RU2591361C1 - Engine of hypersonic aircraft - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: hypersonic aircraft engine comprises engine nacelle, air intake, casing, compressor with rotor compressor, combustion chamber installed after compressor and connected therewith an air duct, gas turbine, jet nozzle and fuel system connected with combustion chamber. Engine unit is made two-shaft. Compressor is made two-staged in form of in-series installed low and high pressure compressors, between which is air path, in which heat exchanger is installed. Inside air path coaxially with first shaft there is steam turbine with inlet and outlet headers. Input header is connected to output of heat exchanger, which input is connected to fuel system operating on water. Output header is connected with electrolytic cell, which first and second outputs are connected to combustion chamber. Downstream of combustion chamber there is gas turbine. Low-pressure compressor rotor is connected by first shaft to the steam turbine and compressor rotor is connected by second shaft with gas turbine.

EFFECT: invention is aimed at improving power performance of engine by increasing compression ratio of compressor, increased thrust and improved specific characteristics of propulsion system.

7 cl, 15 dwg

Description

The invention relates to engine building, specifically to aircraft engines for supersonic and hypersonic aircraft.

Known hydrogen gas turbine engine according to the patent of the Russian Federation for the invention No. 2029118, IPC F02C 3/04, publ. 05/20/1995, with an auxiliary circuit operating on hydrogen, an additional air path was introduced into the auxiliary circuit, connecting the outlet from the free compressor with the auxiliary chamber. Hydrogen in the engine circuit acts as a refrigerant. To cool the turbine of the main circuit, high-pressure air is used, which, after cooling the turbine, is supplied to the intermediate overheating combustion chamber, where liquefied air enters the gaseous state at the same time.

The disadvantage is the low specific characteristics of the engine due to the low degree of air compression in the compressor.

Known hydrogen gas turbine engine according to the patent of the Russian Federation for the invention No. 2320889, IPC F02K 3/04, publ. 03/27/2008 (prototype), which contains a fan, a high-pressure high-speed compressor, a multiplier, a steam-water heater (steam generator), an afterburner, a turboexpander with a heat and mass transfer apparatus. The engine also has a three-stage active-jet turbine, in which the third stage is radial-axial, the flow part of which passes into the critical supersonic section of the Laval nozzle, surrounded by a steam accumulator. The high-speed high-speed compressor is combined with a pressure increase of 60. The engine is designed for thrust of at least 150 tons with an air flow through the first circuit of 600 kg / s, through the second circuit - 1200 kg / s, the gas temperature in front of the turbine is 2000 K. The fan has the outer diameter of the blades of the first row of 4000 mm Inside the housing of the Laval nozzle, atomic hydrogen injectors are installed to burn the unburned oxidizer. The disks of a high-pressure high-speed compressor are made combined - centrifugal pumps are added to the axial steps. The Laval nozzle is provided with a central body, through the openings of which a steam-air mixture is created, creating an external elastic “shell-pillow”, which allows changing the area of the critical section of the Laval nozzle through passage.

Disadvantages - low level of traction, relatively low specific parameters, such as specific fuel consumption, insufficient compression of the compressor.

Low specific parameters are explained by the fact that it is impossible to create a compressor with a compression ratio of more than 30 ... 40, due to the fact that the air temperature at the outlet of it exceeds 800 ° C. In addition, the energy potential of a gas turbine is not enough to drive a more powerful compressor due to gas temperature limitations at the outlet of the turbine with a range of 1700 ... 1800 K, primarily due to a decrease in the resource of working blades of a gas turbine. The working blades of a gas turbine are on a large diameter, rotate at huge peripheral speeds, therefore, they are subject to significant centrifugal loads. The strength properties of materials with increasing temperature deteriorate.

Objectives of the invention: increasing the energy capabilities of a gas turbine engine and improving its specific characteristics.

Achieved technical results: increasing the compression ratio of the compressor, increasing the traction force of the engine and improving its specific characteristics.

The solution of these problems was achieved in the propulsion system of a hypersonic aircraft containing a nacelle, an air intake, a housing, a compressor with a compressor rotor, a combustion chamber installed behind the compressor and connected to it by an air path, a gas turbine, a jet nozzle and a fuel system connected to the combustion chamber, that the engine is twin-shaft, and the compressor is two-stage in the form of sequentially installed low and high pressure compressors, between low and high compressors In addition, an air path has been installed in which a heat exchanger is installed, a steam turbine is installed coaxially with the first shaft inside the air path, which has an input and output collectors, an input collector is connected to the outlet of the heat exchanger, the input of which is connected to a water-based fuel system, and the output is connected to an electrolyzer, the first and second outputs of which are connected to the combustion chamber, a gas turbine is installed behind the combustion chamber, the rotor of the low-pressure compressor is connected to the gas turbine by the first shaft, and the rotor of the high-pressure compressor is connected by a second shaft to a gas turbine.

