CN115535215A - Airplane wingtip with vibration and noise reduction mechanism - Google Patents
Airplane wingtip with vibration and noise reduction mechanism Download PDFInfo
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- CN115535215A CN115535215A CN202211155014.1A CN202211155014A CN115535215A CN 115535215 A CN115535215 A CN 115535215A CN 202211155014 A CN202211155014 A CN 202211155014A CN 115535215 A CN115535215 A CN 115535215A
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- 230000007246 mechanism Effects 0.000 title claims abstract description 35
- 230000009467 reduction Effects 0.000 title claims abstract description 32
- 238000013016 damping Methods 0.000 claims abstract description 4
- 239000000758 substrate Substances 0.000 claims description 17
- 230000006698 induction Effects 0.000 claims description 9
- 238000009434 installation Methods 0.000 claims description 9
- 230000005540 biological transmission Effects 0.000 claims description 3
- 238000004891 communication Methods 0.000 claims description 3
- 239000010720 hydraulic oil Substances 0.000 claims description 3
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- 238000000034 method Methods 0.000 description 7
- 230000008569 process Effects 0.000 description 6
- 238000012423 maintenance Methods 0.000 description 3
- 238000001514 detection method Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/36—Structures adapted to reduce effects of aerodynamic or other external heating
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Abstract
The invention discloses an airplane wingtip with a vibration and noise reduction mechanism, and relates to the technical field of airplane wing structure design. The wing tip counterweight mechanism comprises a wing tip base plate, a connecting framework, a supporting plate and a counterweight mechanism, wherein the counterweight mechanism comprises a front driving cylinder and a rear driving cylinder which are both electric hydraulic cylinders, and further comprises an adjusting motor, an adjusting screw rod, a front counterweight block and a rear counterweight block. The invention adopts a mode of adding a counterweight structure to change the gravity center of the wing so as to realize the functions of vibration reduction and noise reduction by means of stress balance; in actual assembly, the wing is divided into a front wing area and a rear wing area according to the specific structure and stress condition of the wing or the wing tip, the front wing area is greatly compressed, the amplitude is large, and therefore the balancing weight for vibration reduction is large; in addition, the front driving cylinder, the rear driving cylinder, the adjusting motor and the adjusting screw rod are arranged and respectively used as a longitudinal position moving mechanism and a transverse position moving mechanism for the front balancing weight and the rear balancing weight, so that the front balancing weight and the rear balancing weight can be matched with each other to be used at a high-frequency vibration site, and the damping effect on the site is realized.
Description
Technical Field
The invention belongs to the technical field of airplane wing structure design, and particularly relates to an airplane wingtip with a vibration and noise reduction mechanism.
Background
When the flying speed of an airplane reaches a certain critical value in the process of flying, the surface of the wing generates a flutter phenomenon under the coupling action of the elastic force, the inertia force and the aerodynamic force of a structure, namely commonly called vibration. The fundamental reason for the wing vibration is that when the wing slides in the air, the center of the wing deviates due to unbalanced stress, and the two phenomena are mutually superposed to finally generate flutter with amplitude difficult to attenuate; in order to solve the flutter phenomenon, the most common method in the prior art is to add a group of ailerons on one side of the wing, the ailerons can swing according to air flow, and the gravity center and the balance weight of the wing can be adjusted constantly, but the mode has larger damage to the wing, the long-time swing torsion ailerons are easy to wear and collide with the wing, and the maintenance cost is higher; on the other hand, the mode is difficult to fundamentally solve the problem of flutter, introduces new maintenance cost and the problem of part damage, and simultaneously swings and twists the ailerons for a long time to generate larger noise during flight; therefore, in order to improve the prior art, the aircraft with the vibration and noise reduction mechanism is designed.
Disclosure of Invention
The invention aims to provide an aircraft wingtip with a vibration and noise reduction mechanism, which solves the problems of easy damage, high maintenance cost and noise generation of the existing wings.
