RU2447308C2 - Bypass turbojet engine with airflow redistribution at inlet - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises inlet device, low- and high-pressure compressors, combustion chamber, low- and high-pressure turbines, mixer of circuit flows, and nozzle. Engine inlet accommodates two flow-twist-controlled-diffusers and incoming flow power redistribution in radial direction. Cylindrical combustion chamber is arranged between first and second diffusers. Second diffuser is arranged directly at low-pressure compressor inlet. Multi-vane diffusers have their vanes arranged radially. Fixed parts of first diffuser vanes thrust by their one end against central body while another end thrust against chamber channel inner surface front section. Fixed parts of second diffuser vanes thrust by their one end low-pressure compressor cone while another end thrust against chamber channel inner surface rear section. First diffuser is equipped with rear shaped turning sections of vanes to deflect airflow. Second diffuser is equipped with front shaped turning sections of vanes to deflect airflow coupled with appropriate drives and with fixed sections of vanes via moving joints. Drive are incorporated with common engine control system.

EFFECT: increased thrust, optimised engine operating conditions.

5 dwg

Description

The invention relates to the field of aviation technology, in particular to a turbojet double-circuit engine (turbofan) of an aircraft, and can be used as a power plant in other industries.

A known method of short-term boosting a turbojet engine by injection using nozzles of liquid (water) at its entrance (Theory and calculation of jet engines / Edited by S.M. Shlyakhtenko. M.: Engineering. 1987. pp. 374 ... 376) .

The disadvantage is the significant weight and flow rate of the water to reduce the temperature at the engine inlet, which negatively affects the range and duration of the aircraft and can therefore only be used on take-off.

The vortex effect is known (M.E.Deich. Technical gas dynamics. M .: Energy. 1974, p. 460 ... 470), used for separation into cold and hot air flows. However, this effect is mainly used in refrigeration and heat power plants and heat supply systems.

The prototype is a turbojet dual-circuit engine with an axial compressor (Prototype diagram. Yu.N. Nechaev P.M. Fedorov. Theory of aircraft gas turbine engines. Part II. M.: Mechanical engineering. 1978, Page 233. Fig. 17.2.)

The prototype has significant features: low and high pressure compressors, a combustion chamber, low and high pressure turbines, a mixer for flow flows, a nozzle.

The disadvantage is the absence at the engine inlet of the redistribution of energy of the incoming air flow in the radial direction and the provision of additional optimization of the temperature and pressure at the inlet of the air flow compressor to increase the traction characteristics of the turbojet bypass engine at maximum operating speed, when the flight speed (free flow) changes, and when launched on the ground and in flight.

The technical result to which the invention is directed is:

- to increase the traction characteristics of a turbojet bypass engine at maximum operating mode;

- providing the opportunity to optimize the degree of separation of the incident air flow by energy when changing the flight speed, engine operating mode, as well as when starting on the ground and in flight.

To achieve this technical result, in a turbojet two-circuit engine with redistribution of energy of the air flow at the inlet containing low and high pressure compressors, a combustion chamber, low and high pressure turbines, a circuit flow mixer, a nozzle, two guides adjustable in angle of air flow twist are installed at the inlet apparatus (ON) and a chamber for redistributing the energy of the incoming air flow in the radial direction. The camera is made in the form of a cylindrical channel, and this camera is installed between the first and second HA. The guide vanes are multi-blade, the blades of which are mounted radially. The fixed parts of the blades of the 1st guide vane abut at one end against the central body, the other against the inner surface of the chamber channel in its front part, and the second guide vane abut against one side of the stationary coke of the low pressure compressor, and the other against the inner surface of the chamber channel the end, and the first ON is equipped with rear profiled rotary parts of the blades deflecting the air flow, and the second ON is equipped with front profiled rotary parts of the blades deflecting the air flow d associated with the fixed parts of the blades of the guide vanes by means of movable joints and with the corresponding drives. In this case, the drives are included in the general engine control system, depending on the flight conditions and the operating modes of the latter.

