Combustion chamber of an aeroengine and an aeroengine

Abstract

The present invention discloses a combustion chamber of an aeroengine and an aeroengine. The combustion chamber of the aeroengine includes: a combustion chamber casing, including an outer combustion chamber casing and an inner combustion chamber casing; a flame tube disposed between the outer combustion chamber casing and the inner combustion chamber casing, including a tube wall, and the tube wall forms an inner cavity for fuel combustion; wherein, a plurality of cooling holes communicating the inner cavity and the outside of the flame tube are provided on the tube wall, and the cooling holes are used for introducing cooling gas into the inner cavity to cool the tube wall during fuel combustion, at least one partition is provided between the tube wall and the combustion chamber casing, and on the axis of the flame tube, the projection of the partition covers the projection of at least one of the cooling holes. The aeroengine applying the combustion chamber of the aeroengine helps to weaken the oscillating combustion during the combustion of the combustion chamber and makes the combustion more stable.

Description

Technical Field

[0001] The present invention relates to the field of aviation technology, and particularly relates to a combustion chamber of an aeroengine and an aeroengine. Background Art

[0002] As a power device, aeroengines are widely used in the fields of aviation power, electricity, industry, ships, and land-based national defense. Among the technical difficulties in this power device field, oscillatory combustion is a key technical challenge. During the operation of an aeroengine, the combustion chamber of the aeroengine sometimes exhibits unstable combustion phenomena with large-amplitude pressure pulsations. For example, in some aeroengines that adopt lean premixed prevaporized combustion technology to reduce the emissions of pollutants mainly composed of nitrogen oxides (NO X ), during combustion, the flame in the combustion chamber often operates under working conditions close to extinction, and is easily affected by external disturbances. When thermoacoustic coupling occurs, it is likely to trigger oscillatory combustion, which will, in the lightest case, affect the stable combustion working range of the power device, and in the severest case, cause hardware ablation of the hot-end components of the power device combustion chamber or damage to other component hardware, and seriously threaten the operation safety of the entire power device. Due to the unstable combustion phenomenon caused by oscillatory combustion, it is easy to cause engine vibration, increased pollutant emissions, nozzle ablation, etc. Therefore, oscillatory combustion is a problem that must be avoided as much as possible in the combustion chamber of an aeroengine.

[0003] To solve the problem of oscillatory combustion in aeroengines, there are currently two main implementation schemes, namely active control and passive control. Active control is to use an active actuator to monitor the signal of the pulsating pressure or other pneumatic parameters in the combustion device in real time, and according to the oscillation frequency of the monitored pulsating pressure, etc., actively apply a corresponding anti-phase excitation to the inlet gas path or fuel supply through the external excitation of the control system to reduce the pulsating pressure in the combustion device and suppress or eliminate thermoacoustic coupling. However, this method requires adding additional control systems and actuators, and also requires a more complex control system, thus resulting in an increase in additional technical difficulties and costs. Passive control is to deeply understand its generation law and physical mechanism, identify the conditions or generation mechanism for the occurrence of oscillatory combustion through experiments, and according to the actual situation of oscillatory combustion, add certain fixed devices in the combustion device or improve the design of the combustion device structure, so as to suppress or eliminate the oscillatory combustion phenomenon.

[0004] The present application proposes an aeroengine, a combustion chamber of a gas turbine, and a method for suppressing oscillatory combustion, which can suppress the oscillatory combustion phenomenon of the combustion device while ensuring the service life of the combustion chamber and the performance of the power device. Summary of the Invention

[0005] The object of the present invention is to provide a combustion chamber of an aeroengine and an aeroengine. The aeroengine applying the combustion chamber of the present invention helps to weaken oscillatory combustion during the combustion in the combustion chamber, making the combustion more stable.

