An aero turbine engine rotating detonation augmentor combustor

Abstract

The application belongs to the technical field of aviation turbine engine afterburner design, and particularly relates to an aviation turbine engine rotating detonation afterburner, which comprises an outer wall with an annular opening, a splitter ring arranged in the outer wall and located at an inlet of the outer wall to form an outer duct with the outer wall, an inner cone arranged in the splitter ring to form an inner duct with the splitter ring, a rear end of the inner cone extending out of an outlet of the splitter ring, an outward annular protrusion on the outer wall, the outward annular protrusion being located behind the outlet of the splitter ring, a sliding casing with a front end sleeved on the outer wall, a rear end extending into the outer wall through the annular opening, and an inner surface of the outer wall, the sliding casing being in sliding connection with the outer wall and capable of sliding along the outer wall in an axial direction, and an inward annular protrusion on the sliding casing, the inward annular protrusion being located behind the outward annular protrusion.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of afterburner design of an aero turbine engine, and particularly relates to a rotating detonation afterburner of an aero turbine engine. BACKGROUND

[0002] The existing aero turbine engine generally adopts slow-burning isobaric combustion, and the slow-burning wave propagation speed is slow, which is a large entropy-increasing process, and the thermal efficiency is not high. With the progress of technology, the aerodynamic performance and combustion efficiency of the aero turbine engine under the slow-burning isobaric combustion have approached the limit, and it is very difficult to further improve the performance.

[0003] Detonation combustion is a kind of supercharged combustion, and the combustion speed is fast, the supercharging effect is strong, the entropy increase is small, and the thermal efficiency is high. In the rotating detonation afterburner of the aero turbine engine, the afterburner is improved into a rotating detonation afterburner, which can greatly improve the efficiency of the aero turbine engine, reduce the number of compressor stages, improve the upper limit of the aero turbine engine, reduce the axial length of the afterburner, reduce the weight of the aero turbine engine, and increase the thrust-to-weight ratio of the aero turbine engine.

[0004] At present, the research on detonation combustion mostly focuses on the research on the mechanism of rotating detonation initiation and control, fuel atomization and evaporation and mixing, and stable working boundary. From the requirements of preventing reverse pressure conduction of detonation wave and rapidly establishing a combustible filling area, a converging-diverging channel is constructed at the inlet of the rotating detonation afterburner, so that the airflow entering the rotating detonation afterburner reaches supersonic speed at the inlet, forms the ignition starting condition of rotating detonation afterburner, and can prevent the high-pressure products after downstream detonation combustion from conducting to the upstream, and can accelerate the filling of combustible. However, the technical scheme has the following defects:

[0005] Under the non-afterburning state, the inlet airflow of the rotating detonation afterburner of the aero turbine engine and the outlet airflow of the nozzle reach supersonic speed at the same time, which will increase the total pressure loss, affect the thrust of the aero turbine engine, and weaken the regulation ability of the nozzle to the state of the aero turbine engine.

[0006] The present application is proposed in view of the above technical defects.

[0007] It should be noted that the disclosure of the above background art content is only used to assist in understanding the inventive concept and technical solutions of the present application, and it does not necessarily belong to the prior art of the present application. In the absence of explicit evidence that the above content has been disclosed on the filing date of the present application, the above background art should not be used to evaluate the novelty and inventiveness of the present application. SUMMARY

[0008] The purpose of the present application is to provide an aero turbine engine rotating detonation ramjet combustor to overcome or alleviate at least one of the technical defects known to exist.

