CN105953266B - A kind of oblique flow chamber structure - Google Patents

Abstract

A structure of a diagonal flow combustion chamber. The invention relates to a combustion chamber structure, which can realize the non-axial air intake of the combustion chamber. The combustion chamber of the present invention adopts an oblique-flow flame cylinder structure, that is, the head of each flame cylinder deflects at a certain angle relative to the axis of the combustion chamber. The heads of different flame tubes are connected by the head side plate, the side plate of the flame tube head is provided with cooling holes, and the side plate of the flame tube head is perpendicular to the head panel. There are secondary air holes on the inner and outer walls of the flame tube. The secondary air holes are divided into several groups according to the number of flame tube heads. The arrangement direction of each group of secondary air holes is parallel to the outer edge of the corresponding flame tube head panel. The oblique flow combustor of the present invention can cancel the outlet guide vane of the upstream axial flow compressor, shorten the length of the whole machine of the aero-engine, reduce the quality of the whole machine, and improve the thrust-to-weight ratio of the whole machine; in addition, the oblique flow combustor of the present invention can also reduce The deflection angle of the turbine inlet guide vane downstream of the small combustion chamber reduces the difficulty of designing the turbine inlet guide vane and improves turbine efficiency.

Description

A Diagonal Flow Combustion Chamber Structure

technical field

The invention relates to the design field of an aero-engine combustion chamber, and relates to a combustion chamber structure capable of realizing non-axial air intake in an aero-engine combustion chamber, specifically a combustion chamber structure with an oblique-flow flame cylinder.

Background technique

In an aero-engine, the air is first compressed by the compression system, and the compressed high-pressure airflow enters the combustion chamber. The fuel injection system injects fuel into the high-pressure airflow, and the fuel is fully and effectively burned in the combustion chamber. After combustion, high-temperature, high-pressure gas is formed. The gas drives a turbine to provide the work required by the compression system. For the whole gas turbine, shorter compressor, combustor and turbine length means higher thrust-to-weight ratio. The main way to shorten the axial length of the compressor and turbine under the premise of ensuring the same performance of the compressor is to increase the stage load, but how much the stage load will make the design of the blade shape more difficult and the strength of the blade will be reduced. For the combustion chamber, the length of the combustion chamber needs to be sufficient to ensure that the fuel has a sufficient residence time in the combustion chamber, thereby maintaining a high combustion efficiency. At present, the aspect ratio of the combustion chamber has been developed from 6 in the early stage to about 2. If the length of the combustion chamber is further shortened under the premise of ensuring a certain height of the head of the flame tube, it is necessary to introduce more advanced combustion technology, otherwise the performance of the combustion chamber will be seriously weakened. Therefore, engineers and technicians need to innovate in the structural design of aero-engines to further increase the thrust-to-weight ratio of aero-engines.

As we all know, the air intake direction of the conventional combustion chamber is axial, affected by the rotation of the last stage of the compressor, the outlet airflow of the last stage of the rotor blade is at a certain angle to the axial direction, in order to adjust the direction of the outlet airflow of the rotor blade to the axial direction, In order to meet the intake requirements of the combustion chamber, it is necessary to add a first-stage guide vane downstream of the last-stage rotor blade of the compressor to adjust the airflow direction. Therefore, if the outlet guide vane of the compressor can be canceled by some technical means, and the seamless connection between the last stage rotor blade of the compressor and the combustion chamber can be realized, the overall mass of the engine can be effectively reduced, the axial length of the entire engine can be shortened, and the thrust-to-weight ratio can be increased. .

Contents of the invention

The invention relates to the design field of an aero-engine combustion chamber, and relates to a combustion chamber structure capable of realizing non-axial air intake of the aero-engine combustion chamber. The oblique flow combustor of the present invention can cancel the outlet guide vane of the upstream axial flow compressor, effectively shorten the overall length of the aero-engine, and reduce the quality of the complete machine; in addition, the oblique flow combustor of the present invention can also reduce The turning angle of the turbine inlet guide vane reduces the design difficulty of the turbine inlet guide vane and improves the turbine efficiency.

