Jet engine for an aircraft

Abstract

A jet engine for propelling an aircraft capable of providing thrust from standstill to high speed. It is characterized in that it employs a liquid oxidizer and two liquid fuels of different molecular weights. The engine has an axial compressor (16) or several axial compressors located in the same plane, but unlike the conventional gas turbine, the turbine is driven by a gas generator. At the outlet of the turbine there is a vaporization chamber (23) into which more fuel is injected. The combustion of the gases from the vaporization chamber takes place in two combustion chambers (18) of rectangular section, separated by a central body (10). The discharge of the gases takes place in nozzles (19) and (21) with square converging / diverging sections. The section of the throat (26) can be adjusted by means of two movable elements (20). The last part of the central body (10) forms a wedge (27) that enables the exhaust gases to continue to expand.

Description

[0001] The present invention relates to a heat engine that provides thrust to an aircraft by accelerating air from the atmosphere through a duct. As a result of mechanical elements, the present invention relates to a turbine. However, the present invention differs from a typical gas turbine construction as it uses a gas generator to drive the turbine.

[0002] Most aircraft that fly at high speed use a gas turbine as an engine (where high speed is understood to be supersonic). The most typical construction is the so-called turbofan. Usually a low bypass ratio turbofan with afterburner is used. This type of engine requires a huge development and manufacturing cost.

[0003] As additional information, two other alternatives are used in some applications. The first alternative is the rocket engine. This type of engine has at least two problems: low specific impulse and short service life, usually difficult to reuse and expensive to reuse after using it only once. The second alternative is the ramjet. The main problem of this ramjet is that they do not provide thrust at low speed, so other propulsion systems are needed to reach the operating speed.

[0004] The present invention relates to a jet engine that is more economically developed than the traditional engine described above (i.e. low bypass turbofan with afterburner). Figure 1 A schematic view of the present invention (hereinafter engine) is shown. In particular, two sections are shown in planes that contain the shaft (13) of the engine and that are perpendicular to each other. The air flow is represented by arrows.

[0005] Although Figure 1 The engine is depicted in a schematic way, but the most important elements are shown in dimensions that represent a real application. As the dimensions of these components differ greatly, in Figure 2 the area in Figure 1 is reproduced in a larger scale and in Figure 3 the area in Figure 2 is reproduced in a larger scale.

[0006] The engine comprises a central body (10) in which, among other auxiliary elements, the following elements are located for driving the compressor (16):

[0007] - a gas generator (11) in which liquid fuel and liquid oxidizer are injected at a rich fuel ratio at high pressure into the gas generator (11) so that not all the fuel burns in this chamber. As its name indicates, the function of the gas generator is to generate gas at high pressure.

[0008] - a turbine (12) that is driven by the gas from the gas generator and that moves a high speed shaft (13).

[0009] - a gearbox (14) in which the rotation speed is reduced to adapt to the needs of the compressor (16).

[0010] The jet engine itself is composed of the following elements:

[0011] - an air intake (15) according to the state of the art. In Figure 1 the figure only one example is depicted, but the air intake can also be in various forms. In the last 60 years different configurations of elements, usually with variable geometry, have been successfully tested. Like any air intake of a jet engine, its function is to regulate the air consumption and the air entry speed for the needs of the engine. In most of the flight envelope it will act as a diffuser; in particular, when the flight speed is supersonic, one or more shock waves are generated to reduce the air speed, where the speed entering the compressor (16) is always subsonic.

[0012] - an axial compressor (16) with at least one compression stage. The axial compressor is composed of successive rows of blades arranged perpendicular to the compressor axis. The rows of blades that remain fixed are called stators, while the rows of blades that rotate with the shaft are called rotors. Usually one or even both stators have orientable blades; in the case where this solution is adopted, the adjustable stator is located in the first row of the compressor and / or in the last row of the compressor.

[0013] - a transition zone (17) from the outlet of the axial compressor to two combustion chambers (18) with rectangular section. The function of this transition zone is to keep the head losses at reasonable values, therefore, as far as necessary, it will perform this task.

[0014] - two combustion chambers (18) with rectangular section separated by the central body (10). For the sake of clarity, in the attached figures the typical flame holders of many combustion chambers are not depicted, but this does not mean that they can be unnecessary. In fact, a typical perforated plate is depicted with dashed lines, which separates the cold flow in contact with the chamber walls from the hot flow in the rest of the chamber.

[0015] - two convergent nozzles (19) that allow to adjust the section of the throat (26). Depending on the engine operation and the flight conditions, the angle of the nozzles is adjusted by means of two hinged movable elements (20). The central body (10) and the two movable elements (20) form two divergent nozzles (21) with small expansion ratio when reaching the throat (26) of the nozzles.

[0016] - finally, the central body (10) narrows at the end of the engine to form a wedge (27) that allows a greater exhaust expansion when necessary, i.e. it acts as an open divergent nozzle.

