A new structure of a cold turbine blade for spraying combustion oil

Abstract

The application discloses a new structure of a cold turbine blade for jet combustion oil, which comprises a hollow turbine guide vane, a heat-conducting copper plate, a cooling channel, a needle valve, an oil collecting cavity and an oil injection hole, wherein the hollow turbine guide vane is in a hollow columnar shape, the heat-conducting copper plate is embedded in the hollow cavity of the hollow turbine guide vane, the cooling channel is embedded in the heat-conducting copper plate and winds from the trailing edge of the blade to the leading edge of the blade, the oil collecting cavity is embedded in the blade and located at the leading edge of the blade, the oil collecting cavity extends out of the blade through the oil injection hole, the oil inlet of the cooling channel is arranged close to the trailing edge of the blade, the oil outlet is arranged close to the leading edge of the blade, and the oil outlet is communicated with the oil collecting cavity through the needle valve. The cooling structure of the hollow embedding type cooperated with the heat-conducting copper plate can guarantee the uniformity of the blade wall surface temperature, reduce the thermal stress of the blade, aviation oil is heated and warmed up in the cooling channel, becomes superheated vapor state after pressure adjustment of the needle valve, and is more beneficial to full combustion, improves the combustion efficiency, and further improves the engine thrust.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of turbine blades of aviation gas turbine engine, and particularly relates to a new structure of oil-cooled turbine blade with jet combustion. BACKGROUND

[0002] The improvement of the performance of the aviation engine, the realization of higher thrust-to-weight ratio, the reduction of fuel consumption and the improvement of stability are mainly realized by increasing the turbine inlet temperature and improving the inlet compression ratio on this basis from the perspective of thermodynamics, and the core is to increase the turbine inlet temperature T3*. Although the turbine inlet temperature of the aviation engine is high, after the gas passes through the high-pressure turbine to do work, the temperature of the gas is greatly reduced, and the work capacity is also reduced, and under the current material technology and cooling technology, the temperature resistance of the low-pressure turbine blade has a large margin compared with the thermal environment it faces. If the gas can be supplemented with oil and combusted before the low-pressure turbine, the temperature of the low-pressure turbine inlet gas can be increased, and the work capacity of the gas can be increased, which can bring considerable thrust gain to the engine. This technology is called inter-stage combustion technology, and is one of the research hotspots in the field of aviation engines in recent years. The turbine inter-stage combustion chamber designed with this concept shows that the engine unit thrust can be increased by 20-50% under ideal cycle conditions.

[0003] Under high Mach number flight, the aviation fuel carried by the aircraft is the ideal cold source available for the engine. Directly using aviation fuel to cool the turbine blade can fully utilize the heat sink of the aviation fuel, and at the same time of cooling the turbine blade, the fuel absorbs heat and increases in temperature, which is more conducive to the combustion of the fuel and can improve the energy utilization. Combining the inter-stage combustion technology with the oil-cooled turbine blade technology, first, the aviation fuel is used to cool the low-pressure turbine guide vane, and then the fuel is injected into the main flow gas from the turbine blade to participate in combustion after being heated, which can improve the work capacity of the gas and increase the engine thrust under the premise of ensuring the safe operation of the turbine blade.

[0004] Based on the above advantages, the present application proposes a new structure of oil-cooled turbine blade with jet combustion, which combines the inter-stage combustion technology with the oil-cooled turbine technology to solve the problems of turbine blade thermal protection and engine performance improvement under high Mach number. SUMMARY

[0005] The technical problem to be solved by the present application is to overcome the shortcomings of the prior art and provide a new structure of oil-cooled turbine blade with jet combustion to realize the organic combination of the inter-stage combustion technology and the oil-cooled turbine technology, and to improve the engine thrust under the premise of meeting the cooling demand of the turbine blade.

[0006] To solve the above technical problems, the basic idea of the technical solution of the present application is:

[0007] The novel structure of the oil injection combustion oil cooling turbine blade comprises a hollow turbine guide vane, a heat-conducting copper plate, a cooling channel, a needle valve, an oil collecting cavity and an oil injection hole.

[0008] Further, the heat-conducting copper plate is in an arc shape and is attached to the tail edge of the blade, and the remaining part is kept in a gap with the inner wall of the hollow cavity of the blade.

[0009] Further, the thickness of the heat-conducting copper plate near the middle front part of the blade is 3.6 mm, and the gap with the inner wall of the blade is 0.2 mm.

[0010] Further, the cross section of the cooling channel is circular with an inner diameter of 1 mm, and the cooling channel is in a serpentine shape and is distributed along the arc surface of the heat-conducting copper plate.

[0011] Further, the oil collecting cavity is in a cylindrical shape with an inner diameter of 3 mm.

[0012] Further, the oil injection hole is five in number and is arranged at the leading edge of the blade, and the diameter of the oil injection hole is 0.4 mm.

[0013] Further, the central axis of the oil injection hole is at an angle of 135° with the direction of the main flow of the gas.

[0014] After the above technical solution is adopted, the present application has the following beneficial effects compared with the prior art.

