RU2424439C1 - Method for protecting turboejector engine against stall - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: bypass to low pressure channel of gas ejector at closed inlet of the channel is performed. At total degree of pressure increase in conditions of takeoff 3.5÷4.0 of pressure increase degree in compressor before and after air takeoff are preferably equal. Bypass is preferably performed at flight speeds of less than two Mach numbers. Fuel can be supplied to low pressure channel of gas ejector.

EFFECT: possibility of excluding emission of hot gas to compressor inlet without deteriorating the engine characteristics.

4 cl, 2 dwg

Description

The invention relates to aircraft engine manufacturing.

Known turbojet engine (patent RU 2190772, IPC 7 F02C 3/32, 1999) containing an input device, a compressor, a main combustion chamber, a turbine, a gas ejector, the high pressure channel of which is connected on one side to the compressor through the main combustion chamber, and on the other hand, with a turbine through the mixing chamber, the low pressure channel is connected to the atmosphere on the one hand through the inlet device, and on the other hand, with the turbine through the mixing chamber, afterburner, and the outlet device.

In a turbojet engine (TRDE), there is a possibility of hot gases entering the compressor inlet (through the low pressure channel of the gas ejector), which can lead to surging of the engine. The gas ejection occurs when the ejector is blocked, when the dimensions of the jet of ejection gas become equal to the dimensions of the cross section of the mixing chamber (ejection coefficient m = 0).

The aim of the invention is to eliminate the possibility of surge TRDE due to the ingress of hot gases at the inlet to the compressor.

A known jet engine (certificate of utility model RU 22311, 2002), which uses a fan, part of the air through a low pressure channel of the gas ejector is transferred to the mixing chamber.

A known method of controlling an axial compressor by air bypass from the middle stages (Theory and calculation of jet engines. / Under the editorship of S.M. Shlyakhtenko. M.: Engineering, 1987, Fig. 4.13, p.119). Air bypass (air exhaust) is carried out at reduced rotor speeds (start, low gas). The method allows to increase the stability margins of the compressor.

The goal is achieved by the fact that at flight speeds of less than two Mach numbers, air from the middle stages of the compressor is transferred to the low pressure channel of the gas ejector, the entrance to which is closed.

Preferably, with a total degree of pressure increase under take-off conditions of 3.5 ÷ 4.0, the degrees of pressure increase in the compressor before and after air sampling are equal.

Preferably, the bypass is carried out at flight speeds of less than two Mach numbers.

Preferably, fuel is supplied to the low pressure channel of the gas ejector.

The essence of the invention lies in the fact that the differential pressure between the high and low pressure channels of the gas ejector decreases when the air is bypassed (if the degree of increase in pressure in the compressor is equal before and after selection, the pressure drop becomes less critical), which leads to an increase in the ejection coefficient (more than twice) - excludes the possibility of locking the mixing chamber (emission of hot gases).

At flight speeds M> 2, the ejection coefficient in the turbofan engine becomes more than 10%, at which the mixing chamber does not lock.

To increase engine thrust, additional fuel can be supplied to the low pressure channel of the gas ejector. In this case, the combustion of fuel will occur on both sides of the combustion chamber.

Figure 1 shows a diagram of TRDE, illustrating a method of protecting the engine from surging.

Figure 2 shows the characteristic of the gas ejector in the TRDE system.

TRDE consists of an input device 1, a compressor 2, a low pressure channel of a gas ejector 3 (second circuit), a damper 4, a main combustion chamber 5, a mixing chamber 6, a turbine 7, an output device 8. The degree of increase in pressure in the compressor under take-off conditions is 3 5 ÷ 4.0 (Written V.L.Concept of a gas turbine engine for hypersonic flight speeds. // Polet, 2009, No. 8, p.19-23).

The method is as follows.

At flight speeds M <2, the shutter 4 blocks the entrance to the channel 3 and opens the air from the middle stages of the compressor to the same channel 3 (Fig. 1, top view). In this case, the air flow through the channel 3 increases. The increase in air flow rate is associated with an increase in the productivity of the first stages of the compressor, which turns out to be higher than that which the engine had before opening the damper 4. An increase in air flow through channel 3 leads to an increase in the ejection coefficient to 10% or more, at which the mixing chamber does not lock.

Of course, bypassing air at maximum rotational speeds degrades the compressor's performance (puts the compressor on an off-design mode of operation), but this practically does not affect the characteristics of the engine up to flight speeds of M≈2. The deterioration of compressor characteristics is compensated by an increase in air flow rate and a decrease in losses in the mixing chamber (subsonic flow rates of gas from the first and second circuits).

At speeds M> 2, the compensation effect disappears, but at the same time, the ejection coefficient of the TRDE is already higher than 10%, which excludes locking of the mixing chamber. The damper 4 is returned to its original position (figure 1, bottom view).

Figure 2 shows the characteristic of the ejector in the TRDE system (π _ej is the degree of increase in pressure in the ejector, π _to is the degree of increase in pressure in the compressor, m is the ejection coefficient). Here the curve rg connecting the limit points of the lines π _to = const is the line of critical regimes, the curve rr is the working line. The closing moment of the shutter 4 is marked on the characteristic by a horizontal line. It can be seen that the air bypass allows to exclude the operation modes of the ejector with low stability margins (dash-dotted lines).

The application of the method improves flight safety when using turbojet engines.

Claims (4)

1. A method of protecting a turbojet engine from surging, which consists in bypassing air from the middle stages of the compressor, characterized in that the bypass is carried out in the low pressure channel of the gas ejector with a closed entrance to the channel. 2. The method of protecting the turbojet engine from surge according to claim 1, characterized in that, with a total degree of pressure increase under take-off conditions of 3.5 ÷ 4.0, the degrees of pressure increase in the compressor before and after air sampling are equal. 3. A method of protecting a turbojet engine from surge according to claim 1, characterized in that the bypass is carried out at flight speeds of less than two Mach numbers. 4. A method of protecting a turbojet engine from surge according to claim 1, characterized in that fuel is supplied to the low pressure channel of the gas ejector.

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