CN112648109A

CN112648109A - Airframe external air entraining device based on cooling of jet pipe of hypersonic aircraft

Info

Publication number: CN112648109A
Application number: CN202011606716.8A
Authority: CN
Inventors: 王俊伟; 朱伟; 刘艳; 金凤新
Original assignee: Shenyang Aircraft Design and Research Institute Aviation Industry of China AVIC
Current assignee: Shenyang Aircraft Design and Research Institute Aviation Industry of China AVIC
Priority date: 2020-12-28
Filing date: 2020-12-28
Publication date: 2021-04-13
Anticipated expiration: 2040-12-28
Also published as: CN112648109B

Abstract

The application belongs to the technical field of hypersonic aircraft design, and particularly relates to an external airframe air entraining device based on hypersonic aircraft spray pipe cooling. The lower wall surface of the airplane body is provided with a lower web plate so as to form a cooling passage with the airplane body, the front end of the cooling passage is provided with an air inlet, the rear end of the cooling passage is an air outlet, an air inlet compression throat is arranged at the air inlet, the lower wall surface of the airplane body forms the upper wall surface of the air inlet compression throat, the lower web plate forms the lower wall surface of the air entraining device of the air inlet compression throat at the front end, the upper wall surface of the air outlet is the outer wall surface of a spray pipe, and the lower wall surface of the air outlet is formed by the rear end. The application provides an air entraining device outside organism can provide high pressure, low temperature gas through a compression pipeline that admits air, and the gas that introduces passes through the outer wall of high temperature spray tube, gives spray tube cooling protective spraying pipe on the one hand, and on the other hand gas produces high temperature gas through the heating and gets rid of, increases carminative mach number and then produces exhaust thrust.

Description

Airframe external air entraining device based on cooling of jet pipe of hypersonic aircraft

Technical Field

The application belongs to the technical field of hypersonic aircraft design, and particularly relates to an external airframe air entraining device based on hypersonic aircraft spray pipe cooling.

Background

The spray tube is the necessary exhaust system of aircraft advancing device, is the essential aircraft essential element of assurance engine normal work, simultaneously because the spray tube is connected at high temperature, highly compressed combustion chamber rear portion, its operational environment is very abominable, and repeatedly usable spray tube generally adopts superalloy and regenerative cooling mode cooling spray tube, and disposable spray tube generally adopts the carbon-carbon material of ablating the material. Due to the cost requirement and the repeated use requirement of the airplane, an active cooling mode is generally adopted, the temperature of the wall surface of the spray pipe is reduced, and the service life of the spray pipe is ensured. And along with the improvement of flying speed, the heat that the spray tube produced increases, and the heat that especially ramjet produced is bigger, and the cooling requirement to the spray tube is also higher and higher. The general active cooling comprises regenerative cooling and film cooling, wherein the regenerative cooling is to use fuel oil of an airplane as a cooling medium, the fuel oil reaches an inlet of a combustion chamber from the tail end of a spray pipe through a certain channel, and the fuel oil is combusted.

The film cooling comprises liquid film cooling and gas film cooling, wherein the gas film cooling generally comprises the steps of discharging gas with certain pressure and certain temperature through an opening on the inner wall surface of a spray pipe, and forming a thin low-temperature protective layer on the inner wall surface of the spray pipe; liquid film cooling is similar to gas film cooling, except that the cooling medium is changed from low-temperature gas to low-temperature liquid, and a protective liquid film with lower temperature is formed on the inner wall surface of the spray pipe. The air film cooling mode can well achieve the cooling effect, but the air film cooling needs certain pressure, air is required to be introduced from devices such as an air inlet channel and the like, the air introduction amount is difficult to ensure, and the performances of other parts are influenced; the use of liquid film cooling is limited by the need to carry additional cooling media, among other reasons.

Disclosure of Invention

In order to solve the technical problem, the application provides an outer bleed device of organism based on cooling of hypersonic aircraft spray tube is provided with down the web at the lower wall face of aircraft organism, down the web with be provided with the route between the aircraft organism, the route front end sets up to the air inlet, and the rear end is the gas vent, and air inlet department is provided with the compression throat that admits air, and the air inlet sets up in the rear of the intake duct of aircraft organism, and the wall is the cambered surface from the intake duct import department to the import department of the compression throat that admits air and extends to form the upper wall of the compression throat that admits air, and lower web is the wall under the bleed device of the compression throat that admits air at the front end down, and the upper wall of gas vent is the outer wall of spray tube, and the lower wall of gas vent comprises the rear end of lower web.