[0009] The technical solution of the present application is:

[0010] An aero turbine engine rotating detonation ramjet combustor comprises:

[0011] An outer wall having an annular opening thereon;

[0012] A splitter ring arranged in the outer wall and located at the inlet of the outer wall to form an outer duct with the outer wall;

[0013] An inner cone arranged in the splitter ring to form an inner duct with the splitter ring, the rear end of the inner cone extending out of the outlet of the splitter ring, and the outer wall having an outward annular protrusion; the outward annular protrusion is located behind the outlet of the splitter ring;

[0014] A sliding casing, the front end of which is sleeved on the outer wall and abuts against the outer surface of the outer wall, and the rear end of which extends into the outer wall through the annular opening and abuts against the inner surface of the outer wall to be in sliding connection with the outer wall, so that the sliding casing can slide along the outer wall in the axial direction, and the sliding casing has an inward annular protrusion; the inward annular protrusion is located behind the outward annular protrusion.

[0015] According to at least one embodiment of the present application, in the aero turbine engine rotating detonation ramjet combustor described above, the sliding casing and the outer wall are connected through a sliding rail.

[0016] According to at least one embodiment of the present application, in the aero turbine engine rotating detonation ramjet combustor described above, the cross section of the outward annular protrusion and the inward annular protrusion is in the shape of a circular arc.

[0017] According to at least one embodiment of the present application, in the aero turbine engine rotating detonation ramjet combustor described above, further comprising:

[0018] An actuating cylinder connected between the turbine outer casing and the sliding casing to drive the sliding casing to slide along the outer wall in the axial direction. BRIEF DESCRIPTION OF DRAWINGS

[0019] Fig. 1 is a schematic diagram of the aero turbine engine rotating detonation ramjet combustor in the ramjet state provided by the embodiment of the present application;

[0020] Fig. 2 is a schematic diagram of the aero turbine engine rotating detonation ramjet combustor in the non-ramjet state provided by the embodiment of the present application;

[0021] Wherein:

[0022] 1 - outer wall; 2 - splitter ring; 3 - inner cone; 4 - sliding casing; 5 - actuator cylinder; 6 - turbine; 7 - main combustion chamber; 8 - compressor; 9 - fan; 10 - afterburner; 11 - nozzle.

[0023] In order to better illustrate the present embodiment, some components in the drawings can be omitted, enlarged or reduced, and do not represent the actual size of the product. In addition, the drawings are only used for illustrative description and cannot be understood as a limitation of the present application. DETAILED DESCRIPTION

[0024] In order to make the technical solutions of the present application and its advantages clearer, the technical solutions of the present application will be further clearly and completely described in detail below in combination with the drawings. It should be understood that the specific embodiments described herein are only part of the embodiments of the present application, which are only used to explain the present application, but not to limit the present application. It should be noted that, for the purpose of description, only parts related to the present application are shown in the drawings, and other related parts can be referred to the general design. In the case of no conflict, the embodiments in the present application and the technical features in the embodiments can be combined to obtain new embodiments.

[0025] In addition, unless otherwise defined, the technical terms or scientific terms used in the present application description should be the general meaning understood by the general technical personnel in the field of the present application. The words indicating the relative direction or position relationship, such as "up", "down", "left", "right", "center", "vertical", "horizontal", "inner", "outer" and the like used in the present application description, are only used to indicate the relative direction or position relationship, and not to imply that the device or element must have a specific orientation, be constructed and operated in a specific orientation, and the relative position relationship may also change accordingly when the absolute position of the described object changes, therefore, it cannot be understood as a limitation of the present application. The "first", "second", "third" and similar terms used in the present application description are only for the purpose of description, to distinguish different components, and cannot be understood as indicating or implying relative importance. The "one", "an" or "the" and similar terms used in the present application description should not be understood as an absolute limitation on the quantity, but should be understood as the existence of at least one. The "include" or "contain" and similar terms used in the present application description mean that the elements or objects appearing before the word cover the elements or objects listed after the word and their equivalents, and do not exclude other elements or objects.

[0026] In addition, it needs to be explained that, unless otherwise explicitly specified and limited, the "installation", "connection", "connection" and similar words used in the description of the application should be understood broadly, for example, the connection can be fixed connection, or detachable connection, or integrally connected; can be mechanical connection, or electrical connection; can be directly connected, or indirectly connected through intermediate medium, or the internal connection of two elements, and those skilled in the art can understand the specific meaning of the application according to the specific circumstances.

