JET ENGINE FOR AIRCRAFT

Abstract

A jet engine for propelling aircraft, capable of providing thrust from rest to high speeds is provided. The engine has an axial compressor (16) or several axial compressors located on the same plane and is driven by a gas generator. At the outlet of the turbine there is a gasification chamber (23) into which more fuel is injected. Combustion of the gases from the gasification chamber is performed in two combustion chambers (18) with a rectangular cross-section, separated by a central body (10). The exhaust of the gases is performed in nozzles, each with a square convergent/divergent cross-section (19) and (21). The cross-section of the throats (26) can be adjusted by means of two mobile elements (20). The final section of the central body (10) forms a wedge-shape (27), enabling the continued expansion of the exhaust gases.

Claims

1-8. (canceled)

9. A jet engine for propelling an aircraft comprising a gas turbine driving a compressor; the compressor also driven by a gas generator; wherein combustion chambers having rectangular cross-sections and are separated by a central body; an exhaust external to the engine is demarcated by two mobile elements pivoting on a point and ending in a sharp edge; the central body is adapted to widen as exhaust gases progress therethrough until the exhaust gases overcome the two mobile elements, the central body thereafter narrows to form a wedge-shape, an angle of the two mobile elements being adjustable, the cross-section of the combustion chambers being variable such that when the two mobile elements are separated from the central body, a substantially constant cross-section of the two exhausts exists which is similar to the cross-section of the combustion chambers upstream of the exhaust is achieved and when the two mobile elements approach the central body, the two mobile elements together with the central body form two converging nozzles, each having a throat at which the central body, and the two mobile elements are rounded so as to form two diverging nozzles with a small expansion ratio after each of the two throats results in a cross-section of the combustion chambers that always varies smoothly.

10. The jet engine of claim 9 wherein the combustion chambers are two parallel and symmetrical chambers.

11. A method for propelling an aircraft comprising: providing hot gases coming from a gas generator at an outlet of a turbine; reducing speed of gases with a diffuser installed in the outlet; discharging the diffuser into a gasification chamber; injecting additional fuel in liquid state into the gasification chamber; vaporizing the fuel with heat from the hot gases from the gas generator, wherein the gasification chamber discharges into a series of conduits that lead the hot gases to an injector array for combustion in the combustion chamber or a plurality of combustion chambers.

12. The method of claim 11 further comprising using two liquid fuels with different molecular weight at the same time.

13. The jet engine of claim 9 wherein three differentiated areas are provided for injecting fuels: the gas generator the gasification chamber, and the combustion chambers.

14. The method of claim 12 wherein the fuel is mixture rich in fuel with a low molecular weight with an oxidant injected into the gas generator.

15. The method of claim 12 further comprising injecting lean mixtures in the two parallel and symmetrical combustion chambers combustion chambers are provided that are rich in fuel with a high molecular weight during normal mode operation with the proviso that the injectors located in the two parallel and symmetrical combustion chambers combustion chambers are not used.

16. The method of claim 15 further comprising burning rich mixtures in the combustion chambers during exceptional operating condition, the fuel injected into both the gasification chamber and the combustion chambers.

17. The method of claim 16 wherein either of the two fuels areas are injected into injected into both the gasification chamber and the combustion chambers.

18. The method of claim 16 wherein both of the two fuels areas are injected into injected into both the gasification chamber and the combustion chambers.

19. The jet engine of claim 9 further comprising an additional number of axial compressors to provide in total an even number of compressors.

20. The jet engine of claim 19 wherein the compressors are located symmetrically in the same plane, and such that half of the compressors feed each of two of the combustion chambers.

21. The jet engine of claim 19 wherein the gas generator comprises a plurality of gas generators wherein at an the outlet of each of the plurality of gas generators, a plurality of turbines are driven, and the plurality of turbines being mechanically connected with the compressors by transmission boxes and shafts.

