Jet engine for aircraft

Description

[0001] The invention relates to a heat engine for providing thrust to an aircraft by accelerating atmospheric air through a conduit. As a result of the mechanical elements, the invention relates to a turbomachine. However, the invention differs from typical gas turbine configurations due to the use of a gas generator to drive the turbine.

[0002] Most aircraft that fly at high speed use gas turbines as engines (with high speeds being understood as supersonic speeds). The most typical configuration is what is known as a turbofan. Low-bypass-ratio turbofans with an afterburner are often used. Engines of this type require enormous development and manufacturing costs.

[0003] As additional information, two other alternatives are used in some applications. The first alternative is the rocket engine. Engines of this type present at least two problems: a low specific impulse and a short service life, often being difficult and expensive to reuse after just one use. The second alternative is the ramjet. The main problem with ramjets is that they do not provide thrust at low speed, so the use of another propulsion system is required to reach their operating speed.

[0004] Aircraft jet engines are known from documents US 2 968 146 A, US 3 002 340 A, US 5 119 626 A, US 2006 / 242962 A1, US 2007 / 095972 A1, DE 41 11 396 A1 and US 2014 / 325958 A1.

[0005] The present invention relates to a jet engine according to claim 1 or to claim 4, which can be more economical to develop than the aforementioned traditional engine, namely the low-bypass turbofan with an afterburner. Figure 1 shows a diagram of the invention, hereinafter the ENGINE. In particular, two cross-sections are shown in planes that contain the shaft (13) of the ENGINE and are perpendicular to one another. The air flow is represented by arrows.

[0006] Although Figure 1 depicts the ENGINE in a schematic manner, the most significant elements are shown in a size representative of a real application. As the sizes of these components are very different, in Figure 2 an area of Figure 1 is reproduced on a larger scale, and in Figure 3 an area of Figure 2 is reproduced on a larger scale.

[0007] 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): A gas generator (11) into which a liquid fuel and a liquid oxidant are injected at high pressure in a fuel rich ratio, 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. A turbine (12) that is driven with the gases from the gas generator and moves a high rotational speed shaft (13). A gearbox (14) in which the rotational speed is reduced in order to adapt to the needs of the compressor (16).

[0008] The jet engine itself consists of the following elements: An air intake (15) according to the state of the art. Only one example is depicted in Figure 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 flow rate 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. 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. 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. 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. 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. 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.

[0009] 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 Figure 4. In contrast, the expansion of the exhaust gases is adjusted automatically as a result of the wedge-shape (27).

[0010] The ENGINE has three operating modes. The main mode is referred to as the normal mode, and is characterized by the following elements: A diffuser (22) at the outlet of the turbine. 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.

[0011] 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).

[0012] 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.

[0013] 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).

[0014] 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 (25) into the combustion chamber 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.

[0015] 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.

[0016] Like any other heat engine, a series of auxiliary elements are required, the following being the most important: An electrical machine in engagement with the compressor (16), or with one of the stages of the gearbox (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.

[0017] 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.

[0018] 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 through injectors (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.

[0019] 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.

[0020] For its building, 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.

[0021] Besides the elements already mentioned, the ENGINE will also require the following typical elements of a gas turbine for itsbuilding: 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.

[0022] 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 through injectors (25).

[0023] 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.

[0024] As can be seen in Figure 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 Figure 5. This figure shows a cross section perpendicular to the shaft of the ENGINE and in the position 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.

[0025] Figure 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 gearbox (29) in each compressor and a main gearbox with multiple outputs (28).

[0026] Figure 7 shows a diagram of the second option. In Figure 7, the direction of the flow of the gas generator has been intentionally inverted. In Figures 3 and 6 the flow is opposite the forward motion, in Figure 7 the flow of the gas generator is in the same senseas the forward motion. It should be noted that the ENGINE can work with the gas generator oriented in the forward direction or in the opposite direction.

[0027] The ENGINE in the configuration shown in Figure 7 can be simplified by removing the half, as shown in Figure 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. On aircraft performances this configuration introduces a moment that varies at different flight envelope points.

