RU2724559C1 - Turbojet aircraft engine - Google Patents

Abstract

FIELD: engine building.SUBSTANCE: invention relates to aircraft engine building and can be used in development of jet engines intended for flight of aircrafts in atmosphere due to implementation of detonation thermodynamic cycle with high repetition rate of pulses and self-maintenance of fuel detonation process. In the known turbojet aircraft engine comprising an air intake, a gas generator, a nozzle and a detonation device adjacent to the outer engine housing, upon a proposal, detonation device is made in the form of a longitudinal annular combustion chamber with an inlet air channel from the engine air intake and an axial swirler at its inlet, flow part of the chamber is separated from the flow part of the gas generator, and on the inner side of the external wall of the annular combustion chamber there are serially arranged along the longitudinal axis of the engine belts of fuel and air mixers with ignition devices in them, wherein perimeter of each belt there are at least three air and fuel mixers with ignition devices, made in the form of tube sections, the inputs and outputs of which are directed along the air flow direction coming from the axial swirler, and axes of fuel nozzles in mixer are directed at an angle to direction of air flow in it, besides, each separate belt of mixers is tuned to its normalized mode of pulse detonation.EFFECT: application of the engine allows increasing fuel efficiency up to 30 %, qualitatively increasing flight speed up to 5 M, providing constant specific fuel pulse 2,000–2,500 s at operation on hydrocarbon fuel without significant increase of dimensions and weight, increasing efficiency, reducing cost due to simplification of design.6 cl, 3 dwg

Description

The invention relates to the field of aircraft engine manufacturing and can be used to create jet engines designed for flying aircraft in the atmosphere due to the implementation of a detonation thermodynamic cycle with a high pulse repetition rate and self-sustaining fuel detonation process.

Known turbojet engine containing a gas generator, a nozzle and afterburner combustion chamber located between the gas generator and the nozzle.

(G.S. Skubachevsky. Aircraft gas turbine engines. Design and calculation of parts. M: Mechanical Engineering, 1969, p. 5, Fig. 1.01). /1/

The afterburning of fuel in the afterburner of the combustion chamber occurs at low pressure and therefore has low efficiency. Therefore, to obtain additional traction, you have to spend a huge amount of fuel.

The closest in technical essence and the achieved result is a known turbojet aircraft engine containing an air intake, a gas generator, a nozzle and a detonation device adjacent to the body of the outer circuit of the engine.

/ RU No. 2277181 IPC F02K 3/10 Published on May 27, 2006 /. / 2 /

The detonation device according to this invention allows fuel to be burned in gas-dynamic resonators, providing detonation combustion in them having maximum efficiency. Such an engine has a fairly high specific parameters. However, due to the fact that the resonators are parallel to each other, the transverse dimensions of the detonation device increase, which increases the dimensions of the aircraft and its aerodynamic losses. In addition, the throttle control capabilities of the detonation device are very limited, since each particular device reliably operates in a narrow range of air pressure supplied to it, and this depends on the flight speed of the aircraft on which this engine is installed.

Thus, the disadvantage of the known engine is the low cost of fuel combustion at low pressure in forced operation, significant dimensions and weight of the structure, which leads to an increase in engine resistance and does not provide a sufficient supply of traction without increasing the required fuel consumption.

The problem to which the invention is directed, is to increase the efficiency of a turbojet engine at supersonic flight speeds in a wide range of changes.

The expected technical result is an increase in the efficiency of fuel combustion and a decrease in its consumption, reduction of harmful emissions, a significant increase in specific thrust without a significant increase in size and weight.

The specified technical result is achieved by the fact that in the known turbojet aircraft engine containing an air intake, a gas generator, a nozzle and a detonation device adjacent to the housing of the outer circuit of the engine, on the proposal, the detonation device is made in the form of a longitudinal annular combustion chamber with an air inlet from the engine air intake and axial swirl at its entrance, the flow part of the chamber is separated from the flow part of the gas generator, and on the inner side of the outer wall of the annular combustion chamber are sequentially placed along the longitudinal axis of the engine of the belt of the fuel and air mixers with ignition devices in them, while at least around the perimeter of each belt three air and fuel mixers with ignition devices, made in the form of pipe sections, the inputs and outputs of which are directed along the direction of the air flow coming from the axial swirler, and the axis of the fuel nozzles in the mixer are directed at an angle to the regulation of the air flow in it, with each individual belt of the mixers configured to its normalized mode of pulse detonation. For a double-circuit or three-circuit turbojet engine, the flow part of the second or subsequent engine circuits can be selected as the external circuit. A turbojet aircraft engine can be equipped with a system for switching the subsequent belt of fuel and air mixers into operation after switching off the previous belt, and the air channel inlet from the engine air intake into the detonation combustion chamber is equipped with a controlled shutter. A longitudinal annular combustion chamber may be provided with a supersonic jet nozzle at the outlet, and the engine is provided with an annular shell mounted between the body of the external circuit of the engine and a longitudinal annular combustion chamber, forming a slit-like gap for supplying cooling air from the engine intake.

