Low-thrust turbofan engine

Abstract

A turbofan engine consisting of a convergent inlet duct, a fan, a starter-generator, a core in which the fan supplying an airflow to the engine acts as a low-pressure compressor, a high-pressure compressor, a combustion chamber, a turbine, a nozzle assembly for the core, and a bypass duct with a nozzle assembly. The starter-generator is disposed inside the blade wheel of the fan, wherein the blades of the fan have a profile adapted for operation at supersonic speed, the fan, the starter-generator, the high-pressure compressor and the turbine are mounted on a single rotor with no reduction gearing, fuel is injected directly into an annular hot chamber via special ports by injection nozzles having a continuous spray angle, said injection nozzles being disposed inside the housing of the combustion chamber and being blown with a cooling airflow, wherein the housing of the injection nozzles does not touch the walls of the hot chamber. The invention makes it possible to produce a structurally compact engine and thus achieve a high power-to-weight ratio, while maintaining good power plant efficiency.

Description

[0001] Low thrust twin-screw turbofan engine

[0002] The invention relates to turbojet dual-circuit fan engines. Its application is most effective in low-power turbojet dual-circuit fan engines with a thrust of 1000 to 5000 N.

[0003] A large number of different designs of turbojet fan double-circuit jet engines are known. Their designs are usually improved by improving one of the critical units of the product for the local task being solved, while insufficient attention is paid to the remaining units of the product, which reduces the design time, but inevitably leads to a non-optimal design of the product as a whole. The present invention is the result of searching for the optimal design of a turbojet double-circuit fan engine of low thrust by selecting the most advantageous technical solutions for the main units of the engine to solve the task.

[0004] For example, a two-circuit turbojet engine is known (patent RU 2701034 C1). A two-circuit turbojet engine consists of an inlet device, a fan, an inner circuit, and an outer circuit. Inside the inner circuit are a compressor, a combustion chamber, a turbine, gas channels of a heat exchanger-regenerator, and an exhaust pipe. Inside the outer circuit are air channels of a heat exchanger-regenerator, a nozzle. The exhaust pipe is located at the inlet of the fan.

[0005] The disadvantages of the known solution are high gas-dynamic losses in the heat exchanger-regenerator, relatively large dimensions and, as a result, low specific power, as well as the complexity of the design, which limits the possibility of using this solution for low-power engines.

[0006] A dual-circuit gas turbine engine is known (patent RU 2384723 C2). The engine comprises an outer circuit formed by the outer engine casing and the compressor casing, a casing installed in the outer circuit, covering part of its flow section and forming a cavity with one of the casings, a heat exchanger with fins located in this cavity and communicating with high- and low-pressure sources by coolant supply and discharge channels. The cavity is formed between the outer casing and the casing, and the heat exchanger fins are made on one of them. The coolant discharge channel communicates with the low-pressure source to the cavity in front of the heat exchanger fins, and the coolant supply channel communicates with the high-pressure source to the cavity behind the heat exchanger fins. The heat exchanger fins are annular, and at least one axial groove is made in each of its fins. The grooves of adjacent ribs are offset relative to each other.The invention is aimed at solving the problems of heat exchange between engine circuits.

[0007] The disadvantage of this solution is the complexity of the heat exchanger design. The invention also does not consider the layout of the remaining elements of the engine design relative to the heat exchanger.

[0008] A bypass turbojet engine is known (patent RU 2353790 C1) accepted for a number of features as the closest analogue. An aircraft bypass turbojet engine contains a fan, a high-pressure compressor, a combustion chamber, high- and low-pressure turbines, a mixer and an afterburner and a nozzle common to both circuits. Behind the first stage of the high-pressure compressor, a constantly open annular channel with a straightening grid is made, through which, in all engine operating modes, a part of the air is bypassed from behind the stage into the cocurrent air flow of the outer circuit behind the fan. The invention increases the bypass ratio of the engine and improves its efficiency in non-afterburner operating modes. The engine has one nozzle block for the outer and inner circuits. The scheme of existing engines RD-33, AL-31F, EJ-200, M.88, F404, F414, F125 is adopted as a basis for the invention.

[0009] The disadvantages of this design are: efficiency losses due to the bypass of part of the air flow after the compressor stage into the external circuit, relatively low specific power due to the need to use a large external engine shell.

[0010] The aim of this invention is to optimize the main design components of a low-thrust dual-circuit turbojet engine operating in the thrust range from 1000 to 5000 N, in order to obtain a compact design with high specific power at an acceptable overall engine efficiency.

[0011] The essence of the invention is explained by the attached drawing of a longitudinal section of the engine, see Fig. 1, on which the following are marked with positions: 1 - inlet confuser; 2 - fan for supplying air to both circuits, which also functions as a compressor for the inner circuit of the engine; 3 - fairing of the central part of the fan; 4 - starter-generator, located in the inner part of the fan blade wheel; 5 - position designates the racks of the inner and outer circuits, the internal cavities of which are intended to accommodate communications of the engine control system and fuel supply to the combustion chamber; 6 - axial multi-stage high-pressure compressor with alternate arrangement of working wheels and fixed air flow straightening grates; 7 - flow part of the outer circuit of the engine; 8 - manifold for supplying fuel to the combustion chamber with a block of injectors; 9 - annular-type combustion chamber; 10 - combustion chamber; 11 - nozzle block of the outer circuit; 12 - turbine;13 - inner circuit nozzle block; 14 - central fairing of the inner circuit nozzle block; 15 - rotor shaft, 16 - rotor bearing supports.

