RU2271460C2 - Gas-turbine turbofan engine - Google Patents

Abstract

FIELD: aircraft industry.

SUBSTANCE: proposed gas-turbine turbofan engine has fan, high-speed compressor, combustion chamber with ball-torus convector, two stage turbine connected through shaft with fan and through step-up gear with high-speed compressor, reaction nozzle and reheat unit. Engine is furnished also with air-hydrogen heat exchanger of torus design, turboexpander, steam generator and liquid-propellant engine. Torus heat exchanger is arranged between first and second stages of high-speed compressor. Turboexpander is arranged between middle support and combustion chamber. Steam generator is made integral with turbine guide nozzle assembly. Liquid-propellant engine is arranged after turbine along engine axis and it contains ball-type steam-gas generator inside combustion chamber.

EFFECT: increased thrust at all heights and duties.

14 cl, 11 dwg

Description

The invention relates to the field of engine manufacturing, in particular to aircraft turbojet fan engines. The engine is intended for use as a power unit of the Rossiya carrier aircraft - the Universal Aviation and Space Transport System (UAKTS) of reusable use, horizontal take-off from the water surface and landing of the carrier aircraft on the water surface. After working out the resource, the engine can be used in the gas and oil pumping, water pumping, energy producing sectors of the national economy.

A close technical solution to the invention is a dual-circuit turbojet engine containing an inlet common fan for two circuits and a compressor sequentially located in the internal circuit, a combustion chamber, a two-stage turbine connected to the compressor and fan by means of a shaft and a reducer, a jet nozzle, a heat exchanger, and a steam-water heater (steam regenerator, see 1045686, IPC 7 F 02 K 3/4, 1994).

The disadvantages of this engine are low effective efficiency and low specific thrust.

The closest technical solution (prototype) to the invention (see Russian patent N 2209329) is a double-circuit gas turbine fan engine (DGVD) containing a double-row fan, high-speed compressor (VSK), a combustion chamber with a charotor convector, a two-stage turbine connected by a shaft with a fan and through a multiplier with VSK to ensure joint rotation, a steam-water heater located behind the turbine along the engine axis, an air-hydrogen, nitrogen-oxygen heat exchanger located between the first and second stages of VSK, afterburner, Laval nozzle.

The disadvantages of this engine are: a limited decrease in the enthalpy of gaseous air, a decrease in traction with increasing flight altitude.

The main objective of the invention is to increase traction at all operational heights and engine operating modes. The problem is solved by the following constructive measures:

- behind the first stage of the VSK installed an air-hydrogen heat exchanger torus design;

- a turboexpander with a heat and mass transfer apparatus is installed between the middle support and the combustion chamber, in which liquid air is separated into gaseous nitrogen and liquid oxygen;

- a steam generator is installed on the engine, made in one piece with the guide nozzle apparatus of the turbine, and a liquid-rocket engine located behind the turbine along the axis of the engine;

- a liquid-rocket engine located behind the turbine along the axis of the engine contains inside the combustion chamber a ball-shaped steam-air generator;

- on the wheel rim of the last stage of the high-speed compressor are located the working blades of the “starter-turbine”, which has a channel for supplying compressed air from a stationary aerodrome installation and from a neighboring power plant when the engine is started in air;

- a ball-shaped steam-air generator located inside the LRE combustion chamber is connected by pipelines to a cavity behind a high-speed compressor, through which compressed air is supplied to accelerate the formation of a steam-air mixture;

- the first stage of the high-speed compressor is made combined, the centrifugal compression stage is added to the axial stage, the stiffeners are made on the disks, and the disks of the second and subsequent axial stages have holes for air to pass from the centripetal expanding stage of the turboexpander to the centrifugal compression stage of the first combined stage of the VSK ;

- the combustion chamber of a liquid-rocket engine has a cooling jacket for the combustion of gaseous hydrogen, and the critical section of the nozzle is additionally cooled by gaseous hydrogen;

- pipes for air passage from the expansion centripetal stage of the turboexpander to the centrifugal compression stage of the first combined stage of the VSK are mounted in the middle engine mount;

- in the steam collector located on the body of the combustion chamber, two rows of nozzles for supplying steam to the combustion chamber are installed for intensive mixing with the cold air stream of the engine primary circuit and hydrogen combustion products leaving the charotor convector;

