Aircraft engine protection device

Abstract

The present invention relates to the technical field of aircraft engine protection, and in particular to an aircraft engine protection device, comprising a fuselage, the fuselage is placed in a horizontal state, a center column is fixedly installed at the fuselage axis position, a turbine fan blade group is sleeved on the outside of the center column, a protection component is fixedly installed at one end of the fuselage outer side, the protection component comprises a support frame fixedly installed on the outer side of the fuselage end, the support frame is fixedly connected to the fixed frame through a plurality of connecting plates, a plurality of limit rings are fixedly connected to the outside of the fixed frame, a sliding sleeve is slidably installed on the outer side of the limit ring, a receiving plate is fixedly connected to the outer side of the sliding sleeve, an arc plate is hinged at one end of the receiving plate away from the sliding sleeve, a driving component is installed in the fixed frame, and a plurality of limit rings are driven by the driving component. Compared with the prior art, the present invention has the following beneficial effects: the overall operation process is simple, the fuselage opening is protected as a whole, and at the same time, when the aircraft engine needs to be used, the stability of the device is effectively improved, and the safety performance of the device is improved.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft engine protection, in particular to an aircraft engine protection device. Background Art

[0002] An aircraft engine is a highly complex and sophisticated thermal machine. As the heart of an aircraft, it is not only the driving force for the aircraft to fly, but also an important driving force for the development of the aviation industry. Every important change in human aviation history is closely related to the technological progress of aircraft engines.

[0003] After more than a hundred years of development, aircraft engines have developed into mature products with extremely high reliability. The aircraft engines in use include various types such as turbojet / turbofan engines, turboshaft / turboprop engines, ramjet engines and piston engines. They are not only used as power for military and civilian aircraft, drones and cruise missiles for various purposes, but also gas turbines derived from aircraft engines are widely used in ground power generation, marine power, mobile power stations, natural gas and oil pipeline pumping stations and other fields.

[0004] Under the existing technical requirements, in order to ensure that the aircraft engine has enough power to drive the aircraft to fly, no airflow interference can be made to the aircraft engine when it is started. Therefore, the aircraft engine is mostly in an open state in the open air. When the aircraft engine is not working, its relatively open state makes it easy for it to be invaded by birds, humans or other natural impurities, so that foreign matter exists inside the aircraft engine and causes serious accidents. Therefore, the aircraft engine should be protected when it is not in use to avoid the occurrence of the above problems. However, the aircraft engine cannot have any behavior that interferes with the airflow during use. In view of this type of problem, it is urgently necessary to design an aircraft engine protection device. Summary of the invention

[0005] The purpose of this section is to summarize some aspects of embodiments of the present invention and briefly introduce some preferred embodiments. Some simplifications or omissions may be made in this section and the specification abstract and the invention title of this application to avoid blurring the purpose of this section, the specification abstract and the invention title, and such simplifications or omissions cannot be used to limit the scope of the present invention.

[0006] In view of the above problems or problems existing in the prior art, the present invention is proposed.

[0007] Therefore, an object of the present invention is to provide an aircraft engine protection device.

[0008] In order to solve the above technical problems, the present invention provides the following technical solutions: an aircraft engine protection device, comprising a fuselage, the fuselage is placed in a horizontal state, a center column is fixedly installed at the axis position of the fuselage, a turbine fan blade group is sleeved on the outside of the center column, and a protection component is fixedly installed at one end of the outer side of the fuselage;

[0009] The protection component includes a support frame fixedly installed on the outside of the end of the machine body, the support frame is fixedly connected to the fixed frame through a plurality of connecting plates, a plurality of limit rings are fixedly connected to the outside of the fixed frame, a sliding sleeve is slidably installed on the outside of the limit ring, a receiving plate is fixedly connected to the outside of the sliding sleeve, an arc plate is hinged at one end of the receiving plate away from the sliding sleeve, a driving component is installed in the fixed frame, and the plurality of limit rings are driven by the driving component.

