Detonation supercharging turbine engine
An aero-engine and detonation technology, applied in the direction of engine control, machine/engine, mechanical equipment, etc., can solve the problem of further increasing the temperature of the turbine inlet gas, and achieve the effect of reducing weight, reducing difficulty, and increasing thrust-to-weight ratio
Active Publication Date: 2021-12-31
SHANGHAI JIAO TONG UNIV
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Problems solved by technology
Also constrained by material technology and cooling technology, it becomes more and more difficult to further increase the gas temperature at the turbine inlet
Method used
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[0019] The present invention will be further described below in conjunction with the embodiments and accompanying drawings, but the protection scope of the present invention should not be limited thereto.
[0020] see figure 2 , figure 2 It is a structural schematic diagram of a detonation supercharged engine of the present invention, as shown in the figure, including a compressor 1, an intake rotary valve 2, a detonation tube combustion chamber 3, an outlet rotary valve 4, a turbine 5, a casing 6, and a rotor shaft 7 and speed regulating device 8.
[0021] The detonation tube combustion chamber 3 is composed of a plurality of detonation tubes 31 , and each detonation tube 31 is provided with a fuel nozzle 32 and an igniter 33 . In this embodiment, 16 detonation tubes 31 are fixed on the engine casing 6 and distributed in a radial circumferential array to form 16 channels extending axially along the detonation tube combustion chamber 3 .
[0022] The air intake rotary val...
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A detonation supercharging turbine engine comprises a gas compressor, a gas inlet rotary valve, a detonation pipe combustion chamber, a gas outlet rotary valve, a turbine, a case, a rotor shaft and a speed adjusting device, the front end and the rear end of the detonation pipe combustion chamber are provided with a gas inlet rotary valve dynamic seal and a gas outlet rotary valve dynamic seal respectively, and the gas inlet rotary valve and the gas outlet rotary valve are fixed to the rotor shaft through the speed adjusting device. Along with rotation of a rotor, a detonation pipe is fixed to the engine case, a plurality of channels are axially formed and distributed in a radial circumferential array mode, and after gas enters the detonation pipe and is mixed with fuel oil sprayed out of a fuel oil nozzle, the gas is ignited and combusted through an igniter. A gas inlets and a gas outlet are distributed in a staggered mode, and it is guaranteed that gas is detonated in the detonation pipe for constant-volume combustion. The gas inlet rotary valve is located downstream of the gas compressor, and the gas outlet rotary valve is located upstream of the turbine. The combustion chamber with the large slenderness ratio is used and matched with the inlet and outlet rotary valves, constant-volume detonation combustion is achieved, and compared with an isobaric combustion engine, the engine circulation efficiency is improved, the gas compressor and turbine design is simplified, and the engine performance is improved.
Description
technical field [0001] The invention relates to the technical field of aviation turbine engines and gas turbines, in particular to a single-rotor detonation supercharged turbine engine. Background technique [0002] In conventional axial-flow aero-engines, increasing the boost ratio of the compressor and the gas temperature before the turbine is an important way to increase the ideal cycle power of the engine and improve the thermal efficiency. design to achieve. Restricted by material technology, the higher the speed, the greater the centrifugal force on the rotor blades, so it is impossible to increase the speed infinitely; and the tip of the blade is easy to form a supersonic region to generate shock waves, resulting in shock wave loss. There are also theoretical limits to advanced airfoil designs. Also constrained by material technology and cooling technology, it becomes more and more difficult to further increase the gas temperature at the turbine inlet. [0003] ref...
Claims
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IPC IPC(8): F02C3/04F02C3/14F02C9/16F23R7/00
CPCF02C3/04F02C3/14F02C9/16F23R7/00F05D2270/05Y02T10/12
Inventor 孟宇屈应保陈方谢敏骐
Owner SHANGHAI JIAO TONG UNIV