22. The jet engine of claim 19 wherein the engine has an asymmetric exhaust configuration.

23. A jet engine for propelling an aircraft comprising a gas turbine driving a compressor located on one side of plane of the engine; the compressor also driven by a gas generator; wherein a single combustion chamber has a rectangular cross-section; a central body thereafter narrows to form a wedge-shaped ramp; an exhaust external to the engine is demarcated by a single mobile element pivoting on a point and ending in a sharp edge; the central body is adapted to widen as exhaust gases progresses therethrough until exhaust gases overcome the single mobile elements, an angle of the single mobile elements being adjustable, the cross-section of the single combustion chamber being variable such that when the single mobile element is separated from the central body, a substantially constant cross-section of the exhaust exists which is similar to the cross-section of the combustion chamber upstream of the exhaust is achieved and when the single mobile elements approaches the central body, the single mobile elements together with the central body form a converging nozzle with a throat at which, the central body, and the single mobile element are rounded so as to form a diverging nozzle with a small expansion ratio after the throat resulting in a cross-section of the combustion chamber that always varies smoothly wherein the turbine is mechanically connected with the compressor by a transmission box and shaft.

Description

[0006] The ENGINE consists of a central body (10) in which, in addition to other auxiliary elements, the following elements are located to drive the compressor (16): [0007] A gas generator (11) into which a liquid fuel and a liquid oxidant are injected at high pressure in a ratio rich in fuel, such that not all the fuel is burned in this chamber. As its name suggests, the function of the gas generator is to generate gas at high pressure. [0008] A turbine (12) that is driven with the gases from the gas generator and moves a high rotational speed shaft (13). [0009] A transmission box (14) in which the rotational speed is reduced in order to adapt same to the needs of the compressor (16).

[0010] The jet engine itself consists of the following elements: [0011] An air intake (15) according to the state of the art. Only one example is depicted in FIG. 1, but it can be in various forms. Different configurations typically having an element of variable geometry have been successfully tested in the last 60 years. Like any intake of a jet engine, the function thereof is to adjust air consumption and air entry speed to the ENGINE's needs. 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 speed of the air, with the entry speed to the compressor (16) being always subsonic. [0012] An axial compressor (16) with at least one compression stage. Axial compressors are made up of successive stages of blade rows located perpendicular to the shaft of the compressor. The blade rows that remain fixed are called stators, and the blade rows that rotate integrally with the shaft are called rotors. It is common for one of the stators, or even two stators, to have orientable blades; in the event that this technical solution is adopted, the adjustable stator will be located in the first row of the compressor and/or in the last row of the compressor. [0013] A transition area (17) from the outlet of the axial compressor to two combustion chambers with a rectangular cross-section (18). The function of this transition area is to keep head losses at reasonable values, therefore, it will be as long as necessary to perform this task. [0014] Two combustion chambers (18) with a rectangular cross-section separated by the central body (10). For clarification, the figure does not depict the typical flame stabilizers of many combustion chambers, which does not mean that they may not be necessary. The typical perforated sheet that separates the colder flow in contact with the chamber wall from the hot flow in the rest of the chamber is indeed depicted with a dashed line. [0015] Two converging nozzles (19). To allow adjusting the cross-section of the throats (26). Depending on the ENGINE operation and flight conditions, the angles of the nozzles are adjusted by means of two articulated mobile elements (20). Once reaching the throats of the nozzles (26), the central body (10) and the two mobile elements (20) form two divergent nozzles (21) with a small expansion ratio. [0016] Finally, the central body (10) narrows at the end of the ENGINE, forming a wedge-shape (27) that allows, when necessary, a greater expansion of the exhaust gases, that is, it acts as an open diverging nozzle.

[0017] Depending on the angle of the mobile elements (20), an exhaust with a more or less constant cross-section, that is, without forming a converging nozzle, can be achieved. Or it is possible to obtain a nozzle with a very pronounced narrowing in the throat (26), as shown in FIG. 4. In contrast, the expansion of the exhaust gases is adjusted automatically as a result of the wedge-shape (27).