[0006] The present invention discloses a combustion chamber of an aeroengine in a first aspect, comprising:

[0007] A combustion chamber casing, including an outer combustion chamber casing and an inner combustion chamber casing;

[0008] A flame tube, disposed between the outer combustion chamber casing and the inner combustion chamber casing, including a tube wall, and the tube wall forms an inner cavity for fuel combustion;

[0009] Wherein, a plurality of cooling holes communicating the inner cavity and the outside of the flame tube are provided on the tube wall, and the cooling holes are used to introduce cooling gas into the inner cavity to cool the tube wall during fuel combustion. At least one partition is provided between the tube wall and the combustion chamber casing, and on the axis of the flame tube, the projection of the partition covers the projection of at least one of the cooling holes.

[0010] In some embodiments, the partition extends along the extending direction of the axis of the flame tube.

[0011] In some embodiments, along the radial direction of the flame tube, both ends of the partition are respectively connected to the combustion chamber casing and the tube wall.

[0012] In some embodiments, on the axis of the flame tube, the projection of the partition covers the projection of all the cooling holes.

[0013] In some embodiments, the combustion chamber of the aeroengine is an annular combustion chamber. The tube wall includes an outer tube wall coaxial with the outer combustion chamber casing and the inner combustion chamber casing and closer to the outer combustion chamber casing relative to the inner combustion chamber casing, and an inner tube wall coaxial with the outer tube wall and closer to the inner combustion chamber casing relative to the outer combustion chamber casing. A plurality of cooling holes are provided on both the outer tube wall and the inner tube wall, and at least one of the partitions is provided between the outer tube wall and the outer combustion chamber casing and between the inner tube wall and the inner combustion chamber casing.

[0014] In some embodiments, a plurality of the partitions are provided between the outer tube wall and the outer combustion chamber casing, and the plurality of partitions are evenly arranged along the circumferential direction of the flame tube; and / or, a plurality of the partitions are provided between the inner tube wall and the inner combustion chamber casing, and the plurality of partitions are evenly arranged along the circumferential direction of the flame tube.

[0015] In some embodiments, it further comprises:

[0016] A pre - diffuser is arranged upstream of the flame tube along the intake direction and is used to reduce the pressure and velocity of the gas flowing into the flame tube.

[0017] A plurality of combustion components are arranged circumferentially around the inlet of the flame tube. The combustion component includes a pre - combustion stage fuel nozzle orifice in the middle, a main combustion stage fuel nozzle orifice surrounding the pre - combustion stage fuel nozzle orifice, a first swirler and a second swirler arranged radially inside the main combustion stage fuel nozzle orifice and surrounding the pre - combustion stage fuel nozzle orifice, a third swirler located between the main combustion stage fuel nozzle orifice and the second swirler, and a fourth swirler surrounding the main combustion stage fuel nozzle orifice.

[0018] In some embodiments, the first swirler and the second swirler are radial swirlers, the third swirler is an axial swirler, and the fourth swirler is a radial swirler.

[0019] The second aspect of the present invention discloses an aero - engine, including the combustion chamber of any one of the above - mentioned aero - engines.

[0020] Based on the combustion chamber of the aero - engine provided by the present invention, by arranging a partition between the barrel wall of the flame tube and the combustion chamber casing, and the partition covers at least part of the cooling holes along the axial direction of the flame tube. When thermo - acoustic coupling oscillation occurs during the combustion of the combustion chamber, the partition can change the waveform of the sound wave between the barrel wall and the combustion chamber casing near the covered cooling holes, so that the waveform of the sound wave in the flame tube and the waveform near the cooling holes change from the same direction to the opposite direction, increasing the pulsating pressure difference of the sound waves inside and outside the flame tube, thereby increasing the dissipation of the sound energy of the flame tube through the cooling holes, further weakening the oscillatory combustion, and improving the combustion stability of the combustion chamber.