- the main shaft (fan-turbine) of the engine is mounted on sliding bearings, which increases its longitudinal stiffness;

- the engine has separate bonding of the input and output devices, integrated into one by a trellised intake entering the power circuit of the engine, through which external air enters the afterburner at the time when the valves at the engine inlet open, bypassing part of the intake air from the secondary circuit the first when climbing to a height;

- the starting fuel manifold of the engine is made of two channels, one channel serves for starting fuel, the second - for supplying the steam-air mixture in nominal mode;

- the fifth rocket engine is attached to the frame, and its nozzle, with its output flange, is attached to the rear flange of the Laval nozzle stacker;

- the fan is made with an even number of blades arranged in two rows with a partial entry of the second row of blades into the inter-blade cavities of the first row of blades;

- in the steam generator, made in one piece with the guide nozzle apparatus of the turbine, there are holes designed to organize film cooling of the shelves and legs of the blades of the first stage of the turbine, as well as holes connecting the annular cavity with the steam collector.

Figure 1 shows a longitudinal section of the proposed engine.

Figure 2 shows the combustion chamber of the engine.

Figure 3 shows the plate of the combustion chamber.

Figure 4 shows the head of the combustion chamber with a steam generator.

Figure 5 shows a turboexpander with a middle engine mount.

Figure 6 shows element A of figure 1.

In Fig.7 is a section A ₁ -A ₁ Fig.6.

On Fig shows a longitudinal section of the rocket engine.

In Fig.9 is a view B of Fig.8.

Figure 10 shows a section along the blades of the first and second row of the fan.

11 shows an input device to a hydrogen engine.

The double-circuit gas turbine fan engine contains a double-row sixteen-blade variable-pitch fan 1, a high-speed compressor (VSC) 2, a combustion chamber 3, a two-stage turbine 4, which rotates the fan and through a multiplier 5 VSC2, a jet nozzle 6. To increase the power supply and engine efficiency on it installed:

- air-hydrogen heat exchanger 7 torus design, located between the first and second stages of the VSC;

- a turboexpander 8 with a heat and mass transfer apparatus 35 located between the middle support 26 and the combustion chamber 3;

- liquid-rocket engine (LRE) 10, located behind the turbine 4 along the axis of the engine, contains inside the combustion chamber a gas-vapor generator of spherical shape.

On the rim of the disk of the last stage of the VSK are located the rotor blades 11 of the "starter-turbine" (see Fig. 1), having a channel 12 for supplying compressed air from a stationary aerodrome installation or from a neighboring engine when it is launched in the air.

The combustion chamber 3 (FIG. 2) is welded, inside it contains a convector of the charotor structure 15, the annular one is intended for the formation and combustion of a fuel-air mixture. It contains a power casing 14, a convector of the ball joint design. The cavity formed between the casing 14 and the charotor convector 15 is an active mixing zone, where convective heat transfer of burning gases leaving the charotor convector 15 with the main air stream of the VSK is carried out, thereby achieving equalization of the temperature field in front of the turbine 4. On the casing of the combustion chamber 14 is located the starting two-channel fuel manifold 16. Starting fuel is supplied to the channel 17. The collector 19 has two rows of nozzles 20 for supplying steam to the combustion chamber 3. The holes 21 of the steam generator 13 (see figure 4) serve for the organization of film cooling of the shelves and legs of the blades of the first stage of the turbine 4, and the holes 83 connect the outer annular cavity to the steam collector 82, covering the outer annular cavity of the steam generator. The power housing 14 of the combustion chamber is connected with the power frame 22 of the charotor convector 15. On the starting fuel manifold 16, starting nozzles 23 are installed. The plate 24 of the charotor convector, on which the nozzles 25 are located, is made detachable from the case of the charotor convector 15.