An electric generator can be connected to the first shaft, which is installed in the input fairing and the output of which is connected to the electrolyzer by electric wires. The nacelle can be made with its cooling system, which is connected to the fuel system and to the heat exchanger. The inlet cowl is made cooled and connected in series with the fuel system and the heat exchanger. A propulsion system may contain a second fuel system, for example, operating on hydrocarbon fuel, and the combustion chamber is made with two groups of nozzles. The air intake can be made supersonic. The jet nozzle can be made supersonic.

The invention is illustrated in FIG. 1 ... 15, where:

- in FIG. 1 shows a first diagram of a gas turbine engine,

- in FIG. 2 shows a second diagram of a gas turbine engine,

- in FIG. 3 shows a third diagram of a gas turbine engine,

- in FIG. 4 shows a diagram of a cooled air intake,

- in FIG. 5 shows a diagram of a cooled input fairing,

- in FIG. 6 shows a fourth diagram of a gas turbine engine,

- in FIG. 7 shows a diagram of a steam turbine,

- in FIG. 8 shows a diagram of an electrolyzer,

- in FIG. 9 shows a first diagram of a combustion chamber,

- in FIG. 10 shows a second diagram of a combustion chamber,

- in FIG. 11 shows a connection diagram of a heat exchanger and a nacelle,

- in FIG. 12 shows a connection diagram of a heat exchanger, a nacelle and an air intake,

- in FIG. 13 shows a connection diagram of a heat exchanger, a nacelle, an air intake and an intake fairing,

- in FIG. 14 shows a supersonic air intake,

- in FIG. 15 shows a supersonic jet nozzle.

The proposed solution (Fig. 1 ... 15) contains a nacelle 1, an air intake 2, an inlet cowl 3, a housing 4, a low pressure compressor 5, an air duct 6, a heat exchanger 7, a high pressure compressor 8, a combustion chamber 9, a gas turbine 10, and a jet nozzle 11. The jet nozzle 11 is preferably supersonic. In the jet nozzle 11 mounted fairing 12.

The low-pressure compressor 5 contains a stator 13 and a rotor 14. The high-pressure compressor 8 contains a stator 15 and a rotor 16. The combustion chamber 9 contains a heat pipe 17 and nozzles 18. A gas turbine 10 contains a stator 19 and a rotor 20.

The engine contains two shafts 21 and 22. The first shaft 21 is mounted on the supports 23 and 24. The second shaft 22 is mounted on the supports 25 and 26. Inside the air path 6 concentrically to the first shaft 21 is installed a steam turbine 27 operating on superheated water (steam). The steam turbine 27 has an outer diameter smaller than the inner diameter of the air path 6 so as not to clutter it. In addition, the small diametrical dimensions of the steam turbine 27 reduce centrifugal loads on its rotating parts. The steam turbine 27 comprises an outer casing 28, input and output manifolds 29 and 30, respectively, and a rotor 31 (Fig. 1 and 2). In FIG. 2 shows in more detail the design of the steam turbine 27.

The engine (Fig. 1) contains a fuel supply system 32 having a tank 33 for storing the first fuel (water), a low pressure fuel line 34 connected to the outlet of the tank 33. A pump 35, a high pressure fuel line 36, a regulator are connected to the low pressure fuel line 34. 37 and a shut-off valve 38. A heat exchanger 7 is connected to the fuel supply system 32, to the outlet of which a steam turbine 27 is connected by a pipe 39, to the outlet of which an electrolyzer 41, designed to decompose water and having two outputs, is connected by a pipe 40, the first 42 for hydrogen and the second 43 for oxygen. The first exit 42 by a pipe 44 is connected to the manifold 45 and then to the nozzles 18. The second exit 42 by a pipe 46 is connected to a cavity 47 in front of the combustion chamber 9.

An electric generator 48 is connected to the first shaft 21, to the output of which electric wires 49 are connected, the outputs of which are connected to the anode 50 and the cathode 51 of the electrolyzer 41.

The nacelle 1 (Fig. 2) can be made of the outer shell 52, the inner shell 53 with a gap 54 between them, the inlet manifold 55 and the outlet manifold 56, which is connected by a pipe 57 to the inlet to the heat exchanger 7, the outlet of which is connected by a pipe 58 to the inlet electrolyzer 41.

In FIG. 3 shows the third engine diagram.

