CN106930864B - A kind of supersonic speed detonation engine and its propulsion system - Google Patents

Abstract

The invention discloses a propulsion system, which comprises an air inlet, a combustion chamber and an expansion nozzle connected with the combustion chamber. The second hot jet nozzle ejected by the hot jet, the first hot jet nozzle and the second hot jet nozzle can induce the supersonic incoming flow after ejecting the hot jet to form two oblique detonation waves, and the two oblique detonation waves can be formed in the expansion nozzle intersects in the middle. In the propulsion system provided by the present invention, the first and second hot jet nozzles are provided in the combustion chamber, and the hot jet flow ejected from the two hot jet nozzles induces the incoming flow to form a stable oblique detonation wave, and the oblique detonation wave is regularly reflected to avoid The effect of the detonation wave and the boundary layer ensures the full combustion of the incoming premixed gas. After burning through the double oblique detonation wave, heat is released rapidly, and the expansion generates thrust, which improves the combustion efficiency. The invention also discloses a supersonic detonation engine comprising the propulsion system.

Description

A supersonic detonation engine and its propulsion system

technical field

The invention relates to the field of detonation engine equipment, in particular to a propulsion system. In addition, the present invention also relates to a supersonic detonation engine comprising the above propulsion system.

Background technique

With the development of hypersonic vehicles, the efficiency of the propulsion system has become one of the bottlenecks restricting the further improvement of engine thrust, while the detonation combustion thermodynamic cycle has high efficiency and fast heat release, and the supersonic propulsion system based on detonation combustion has a compact layout and a simple structure , therefore, detonation-based engines can be used as a potential solution for hypersonic propulsion systems.

Under hypersonic conditions, the detonation combustor organizes combustion through the slope-induced detonation method. However, since the existing detonation-based engine considers the slope-induced detonation wave formation method to organize the detonation combustion, the oblique detonation wave directly Acting on the inner wall of the expanding nozzle, it interacts with the boundary layer on the wall of the expanding nozzle, causing the oblique detonation wave to propagate forward along the boundary layer, and the boundary layer has a greater influence on the oblique detonation wave. detonation, resulting in greater aerodynamic drag and aerodynamic heat.

At the same time, in the propulsion system in the prior art, the combustion chamber and the expansion nozzle are arranged separately. Since the combustion chamber needs to mix fuel, the length of the combustion chamber is generally long, resulting in a large volume of the entire propulsion system and a complex structure of the engine. .

Therefore, how to reduce the impact of the boundary layer on the oblique detonation wave in the propulsion system and at the same time ensure the full combustion of the fuel is a technical problem to be solved by those skilled in the art.

Contents of the invention

The object of the present invention is to provide a kind of propulsion system, this propulsion system utilizes two hot jets to detonate the incoming premixed gas, two oblique detonation waves intersect in the middle of the expansion nozzle, effectively reducing the impact of the boundary layer on the oblique detonation shock wave impact while ensuring full combustion of the fuel. Another object of the present invention is to provide a supersonic detonation engine comprising the above propulsion system.

To achieve the above object, the present invention provides the following technical solutions:

A propulsion system, comprising an air inlet, a combustion chamber and an expansion nozzle connected to the combustion chamber, the upper wall of the combustion chamber is provided with a first hot jet nozzle for hot jet ejection, and the lower wall is provided with a The second hot jet nozzle ejected by the hot jet, the first hot jet nozzle and the second hot jet nozzle can induce the supersonic incoming flow after ejecting the hot jet to form two oblique detonation shock waves, and the two oblique explosions The shock waves may intersect in the middle of the expanding nozzle.

Preferably, the opening of the first heat jet nozzle is downward and inclined away from the air inlet; the opening of the second heat jet nozzle is upward.

Preferably, both the first hot jet nozzle and the second hot jet nozzle are arranged at one end of the combustion chamber close to the expansion nozzle.