In order to solve the technical problems, the invention is realized by the following technical scheme:
the invention relates to an aircraft wing tip with a vibration and noise reduction mechanism, which comprises a wing tip substrate, a connecting framework, a supporting plate and a counterweight mechanism, wherein the wing tip substrate is of a hollow structure, and the counterweight mechanism, the connecting framework and the supporting plate are all arranged in the wing tip substrate; the upper surface and the lower surface of the supporting plate are welded and fixed with the inner surface of the wingtip substrate, and the inside of the wingtip substrate is respectively provided with a front winged area and a rear winged area through the supporting plate, wherein aiming at the vibration reduction of the wingtip of the airplane, the front winged area and the rear winged area respectively correspond to the front half part and the rear half part of the wing, and the frequencies of the two areas generating vibration during flying usually have difference, so the vibration reduction measures are correspondingly separated;
the technical scheme is mainly characterized in that the vibration reduction and noise reduction related structure is set according to the reason that the wings of the airplane vibrate during flying, namely, the whole wings or the wing tips are stressed uniformly and balanced in a mode of adding local counterweights;
the counterweight mechanism comprises a front driving cylinder, a rear driving cylinder and a supporting cylinder, wherein the front driving cylinder and the rear driving cylinder are both electric hydraulic cylinders; because the front and the back of the wing are subjected to different atmospheric pressures in the flying process, the frequency of the wing tip when the wing tip vibrates is in direct proportion to the external pressure, and the specifications of the corresponding front driving cylinder and the rear driving cylinder are different, the power of the front driving cylinder is greater than that of the rear driving cylinder; the front driving cylinder and the supporting cylinder are both arranged in the front wing area, and the rear driving cylinder is arranged in the rear wing area; the front driving cylinder, the rear driving cylinder and the supporting cylinder are identical in structure and comprise mounting sleeves and pushing rods, and the pushing rods are clamped with the mounting sleeves in a sliding mode; the support cylinder is used as an auxiliary structure of the front drive cylinder, and an independent drive source is not needed, so that the support cylinder is only of a common hydraulic cylinder structure;
the connecting framework comprises a plurality of pipe frames which are mutually welded and communicated; the bottom surface of the mounting sleeve is in bolted communication with the pipe frame, and hydraulic oil is filled in the mounting sleeve and the pipe frame; the front driving cylinder and the supporting cylinder are mutually communicated through a pipe frame, and a communicating device structure based on a hydraulic transmission principle is formed; the front driving cylinder and the rear driving cylinder are mutually isolated through a supporting plate; the counterweight mechanism also comprises a front adjusting frame, a rear adjusting frame, a front counterweight block and a rear counterweight block, wherein the front adjusting frame is welded and fixed between two groups of push rods which are opposite to each other and are arranged between the front driving cylinder and the supporting cylinder; the front adjusting frame and the rear adjusting frame are both of frame structures, and the front balancing weight is clamped with the inner surface of the front adjusting frame in a sliding manner; one end of the front adjusting frame is fixedly bolted with an adjusting motor, and one end of an output shaft of the adjusting motor is welded with an adjusting screw rod; the adjusting screw rod penetrates through the front balancing weight and is in rotary shaft connection with the front adjusting frame, and a screw rod structure is formed between the adjusting screw rod and the front balancing weight by arranging a thread groove; by combining the structure, in order to relieve the vibration generated by high pressure, the weight of the corresponding front balancing weight is larger than that of the rear balancing weight, so that the supporting cylinder is required to perform auxiliary support when the position of the front balancing weight is adjusted, and the structure of the communicating vessel enables the whole structure to be more labor-saving; during actual installation, the installation positions of the front driving cylinder and the supporting cylinder are opposite, and when the position of the front balancing weight needs to be adjusted, the front driving cylinder and the supporting cylinder are matched with each other, so that the adjusting process is more labor-saving;
in the structure, the front driving cylinder and the supporting cylinder are matched with each other to be used for adjusting the longitudinal position of the front balancing weight, the adjusting motor and the adjusting screw can adjust the transverse position of the front balancing weight, the front balancing weight and the adjusting screw are combined, so that the high-frequency point position of the surface vibration of the wing tip base plate is realized, the front balancing weight is moved to the position, and the vibration effect can be effectively weakened.
Furthermore, the assembly structure between the rear adjusting frame and the rear balancing weight is the same as that of the front wing area, and an adjusting motor and an adjusting screw rod are also arranged; a screw rod structure is formed between the rear balancing weight and the adjusting screw rod through arranging a thread groove; the front driving cylinders, the rear driving cylinders and the supporting cylinders are the same in number, and the installation positions are opposite.