[0018] The ENGINE has three operating modes. The main mode is referred to as the normal mode, and is characterized by the following elements: [0019] A diffuser at the outlet of the turbine (22). [0020] A gasification chamber (23) into which the diffuser (22) discharges. The speed of the exhaust gases in this chamber is relatively low, and there is injected more fuel which, as a result of the effect of temperature, is gasified.

[0021] Gases from the gasification chamber are led to an injector array (24) which distributes the gases between the two combustion chambers (18). The injectors are sized and placed such that gases are not injected into the region close to the walls, thereby ensuring a lower temperature on the walls of the combustion chambers (18) and the walls of the nozzles (19) and (21).

[0022] In addition, it is possible to operate the ENGINE in an alternative mode referred to as the super-thrust mode. In this mode, the objective is to obtain a thrust greater than the normal mode. To that end, fuel in excess of the oxygen available is injected into the combustion chambers (18), that is, the ENGINE works with a mixture rich in fuel, unlike the normal operating mode mentioned above in which not all the available oxygen is consumed.

[0023] In this mode, not all the fuel is injected into the gas generator (11) and the gasification chamber (23). In addition, part of the fuel is injected directly into the combustion chambers (18), in a series of injectors (25) placed 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 thrust, ensures a lower temperature on the walls of the combustion chamber (18) and the walls of the nozzles (19) and (21).

[0024] In addition to a liquid oxidant to react in the gas generator (11), another differentiating feature of the ENGINE is the use of two liquid fuels. Preferably, a gaseous fuel with a low molecular weight that is stored liquefied in a pressurized tank, and depending on the choice of fuel, the tank will also be cryogenic. And a fuel with a higher molecular weight that will preferably be liquid at ambient conditions. The fuel with a lower molecular weight will preferably be injected into the gas generator (11), and the fuel with a higher molecular weight will preferably be used in the gasification chamber (23). Both fuels can also be used for direct injection into the combustion chamber (25) in the super-thrust mode. However, depending on the needs of thrust and the speed of the aircraft, a more or less proportion of each of the two fuels can be chosen.

[0025] It is emphasized that the use of two fuels is considered the optimal solution, but if simplicity is desired, the ENGINE can also work with a single type of fuel, in which case it will be a fuel with a low molecular weight.

[0026] Like any other heat engine, a series of auxiliary elements are required, the following being the most important:

[0027] An electrical machine in engagement with the compressor (16), or with one of the stages of the transmission box (14). This electrical machine can work as a generator, providing electrical power to power the aircraft and the auxiliary systems of the ENGINE. The electrical machine can also work as an electric motor, for example, to facilitate starting the ENGINE.

[0028] Finally, the electrical machine driving the compressor (16) without the help of the gas generator (11) may be of interest. This would be the third operating mode and is referred to as the electric mode. In this mode, the ENGINE requires an external power supply, for example, batteries. In the electric mode, neither fuel nor oxidant is injected, so the thrust is quite low, but it can be useful in certain circumstances, for example, as emergency in the event of a breakdown, or to fly at low speed, or on descending trajectories, or to generate less noise.

[0029] Other notable auxiliary elements are the pumps. High pressure pumps are required to feed the gas generator (11). Low-pressure pumps are required to inject fuel into the gasification chamber (23) or into the combustion chamber (25). The pumps can be driven mechanically by connecting them to one of the shafts of the ENGINE, or can be driven by electric motors, or a combination of both options.

[0030] Like any thermal engine, a cooling system is required so that the components of the ENGINE with thermal load do not exceed the design temperatures; the most noteworthy ENGINE cooling system is the heat sink. For applications with flight speeds up to, for example, Mach 2.5, heat exchange with atmospheric air can be carried out. For applications with higher speeds, fuel and oxidant (or any other fluid carried by the aircraft, such as water, for example) can be used as a heat sink. The first option is the conventional solution for cooling gas turbines; the second option is the conventional solution for cooling rocket engines.

[0031] For its development, the ENGINE requires the same elements and subsystems as any gas turbine. The parts will be mostly metal parts, and the alloys will be chosen, among other reasons, based on the working temperatures. In colder areas, steels and titanium alloys can be used, for example. In areas with moderate temperatures, stainless steels and refractory steels can be used, for example. Finally, in areas with higher temperatures, specific high-temperature alloys, for example, nickel-based alloys, must be used.