[0015] The present application combines the inter-stage combustion technology and the oil cooling turbine blade technology, uses the aviation fuel of the aircraft as the cooling medium to cool the low-pressure turbine guide vane, can make up for the lack of cold source under high Mach number flight, expand the flight envelope of the aero-engine, and meet the cooling demand of the hot end components of the engine under more extreme working conditions; through the hollow nested cooling structure and the heat-conducting copper plate, the wall surface temperature of the turbine blade can be kept in a relatively uniform state, and the thermal stress of the blade can be reduced; the aviation fuel is heated and warmed up in the cooling channel, becomes superheated vapor after the pressure is adjusted by the needle valve, and is more conducive to its full combustion; the oil is injected from the leading edge of the blade into the main flow of the gas to participate in combustion, which can improve the temperature of the low-pressure turbine pre-combustion gas, improve the gas power, and further improve the engine thrust.

[0016] The specific embodiments of the present application will be further described in detail below with reference to the accompanying drawings. BRIEF DESCRIPTION OF DRAWINGS

[0017] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate exemplary embodiments of the application and together with the description, serve to explain the application. Obviously, the drawings in the following description are only some embodiments of the application and based on these drawings, other drawings can be obtained by those skilled in the art without any creative effort. In the drawings:

[0018] Fig. 1 The schematic diagram of the overall structure of an embodiment of the application is shown in the figure.

[0019] Fig. 2 The perspective view of the structure of an embodiment of the application is shown in the figure.

[0020] Fig. 3 The schematic diagram of the oil injection combustion of the application is shown in the figure.

[0021] In the figure: 1 - hollow turbine guide vane, 2 - heat-conducting copper plate, 3 - cooling channel, 4 - needle valve, 5 - oil collection cavity, 6 - oil injection hole.

[0022] It should be noted that the drawings and the written description are not intended to limit the scope of the inventive concept in any way, but to illustrate the inventive concept to those skilled in the art by referring to specific embodiments. DETAILED DESCRIPTION

[0023] In order to make the purpose, technical scheme and advantages of the embodiments of the application clearer, the technical scheme in the embodiments will be described clearly and completely below in combination with the drawings of the embodiments of the application. The following embodiments are used to illustrate the application, but not to limit the scope of the application.

[0024] In the description of the application, it should be noted that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "inner", "outer" and the like indicate the orientation or positional relationship based on the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the application and simplifying the description, and therefore cannot be understood as indicating or implying that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as limiting the application.

[0025] In the description of the application, it should be noted that unless otherwise explicitly specified and limited, the terms "mounting", "connection", "connection" should be understood broadly, for example, it can be fixed connection, or detachable connection, or integrally connected; it can be mechanical connection, or electrical connection; it can be directly connected, or indirectly connected through an intermediate medium. For those skilled in the art, the specific meaning of the above terms in the application can be understood according to the specific circumstances.

[0026] Embodiment one

[0027] AsFigs. 1-3 As shown, the new structure of the oil-injection cooling turbine blade of the present embodiment comprises a hollow turbine guide vane 1, a heat-conducting copper plate 2, a cooling channel 3, a needle valve 4, an oil collecting cavity 5 and an oil injection hole 6. The hollow turbine guide vane 1 is in a hollow columnar shape, the heat-conducting copper plate 2 is embedded in the hollow cavity of the hollow turbine guide vane 1, the cooling channel 3 is embedded in the heat-conducting copper plate 2 and extends from the trailing edge of the blade to the leading edge of the blade, the oil collecting cavity 5 is embedded in the blade and located at the leading edge of the blade, the oil collecting cavity 5 extends out of the blade through the oil injection hole 6, the oil inlet of the cooling channel 3 is located close to the trailing edge of the blade, the oil outlet is located close to the leading edge of the blade, and the oil outlet is communicated with the oil collecting cavity 5 through the needle valve 4.

[0028] Specifically, the hollow turbine guide vane 1 is a low-pressure turbine guide vane, adopts a layered nested cooling structure, and the heat-conducting copper plate 2 is inserted and embedded in the hollow cavity. The heat-conducting copper plate 2 is in an arc-shaped strip shape, the front part has a thickness of 3.6 mm and has a gap of 0.2 mm with the inside of the blade, and the tail part is designed to be conformal to the trailing edge of the blade, that is, the tail part is conformal to the trailing edge of the blade, as shown in the accompanying drawing. The cooling channel 3 is arranged in the heat-conducting copper plate 2, the cooling channel 3 is in a serpentine shape and is distributed along the arc surface of the heat-conducting copper plate, and the cross section of the cooling channel 3 is circular with an inner diameter of 1 mm.

[0029] The oil collecting cavity 5 is arranged in the leading edge of the blade, the oil collecting cavity 5 is in a cylindrical shape with an inner diameter of 3 mm and a height not exceeding the height of the blade, and the oil collecting cavity 5 is communicated with the external main stream of the gas through the oil injection hole 6; the oil injection hole 6 is arranged at the leading edge of the blade, the number of the oil injection hole 6 is five, the diameter of the oil injection hole 6 is 0.4 mm, and the central axis of the oil injection hole 6 is at an angle of 135° with the injection direction of the main stream of the gas, as shown in the accompanying drawing.