[0021] Through the following detailed description of the exemplary embodiments of the present invention with reference to the accompanying drawings, other features and advantages of the present invention will become clear. Description of the Drawings

[0022] The accompanying drawings described herein are used to provide a further understanding of the present invention, form a part of this application, and the schematic embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation to the present invention. In the drawings:

[0023] Figure 1 It is a schematic structural diagram of the combustion chamber of the aero - engine according to the embodiment of the present invention;

[0024] Figure 2 It is Figure 1 A partially enlarged structural diagram of the combustion component of the combustion chamber of the aero - engine shown;

[0025] Figure 3Schematic diagram of a partial sectional view of a combustion chamber of an aeroengine according to another embodiment of the present invention;

[0026] Figure 4 is Figure 3 Schematic diagram of the AA-direction sectional view of the combustion chamber of the aeroengine shown in the figure. Specific embodiments

[0027] The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. The following description of at least one exemplary embodiment is actually illustrative only and in no way limits the present invention and its application or use. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the scope of protection of the present invention.

[0028] Unless otherwise specifically stated, the relative arrangements of components and steps, numerical expressions and values set forth in these embodiments do not limit the scope of the present invention. At the same time, it should be understood that for the sake of description, the dimensions of the various parts shown in the drawings are not drawn in actual proportional relationships. Technologies, methods and devices known to those of ordinary skill in the relevant art may not be discussed in detail, but where appropriate, the technologies, methods and devices should be regarded as part of the authorization specification. In all the examples shown and discussed here, any specific value should be construed as merely exemplary and not as a limitation. Therefore, other examples of the exemplary embodiments may have different values. It should be noted that like reference numerals and letters denote like items in the following drawings, and thus, once an item is defined in one drawing, it does not need to be further discussed in subsequent drawings.

[0029] For the sake of description, spatial relative terms such as "above...", "over...", "on the upper surface of...", "above" and the like may be used herein to describe the spatial positional relationship between one device or feature and other devices or features as shown in the figures. It should be understood that the spatial relative terms are intended to encompass different orientations in use or operation in addition to the orientation depicted in the figures. For example, if the device in the figure is inverted, the device described as "above other devices or structures" or "over other devices or structures" will then be positioned "below other devices or structures" or "under other devices or structures". Thus, the exemplary term "above..." can include both the orientation of "above..." and "below...". The device may also be positioned in other different ways (rotated 90 degrees or in other orientations), and corresponding interpretations should be made for the spatial relative descriptions used herein.

[0030] As shown Figures 1 to 4 in the figure, the combustion chamber of the aero-engine in this embodiment includes a combustion chamber casing and a flame tube.

[0031] The combustion chamber casing includes an outer combustion chamber casing 11 and an inner combustion chamber casing 12. With respect to the axis of the aero-engine, the outer combustion chamber casing 11 is the part of the combustion chamber casing located on the radially outer side, and the inner combustion chamber casing 12 is the part of the combustion chamber casing located on the radially inner side.

[0032] The flame tube is arranged between the outer combustion chamber casing 11 and the inner combustion chamber casing 12. The flame tube includes a tube wall, and the tube wall forms an inner cavity 23 for fuel combustion. The combustion chamber can be Figures 1 to 4 as shown in the figure an annular combustion chamber. At this time, with respect to the axis of the aero-engine, the tube wall includes an outer tube wall 21 located on the radially outer side and an inner tube wall 22 located on the radially inner side. The outer combustion chamber casing 11, the inner combustion chamber casing 12, the outer tube wall 21 and the inner tube wall 22 are all coaxial annular structures, and the tube wall is a double-annular structure including an annular inner tube wall 22 and an annular outer tube wall 21. In some embodiments not shown in the figure, the combustion chamber can also be a single-tube combustion chamber. At this time, the tube wall of the flame tube is a structure integrally in the shape of a cylinder. In some embodiments, the combustion chamber can also be a can-annular combustion chamber.

[0033] Wherein, a plurality of cooling holes 212 communicating the inner cavity 23 and the outside of the flame tube are provided on the tube wall. The cooling holes 212 are through-hole structures. The cooling holes 212 are used to introduce cooling gas into the inner cavity 23 to cool the tube wall during fuel combustion, and perform film cooling on the tube wall. At least one partition 6 is provided between the tube wall and the combustion chamber casing. On the axis of the flame tube, the projection of the partition 6 covers the projection of at least one cooling hole 212.