Preferably, the aircraft body is designed in one piece with the lower web.

Preferably, the lower wall of the bleed air device is tangential to the incoming flow direction of the cooling air.

Preferably, the part of the aircraft body extending from the air inlet channel to the lower web plate and the outer surface of the lower web plate are in streamline design.

Preferably, the exhaust port is provided as a convergent-type outlet in the flow direction of the gas stream.

Preferably, a passage between the intake compression throat and the exhaust port is provided as a diffuser pipe, and the diffuser pipe is provided as an arc-shaped structure protruding toward the interior of the aircraft body relative to the compression throat and the exhaust port.

The application provides an air entraining device outside organism can provide high pressure, low temperature gas through a compression pipeline that admits air, and the gas that introduces passes through the outer wall of high temperature spray tube, gives spray tube cooling protective spraying pipe on the one hand, and on the other hand gas produces high temperature gas through the heating and gets rid of, increases carminative mach number and then produces exhaust thrust.

Drawings

Fig. 1 is a front view of an external bleed air installation according to the invention based on hypersonic aircraft nozzle cooling.

Figure 2 is a schematic view of the bleed air arrangement of the embodiment of the invention shown in figure 1.

The method comprises the following steps of 1-an aircraft body, 2-wings, 3-spray pipes, 4-air inlet channels, 5-air inlets, 6-air outlets, 7-a lower wall surface of the aircraft body, 8-air inlet compression throats, 9-diffusion pipelines, 10-outer wall surfaces of the spray pipes, 11-a lower wall surface of a gas-entraining device and 12-a lower web plate.

Detailed Description

In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the accompanying drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all embodiments of the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application, and should not be construed as limiting the present application. All other embodiments obtained by a person of ordinary skill in the art without any inventive work based on the embodiments in the present application are within the scope of protection of the present application. Embodiments of the present application will be described in detail below with reference to the drawings.

The utility model aims at providing a through the outer bleed cooling high temperature spray tube scheme of organism, as shown in figure 1 and figure 2, the aircraft of this application mainly indicates hypersonic aircraft, is provided with wing 2 on aircraft organism 1, and the place ahead is provided with intake duct 4, and the rear is provided with spray tube 3, and whole organism is streamlined design, and this application is provided with web 12 down at the lower wall of aircraft organism 1, be provided with the route between web 12 down and the aircraft organism 1, the route front end sets up to air inlet 5, and the rear end is gas vent 6, and air inlet 5 department is provided with air intake compression throat 8, and air inlet 5 sets up the rear in the intake duct 4 of aircraft organism 1, and organism lower wall 7 of aircraft organism is the cambered surface from intake duct 4 import department to the department of air intake compression throat 8 and extends to the lower wall of the formation air intake compression throat 8, and lower web 12 forms air intake device lower wall 11 of air intake compression throat 8 at the front end, the upper wall surface of the exhaust port 6 is a nozzle outer wall surface 10, and the lower wall surface of the exhaust port 6 is formed by the rear end of a lower web 12.

The upper wall surface of the exhaust port 6 is the outer wall surface of the spray pipe, and besides the temperature reduction of the spray pipe, air with lower temperature flows to contact with the wall surface of the high-temperature spray pipe, so that the incoming air is heated, after the incoming air is heated, the Mach number of the exhaust gas is increased, and larger thrust can be generated, and the thrust can offset or be higher than the air resistance brought by the lower guide plate 12.

In some alternative embodiments, the aircraft body 1 is designed in one piece with said lower web 12.

In some alternative embodiments, the lower wall 11 of the bleed air device is tangential to the incoming flow direction of the cooling air.

In some alternative embodiments, the portion of the aircraft body 1 extending from the air inlet 4 to the lower web 12, and the outer surface of the lower web 12 are of streamlined design.

In some alternative embodiments, the exhaust port 6 is arranged as a convergent outlet in the direction of flow of the gas stream. In this embodiment, the outlet nozzle of the external bleed air device is made slightly convergent, on the one hand for discharging the air and on the other hand generating a small thrust gain to counteract the drag generated by the external device.

In some alternative embodiments, the passage between the intake compression throat 8 and the exhaust port 6 is provided as a diffuser pipe 9, and the diffuser pipe 9 is provided as an arc-shaped structure that protrudes toward the interior of the aircraft body with respect to the compression throat 8 and the exhaust port 6.

