RU2805947C1 - Birotating bypass gas turbine engine - Google Patents

Abstract

FIELD: aviation.

SUBSTANCE: invention is related to drives for aviation, specifically to bypass gas turbine engines. A two-circuit birotative gas turbine engine is proposed, containing a non-rotating combustion chamber, a jet nozzle and a birotative turbine with a birotative compressor. The birotative compressor is made inside a birotative turbine; a fairing is installed concentrically with the outer rotor of the turbine, inside of which at least one working part of the propfan is installed with the guide part of the propfan behind it. The non-rotating combustion chamber is made cylindrical at the junction of the birotation compressor with the birotation turbine with the ability to rotate the air flow from the birotation compressor by 180°. The outlet from the birotation turbine is connected by turning the flow of combustion products by 180° with a central cylindrical channel inside the bi-rotative 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. The internal shaft of a birotative compressor can be connected to the first working part of a propfan. The external shaft of the birotation turbine can be connected to the second working part of the propfan. A birotative gas turbine engine may include front, middle and rear bearings. The front bearing can be installed under the first guide part of the propfan, the middle bearing under the second guide part of the propfan, and the rear bearing at the outlet of the internal rotor of the birotative compressor. 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.

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

7 cl, 11 dwg

Description

The invention can be used for aviation gas turbine engines (gas turbine engines) for any purpose, including military and civilian, 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 bi-rotative 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 shaft with the possibility of rotation, containing a compressor for compressing the oxidizing working fluid and a reaction turbine made in the form of a Segner wheel, and also containing the 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, rigidly mounted on its shaft with the possibility of rotation, containing double-flow closed centrifugal impeller.

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 bi-rotative gas turbine engine containing a fixed combustion chamber and 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 and simplifying the design.

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, and also because 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 at least one working stage of the propfan with a guide stage of the propfan 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 products combustion at 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 birotation turbine with each other and with 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 birotative gas turbine engine may include front, middle and rear bearings.

The front bearing can be installed under the first guide stage of the propfan, the middle bearing - under the second guide stage of the propfan, and the rear bearing - at the outlet of the internal rotor of the birotative compressor.

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.

Conventions used in the description:

combustion chamber 1,

jet nozzle 2,

birotative compressor 3.

internal compressor rotor 4,

intermediate rotor 5,

compressor guide vanes 6

compressor blades 7,

birotative turbine 8,

external turbine rotor 9,

turbine nozzles 10,

turbine impeller 11,

entrance fairing 12,

air intake 13,

internal cavity of chamber 14,

air cavity 15,

combustion compartment 16,

nozzle plate 17,

fuel injector 18,

flame tube 19,

connection protection unit 20,

seals 21,

plain 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,

first working stage of propfan 30,

the first working blade of the propfan 31,

second working stage of propfan 32,

second rotor blade of the propfan 33,

first guide stage of the propfan 34,

first guide vane of the propfan 35,

second guide stage of the propfan 36,

second guide blade of the propfan 37,

internal drive cavity 38,

fairing 39,

fairing windows 40,

rear fairing rib 41,

fairing outlet channel 42,

air outlet channel 43,

front support 44,

middle support 45,

rear support 46,

fuel tank 47,

oil tank 48,

oil line 49.

oil cavity 50,

fuel manifold 51,

flame tube hole 52,

internal inlet hole 53,

input channel 54,

inner wall of the flame tube 55,

outer wall of the flame tube 56,

output channel 57,

radial rib 58,

2nd circuit path 59,

internal drive 60,

internal drive drive gear 61,

internal drive driven gear 62,

external drive 63,

external drive drive gear 64,

external drive driven gear 65,

external drive box 66,

sound-absorbing panel 67.

The essence of the invention is illustrated in the drawings (Fig. 1-11), where

Fig. 1 shows the appearance of a bi-rotation gas turbine engine with one first working rotor of an external propfan,

Fig. 2 shows the appearance of a bi-rotation gas turbine engine with two working rotors of an external propfan,

Figure 3 shows a view along A,

Figure 4 shows the connection node,

Fig. 5 shows the combustion chamber and the rotating jet nozzle,

Fig.6 shows a rear view of the engine, Fig. IN,

Figure 7 shows a section C - C of the engine from the rear,

Figure 8 shows the internal drive of the rotor fan blades,

Figure 9 shows the external drive of the propeller fan guide blades,

Figure 10 shows a section D - D;

Fig. 11 shows a fairing with a sound-absorbing panel.

The essence of the invention is illustrated in Figures 1...11.

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 rotor of the compressor 4, an intermediate rotor 5, guide vanes of the compressor 6, working blades of the compressor 7; birotation turbine 8 contains an external turbine rotor 9, turbine nozzles 10, a turbine impeller 11, an inlet fairing 12, an air intake 13 (Figs. 1-3).