Preferably, the first heat jet nozzle and the second heat jet nozzle are respectively connected with a first heat jet delivery pipeline and a second heat jet delivery pipeline.

Preferably, it also includes a controller connected to the first heat jet nozzle and the second heat jet nozzle, the controller is used for:

Obtain the flow velocity of the incoming flow;

According to the flow velocity of the incoming flow, calculate the first hot jet optimum injection intensity and the second hot jet optimum injection intensity of the first hot jet nozzle and the second hot jet nozzle respectively;

The intensities of the hot jets ejected from the first hot jet nozzle and the second hot jet nozzle are respectively adjusted to the optimum jet intensity of the first hot jet and the optimum jet intensity of the second hot jet.

Preferably, the combustion chamber and the expansion nozzle form an integrated combustion chamber expansion nozzle.

Preferably, the air inlet communicates with the combustion chamber through an isolation section.

Preferably, the front end of the intake passage is provided with a precursor fuel injection port, and the precursor fuel injection port is used to inject fuel into the intake passage to form premixed fuel.

The present invention also provides a supersonic detonation engine, including the propulsion system described in any one of the above.

The propulsion system provided by the present invention includes an air inlet, a combustion chamber and an expansion nozzle connected with the combustion chamber, the upper wall of the combustion chamber is provided with a first hot jet nozzle for the ejection of the heat jet, and the lower wall is There is a second hot jet nozzle for hot jet ejection, the first hot jet nozzle and the second hot jet nozzle can induce supersonic incoming flow after ejecting hot jet to form two oblique detonation waves, and the two The oblique detonation waves may intersect in the middle of the expanding nozzle. The propulsion system, by opening the first hot jet nozzle and the second hot jet nozzle in the combustion chamber, the hot jets ejected from the two hot jet nozzles respectively induce the incoming flow to form a stable oblique detonation shock wave, and the oblique detonation The regular reflection of the shock wave avoids the effect of the detonation wave and the boundary layer, and at the same time ensures the full combustion of the incoming premixed gas. After the premixed gas is burned through the double oblique detonation wave, it quickly releases heat and expands to generate thrust, which effectively improves the combustion efficiency. .

The supersonic detonation engine provided by the present invention is provided with the above-mentioned propulsion system. Since the propulsion system has the above-mentioned technical effects, the supersonic detonation engine provided with the propulsion system should also have corresponding technical effects.

Description of drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings that need to be used in the description of the embodiments or the prior art. Obviously, the accompanying drawings in the following description are only These are some embodiments of the present invention. Those skilled in the art can also obtain other drawings based on these drawings without creative work.

Fig. 1 is a structural schematic diagram of a specific embodiment of the propulsion system provided by the present invention;

Among them: 1-precursor fuel injection port, 2-premixed fuel, 3-intake inlet, 5-first hot jet nozzle, 4-second hot jet nozzle, 7-first oblique detonation wave, 6- The second oblique detonation wave, A-intake port, B-isolating section, C-integrated combustion chamber expansion nozzle.

Detailed ways

The core of the invention is to provide a propulsion system, which can significantly reduce the impact of the boundary layer on the oblique detonation wave, and at the same time ensure sufficient combustion of fuel. Another core of the present invention is to provide a supersonic detonation engine comprising the above propulsion system.

The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only some, not all, embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by persons of ordinary skill in the art without making creative efforts belong to the protection scope of the present invention.

Please refer to FIG. 1 , which is a schematic structural diagram of a specific embodiment of the propulsion system provided by the present invention.

In this embodiment, the propulsion system includes an air inlet A, a combustion chamber, and an expansion nozzle connected to the combustion chamber. The combustion chamber is located between the air intake A and the expansion nozzle. The first hot jet nozzle 5 ejected, the lower wall is provided with the second hot jet nozzle 4 for hot jet ejection, the first hot jet nozzle 5 and the second hot jet nozzle 4 can induce supersonic incoming flow after ejecting the hot jet Two oblique detonation waves are formed, and the two oblique detonation waves may intersect in the middle of the expanding nozzle.