[0010] An aircraft engine protection device is a preferred embodiment of the present invention, wherein: the drive assembly includes two gear rings slidably mounted on the inner side of a fixed frame, a motor is fixedly mounted on the side of the fixed frame, and the output end of the motor passes through the side wall of the fixed frame and is sleeved with a gear, a recessed groove is provided at the location where the motor is mounted on the fixed frame, an avoidance groove is provided at the center of one of the connecting plates, and a gear is mounted at the junction of the avoidance groove and the recessed groove, and the two gear rings are meshed with the gear.

[0011] An aircraft engine protection device is a preferred embodiment of the present invention, wherein: a plurality of mounting grooves are provided at the outermost portions of the two gear rings, and the mounting grooves on the two gear rings are symmetrically distributed, steel balls are installed in the mounting grooves, a plurality of inclined grooves are provided on the inner side wall of the sliding sleeve, and the steel balls roll in the inclined grooves.

[0012] An aircraft engine protection device is a preferred embodiment of the present invention, wherein: a connecting rod is fixedly installed at the hinge between the arc plate and the receiving plate, a double torsion spring is sleeved on the outside of the connecting rod, and the two ends of the double torsion spring are set against the side walls of the arc plate and the receiving plate.

[0013] An aircraft engine protection device is a preferred solution of the present invention, wherein: both ends of the bottom of the receiving plate are formed with limited flanges, and the top of the arc plate is fixedly installed with a limited stop at the hinge between the receiving plate and the receiving plate.

[0014] An aircraft engine protection device is a preferred embodiment of the present invention, wherein: a limit plate is fixedly installed on the outside of the front end of the center column, a side wall of the body on which one end of the support frame is installed is provided with a plurality of mounting grooves, and the support frame is provided with a plurality of limit grooves that cooperate with the mounting grooves.

[0015] An aircraft engine protection device is a preferred solution of the present invention, wherein: the side wall of the receiving plate is provided with a mouth-shaped groove for fixing the double torsion spring, and the side wall of the arc plate is provided with a straight groove for the double torsion spring to slide.

[0016] The present invention has a beneficial effect: first, when the aircraft engine needs to be protected, the present invention turns on the motor, drives the receiving plate and the arc plate to rotate synchronously through the driving assembly, until the side of the arc plate contacts the limit plate, and the overall closed protection process can be completed. At the same time, when the aircraft engine needs to be used, it is only necessary to reverse the motor, rotate the arc plate and the receiving plate to be close to the side wall of the body shell through the sliding sleeve, and then the aircraft engine can be operated normally. The overall operation process is simple, and several groups of arc plates are used to cooperate with each other to completely close the front end of the body, and the overall opening of the body is protected, which effectively reduces the possibility of damage to the aircraft engine body when it is in standby mode. At the same time, when the aircraft engine needs to be used, the arc plate is flipped to the body shell, and the arc plate is locked by the elastic force generated by the deformation of the double torsion spring, which effectively improves the stability of the device and further improves the safety performance of the device. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the following briefly introduces the drawings required for describing the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative labor. Among them:

[0018] Figure 1 It is a schematic diagram of the closed overall structure of the present invention;

[0019] Figure 2 The overall structural diagram of the present invention is opened;

[0020] Figure 3 This is a schematic diagram of the overall structure of the present invention after the body is removed;

[0021] Figure 4 It is a schematic diagram of some structures in the present invention;

[0022] Figure 5 for Figure 4 A schematic diagram of the enlarged structure in the middle;

[0023] Figure 6 for Figure 3 Schematic diagram of the overall structure of some components in the middle;

[0024] Figure 7 It is a schematic diagram of the structure of the sliding sleeve and the receiving plate in the present invention;

[0025] Figure 8 It is a schematic diagram of the arc plate and double torsion spring structure in the present invention.

[0026] In the figure: 1. body; 2. center column; 3. turbine fan blade group; 4. support frame; 5. connecting plate; 6. fixing frame; 7. limiting ring; 8. sliding sleeve; 9. receiving plate; 10. arc plate; 11. gear ring; 12. motor; 13. gear; 14. sink groove; 15. avoidance groove; 16. installation groove; 17. steel ball; 18. inclined groove; 19. connecting rod; 20. double torsion spring; 21. limiting flange; 22. limiting stop; 23. limiting plate; 24. placement groove; 25. limiting groove; 26. mouth groove; 27. straight groove. DETAILED DESCRIPTION

[0027] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the specific implementation methods of the present invention are described in detail below in conjunction with the accompanying drawings.