[0027] The aviation turbine engine comprises a fan, a compressor, a main combustion chamber, a turbine, a booster combustion chamber and a nozzle which are sequentially connected, wherein the booster combustion chamber comprises an outer wall, a splitter ring arranged in the outer wall, an inner cone arranged in the splitter ring, an outer channel formed between the outer wall and the splitter ring, and an inner channel formed between the splitter ring and the inner cone; the inner and outer channel airflow from the turbine outlet enters the inner and outer channel of the booster combustion chamber respectively, and mixing occurs at the outlet of the splitter ring; fuel injection and ignition are performed, expansion combustion is performed along the inner cone, and the aviation turbine engine rotary knock booster combustion chamber is provided on the basis. Figs. 1-2 The application is further described in detail.

[0028] The aviation turbine engine rotary knock booster combustion chamber comprises:

[0029] The outer wall 1 has an annular opening thereon;

[0030] The splitter ring 2 is arranged in the outer wall 1 and located at the inlet of the outer wall 1 to form an outer channel with the outer wall 1;

[0031] The inner cone 3 is arranged in the splitter ring 2 to form an inner channel with the splitter ring 2, and the rear end of the inner cone 3 extends out of the outlet of the splitter ring 2, and the outer wall has an outward annular protrusion; the outward annular protrusion is located behind the outlet of the splitter ring 2;

[0032] The sliding casing 4 is sleeved on the outer wall 1 and abuts against the outer surface of the outer wall 1 at the front end, and is sleeved into the outer wall 1 through the annular opening at the rear end and abuts against the inner surface of the outer wall 1 to be slidably connected with the outer wall 1, so that the sliding casing 4 can slide along the outer wall 1 in the axial direction, and the sliding casing 4 has an inward annular protrusion; the inward annular protrusion is located behind the outward annular protrusion.

[0033] The aviation turbine engine rotary knock booster combustion chamber disclosed in the above embodiment can drive the sliding casing 4 to slide along the outer wall 1 in the axial direction, so that the booster combustion chamber 10 is in:

[0034] The booster state, the sliding casing 4 is driven to slide along the outer wall 1 in the axial direction, so that the inward annular protrusion approaches the outward annular protrusion, the throat area A65 between the inward annular protrusion and the outward annular protrusion is reduced, and the booster combustion chamber 10 is in the booster state;Fig. 1 As shown, the mixed gas flow at the outlet of the splitter ring 2 is made to reach supersonic speed when flowing through the outlet, so that the igniting start condition of the rotating detonation force can be quickly established, and the reverse pressure conduction of the detonation wave is prevented, and then the mixed gas flow is ejected through the nozzle 11, so that the efficiency of the aero turbine engine can be greatly improved, the number of compressor stages can be reduced, the upper limit of the working capacity of the aero turbine engine can be improved, the axial length of the afterburner can be reduced, the weight of the aero turbine engine can be reduced, and the thrust-to-weight ratio of the aero turbine engine can be increased.

[0035] In the non-afterburning state, the sliding casing 4 is driven to slide axially along the outer wall 1, so that the inward annular protrusion is away from the outward annular protrusion, and the throat area A65 between the inward annular protrusion and the outward annular protrusion is increased, as shown in Fig. 2 As shown, the mixed gas flow at the outlet of the splitter ring 2 is made to reach supersonic speed when flowing through the outlet, so that the igniting start condition of the rotating detonation force can be quickly established, and the reverse pressure conduction of the detonation wave is prevented, and then the mixed gas flow is ejected through the nozzle 11, so that the efficiency of the aero turbine engine can be greatly improved, the number of compressor stages can be reduced, the upper limit of the working capacity of the aero turbine engine can be improved, the axial length of the afterburner can be reduced, the weight of the aero turbine engine can be reduced, and the thrust-to-weight ratio of the aero turbine engine can be increased.