The propulsion system adopts dual-jet detonation, and the design state is a hypersonic state. The incoming flow velocity in the combustion chamber is much higher than the flame forward velocity of the oblique detonation combustion wave, so that the oblique detonation shock wave can be stabilized in the combustion chamber.

Specifically, under hypersonic flight conditions, fuel is injected into the precursor of inlet port A, and the fully mixed fuel enters the combustion chamber at supersonic speed after passing through a short isolation section B, and a jet of fuel is injected into the upper and lower walls of the combustion chamber. Hot jets, i.e. the first hot jet nozzle 5 and the second hot jet nozzle 4 each eject a hot jet, the hot jet ejected in the first hot jet nozzle 5 induces incoming flow to form the first oblique detonation wave 7, the second The hot jet flow ejected from the second hot jet nozzle 4 induces the incoming flow to form the second oblique detonation wave 6, the first oblique detonation wave 7 and the second oblique detonation wave 6 make the fuel burn and release heat rapidly, the first oblique detonation wave 7 and the second oblique detonation wave 6 After intersecting with the second oblique detonation wave 6, reflection occurs on the profile of the expansion nozzle immediately behind the combustion chamber, that is, the wall surface of the expansion nozzle, and the combustion products after the formation of the oblique detonation wave expand rapidly to generate thrust.

The propulsion system, by opening the first hot jet nozzle 5 and the second hot jet nozzle 4 in the combustion chamber, the hot jets ejected from the two hot jet nozzles respectively induce the incoming flow to form a stable oblique detonation wave, and the oblique detonation wave occurs The regular reflection avoids the effect of the detonation wave and the boundary layer, and at the same time ensures the full combustion of the incoming premixed gas. After the premixed gas burns through the double oblique detonation wave, it quickly releases heat and expands to generate thrust, which effectively improves the combustion efficiency.

Further, the opening of the first heat jet nozzle 5 is downward and inclined away from the air inlet A; the opening of the second heat jet nozzle 4 is upward. The above arrangement can facilitate the formation of two bow-shaped shock waves induced by the hot jets ejected from the two hot jet nozzles to form two bow-shaped shock waves, and then rapidly ignite the incoming combustible gas to form an oblique detonation shock wave.

On the basis of the above-mentioned embodiments, the first hot jet nozzle 5 and the second hot jet nozzle 4 are both arranged at one end of the combustion chamber close to the expanding nozzle, so as to facilitate rapid combustion in the expanding nozzle during oblique detonation waves.

On the basis of the above embodiments, the first heat jet nozzle 5 and the second heat jet nozzle 4 are respectively connected with a first heat jet delivery pipeline and a second heat jet delivery pipeline.

Specifically, the hot jet is delivered to the first hot jet nozzle 5 through the first hot jet delivery pipeline, and the hot jet is delivered to the second hot jet nozzle 4 through the second hot jet delivery pipeline. The intensity of the hot jet can be controlled separately. Specifically, the opening or closing of the hot jet and the adjustment of the intensity of the hot jet can be controlled by valves respectively installed on the two hot jet delivery pipelines.

It should be noted here that the injection angle of the hot jets ejected in the first hot jet nozzle 5 and the second hot jet nozzle 4 can also be adjusted according to the change of the flow condition, so that the intensity and height of the two hot jets It can match the incoming flow and burn stably.