The second ON is installed immediately before entering the low-pressure compressor.

This embodiment of the claimed engine will allow to increase the traction characteristics of the engine at maximum operating mode while maintaining the possibility of its optimal operation in the entire range of operation.

List of figures in the drawings

The invention is illustrated by drawings, in which:

figure 1 shows a turbojet dual-circuit engine with energy redistribution at the inlet (longitudinal section), where the arrows show the direction of flow of "hot" air (Tg) and "cold" (Tx);

figure 2 shows the guide apparatus (ON), front view of figure 1;

figure 3 is a diagram of the deviation of the adjustable blades of the 1st and 2nd guide vanes to create a stationary vortex, where V is the direction of air flow, 1-PK is the first impeller of the low pressure compressor, U is the direction of its rotation;

figure 4 is a curve of the temperature of the gas in sections along the engine path;

figure 5 is a curve of the change in gas pressure in sections along the engine path.

In Figs. 4 and 5, the broken n-a1-b1-b1-gd3-e2-g corresponds to the internal circuit, n-a2-b2-b2-gd3-e1-g corresponds to the internal circuit in the absence of a stationary vortex at the input , n-a3-b3-d2-e2-g - to the external circuit, n-a2-b2-d1-e1-g - to the external circuit in the absence of a stationary vortex at the inlet.

Information confirming the possibility of carrying out the invention

A turbojet dual-circuit engine with redistribution of energy of the air flow at the inlet contains an input device that includes a central body - 1 with a shell - 17 and a 1st guiding apparatus with front stationary - 2 and rear profiled rotary parts of the blades deflecting the air flow - 18, and rotation drive - 3, 2nd guide vane with rear fixed - 5 and front deflecting air flow profiled rotary parts of the blades - 19, controlled by the drive - 6, camera for energy aspredeleniya formed into a cylindrical channel - 4 disposed between the first and second ON. The 2nd guiding apparatus in the direction of air flow is installed behind the chamber - 4, directly at the inlet to the low-pressure compressor (LPC) - 7, with a stationary coke (central body) - 20. The cylindrical housing - 8 divides the channels - 9 and - 10, respectively external and internal contours. Inside the housing - 8 there is a high-pressure compressor (HPC) - 11, a combustion chamber 12, low-pressure turbines - 13 and high - 14 pressures, and outside the external circuit of the engine is in the form of an annular channel. The output device contains a mixer of gas flows of the circuits - 15 and the nozzle - 16, (see figure 1, 2).

The guide vanes are multi-blade, the blades of which are mounted radially. The fixed parts of the blades of the 1st guiding apparatus - 2 abut with one end against the central body - 1, and the other against the inner surface of the channel of the chamber - 4 in its front part, and the fixed blades of the 2nd guiding apparatus - 5 abut with one end against the stationary coke - 20, and others into the inner surface of the channel of the camera - 4 at its end. Deflected air flow profiled rotary parts of the blades - 18 and - 19 are connected with the respective drives and with the fixed parts of the blades of the guide vanes by means of movable joints.

The essential features characterizing the invention and different from the essential features of the prototype is the following.

1. Adjustable in swirl angle from 0 to max of the 1st ON to translate the incoming air flow from the axial direction to the rotational one.

2. A chamber for the formation of a stationary vortex and redistribution of flow energy in the radial direction.

3. Adjustable in the swirl angle from 0 to max of the 2nd ON to transfer the incoming air flow from the rotational direction to the axial one optimal for the operation of the 1st impeller of the low pressure compressor.

4. The presence of the regulation system of the 1st and 2nd guiding devices, depending on the flight conditions and engine operating conditions.

4. The presence of the regulation system of the 1st and 2nd guiding devices, depending on the flight conditions and engine operating conditions.

Work turbofan.