In it, the input device 2 and the input fairing 3 are made cooled (Figs. 4 and 5). The air intake 2 comprises an outer wall 59, an inner wall 60 with a gap 61 between them. In addition, it contains an input manifold 62 and an output manifold 63. A pipe 64 is connected to the input manifold 62, and a pipe 65 is connected to the output manifold 63.

In FIG. 5 shows a cooled inlet fairing 3, which comprises an outer wall 66, an inner wall 67, a gap 68 between them, an input manifold 69 and an output manifold 70 to which the pipe 71 is connected. An electric generator 48 is installed inside the input fairing 3. This helps to cool it.

In FIG. 6 shows a diagram of a dual-fuel engine.

The use of a dual-fuel engine is expedient for accelerating a hypersonic aircraft, since for its take-off if water is used as fuel, there will be insufficient energy from burning hydrogen decomposed into hydrogen and oxygen in the electrolyzer 41.

The second fuel supply system 72 comprises a second fuel storage tank 73, a low pressure fuel line 74 connected to the outlet of the tank 73. A pump 75, a high pressure fuel line 76, a flow regulator 77, and a shutoff valve 78 are connected to the low pressure line 74. the second fuel 72 is connected to a heat exchanger 7 and a cooled air intake 2 and an intake fairing 3, then a steam turbine 27 and a second manifold 79 with a second group of nozzles 80.

The construction of the steam turbine 27 is described in more detail below (FIG. 7).

The housing 28 includes end caps 81 and 82, on which the input manifold 29 and the output manifold 30 are located. The holes 83 and 84 are made on the end caps 81 and 82 under the collectors 29 and 30. The housing 31 is mounted on the stator part 85.

The rotor 31 comprises a housing 86 in the form of a hollow truncated cone to which end walls 87 and 88 are attached. A first shaft 25 is connected to end wall 87, and sleeve 89 is connected to end wall 88. Sleeve 89 is mounted on support 90.

Nozzle vanes 91 are installed on the housing 28 on the inside, and rotor blades 92 are installed on the housing 86 of the rotor 31 on the outside. The rotor 31 is mounted on the supports 24 and 25 and sealed relative to the stator 28 by seals 93 ... 96.

The second shaft 22 connects the rotor 18 of the high-pressure compressor 6 and the rotor 24 of the gas turbine 10, respectively, and is mounted on supports 25 and 26.

In FIG. 8 shows the electrolyzer.

In FIG. 9 shows a first embodiment of the combustion chamber 9. An annular combustion chamber is used. It contains a nozzle plate 97 with nozzles 18, a flame tube 17 with holes 98, a casing 99 under the flame tube 17, mounted with the formation of an annular cavity 100 between it and the flame tube 17. Inside the casing 99, a cavity 101 is made.

Figure 10 shows a second embodiment of the combustion chamber. A cavity 103 is formed between the flame tube 17 and the outer wall 102.

The heat exchanger 7 can be connected in series with a steam turbine (Fig. 11) or in series with a cooled air intake 2 (Fig. 10) or additionally with a cooled inlet cowl 3 (Fig. 11)

It is possible to perform the suction 2 supersonic (Fig. 12). In this case, it contains a supersonic part 104, a subsonic part 105, an input collector 106, an output collector 107.

The execution of the jet nozzle 11 is supersonic. This is suitable for supersonic aircraft. In this case, it contains a subsonic part 108, a supersonic part 109, an input collector 110, an output collector 111.

ENGINE OPERATION

When the propulsion system is operating (Fig. 1 ... 15), it is started by supplying electricity to the starter from an external energy source (in Fig. 1 ... 15, the starter and the energy source are not shown). Then they turn on the pump 35 and the water from the tank 33 is supplied to the heat exchanger 7, then through the pipe 39 to the steam turbine 27 and then through the pipe 40 to the electrolyzer 41, where it decomposes into hydrogen and oxygen. The obtained hydrogen is fed through a pipe 44 to a manifold 45 and then to the nozzles 18 of the combustion chamber 9. At the same time, a pipe 46 is supplied with oxygen into a cavity 47, where it is mixed with air (Figs. 1 and 9). Hydrogen is ignited by the ignition device (in Fig. 1 ... 15 the ignition device is not shown). In this mode, the engine can work if the plane is already flying at a transonic speed and the air heating at its inlet reaches a significant value. Acceleration of the aircraft can be performed using other engines or using a second fuel.

In the second embodiment of the combustion chamber (Fig. 10), the engine is started and the first stage is operated using a second fuel, which is supplied from the tank 73 (Fig. 6) through the second fuel supply system 72 to the second manifold 79 and then to the second group of nozzles 80 After acceleration of the aircraft to M = 0.8 ... 1.0, they switch to the first fuel (water), the energy potential of which will be sufficient for the autonomous operation of the engine.