In this embodiment, the lower web 12 is connected to the lower wall 7 of the machine body, the inner portion is a circular arc convex surface for compressing the incoming air, the diffuser 9 is used for increasing the pressure for capturing the incoming air, and the high total pressure of the incoming air is beneficial to the air flowing on the profile of the nozzle. The lower web plate 12 and the front end of the machine body form an air inlet compression throat 8, the lower wall surface 11 of the air entraining device is connected with the wall surface of the machine body and tangent to the incoming flow direction for capturing the incoming flow air, and the air entraining device outside the machine body is a streamline device for reducing drag.

The application provides an air entraining device outside organism can provide high pressure, low temperature gas through a compression pipeline that admits air, and the gas that introduces passes through the outer wall of high temperature spray tube, gives spray tube cooling protective spraying pipe on the one hand, and on the other hand gas produces high temperature gas through the heating and gets rid of, increases carminative mach number and then produces exhaust thrust. The air entraining mode is suitable for Ma 0-Ma 4 airplanes. Taking the condition of Ma4 as an example, the atmospheric ambient temperature is low, the temperature of the air stream is 216.65K at 20Km, and the stopping temperature is 900K calculated according to the following formula.

And the wall surface temperature of the spray pipe is about 2300K at the moment, the temperature of the spray pipe can be greatly reduced after air entraining, the spray pipe is protected, and the aim of reusing the spray pipe is fulfilled.

The beneficial effect who adopts above scheme: 1) the cooling device of the air introduced from the outside of the engine body can cool the high-temperature spray pipe without being limited by fuel oil or other heat sinks; 2) the cooling device for the air introduced from the outside of the machine body can prolong the service life of the spray pipe, and air introduced by the air source is not influenced by the air introducing amount, so that the cooling device has a better application prospect; 3) the air cooled by the air-bleed cooling device outside the machine body is discharged through the contraction channel to generate additional thrust, so that the resistance of the air-bleed device can be offset or higher; 4) the cooling device for the air introduced from the outside of the machine body is simple in structure and easy to realize.

Having thus described the present application in connection with the preferred embodiments illustrated in the accompanying drawings, it will be understood by those skilled in the art that the scope of the present application is not limited to those specific embodiments, and that equivalent modifications or substitutions of related technical features may be made by those skilled in the art without departing from the principle of the present application, and those modifications or substitutions will fall within the scope of the present application.

Claims

1. An external bleed device based on hypersonic aircraft nozzle cooling is characterized in that a lower web plate (12) is arranged on the lower wall surface of an aircraft body (1), a passage is arranged between the lower web plate (12) and the aircraft body (1), the front end of the passage is provided with an air inlet (5), the rear end of the passage is provided with an exhaust port (6), an air inlet compression throat (8) is arranged at the air inlet (5), the air inlet (5) is arranged behind an air inlet channel (4) of the aircraft body (1), the lower wall surface (7) of the aircraft body extends backwards and downwards in a cambered surface manner from the inlet position of the air inlet channel (4) to the inlet position of the air inlet compression throat (8) and forms an upper wall surface of the air inlet compression throat (8), the lower web plate (12) forms a lower wall surface (11) of the bleed device of the air inlet compression throat (8) at the front end, and the upper wall surface of the exhaust port (6) is an external nozzle wall surface (, the lower wall surface of the exhaust port (6) is formed by the rear end of the lower web (12).

2. The extrabody bleed air device based on hypersonic aircraft nozzle cooling according to claim 1, characterized in that the aircraft body (1) is designed in one piece with the lower web (12).

3. The extrabody bleed air device based on jet pipe cooling of a hypersonic aircraft according to claim 1, characterized in that the lower wall surface (11) of the bleed air device is tangential to the incoming flow direction of the cooling air.

4. The airframe external bleed air device based on hypersonic aircraft nozzle cooling as claimed in claim 1, characterized in that the aircraft airframe (1) extends from the air intake duct (4) to the lower web (12), and the outer surface of the lower web (12) is of streamlined design.

5. The extrabody bleed air device based on jet pipe cooling of a hypersonic aircraft according to claim 1, characterized in that the exhaust opening (6) is arranged as a convergent outlet in the direction of flow of the air stream.

6. The extrabody bleed air device based on jet cooling of a hypersonic aircraft according to claim 1, characterized in that the passage between the inlet compression throat (8) and the exhaust port (6) is provided as a diffuser pipe (9), the diffuser pipe (9) being provided as an arc-shaped structure which is convex towards the interior of the aircraft body with respect to the compression throat (8) and the exhaust port (6).