[0028] In the following description, many specific details are set forth to facilitate a full understanding of the present invention, but the present invention may also be implemented in other ways different from those described herein, and those skilled in the art may make similar generalizations without violating the connotation of the present invention. Therefore, the present invention is not limited to the specific embodiments disclosed below.

[0029] Secondly, the term "one embodiment" or "embodiment" as used herein refers to a specific feature, structure, or characteristic that may be included in at least one implementation of the present invention. The term "in one embodiment" that appears in different places in this specification does not necessarily refer to the same embodiment, nor does it refer to a separate or selective embodiment that is mutually exclusive with other embodiments.

[0030] Example

[0031] Reference Figures 1 to 8 An aircraft engine protection device comprises a body 1, the body 1 is placed in a horizontal state, a center column 2 is fixedly installed at the axis position of the body 1, a turbine fan blade group 3 is sleeved on the outside of the center column 2, and a protective component is fixedly installed at one end of the outer side of the body 1;

[0032] The protection component includes a support frame 4 fixedly installed on the outside of the end of the body 1, the support frame 4 is fixedly connected to the fixed frame 6 through a plurality of connecting plates 5, a plurality of limit rings 7 are fixedly connected to the outside of the fixed frame 6, a sliding sleeve 8 is slidably installed on the outside of the limit ring 7, a receiving plate 9 is fixedly connected to the outside of the sliding sleeve 8, and an arc plate 10 is hinged at one end of the receiving plate 9 away from the sliding sleeve 8, a driving component is installed in the fixed frame 6, and the plurality of limit rings 7 are driven by the driving component.

[0033] Furthermore, the driving assembly includes two gear rings 11 slidably mounted on the inner side of the fixed frame 6, a motor 12 is fixedly mounted on the side of the fixed frame 6, and the output end of the motor 12 passes through the side wall of the fixed frame 6 and is sleeved with a gear 13, a groove 14 is provided at the fixed frame 6 where the motor 12 is installed, one of the connecting plates 5 has an avoidance groove 15 in the center, and the gear 13 is installed at the junction of the avoidance groove 15 and the sinking groove 14, and the two gear rings 11 are meshed with the gear 13.

[0034] Preferably, a plurality of mounting grooves 16 are provided at the outermost portions of the two gear rings 11 , and the mounting grooves 16 on the two gear rings 11 are symmetrically distributed, steel balls 17 are installed in the mounting grooves 16 , a plurality of inclined grooves 18 are provided on the inner side wall of the sliding sleeve 8 , and the steel balls 17 roll in the inclined grooves 18 .

[0035] The spacing of the steel balls 17 is determined according to the length of the chute 18. To further explain, when the first steel ball 17 slides in and out from one end of the chute 18, the second steel ball 17 follows and enters the next chute 18, thereby completing the integrity of the overall rotation process.

[0036] Furthermore, a connecting rod 19 is fixedly installed at the hinge between the arc plate 10 and the receiving plate 9, and a double torsion spring 20 is sleeved on the outside of the connecting rod 19, and the two ends of the double torsion spring 20 are set against the side walls of the arc plate 10 and the receiving plate 9.

[0037] Preferably, both ends of the bottom of the receiving plate 9 are formed with limit flanges 21 , and the top of the arc plate 10 and the receiving plate 9 are hingedly fixed with a limit stop 22 .

[0038] When the limiting flange 21 contacts the limiting stop 22 , the corners of the limiting flange 21 are polished and rounded to avoid the limiting flange 21 from rotating in a too large range, thereby preventing the limiting flange 21 from causing interference and jamming.

[0039] Furthermore, a limiting plate 23 is fixedly installed on the outside of the front end of the center column 2, a side wall of the body 1 on which the support frame 4 is installed is provided with a plurality of placement grooves 24, and the support frame 4 is provided with a plurality of limiting grooves 25 that cooperate with the placement grooves 24.

[0040] The diameter of the limit plate 23 installed on the center column 2 should be slightly larger than the center column 2, but not too large, otherwise it will cause certain interference to the airflow at the front end of the body 1 during the operation of the aircraft engine, thereby affecting the performance of the aircraft engine.