[0030] The oil inlet of the cooling channel 3 is close to the trailing edge of the blade, the oil outlet is close to the leading edge of the blade, and the oil outlet is communicated with the oil collecting cavity 5 through the needle valve 4, that is, the downstream of the needle valve 4 is connected to the oil collecting cavity 5 in the blade and is used for adjusting the injection pressure of the fuel. In the oil supply system of the engine, the fuel pressure is usually greater than the critical pressure, so if the fuel with such high pressure is directly injected through the oil injection hole, the injection distance will be too far, which is not conducive to combustion. Therefore, the aviation fuel is heated and warmed up in the cooling channel, then the pressure is reduced to the ideal injection pressure through the needle valve, and then enters the oil collecting cavity, and finally is injected out of the oil injection hole into the main stream of the gas to participate in combustion.

[0031] The aviation fuel is pressurized to supercritical pressure by a fuel pump and enters the blade cooling channel, the fuel flows in the serpentine cooling channel, absorbs the heat of the heat-conducting copper plate 2 through convection heat exchange, and fully cools the heat-conducting copper plate 2, and the heat-conducting copper plate cools and cools the turbine blade through radiation and heat conduction, etc., to ensure its safe operation, and because the heat-conducting copper plate 2 has high heat conductivity, the overall temperature of the heat-conducting copper plate is relatively close, and the blade temperature uniformity can be improved. The heated fuel is depressurized to the injection pressure by the needle valve 4, at this time the fuel temperature is high, the pressure is low, it is in a superheated vapor state, which is beneficial to full combustion, and the depressurized fuel enters the oil collecting cavity 5, and finally is injected into the main flow gas through the oil injection hole 6 opened on the oil collecting cavity 5. Here, the fuel injection direction and the main flow gas direction form an angle of 135°, and the fuel is in a superheated vapor state, which can enhance the mixing of the superheated vapor state fuel and the main flow gas, and is beneficial to the full combustion of the fuel, so as to improve the combustion efficiency.

[0032] The present application can meet the turbine guide vane cooling demand, achieve the effect of turbine inter-stage oil injection afterburning, improve the main flow gas temperature and improve its work capacity, through the use of the hollow nested type new cooling structure, the blade wall temperature uniformity can be guaranteed, and the blade thermal stress can be reduced, and through the use of the needle valve to depressurize the fuel, the fuel injection distance can be shortened, the fuel is changed into a superheated vapor state, which is beneficial to its combustion.

[0033] The above only describes the preferred embodiments of the present application, and does not limit the present application in any form, although the present application has been disclosed as above, however, it is not intended to limit the present application, any person skilled in the art can make some changes or modifications to the above-mentioned technical content without departing from the scope of the technical scheme of the present application, as long as it does not depart from the content of the technical scheme of the present application, any simple modification, equivalent change and modification of the above embodiments according to the technical essence of the present application, all still belong to the scope of the present application.

Claims

1. A new structure of a cold turbine blade for injection combustion oil, characterized by: The hollow turbine guide vane (1), heat-conducting copper plate (2), cooling channel (3), needle valve (4), oil collecting cavity (5) and oil injection hole (6) are included. The hollow turbine guide vane (1) is hollow columnar. The heat-conducting copper plate (2) is embedded in the hollow cavity of the hollow turbine guide vane (1). The heat-conducting copper plate (2) is arc-shaped and strip-shaped. The tail part is attached to the tail edge of the blade, and the remaining part keeps a gap with the inner wall of the blade hollow cavity. The cooling channel (3) is embedded in the heat-conducting copper plate (2) and winds from the tail edge of the blade to the leading edge of the blade. The oil collecting cavity (5) is embedded in the blade and located at the leading edge of the blade. The oil collecting cavity (5) extends out of the blade through the oil injection hole (6). The oil inlet of the cooling channel (3) is arranged close to the tail edge of the blade, and the oil outlet is arranged close to the leading edge of the blade. The oil outlet is communicated with the oil collecting cavity (5) through the needle valve (4).

2. A new structure of a cold turbine blade of a fuel injection combustion according to claim 1, characterized in that: The thickness of the heat-conducting copper plate (2) close to the middle front part of the leading edge of the blade is 3.6 mm, and the gap with the inner wall of the blade is 0.2 mm.

3. A new structure of a cold turbine blade of a spray combustion oil according to claim 1, characterized in that: The cross section of the cooling channel (3) is circular, the inner diameter is 1 mm, and the cooling channel (3) is serpentine and distributed along the arc surface of the heat-conducting copper plate (2).

4. A new structure of a cold turbine blade of a spray combustion oil according to claim 1, characterized in that: The oil collecting cavity (5) is cylindrical, and the inner diameter is 3 mm.

5. A new structure of a cold turbine blade of a fuel injection combustion according to claim 1, characterized in that: The oil injection hole (6) is five, arranged at the leading edge of the blade, and the diameter of the oil injection hole (6) is 0.4 mm.

6. A new structure of a cold turbine blade of a jet combustion oil according to claim 5, characterized in that: The central axis direction of the oil injection hole (6) is 135° with the direction of the main flow of the gas.

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