Method of and turbojet engine for creating reactive thrust

FIELD: turbojet engines.

SUBSTANCE: proposed method of creating reactive thrust in turbojet engine provided with compressor connected with turbine is implemented by preliminary compression of air delivered together with fuel into combustion chamber. Gas received at combustion of fuel and air mixture is used to drive turbine. Additional fuel is combustion in second combustion chamber installed after turbine. Gas formed in combustion chambers is directed to nozzle to create reactive thrust. Ring-shaped flow of gas coming out of turbine is formed after turbine uniformly over circumference. Direction of movement of said gas flow is changed by directing it to engine axis line into second combustion chamber after turbine. Radial concentric flows of gas are formed which collide in center of second combustion chamber with relative braking and conversion of kinetic energy of gas into heating and compressing. Additional fuel is combustion in said higher gas compression area. Gas with sufficient amount of oxygen is delivered into second combustion chamber for combustion of additional fuel.

EFFECT: increased reactive thrust.

4 cl, 1 dwg

 

The invention relates to jet propulsion installations and is intended for use on aircraft.

There is a method to create thrust used in ramjet engines (RAMJET), in which the compression of the air and produce high-speed pressure during the flight of the aircraft, compressed air direct flow fed to the combustion chamber, and combustion products are directed into the nozzle (see Polytechnical dictionary edited Awesonme. M., Soviet encyclopedia. - 1980, pages 420-421) (1).

The disadvantage of this way of generating thrust using a RAMJET is the possibility of its implementation only when flight speed, 2-3 times the speed of sound.

The closest set of features of the claimed invention is a method to create thrust used in turbojet engines (turbojet), in which the pre-compression of the air supplied to the combustion chamber, is made by means associated with the turbine compressor, and use the opportunity transient increases engine power by burning additional fuel in the afterburner to direct the flow of air passing through the engine (see (1), pp. 544-545).

The disadvantage of this method is to create thrust is the difficulty of ensuring the sufficiency of the internal largest pre-compression ratio supplied to the combustion air, within the motor has a straight character movement is predominantly in the axial direction. While multiple, complicated structure and a large mass of the applied axial compressor mainly related to difficulties compression free flow of air in the direction of its flow movement.

Known ramjet engine (RAMJET), comprising a housing in the form of profiled tubes, front part of which is made in the form of a cone, contributing to the compression of the incoming air, and the rear part of the tube is in the form of a nozzle, in the middle part of the tube is placed in the combustion chamber (see (1), pp. 420-421).

The disadvantage of RAMJET is that efficient operation is possible only when the speed is 2-3 times the speed of sound, and that for an aircraft with RAMJET need additional starting device.

Closest to the claimed invention by a combination of traits is a turbojet engine (turbojet), comprising a housing with an inlet, a compressor, a turbine, a combustion chamber disposed between the compressor and turbine, an additional afterburner combustion chamber with intermittent mode of operation and the nozzle (see(1), pp. 544-545).

The lack of a turbojet engine is a complicated device and a large mass primarily in connection with the necessity to the Yu of multistage compressor and turbine high power to achieve the necessary amount of pre-compression of air, supplied to the combustion chamber.

The present invention by way of create thrust and device for its implementation in the form of a two-stage turbojet engine allows to obtain the technical result consists in the simplification of the device and reducing the mass of the engine by reducing levels of compression in the compressor and reducing the power of the turbine at normal for existing engines, the amount of compression of the air or increase the compression ratio of the used air without any considerable variation in the complexity and weight of the currently used engines. This meant that increasing the compression ratio of the used air increases generated by the thrust without increasing fuel consumption.

This technical result on the way to create thrust is achieved by the turbojet engine use is associated with turbine compressor, which produce pre-compression of air, which is fed to the combustion chamber, resulting from the combustion of fuel gas is used to drive a turbine. In addition, by exploiting the combustion of additional fuel in the second combustion chamber installed in the turbine. Produced in combustion chambers of gas is directed into the nozzle and make the jet thrust. With the according to the invention, for turbine form a uniform around the circumference of the annular flow released from the gas turbine, the direction of which change and direct it to the line axis of the engine is placed behind the second turbine combustor. This creates a counter-radial concentric gas streams that pit in the center of the second combustion chamber with their mutual inhibition and the transition of the kinetic energy of the gas in the heating and compression in this area increased gas compression exercised by the combustion of additional fuel. Due to this further increase the energy saturation and compression of gas, which, with a corresponding increase in the rate for the above reasons, is sent to the nozzle and creates increased the amount of jet thrust. While the second combustion chamber serves gas with a sufficient amount of oxygen for combustion of additional fuel.

