby avoiding the direct action of the external force, and achieving the technical effect of resisting breakage, wherein the technical effect is that the partial loosening is still connected on the whole.

As described in further detail below.

As shown in fig. 13-16, in this embodiment, the aircraft also includes an anti-crash antenna structure. The antenna structure includes: an antenna 11, an antenna mount 12, and a notch assembly 22 disposed on the fuselage of the aircraft, wherein the antenna 11 is secured to the antenna mount 12, and the antenna mount 12 is releasably disposed in a tail position of the aircraft by the notch assembly 22, as shown in fig. 20 and 21.

Specifically, in the present embodiment, two multi-angle engaging grooves 14 are disposed below the antenna bracket 12, a rotating bracket shaft 17 is disposed on the multi-angle engaging grooves 14, and the groove component 22 includes a rotating bracket shaft mounting groove 19 whose position and size are matched with the rotating bracket shaft 17. The antenna support 12 is fixed but rotatable by its swivel support shaft 17 being inserted into a swivel support shaft mounting groove 19 on the fuselage, in which case, if an external force is applied, the antenna support 12 can be turned over a certain angle around the swivel support shaft 17 to avoid breaking directly by the action of the external force, while at the same time the swivel support shaft 17 can still remain embedded in the swivel support shaft mounting groove 19, in other words, the antenna support 12 is releasably disposed in the swivel support shaft mounting groove 19 at the tail position of the aircraft by the swivel support shaft 17.

In this embodiment, the multi-angle engaging groove 14 is provided with a first pin 15 and a second pin 16, and the fuselage of the aircraft is provided with a main body engaging and fixing groove 21 and a folding angle limiting groove 18, which are matched with the positions and sizes of the first pin 15 and the second pin 16. The first pin 15 and the second pin 16 can be inserted into the main body clamping and fixing groove 21 and the folding angle limiting groove 18 respectively to play a further fixing role.

Further, in the present embodiment, the first pin 15 and the second pin 16 form at least 2 grooves engaging at different angles. Specifically, the inner edges of the first pin 15 and the second pin 16 are provided with 2 bending portions, so that the first pin 15 and the second pin 16 can be engaged with the main body engaging and fixing groove 21 and the folding angle limiting groove 18 at different angles. In other words, since the inner side edges of the first pin 15 and the second pin 16 are provided with 2 bending portions, when the first pin 15 and the second pin 16 are engaged with the main body engaging and fixing groove 21 and the folding angle limiting groove 18 at different angles, the postures of the antenna support 12 are different, and thus the direction of the antenna 11 is different, so that the antenna can flexibly adapt to flying states at different angles, and the performance of the antenna is better. Although the first pin 15 and the second pin 16 are provided with 2 bent portions at the inner edges thereof in the present embodiment, the present invention is not limited thereto, and in other embodiments, 3 or more bent portions may be provided at the inner edges of the first pin 15 and the second pin 16, so that the antenna 11 can be adjusted to more different angles.

In the present embodiment, the antenna 11 is a patch antenna, but the present invention is not limited thereto, and in other embodiments, the antenna 11 may be other types of antennas, for example, a wire antenna, a helical antenna, or the like.

In the present embodiment, the antenna slot 13 is disposed on the antenna bracket 12 for accommodating the antenna 11. The tail location is the position of the tail of the aircraft.

The above-mentioned anti-break antenna structure includes the following main advantages:

1. has stronger impact resistance.

2. The maintenance is convenient, and the maintenance cost is low.

3. Simple structure, easy manufacture and low manufacturing cost.

To sum up, the utility model provides an antenna structure is showing the ability that has improved antenna emission and received signal to simple structure, low in manufacturing cost consequently has very wide application prospect in the aircraft field.

5. Circuit board stack

The present inventors have made extensive and intensive studies and as a result, have proposed a new stacked structure for fixing an image board, a main control board, and a battery in an aircraft, and a camera to a main frame body in a stacked manner, wherein the camera is disposed in front. The whole thickness of heap structure can further be reduced to this kind of setting, makes the fuselage can make more thinly on the whole, has reduced the holistic focus of fuselage to the focus is concentrated, improves flight stability and controllability.