The engine can operate simultaneously on two fuels.

The rotor 31 of the steam turbine 27 spins and spins through the first shaft 21 the rotor 14 of the low pressure compressor 4. The rotor 16 of the high pressure compressor 8 is driven by the rotor 20 of the gas turbine 10 through the second shaft 22. The low pressure compressor 5 provides a compression ratio of up to 30 ... 40, while the air temperature at its outlet can reach 800 ° K. It is next to impossible to compress such hot air in the compressor further. In the air path 6, the air is cooled to 300 ... 400 K and becomes again suitable for compression. During the combustion of fuel (hydrogen) in the combustion chamber 9, the temperature of the exhaust gas rises to 1800 ... 2000 ° C. A gas having a high temperature and pressure has a significant energy potential for its operation on the gas turbine 10 and to ensure the effective operation of the jet nozzle 11.

The use of a steam heat exchanger 7, as noted earlier, will reduce the air temperature at the outlet of them from 700 ... 800 K to a temperature of 250 ... 300 K in front of the low and high pressure compressors 4 and 6, which will allow both compressors 5 and 8 to compress the combustion products to 100 ... 150 kgf / cm (in terrestrial conditions), i.e. to a pressure commensurate with the pressure in modern rocket engines. Without pre-cooling, low and high pressure compressors 5 and 8 would in principle be inoperative. The high pressure in front of the combustion chamber 9 allows for a pressure differential across the gas turbine 10 and the expiration of the combustion products from the jet nozzle 11 at supersonic speeds, thereby creating a large jet thrust. Very high thrust with small engine dimensions allows you to reach aircraft equipped with such an engine speed M = 10 ... 20 and significantly increase the altitude of the engine.

The thrust force is regulated by the flow regulator 37. When the jet engine is stopped, all operations are carried out in the reverse order, i.e. close the shut-off valve 38.

The application of the invention allowed:

1. To increase the compression ratio of the compressors of the gas turbine engine through the use of two compressors, a gas turbine and a steam turbine running on steam, as well as cooling the combustion products in front of the second compressor. The steam turbine has small diametric dimensions, therefore, less centrifugal forces act on its working blades. The use of water vapor as a working fluid significantly increases the energy potential of this working fluid and ensures safety. In addition, the electrolysis of water and the use of hydrogen as fuel increases the energy potential of the engine. The compression ratio of the compressors can be increased through the use of a heat exchanger installed between them.

2. Ensure that aircraft equipped with these engines achieve hypersonic speeds M = 10 ... 20.

3. Increase engine altitude by generating oxygen from water.

4. Increase engine reliability due to the small dimensions of the steam turbine and reduce centrifugal loads on its blades.

Claims (7)

1. The propulsion system of a hypersonic aircraft containing a nacelle, an air intake, a housing, a compressor with a compressor rotor, a combustion chamber installed behind the compressor and connected to it by an air path, a gas turbine, a jet nozzle and a fuel system connected to the combustion chamber, characterized in that the propulsion system is twin-shaft, and the compressor is two-stage in the form of sequentially installed low and high pressure compressors, between low and high pressure compressors an air path is made in which a heat exchanger is installed, a steam turbine is installed inside the air path coaxially to the first shaft, which has an inlet and an outlet manifold, the inlet manifold is connected to the outlet of the heat exchanger, the inlet of which is connected to the fuel system running on water, and the outlet is connected to the electrolyzer the first and second exits of which are connected to the combustion chamber, a gas turbine is installed behind the combustion chamber, the rotor of the low-pressure compressor is connected to the steam turbine by the first shaft, and the rotor the high pressure compressor is connected by a second shaft to a gas turbine. 2. The propulsion system of a hypersonic aircraft according to claim 1, characterized in that an electric generator is connected to the first shaft, which is installed in the input fairing and the output of which is connected to the electrolyzer by electric wires. 3. A propulsion system according to claim 1 or 2, characterized in that the fuselage is made with its cooling system, which is connected to the fuel system and to the heat exchanger. 4. The propulsion system of a hypersonic aircraft according to claim 1 or 2, characterized in that the inlet fairing is made cooled and connected in series with the fuel system and the heat exchanger. 5. The propulsion system of a hypersonic aircraft according to claim 1 or 2, characterized in that it contains a second fuel system, for example, operating on hydrocarbon fuel, and the combustion chamber is made with two groups of nozzles. 6. The propulsion system of a hypersonic aircraft according to claim 1 or 2, characterized in that the air intake is made supersonic. 7. The propulsion system of a hypersonic aircraft according to claim 1 or 2, characterized in that the jet nozzle is made supersonic.

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