CN114704384A - A method and structure for cooling a turbine engine compartment of a super aircraft - Google Patents

Abstract

The application provides a method for cooling a turbine engine nacelle of a hypersonic vehicle, comprising the following steps: and when the flying speed of the super aircraft is higher than the preset Mach number and the turbine engine is closed and is not providing flying power, the cooling air from the rear side of the ram air turbine is guided to the aircraft inlet channel to enter the engine compartment along the aircraft inlet channel, so that the heat of the engine compartment is taken away, the temperature of the engine compartment is controlled, and the engine compartment of the turbine engine is cooled. The ram air after the ram air turbine expansion cooling cools the turbine engine compartment in the pressure turbine power generation system, so that the accessory cables and the like in the turbine engine compartment can be guaranteed to have safe and reliable thermal environments, and the failure or damage of the high-temperature environment is avoided.

Description

A method and structure for cooling a turbine engine compartment of a super aircraft

technical field

The present application belongs to the technical field of aero-engine thermal managers, and in particular, relates to a method and structure for cooling a turbine engine compartment of a super aircraft.

Background technique

The engine compartment is the environmental space where the engine is installed on the aircraft. The temperature of the engine compartment will rise due to the heat release of the engine surface. Therefore, it is necessary to control the temperature in the engine compartment through cold air, so that the accessories and cables installed on the engine surface are safe. working environment, and at the same time, avoiding the engine transferring a lot of heat to the aircraft, a process known as engine compartment thermal management.

FIG. 1 is a schematic diagram showing the cooling of the engine compartment of a typical small and medium bypass ratio turbine engine. The engine compartment 12 is cooled by the inlet ram air in the air intake 11 of the aircraft, that is, most of the air at the inlet of the turbine engine 10 enters The compression and combustion in the engine flow channel do work, and a small part of the air enters the engine compartment to take away the heat of the engine compartment and control the temperature of the engine compartment.

However, with the increase of the speed of the aircraft, especially for the high-level aircraft with combined power (powered by the ramjet engine and the turbine engine), the turbine engine is turned off and the ramjet engine works at about Ma2.5 ~ Ma3.

Therefore, for the combined power super aircraft, the cooling of the nacelle of the turbine engine has the following problems:

1) The existing scheme of using the ram air at the engine inlet to cool the turbine engine compartment can only be used under the flight Mach number Ma3. When the super aircraft is flying at a high speed higher than Ma3, the temperature of the ram air exceeds 200 ℃, exceeding the turbine engine. The allowable temperature range of external accessories and cables, as shown in Figure 2;

2) When the high-speed aircraft flies at high speed, the temperature of the incoming flow reaches a very high level due to the ramming effect. Although the inlet of the turbine engine is closed at this time, it is inevitable that air leaks into the flow passage of the turbine engine. The components of the fan part are in high temperature, which seriously endangers its life and safety.

SUMMARY OF THE INVENTION

The purpose of the present application is to provide a method and structure for cooling a turbine engine compartment of a hyper-aircraft, so as to solve or alleviate at least one of the problems in the background art.

In one aspect, the present application provides a method for cooling a turbine nacelle of a hyper-aircraft, the method comprising:

The cooling air is led from the rear side of the ram air turbine of the ram turbine power generation system to the aircraft intake port of the turbine engine. When the flying speed of the super aircraft is greater than the predetermined Mach number and the turbine engine is turned off and the flight power is not provided, the air from the rear of the ram air turbine The cooling air from the side bleed air to the aircraft intake duct enters the engine compartment along the aircraft intake duct, thereby taking away the heat of the engine compartment, thereby controlling the temperature of the engine compartment and cooling the engine compartment of the turbine engine.

Further, the predetermined Mach number is Ma2.5˜Ma3.

Further, a control valve is provided on the cooling channel between the rear side of the ram air turbine and the air intake of the aircraft, and the circulation of the cooling channel is controlled by opening and closing the control valve.