Two versions of the engine are possible: with one working stage of the propfan (Fig. 1) and with two or more working stages of the propfan (Fig. 2).

In the first version (Fig. 1), the engine contains in the front part of the engine the first working stage of the propfan 30, with the first working blades of the propfan 31, the first guide stage of the propfan 34, with the first guide blades of the propfan 35.

The first stage of the propfan 30 is attached to the inner rotor of the compressor 4, and the first guide stage 34 is installed between the fairing 39 and the outer rotor of the turbine 9 on the front support 44.

In the second embodiment of the engine (Fig. 2), it additionally contains a second working stage of the propfan 32 with second working blades of the propfan 33, a second guide stage of the propfan 36 with second guide vanes of the propfan 37. It is possible to use three or more working stages of the propfan.

The second working stage of the propfan 32 is attached to the outer rotor of the turbine 9, and the second guide stage 36 is installed between the fairing 39 and the outer rotor of the turbine 9 on the middle support 45.

The combustion chamber 1 contains (Fig. 5) the internal cavity of the chamber 14, which is divided into an air cavity 15 and a combustion compartment 16 with a nozzle plate 17. On the nozzle plate 17, fuel injectors 18 are installed on the side of the combustion compartment 16, and on the side of the air cavity 15 - collector 19.

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. 4).

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.

The fuel line 24 (Figs. 1 and 2) contains a pump 25 with a shut-off valve 26 and is connected to a manifold 19 to supply fuel to the fuel injectors 18.

A fixed front panel 27 and a front panel 28 are installed in the front part of the engine.

At the rear of the engine (Figs. 6 and 7) there is a rotating nozzle 2, which has longitudinal partitions 58 inside, which swirl the exhaust combustion products, and this contributes to the swirling of the air leaving the 2nd circuit path (Fig. 7).

8 shows an internal drive 60 that includes an internal drive gear 61 and an internal drive driven gear 62.

Figure 9 shows an outer drive 63 that includes an inner drive drive gear 64 and an inner drive drive gear 65.

The use of drives 60 and 63 (Figs. 6 and 8) made it possible to change the angles of inclination of the propeller fan blades 31 and 33 depending on the operating mode of the engine.

Figure 10 shows a section D - D.

Figure 11 shows a fairing 39 with a sound-absorbing panel 67. It is possible to use other methods of reducing propfan noise. Known, for example, are the works of the TsIAS im. P.I. Baranov, including from the utility model “Birotative propfan”, utility model No. 156245, MPK F02R 3/052, publ. November 10, 2015

OPERATION OF DOUBLE-CIRCUIT BIROTATIVE GAS TURBINE ENGINE

A two-circuit birotative gas turbine engine, as mentioned earlier, contains (Fig. 1...11) a non-rotating combustion chamber 1 and a jet nozzle 2 (Fig. 1 and 2). 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 (Figs. 1 and 2), 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 rotor blade of the propfan 31.

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

Passing under the outer rotor of turbine 9 through the nozzle devices of turbine 10 and the impellers of turbine 11, the combustion products give power to turbine 8 to drive compressor 3 and to create thrust in the jet nozzle 2 (Figs. 6 and 7).

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 3.

The non-rotating combustion chamber 1 (Fig. 1 and 2) contains the internal cavity of the chamber 14 (Fig. 5), which is divided into an air cavity 15 and a combustion compartment 16 with a nozzle plate 17. Fuel injectors 18 are installed on the nozzle plate 17 from the side of the combustion compartment 16 , and on the side of the air cavity 15 there is a collector 19. In addition, all rotors 4, 5 and 9 have a connection protection unit 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 (Figs. 1 and 2) contains a pump 25 with a shut-off valve 26 and is connected to a manifold 19.

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

Figures 8 and 9 show the rear of the engine.

Figure 8 shows a rear view of the engine (Figure 8), view B. Figure 9 shows a section C - C of the nozzle 2 and the rear of the stationary combustion chamber 2.

Figure 10 shows a section D - D. The walls of the flame tube 55 and 56 are installed with radial clearances δ. Corrugations 55 are installed in the radial gaps δ.

The application of the invention allowed;

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

- 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 circuits at the exhaust to increase thrust;

- reduce the noise of the propfan due to its hood and the use of sound-absorbing panels.

Claims (7)

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 at least one working stage of a propfan with the guide stage of the propeller fan 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 the central cylindrical a 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 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. 5. A double-circuit birotative gas turbine engine according to claim 4, characterized in that the front bearing is installed under the first guide stage, the middle bearing is installed under the second guide stage of the propfan, and the rear bearing is installed at the outlet of the internal rotor of the birotative compressor. 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, or 4, characterized in that the jet nozzle is designed to rotate from the internal rotor of the birotative compressor.