Thermal protection structure of thrust chamber of oxygen-enriched turbine engine

Abstract

The invention relates to the technical field of turbine engines, in particular to an oxygen-enriched turbine engine thrust chamber thermal protection structure which comprises a thrust chamber body, an air film cooling mechanism for sucking low-temperature air to achieve cooling is arranged in the side wall of the thrust chamber body, and an injector and a flow dividing hole channel are arranged at the air inlet end of the thrust chamber body. Compressed airflow is arranged on the side face of the thrust chamber body, a regeneration cooling mechanism is arranged on the thrust chamber body, and a flat seam type air film hole, a second air film hole and a third air film hole are sequentially formed in the inner side wall of the thrust chamber body from left to right. The gas film cooling mechanism is arranged, a flowing gas film can be formed on the inner wall of the thrust chamber body through mutual cooperation of the flat seam type gas film holes and the gas film channels, so that the thrust chamber body is cooled, compared with a liquid film cooling mode of a rocket engine in the prior art, liquid fuel does not need to be consumed as a cold source, and the cost is reduced. And the energy consumption of the oxygen-enriched turbine engine is effectively reduced.

Description

Technical Field

[0001] The invention relates to the technical field of turbine engines, and in particular to a thermal protection structure of a thrust chamber of an oxygen-rich turbine engine. Background Art

[0002] With the development of aerospace technology, the thrust chamber of hypersonic aircraft faces a harsh working environment of high temperature, high pressure and high heat flux density. In order to prevent the thrust chamber from overheating during engine operation, thermal protection measures are usually used to cool the thrust chamber.

[0003] In the prior art, most rocket engines use liquid film cooling to cool the thrust chamber, that is, consuming part of the liquid fuel to arrange several liquid films as a thermal protection measure. Obviously, this will lead to an increase in the fuel consumption of the engine, which will not only affect the overall performance of the engine, but also be detrimental to the weight reduction of the engine. Summary of the invention

[0004] In view of the deficiencies in the prior art, the present invention provides a thermal protection structure for the thrust chamber of an oxygen-rich turbine engine, which solves the technical problem in the prior art that the rocket engine needs to consume liquid fuel when cooling the engine thrust chamber, which is not conducive to engine weight reduction. The structure has the advantage of being able to combine regenerative cooling and film cooling, which can not only reduce the weight of the engine, but also improve the overall performance of the engine.

[0005] To solve the above technical problems, the present invention provides the following technical solutions: a thermal protection structure of a thrust chamber of an oxygen-rich turbine engine, comprising a thrust chamber body, a film cooling mechanism for inhaling low-temperature gas to achieve cooling is arranged inside the side wall of the thrust chamber body, an injector and a diversion channel are arranged at the air inlet end of the thrust chamber body, a compressed air flow is arranged on the side of the thrust chamber body, and a regeneration cooling mechanism is arranged on the thrust chamber body. When the engine is running, most of the external air will directly enter the interior of the thrust chamber body as a combustion oxidant after being compressed by the compressor, and a small part will enter the interior of the film cooling mechanism under the action of the diversion channel, the film cooling mechanism includes an air film channel opened inside the side wall of the thrust chamber body, and a flat-slit air film hole, a second air film hole and a third air film hole are arranged on the inner side wall of the thrust chamber body from left to right in sequence, a part of the gas diverted by the diversion channel will be ejected through the flat-slit air film hole, and the other part will enter the interior of the air film channel.

[0006] Preferably, the thrust chamber body includes a straight section and a contraction-expansion section, the flat-slit air film hole is parallel to the straight section, and after the airflow is ejected through the flat-slit air film hole, it moves along the side wall of the straight section, thereby forming an air film.

[0007] Preferably, the second and third air film holes are both in the form of rows of holes, and are both located in the contraction and expansion section, and the second air film hole is located near the position where the thrust chamber body line begins to contract. The third air film hole is to cover the area with the highest absolute value of the throat heat flux density and the highest wall temperature.

[0008] Preferably, the compressed airflow includes a combustion airflow and a cold source airflow, the cold source airflow includes a first cold airflow and a second cold airflow, the combustion airflow will enter the thrust chamber body under the action of the injector, and the cold source airflow will enter the air film channel under the action of the diversion channel.

[0009] Preferably, the regenerative cooling mechanism includes a cooling channel opened inside the side wall of the thrust chamber body, and a fuel injection port and a fuel discharge port are provided on the thrust chamber body. The fuel injection port and the fuel discharge port are respectively connected to the cooling channel. When the engine is running, liquid fuel will enter the cooling channel through the fuel injection port and then be discharged through the fuel discharge port.