Furthermore, signal receivers are arranged on the surfaces of the front balancing weight and the rear balancing weight; a plurality of vibration induction grooves are formed in the wall plate of the wingtip substrate, piezoelectric sensors are welded on the inner surfaces of the vibration induction grooves, and conduction springs are bonded between the piezoelectric sensors and the outer wall of the wingtip substrate; the piezoelectric sensor is electrically connected with the signal receiver; the vibration induction grooves and the piezoelectric sensors are distributed all over each aircraft wallboard, when vibration occurs, the outer wall of each aircraft wallboard can continuously apply pressure to the piezoelectric sensors at high-frequency vibration through the conduction springs, and meanwhile, the piezoelectric sensors transmit signals to the signal receiver, so that accurate coordinates of vibration sites can be determined in the subsequent process.
Furthermore, a microcontroller is arranged in the support plate, the front driving cylinder, the rear driving cylinder and the adjusting motor are electrically connected with the microcontroller, and the front driving cylinder and the rear driving cylinder are electrically connected in parallel; the adjusting motor in the front wing area and the adjusting motor in the rear wing area are electrically connected in parallel, the signal receiver is electrically connected with the microcontroller, and the front driving cylinder, the rear driving cylinder and the adjusting motor are started through the microcontroller; after receiving the piezoelectric signal, the microcontroller accurately positions the coordinates of the high-frequency vibration site by using a built-in program, then respectively controls the adjusting motor and the front driving cylinder or the rear driving cylinder, and moves the front balancing weight or the rear balancing weight to the monitored site for balancing and vibration reduction.
In the technical scheme, the adjustability of the front balancing weight and the rear balancing weight is crucial and different from that of the traditional hinged wing, the scheme is that the balancing weight structure is arranged inside the wing tip base plate and is divided into the front balancing weight and the rear balancing weight according to the reasons that the front stress and the rear stress of the wing are different, and the front balancing weight and the rear balancing weight are matched with each other, so that the high-frequency vibration point can be directly accurately positioned and damped; for example, vibration occurs at a point in the front wing area of the wingtip, positioning and detection are performed on structures such as a piezoelectric sensor, a signal receiver and the like, and finally the front balancing weight is moved to reach the corresponding point; at the moment, although the vibration at the position is relieved, the unbalance phenomenon probably occurs on the whole wing, the vibration occurs again, the rear balancing weight plays a role at the moment and is matched with the front balancing weight, and the wing tip and even the whole wing of the airplane can achieve dynamic balance.
The invention has the following beneficial effects:
the invention adopts a mode of adding a counterweight structure to change the gravity center of the wing so as to realize the functions of vibration reduction and noise reduction by means of stress balance; in actual assembly, the wing is divided into a front wing area and a rear wing area according to the specific structure and stress condition of the wing or the wing tip, the front wing area is greatly compressed and has large amplitude, so that the balancing weight for vibration reduction is larger, and the weight of the balancing weight of the rear wing area is smaller than that of the front balancing weight correspondingly; in addition, the front driving cylinder, the rear driving cylinder, the adjusting motor and the adjusting screw rod are respectively used as a longitudinal and transverse position moving mechanism for the front balancing weight and the rear balancing weight, so that the front balancing weight and the rear balancing weight can be matched with each other to be used at a high-frequency vibration site, and the damping effect on the site is realized; on the other hand, because two balancing weights in front and back all set up inside the wingtip base plate, can not receive outside air pressure, consequently can further reduce the production of noise when actual work, this compares with aileron damping structure among the prior art and has unexpected effect.
Of course, it is not necessary for any product to practice the invention to achieve all of the above-described advantages at the same time.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings used in the description of the embodiments will be briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art that other drawings can be obtained according to the drawings without creative efforts.
FIG. 1 is a partial block diagram of an aircraft wingtip with a vibration and noise reduction mechanism according to the present invention;
FIG. 2 is a front view of an aircraft wingtip with a vibration and noise reduction mechanism of the present invention;
FIG. 3 isbase:Sub>A schematic structural view of section A-A in FIG. 2;
FIG. 4 is a partial display view of portion C of FIG. 3;
FIG. 5 is a schematic structural view of section D-D in FIG. 3;
FIG. 6 is a partial display view of portion E of FIG. 5;
fig. 7 is a schematic structural view of a section B-B in fig. 2.