[0041] Preferably, the side wall of the receiving plate 9 is provided with a mouth-shaped groove 26 for fixing the double torsion spring 20 , and the side wall of the arc plate 10 is provided with a straight groove 27 for the double torsion spring 20 to slide.

[0042] The purpose of opening the mouth-shaped groove 26 on the receiving plate 9 is to fix one end of the double torsion spring 20 to prevent the double torsion spring 20 from separating from the receiving plate 9 and the arc plate 10 during the deformation process. At the same time, a straight groove 27 is opened on the side wall of the arc plate 10, so that the double torsion spring 20 can slide easily in the straight groove 27 during the deformation process, thereby fitting the arc plate 10 more closely, thereby further limiting the arc plate 10.

[0043] Working principle: First, when the aircraft engine needs to be protected, the motor 12 is turned on, and the output end of the motor 12 drives the gear 13 to rotate. Since the ring gear 11 and the gear 13 are meshed and transmitted, the ring gear 11 rotates and slides inside the fixed frame 6 driven by the gear 13. Since a number of steel balls 17 are fixedly installed in the ring gear 11, when the ring gear 11 rotates, the steel balls 17 pass through the inclined grooves 18 in the sliding sleeve 8 at a uniform speed, so that the sliding sleeve 8 can rotate around the limit ring 7, and at the same time drive the receiving plate 9 and the arc plate 10 to rotate synchronously, and the double torsion spring 20 arranged at the hinge of the receiving plate 9 and the arc plate 10 generates elastic force on the receiving plate 9 and the arc plate 10 after deformation, so that the limit flange 21 in the receiving plate 9 can always press against the limit stop 22 , so that the receiving plate 9 and the arc plate 10 are kept in the same plane, until the side of the arc plate 10 hits the limit plate 23, the overall closed protection process can be completed; when the aircraft engine needs to be used, it is only necessary to reverse the motor 12, and again through the cooperation of the steel ball 17 and the inclined groove 18, the sliding sleeve 8 drives the receiving plate 9 and the arc plate 10 to rotate and move, until the surface of one side of the arc plate 10 hits the outer wall of the body 1, and then continue to operate the motor 12, so that the receiving plate 9 continues to rotate with the sliding sleeve 8, and the double torsion spring 20 is deformed due to the continued rotation of the receiving plate 9. The double torsion spring 20 generates a force that causes the arc plate 10 to hit the outer wall of the body 1, so as to achieve the effect of locking the arc plate 10, and the aircraft engine can be operated normally at this time.

[0044] Importantly, it should be noted that the construction and arrangement of the present application shown in a plurality of different exemplary embodiments are only exemplary. Although only a few embodiments are described in detail in this disclosure, it should be readily understood by those who refer to this disclosure that many modifications are possible (e.g., the size, scale, structure, shape and proportion of various elements, and parameter values ​​(e.g., temperature, pressure, etc.), installation arrangement, use of materials, color, directional changes, etc.) without substantially departing from the novel teachings and advantages of the subject matter described in the application. For example, the element shown as integrally formed can be composed of multiple parts or elements, the position of the element can be inverted or otherwise changed, and the nature or number or position of the discrete element can be changed or changed. Therefore, all such modifications are intended to be included in the scope of the present invention. The order or sequence of any process or method steps can be changed or reordered according to alternative embodiments. In the claims, any "device plus function" clause is intended to cover the structure of performing the function described herein, and is not only structurally equivalent but also equivalent structure. Without departing from the scope of the present invention, other replacements, modifications, changes and omissions can be made in the design, operating conditions and arrangement of the exemplary embodiments. Therefore, the invention is not limited to a specific embodiment, but extends to numerous modifications still falling within the scope of the appended claims.

[0045] Additionally, in order to provide a concise description of exemplary embodiments, all features of an actual embodiment (ie, those features that are not relevant to the best mode presently contemplated for carrying out the invention or those that are not relevant to implementing the invention) may not be described.