As described in further detail below.

As shown in fig. 5, 6, 7, and 22, in the present embodiment, the stacked structure of the aircraft includes: a frame body 611, an image board 51, a main control board 52, and a battery 53, wherein the image board 51, the main control board 52, and the battery 53 are fixed in an up-down stacked manner on the fuselage 1 of the aircraft, i.e., the frame body 611.

Specifically, the image board 51 is a circuit board for image processing. The main control board 52 may also be referred to as a main circuit board, and is provided with connection points for connecting to a processor, a memory, and an external device. Typically, these components may be directly inserted into the socket or connected by wires.

By stacking up and down, the image board 51, the main control board 52 and the battery 53 are arranged in a concentrated manner, so that the aircraft has the advantages of reducing the volume and concentrating the center of gravity of the aircraft, and further improving the flight performance.

In the present embodiment, the image board 51, the main control board 52, and the battery 53 are fixed to the fuselage 1 of the aircraft, i.e., the rack body 611, by screws and screw holes in a vertically stacked manner. For example, screw holes may be provided at corresponding positions on the main control board 52 so that the image board 51 and the battery 53 may be fixed on the main control board 52 by screws. However, the present invention is not limited thereto, and in other embodiments, the image board 51, the main control board 52 and the battery 53 may be fixed on the frame main body 611 of the aircraft in an up-down stacking manner through a buckle or a plug and a socket.

In the present embodiment, the image board 51 and the main control board 52 are disposed on the upper side of the housing main body 611, and the battery 53 is disposed on the lower side of the housing main body 611. This way is favorable to providing a comparatively safe accommodation space for image board 51, main control panel 52 and battery 53, reduces the probability that receives the impact in the flight process, ensures flight safety. However, the present invention is not limited thereto, and the image board 51, the main control board 52 and the battery 53 may be sequentially stacked, or may be stacked in any order, and may be all stacked on the upper side of the frame main body 611, some may be stacked on the upper side of the frame main body 611, and some may be stacked on the lower side of the frame main body 611.

If the image board 51, the main control board 52, and the battery 53 are sequentially stacked, the battery 53 is located under the image board 51 and the main control board 52, which is advantageous in improving flight stability, and the battery 53 can be conveniently mounted and dismounted.

In the present embodiment, the stacked structure further includes a camera 54, wherein the camera 54 is disposed at the front side of the image board 51, the main control board 52, and the battery 53, that is, at the front side of the rack main body 611. In other words, the camera 54 is arranged on the front side of the fuselage 1 of the aircraft. The thickness of stack structure can further be reduced to this kind of setting, makes the fuselage can make more thinly on the whole, has reduced the holistic focus of fuselage, improves flight stability and controllability. More specifically, the camera 54 may be rotatably connected to the housing main body 611 by a screw, or may be connected thereto by a method such as a snap.

In the present embodiment, the camera 54 includes a connecting wire for electrically connecting with the image board 51 and transmitting information such as image signals. However, the present invention is not limited thereto, and in other embodiments, the camera 54 may be fixed on the image board 51 by a mounting bracket.

In this embodiment, the stacked structure further includes an upper cover 62, the upper cover 62 is disposed on the upper side of the rack main body 611 in a stacked manner, and forms a space for accommodating the image board 51 and the main control board 52 with the rack main body 611. In this embodiment, the upper cover 62 is fixed on the upper side of the frame main body 611 by screws, but the present invention is not limited thereto, and in other embodiments, the upper cover may be provided by means of a snap, a plug, a socket, or the like.

The main advantages of the above-described stacked structure include:

1. the whole body is made thinner.

2. The center of gravity is more concentrated, the integral center of gravity of the airplane body is reduced, and the flying stability and the controllability are improved.

3. Simple structure, easy manufacture and low manufacturing cost.

To sum up, the utility model provides a stack structure has very wide application prospect in the aircraft field.