In another aspect, the present application provides a cooling structure for a turbine engine compartment of a hyper-aircraft, the structure comprising:

a cooling channel, the cooling channel connects the rear side of the ram air turbine of the ram turbine power generation system with the aircraft air intake of the turbine engine, and is used for introducing the cooling airflow from the rear side of the ram air turbine of the ram turbine power generation system to the aircraft of the turbine engine Intake port, wherein, when the flying speed of the high-altitude aircraft is greater than the predetermined Mach number and the turbine engine is turned off and does not provide flight power, the cooling air bleed from the rear side of the ram air turbine to the aircraft intake port enters the engine along the aircraft intake port. The nacelle, thereby taking away the heat of the engine compartment, so as to control the temperature of the engine compartment and realize the cooling of the engine compartment of the turbine engine.

Further, the predetermined Mach number is Ma2.5˜Ma3.

Further, it also includes a control valve, the control valve is arranged in the cooling channel between the rear side of the ram air turbine and the air intake of the aircraft, and the circulation of the cooling channel is controlled by opening and closing the control valve.

The method and structure for cooling the turbine engine compartment of a super aircraft provided by the present application cool the turbine engine compartment through the ram air after the expansion and cooling of the ram air turbine in the pressure turbine power generation system. The atmospheric pressure is low, and the temperature of the ram air after expansion can be controlled to a very low level. This cold air enters the turbine engine compartment, which can ensure a safe and reliable thermal environment for accessory cables in the turbine engine compartment, avoiding failure or damage due to high temperature environment. .

Description of drawings

In order to more clearly illustrate the technical solutions provided by the present application, the accompanying drawings will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application.

FIG. 1 is a schematic diagram of engine compartment cooling in the prior art.

Figure 2 shows the change curve of the total temperature of the incoming flow with the flight Mach number.

FIG. 3 is a schematic diagram of the cooling scheme of the turbine engine compartment of the super aircraft in the application

Figure 4 shows the variation of the total pressure of the engine inflow with the flight Mach number.

Detailed ways

In order to make the implementation purpose, 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 order to solve the problem of cooling the engine compartment of the turbine engine when flying at high speeds above Ma3, the present application proposes a method and structure for cooling the turbine engine compartment of a hyper-aircraft, which are used to create a safe and reliable cooling system for the turbine engine fan, surface-mounted external accessories, cables, etc. working environment to avoid damage due to high temperature failure.

As shown in FIG. 3 , the method for cooling a turbine engine compartment of a high-speed aircraft provided by the present application is as follows: from the rear side of the ram air turbine 21 of the ram turbine power generation system (a part of the ram engine), a cooling gas is introduced to the aircraft air intake of the turbine engine 10. 11. The inlet cold air introduced into the aircraft air intake 11 enters the engine compartment 12 along the flow channel, and takes away the heat of the engine compartment 12, thereby controlling the temperature of the engine compartment 12 and cooling the engine compartment of the turbine engine;

Wherein, when the air is bleed from the rear side of the ram air turbine 21 to the air inlet 11 of the aircraft, the flying speed of the hyper-aircraft must be greater than a predetermined Mach number, and the predetermined Mach number is usually Ma2.5˜Ma3. At this Mach number, the turbine engine is turned off and is no longer providing flight power.

Preferably, a control valve (not shown) may be provided on the cooling passage between the rear side of the ram air turbine 21 and the air intake 11 of the aircraft, and the circulation of the cooling passage is controlled by opening and closing the control valve.

In the method of the present application, when the high-speed aircraft is used to fly at a high Mach number, the aircraft is at high altitude, the ambient pressure at the expansion outlet of the ram air turbine is very low, the pressure before and after the expansion turbine is very high, the ram air in the ram turbine power generation system has pressure energy, and the ram air The expansion work is performed in the ram air turbine 21, and the expansion process design can make the air after expansion reach a very low level, as shown in FIG. 4 .