[0010] Preferably, the fuel discharge port is connected to the interior of the thrust chamber body, and the liquid fuel participating in cooling will enter the interior of the thrust chamber body for combustion after being discharged through the fuel discharge port.

[0011] Preferably, a plurality of cooling channels are provided at equal intervals, two air film channels are arranged between every two cooling channels, and the outlet pressure of each air film is higher than the chamber pressure of the thrust chamber body.

[0012] By means of the above technical solution, the present invention provides a thermal protection structure for the thrust chamber of an oxygen-rich turbine engine, which has at least the following beneficial effects:

[0013] 1. The present invention adopts an air film cooling mechanism and a regenerative cooling mechanism, inhales external air as a combustion oxidant and a cold source for air film cooling, and uses liquid fuel for cooling, perfectly combining regenerative cooling and air film cooling, which can not only reduce the weight of the engine, but also improve the overall performance of the engine.

[0014] 2. The present invention sets up an air film cooling mechanism and utilizes the mutual cooperation between the flat-seam air film holes and the air film channels to form a flowing air film on the inner wall of the thrust chamber body, thereby cooling the thrust chamber body. Compared with the liquid film cooling method in the prior art, it does not require the consumption of liquid fuel as a cooling source, thereby effectively reducing the energy consumption of the oxygen-rich turbine engine.

[0015] 3. The present invention provides a regenerative cooling mechanism and utilizes the cooperation between the fuel injection port and the fuel discharge port, so that the thrust chamber body can be cooled by liquid fuel without consuming additional coolant, and the liquid fuel can be burned and utilized, which can effectively reduce resource waste;

[0016] 4. The present invention provides a regenerative cooling mechanism. When the liquid fuel flows through the cooling channel, it will exchange heat with the interior of the thrust chamber body, thereby increasing the temperature of the liquid fuel itself, ensuring more complete combustion of the liquid fuel and improving the overall performance of the engine.

[0017] 5. The present invention provides an air film cooling mechanism, wherein the second air film hole is located near the position where the thrust chamber linear shape begins to shrink, and the third air film hole is located near the position where the thrust chamber linear shape begins to expand. This enables the air film to be more firmly attached to the inner wall of the contraction and expansion section, thereby covering the area with the highest absolute value of the throat heat flux density and the highest wall temperature, and can effectively improve the cooling effect. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] The drawings described herein are used to provide a further understanding of the present application and constitute a part of the present application. The illustrative embodiments of the present application and their descriptions are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0019] Figure 1 It is a front view of the overall structure of the present invention;

[0020] Figure 2 It is a structural schematic diagram of the air film cooling mechanism in the present invention;

[0021] Figure 3 It is a schematic diagram of the structure of the flat seam type air film hole in the present invention;

[0022] Figure 4 It is a structural schematic diagram of the regenerative cooling mechanism in the present invention.

[0023] In the figure: 1. thrust chamber body; 101. straight section; 102. contraction and expansion section; 2. air film cooling mechanism; 201. flat-seam air film hole; 202. air film channel; 203. second air film hole; 204. third air film hole; 3. injector; 4. diversion channel; 5. compressed air flow; 501. combustion air flow; 502. cold source air flow; 503. first cold air flow; 504. second cold air flow; 6. regeneration cooling mechanism; 601. fuel injection port; 602. fuel discharge port; 603. cooling channel. DETAILED DESCRIPTION

[0024] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0025] Embodiment 1

[0026] In the prior art, most rocket engines use liquid film cooling to cool the thrust chamber, that is, consuming part of the liquid fuel to arrange several liquid films as a thermal protection measure. Obviously, this will increase the fuel consumption of the engine, which will not only affect the overall performance of the engine, but also be unfavorable to the weight reduction of the engine. In order to solve this technical defect in the prior art, Figure 1 , Figure 2 as well as Figure 3 As shown, this embodiment proposes a heat protection structure of the thrust chamber of an oxygen-rich turbine engine, which can form a flowing air film on the inner wall of the thrust chamber body 1, so as to cool the thrust chamber body 1. An air film cooling mechanism 2 for inhaling low-temperature gas to achieve cooling is provided inside the side wall of the thrust chamber body 1. An injector 3 and a diverter channel 4 are provided at the air inlet end of the thrust chamber body 1. A compressed air flow 5 is provided on the side of the thrust chamber body 1. The compressed air flow 5 includes a combustion air flow 501 and a cold source air flow 502. The cold source air flow 502 includes a first cold air flow 503 and a second cold air flow 504. The combustion air flow 501 will enter the thrust chamber body 1 under the action of the injector 3, and the cold source air flow 502 will enter the air film channel 202 under the action of the diverter channel 4. A regenerative cooling mechanism 6 is provided on the thrust chamber body 1. When the engine is running, most of the external air will directly enter the interior of the thrust chamber body 1 as a combustion oxidant after being compressed by the compressor, and a small part will enter the interior of the air film cooling mechanism 2 under the action of the diverter channel 4.