[0017] Depending on the angle of the movable element (20), it is possible to achieve an exhaust with approximately constant section, i.e. without forming a converging nozzle, or to obtain a nozzle with a significant narrowing in the throat (26), as shown in Figure 4 Conversely, thanks to this wedge (27), the exhaust expansion is automatically regulated.

[0018] The engine has three operating modes. The main mode is called normal mode, which is characterized by the following elements:

[0019] - a diffuser (22) at the outlet of the turbine.

[0020] - a gasification chamber (23) into which the diffuser (22) discharges. The speed of the exhaust in this chamber is relatively low and more fuel is injected, which is gasified due to the effect of the temperature.

[0021] The gas from the gasification chamber is directed to an array of injectors (24) which distribute the gas between the two combustion chambers (18). The size and placement of the injectors are such that the gas is not injected into the areas close to the walls, thus ensuring lower temperatures on the walls of the combustion chambers (18) and on the walls of the nozzles (19) and (21).

[0022] In addition, the engine can be operated in an alternative mode called super-thrust mode. In this mode, the aim is to obtain a thrust greater than in the normal mode. To do this, fuel containing more than the available oxygen is injected into the combustion chambers (18), that is to say, unlike the normal operating mode described above, in which not all the available oxygen is consumed, the engine works with a fuel-rich mixture.

[0023] In this mode, not all the fuel is injected into the gas generator (11) and the gasification chamber (23). In addition, a part of the fuel is injected directly into the combustion chambers (18) in a series of injectors (25) located on the periphery of each combustion chamber (18) and, if necessary, on the periphery of each nozzle (19) and (21). This additional fuel, in addition to helping to increase the thrust, ensures lower temperatures on the walls of the combustion chambers (18) and on the walls of the nozzles (19) and (21).

[0024] Another distinctive feature of the engine, in addition to the liquid oxidizer that reacts in the gas generator (11), is the use of two liquid fuels. Preferably, a gaseous fuel with low molecular weight is stored liquefied in a pressurized tank, and according to the chosen fuel, the tank can also be cryogenic. A fuel with higher molecular weight is preferably liquid at ambient conditions. The fuel with lower molecular weight is preferably injected into the gas generator (11), while the fuel with higher molecular weight is preferably used in the gasification chamber (23). In the super-thrust mode, both fuels can also be used for direct injection into the combustion chamber (25). However, according to the thrust required and the speed of the aircraft, the approximate proportion of each of the two fuels can be chosen.

[0025] It is emphasized that the use of two fuels is considered the best solution, but if simplification is required, the engine can also work with a single type of fuel, which in this case can be the fuel with low molecular weight.

[0026] Similar to any other thermal engine, a series of auxiliary elements are required, the most important of which are:

[0027] An electric machine, which is coupled to the compressor (16), or to one of the stages of the transmission box (14). The electric machine can work as a generator, thus providing the electrical power to power the auxiliary systems of the aircraft and of the engine. The electric machine can also work as an electric motor, for example, in order to start the engine.

[0028] Finally, the electric machine driving the compressor (16) without the help of the gas generator (11) can be of interest. This would be a third operating mode, and is called electric mode. In this mode, the engine requires an external power source, for example a battery. In the electric mode, neither fuel nor oxidizer is injected, so the thrust is very low, but this will be useful in certain situations, for example, in emergency situations in case of failure, or to fly at low speed, or to fly on a descending trajectory, or to generate less noise.

[0029] Another important auxiliary element is the pump. A high-pressure pump is required to feed the gas generator (11). A low-pressure pump is required to inject the fuel into the gasification chamber (23) or into the combustion chamber (25). The pump can be driven mechanically by coupling it to one of the shafts of the engine, or it can be driven by an electric motor, or by a combination of the two options.

[0030] As for any heat engine, a cooling system is required so that the engine components under thermal load do not exceed the design temperature; the most interesting engine cooling system is the radiator. For applications with flight speeds, for example, up to 2.5 Mach, heat exchange with the air in the atmosphere is possible. For applications with higher speeds, the fuel and oxidizer (or any other fluid carried by the aircraft, such as water) can be used as a radiator. The first option is the conventional solution for cooling gas turbines; the second option is the conventional solution for cooling rocket engines.

[0031] For the development of the engine, the same elements and subsystems as for any gas turbine are required. These components will be mainly metallic components, and the alloy is chosen according to the operating temperature and other reasons. For example, in colder areas, steel and titanium alloys can be used. For example, in areas with moderate temperatures, stainless steel and refractory steel can be used. Finally, in areas with higher temperatures, specific high-temperature alloys must be used, such as nickel-based alloys.

[0032] In addition to the elements already mentioned, for the development, the engine will also require the following typical elements of a gas turbine: various types of sensors, actuators, valves, electronic control systems, and ignition systems. It is also necessary to ensure reasonable low friction between movable elements through bearings, bushings, and lubrication systems. Hydraulic systems and / or electric servos will be necessary to drive the actuators.

