RU2803681C1 - Birotating bypass gas turbine engine - Google Patents

Abstract

FIELD: gas turbine engines.

SUBSTANCE: birotating bypass gas turbine engine containing a non-rotating combustion chamber, a jet nozzle, and a birotating turbine with a birotating compressor. The birotating compressor is made inside the birotating turbine, a fairing is installed concentrically with the outer rotor of the turbine, inside which one working stage of the fan is installed with a fan guide stage behind it, and the non-rotating combustion chamber is made cylindrical, at the junction of the birotating compressor with the birotating turbine with the possibility of turning the air flow from birotating compressor at 180°, and the outlet from the birotating turbine is connected by turning the flow of combustion products by 180° - with a central cylindrical channel inside the birotating compressor and with a jet nozzle at the outlet of the engine. All connections of the stators and rotors of the birotating compressor and birotating turbine to each other and to the combustion chamber are protected by connection protection units.

EFFECT: increased efficiency of the compressor and turbine and simplified design.

8 cl, 9 dwg

Description

The invention can be used for aviation gas turbine engines (gas turbine engines) for any purpose, including military and civilian ones with subsonic flight speed and very low fuel consumption.

A birotative gas turbine engine is known according to RF patent for invention No. 2702317, IPC F02C 3/16, publ. 10/07/2019

This birotative gas turbine engine contains a housing on which pipelines for supplying oxidizing and fuel working fluids to the impeller of the first rotor are rigidly mounted, mounted rigidly on a rotatable shaft, containing a compressor for compressing the oxidizing working fluid, and a reaction turbine made in the form of a Segner wheel, and also containing an impeller of the second rotor, installed coaxially and coaxially around the impeller of the first rotor, with the possibility of independent rotation on its shaft in the direction opposite to the first rotor, the impeller of the first rotor is made in the form of a monoblock, mounted rigidly on its shaft with the possibility of rotation, containing a double-flow closed centrifugal wheel.

Disadvantage: complex design of gas turbine engines due to the use of centrifugal compressors and turbines.

A birotative gas turbine engine is known according to RF patent No. 2347179, IPC E45B 25/00, publ. 02/20/2009, prototype.

This bi-rotative gas turbine engine contains a non-rotating combustion chamber, a jet nozzle and a bi-rotative turbine with a bi-rotative compressor, placed in series one behind the other.

Disadvantages: the large axial size of the engine and the suboptimal operation of its components: the compressor and turbine and their heavy weight.

Objectives of creating the invention: increasing the efficiency of the compressor and turbine, simplifying the design and reducing weight and dimensions.

Achieved technical results; increasing the efficiency of the compressor and turbine, simplifying the design and reducing weight and dimensions.

The solution to these problems is achieved due to the fact that a two-circuit bi-rotative gas turbine engine containing a non-rotating combustion chamber, a jet nozzle and a bi-rotative turbine with a bi-rotative compressor, the fact that the bi-rotative compressor is made inside the bi-rotative turbine, a fairing is installed concentrically with the outer rotor of the turbine, inside which is installed one the working stage of the fan with the fan guide stage behind it, and the non-rotating combustion chamber is made cylindrical, at the junction of the birotation compressor with the birotation turbine with the possibility of rotating the air flow from the birotation compressor by 180°, and the outlet from the birotation turbine is connected by rotating the flow of combustion products by 180 ° - with a central cylindrical channel inside the birotation compressor and with a jet nozzle at the engine outlet, and all connections of the stators and rotors of the birotation compressor and the birotation turbine to each other and to the combustion chamber are protected by connection protection units.

The internal shaft of a birotative compressor can be connected to the first working stage of a propfan.

The external shaft of the birotation turbine can be connected to the second working stage of the propfan.

A bi-rotative gas turbine engine may contain front, middle and rear bearings, which contain lubrication systems.

The front bearing can be installed under the first guide stage of the fan, the middle bearing - under the non-rotating chamber between the outer shaft of the birotation compressor and the intermediate rotor, and the rear bearing - at the outlet of the inner rotor of the birotation compressor from under the non-rotating chamber between the outer shaft of the birotation compressor and the intermediate rotor,

Connection protection units can be equipped with seals, plain bearings and thrust bearings.