[0046] It will be appreciated that in the development of any actual implementation, as in any engineering or design project, numerous implementation-specific decisions may be made. Such a development effort may be complex and time-consuming, but will be a routine task of design, fabrication, and production for those of ordinary skill having the benefit of this disclosure without undue experimentation.

[0047] It should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention rather than to limit it. Although the present invention has been described in detail with reference to the preferred embodiments, those skilled in the art should understand that the technical solutions of the present invention may be modified or replaced by equivalents without departing from the spirit and scope of the technical solutions of the present invention, which should all be included in the scope of the claims of the present invention.

Claims

1. An aircraft engine protection device, comprising an engine body (1), Features: The machine body (1) is placed in a horizontal state, a center column (2) is fixedly mounted at the axis position of the machine body (1), a turbine blade group (3) is sleeved on the outside of the center column (2), and a protective component is fixedly mounted on one end of the outer side of the machine body (1); The protection component comprises a support frame (4) fixedly mounted on the outer side of the end of the machine body (1); the support frame (4) is fixedly connected to a fixed frame (6) via a plurality of connecting plates (5); a plurality of limiting rings (7) are fixedly connected to the outside of the fixing frame (6); a sliding sleeve (8) is slidably mounted on the outer side of each limiting ring (7); a receiving plate (9) is fixedly connected to the outer side of each sliding sleeve (8); an arc plate (10) is hingedly connected to one end of each receiving plate (9) away from the sliding sleeve (8); and a driving component is installed inside the fixing frame (6); The driving assembly comprises two gear rings (11) slidably mounted on the inner side of a fixing frame (6); a motor (12) is fixedly mounted on the side of the fixing frame (6); an output end of the motor (12) penetrates through the side wall of the fixing frame (6) and is sleeved with a gear (13); a sink groove (14) is provided at the fixing frame (6) where the motor (12) is mounted; a avoidance groove (15) is provided at the center of one of the connecting plates (5); and the gear (13) is mounted at the junction of the avoidance groove (15) and the sink groove (14); and the two gear rings (11) are meshed with the gear (13); A plurality of mounting grooves (16) are formed at the outermost portions of the two gear rings (11), and the mounting grooves (16) on the two gear rings (11) are symmetrically distributed. Steel balls (17) are installed in the mounting grooves (16). A plurality of inclined grooves (18) are formed on the inner side wall of the sliding sleeve (8), and the steel balls (17) roll in the inclined grooves (18); A connecting rod (19) is fixedly installed at the hinged joint between the arc plate (10) and the receiving plate (9), a double torsion spring (20) is sleeved on the outside of the connecting rod (19), and two ends of the double torsion spring (20) are arranged to press against the side wall of the arc plate (10) and the side wall of the receiving plate (9); When the aircraft engine needs to be protected, the motor (12) is turned on, and the output end of the motor (12) drives the gear (13) to rotate. Since the ring gear (11) and the gear (13) are meshed and transmitted, the ring gear (11) rotates and slides inside the fixed frame (6) driven by the gear (13). Since a plurality of steel balls (17) are fixedly installed inside the ring gear (11), when the ring gear (11) rotates, the steel balls (17) pass through the inclined grooves (18) in the sliding sleeve (8) at a uniform speed, so that the sliding sleeve 8 can rotate around the limit ring (7), and at the same time drive the receiving plate (9) and the arc plate (10) to rotate synchronously.

2. An aircraft engine protection device according to claim 1, Features: Both ends of the bottom of the receiving plate (9) are formed with limited flanges (21), and the top of the arc plate (10) is fixedly installed with a limited stop (22) at the hinged connection with the receiving plate (9).

3. An aircraft engine protection device according to claim 1, Features: A limiting plate (23) is fixedly installed on the outside of the front end of the center column (2), and a side wall of one end of the machine body (1) on which a support frame (4) is installed is provided with a plurality of placement grooves (24), and the support frame (4) is provided with a plurality of limiting grooves (25) that cooperate with the placement grooves (24).

4. An aircraft engine protection device according to claim 1, Features: The side wall of the receiving plate (9) is provided with a mouth-shaped groove (26) for fixing the double torsion spring (20), and the side wall of the arc plate (10) is provided with a straight groove (27) for the double torsion spring (20) to slide.

Images