In the drawings, the reference numbers indicate the following list of parts:
1. a tip substrate; 2. connecting the framework; 3. a support plate; 4. an anterior wing region; 5. a posterior wing region; 6. a front drive cylinder; 7. a rear drive cylinder; 8. a support cylinder; 9. installing a sleeve; 10. a push rod; 11. a front adjusting bracket; 12. a rear adjusting bracket; 13. a front balancing weight; 14. a rear balancing weight; 15. adjusting the motor; 16. adjusting a screw rod; 17. a signal receiver; 18. a vibration induction tank; 19. a piezoelectric sensor; 20. a conductive spring.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
In the description of the present invention, it is to be understood that the terms "upper", "middle", "outer", "inner", and the like, indicate orientations or positional relationships, are used for convenience in describing the present invention and simplifying the description, but do not indicate or imply that the referenced components or elements must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be construed as limiting the present invention.
Referring to fig. 1 to 7, the invention relates to an aircraft wing tip with a vibration and noise reduction mechanism, which comprises a wing tip substrate 1, a connecting framework 2, a supporting plate 3 and a counterweight mechanism, wherein the wing tip substrate 1 is of a hollow structure, and the counterweight mechanism, the connecting framework 2 and the supporting plate 3 are all arranged inside the wing tip substrate 1, wherein the connecting framework 2 is of a grid structure and is fixedly connected with the inner surface of the wing tip substrate 1, the basic constituent unit of the wing tip substrate 1 is an aircraft panel with various dimensions, and the aircraft panel is fixedly connected with the connecting framework 2 in a welding and riveting manner; the upper surface and the lower surface of the support plate 3 are welded and fixed with the inner surface of the wingtip substrate 1, and the inside of the wingtip substrate 1 is respectively provided with a front wing area 4 and a rear wing area 5 through the support plate 3, wherein aiming at the vibration reduction of the wingtip of the airplane, the front wing area 4 and the rear wing area 5 respectively correspond to the front half part and the rear half part of the wing, and the frequencies of the two areas which generate vibration during flying usually have difference, so the vibration reduction measures are correspondingly separated;
the technical scheme is mainly characterized in that the vibration reduction and noise reduction related structure is set according to the reason that the wings of the airplane vibrate during flying, namely, the whole wings or the wing tips are stressed uniformly and balanced in a mode of adding local counterweights;
the counterweight mechanism comprises a front driving cylinder 6, a rear driving cylinder 7 and a supporting cylinder 8, wherein the front driving cylinder 6 and the rear driving cylinder 7 are both electric hydraulic cylinders; because the front and the back of the wing are subjected to different atmospheric pressures in the flying process, the frequency of the wing tip when the wing tip vibrates is in direct proportion to the external pressure, and the specifications of the corresponding front driving cylinder 6 and the corresponding back driving cylinder 7 are different, the power of the front driving cylinder 6 is greater than that of the back driving cylinder 7; the front driving cylinder 6 and the supporting cylinder 8 are both arranged in the front wing area 4, and the rear driving cylinder 7 is arranged in the rear wing area 5; the front driving cylinder 6, the rear driving cylinder 7 and the supporting cylinder 8 have the same structure and structure, and respectively comprise an installation sleeve 9 and a pushing rod 10, and the pushing rod 10 is clamped with the installation sleeve 9 in a sliding manner; the supporting cylinder 8 is used as an auxiliary structure of the front driving cylinder 6, and an independent driving source is not needed, so that the supporting cylinder 8 is only of a common hydraulic cylinder structure;