[0032] Besides the elements already mentioned, the ENGINE will also require the following typical elements of a gas turbine for its development: sensors of various types, actuators, valves, an electronic control system, and an ignition system. It will also be necessary to ensure a reasonably low friction between the mobile elements by means of bearings, bushings, and a lubrication system. A hydraulic system and/or electric servos will be necessary to drive the actuators.

[0033] The rotational speed of the ENGINE is controlled by acting on the pumps that inject the fuel and the oxidant into the gas generator (11). It is possible to vary both the flow rate/pressure of the liquids and their proportion. Thrust is also regulated by the amount of fuel that is injected into the gasification chamber (23), and if the super-thrust mode is used with the amount of fuel that is injected directly into the combustion chamber (25).

[0034] In exhaust control, the objective is to adjust the cross-section of the throats (26) in order to reach critical conditions, and therefore supersonic speeds are reached after the throats (26). In the area of the flight envelope where the pressure in the combustion chamber (18) is not high enough to allow critical conditions, then the cross-section of the throats (26) is adjusted so that the pressure at the exhaust outlet is similar to atmospheric pressure.

[0035] As can be seen in FIG. 1, a weakness of the ENGINE depicted is that space is required for the transition area (17) to ensure that the axial symmetry flow at the outlet of the compressor adapts to the two combustion chambers (18) with a rectangular section. The length of the transition area (17) can be reduced if instead of a single compressor, two, four, or even six compressors are placed on the same plane, working in parallel, such that half of the compressors feed one combustion chamber (18) and the other half feed the other combustion chamber (18). This alternative architecture is shown in FIG. 5. This figure shows a section perpendicular to the shaft of the ENGINE. The section is made at the height of the compressors, and in particular the option of four compressors is shown. The price to pay for this solution is a more complex system. All the compressors can be driven by a single turbine, or each compressor can be driven by its turbine with its gasification chamber.

[0036] FIG. 6 shows a diagram of the first option, with a turbine (12) in the central body (10) driving several compressors. In this case, the transmission system is complicated, where there is a need for an additional secondary transmission box (29) in each compressor and a main transmission box with multiple outputs (28).

[0037] FIG. 7 shows a diagram of the second option. In FIG. 7, the direction of the flow of the gas generator has been intentionally inverted. In FIGS. 3 and 6 the flow is opposite the forward flow, in FIG. 7 the flow of the gas generator is in the same direction as the forward flow. It should be noted that the ENGINE can work with the gas generator oriented in the forward direction or in the opposite direction.

[0038] The ENGINE in the configuration shown in FIG. 7 can be simplified by removing the half, as shown in FIG. 8. This configuration is also viable. In this configuration, the central body (10) is reduced to one of the sides of the ENGINE, and the wedge-shape (27) that allows open expansion becomes a ramp (31). On the other side of the nozzle, the two mobile elements (20) are reduced to a single mobile element (30). The notable drawback of this configuration is that the symmetry in the nozzle is lost, therefore, depending on the operating conditions, the thrust angle will vary. For aircraft purposes, this configuration introduces a moment that varies at different points in the flight envelope.

LIST OF ELEMENTS INDICATED IN THE DRAWINGS

[0039] 10 central body [0040] 11 gas generator [0041] 12 turbine [0042] 13 high rotational speed shaft [0043] 14 transmission box [0044] 15 air intake [0045] 16 axial compressor [0046] 17 transition area [0047] 18 combustion chambers [0048] 19 converging nozzles [0049] 20 mobile elements [0050] 21 diverging nozzles [0051] 22 turbine outlet diffuser [0052] 23 gasification chamber [0053] 24 injector array [0054] 25 injectors for the super-thrust mode [0055] 26 throats [0056] 27 wedge-shape [0057] 28 transmission box with several outputs [0058] 29 secondary transmission box [0059] 30 a mobile element [0060] 31 ramp