Two-stage turbofan engine includes a housing with an inlet, a compressor, a first combustion chamber, a turbine located behind the turbine, a second combustion chamber and nozzle, characterized in that the turbine has a guide liner having a bottom facing the outer surface toward the nozzle, and the side surface symmetric to the axis of the engine. The lateral guide surface of the liner together with the outside from siteline its circular shell form for turbine annular channel for the passage of gas from the first combustion chamber. The end of the annular channel is placed connected with the case of the symmetric axis line engine concave surface, an internal cavity which together with the bottom guide liner to form a second combustion chamber, which in the middle part of the concave surface communicates with the nozzle. On its outer circumference, the second combustion chamber is communicated with the annular channel. The curvature of the inner part of the concave surface is configured to implement a uniform deflection of the gas flow from the annular channel in concentric radial directions toward the Central part of the second combustion chamber with the possibility of mutual inhibition coming from the annular channel from all sides of a circle concentric opposing radial gas flows, ensuring its compression through the conversion of kinetic energy of the gas in the collision in the Central part of the second combustion chamber. Within the second combustion chamber facing the nozzle with spray fuel mainly in the zone of maximum compression of gas in its Central part. The second combustion chamber is made with regard mainly to its constant operation during the operation of the two stage turbojet engine, two-step which is determined by sequentially placed PE is howling and second combustion chambers.

In the two-stage turbine engine between the annular channel and the second combustion chamber established the guiding surfaces in the form of one or more profiled rims, symmetrically covering the second combustion chamber and enabling more accurate concentric direction radial flow of gas from the annular channel in the Central part of the second combustion chamber.

The bottom of the guide pad has a large diameter and area as compared with the diameter and cross-sectional area of the nozzle at the outlet of the second combustion chamber.

In the accompanying drawing in cross section along the axial frontal plane shown in the General implementation of the method to create thrust for example, the corresponding device in the form of a two-stage turbine engine. The arrows in the drawing show the direction of movement of the air and formed from the combustion of fuel gas.

Two-stage turbofan engine includes a housing 1 with an inlet 2, a compressor 3, the first combustion chamber 4, a turbine 5, located outside of the turbine, a second combustion chamber 6 and the nozzle 7. For turbine 5 has a guide insert 8 having a bottom 9, facing the outer surface toward the nozzle 7, and the side surface 10, the symmetric axis line o-O of the engine. The side surface of the guide pad is 8 together with external with respect to its circular shell 11 is formed over the turbine annular channel 12 for the passage of gas from the first combustion chamber 4. The end of the annular channel 12 is placed connected with the housing 1 symmetric axis line o-O of the engine concave surface 13, an internal cavity which together with the bottom 9 of the guiding liner 8 form a second combustion chamber 6, which in the middle part of the concave surface 13 is communicated with the nozzle 7. On its outer circumference, the second combustion chamber 6 is communicated with the annular channel 12. The curvature of the inner part of the concave surface 13 is arranged to implement a uniform deflection of the gas flow from the annular channel 12 in concentric radial directions toward the Central part of the second combustion chamber 6 with the possibility of mutual inhibition coming from the annular channel 12 evenly on all sides of a circle concentric opposing radial gas flows, ensuring its compression through the conversion of kinetic energy of the gas in the collision in the Central part of the second combustion chamber 6. Within the second combustion chamber facing the nozzle 14 with the possibility of spraying fuel mainly in the zone of maximum compression of gas in its Central part. The second combustion chamber 6 is made with regard mainly to its constant operation during the operation of the two stage turbojet engine, two-step which determines the tsya consistently placed first 4 and second 6 combustion chambers.

Between the annular channel 12 and the second combustion chamber 6 is set by the guide surface 15 in the form of one or more profiled rims, symmetrically covering the second combustion chamber and enabling more accurate concentric direction radial flow of gas from the annular channel in the Central part of the second combustion chamber.

The bottom 9 of the guiding insert 8 has a large diameter and area as compared with the diameter and cross-sectional area of the nozzle 7 at the outlet of the second combustion chamber 6.

Shows the device in the form of a two-stage turbine engine carries a suggested way to create thrust in the following way.