The jet nozzle can be configured to rotate from the internal rotor of the birotative compressor.

All bearings can be equipped with an oil cooler.

Conventions used in the description:

non-rotating combustion chamber - 1,

jet nozzle - 2,

birotative compressor - 3,

internal compressor rotor - 4,

intermediate rotor - 5,

compressor guide vane - 6,

compressor blade - 7,

birotative turbine - 8,

external turbine rotor - 9,

turbine nozzle apparatus - 10,

turbine impeller - 11,

entrance fairing - 12,

air intake - 13,

internal cavity of the chamber - 14,

air cavity - 15,

combustion compartment - 16,

nozzle plate - 17,

fuel-air injector - 18,

flame tube - 19,

connection protection unit - 20,

seal - 21,

sliding bearing - 22,

thrust bearing - 23,

fuel line - 24,

main pump - 25,

main shut-off valve - 26,

central front panel - 27,

front panel - 28,

cylindrical central cavity - 29,

the first operating stage of the fan is 30,

first fan blade - 31,

first guide stage of the fan - 32,

first fan guide blade - 33,

second circuit - 34,

fairing - 35,

rear fins of the fairing - 36,

secondary circuit nozzle - 37,

rear support - 38,

middle support - 39,

first air inlet - 40,

second air inlet - 41,

fuel tank - 42,

oil tank - 43,

oil line - 44,

oil cavity - 45,

fuel manifold - 46,

flame tube holes - 47

inner wall of the flame tube - 48,

outer wall of the flame tube - 49,

output channel - 50,

first stage of the propfan - 51,

blade of the first stage of the propfan - 52,

bushing of the first stage of the lorat for propfans - 53,

second stage of the propfan - 54,

blade of the second stage of the propfan - 55,

bushing of the second stage of the propfan blade - 56,

combustion chamber casing - 57,

combustion chamber seal - 58,

internal seal - 59,

oil drain - 60,

heat exchanger - 61,

fan - 62,

oil return pipe - 63,

corrugated sheet - 64,

longitudinal rib - 65,

front bearing - 66.

The essence of the invention is illustrated in the drawings (Fig. 1-9), where:

in fig. Figure 1 shows the appearance of a birotation gas turbine engine with one first working rotor of an external fan,

in fig. Figure 2 shows a view along A at the engine inlet,

in fig. 3 shows the connection node,

in fig. 4 shows the combustion chamber,

in fig. 5 shows a non-rotating combustion chamber and a jet nozzle,

in fig. Figure 6 shows a section C-C along the first holes on the external rotor of the turbine,

in fig. Figure 7 shows the outer turbine rotor with the first air inlet,

in fig. 8 shows the section Ε-Ε,

in fig. 9 shows the rear part of the engine with a jet nozzle and its operation.

The essence of the invention is illustrated in Fig. 1…9.

A birotative gas turbine engine contains (Fig. 1) a non-rotating combustion chamber 1 and a jet nozzle 2, which rotates together with the internal rotor of the compressor 4.

In addition, it contains a birotative compressor 3, an internal compressor rotor 4, an intermediate rotor 5, compressor guide vanes 6, and compressor rotor blades 7.

The birotation turbine 8 contains an external turbine rotor 9, turbine nozzles 10, a turbine impeller 11, an inlet fairing 12, and an air intake 13 (Fig. 1).

The engine can be designed with one fan operating stage (Fig. 1) with a low bypass ratio of 0.2-0.3.

Non-rotating combustion chamber 1 (Fig. 1 and 4) contains an internal chamber cavity 14, an air cavity 15, a combustion compartment 16, a burner plate 17, fuel-air nozzles 18 and a flame tube 19.

In the engine at its inlet (Fig. 1) there is installed the first working stage of the fan 30 with the first working blades of the fan 31, the first guide stage of the fan 32 with the first guide blades of the fan 33.

The first operating stage of the fan 30 is attached to the inner rotor of the compressor 4, and the first guide stage of the fan 32 is installed between the fairing 35 and the outer rotor of the turbine 9.

