RU2716992C2 - Annular combustion chamber of gas turbine engine and method of arrangement of working process therein - Google Patents

Abstract

FIELD: machine building.SUBSTANCE: annular combustion chamber of gas turbine engine includes external and internal walls, which are made in the form of curvilinear surfaces, corner flame stabilizer. On the angular flame stabilizer rear side there is profiling in the form of two curvilinear surfaces. Outlet part of combustion chamber includes gas flow collector, which is a channel formed by collector blades and walls of combustion chamber. Gas collector consists of a cascade of separate outlet branch pipes restricted with housings which, as they approach the turbine, are combined into a single annular channel.EFFECT: invention is aimed at possibility to change excess air factor in annular combustion chamber with transverse system of reverse current zones formation and preventing formation of stagnant zones.5 cl, 7 dwg

Description

The invention relates to a device for burning fuel in jet engines, and in particular to the field of gas turbine engines, in particular to the combustion chambers of gas turbine engines, which can be used both in aerospace and industrial thermal power plants.

To date, a widespread method of organizing a fuel-air mixture using swirls of gas flow, which are often used as front-end devices [1]. A swirler is a device having elements that swirl air or a fuel-air mixture to carry out the combustion process in the combustion chamber. As the air stream passes through the blades of the swirl, it swirls and forms a vortex flow. The swirling flow in a swirling flow leads to the formation of a circulation zone in the central region of the flow if the reported swirling flow becomes large. The circulation zone created in this way provides high-quality mixing of fuel and air, since the rotational velocity components create areas of strong shear of the flow with a high level of turbulence and a high mass transfer rate. This method of organizing a fuel-air mixture is used in practice to increase the stability and intensity of combustion in modern combustion chambers of gas turbine engines.

However, the use of flow swirls dictates the need for many separate front-end devices along the circumference of the annular channel of the combustion chamber with a certain distance between these devices. Thus, between front-end devices of this type there is always an inefficiently used space through which the air supply from the compressor and the organization of the fuel-air mixture are impossible. In addition, the flow swirls form behind them many separate vortex structures having a longitudinal swirl. These structures interfere with each other, which worsens the homogeneity of the gas flow parameters around the perimeter of the annular channel.

A device is known - the combustion chamber of a gas turbine engine of the ring type. In the classic version, this device is an annular space formed by the outer and inner housings arranged coaxially. The closest technical solution to the declared and adopted as a prototype is the device described in the patent [2], where the angular flame stabilizer is used as the front device. The transition from multiple flow swirls to a single front-end device around the entire perimeter of the combustion chamber allows the use of previously inaccessible sections, thereby increasing the throughput of the combustion chamber. In turn, the transition from the longitudinal system of the formation of zones of reverse currents to the transverse system allows you to organize continuous vortex structures along the entire circumference of the combustion chamber and significantly reduce its axial length. However, this option contains disadvantages, such as:

- the inability to set the coefficient of excess air in the combustion chamber in a wide range;

- the presence of a stagnant zone behind the corner flame stabilizer, which, as a result, will lead to the formation of zones of local re-enrichment of the fuel-air mixture and an increase in the level of emission of harmful substances.

The task of the invention is to remedy the above disadvantages.

The task is achieved due to the fact that the annular combustion chamber of a gas turbine engine contains an angular flame stabilizer, an external and an internal wall, which are made in the form of curved surfaces. The difference is that the back side of the corner flame stabilizer has profiling in the form of two curved surfaces to prevent stagnant zones, and the back side of the corner flame stabilizer contains a cascade of additional profiled elements located around the circle to eliminate local vortex structures in this area. The output of the combustion chamber contains a gas flow manifold, which is a channel formed by the blades of the manifold and the walls of the combustion chamber, and the blades of the manifold are corrugated in the form of additional profiling around the circumference of the output edges. Also, the output part of the combustion chamber contains a gas collector, consisting of a cascade of separate outlet pipes, limited by housings, which, as they approach the turbine, are combined into a single annular channel.

The task is also achieved by the method of organizing the working process in the annular combustion chamber, which consists in swirling the air coming from the compressor behind the front device, thereby forming reverse current zones necessary for organizing the combustion process of the fuel-air mixture. The difference is that the flow is swirled behind the front device by a single solid front along the entire annular channel in the transverse direction relative to the longitudinal axis of the engine, in the output region of the combustion chamber the gas flow is divided into two parts, one part is forcibly directed to the reverse current zone, the second part is fed directly to the combustion zone. They prevent the occurrence of stagnant zones and local vortex structures on the back of the front device by geometrical influence on the flow structure. Organize the exit of gases from the combustion chamber in the form of many separate jets, which are then combined into a single stream.

