阅读说明：本技术 一种超声速飞行器发动机防结霜保温方法 (Anti-frosting and heat-insulating method for supersonic aircraft engine ) 是由 孟美芳 于 2020-06-13 设计创作，主要内容包括：本发明公开了一种超声速飞行器发动机防结霜保温方法,涉及发动机保温技术领域,本发明包括以下步骤：步骤一：温度感应器对飞行器发动机外壳温度进行感应,操作人员根据感应温度作出判断；步骤二：基于步骤一,到达结霜温度的阈值时对发动机外壳进行加热；步骤三：若干一级导热管串联,将加热装置的热量传递给发动机外壳；步骤四：二级导热管进行电动推伸,与一级导热管的接触面积增加,提升导热效率；步骤五：将发动机尾端口排出的热量进行回收,于发动机外壳内侧导热内腔内形成热循环。本发明一种超声速飞行器发动机防结霜保温方法,其热传递结构具备可调节性,根据实际使用需求及使用情况,其二级导热管可进行独立配合使用,热传递效率更高。(The invention discloses a frosting prevention and heat preservation method for an engine of a supersonic aircraft, which relates to the technical field of engine heat preservation and comprises the following steps: the method comprises the following steps: the temperature sensor senses the temperature of the shell of the aircraft engine, and an operator makes a judgment according to the sensed temperature; step two: based on the step one, heating the engine shell when the threshold value of the frosting temperature is reached; step three: the first-stage heat conduction pipes are connected in series and transfer the heat of the heating device to the shell of the engine; step four: the second-stage heat conduction pipe is electrically pushed and extended, so that the contact area with the first-stage heat conduction pipe is increased, and the heat conduction efficiency is improved; step five: the heat discharged from the tail end of the engine is recovered, and heat circulation is formed in the heat conduction inner cavity on the inner side of the engine shell. According to the anti-frosting heat preservation method for the supersonic aircraft engine, the heat transfer structure has adjustability, the secondary heat conduction pipe can be independently matched for use according to actual use requirements and use conditions, and the heat transfer efficiency is higher.)

1. A supersonic aircraft engine anti-frosting heat preservation method is characterized by comprising the following steps:

the method comprises the following steps: the temperature sensor (17) senses the temperature of the shell of the aircraft engine, and an operator makes a judgment according to the sensed temperature;

step two: based on the step one, heating the engine shell when the threshold value of the frosting temperature is reached;

step three: a plurality of first-stage heat conduction pipes (9) are connected in series and transfer the heat of the heating device (6) to the shell of the engine;

step four: the second-stage heat conduction pipe (11) is electrically pushed and extended, so that the contact area with the first-stage heat conduction pipe (9) is increased, and the heat conduction efficiency is improved;

step five: recovering heat discharged from an engine tail port (12), and forming heat circulation in a heat conduction inner cavity (13) on the inner side of an engine shell;

step six: the temperature sensor (17) detects the temperature of the engine shell again, and the heating device (6) is closed when the temperature reaches the threshold value of the required temperature, so that the heat preservation effect on the aircraft engine is achieved.

2. The supersonic aircraft engine frost prevention and heat preservation method according to claim 1, wherein: including engine body (1), engine body (1) side surface is provided with mounting panel (2), be connected with a plurality of one-level heat pipe (9) between mounting panel (2) and engine body (1), a plurality of electric putter (10) of mounting panel (2) surface mounting, the output of a plurality of electric putter (10) all is connected with second grade heat pipe (11), the aperture of second grade heat pipe (11) is greater than one-level heat pipe (9), and one-level heat pipe (9) and second grade heat pipe (11) pull cooperation.

3. The supersonic aircraft engine frost prevention and heat preservation method according to claim 1 or 2, characterized in that: heating device (6) are still installed on mounting panel (2) one surface, heating device (6) one side surface is connected with wireless controller (8).