The implementation of the detonation device in the form of a longitudinal annular combustion chamber adjacent to the housing of the outer contour of the engine, allows to reduce the transverse dimensions of the engine by placing all the necessary elements of the device along its longitudinal axis, which has maximum dimensions in it.

The connection of the longitudinal annular combustion chamber with the air channel from the engine intake and the axial swirler at its inlet allows not only high pressure air to be received at the inlet to the detonation combustion chamber, but also twisted in the right circumferential direction for the detonation chamber to work.

Separation of the flow part from the flow part of the gas generator allows their operation to be autonomous, while pulse detonation in the detonation chamber cannot directly affect the operation of the gas generator, for example, without affecting the rather sensitive blades of the gas generator compressor.

Placing on the inner side of the outer wall of the annular combustion chamber sequentially along the longitudinal axis of the engine of the belts of mixers (resonators) of fuel and air with at least three mixers along the perimeter of each belt, they can be placed uniformly with the smallest radial dimensions and ensure reliable operation of gas-dynamic resonators.

The implementation of the mixers in the form of pipe segments, the inputs and outputs of which are directed along the direction of the air flow coming from the axial swirler and the direction of the axes of the fuel nozzles in the mixer at an angle to the direction of air flow in it, allows the resonators to create forces directed at an angle to the longitudinal axis engine.

Setting each individual mixer belt to its normalized mode of pulse detonation allows this belt to operate at its own pressure in the tube, and therefore at its own flight speed of the aircraft with maximum efficiency, and go out when moving away from this pressure.

Besides:

a) for a double-circuit or three-circuit turbojet engine, the flow part of the second or subsequent engine circuits is selected as the external circuit, which makes this device universal;

b) the engine can be equipped with a system for switching on the operation of the subsequent belt of fuel and air mixers after disconnecting the previous belt;

c) the air channel inlet from the engine’s air intake into the detonation combustion chamber can be equipped with a controlled damper, which will allow the turbojet aircraft engine to shut the gas generator out of operation and create thrust in the most efficient way, directing the entire air flow through the detonation device;

d) a longitudinal annular combustion chamber may be provided at the outlet with a supersonic jet nozzle;

e) for cooling the outer loop housings and the detonation chamber housing, a slit-like gap may be formed between the outer loop housing and the longitudinal annular combustion chamber for supplying cooling air there from the engine air intake. In addition, such a gap is an excellent damper from detonation vibrations from a detonation device to a gas generator.

In FIG. 1 shows a longitudinal section of a turbojet engine;

In FIG. 2 shows a longitudinal section of a turbojet engine with a slit-like gap between the casing of the external circuit of the engine and the longitudinal annular combustion chamber.

In FIG. 3 diagram of an annular detonation combustion chamber with a traveling wave.

A turbojet aircraft engine contains an air intake 1, a gas generator 2, a jet nozzle 3, a detonation device 4, adjacent to the housing 5 of the outer circuit 6 of the engine. The detonation device 4 is made in the form of a longitudinal annular combustion chamber 7 with an air channel from the engine air intake and an axial swirler 8 at its inlet, the flow part of which is separated from the gas generator flow part. On the inner side of the outer wall of the annular combustion chamber 7, sequentially placed along the longitudinal axis of the engine, the belts of the mixers (resonators) 9 of fuel and air with ignition devices 10 in them. At least three air and fuel mixers are installed along the perimeter of each belt, made in the form of pipe segments, the inlets and outlets of which are directed along the direction of the air flow coming from the axial swirler 8. The fuel nozzles 11 are mounted on the mixers 9, and their axes are directed at an angle to the direction of the air flow in the mixers, with each individual belt of the mixers configured to its normalized mode of pulse detonation. The entrance of the air channel from the engine intake to the detonation device 4 with the combustion chamber 7 is provided with a controlled shutter 12. An annular shell 13 is installed between the outer loop body 5 of the engine and the longitudinal annular combustion chamber 7, which forms a slit-like gap 14 for supplying cooling air from the engine intake.

To create a traveling wave 17 when implementing a detonation thermodynamic cycle with a high pulse repetition rate, in each individual belt, the air and fuel mixers are tuned to their own mode of pulse detonation and self-sustaining detonation.

The combustion chamber 7 of the detonation device 4, is equipped with several belts 18, at least three mixers 9 with an open input and output are installed along the perimeter of each belt and which have an internal volume and a special configuration sufficient to initiate an instant explosion in them.