The proposed annular combustion chamber of a gas turbine engine is shown in FIG. 1-7. FIG. 1 is a longitudinal section through a combustion chamber made in plane III-III of FIG. 6-7. FIG. 2 is a longitudinal section through a combustion chamber made in plane IV-IV of FIG. 6-7. FIG. 3 is an isometric view of the front of the combustion chamber. FIG. 4 is an isometric view of the rear of the combustion chamber. FIG. 5 is a side view of a combustion chamber. FIG. 6 is a partial isometric view of a cross section of a combustion chamber made in plane I-I shown in FIG. 5. FIG. 7 is a partial isometric view of a cross section of a combustion chamber made in plane II-II of FIG. 5.

In FIG. 1 shows a longitudinal section through a combustion chamber, the secant plane of which coincides with plane III-III in FIG. 6-7. The combustion chamber 1 includes an outer 2 and an inner 3 wall, in the input part of the combustion chamber there is a diffuser part in which an angular flame stabilizer 4 is located, which has a profiling 5 on the back side to prevent the formation of a stagnant zone. The output part of the combustion chamber contains a gas flow manifold consisting of a plurality of external 6 and internal 7 channels formed by external 8 and internal 9 blades of the manifold, as well as external 2 and internal 3 walls of the combustion chamber, respectively. Also, the output of the combustion chamber contains a gas collector, consisting of an external 10 and an internal 11 cascade of individual output pipes, limited by external 12 and internal 13 buildings, and these pipes are combined at the output into one common channel 14. On the back side of the corner flame stabilizer there are additional external 15 and the internal 16 elements of the transverse profiling of the surfaces necessary for the elimination of local vortex structures in the field of connection of return flows to the main stream. There is also corrugation in the form of additional profiling around the circumference of the outer 17 and 18 inner outlet edges of the manifold blades to increase the throughput of the gas flow collector. Additionally, this illustration shows the longitudinal axis 19 of a gas turbine engine. The flow structure in the combustion chamber is schematically represented as arrows, with the visible part of the flow indicated by solid arrows, and the hidden part by dashed arrows.

In FIG. 2 shows a longitudinal section through the combustion chamber, such that the secant plane corresponds to plane IV-IV in FIG. 6-7 and passes through the channels of the collector 6.7, and not through the outlet pipes 10.11 as in FIG. 1.

In FIG. Figure 3 shows an isometric view of the front of the combustion chamber, on which the external 2 and internal 3 walls of the combustion chamber, the corner flame stabilizer 4, as well as the external casings of the gas collecting pipe 12 are visible.

In FIG. Figure 4 shows an isometric view of the rear of the combustion chamber, on which the external 2 and internal 3 walls of the combustion chamber, the external 10 and internal 11 stages of the outlet pipes of the gas collector, the external 12 and internal 13 of the housing of the outlet pipes of the gas collector, and also the output annular channel 14 are visible.

In FIG. 5 shows a side view of the combustion chamber, on which the outer wall 2 of the combustion chamber and the outer casings of the outlet pipes of the gas collector 13 are visible. Also shown in this illustration are planes I-I and II-II crossing the combustion chamber transversely.

In FIG. 6 is a partial isometric cross-sectional view of a combustion chamber formed in plane I-I of FIG. 5. In this illustration, the external 2 and internal 3 walls of the combustion chamber, the corner flame stabilizer 4 and the elements of its profiling 5, 15, 16 are visible. Also shown in this illustration are planes III-III and IV-IV, crossing the combustion chamber along.

In FIG. 7 is a partial isometric view of a cross section of a combustion chamber formed in plane II-II of FIG. 5. In this illustration, the external 2 and internal 3 walls of the combustion chamber are visible, the external 6 and internal 7 channels of the gas flow collector, the external 8 and internal 9 blades of the gas flow collector, the external 10 and internal 11 stages of the outlet pipes of the gas collector, external 12 and internal 13 the body of the outlet pipes of the gas collector, and corrugated in the form of additional profiling around the circumference of the outer 17 and inner 18 outlet edges of the manifold blades. Also shown in this illustration are planes III-III and IV-IV, crossing the combustion chamber along.