In order to achieve the above-mentioned technical purpose, the oblique flow combustion chamber structure of the present invention is realized through the following technical solutions:

A combustion chamber structure, used to realize the non-axial air intake of the combustion chamber of the aero-engine; the aero-engine includes a compressor, a combustion chamber and a turbine; the compressor is an axial flow type; the combustion chamber is a co-flow type, It includes a diffuser, a flame tube, and inner and outer casings, located downstream of the axial flow compressor; the flame tube is a diagonal flow structure, and it is characterized in that: the diagonal flow flame tube is annular and includes a flame tube head And the inner and outer wall surfaces of the flame tube; the head of the flame tube is multiple, uniformly distributed along the circumference of the combustion chamber, including a swirler, a head panel, an oil baffle plate and a head side plate, and each flame tube head is relatively combusted The axis of the chamber is deflected at a certain angle; the head panel has cooling holes; the swirler is installed on the head panel, and the central axis of the swirler is perpendicular to the head panel; the oil deflector is located on the The high temperature side of the head panel, and a certain distance from the head panel; the head side plate is used to connect the flame tube heads; the head side plate is provided with cooling holes; the flame tube The inner and outer walls are connected by the flame tube head panel and the head side plate, and the front edges of the flame tube inner and outer walls are connected with the inner and outer edges of the flame tube head panel and the inner and outer sides of the flame tube head side plate. The edges are flush; there are secondary air holes on the inner and outer walls of the flame tube; the secondary air holes are divided into several groups according to the number of flame tube heads, and the arrangement direction of each group of secondary air holes corresponds to that of the flame tube The outer edges of the head panels are parallel.

Preferably, the angle α between the head of the flame tube and the axis of the combustion chamber is:

Among them, β is the geometric angle of the outlet of the rotor blade of the last stage of the compressor, w is the relative velocity of the outlet airflow, and c is the absolute velocity of the outlet airflow.

Compared with the prior art, the oblique flow combustion chamber structure of the present invention has significant technical effects: it can realize non-axial air intake in the combustion chamber of the aero-engine, effectively shorten the length of the aero-engine, reduce the mass of the aero-engine, and improve the thrust-to-weight ratio of the aero-engine .

Description of drawings

Fig. 1 is a combustion chamber (3/4 sectional view) with a diagonal flow flame tube structure, and the flame tube in the figure is a diagonal flow flame tube.

Fig. 2 is a schematic diagram showing the circumference development of the oblique-flow flame cylinder and the last-stage rotor blade of a certain type of axial-flow compressor according to the present invention.

Fig. 3 is the oblique-flow flame tube of the present invention, (A) oblique front view at 45°, (B) front view, (C) oblique rear view at 45°.

The symbols in the figure are explained as follows:

1. Cyclone; 2. Head panel; 3. Head side panel; 4. Oil retaining plate; 5. Inner wall of flame tube; 6. Outer wall of flame tube; 7. Secondary air hole; 8. Diffusion 9. Flame tube; 10. Inner casing; 11. Outer casing; 12. Last stage moving blade of axial flow compressor.

Detailed ways

In order to make the object, technical solution and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and examples.

Fig. 1 is the annular combustor (3/4 sectional view) that has oblique flow flame tube structure, comprises diffuser 8, flame tube 9 and inner and outer casing 10,11, and described flame tube 9 is oblique flow type flame tube , the head of the flame tube deflects at a certain angle relative to the axis of the combustion chamber. Fig. 2 is a schematic diagram showing the circumference development of the oblique-flow flame cylinder 9 and the last-stage rotor blade 12 of a certain type of axial-flow compressor of the present invention. Among them, the geometric angle β of the rotor blade outlet is 55°, the relative velocity w of the outlet airflow is 210m/s, and the absolute velocity c of the outlet airflow is 150m/s. According to the following formula, the angle α between the head of the flame tube and the axis of the combustion chamber is obtained as 36.58°.