[0027] In order to form a flowing air film on the inner wall of the thrust chamber body 1, an air film cooling mechanism 2 is provided in the present embodiment. Specifically, the air film cooling mechanism 2 includes an air film channel 202 opened inside the side wall of the thrust chamber body 1. A flat-seam air film hole 201, a second air film hole 203 and a third air film hole 204 are sequentially provided on the inner wall of the thrust chamber body 1 from left to right. A part of the gas diverted by the diversion channel 4 will be ejected through the flat-seam air film hole 201, and the other part will enter the interior of the air film channel 202. The thrust chamber body 1 includes a straight section 101 and a contraction-expansion section 102. The flat-seam air film hole 201 is parallel to the straight section 101. After the airflow is ejected through the flat-seam air film hole 201, it will move along the side wall of the straight section 101 to form an air film. The second air film hole 203 and the third air film hole 204 are both in the form of rows of holes. The second air film hole 203 and the third air film hole 204 are both located in the contraction-expansion section 102. The second air film hole 203 is located near the position where the thrust chamber body 1 profile begins to contract. The third air film hole 204 is to cover the area with the highest absolute value of the throat heat flux density and the highest wall temperature.

[0028] According to the above content, when the oxygen-rich turbine engine is working, the compressor compresses the external air to form a compressed air flow 5 and transmits it to the thrust chamber body 1. Next, Figure 2 As shown, the compressed airflow 5 is divided into two parts, namely, a combustion airflow 501 and a cold source airflow 502 , under the action of the flow splitting channel 4 . Moreover, the cold source airflow 502 does not exceed 10% of the total amount of the compressed airflow 5 .

[0029] Subsequently, the combustion airflow 501 will directly enter the interior of the thrust chamber body 1 as a combustion oxidant under the action of the injector 3, and the cold source airflow 502 will enter the interior of the diversion channel 4. Next, the cold source airflow 502 will be divided into a first cold airflow 503 and a second cold airflow 504. The first cold airflow 503 will be ejected through the flat-slit air film hole 201, thereby forming a flowing air film on the inner wall of the straight section 101.

[0030] The second cold air flow 504 enters the interior of the air film channel 202 and is ejected through the second air film holes 203 and the third air film holes 204 , thereby forming a flowing air film on the inner wall of the contraction-expansion section 102 .

[0031] In this embodiment, an air film cooling mechanism 2 is provided, and the cooperation between the flat-seam air film holes 201 and the air film channels 202 is utilized to form a flowing air film on the inner wall of the thrust chamber body 1, thereby cooling the thrust chamber body 1. Compared with the liquid film cooling method in the prior art, there is no need to consume liquid fuel as a cold source, thereby effectively reducing the energy consumption of the oxygen-rich turbine engine. Moreover, in this embodiment, an air film cooling mechanism 2 is provided, and the second air film holes 203 are located near the position where the thrust chamber line begins to contract, and the third air film holes 204 are located near the position where the thrust chamber line begins to expand, so that the air film can be more firmly attached to the inner wall of the contraction and expansion section 102, thereby covering the area with the highest absolute value of the throat heat flux density and the highest wall temperature, thereby effectively improving the cooling effect.

[0032] Embodiment 2

[0033] In order to improve the cooling effect as much as possible and realize the regeneration cycle of liquid fuel, on the basis of embodiment 1, this embodiment is provided with a regeneration cooling mechanism 6. Specifically, the regeneration cooling mechanism 6 includes a cooling channel 603 opened inside the side wall of the thrust chamber body 1. The thrust chamber body 1 is provided with a fuel injection port 601 and a fuel discharge port 602. The fuel discharge port 602 is connected to the interior of the thrust chamber body 1. After the liquid fuel participating in cooling is discharged through the fuel discharge port 602, it will enter the interior of the thrust chamber body 1 for combustion. The fuel injection port 601 and the fuel discharge port 602 are respectively connected to the cooling channel 603. When the engine is running, the liquid fuel will enter the cooling channel 603 through the fuel injection port 601, and then be discharged through the fuel discharge port 602. A plurality of cooling channels 603 are opened at equal intervals. Two air film channels 202 are arranged between every two cooling channels 603. The outlet pressure of each air film is higher than the chamber pressure of the thrust chamber body 1.