[0036] In some alternative embodiments, the aero turbine engine rotating detonation afterburner described above, the sliding casing 4 and the outer wall 1 are connected through a sliding rail.

[0037] In some alternative embodiments, the aero turbine engine rotating detonation afterburner described above, the cross section of the outward annular protrusion and the inward annular protrusion is in the shape of a circular arc, so as to ensure the aerodynamic performance of the afterburner 10.

[0038] In some alternative embodiments, the aero turbine engine rotating detonation afterburner described above, further comprises:

[0039] The actuating cylinder 5 is connected between the turbine 6 outer casing and the sliding casing 4, so as to drive the sliding casing 4 to slide axially along the outer wall 1, and the operation is convenient.

[0040] The embodiments in the description are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts of each embodiment can be referred to each other.

[0041] So far, the technical solution of the present application has been described in combination with the preferred embodiments shown in the drawings. It should be understood by those skilled in the art that the protection scope of the present application is obviously not limited to these specific embodiments. Without deviating from the principles of the present application, those skilled in the art can make equivalent changes or replacements to the related technical features, and the technical solutions after the changes or replacements will fall within the protection scope of the present application.

Claims

1. An aero turbine engine rotating detonation augmentor, characterized by, It comprises: an outer wall (1) with an annular opening on it; a splitter ring (2) arranged in the outer wall (1) and located at the inlet of the outer wall (1) to form an outer channel with the outer wall (1); an inner cone (3) arranged in the splitter ring (2) to form an inner channel with the splitter ring (2), with its rear end extending out of the outlet of the splitter ring (2) and having an outward annular protrusion on the outer wall; the outward annular protrusion is located behind the outlet of the splitter ring (2); a sliding casing (4) with its front end sleeved on the outer wall (1) and its rear end extending into the outer wall (1) through the annular opening to abut against the inner surface of the outer wall (1) and be connected with the outer wall (1) in a sliding manner, so as to slide along the outer wall (1) in an axial direction, and with an inward annular protrusion on it; the inward annular protrusion is located behind the outward annular protrusion; an actuator (5) connected between the turbine (6) outer casing and the sliding casing (4) to drive the sliding casing (4) to slide along the outer wall (1) in an axial direction; by driving the sliding casing (4) to slide along the outer wall (1) in an axial direction, the aviation turbine engine rotating detonation afterburning chamber can be in: an afterburning state, in which the sliding casing (4) is driven to slide along the outer wall (1) in an axial direction forward, so that the inward annular protrusion approaches the outward annular protrusion and the throat area A65 between the inward annular protrusion and the outward annular protrusion is reduced, so that the mixed gas flow at the outlet of the splitter ring (2) can reach supersonic speed when flowing through the place, thereby quickly establishing a combustible filling area to form the ignition starting condition of rotating detonation afterburning, and preventing the reverse pressure conduction of detonation wave, and then being ejected through the nozzle; a non-afterburning state, in which the sliding casing (4) is driven to slide along the outer wall (1) in an axial direction backward, so that the inward annular protrusion moves away from the outward annular protrusion and the throat area A65 between the inward annular protrusion and the outward annular protrusion is increased, so that the mixed gas flow at the outlet of the splitter ring (2) can be reduced to subsonic speed when flowing through the place, avoiding the gas flow at the place and the gas flow at the outlet of the nozzle reaching supersonic speed at the same time.

2. The aviation turbine engine rotating detonation afterburning chamber according to claim 1, wherein the sliding casing (4) and the outer wall (1) are connected through a sliding rail.

3. The aviation turbine engine rotating detonation afterburning chamber according to claim 1, wherein the cross sections of the outward annular protrusion and the inward annular protrusion are in the shape of a circular arc. ​ ​

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