Fig. 3 is a diagonal flow flame tube of the present invention. Wherein, the oblique-flow flame cylinder includes a two-stage radial swirler 1, a head panel 2, a head side plate 3, an oil baffle plate 4, and inner and outer walls 5, 6 of the flame cylinder; The flame tube includes 12 heads, which are evenly distributed along the circumference of the combustion chamber; the central axis of the two-stage radial swirler 1 is perpendicular to the head panel 2; the head panel 2 is perpendicular to the head side plate 3 , there is a cooling hole on it, the diameter of the cooling hole is 1.5mm; the oil deflector 4 is located on the high temperature side of the head panel 2, and is separated from the head panel 2 by 2mm; the inner and outer walls of the flame tube 5 , 6 are connected by the head panel 2 of the flame tube and the head side plate 3; The inner and outer edges of the side plate 3 are flush; the outer wall surface 6 of the flame tube is provided with secondary air holes, of which the main combustion holes have a diameter of 10mm, and 12 groups are distributed along the circumference of the combustion chamber, and each group has 4 holes; the diameter of the mixing holes is 8mm , 12 groups are distributed along the circumference of the combustion chamber, 4 in each group; the wall cooling hole diameter is 1.5 mm, and 12 groups are distributed along the circumference of the combustion chamber, and each group is arranged in 3 rows along the airflow direction, with 24 in each row; The wall 5 is provided with main combustion holes with a diameter of 11 mm, and there are 12 groups distributed along the circumference of the combustion chamber, one for each group. The arrangement direction of all the secondary air holes on the inner and outer walls 5 and 6 of the flame tube is parallel to the outer edge of the corresponding flame tube head panel.

After the air flows out from the last stage rotor blade 12 of the axial flow compressor, it directly enters the combustion chamber at an angle of 36.58° with the axis of the combustion chamber. The axes flow out at the same angle and enter the turbine to do work. Due to the cancellation of the compressor outlet guide vane, the axial length of the whole machine is shortened, the weight of the whole machine is reduced, and the thrust-to-weight ratio of the whole machine is improved; The deflection angle of the turbine inlet guide vane (even the inlet guide vane can be eliminated) is conducive to improving the efficiency of the turbine and further increasing the thrust of the whole machine.

The above descriptions are only preferred embodiments of the present invention, and are not intended to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention shall be included in the present invention. within the range.

Claims (1)

1. a kind of chamber structure, to realize that the non axial air inlet in aeroengine combustor buring room, the aero-engine include pressure Mechanism of qi, combustion chamber and turbine；The compressor is axial-flow type；The combustion chamber be downflow type, including diffuser, burner inner liner and Inside and outside casing is located at the compressor downstream；The burner inner liner is oblique flow structure, it is characterised in that：The burner inner liner is ring Shape, including flame tube head and the inside and outside wall surface of burner inner liner；The flame tube head is multiple, is circumferentially evenly distributed with along combustion chamber, including Cyclone, head panel, oil baffle disc and head side plate, each flame tube head relative combustion chamber axis deflection；The head surface Plate is provided with cooling hole；The cyclone is installed on the head panel, and the cyclone central axis upright is in head panel； The oil baffle disc is located at the head panel high temperature side, and spaced apart with the head panel；The flame tube head Between using the head side plate connect；The head side plate is provided with cooling hole；The inside and outside wall surface of burner inner liner passes through burner inner liner Head panel and the connection of head side plate, the inside and outside wall surface leading edge of burner inner liner and the inside and outside edge of the flame tube head panel and fire Side plate inside and outside edge in flame cylinder head is concordant；The inside and outside wall surface of burner inner liner is provided with auxiliary air hole；The auxiliary air hole according to Flame tube head number is divided into several groups, and the corresponding flame tube head panel outer rim of every group of auxiliary air hole orientation is flat Row；Diagonal flow type combustion chamber eliminates upstream axial flow compressor exit guide blade, reduces combustion chamber downstream turbine inlet guide vane and turns back Angle；Flame tube head is with combustion chamber axis angle α：

Wherein, β is that compressor last stage movable vane exports geometry angle, and w is exit flow relative velocity, and c is exit flow absolute velocity.