[0034] A cavity is formed between the tube wall and the combustion chamber casing. A first cavity 121 is formed between the tube wall and the outer combustion chamber casing 11, and a second cavity 122 is formed between the tube wall and the inner combustion chamber casing 12. In the embodiment where the combustion chamber is an annular combustion chamber, both the first cavity 121 and the second cavity 122 are integral annular cavities. In a single-tube combustion chamber, both the first cavity 121 and the second cavity 122 are semi-annular cavities.

[0035] A partition is provided in the first cavity 121 or the second cavity 122. On the axis of the flame tube, the projection of the partition 6 covers the projection of at least one cooling hole 212, that is, along the axial direction of the flame tube, the axial position of the partition covers the axial position of at least one cooling hole 212. That is, if a three-dimensional coordinate is established with the axis of the flame tube as the x-axis, the coordinate range of at least one cooling hole on the x-axis is located within the coordinate range of the partition on the x-axis. When the combustion chamber is an annular combustion chamber, the axis of the flame tube is collinear with the axis of the aero-engine. When the combustion chamber is a single-tube combustion chamber or a can-annular combustion chamber, the axis of the flame tube is not collinear with the axis of the aero-engine.

[0036] For the combustion chamber of the aero-engine in this embodiment, by arranging a partition plate 6 between the cylinder wall of the flame tube and the combustion chamber casing, and the partition plate 6 covers at least part of the cooling holes 212 along the axial direction of the flame tube. When thermoacoustic coupling oscillation occurs during the combustion of the combustion chamber, the sound waves generated in the inner cavity 23 will be transmitted to the chamber between the cylinder wall and the combustion chamber casing through the upstream end of the flame tube and the cooling holes 212. Many of the sound waves generated in the inner cavity during oscillating combustion are circumferential waves. After arranging the partition plate 6, the partition plate 6 can change the waveform of the sound waves in the chamber between the cylinder wall and the combustion chamber casing near the covered cooling holes 212, so that the waveform of the sound waves in the inner cavity of the flame tube and the waveform of the circumferential waves near the cooling holes 212 outside the flame tube change from the same direction to the opposite direction, increasing the pulsating pressure difference of the sound waves on both sides of the cooling holes 212 covered by the partition plate in the axial position of the flame tube. Therefore, when the sound waves pass through the cooling holes 212 covered by the partition plate in the axial position, the dissipation of sound energy can be further improved, the oscillating combustion can be further weakened, and the combustion stability of the combustion chamber can be improved.

[0037] In some embodiments, such as Figure 3 and Figure 4 shown, the partition plate 6 extends along the extending direction of the axis of the flame tube. That is, the partition plate 6 is a relatively standard plate structure, the longitudinal section of which is parallel to the axis of the flame tube, and the partition plate 6 can be obtained by translating this longitudinal section. The partition plate 6 obtained by this setting is simple to manufacture and has a good effect on weakening oscillating combustion.

[0038] In some embodiments, along the radial direction of the flame tube, both ends of the partition plate 6 are respectively connected to the combustion chamber casing and the cylinder wall. That is, in addition to extending along the axial direction of the flame tube, the partition plate 6 extends to the cylinder wall and the combustion chamber casing respectively in the radial direction of the flame tube. The partition plate set in this way has a large range in the radial direction of the flame tube and has a good effect on changing the waveform of the sound waves and a good effect on weakening oscillating combustion.

[0039] In some embodiments, such as Figure 3 and Figure 4 shown, on the axis of the flame tube, the projection of the partition plate 6 covers the projection of all the cooling holes 212. That is, each partition plate can cover all the cooling holes 212, so that the sound energy dissipation effect of all the cooling holes 212 can be improved, and the effect of further weakening oscillating combustion can be further improved.