The changes in ram air temperature and pressure during expansion correspond to:

T ₂ /T ₁ =1-η _t [1-(P ₂ /P ₁ ) ^(k-1)/k ]

Among them, η _t represents the efficiency of the expansion turbine, T and P represent the total temperature and total pressure parameters of the section, respectively, the angles 1 and 2 represent the inlet and outlet sections, respectively, and k is the coefficient.

The bleed air volume of the ramjet power generation system can be calculated from the heat balance of the turbine engine compartment.

On this basis, the present application also provides a cooling structure for a turbine engine compartment of a super aircraft, the structure comprising:

A cooling channel connecting the rear side of the ram air turbine 21 of the ram turbine power generation system with the aircraft intake port 11 of the turbine engine 10, the cooling channel can introduce the cooling bleed air after the expansion of the rear side of the ram air turbine 21 to the aircraft intake port. 11. The cooling bleed air enters the engine compartment 12 along the engine flow channel, and takes away the heat of the engine compartment 12, thereby controlling the temperature of the engine compartment 12 and cooling the engine compartment of the turbine engine;

Wherein, the cooling channel will bleed air from the rear side of the ram air turbine 21 to the aircraft intake port 11 only when the flying speed of the super aircraft is greater than a predetermined Mach number, and the predetermined Mach number is usually Ma2.5-Ma3. At this Mach number, the turbine engine is turned off and is no longer providing flight power.

The method and structure for cooling the turbine engine compartment of a super aircraft provided by the present application cool the turbine engine compartment through the ram air after the expansion and cooling of the ram air turbine in the pressure turbine power generation system. The atmospheric pressure is low, and the temperature of the ram air can be controlled to a very low level after expansion. This cold air enters the turbine engine compartment, which can ensure a safe and reliable thermal environment for the accessory cables in the turbine engine compartment, and avoid failure or damage due to high temperature environment. .

The above are only specific embodiments of the present application, but the protection scope of the present application is not limited thereto. Any person skilled in the art who is familiar with the technical field disclosed in the present application can easily think of changes or substitutions. All should be covered within the scope of protection of this application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.

Claims (6)

1. A method of cooling a nacelle of a turbine engine of a hypersonic vehicle, the method comprising:

and when the flying speed of the super aircraft is higher than the preset Mach number and the turbine engine is closed and is not providing flying power, the cooling air from the rear side of the ram air turbine is guided to the aircraft inlet channel to enter the engine compartment along the aircraft inlet channel, so that the heat of the engine compartment is taken away, the temperature of the engine compartment is controlled, and the engine compartment of the turbine engine is cooled.

2. The method of cooling a nacelle of a hypersonic vehicle turbine engine as defined in claim 1, wherein the predetermined mach number is from Ma2.5 to Ma 3.

3. A method for cooling a nacelle of a hypersonic vehicle turbine engine as claimed in claim 1, characterized in that a control valve is provided in the cooling channel between the rear side of the ram air turbine and the aircraft air intake, the flow through the cooling channel being controlled by opening and closing the control valve.

4. A high altitude aircraft turbine engine nacelle cooling structure, the structure comprising:

and the cooling channel is communicated with the ram air turbine rear side of the ram turbine power generation system and the turbine engine air inlet channel and is used for introducing cooling air flow from the ram air turbine rear side of the ram turbine power generation system into the turbine engine air inlet channel, wherein when the flying speed of the super aircraft is higher than the preset Mach number and the turbine engine is closed and is not providing flying power, the cooling air from the ram air turbine rear side is guided to the turbine engine inlet channel and enters the engine cabin along the aircraft air inlet channel, so that the heat of the engine cabin is taken away, the temperature of the engine cabin is controlled, and the engine cabin of the turbine engine is cooled.

5. The super aircraft turbine engine nacelle cooling structure as claimed in claim 4, wherein the predetermined Mach number is Ma2.5 to Ma 3.

6. The hypersonic aircraft turbine engine nacelle cooling square of claim 4, further comprising a control valve disposed in the cooling passage between the ram air turbine aft side and the aircraft air intake, flow through the cooling passage being controlled by opening and closing of the control valve.

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