the connecting framework 2 comprises a plurality of pipe frames which are mutually welded and communicated; the bottom surface of the mounting sleeve 9 is in bolted communication with the pipe frame, and the mounting sleeve 9 and the pipe frame are filled with hydraulic oil; the front driving cylinder 6 and the supporting cylinder 8 are communicated with each other through a pipe frame, and a communicating vessel structure based on a hydraulic transmission principle is formed; the front driving cylinder 6 and the rear driving cylinder 7 are mutually isolated by the supporting plate 3; the counterweight mechanism also comprises a front adjusting frame 11, a rear adjusting frame 12, a front counterweight block 13 and a rear counterweight block 14, wherein the front adjusting frame 11 is welded and fixed between two groups of pushing rods 10 which are opposite to each other of the front driving cylinder 6 and the supporting cylinder 8; the front adjusting frame 11 and the rear adjusting frame 12 are both frame structures, and the front balancing weight 13 is clamped with the inner surface of the front adjusting frame 11 in a sliding manner; one end of the front adjusting frame 11 is fixedly bolted with an adjusting motor 15, and one end of an output shaft of the adjusting motor 15 is welded with an adjusting screw rod 16; the adjusting screw rod 16 penetrates through the front balancing weight block 13 and is in rotary shaft connection with the front adjusting frame 11, and a screw rod structure is formed between the adjusting screw rod 16 and the front balancing weight block 13 by arranging a thread groove; in combination with the above structure, in order to alleviate the vibration generated by high pressure, the weight of the corresponding front balancing weight 13 is greater than that of the rear balancing weight 14, so that the supporting cylinder 8 is required to perform auxiliary support when the position of the front balancing weight 13 is adjusted, and the structure of the communicating vessel makes the whole structure more labor-saving; during actual installation, the installation positions of the front driving cylinder 6 and the supporting cylinder 8 are opposite, and when the position of the front balancing weight 13 needs to be adjusted, the front driving cylinder 6 and the supporting cylinder 8 are matched with each other, so that the adjusting process is more labor-saving;
in the structure, the front driving cylinder 6 and the supporting cylinder 8 are matched with each other to adjust the longitudinal position of the front balancing weight 13, the adjusting motor 15 and the adjusting screw 16 can adjust the transverse position of the front balancing weight 13, and the front balancing weight and the adjusting screw are combined to form a high-frequency point position of surface vibration of the wing tip base plate 1, so that the front balancing weight 13 is moved to the position, and the vibration effect can be effectively weakened.
Preferably, the assembly structure between the rear adjusting frame 12 and the rear balancing weight 14 is the same as that of the front wing area 4, and an adjusting motor 15 and an adjusting screw rod 16 are also arranged; a screw rod structure is formed between the rear balancing weight 14 and the adjusting screw rod 16 by arranging a thread groove; the front driving cylinder 6, the rear driving cylinder 7 and the supporting cylinder 8 are the same in number, and the installation positions are opposite.
Preferably, the surfaces of the front weight block 13 and the rear weight block 14 are both provided with a signal receiver 17; a plurality of vibration induction grooves 18 are formed in the wall plate of the wingtip base plate 1, piezoelectric sensors 19 are welded on the inner surfaces of the vibration induction grooves 18, and conduction springs 20 are adhered between the piezoelectric sensors 19 and the outer wall of the wingtip base plate 1; the piezoelectric sensor 19 is electrically connected with the signal receiver 17; the vibration induction groove 18 and the piezoelectric sensor 19 are distributed on each aircraft wall, when vibration occurs, the outer wall of the aircraft wall can continuously apply high-intensity pressure on the piezoelectric sensor 19 through the conductive spring 20 under high-frequency vibration, and simultaneously transmit signals to the signal receiver 17, so that accurate coordinates of a vibration point can be determined subsequently.
Preferably, a microcontroller is arranged in the support plate 3, and the front driving cylinder 6, the rear driving cylinder 7 and the adjusting motor 15 are electrically connected with the microcontroller, wherein the front driving cylinder 6 and the rear driving cylinder 7 are electrically connected in parallel; the adjusting motor 15 in the front wing area 4 and the adjusting motor 15 in the rear wing area 5 are electrically connected in parallel, the signal receiver 17 is electrically connected with the microcontroller, and the front driving cylinder 6, the rear driving cylinder 7 and the adjusting motor 15 are started through the microcontroller; after receiving the piezoelectric signal, the microcontroller accurately positions the coordinates of the high-frequency vibration site by using a built-in program, then respectively controls the adjusting motor 15 and the front driving cylinder 6 or the rear driving cylinder 7, and moves the front balancing weight 13 or the rear balancing weight 14 to the monitored site for balancing and vibration reduction.