4. The supersonic aircraft engine frost prevention and heat preservation method according to claim 1 or 2, characterized in that: be connected with supporting shoe (7) between mounting panel (2) and engine body (1), mounting panel (2) two relative surfaces all are connected with a plurality of mounts (3), a plurality of mounts (3) one end all are connected with connecting piece (5), a plurality of brackets (4) are installed to engine body (1) week side, bracket (4) are connected with connecting piece (5).

5. The supersonic aircraft engine frost prevention and heat preservation method according to claim 1 or 2, characterized in that: the engine comprises an engine body (1), and is characterized in that an output end of the engine body (1) is provided with a tail end port (12), an inner baffle (14) is arranged on one surface of the inner side of the tail end port (12), the inner side of the engine body (1) is of a hollow structure, a heat conduction inner cavity (13) is arranged in the hollow structure of the inner side of the engine body (1), and the inner baffle (14) is matched with the heat conduction inner cavity (13) in position.

6. The supersonic aircraft engine frost prevention and heat preservation method according to claim 2, wherein: and heat conducting wires (15) are connected among the plurality of first-level heat conducting pipes (9).

7. The supersonic aircraft engine frost prevention and heat preservation method according to claim 1 or 2, characterized in that: the surface of the mounting plate (2) is provided with an insulating coating (16).

8. The supersonic aircraft engine frost prevention and insulation method according to claim 1, 2 or 3, wherein: the engine body (1) one side surface mounting has temperature-sensing ware (17), wireless controller (8) and temperature-sensing ware (17) electric connection.

Technical Field

The invention relates to the technical field of engine heat preservation, in particular to a frosting prevention and heat preservation method for an engine of a supersonic aircraft.

Background

The aircraft is an apparatus flying object which is manufactured by human beings, can fly off the ground, flies in space and is controlled by human beings to fly in the atmosphere or in the space (outer space), and the aircraft is divided into 3 types: the aircraft flying in the atmosphere, such as balloons, airships, airplanes and the like, fly by the static buoyancy of air or the aerodynamic force generated by the relative motion of air, and fly in space, such as artificial earth satellites, manned spacecrafts, space detectors, space airplanes and the like, obtain the necessary speed to enter space under the driving of a carrier rocket, and then make orbital motion similar to that of a celestial body by means of inertia, and the aircraft usually needs to be driven by an engine.

At present, a supersonic aircraft engine is lack of a certain protection structure when running in a low-temperature environment, the air temperature is continuously reduced along with the continuous increase of the rising height of the aircraft, the engine is prone to frosting and other phenomena, the normal running of the engine can be affected, and certain potential safety hazards are generated.

Disclosure of Invention

The invention mainly aims to provide a method for preventing frosting and insulating a supersonic aircraft engine, which can effectively solve the problems in the background technology by arranging a controllable and adjustable automatic heat transfer structure to heat the surface of the engine.

In order to achieve the purpose, the invention adopts the technical scheme that: a supersonic aircraft engine anti-frosting heat preservation method comprises the following steps:

the method comprises the following steps: the temperature sensor senses the temperature of the shell of the aircraft engine, and an operator makes a judgment according to the sensed temperature;

step two: based on the step one, heating the engine shell when the threshold value of the frosting temperature is reached;

step three: the first-stage heat conduction pipes are connected in series and transfer the heat of the heating device to the shell of the engine;

step four: the second-stage heat conduction pipe is electrically pushed and extended, so that the contact area with the first-stage heat conduction pipe is increased, and the heat conduction efficiency is improved;

step five: recovering heat discharged from the tail end of the engine, and forming heat circulation in a heat conduction inner cavity on the inner side of the engine shell;

step six: the temperature sensor detects the temperature of the engine shell again, and the heating device is closed when the temperature reaches the threshold value of the required temperature, so that the heat preservation effect on the aircraft engine is formed.

Preferably, including engine body, engine body side surface is provided with the mounting panel, be connected with a plurality of one-level heat pipes between mounting panel and the engine body, a mounting panel surface mounting has a plurality of electric putter, a plurality of electric putter's output all is connected with the second grade heat pipe, the aperture of second grade heat pipe is greater than the one-level heat pipe, and one-level heat pipe and the cooperation of second grade heat pipe pull.

