Distributed turbines for reducing aerodynamic drag in engines with high bypass ratios

Abstract

An aircraft engine comprising at least one core engine (A) and at least one auxiliary engine (B), a propulsor (2) (blower, propeller or compressor), a compressor (3), a combustion chamber (4) and a turbine (5), wherein downstream of the compressor (3) the compressed air is guided via an intermediate pipe (9) into at least a second combustion chamber (4B), downstream of which at least one further turbine (6) is located and directly or indirectly drives at least one propulsor (2).

Description

[0001] The present invention relates to aerospace engineering, in particular to the design and optimization of aircraft engines with high bypass ratios.

[0002] Current technology strives for continuous improvements in specific fuel consumption within the framework of specified flight missions. Specific fuel consumption is comprised of weight, drag, and overall efficiency. Overall efficiency, in turn, consists of thermal efficiency and propulsive efficiency.

[0003] An increase in thermal efficiency can be achieved through improvements in the cycle or through component improvements. However, these improvements are often associated with high development costs and currently usually result in only small efficiency gains. Larger improvements in the cycle, such as combined cycles with energy recovery, require large structural volumes, which can increase drag and weight. This can offset the benefits in the context of the overall flight mission.

[0004] The last significant improvements in overall aircraft engine efficiency have been achieved through higher bypass ratios in propulsive efficiency. A higher bypass ratio means that a greater proportion of the air flowing through the engine is diverted around the core engine rather than flowing through the combustion chamber and turbine. This results in more efficient thrust generation, as the propulsor (e.g., a fan or propeller) accelerates a larger airflow at a lower speed, reducing specific fuel consumption. However, a higher bypass ratio also results in larger engine diameters, which increases weight and drag. It also creates ground clearance issues, particularly for low-wing aircraft, making wing integration more difficult.The present invention aims to minimize these disadvantages by proposing innovative measures to reduce the aerodynamic nacelle and core fairing drags.

[0005] According to one aspect of the present invention, an aircraft engine consists of at least one core engine and at least one auxiliary engine, a propulsor, a compressor, a combustion chamber and a turbine, wherein downstream of the compressor the compressed air is guided via an intermediate pipe into at least one second combustion chamber, downstream of which at least one further turbine is located and drives at least one propulsor directly or indirectly.

[0006] According to another aspect, the flow of the working gases in the secondary combustion chamber and turbine flows counter to the flow of the propulsor.

[0007] According to a further aspect, the mass flow in the turbines can be controlled via at least one mechanical adjustment, such as a valve or an adjustable guide vane, wherein the installation position of the mechanical adjustment is irrelevant.

[0008] According to a further aspect, the flow from the auxiliary engine is partially or completely directed outwards through the bypass flow via at least one pipe.

[0009] According to a further aspect, the sheath flow stators and / or nacelle suspension elements serve as pipes for guiding the exhaust gas flow to the outside.

[0010] According to a further aspect, the tubes serve as heat exchangers to transfer heat from the exhaust gas to the bypass stream.

[0011] According to another aspect, the tubes discharge part or all of the exhaust flow directly into the bypass stream via a dedicated outlet or indirectly through perforated stators.

[0012] According to a further aspect, the exhaust gas flow from the power turbines is directed to the outside via at least one heat exchanger in the bypass flow.

[0013] According to a further aspect, the exhaust gas flow from the power turbines is partially or completely mixed with the bypass flow within the bypass duct.

[0014] According to a further aspect, the outlet temperatures of at least one of the respective turbines of the respective shafts are controlled to enable selective non-catalytic reduction upstream of at least one expansion nozzle, wherein the application of selective non-catalytic reduction requires the integration of further selective non-catalytic reduction components, such as ammonia injection nozzles, between the respective turbines and expansion nozzles.

[0015] According to a further aspect, a selective catalytic reduction system is integrated within the flow guide, wherein the application of selective catalytic reduction requires the integration of further selective catalytic reduction components, such as ammonia injection nozzles, between the respective turbines and catalysts.