When creating the conditions for the detonation thermodynamic cycle, the mixers of each given belt 18 receive the amount of fuel and gaseous oxidizer necessary for the ignition conditions, while the amount of fuel and gaseous oxidizer in the mixer is determined by the speed of arrival of the oxidizer flow 15 or the speed of the aircraft. (Mach number) . After the passage of the swirler 8, the swirling stream 16 enters the mixers of the first belt of the combustion chamber 7, into which fuel is supplied via the fuel nozzles 11. Ignition devices 10 ignite the combustible mixture to initiate combustion, which transforms into an instantaneous explosion. The swirling shock wave 17 created by this explosion travels at a supersonic speed and contributes to the creation of a secondary pressure, temperature and composition of the environment, causing ignition of the fuel and gaseous oxidizer and the initiation of a subsequent instantaneous explosion in the current or next belt of the mixers. The resulting total shock wave 17 improves the specific thrust of the engine.

The proposed engine provides for the implementation of the mixers 9 in the form of pipe segments bent along the arc of the housing of the combustion chamber 7, and installed in it in the direction of the vortex flow at an angle to the longitudinal plane of the combustion chamber. This allows you to assemble engines with a multi-tube design of mixers (resonators) with a length of 700 mm to 1500 mm. The detonation device 4 of the engine allows you to work effectively at a speed of up to 5 Mach, while the proposal does not limit the number of belts with mixers 9 in the combustion chamber 7, but provides for the adjustment of the mixers (resonators) of each belt to the necessary (normalized) Mach number. The invention provides for the installation of an annular shell 13 in the engine, which smooths impulse shock loads on turbine elements and reduces acoustic impacts on the environment.

An engine equipped with a detonation device operates as follows.

At high flight speeds, the oncoming high-speed air flow is inhibited in the curved space of the air intake. After the aircraft reaches a speed of 0.8-1.0 M, the automation system (not shown) opens the controlled shutter 12 and the valve of the fuel injectors 11. After passing through the swirl 8, part of the swirling stream of gaseous oxidizer 16 enters the mixers (resonators) of the first belt tuned to 0.8-1.0 M. The fuel and gaseous oxidizer in the resonators form a mixture sufficient to ignite. The glow plug 10 ignites the mixture and initiates its burning, which transforms into an instantaneous explosion. The shock wave created by this explosion travels at a supersonic speed of the order of 4000 m / s. The temperature in the resonators rises abruptly to approximately 2800 ° C and the pressure rises abruptly in the mixers of the first belt and at a certain distance behind it.

Since the creation of the environment in the resonators causing ignition of the fuel and the gaseous oxidizer after an instant explosion is determined by the temperature, the rate of creation of the secondary pressure and the feed rate of the gaseous oxidizer determined by the speed (Mach number) of the aircraft, the ignition of the fuel and the initiation of the subsequent instant explosion in the resonators the first belt occurs only at a speed of 0.8-1.0 M. If the speed of the aircraft is higher, then the automation closes the fuel supply to the mixers of the first belt and opens the flow to the second belt with resonators tuned to large Mach numbers and so on, up to speeds Mach number 5. A transition from belts with a higher Mach number to belts with a lower value is possible.

The use of the engine makes it possible to increase fuel efficiency by up to 30%, to qualitatively increase flight speed to 5 M, to provide a constant specific impulse for fuel of 2000-2500 s when working on hydrocarbon fuel without a significant increase in size and weight, increase efficiency, and reduce costs by simplifying the design.

Claims (6)

1. A turbojet aircraft engine containing an air intake, a gas generator, a nozzle and a detonation device adjacent to the housing of the external engine circuit, characterized in that the detonation device is made in the form of a longitudinal annular combustion chamber with an air inlet from the engine air intake and an axial swirl at its inlet, the flow part of the chamber is separated from the flow part of the gas generator, and on the inner side of the outer wall of the annular combustion chamber are sequentially arranged along the longitudinal axis of the engine of the belt of fuel and air mixers with ignition devices in them, while at least three air and fuel mixers are installed along the perimeter of each belt ignition devices made in the form of pipe segments, the inputs and outputs of which are directed along the direction of the air flow coming from the axial swirler, and the axis of the fuel nozzles in the mixer are directed at an angle to the direction of air flow in it, and each individual The mixers are tuned to their normalized pulse detonation mode. 2. A turbojet aircraft engine according to claim 1, characterized in that for the dual-circuit or three-circuit turbojet engine, the flow part of the second or subsequent engine circuits is selected as the external circuit. 3. A turbojet aircraft engine according to claim 1, characterized in that the engine is equipped with a system for switching on a subsequent belt of fuel and air mixers after turning off the previous belt. 4. A turbojet aircraft engine according to claim 1, characterized in that the air channel inlet from the engine air intake into the detonation combustion chamber is provided with a controlled damper. 5. A turbojet aircraft engine according to claim 1, characterized in that the longitudinal annular combustion chamber is provided at the outlet with a supersonic jet nozzle. 6. A turbojet aircraft engine according to claim 1, characterized in that it is equipped with an annular shell mounted between the body of the outer contour of the engine and a longitudinal annular combustion chamber, forming a slit-like gap for supplying cooling air there from the engine intake.

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