JPS59211723A - Gas turbine engine - Google Patents

Abstract

PURPOSE:To stabilize the combustion of fuel for after-burning, by bleeding a part of air from a compressor and introducing the part of air as combustion air for fuel for after-burning. CONSTITUTION:Upon take-off of an aircraft, after-burning is performed for acceleration. The amount of unburnt oxygen is less in burnt gas when the inlet gas temperature of a turbine 6 is high. Therefore, the output of the turbine 6 is increased to open a shut-off member 10 for bleeding a part of air from a compressor 4, and therefore, the bled air is led as combustion air to an after-burner 7 through a passage 9. Burnt gas the temperature of which is raised by this after-burning is expanded through a jet nozzle to become high-speed jet stream, thereby high-thrust power can be obtained. With this arrangement combustion of fuel for after-burning may be stabilized.

Description

[Detailed description of the invention] The present invention relates to gas turbine engines.

Among aviation gas turbine engines, turbojet engines are in practical use, for example, for supersonic flight. This turbojet engine uses a compressor to easily compress air that has been compressed to ram pressure at one engine speed, and then leads this compressed air to a combustor to burn fuel. After expanding with a turbine that drives
A jet nozzle inflates it to atmospheric pressure and generates a high-speed jet that provides propulsion for the aircraft.

Generally, in order to increase thermal efficiency in gas turbine engines including turbojet engines, it is necessary to increase the maximum cycle temperature, that is, the turbine inlet temperature. However, if the turbine inlet temperature is limited by the heat resistance of the multi-bin and the turbine blade has a low allowable temperature limit, the gas temperature is lowered by mixing dilution air into the combustion gas in the combustor. Also, if the turbine blade has a high allowable temperature limit, the combustion gas will not be diluted.

Furthermore, combustion in the combustor is performed at a near stoichiometric mixture ratio, and high-temperature combustion gas is guided to the turbine.

On the other hand, with the development of aircraft with fewer gas turbine engines,
It is necessary to temporarily increase thrust during takeoff or when accelerating to break the speed of sound, so Taverning of the 77 is carried out for Jin. This afterburning is achieved by injecting fuel into the combustion gas whose temperature has been lowered by expansion in the turbine, and burning this fuel with unburned oxygen contained in the combustion residue to bring the combustion gas down to 1 degree Celsius. High-temperature combustion gas is expanded through a jet nozzle to generate powerful propulsive force.

However, as mentioned above, in order to increase the thermal efficiency of the gas turbine, as the maximum temperature of the four cycles is increased, that is, the turbine inlet temperature is set to h, the combustion at the combustion base becomes close to the stoichiometric mixture ratio. The amount of unburned oxygen contained in the combustion gas decreases. If the oxygen-containing layer in the combustion gas is small, the combustion of the fuel for afterburning becomes unstable and even becomes impossible.

The present invention has been made in view of the above circumstances, and its object is to provide a gas that enables stable 77 turbaning when the turbine inlet temperature is high and the oxygen layer contained in the combustion gas is small. To provide a turbine engine.

An embodiment of the present invention will be described below with reference to the drawings.

The turbojet engine 1 has an air intake 2° and a rotor 8.
Pressure machine installed on the niT side of 4. Combustor 5, mouth 3
A turbine 6, a taverna 7, and a jet nozzle 8 are provided at the rear of the engine.

The air that has also flowed into the air intake port 2 is compressed by the compressor 4 and sent to the combustor 5. The fuel supplied in the burner 5 is combusted, and the combustion gas of 0 aai pressure generated is expanded in the turbine 6. Drives the turbine 6.

Through the rotor 8 by the turbine 6 the pressure +TIIf1g
V@4 is driven. The combustion gas exiting the turbine 6 expands in the jet nozzle 8 and becomes a high-velocity jet, generating propulsive force.

Reference numeral 9 denotes a flow path for extracting a portion of compressed air from an intermediate stage of the compressor 4 and guiding it to the Ama4I burner 7 at the rear of the engine 1. Reference numeral 10 denotes an opening/closing member, and the front end 9m of the flow path 9 communicates with the bleed part 11 of the compressor 4 via the opening/closing member 10. The rM closed line member 10 is composed of, for example, a sliding shielding plate operated by hydraulic pressure. Rear end 9 of channel 9
b is arranged so as to discharge compressed air into the combustion gas passage 12 of the afterburner 7, or arranged so that the compressed air and the fuel supplied for afterburning are mixed in advance. Ru.

Afterburning is performed by 1J when the aircraft is landing or accelerating beyond the speed of sound, but as mentioned above,
When the inlet gas temperature of the turbine 6 is low, the combustion in the combustor 5 is carried out near the thermal mixture ratio, so the amount of unburned oxygen contained in the combustion gas exiting the turbine 6 is small, and therefore afterburning occurs. When performing this, the output of the turbine 6 is increased, and a part of the compressed air is extracted from the pressure=Xta4 by adjusting the opening/closing member 10, and this extracted compressed air is used as combustion air to be sent to the afterburner 7 via the flow path 9. In the afterburner 7, the compressed air burns the fuel, and this combustion increases the 'lff of the combustion gas exiting the turbine 6.

The combustion gas, whose temperature has been increased by this afterburning, expands in the jet nozzle 8 and becomes a nose-velocity jet, generating a strong propulsive force. When the aircraft flies at cruising speed, there is no need for afterburning, so #1
11 member 10 to stop the extraction of compressed air.

As explained above, in the present invention, a means is provided for extracting a part of the compressed air from the pressure machine and guiding it as combustion air for the afterburner, and afterburning is performed using the extracted compressed air. As a result, as a result of increasing the turbine inlet temperature to increase the heat and efficiency of the gas turbine engine, it is possible to stably perform afterburning when there is a small layer of unburned oxygen in the combustion gas. can.

[Brief explanation of the drawing]

The drawing is a cross-sectional view showing an embodiment of the present invention. l...turbojet engine, 4...pressure: 1sfi,
5. Combustor, 6. Turbine, 7. Afterburner, 8. Jet nozzle, 9. Flow path, 10. Opening/closing means. 11...Air extraction part.

Claims (1)

[Claims] (1) In a gas turbine engine equipped with a compressor, a combustor, a turbine, an afterburner, and a jet nozzle, there is provided a means for extracting a portion of compressed air from the compressor, and a means for extracting the extracted compressed air. and means for directing fuel as combustion air in the afterburner.