[0016] The present invention is explained below by way of example with reference to the accompanying drawings, in which identical reference numerals designate identical or similar components. In the highly schematically simplified figures: Fig. 1 shows a schematic representation of an aircraft engine according to the present invention.

[0017] Today's aircraft engine technology continually strives for improvements in specific fuel consumption within the context of a given flight mission. The specific fuel consumption of an aircraft engine is comprised of the engine weight, its drag, and the overall efficiency of the engine. The overall efficiency, in turn, consists of the thermal efficiency and the propulsive efficiency.

[0018] An improvement in thermal efficiency can be achieved through optimizations in the cycle or through component improvements. However, these improvements require high development costs and are often only achievable today in the decimal percentage range. Larger improvements in the cycle, such as combined cycles with energy recovery, are associated with large structural volumes, which increase drag and weight and can thus offset the benefits in the context of the overall flight mission.

[0019] The last significant improvements in overall aircraft engine efficiency were achieved by increasing the bypass ratio in propulsive efficiency. However, this led to larger engine diameters, increasing weight and drag. It also created ground clearance problems for low-wing aircraft, making wing integration more disadvantageous.

[0020] The present invention aims to minimize the swept nacelle surfaces of an engine, reduce the swept surface of the core engine cowling, counteract weight gain, and mitigate ground clearance issues at higher bypass ratios. Furthermore, the invention enables a potential improvement in thermal efficiency.

[0021] The invention involves separating the core engine from the power turbine, with the auxiliary combustion chambers and auxiliary turbines arranged in counterflow. This arrangement allows for simultaneous shortening of the engine nacelle and the core engine cowling, reducing aerodynamic drag and maximizing the benefits of distributed propulsion systems.

[0022] Additionally, the invention incorporates a control of the mass flow through the core engine via a mechanical adjustment, allowing combustion chamber temperatures to be optimized over a wide operating range. This leads to an increase in thermal efficiency and enables the integration of a selective non-catalytic reduction system downstream of the turbines.

[0023] By arranging the flow of the working gases in the secondary combustion chamber(s) and turbine(s) against the flow of the propulsor, a more compact engine design is possible. This configuration results in a shortened nacelle and core engine cowling, which reduces aerodynamic drag. This contributes to improved propulsive efficiency, as less energy is required to overcome drag.

[0024] The counterflow arrangement of the working gases also enables more efficient use of exhaust energy, as the flow of the working gases can be optimally tailored to the downstream turbines. This leads to improved thermal efficiency and thus a reduction in specific fuel consumption.

[0025] Another advantage of this arrangement is improved component cooling, as the working gases are directed in the opposite direction to the propulsor flow. This can increase the service life of the turbines and combustion chambers and reduce maintenance costs.