The two-stage compressor 3 turbojet engine pre-compression of the air which enters the first combustion chamber 4, which simultaneously serves fuel and burn it. Received in the first chamber of the combustion gas is directed into the turbine 5 and bring it together with the compressor 3 into rotation. From the turbine the gas is directed into the annular channel 12, where it is while driving evenly covers all sides of the lateral surface of the guide liner 8 and enters the concave surface 13 and profiled rims of the guide surfaces 15, which together alter the direction of movement of gas from kolicevo the channel 12 in the radial direction towards the Central part of the second combustion chamber 6, where is the mutual inhibition of the incoming concentric radial colliding gas streams. When this occurs, additional compression of the gas in connection with the transition of the kinetic energy of the flow of gas into thermal energy with a corresponding increase its compression ratio. The greatest compression of the gas is carried out in the Central part of the second combustion chamber 6, where from the nozzle 14 to produce atomization of the fuel, its combustion with maximum temperature rise and establishment of maximum compression and pressure of the combustion products of the surrounding surface of the second combustion chamber 6. This provides an increase in the rate of gas flow from the nozzle 7 with a corresponding increase in thrust.

For efficient operation of the engine in the composition of the gas flowing into the second combustion chamber 6,contains a sufficient amount of oxygen for complete and reliable combustion of the sprayed from the nozzle 14 of the fuel.

To the fullest use of high pressure gas on the surfaces of the second combustion chamber 6 to create thrust the bottom of the guide 9 of the insert 8 has a large area compared to the cross-sectional area of the beginning of the nozzle 7 when the output from the second combustion chamber 6.

The proposed two-stage turbofan engine provides increased thrust in and is used currently known TRD or enables simplification of the device in comparison with the known TRD without reducing the magnitude of the thrust.

1. The way to create thrust, in which the turbojet engine use is associated with turbine compressor, which produce pre-compression of air, which serves at the same time with fuel in the combustion chamber produced by burning the fuel gas is used to drive a turbine, in addition to use the opportunity of burning additional fuel in the second combustion chamber is installed at the turbine produced in combustion chambers of gas is directed into the nozzle and make the jet thrust, characterized in that the turbine form a uniform around the circumference of the annular flow released from the gas turbine, the direction of which change and send it to the axis of the engine placed behind the turbine, a second combustion chamber, this creates a counter-radial concentric gas streams that pit in the center of the second combustion chamber with their mutual inhibition and the transition of the kinetic energy of the gas in the heating and compression in this area increased gas compression exercised by the combustion of additional fuel, this increases the energy saturation and compression of gas, which, with a corresponding increase in the rate for the above reasons, is sent to the nozzle and creates increased the value of the torque rods is, while the second combustion chamber serves gas with a sufficient amount of oxygen for combustion of additional fuel.

2. Turbojet engine, comprising a housing with an inlet, a compressor, a first combustion chamber, a turbine located behind the turbine, a second combustion chamber and nozzle, characterized in that the turbine has a guide liner having a bottom facing the outer surface toward the nozzle, and the side surface symmetric to the axis of the engine, the side surface of the guide liner with external with respect to its circular shell form for turbine annular channel for the passage of gas from the first combustion chamber, the end of the annular channel is placed connected with the case of the symmetric axis line engine concave surface, an internal cavity which together with the bottom guide liner to form a second combustion chamber, which in the middle part of the concave surface communicates with the nozzle on its outer circumference, the second combustion chamber is communicated with the annular channel, the curvature of the inner part of the concave surface is configured to implement a uniform deflection of the gas flow from the annular channel in concentric radial directions toward the Central part of the second combustion chamber with the possibility vzaimnoj the braking coming from the annular channel evenly on all sides of a circle concentric opposing radial gas flows, providing a compression through the conversion of kinetic energy of the gas in the collision in the Central part of the second combustion chamber, within the second combustion chamber facing the nozzle with spray fuel mainly in the zone of maximum compression of gas in its Central part, while the second combustion chamber is made with regard predominantly continuous operation during operation of the engine, two-step which is determined by sequentially placed continuously operating the first and second combustion chambers.

3. Turbojet engine, characterized in that between the annular channel and the second combustion chamber established the guiding surfaces in the form of one or more profiled rims, symmetrically covering the second combustion chamber and enabling more accurate concentric direction radial flow of gas from the annular channel in the Central part of the second combustion chamber.

4. Turbojet engine, characterized in that the bottom of the guide pad has a larger diameter and area as compared with the diameter and cross-sectional area of the nozzle at the outlet of the second combustion chamber.



 

Same patents:

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8 dwg

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