Turbo expander 8 (see Figs. 1 and 5) is located behind the middle support 26. It contains two wheels mounted on bearings that rotate on rolling bearings. The first wheel (left) has two centripetal (expanding) steps 27 and 28. Step 27 delivers part of the air taken after the first stage of the VSK through the bypass windows 29 to the centrifugal (pumping) stage 30, which is integral with the disk 31 of the axial stage 32 of the VSK . Stage 28 delivers most of the liquid air to the heat and mass transfer apparatus 35, where it is separated into gaseous nitrogen and liquid oxygen. The second wheel (right) of the turboexpander also has two stages, one of them 33 - centrifugal (pumping) takes liquid oxygen from the heat and mass exchanger 35, and the centripetal (expanding) stage 34 directs the cooled air to the disks and blades of the turbine 4 with the aim of cooling them. The heat and mass exchanger 35 has a nitrogen gas extraction flange 36. Both wheels are driven into rotation by the blades 37 and 38 operating in turbine mode using the energy of the high-speed air flow of the primary engine circuit. Openings 39 are made in the VSK disks of the second, third, and subsequent stages, through which cooled air from stage 27 enters stage 30. When starting the engine at start, air is supplied from the stationary airfield installation to flange 40. When starting the engine in the air, air is supplied to channel 12 from an adjacent engine. On the disks of all stages of the compressor and turbine, stiffeners 41 are made.

Behind the first combined stage of the HSC, a torus-type heat exchanger 7 is located, having valves 43 above, through which part of the vaporized hydrogen gas enters the flow part of the HSC, further reducing the enthalpy of air passing through the first motor circuit. The heat exchanger 7 on the outer and inner surfaces has guide vanes 44 (see figure 7).

To increase frontal thrust, flight efficiency and reduce noise, a two-row variable-pitch fan (see figure 10) is made with an even number of blades. The blades of the second row 45 to a certain size are included in the inter-blade space of the blades of the first row 46.

In the middle support 26, air bypass pipes 47 are arranged from the centripetal expanding stage of the turboexpander 27 to the centrifugal compression stage 30 of the combined first stage of the VSK.

The main shaft of the fan-turbine of the engine rotates on sliding bearings 48 (see Fig. 5) located in its front, middle and rear bearings, having the following advantages compared to rolling bearings:

- significantly reduces engine noise in throttle mode;

- it is less sensitive to shock loads;

- has low hydraulic resistance at high speeds;

- simple in design;

- increases the longitudinal stiffness of the engine;

- less heat from friction;

- less weight.

The engine has separate cowling of the input and output devices, integrated into one by the lattice air intake 49 included in the power circuit of the engine, and the lattice air intake is capotated (not shown) by shutters opening with hydraulic cylinders. The input device is adjustable (see Fig. 12), which allows you to change the bypass ratio with a rise to a height to reduce the rate of thrust drop created by the first circuit. A high degree of ejection of the jet nozzle allows the external air to be sucked into the second circuit through the open gates of the air intake grill, which provides an additional supply of air necessary for inclusion in the ramjet operation in all modes, including takeoff, in order to obtain greater thrust.

In the engine, the unloading of the high-speed compressor from the action of axial forces is carried out by supplying compressed air through pipelines 50 to the space between the disk 31 and the front engine support housing.

In the dual-circuit gas turbine fan engine behind the turbine, an axis liquid-propellant rocket engine (LRE) 10 is located, inside the combustion chamber of which a steam-gas generator 52 of spherical construction is located. The liquid propellant rocket engine has a cooling jacket 53 cooled by the products of hydrogen combustion, and the critical section of the nozzle is additionally cooled by hydrogen gas, which is supplied through pipes 54 to nozzles 55 located opposite the steam-air mixture supply valves 85. The products of hydrogen combustion are fed into the head of the liquid fuel combustion chamber through openings 56, leaving the turbine 4 and through the windows 57 (see Fig. 9) entering the space under the drain of the Laval nozzle 58. Air is supplied through the pipe 59 to the steam and gas generator 52 because of the last stage of the VSC. The LRE mounts the fifth 60 head of the combustion chamber to the frame 61 (see figure 9), and the output flange 62 of the nozzle of the combustion chamber LRE - to the flange 63 of the inner runner 58 of the Laval nozzle. On the steam and gas generator 52 (along the axis) there is a nozzle 64, which serves to fill the air-fuel mixture of the LRE combustion chamber, which enters into operation with the engine entering the idle mode.

The bypass control system (see Fig. 12) includes sashes 75 and 76 and hydraulic cylinders 77 and 78.