[0026] In one embodiment, the aircraft engine may be configured so that the flow of the working gases in the secondary combustion chamber (4B) and the secondary turbines (6B) is counter to the flow of the propulsor (2), enabling improved efficiency and performance. A possible variation of this embodiment could include the use of different materials for the secondary combustion chamber and the secondary turbines to increase thermal resistance and component lifespan. Another embodiment could involve the integration of additional cooling mechanisms to regulate the temperature of the working gases and thus improve the efficiency of the entire system. Furthermore, in another embodiment, the aircraft engine could be designed so that the flow of the working gases is controlled by a special arrangement of guide vanes or valves to precisely regulate the mass flow and optimize the propulsor performance.Another possible variation could be the implementation of a modular design in which the auxiliary combustion chamber and auxiliary turbines can be easily replaced or serviced to increase the maintainability and adaptability of the aero engine. Finally, the aero engine could also be configured to optimize the flow of the working gases through a combination of mechanical and electronic control systems to ensure maximum efficiency and performance under different operating conditions. In one embodiment, the aero engine may be designed so that the tubes (11) serve as heat exchangers to transfer heat from the exhaust gas to the bypass stream (7). These tubes can be made of various materials that have high thermal conductivity, such as aluminum or copper, to maximize the efficiency of heat transfer.In another embodiment, the tubes (11) could have a special coating that is corrosion-resistant and at the same time improves heat transfer. Another possible embodiment could include the integration of flexible tubes that can adapt to different operating conditions in order to optimize heat transfer. In addition, the tubes (11) could be designed in various geometric shapes, such as spiral or wave-shaped, in order to increase the surface area for heat transfer. In a further embodiment, the tubes (11) could be equipped with internal flow deflectors that direct the exhaust gas flow and thus increase the efficiency of heat transfer. Finally, the tubes (11) could also be modular so that they can be easily replaced or maintained in order to extend the service life of the aircraft engine and reduce maintenance costs.In one embodiment, the aircraft engine includes a selective catalytic reduction (SCR) system integrated within the flow guide. This system can have various configurations to maximize efficiency and adaptability. One possible embodiment provides for the SCR components, such as ammonia injectors, to be arranged between the respective turbines and catalysts to ensure optimal reduction of nitrogen oxides. In another embodiment, the catalysts could be designed in various geometric shapes and materials to increase the reaction surface and enhance catalytic efficiency. A further embodiment could include the integration of additional sensors and control systems that adjust the ammonia injection rate in real time based on exhaust gas parameters to enable precise and efficient emission reduction.Additionally, the system could be configured to operate in different modes to adapt to different flight conditions, for example, by switching between a high-performance mode during takeoff and a low-power mode during cruise. These different embodiments demonstrate the versatility and adaptability of the SCR system within the aircraft engine to meet regulatory emissions requirements while maximizing engine efficiency.

Claims

[1] An aircraft engine, consisting of at least one core engine (A) and at least one auxiliary engine (B), a propulsor (2) (blower, propeller or compressor), a compressor (3), a combustion chamber (4) and a turbine (5), wherein downstream of the compressor (3) the compressed air is guided via an intermediate pipe (9) into at least one second combustion chamber (4B), downstream of which at least one further turbine (6) is located and drives at least one propulsor (2) directly or indirectly. [2] The aircraft engine according to claim 1, wherein the flow of the working gases in the secondary combustion chamber(s) (4B) and turbine(s) (6B) flows counter to the flow of the propulsor (2). [3] The aircraft engine according to claim 1, wherein the mass flow in the turbines can be controlled via at least one mechanical adjustment (10), such as a valve or an adjustable guide vane, wherein the installation position of the mechanical adjustment is irrelevant. [4] The aircraft engine according to claims 1 or 2, wherein the flow from the auxiliary engine(s) (B) is partially or completely guided outwards via at least one pipe (11) through the bypass flow (7). [5] The aircraft engine according to claim 4, wherein the bypass stators and / or nacelle suspension elements serve as pipes (11) for guiding the exhaust gas flow to the outside. [6] The aircraft engine according to claim 4 or 5, wherein the tubes (11) serve as heat exchangers to transfer heat from the exhaust gas into the bypass flow (7). [7] The aircraft engine according to claim 4 or 5, wherein the tubes (11) discharge part or all of the exhaust gas flow directly into the bypass flow (7) via a dedicated outlet (12) or indirectly through perforated stators. [8] The aircraft engine according to claims 1 or 2, wherein the exhaust gas flow from the power turbines (6) is directed to the outside via at least one heat exchanger (11) in the bypass flow (7). [9] The aircraft engine according to claims 1 or 2, wherein the exhaust gas flow from the power turbines (6) is partially or completely mixed with the bypass flow (7) within the bypass flow channel. [10] The aircraft engine according to claims 1 or 3, wherein the outlet temperatures of at least one of the respective turbines (5 and 6) of the respective shafts are controlled so as to enable a selective non-catalytic reduction (SNCR - 13) upstream of at least one expansion nozzle (8), wherein the application of an SNCR requires the integration of further SNCR components (13), such as ammonia injection nozzles, between the respective turbines and expansion nozzles. [11] The aircraft engine according to claims 1 or 2, wherein a selective catalytic reduction system (SCR - 13) is integrated within the flow guide, wherein the application of an SCR requires the integration of further SCR components (13), such as ammonia injection nozzles, between the respective turbines and catalysts.

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