Single-shaft turbojet engine

Abstract

The invention discloses a single-shaft turbojet engine. Comprising an air inlet necking used for initial acceleration of air flow, a fan provided with a blade profile suitable for the supersonic speed of the air flow, a starting-power generator arranged in the center of a fan wheel and based on a brushless direct current motor, a multi-stage axial flow air compressor used for compressing the air flow of a core machine, and an annular combustion chamber. And a turbine having active blades and having one or more wheels. The fan, the starter-generator, the axial flow compressor and the turbine are all mounted on a single shaft without reduction gears. Parameters of the blades, the fan, the compressor and the turbine are optimized, and the gear train does not need to be decelerated. The technical scheme can be used for engines with the bypass ratio as high as 8. According to the invention, the characteristics of weight and size of the turbojet engine can be significantly improved while good efficiency is maintained.

Description

[0001] This invention relates to turbojet engines. Its most effective application is in low-thrust (1000 to 5000 N) turbojet engines.

[0002] A low-power, single-shaft, single-jet engine is known (Kurt Schreckling, Model turbines 2005, Traplet Publications Ltd). This engine comprises a centrifugal compressor and a single-wheel turbine mounted on the same shaft, featuring a simple and compact structure. It is started using a detachable starter. This engine has become a prototype for most low-power propulsion systems used in ultralight unmanned aerial vehicles and model aircraft. The disadvantages of this known design include low propulsion efficiency, low reliability, and long pre-start preparation time.

[0003] A low-power single-shaft gas turbine engine (USSR Invention Certificate 165039) is known. This engine is equipped with a centrifugal compressor and a radial turbine mounted on the same shaft, and the starter is connected via a single-stage gear drive and a freewheel clutch. The disadvantages of this known design are: low power plant efficiency due to the use of a single-stage centrifugal compressor and a single-stage radial turbine; and reduced reliability of the propulsion system due to the installation of the starter via a freewheel clutch and gear drive.

[0004] A single-shaft turbojet engine (patent RU2674172C1) is known and is considered the closest to the prior art based on several features. This engine includes a compressor, a turbine, a main continuous detonation combustion chamber with a fuel supply passage, fuel nozzles, and a detonation initiator, an aerodynamic damper, a nozzle assembly, and a turbine. Four continuous detonation combustion chambers are arranged within the engine's after-vortex chamber, each in a closed annular sector shape, with its inlet connected to the outer bypass air duct. The main combustion chamber inlet is connected to the air duct of the engine core. The engine bypass ratio is equal to or greater than 1. An electric motor is mounted on the engine shaft and connected to a battery pack. Additionally, an annular piezoelectric generator (divided into two parts by a damper) is also connected to the battery pack. Annular cooling jackets, equipped with penetrating matrix nozzles and fuel nozzle assemblies, are installed around the inner surfaces of the main combustion chamber and the after-vortex combustion chamber. Annular cooling chambers, equipped with matrix nozzles and fuel nozzle assemblies, are installed to cool the outer surfaces of the four after-vortex combustion chambers in contact with the turbine exhaust heat flow. This invention aims to improve engine thermal efficiency, thrust, reduce weight and size, and expand the thrust vector variation range.

[0005] The known drawbacks of this design include: low reliability due to structural complexity, vibration caused by the unsteady process of detonation combustion, short lifespan, low durability of piezoelectric elements, and the need for additional pneumatic dampers to worsen weight and size characteristics.

[0006] The purpose of this invention is to develop a small-thrust turbojet engine design with a thrust of 1000 to 5000 N, which is as compact as possible, easy to manufacture, and suitable for mass production, applicable to ultralight aircraft and similar products.

[0007] The way to achieve the above goal is to optimize the design of the fan, axial compressor, turbine impeller and starter-generator so that they operate at the same speed, so that they can be installed on the same rotor shaft without the need for speed-increasing or speed-reducing transmissions.

[0008] To address this task, an intake duct is installed at the engine turbine inlet, reducing the fan impeller diameter and thus lowering the circumferential velocity of the fan blades. The blade shape is designed to work not only for subsonic but also for supersonic airflow. A multi-stage axial compressor is installed within the core engine. Compared to centrifugal compressors, the axial compressor design reduces the engine diameter and eliminates the need for a large bypass rectifier (i.e., a "volute"), allowing for a bypass ratio as high as 8 within a compact size. A brushless DC motor is used as the starter-generator, employing high-energy-density magnets to achieve high specific power. Its high efficiency eliminates the need for a complex, large cooling system; the absence of contact points grants it a long lifespan. The starter-generator rotor is rigidly fixed to the engine shaft. The turbine impeller uses active blades, operating at speeds lower than comparable reactive or impulse-reactive turbines, allowing for fan-matched speeds and coaxial mounting without a reduction gear system. While the use of active blades results in a slight efficiency loss, it allows for low circumferential speed operation, and the unused internal energy in combustion products is converted into useful work in the nozzle assembly.

[0009] The essence of the invention is illustrated by a longitudinal cross-sectional view of the engine (see...) Figure 1 ): Position 1: Intake bypass duct for accelerating airflow, used to supply air to fan wheel (position 2); fan wheel is fitted onto starter-generator rotor (position 3); stator is fixed to core engine bracket (position 5); rotor is mounted on engine hollow shaft (position 6); compressor stage wheel (position 7) and intermediate airflow rectifier (position 8) are installed on the shaft in sequence; shaft passes through annular combustion chamber (position 9), its working area is isolated by inner shell (position 10); turbine wheel (position 11) is installed at shaft end; combustion products enter wheel via guide (position 12); airflow enters engine core nozzle assembly (position 14) containing central body (position 15) via rectifier (13); airflow forced into outer bypass duct (position 16) by fan passes through outer bypass duct nozzle assembly (position 17).

[0010] This technical solution can achieve: maximize structural compactness; maintain relatively high efficiency; simplify design to the greatest extent; ensure high reliability (especially critical for small thrust engines); and eliminate complex process nodes, facilitating rapid mass production.

Claims

1. A single-shaft turbojet engine, comprising an inlet constriction, a fan with a supersonic blade profile, a brushless DC starter-generator integrated at the center of the fan impeller, a multi-stage axial compressor for the core engine, a combustion chamber, and a turbine with active blades and containing one or more stages, characterized in that, To ensure a compact design, the fan supplying air to the dual-duct system, starter-generator, axial compressor, and turbine stage are all mounted on a single shaft, without the use of any gearboxes.

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