[0040] In some embodiments, such as Figure 3 and Figure 4As shown in the figure, the combustion chamber of the aero-engine is an annular combustion chamber. The cylinder wall includes an outer cylinder wall 21 that is coaxial with the outer casing 11 of the combustion chamber and closer to the outer casing 11 of the combustion chamber than to the inner casing 12 of the combustion chamber, and an inner cylinder wall 22 that is coaxial with the outer cylinder wall 21 and closer to the inner casing 12 of the combustion chamber than to the outer casing 11 of the combustion chamber. A plurality of cooling holes 212 are provided on both the outer cylinder wall 21 and the inner cylinder wall 22. At least one partition 6 is provided between the outer cylinder wall 21 and the outer casing 11 of the combustion chamber and between the inner cylinder wall 22 and the inner casing 12 of the combustion chamber. In the annular combustion chamber, through experiments, it is found that setting partitions can significantly weaken the effect of oscillatory combustion.

[0041] In some embodiments, as Figure 3 and Figure 4 shown, a plurality of partitions 6 are provided between the outer cylinder wall 21 and the outer casing 11 of the combustion chamber, and the plurality of partitions 6 are evenly arranged along the circumferential direction of the flame tube; and / or, a plurality of partitions 6 are provided between the inner cylinder wall 22 and the outer casing 11 of the combustion chamber, and the plurality of partitions 6 are evenly arranged along the circumferential direction of the flame tube. Setting a plurality of partitions can further improve the sound energy dissipation effect of the cooling holes 212. In the embodiment shown in Figure 4 the figure, three partitions 6 are provided in the first chamber 121 between the outer cylinder wall 21 and the outer casing 11 of the combustion chamber, and the three partitions 6 are evenly arranged along the circumferential direction. Three partitions 6 are provided in the second chamber 1212 between the inner cylinder wall 22 and the inner casing 12 of the combustion chamber, and the three partitions 6 are evenly arranged along the circumferential direction. This setting can make the effect of weakening oscillatory combustion more uniform and effective.

[0042] In some embodiments, as Figures 1 to 4 shown, the combustion chamber of the aero-engine further includes a pre-stage diffuser 3 and a plurality of combustion components.

[0043] The pre-stage diffuser 3 is arranged upstream of the flame tube along the intake direction. The pre-stage diffuser 3 is used to expand and decelerate the gas flowing into the flame tube. When the aero-engine is operating, the compressed air coming from the high-pressure compressor flows towards the flame tube after the expansion and deceleration effect of the pre-stage diffuser 3.

[0044] A plurality of combustion components are arranged along the circumferential direction of the flame tube at the entrance of the flame tube. The combustion component includes a pre-combustion stage fuel nozzle orifice 50 located in the middle, a main combustion stage fuel nozzle orifice 55 surrounding the pre-combustion stage fuel nozzle orifice, a first swirler 51 and a second swirler 52 located radially inside the main combustion stage fuel nozzle orifice 55 and surrounding the pre-combustion stage fuel nozzle orifice 50, a third swirler 53 located between the main combustion stage fuel nozzle orifice 55 and the second swirler 52, and a fourth swirler 54 surrounding the main combustion stage fuel nozzle orifice 55. In this embodiment, as Figure 2The air flow path shown by the dashed line in the figure. When the aero-engine is operating, one stream of compressed air passes through the first swirler 51, and another stream of compressed air passes through the second swirler 52 and then mixes with the fuel passing through the nozzle 50 of the pre-combustion stage fuel nozzle. One stream of compressed air passes through the third swirler 53, and another stream of compressed air passes through the fourth swirler 54 and then mixes with the fuel passing through the nozzle 50 of the main combustion stage fuel nozzle. Finally, all the fuel-air mixture gases converge and mix in the inner cavity of the combustion chamber. The combustion assembly of this embodiment can achieve lean-premixed pre-vaporization combustion, which can effectively reduce nitrogen oxides, but is more likely to generate oscillating combustion. With the arrangement of the partition plate, this embodiment can make the combustion more stable while effectively reducing nitrogen oxides.

[0045] In some embodiments, the first swirler 51 and the second swirler 52 are radial swirlers, the third swirler 53 is an axial swirler, and the fourth swirler 54 is a radial swirler. This arrangement can improve the combustion effect of lean-premixed pre-vaporization combustion.