The engine operates as follows: at the start, hot compressed air is supplied to the flange 40 of the collector 50 (see Fig. 5) from the airfield installation, which enters the working blades of the "starter-turbine", and the starting fuel-gas hydrogen is fed into the channel 17 of the fuel manifold (see .Fig. 2) and a two-channel nozzle 23 - the engine goes to idle. Next, the fuel is supplied to the fitting 79 (see figure 3) and the channel of the plate 24 of the sharotor convector. At the same time, steam is supplied to the flange 80 of the manifold 19, and through the nozzles 20 it goes further to the combustion chamber, and through the nozzles to the charotor convector 15. Steam to the flange 80 comes through pipes 81 from the steam collector 82, openings 83 connected to the steam generator 13 made in one piece with the hollow guide vanes 84 of the turbine 4. In the sharotor convector, the combustion of hydrogen begins in a mixture of air coming from the VSC and the steam coming from the steam generator 13. The steam is also fed into the channel 19 of the starting manifold and nozzles 23. The engine goes to nom tional regime. On take-off, fuel is supplied to the nozzles of the LRE combustion chamber and the nozzles of the afterburner of the engine. At the same time, the vapor-air mixture through the valves 85 enters the combustion chamber of the rocket engine. On take-off, the engine develops a thrust of at least 150 tons. With a rise to a height, the shutters 75 and 76 open and air from the second circuit enters the first circuit, thereby reducing the rate of decrease in traction created by the first circuit, and to open the afterburner’s starvation with a rise to a height, they open leaflets of the trellised intake (not shown), through which the intake of external air into the afterburner takes place.

Claims (14)

1. A double-circuit gas turbine fan engine containing a fan, a high-speed compressor (VSK), a combustion chamber with a charotor convector, a two-stage turbine connected by a shaft with a fan and through a multiplier with a high-speed compressor, a jet nozzle, an afterburner, characterized in that it is equipped with an air - a torus-type hydrogen heat exchanger located between the first and second stages of a high-speed compressor, a turboexpander located between the middle support combustion chamber, a steam generator, formed integrally with the guide nozzle assembly of the turbine, and a liquid-propellant rocket engine, arranged downstream of the turbine of the motor axis. 2. The engine according to claim 1, characterized in that the liquid-rocket engine located behind the turbine along the axis of the engine contains inside the combustion chamber a steam-gas generator of a spherical shape. 3. The engine according to claim 1 or 2, characterized in that on the wheel rim of the last stage of the high-speed compressor are located the working blades of the “starter-turbine”, which has a channel for supplying compressed air from a stationary aerodrome installation or from a neighboring engine when it is launched in the air. 4. The engine according to p. 3, characterized in that the gas generator located inside the combustion chamber of a liquid-rocket engine is connected to a high-speed compressor. 5. The engine according to claim 4, characterized in that the first stage of the high-speed compressor is made combined, a centrifugal stage is added to the axial stage, and stiffeners are made on the disk, and in the disks of the second and subsequent stages there are openings for air to pass from the centripetal expanding stage turboexpander to the centrifugal compression stage VSK. 6. The engine according to claim 5, characterized in that the combustion chamber of a liquid-propellant rocket engine has a cooling jacket for the combustion of gaseous hydrogen, and the critical section of the nozzle is further cooled by gaseous hydrogen. 7. The engine of claim 6, characterized in that pipes for air passage from the expansion centripetal stage of the turboexpander to the centrifugal compression stage of the first combined stage of the high-speed compressor are mounted in the middle support. 8. The engine according to claim 7, characterized in that in the steam manifold located on the housing of the combustion chamber, two rows of nozzles for supplying steam to the combustion chamber are installed. 9. The engine according to claim 8, characterized in that the main shaft (fan-turbine) is mounted on sliding bearings. 10. The engine according to p. 9, characterized in that it has separate bonding of the input and output devices, combined into a single trellised intake, included in the power circuit of the engine. 11. The engine according to claim 10, characterized in that the starting fuel manifold is made of two channels, one channel serves for starting fuel, and the second for supplying a steam-air mixture in nominal mode. 12. The engine according to claim 11, wherein the fifth rocket engine is attached to the frame, and its nozzle is attached with its output flange to the rear flange of the Laval nozzle stacker. 13. The engine according to claim 11, characterized in that the fan is made with an even number of blades arranged in two rows with a partial entry of the second row of blades into the interscapular cavities of the first row of blades. 14. The engine according to p. 13, characterized in that the steam generator, made in one piece with the guide nozzle apparatus of the turbine, has openings designed to organize film cooling of the shelves and legs of the blades of the first stage of the turbine, as well as openings connecting the annular cavities with steam collector.

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