6. Angle adjustment of lens

The inventor of the invention provides a novel angle adjusting structure of a lens through extensive and intensive research, and the structure has higher utilization rate of a space structure through the mutual matching and connection of a frame, the lens and an upper cover of a machine body. In this structure, the lens is provided with rotational positioning shafts on both sides thereof, and the lens is rotatably connected with the support member by the rotational positioning shafts. Specifically speaking, on the one hand through setting up the support piece on the frame rotatably connect the rotation location axle, make the camera lens can be through rotating the rotation angle of adjustment of location axle for support piece, meanwhile, rotate the location axle further with fuselage upper cover fixed connection to as fixed connection between them, make the structure more reasonable from this, thereby saved the structural space of fuselage, make overall structure littleer. And the damping gear is arranged on the contact surface of the rotary positioning shaft and the supporting piece, so that the angle adjustment of the lens connected with the supporting piece in the mode can be conveniently and finely realized, and the adjusted angle of the lens can be more easily maintained in the flying process of the aircraft.

The following describes the angle adjustment structure of the lens with reference to the accompanying drawings.

As shown in fig. 23, 24 and 25, in the present embodiment, the angle adjusting structure of the lens includes: a lens 54 and a supporting member 64 for supporting the lens 54, wherein, the lens 54 is provided with a rotational positioning shaft 63 at both sides, and the lens 54 is rotatably connected with the supporting member 64 by the rotational positioning shaft 63. Specifically, the rotational positioning shafts 63 provided on both sides of the lens 54 are rotatably provided on the support member 64, whereby the angle of the lens 54 can be adjusted by the rotation of the rotational positioning shafts 63 with respect to the support member 64, thereby achieving a lens angle adjusting function by a relatively simpler structure, and this structure effectively reduces the structural space of the aircraft and the overall volume.

Further, as described above, the lens 54 is angularly adjusted by rotating the positioning shaft 63 and the support member 64 relative to each other, in this case, the up-down rotation angle of the lens 54 is adjusted. However, the present invention is not limited to this, in other embodiments, the lens 54 further includes a fixing ring, and after the rotation positioning shaft 63 is fixedly connected to both sides of the fixing ring, the lens 54 is rotatably connected to the supporting member 64, and simultaneously, the upper and lower ends of the lens 54 are rotatably connected to the fixing ring through the second rotating shaft, in this case, the lens 54 not only can rotate the rotation positioning shaft 63 to adjust the vertical angle, but also can adjust the horizontal angle through the second rotating shaft.

In the present embodiment, the support 64 is provided on the frame 61 of the aircraft by means of integral molding. However, the present invention is not limited thereto, and in other embodiments, the supporting member 64 may be disposed on the frame 61 of the aircraft by means of screws and screw holes, inserts and sockets, or buckles.

In the present embodiment, the support 64 is connected to the upper cover 62 of the aircraft. More specifically, the front end of the upper cover 62 is provided with an area for accommodating the lens 54, which is sized and shaped to match the size of the lens 54 and its requirements for the shooting angle of view in the case of rotation. In this case, the support member 64 can both rotatably support the lens 54 and fixedly connect the frame 61 and the upper cover 62, thereby further reducing the structural space required by the aircraft and making the aircraft smaller.

In the present embodiment, the supporting member 64 is provided with a through hole, the upper cover 62 is correspondingly provided with a screw hole, and the supporting member 64 and the upper cover 62 are fixedly connected through the through hole, the screw hole and the screw. However, the present invention is not limited to this, and in other embodiments, the supporting member 64 and the upper cover 62 may also be fixedly connected through a plug-in and a plug hole, or a buckle, or other methods, which are not described herein.

In the present embodiment, the support member 63 is provided with a recess matching in position, shape and size to the rotational positioning shaft 63, and the rotational positioning shaft 64 is rotatably provided in the recess. In this case, the lens 54 can be easily attached and detached. However, the present invention is not limited to this, and in other embodiments, a through hole whose position, shape, and size match the rotational positioning shaft 63 may be provided on the support member 63, and the rotational positioning shaft 64 may be rotatably provided in the through hole. Alternatively, a concave portion whose position, shape and size are matched with the rotary positioning shaft 63 may be provided on the supporting member 63, so that both ends of the rotary positioning shaft 64 are rotatably inserted and fixed in the concave portion, and the like, which will not be described herein.