Preferably, heating device is still installed on a surface of the mounting plate, a wireless controller is connected to one side surface of the heating device, and the wireless controller is utilized to electrically control the heating device, so that the operation is convenient and fast.

Preferably, be connected with the supporting shoe between mounting panel and the engine body, the two relative surfaces of mounting panel all are connected with a plurality of mounts, a plurality of mount one end all are connected with the connecting piece, a plurality of brackets are installed to engine body week side, the bracket is connected with the connecting piece, and this mounting panel forms detachable construction, is convenient for overhaul it and maintain.

Preferably, the output of engine body is provided with the tail port, the inboard surface mounting of tail port has interior baffle, engine body inboard is hollow structure, and is provided with the heat conduction inner chamber in the inboard hollow structure of engine body, interior baffle suits with heat conduction inner chamber position, and interior baffle blocks the hot gas of turning back to the engine derivation, goes to the heat conduction inner chamber.

Preferably, all be connected with the heat conduction silk between a plurality of one-level heat pipes, form the heat conduction series connection between the one-level heat pipe, the heat conduction area enlarges, generates heat more evenly.

Preferably, an insulating coating is arranged on one surface of the mounting plate, and the insulating coating specifically uses an alumina ceramic coating, so that the use safety is effectively improved.

Preferably, engine body one side surface mounting has the temperature-sensing ware, wireless controller and temperature-sensing ware electric connection, the temperature control effect is stronger, and operating personnel controls more accurately.

The invention has the following beneficial effects:

according to the invention, the heat transfer structure has adjustability, the secondary heat conduction pipes can be independently matched for use according to actual use requirements and use conditions, the environment response capability is stronger, and the heat transfer efficiency is higher when the secondary heat conduction pipes are synchronously used.

In the invention, hot gas exhausted from the tail end of the engine can be blocked and recovered, and heat conduction circulation is formed on the inner surface of the shell of the engine, so that the heat insulation effect of the engine is further improved.

Drawings

FIG. 1 is a schematic diagram of the overall structure of a supersonic aircraft engine anti-frosting and heat-preserving method of the present invention;

FIG. 2 is a schematic diagram of the position structure of a temperature sensor of the supersonic aircraft engine anti-frosting heat preservation method of the present invention;

FIG. 3 is a schematic side view of the anti-frosting and heat-insulating method for the supersonic aircraft engine according to the present invention;

FIG. 4 is a schematic diagram of a heat-conducting cavity position structure of a supersonic aircraft engine anti-frosting heat preservation method of the present invention;

FIG. 5 is a schematic top view of the anti-frosting and heat-insulating method for supersonic aircraft engine according to the present invention;

FIG. 6 is a schematic diagram of the position structure of an inner baffle of the supersonic aircraft engine anti-frosting and heat-preserving method of the present invention.

In the figure: 1. an engine body; 2. mounting a plate; 3. a fixed mount; 4. a bracket; 5. a connecting member; 6. a heating device; 7. a support block; 8. a wireless controller; 9. a first-stage heat conduction pipe; 10. an electric push rod; 11. a secondary heat conduction pipe; 12. a tail port; 13. a heat conducting inner cavity; 14. an inner baffle; 15. heat conducting wires; 16. an insulating coating; 17. a temperature sensor.

Detailed Description

In order to make the technical means, the creation characteristics, the achievement purposes and the effects of the invention easy to understand, the invention is further described with the specific embodiments.

In the description of the present invention, it should be noted that the terms "upper", "lower", "inner", "outer", "front", "rear", "both ends", "one end", "the other end", and the like indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience of description and simplicity of description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.

In the description of the present invention, it is to be noted that, unless otherwise explicitly specified or limited, the terms "mounted," "disposed," "connected," and the like are to be construed broadly, such as "connected," which may be fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.