[0028] List of elements indicated in the drawings: 10 central body 11 gas generator 12 turbine 13 high rotational speed shaft 14 gearbox 15 air intake 16 axial compressor 17 transition area 18 combustion chambers 19 converging nozzles 20 mobile elements 21 diverging nozzles 22 turbine outlet diffuser 23 gasification chamber 24 injector array 25 injectors for the super-thrust mode 26 throats 27 wedge-shape 28 gearbox with several outputs 29 secondary gearbox 30 a mobile element 31 ramp

Claims

1. A jet engine for propelling an aircraft comprising an air intake (15), an axial compressor (16) downstream of the air intake, two combustion chambers (18) with a rectangular cross-section and a transition area (17) from the outlet of the compressor to the two combustion chambers, an exhaust downstream of the combustion chambers, a central body (10) in the center of the engine, two mobile elements (20), a gas generator (11) and a turbine (12), whereby the engine is configured such that liquid fuel and liquid oxidant are injected at high pressure into the gas generator; the turbine drives the compressor and is in turn driven by the gas generator; the exhaust is defined by two parallel and symmetrical conduits, the conduits not being of axial symmetry but having rectangular cross-sections ; on the outside of the engine, the exhaust is demarcated by the two mobile elements pivoting on a point and ending in a sharp edge ; in the center of the engine, there is the central body that separates both conduits; said central body widens as the exhaust progresses until the exhaust gases overcome the two mobile elements, from this point it narrows to form a wedge-shape ; by adjusting the angle of the two mobile elements, the cross-section of the two conduits can be varied to a great extent ; when the two mobile elements are separated from the central body, a more or less constant cross-section of the two exhausts which is similar to the cross- section of the two combustion chambers upstream of the exhaust is achieved; when the two mobile elements approach the central body, the two mobile elements together with the central body form two converging nozzles (19) ; by adjusting the angle of the two mobile elements, a small narrowing or a very pronounced narrowing with respect to the cross-section of the two combustion chambers can be achieved; in the area where the narrowing is maximum, that is, the throats (26) of the two nozzles, the central body, and the two mobile elements are rounded so as to form two diverging nozzles (21) with a small expansion ratio after the two throats; and as a result of the aforementioned rounding, the cross-section of the conduits always varies smoothly.

2. The jet engine as described in claim 1, comprising two, four, or six compressors (16) placed on the same plane, all of them driven with the turbine (12), which is a single turbine; the jet engine further comprising a main transmission box (28) that has an input shaft coming from the turbine, as many output shafts as the number of compressors installed and a secondary transmission box (29) in each one of the compressors.

3. The jet engine as described in claim 1, comprising two, four, or six compressors (16) placed on the same plane, whereby each compressor is driven with an independent turbine (12), and each turbine has its own gas generator (11).

4. A jet engine for propelling an aircraft comprising an air intake (15), one, two or three axial compressors (16) downstream of the air intake, a single combustion chamber (18) with a rectangular cross-section and a transition area (17) from the outlet of the one, two or three compressors to the single combustion chamber, an exhaust downstream of the combustion chamber, a single fixed ramp (31), a single mobile element (30), one, two or three gas generators (11) and one, two or three turbines (12); whereby the engine is configured such, that liquid fuel and liquid oxidant are injected at high pressure into the one, two or three gas generators; the one, two or three turbines drive the one, two or three compressors respectively and are in turn driven by the one, two or three gas generators respectively; the exhaust is defined by a single conduit, the conduit being not of axial symmetry, but having a rectangular cross-section; on the outside of the engine, the exhaust is demarcated by the single mobile element pivoting on a point and ending in a sharp edge; said single fixed ramp widens as the exhaust progresses until the exhaust gases overcome the single mobile element, from this point it narrows to form a wedge-shape; by adjusting the angle of the single mobile element, the cross-section of the single conduit can be varied to a great extent; when the single mobile element is separated from the single fixed ramp, a more or less constant cross-section of the exhaust which is similar to the cross-section of the single combustion chamber upstream of the exhaust is achieved; when the single mobile element approaches the single fixed ramp, the single mobile element together with the single fixed ramp forms a single converging nozzle (19), by adjusting the angle of the single mobile element, a small narrowing or a very pronounced narrowing with respect to the cross-section of the single combustion chamber can be achieved; in the area where the narrowing is maximum, that is, the throat (26) of the nozzle, the single fixed ramp, and the single mobile element are rounded so as to form a single diverging nozzle (21) with a small expansion ratio after the throat; and as a result of the aforementioned rounding, the cross-section of the single conduit always varies smoothly.