[0034] According to the above content, when the oxygen-rich turbine engine is working, a part of the liquid fuel (kerosene) will enter the cooling channel 603 through the fuel injection port 601 and flow inside the cooling channel 603 at a certain rate, thereby cooling the inner wall of the thrust chamber body 1.

[0035] Next, the liquid fuel in the cooling channel 603 will pass through the fuel outlet 602 and enter the thrust chamber body 1 for combustion, thereby avoiding the waste of this part of the liquid fuel.

[0036] Moreover, when the liquid fuel flows in the cooling channel 603, the heat inside the thrust chamber body 1 will preheat the liquid fuel, which can make the fuel burn more completely and improve the overall performance of the engine to a certain extent.

[0037] In this embodiment, by providing a regenerative cooling mechanism 6 and utilizing the mutual cooperation between the fuel injection port 601 and the fuel discharge port 602, the thrust chamber body 1 can be cooled by utilizing liquid fuel without consuming additional coolant, and this part of the liquid fuel can be burned and utilized, which can effectively reduce resource waste; moreover, in this embodiment, by providing a regenerative cooling mechanism 6, the liquid fuel will exchange heat with the inside of the thrust chamber body 1 during the flow of the liquid fuel from the cooling channel 603, thereby increasing the temperature of the liquid fuel itself, which can ensure more complete combustion of the liquid fuel and improve the overall performance of the engine; in addition, in this embodiment, by providing an air film cooling mechanism 2 and a regenerative cooling mechanism 6, external air is inhaled as a cold source for combustion oxidant and air film cooling, and liquid fuel is used for cooling, thereby perfectly combining regenerative cooling and air film cooling, which can reduce the weight of the engine and improve the overall performance of the engine.

[0038] It should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device.

[0039] Although embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions and variations may be made to the embodiments without departing from the principles and spirit of the present invention, and that the scope of the present invention is defined by the appended claims and their equivalents.

Claims

1. A thrust chamber thermal protection structure for an oxygen-rich turbine engine, comprising a thrust chamber body (1), characterized in that: An air film cooling mechanism (2) for inhaling low-temperature gas to achieve temperature reduction is provided inside the side wall of the thrust chamber body (1); an injector (3) and a flow diversion channel (4) are provided at the air inlet end of the thrust chamber body (1); a compressed air flow (5) is provided on the side of the thrust chamber body (1); and a regenerative cooling mechanism (6) is provided on the thrust chamber body (1); The air film cooling mechanism (2) comprises an air film channel (202) opened inside the side wall of the thrust chamber body (1), and the inner side wall of the thrust chamber body (1) is provided with a flat seam air film hole (201), a second air film hole (203) and a third air film hole (204) in sequence from left to right.

2. The thermal protection structure of the thrust chamber of an oxygen-rich turbine engine according to claim 1, characterized in that: The thrust chamber body (1) comprises a straight section (101) and a contraction-expansion section (102), and the flat-seam type air film hole (201) is parallel to the straight section (101).

3. The thermal protection structure of the thrust chamber of an oxygen-rich turbine engine according to claim 2, characterized in that: The second row of air film holes (203) and the third row of air film holes (204) are both in the form of rows of holes, and the second row of air film holes (203) and the third row of air film holes (204) are both located in the contraction and expansion section (102).

4. The thermal protection structure of the thrust chamber of an oxygen-rich turbine engine according to claim 1, characterized in that: The compressed air flow (5) includes a combustion air flow (501) and a cold source air flow (502), and the cold source air flow (502) includes a first cold air flow (503) and a second cold air flow (504).

5. The thermal protection structure of the thrust chamber of an oxygen-rich turbine engine according to claim 1, characterized in that: The regenerative cooling mechanism (6) comprises a cooling channel (603) opened inside the side wall of the thrust chamber body (1); a fuel injection port (601) and a fuel discharge port (602) are provided on the thrust chamber body (1); the fuel injection port (601) and the fuel discharge port (602) are respectively connected to the cooling channel (603).

6. The thermal protection structure of the thrust chamber of an oxygen-rich turbine engine according to claim 5, characterized in that: The fuel discharge port (602) is connected to the interior of the thrust chamber body (1).

7. The thermal protection structure of the thrust chamber of an oxygen-rich turbine engine according to claim 5, characterized in that: A plurality of cooling channels (603) are provided at equal intervals, and two air film channels (202) are arranged between every two cooling channels (603).

Images