The operation of the device, explaining the proposed method of organizing the working process in the combustion chamber, is as follows: the air from the compressor (not shown) is carried out in an expanding channel formed by the external 2 and internal 3 walls of the combustion chamber 1, where, as a result of flow around the corner flame stabilizer 4 , the air flow is directed to the gas flow manifold. The collector is necessary for the intensification of return flows in the zones of reverse currents, as well as to prevent the mutual influence of gas flows in the zones of reverse currents and at the exit of the combustion chamber. The collector also provides control over the coefficient of excess air in the combustion chamber.

Then the flow is divided into two parts by the external 8 and internal 9 collector blades, one part enters the collector outer 6 and internal 7 channels, where it is twisted in the transverse direction relative to the longitudinal axis of the engine 19 and organizes the return flow (B in Fig. 1-2), directed to the side of the profiling elements of the corner flame stabilizer 5. Then the flow swirls in the opposite direction and enters the combustion zone of the fuel-air mixture (G in Fig. 1-2), where, taking into account the effects of additional external 15 and internal 16 lementov profiling is attached to the second portion of air, the separated blades reservoir and trapped in the area of reverse currents. After that, hot gas flows enter the external 10 and internal 11 stages of the outlet pipes of the gas collector, limited by the external 12 and internal 13 buildings, where, ultimately, they are combined in a continuous annular channel 14 and enter the turbine (not shown) of the gas turbine engine.

This gas collector, unlike standard ones, allows you to combine individual gas flows at the outlet of the combustion chamber in a single annular channel, taking into account the presence of a gas flow collector.

The location of the manifold blades determines how much air will flow into the reverse current zone B, and how much air will pass this zone. That is, the collector allows you to set the coefficient of excess air in the combustion chamber over a wide range by moving the external 8 and internal 9 blades of the manifold relative to the external 2 and internal 3 walls of the combustion chamber, respectively.

In turn, the presence of elements of the internal profiling of the corner flame stabilizer 5, by geometrical influence on the flow structure, prevents the occurrence of stagnation and, as a result, zones of local re-enrichment of the fuel-air mixture, which will positively affect the completeness of combustion of the fuel and the amount of harmful substances.

Additional elements of the corner flame stabilizer 15, 16 and the manifold blades 17, 18 increase the ejection of the return flow to the incoming stream and prevent the occurrence of local vortex structures, which will favorably affect the intensity of mixing the fuel with the oxidizing agent.

The results of hydrodynamic experiments carried out with a physical model of the combustion chamber of the proposed design confirm the efficiency and prospects of the above method of organizing the working process in the annular combustion chamber of a gas turbine engine.

Literature

1. Inozemtsev, A.A. Fundamentals of designing aircraft engines and power plants. - M.: Mechanical Engineering, 2008 .-- 366 p.

2. RF patent No. 2651692, IPC F23R 3/26, published on April 23, 2018.

Claims (5)

1. An annular combustion chamber of a gas turbine engine, containing the outer and inner walls, which are made in the form of curved surfaces, an angular flame stabilizer, characterized in that on its back side there is profiling in the form of two curved surfaces, the output of the combustion chamber contains a gas flow collector, which is a channel formed by the blades of the manifold and the walls of the combustion chamber, a gas collector consisting of a cascade of individual outlet pipes bounded by bodies, a cat The others as they approach the turbine are combined into a single annular channel. 2. The annular combustion chamber of a gas turbine engine according to claim 1, characterized in that the corner flame stabilizer on the back side has a cascade of additional profiled elements located around the circumference to eliminate local vortex structures. 3. The annular combustion chamber of a gas turbine engine according to claim 1, characterized in that the manifold blades have corrugation in the form of additional profiling around the circumference of the outlet edges to increase the throughput of the gas flow manifold. 4. A gas turbine engine containing an annular combustion chamber according to any one of the preceding paragraphs. 1-3. 5. A method of organizing a working process in an annular combustion chamber, which consists in twisting the gas flow coming from the compressor behind the front-end device, thereby forming reverse current zones for organizing the combustion process of the fuel-air mixture, characterized in that they swirl the gas flow behind the front device with a single solid front along the entire annular channel in the transverse direction relative to the longitudinal axis of the engine; in the output region of the combustion chamber, the gas flow is divided into two parts One part is forcibly directed to the reverse current zone, the second part is fed directly to the combustion zone, stagnation zones are prevented from occurring on the back of the front device by geometrical influence on the flow structure, gas exit from the combustion chamber is organized in the form of many separate jets, which are then combined into a single stream.

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