In the technical scheme, the adjustability of the front balancing weight 13 and the rear balancing weight 14 is crucial and different from that of the traditional hinged wing, the balancing weight structure is arranged inside the wing tip base plate 1, the balancing weight structure is divided into the front balancing weight 13 and the rear balancing weight 14 according to the reasons that the front stress and the rear stress of the wing are different, and the front balancing weight 13 and the rear balancing weight are matched with each other, so that the high-frequency vibration point can be directly and accurately positioned and damped; for example, vibration occurs at a point in the front wing area 4 of the wingtip, positioning and detection are carried out on structures such as the piezoelectric sensor 19 and the signal receiver 17, and finally the front balancing weight 13 is moved to reach a corresponding point; at this time, although the vibration at this position is relieved, the unbalance phenomenon may occur on the whole wing, and the vibration occurs again, at this time, the rear balancing weight 14 plays a role and is matched with the front balancing weight 13, so that the dynamic balance of the wing tip and even the whole wing of the airplane can be achieved.
In the description herein, references to the description of "one embodiment," "an example," "a specific example" or the like are intended to mean that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the invention. In this specification, the schematic representations of the terms used above do not necessarily refer to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples.
The preferred embodiments of the invention disclosed above are intended to be illustrative only. The preferred embodiments are not intended to be exhaustive or to limit the invention to the precise embodiments disclosed. Obviously, many modifications and variations are possible in light of the above teaching. The embodiments were chosen and described in order to best explain the principles of the invention and the practical application, to thereby enable others skilled in the art to best utilize the invention. The invention is limited only by the claims and their full scope and equivalents.
Claims (7)
1. Aircraft wingtip with mechanism of making an uproar falls in damping, including wingtip base plate (1), connection skeleton (2), backup pad (3) and counter weight mechanism, its characterized in that: the wing tip base plate (1) is of a hollow structure, and the counterweight mechanism, the connecting framework (2) and the supporting plate (3) are all arranged inside the wing tip base plate (1), wherein the connecting framework (2) is of a grid structure and is fixedly connected with the inner surface of the wing tip base plate (1); the upper surface and the lower surface of the supporting plate (3) are welded and fixed with the inner surface of the wingtip base plate (1), and a front wing area (4) and a rear wing area (5) are arranged in the wingtip base plate (1) through the supporting plate (3);
the counterweight mechanism comprises a front driving cylinder (6), a rear driving cylinder (7) and a supporting cylinder (8), wherein the front driving cylinder (6) and the rear driving cylinder (7) are both electric hydraulic cylinders, the front driving cylinder (6) and the supporting cylinder (8) are both arranged in the front wing area (4), and the rear driving cylinder (7) is arranged in the rear wing area (5); the front driving cylinder (6), the rear driving cylinder (7) and the supporting cylinder (8) are identical in structure and comprise mounting sleeves (9) and pushing rods (10), and the pushing rods (10) are clamped with the mounting sleeves (9) in a sliding mode;
the connecting framework (2) comprises a plurality of pipe frames which are mutually welded and communicated; the bottom surface of the mounting sleeve (9) is in bolted connection and communication with the pipe frame, and hydraulic oil is filled in the mounting sleeve (9) and the pipe frame; the front driving cylinder (6) and the supporting cylinder (8) are communicated with each other through a pipe frame, and a communicating vessel structure based on a hydraulic transmission principle is formed; the front driving cylinder (6) and the rear driving cylinder (7) are isolated from each other through the supporting plate (3).
2. The aircraft wing tip with the vibration and noise reduction mechanism is characterized in that the counterweight mechanism further comprises a front adjusting frame (11), a rear adjusting frame (12), a front counterweight block (13) and a rear counterweight block (14), wherein the front adjusting frame (11) is welded and fixed between two groups of pushing rods (10) opposite to the front driving cylinder (6) and the supporting cylinder (8); the front adjusting frame (11) and the rear adjusting frame (12) are both of a frame structure, and the front balancing weight (13) is clamped with the inner surface of the front adjusting frame (11) in a sliding manner; one end of the front adjusting frame (11) is fixedly connected with an adjusting motor (15) in a bolted mode, and one end of an output shaft of the adjusting motor (15) is welded with an adjusting screw rod (16); the adjusting screw rod (16) penetrates through the front balancing weight (13) and is in rotary shaft connection with the front adjusting frame (11), and a screw rod structure is formed between the adjusting screw rod (16) and the front balancing weight (13) by arranging a thread groove.