In addition, all rotors 4, 5 and 9 have a connection protection unit 20 on both sides, which contains seals 21, a plain bearing 22 and a thrust bearing 23 (Fig. 3).

An important component is the seals 21. They must operate at high temperatures for a long time.

An example of such a seal can be considered the invention of the Russian Federation No. 2695874, IPC F16J 15/08, publ. 07/29/2019, developed by the Russian Academy of Sciences (A.A. Blagonravov Institute).

It is proposed to make the seal in the form of a silica hollow cord.

Fuel line 24 (FIG. 1) contains a pump 25 with a main shut-off valve 26 and is connected to a fuel manifold 46 to supply fuel to fuel injectors 18.

A fixed central front panel 27 and a front panel 28 are installed in the front part of the engine; a cylindrical central cavity 29 is made inside the internal birotative compressor 3.

Air is supplied to the unaffected combustion chamber 1 through the first inlet openings 40 and through the second inlet openings 41, shown in more detail in FIG. 9 and 10. Most of the air (about 90% enters the air cavity 15 to power the fuel-air injectors 18 mounted on the nozzle plate 17 and communicating with the fuel manifold 46 (Fig. 1 and 4).

The front bearing 66 can be installed under the first guide stage of the fan 30, the middle bearing 39 - under the non-rotating chamber 2 between the outer shaft of the birotation compressor 4 and the intermediate rotor 9, and the rear bearing 38 - at the exit of the internal rotor of the birotation compressor 4 from under the non-rotating chamber 2 between the outer shaft of the birotative compressor 4 and the intermediate rotor 9,

All bearings 66, 39 and 38 are equipped with lubrication and oil cooling systems. This system will be shown below in more detail (oil pumps and filters and other fittings for the lubrication and oil cooling system are not shown in Fig. 1-9),

The non-rotating combustion chamber 1 contains a flame tube 19 with an inner wall of the flame tube 48 and an outer wall of the flame tube 49. In these walls, holes are made in the flame tube 47 for its cooling. (Fig. 1 and 4).

The outlet channel of the flame tube 40 is made cylindrical in the form of a hollow cylinder and is protected by a corrugated sheet 64 to prevent burning (Fig. 8). At the rear of the engine (Figs. 1 and 9) there is a jet nozzle 2, which is designed to rotate together with the internal rotor of the compressor 4 and has longitudinal ribs 65 inside, which swirl the exhaust combustion products, and this contributes to the swirling of cold air leaving the duct 2- th circuit 34 (Fig. 9). This helps to increase engine thrust and reduce specific fuel consumption.

OPERATION OF DOUBLE-CIRCUIT BIROTATIVE GAS TURBINE ENGINE

A double-circuit birotative gas turbine engine, as mentioned earlier, contains (Fig. 1...9) a non-rotating combustion chamber 1 and a jet nozzle 2 (Fig. 1). Non-rotating combustion chamber 1 is attached to the aircraft wing.

In the birotative compressor 3, passing between the guide vanes of the compressor 6 and the working blades of the compressor 7 (Fig. 1), the air is compressed up to 30...40 times or more. At the same time, rotation from the internal rotor of the compressor 4 is transmitted to the first working blades of the fan 31 of the first working stage of the fan 30.

Then the flow of compressed air from the primary circuit turns 180 to enter the non-rotating combustion chamber 1 and then the combustion products go into the bi-rotating turbine 8.

Passing under the outer rotor of the turbine 9 through the nozzle devices of the turbine 10 and the impellers of the turbine 11, the combustion products give power to the birotation turbine 8 to drive the compressor 3 and to create thrust in the jet nozzle 2 (Figs. 1 and 4)

To do this, the flow of combustion products is once again rotated by 180° in the inlet fairing 12 to be fed into the cylindrical central cavity 29 (details shown in Figs. 1 and 2). The flow of combustion products passes between the air intakes 13 for further advancement into the cylindrical central cavity 29 and then into the jet nozzle 2. The jet nozzle 2 is designed to rotate together with the internal rotor of the compressor 4.