[0046] In some embodiments, an aero-engine is also disclosed, which includes the combustion chamber of any of the above-mentioned aero-engines.

[0047] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and not to limit them; although the present invention has been described in detail with reference to the preferred embodiments, those of ordinary skill in the art should understand that: it is still possible to modify the specific implementation manners of the present invention or perform equivalent replacements for some technical features; without departing from the spirit of the technical solutions of the present invention, they should all be covered within the scope of the technical solutions claimed by the present invention.

Claims

1. A combustion chamber of an aeroengine, characterized in that, Comprising: A combustion chamber casing, including an outer combustion chamber casing (11) and an inner combustion chamber casing (12); A flame tube, disposed between the outer combustion chamber casing (11) and the inner combustion chamber casing (12), including a tube wall, and the tube wall forms an inner cavity (23) for fuel combustion; Wherein, a plurality of cooling holes (212) communicating the inner cavity (23) and the outside of the flame tube are provided on the tube wall, and the cooling holes (212) are used to introduce cooling gas into the inner cavity (23) to cool the tube wall during fuel combustion. At least one partition (6) is provided between the tube wall and the combustion chamber casing. On the axis of the flame tube, the projection of the partition (6) covers the projection of at least one of the cooling holes (212). The partition (6) extends along the extending direction of the axis of the flame tube. Radially along the flame tube, both ends of the partition (6) are respectively connected to the combustion chamber casing and the tube wall.

2. The combustion chamber of an aeroengine according to claim 1, characterized in that On the axis of the flame tube, the projection of the partition (6) covers the projections of all the cooling holes (212).

3. The combustion chamber of an aero-engine according to claim 1, characterized in that, The combustion chamber of the aeroengine is an annular combustion chamber. The tube wall includes an outer tube wall (21) coaxial with the outer combustion chamber casing (11) and the inner combustion chamber casing (12) and closer to the outer combustion chamber casing (11) relative to the inner combustion chamber casing (12), and an inner tube wall (22) coaxial with the outer tube wall (21) and closer to the inner combustion chamber casing (12) relative to the outer combustion chamber casing (11). A plurality of cooling holes (212) are provided on both the outer tube wall (21) and the inner tube wall (22). At least one of the partitions (6) is provided between the outer tube wall (21) and the outer combustion chamber casing (11) and between the inner tube wall (22) and the inner combustion chamber casing (12).

4. The combustion chamber of an aeroengine according to claim 3, characterized in that, A plurality of the partitions (6) are provided between the outer tube wall (21) and the outer combustion chamber casing (11), and the plurality of partitions (6) are evenly arranged along the circumferential direction of the flame tube; and / or, a plurality of the partitions (6) are provided between the inner tube wall (22) and the inner combustion chamber casing (12), and the plurality of partitions (6) are evenly arranged along the circumferential direction of the flame tube.

5. The combustion chamber of an aeroengine according to claim 3, characterized in that, Further comprising: A pre-stage diffuser, disposed upstream of the flame tube along the intake direction, for diffusing and reducing the speed of the gas flowing into the flame tube; A plurality of combustion assemblies, disposed at the inlet of the flame tube along the circumferential direction of the flame tube. The combustion assembly includes a pre-combustion stage fuel nozzle orifice (50) located in the middle, a main combustion stage fuel nozzle orifice (55) surrounding the pre-combustion stage fuel nozzle orifice, a first swirler (51) and a second swirler (52) located radially inside the main combustion stage fuel nozzle orifice (55) and surrounding the pre-combustion stage fuel nozzle orifice (50), a third swirler (53) located between the main combustion stage fuel nozzle orifice (55) and the second swirler (52), and a fourth swirler (54) surrounding the main combustion stage fuel nozzle orifice (55).

6. The combustion chamber of an aeroengine as claimed in claim 5, characterized in that, The first cyclone (51) and the second cyclone (52) are radial cyclones, the third cyclone (53) is an axial cyclone, and the fourth cyclone (54) is a radial cyclone.

7. An aeroengine, characterized in that, A combustor of an aeroengine according to any one of claims 1 to 6.

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