3. The aircraft wingtip with the vibration and noise reduction mechanism is characterized in that the assembly structure between the rear adjusting frame (12) and the rear balancing weight (14) is the same as that of the front wing area (4), and an adjusting motor (15) and an adjusting screw rod (16) are also arranged; and a screw rod structure is formed between the rear balancing weight (14) and the adjusting screw rod (16) by arranging a thread groove.
4. The aircraft wing tip with the vibration and noise reduction mechanism according to claim 3, characterized in that the front driving cylinders (6), the rear driving cylinders (7) and the supporting cylinders (8) are the same in number, and the installation positions are opposite.
5. The aircraft wing tip with the vibration and noise reduction mechanism according to claim 4, characterized in that the surfaces of the front counterweight block (13) and the rear counterweight block (14) are provided with signal receivers (17); a plurality of vibration induction grooves (18) are formed in the wall plate of the wingtip substrate (1), piezoelectric sensors (19) are welded on the inner surfaces of the vibration induction grooves (18), and a conduction spring (20) is adhered between each piezoelectric sensor (19) and the outer wall of the wingtip substrate (1).
6. The aircraft wing tip with the vibration and noise reduction mechanism according to claim 5, characterized in that a microcontroller is arranged in the support plate (3), the front driving cylinder (6), the rear driving cylinder (7) and the adjusting motor (15) are electrically connected with the microcontroller, wherein the front driving cylinder (6) and the rear driving cylinder (7) are electrically connected in parallel; the adjusting motor (15) in the front wing area (4) is electrically connected in parallel with the adjusting motor (15) in the rear wing area (5).
7. The aircraft wing tip with vibration and noise reduction mechanism according to claim 6, characterized in that the piezoelectric sensor (19) is electrically connected with a signal receiver (17), and the signal receiver (17) is electrically connected with a microcontroller, and the front driving cylinder (6), the rear driving cylinder (7) and the adjusting motor (15) are started by the microcontroller.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202211155014.1A CN115535215A (en) | 2022-09-22 | 2022-09-22 | Airplane wingtip with vibration and noise reduction mechanism |
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| CN202211155014.1A CN115535215A (en) | 2022-09-22 | 2022-09-22 | Airplane wingtip with vibration and noise reduction mechanism |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115857052A (en) * | 2023-02-22 | 2023-03-28 | 中国民用航空飞行学院 | Navigation meteorological information receiving equipment with noise reduction function and broadcasting device |
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| US20200001987A1 (en) * | 2016-03-14 | 2020-01-02 | Bell Helicopter Textron, Inc. | Rotor hub vibration attenuator |
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2022
- 2022-09-22 CN CN202211155014.1A patent/CN115535215A/en active Pending
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| US3836098A (en) * | 1971-03-08 | 1974-09-17 | Kawasaki Heavy Ind Ltd | Vibration absorbing device for rotorcraft |
| JPH05193580A (en) * | 1992-01-20 | 1993-08-03 | Mitsubishi Heavy Ind Ltd | Vibration damping device for helicopter |
| WO2003011689A1 (en) * | 2001-08-02 | 2003-02-13 | Bell Helicopter Textron Inc. | Helicopter main rotor blade balance weight retention assembly |
| US20040222329A1 (en) * | 2002-08-02 | 2004-11-11 | Norman Kuhns | Helicopter main rotor blade balance weight retention assembly |
| US20170259911A1 (en) * | 2016-03-14 | 2017-09-14 | Bell Helicopter Textron, Inc. | Rotor hub vibration attenuator |
| US20200001987A1 (en) * | 2016-03-14 | 2020-01-02 | Bell Helicopter Textron, Inc. | Rotor hub vibration attenuator |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115857052A (en) * | 2023-02-22 | 2023-03-28 | 中国民用航空飞行学院 | Navigation meteorological information receiving equipment with noise reduction function and broadcasting device |
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