[0012] The stated objective is achieved by the following design solutions used in the design of a two-circuit turbojet fan engine: an inlet confuser used for preliminary acceleration of the air flow supplied to the fan impeller, a fan containing wide-chord blades in two profiling options either for movement at a supersonic air flow speed or for movement at speeds close to supersonic depending on the planned area of ​​application of the engine, a starter-generator located inside the fan impeller with a rotor fixedly connected to the impeller hub, the fan supplies air to both circuits and is used for the inner circuit as a low-pressure compressor, then the inner circuit consists of an axial multi-stage compressor between the stages of which flow straighteners are located,annular combustion chambers, fuel injection into which is carried out by means of a fuel manifold with injectors with a continuous spray sector, injection is carried out through special holes directly into the combustion chamber, while the injectors do not come into contact with the walls of the combustion chamber, in addition, a part of the air flow flowing around the combustion chamber from its outer side is used for their cooling, to improve the quality of mixing of the atomized fuel, if permissible within the scope of application of the engine, it is possible to use high-pressure injectors designed for operation with a pressure drop from 2.5 to 20 MPa, a turbine consisting of several working wheels or it is possible to use one working wheel with active-type blades, in the case of using a single-stage turbine, the nozzle block of the inner circuit is made with a central body serving for efficient operation of the nozzle block in a wide range of pressure drops and reducing the overall length of the engine,the outer contour of the engine has its own nozzle block, which also allows to reduce the length of the outer contour shell, if appropriate, depending on the scope of application of the engine, the nozzle blocks are combined into one. In addition, the engine uses a gearless installation scheme for the turbine, compressor stages, fan and starter-generator, which allows for a compact engine layout and relieves the rotor bearing supports from axial loads. A brushless DC electric machine is used as a starter-generator, the main advantages of which are: high efficiency of at least 94%, high reliability and high specific power, placing a starter-generator of another type in the fan impeller is impossible due to its relatively large dimensions and rapid failure in the presence of contact units.

[0013] The engine operates as follows. In the starting mode, the starter-generator 2 spins the rotor shaft 15 to the speed corresponding to the start of fuel supply to the combustion chamber. Then, using the fuel pump, the fuel is pumped into the fuel manifold 8 and injected into the combustion chamber 10 through the nozzles; the fuel is ignited using ceramic glow plugs. The starter-generator continues to spin the engine rotor to the minimum operating speed, after which its power supply is disconnected. Further spinning of the rotor is performed using the turbine 12 due to the energy of fuel combustion. The electronics of the starter control system, upon reaching the idle speed of the rotor, switch the windings of the starter-generator to the rectifier and subsequently, during engine operation, it is used to generate electrical energy for the on-board power supply of the aircraft. During engine operation, the fan 2 pumps air flow into the outer and inner circuits.The fan bushing is covered with a fairing 3 to reduce engine resistance to the oncoming air flow. The air flow is preliminarily accelerated in the confuser 1 before being fed to the fan. The communications of the control system and the fuel supply system in the engine are pulled through hollow fairings 5. If it is necessary to increase the engine thrust, the volume of fuel supplied to the combustion chamber 9 is increased in the afterburner mode, it is possible to switch the starter-generator electronics to the rotor spin-up mode, the thrust can be briefly increased to 10% of the maximum operating value. The engine is stopped by stopping the fuel supply to the fuel manifold, it is also possible to use the starter in the rotor braking mode by generating a magnetic field with a reverse direction of rotation on its winding using the electronic control system.

[0014] The technical solutions adopted in the present invention make it possible to design the engine compactly, thus obtaining a high specific thrust of the engine while maintaining good efficiency of the power plant.

Claims

6 Invention formula 1. A two-circuit turbofan engine consisting of an inlet confuser, a fan, a starter-generator, an inner circuit in which a fan is used as a low-pressure compressor, supplying air flow to the engine, a high-pressure compressor, a combustion chamber, a turbine, a nozzle block of the inner circuit, an outer circuit with a nozzle block, characterized in that the starter-generator is located inside the fan blade wheel, and the fan blades have a profile operating at supersonic speed, the fan, starter-generator, high-pressure compressor and turbine are mounted on a single rotor without the use of a gearbox, fuel injection is carried out directly into the combustion chamber of the annular fuel through special openings by nozzles with a continuous spray sector located in the combustion chamber body and blown by a cooling air flow, while the nozzle body does not touch the walls of the combustion chamber.

2. A two-circuit turbofan engine according to claim 1, characterized in that the fan wheel is made with blades with a profile designed for operation at subsonic speeds.

3. A two-circuit turbofan engine according to any of paragraphs 1 and 2, characterized in that the turbine contains one stage with active-type blades.

4. A two-circuit turbofan engine according to any of paragraphs 1, 2 and 3, characterized in that high-pressure injectors from 2.5 to 20 MPa are used to supply fuel to the combustion chamber, depending on the purpose of the engine.

5. A two-circuit turbofan engine according to any of paragraphs 1, 2, 3 and 4, characterized in that it has one common nozzle block for the outer and inner circuits.

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