The non-rotating combustion chamber 1 (Fig. 1) contains the internal cavity of the chamber 14 (Fig. 4), which is divided into an air cavity 15 and a combustion compartment 16 by a nozzle plate 17. On the nozzle plate 17 from the side of the combustion compartment 16, fuel = air nozzles 18 are installed, and on the side of the air cavity 15 - fuel manifold 19,

In a non-rotating combustion chamber 1

In addition, all rotors 4, 5 and 9 have connection protection units 20 on both sides. The connection protection unit 20 contains a seal 21, a plain bearing 22, and a thrust bearing 23 (Fig. 5).

Fuel line 24 (Fig. 1) contains a pump 25 with a shut-off valve 26 and is connected to the fuel manifold 46.

At the front of the engine, a central front panel 27 and a front panel 28, which rotates, are installed.

In fig. Figure 8 shows the section C-C. The walls of the flame tube 48 and 49 are installed with radial clearances δ. Corrugated sheets 64 are installed in the radial gaps δ.

In fig. Figure 9 shows the jet nozzle 2 and its operation.

The jet nozzle 2 is designed to rotate together with the external rotor of the birotative compressor 4. Inside the jet nozzle 2, longitudinal ribs 65 are installed, which spin the exhaust flow of the primary circuit

As a result, the hot combustion products and cold air are mixed after the propfan stages 51 and 53.

This increases engine thrust and reduces specific fuel consumption.

The use of the invention allowed:

- create a two-circuit birotative gas turbine engine with the maximum number of rotating rotors,

- abandon the gearbox to reduce the weight of the amount of equipment,

- significantly improve strength and aerodynamic parameters and reduce the length of the engine by almost 50% and its weight by 3...4 times,

- reduce stress from the action of centrifugal forces on rotating parts,

- ensure subsonic flights of civil and military aircraft with very low fuel consumption,

- reliably mix the flows of the first and second exhaust circuits to increase engine thrust and reduce specific fuel consumption.

Claims (8)

1. A two-circuit birotative gas turbine engine containing a non-rotating combustion chamber, a jet nozzle and a birotative turbine with a birotative compressor, characterized in that the birotative compressor is made inside the birotative turbine, a fairing is installed concentrically with the outer rotor of the turbine, inside which one working stage of the fan with a fan guide stage is installed behind it, and the non-rotating combustion chamber is made cylindrical, at the junction of the birotation compressor with the birotation turbine with the possibility of rotating the air flow from the birotation compressor by 180°, and the outlet from the birotation turbine is connected by rotating the flow of combustion products by 180° - with a central cylindrical channel inside birotative compressor and with a jet nozzle at the engine outlet, and all connections of the stators and rotors of the birotative compressor and birotative turbine with each other and with the combustion chamber are protected by connection protection units. 2. Double-circuit birotative gas turbine engine according to claim 1, characterized in that the internal shaft of the birotative compressor is connected to the first working stage of the propfan. 3. A two-circuit birotative gas turbine engine according to claim 1, characterized in that the outer shaft of the birotative turbine is connected to the second working stage of the propfan. 4. Double-circuit birotative gas turbine engine according to claim 1, characterized in that it contains front, middle and rear bearings, which contain lubrication systems. 5. Double-circuit birotative gas turbine engine according to claim 4, characterized in that the front bearing is installed under the first guide stage of the fan, the middle bearing is installed under the non-rotating chamber between the outer shaft of the birotative compressor and the intermediate rotor, and the rear bearing is installed at the exit of the internal rotor of the birotative compressor from -under the non-rotating chamber between the outer shaft of the birotative compressor and the intermediate rotor. 6. Double-circuit birotative gas turbine engine according to claim 1, or 2, or 3, characterized in that the connection protection units are equipped with seals, plain bearings and thrust bearings. 7. Double-circuit birotative gas turbine engine according to claim 1, or 2, or 3, characterized in that the jet nozzle is made rotating from the internal rotor of the birotative compressor. 8. Double-circuit birotative gas turbine engine according to claim 4, characterized in that all bearings are equipped with an oil system with oil cooling.

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