At the maximum operating mode, an air stream having a temperature T * n and pressure P * n (take-off or flight conditions), which enters the engine inlet through the input device, is swirled using the blades of the 1st ON (see Figs. 1, 2), moreover, the twist angle using the control system is set so as to provide the maximum possible for the given flight conditions and engine operating conditions in the chamber 4 redistribution of the total energy of the air in the radial direction (the so-called vortex effect), namely in the peripheral region of the chamber s temperature and pressure increased by ΔTg, ΔRg and, respectively, in the axial region they fall on the camera ΔTh, ΔRh. From the energy redistribution chamber, the air flow by controlling the 2nd HA is transferred from the rotational direction to the axial direction, optimal for the operation of the 1st impeller of the low pressure compressor (LPC) 7, (see figure 3). A further process of compressing the air flowing through the low-pressure compressor 7 preserves the flow structure along the radius, since the flow lines are parallel to the axis of the engine. At the exit of the KND 7, the peripheral ("hot") air flow enters the external circuit 9, while its pressure and temperature are:

P * ext. = P * n + ΔPr + ΔPknd,

T * ext. = T * n + ΔTg + ΔTknd,

Where:

P * external - pressure in the external circuit - 9;

P * n - free flow pressure;

ΔРг - the magnitude of the pressure increase at the output of the peripheral zone of the flow energy redistribution chamber — 4;

ΔРкнд, ΔТкнд - increase in pressure and temperature in the low-pressure compressor - 7;

while the prototype, these parameters are less by ΔТг, ΔРг.

Thus, an increase in pressure and temperature in the external circuit - 9 will lead to an increase in these parameters behind the mixer - 15. At the same time, the degree of expansion of the gas in the nozzle - 16 will increase, which will lead to an increase in the outflow rate and ultimately to an increase in the thrust of such an engine. At the outlet of the KND - 7, the axial "cold" air stream enters the internal circuit - 10, where it is additionally compressed with a high pressure compressor (HPC) - 11 to Рmax and then in the combustion chamber - 12, at the pressure indicated above, it is heated to a temperature Tmax, limited by the strength characteristics of the blades of the high-pressure turbine - 14, (see figures 4 and 5). The work of the inner loop cycle is increased due to the possibility of a greater degree of heating, i.e. supplying additional fuel to the combustion chamber and increasing the degree of pressure increase through the use of more pressure HPC - 11. In cruising flight modes, when engine boost is not required, the separation of the input stream by energy is intermediate. In this case, the flow swirl in the chamber may be partial or absent. This feature is provided by the deviation of the profiled rotary parts of the blades - 18 and - 19 of the 1st and 2nd guide vanes, synchronously rotated using the drives - 3 and - 6 of the control system for the swivel angle from max to 0. In the engine starting modes on the ground and in flight (in autorotation modes), the 1st and 2nd guiding devices are switched to the position where there is no flow swirl to provide increased pressure and air temperature in the combustion chamber to ignite the fuel.

Claims (1)

A turbojet dual-circuit engine with redistribution of energy of the air flow at the inlet containing the input device, low and high pressure compressors, a combustion chamber, low and high pressure turbines, a circuit flow mixer, a nozzle, characterized in that two flow-controlled swirl angles are installed at its input device guides (ON) and a camera for redistributing the energy of the incoming air flow in the radial direction, and this camera is made in the form of a channel of cylindrical shape and and between the first and second ON, the second ON is installed directly in front of the low-pressure compressor inlet, the guide vanes are multi-vane, the blades of which are mounted radially, the fixed parts of the vanes of the 1st guide vane abut one end on the central body, the other on the inner surface of the chamber channel in it of the front part, and the 2nd guide vane abut at one end into the stationary coke of the low-pressure compressor, with the other into the inner surface of the chamber channel at its end, and the 1st the rear deflecting air flow profiled rotary parts of the blades, and the 2nd AN is equipped with front deflecting air flow profiled rotary parts of the blades associated with the respective drives and the fixed parts of the blades of the guide vanes by means of movable joints, while the drives are included in the general engine control system in depending on the flight conditions and operating modes of the latter.

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