In this embodiment, the surface of the rotational positioning shaft 63 is provided with a damping gear for damping with a recess, or through hole surface, of the support 64. In this case, not only can fine adjustment of the angle of the lens 54 be more convenient, but also the angle of the lens 54 can be more easily maintained without slipping or shifting during flight or carrying of the aircraft. The utility model discloses be not limited to this, in other embodiments, also can set up damping gear on the surface of notch, or through-hole for with the surperficial damping that produces of rotational positioning axle 63, can play convenient fine adjustment angle and effectively keep the effect of angle equally, do not do here and describe repeatedly.

The angle adjusting structure of the lens has the following advantages:

1. the structure space layout is more reasonable and effective, and the whole volume of the aircraft is smaller.

2. The lens can be conveniently installed and detached, can be freely and finely adjusted, and can keep the stability of the angle adjustment in the flying and carrying processes of the aircraft.

3. The structure is simplest, the manufacture is easy, and the manufacturing cost is low.

To sum up, the utility model provides an angle modulation structure of camera lens has very wide application prospect in the aircraft field.

It should be noted that, the utility model discloses in the various new improvements that propose, both can independently realize, also can combine each other to realize, can also combine to realize as a new aircraft structure, also can use in combination each other between each technical characteristic, do not do here and describe in detail.

It should be noted that all references mentioned in the present application are incorporated by reference in the present application as if each reference were individually incorporated by reference. Furthermore, it should be understood that various changes and modifications of the present invention may be made by those skilled in the art after reading the above teachings of the present invention, and these equivalents also fall within the scope of the appended claims.

Also, in the claims and the description of the present patent, relational terms such as first and second, and the like may be used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions. Also, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus. Without further limitation, the use of the verb "comprise a" to define an element does not exclude the presence of another, same element in a process, method, article, or apparatus that comprises the element. In the claims and the specification of this patent, if it is mentioned that a certain action is performed according to a certain element, it means that the action is performed at least according to the element, and two cases are included: performing the action based only on the element, and performing the action based on the element and other elements.

While the invention has been shown and described with reference to certain preferred embodiments thereof, it will be understood by those skilled in the art that various changes in form and details may be made therein without departing from the spirit and scope of the invention.

Claims (10)

1. An airframe structure for an aircraft, comprising: a housing main body (611), and an upper cover (62) connected to the housing main body (611); wherein,

side heat radiating members (102) are arranged on two sides of the upper cover (62), so that air flow can enter the side heat radiating members (102) when the aircraft flies.

2. The airframe structure of claim 1 wherein said side heat sink member (102) comprises: the side air inlet (106) is matched with the side shape of the upper cover (62), and the blade (105) is arranged inside the side air inlet (106).

3. The frame structure of the aircraft according to claim 2, characterized in that the louvers (105) are arranged at an inclination such that the air flow can enter the side intakes (106) when the aircraft is in flight.

4. Airframe structure as claimed in claim 3, characterized in that said blades (105) are arranged in an inclined direction with a high front and a low rear in the direction of advance of said aircraft, wherein the fuselage (1) is forward in the head position and the fuselage (1) is aft in the tail position.

5. Airframe structure as claimed in claim 4, characterized in that said blades (105) have an angle with respect to the horizontal which decreases progressively from the head of the fuselage (1) towards the tail.

6. The frame structure of an aircraft according to claim 2, characterized in that the side heat-radiating part (102) further comprises a reinforcing bar (104) which is fixedly arranged at both ends in the side air inlets (106), the reinforcing bar (104) being connected to each of the louvers (105).

7. The airframe structure as recited in claim 1, further comprising a middle heat sink member (101) disposed at a middle position of said upper cover (62).

8. The frame structure of an aircraft according to claim 7, characterized in that the intermediate heat sink (101) comprises intermediate air intakes (107) and louvers (105), wherein,

the blades (105) are obliquely arranged in the middle air inlet (107) so that air flow can enter the middle air inlet (107) when the aircraft flies.

9. The frame structure of an aircraft according to claim 8, characterized in that the central heat-dissipating component (101) further comprises a reinforcing bar (104) which is arranged at both ends in the central air inlet opening (107), the reinforcing bar (104) connecting each of the louvers (105).

10. An aircraft, characterized in that it comprises a frame structure according to any one of claims 1-9.