[0033] The engine speed is controlled by the pumping action of the fuel and oxidizer into the gas generator (11). The flow rate / pressure of the liquids and their ratio can be varied. The thrust can also be adjusted by the amount of fuel injected into the vaporization chamber (23), and if the super-thrust mode is used, by the amount of fuel injected directly into the combustion chamber (25).

[0034] In the exhaust control, the objective is to adjust the cross section of the throat (26) to reach the critical condition, thus reaching supersonic speed after the throat (26). In the flight envelope area where the pressure in the combustion chamber (18) is not enough to allow the critical condition, the cross section of the throat (26) is adjusted so that the pressure at the exhaust is similar to atmospheric pressure.

[0035] In Figure 1 It can be seen in Fig. 1 1 that the depicted engine has the disadvantage that the transition zone (17) requires space to ensure that the axisymmetric flow at the compressor outlet adapts to the two combustion chambers (18) with rectangular cross section. If instead of a single compressor two, four, or even six compressors are placed in the same plane to work in parallel, so that half the compressors feed one combustion chamber (18) and the other half feed the other combustion chamber (18), the length of the transition zone (17) can be reduced. Figure 5This alternative architecture is illustrated. The figure shows a cross-section perpendicular to the engine shaft. This cross-section is formed at the height of the compressors and specifically shows the option of four compressors. The trade-off for this solution is a more complex system. All compressors can be driven by a single turbine, or each compressor can be driven by a turbine with a vaporization chamber.

[0036] Figure 6 A schematic diagram of the first option is shown, in which a turbine (12) in the central body (10) drives several compressors. In this case, the transmission system is complex, requiring an additional secondary drivebox (29) and a main drivebox (28) with multiple outputs in each compressor.

[0037] Figure 7 A diagram illustrating the second option is shown. Figure 7 In this process, the flow direction within the gas generator is intentionally reversed. Figure 3 and Figure 6 In the middle, the flow is opposite to the forward flow, in Figure 7 In this configuration, the flow in the gas generator is in the same direction as the forward flow. It should be noted that the engine can operate with a gas generator oriented in either the forward or reverse direction.

[0038] Simplify by removing half Figure 7 The engine in the structure shown, such as Figure 8 As shown. This configuration is also feasible. In this configuration, the central body (10) is reduced to one side of the engine, and the open-expansion wedge (27) is allowed to become a ramp (31). On the other side of the nozzle, the two movable elements (20) are reduced to a single movable element (30). A significant drawback of this configuration is the loss of nozzle symmetry, and therefore, the thrust angle will vary depending on the operating conditions. For the purposes of the aircraft, this configuration introduces a torque that varies at different points in the flight envelope.

[0039] List of components shown in the attached diagram:

[0040] 10. Central body; 11. Gas generator; 12. Turbine; 13. High-speed shaft; 14. Transmission box; 15. Inlet; 16. Axial compressor; 17. Transition zone; 18. Combustion chamber; 19. Converging nozzle; 20. Movable element; 21. Diverging nozzle; 22. Turbine outlet diffuser; 23. Vaporization chamber; 24. Injector array; 25. Injector for super-thrust mode; 26. Throat; 27. Wedge; 28. Transmission box with several outputs; 29. ​​Secondary transmission box; 30. Movable element; 31. Ramp.

Claims

1. A jet engine capable of providing thrust to an aircraft by accelerating air in the atmosphere through a duct, said engine comprising: An axial compressor (16) connected to a high-speed shaft (13) via a transmission housing, the high-speed shaft (13) driving the axial compressor (16); an air inlet (15) regulating the air speed to subsonic and directing the air to the axial compressor (16); a gas generator (11) in which a fuel-rich mixture of fuel and oxidant is burned under high pressure; a turbine (12) connected to the gas generator and to the high-speed shaft (13) that drives the compressor; a diffuser (22) at the outlet of the turbine (12) discharging into a vaporization chamber (23), into which more fuel is injected, wherein the compressor is adapted to discharge into a transition zone (17), in which... In the engine, airflow is directed to two combustion chambers (18) having rectangular cross-sections separated by a central body (10); a piping system that guides gas from the vaporization chamber (23) to an injector array (24) in the combustion chambers (18); two converging nozzles (19) having rectangular cross-sections connected to each combustion chamber (18) and having two throats (26) with rectangular cross-sections, the cross-sections of the throats (26) being variable due to two movable elements (20) on opposite sides of the central body (10); two diverging nozzles having rectangular cross-sections connected to the throats (26), wherein the central body (10) terminates in a wedge-shaped (27) geometry; and a motor connected to any shaft of the engine, the motor operating as an electric motor and generator.

2. The jet engine according to claim 1, comprising two, four or six compressors arranged in the same plane, all compressors being driven by a single turbine, and including a main drive housing (28) and a secondary drive housing (29) in each compressor, the main drive housing (28) having an input shaft from the turbine and an output shaft in the same number as the compressors.

3. The jet engine of claim 1, comprising two, four or six compressors arranged in the same plane, wherein each compressor is driven by an independent turbine, and each turbine has its own vaporization chamber.

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