On the basis of the above-mentioned embodiments, the propulsion system also includes a controller connected to the first heat jet nozzle 5 and the second heat jet nozzle 4, the controller is used for:

Obtain the flow velocity of the incoming flow;

According to the flow velocity of incoming flow, respectively calculate the first hot jet optimum injection intensity and the second hot jet optimum injection intensity of the first hot jet nozzle 5 and the second hot jet nozzle 4, specifically, the first hot jet of above-mentioned hot jet The principle of determining the optimal injection intensity and the optimal injection intensity of the second hot jet is to make the two oblique detonation waves formed after the incoming flow ignite converge in the middle of the expanding nozzle. The relationship between the optimal injection intensity of the first hot jet and the optimal injection intensity of the second hot jet of the jet, and use this as the data reference basis for the controller;

The intensity of the hot jets ejected from the first hot jet nozzle 5 and the second hot jet nozzle 4 is respectively adjusted to the optimum jet intensity of the first hot jet and the optimum jet intensity of the second hot jet.

The setting of the above-mentioned controller can not only control the opening and closing of the heat jet, but also adjust the intensity of the heat jet to the optimum value according to the flow velocity of the incoming flow, so as to ensure the working stability of the engine.

On the basis of the above-mentioned embodiments, the end of the combustion chamber away from the intake port A is provided with an expansion nozzle, and the combustion chamber and the expansion nozzle are integrally structured, and the combustion chamber and the expansion nozzle together form an integrated combustion chamber expansion nozzle. C, Combining the oblique detonation combustor and the thrust expansion nozzle can effectively reduce the reflection angle between the oblique detonation wave and the wall of the expansion nozzle.

On the basis of the above-mentioned embodiments, the air intake port A communicates with the combustion chamber through the isolation section B, which constitutes the isolation section B of the propulsion system, and the isolation section B is located between the air intake port A and the integrated combustion chamber. Between the expansion nozzles C, the isolation section B not only has the function of isolating the mutual interference between the intake port A and the combustion chamber, the fuel can be further mixed in the relatively long distance of the isolation section B, and at the outlet of the isolation section B , that is, the entrance of the combustion chamber, forming a uniform premixed combustible gas that is fully mixed.

On the basis of the above-mentioned embodiments, the front end of the intake port A is provided with a precursor fuel injection port 1 for injecting fuel into the intake port A to premix the fuel 2 . Specifically, the intake port A of the propulsion system is mainly responsible for capturing the incoming flow, so that more incoming flow enters the combustion chamber after being compressed by the precursor of the intake port A; the intake port A also has the function of fuel blending. There is a precursor fuel injection port 1 at the precursor position of the air channel A, and hydrogen fuel is injected, and the injected fuel has a certain penetration depth; the precursor distance of the air inlet channel A is long, and the fuel is basically mixed when reaching the isolation section B Evenly, the fuel is just able to fully enter the intake port 3 without overflowing from the lower lip, i.e. the intake port 3.

In this propulsion system, since the front end of the intake port A is provided with a precursor fuel injection port 1, and an isolation section B is provided between the intake port A and the combustion chamber, the fuel can be discharged before the intake port A. When it reaches the combustion chamber, it has been mixed evenly, the length of the combustion chamber can be shortened as much as possible, and then forms an integrated structure with the expansion nozzle. The integrated combustion chamber expansion nozzle C referred to here refers to One end of the expansion nozzle can be provided with a length sufficient to form an oblique detonation wave, so that the total length of the integrated combustion chamber expansion nozzle C is significantly reduced compared to the length of the combustion chamber and the expansion nozzle in the prior art, so that the engine The length of the motor is reduced, the structure is simplified, the production cost is saved, the weight is reduced, and the performance of the engine is improved.

The propulsion system provided in this embodiment adopts the pre-injection of the premixed fuel 2 at the precursor position of the intake port A, the combination of the combustion chamber and the expansion nozzle, and the detonation of the double hot jets to form intersecting oblique detonation waves. As shown in Figure 1, the integrated thermal jet detonation oblique detonation propulsion system is mainly composed of an intake port A, an isolation section B, an integrated combustion chamber expansion nozzle C, a precursor fuel injection port 1, and a first thermal jet nozzle 5. Composition of the second hot jet nozzle 4.

The propulsion system adopts dual hot jets for ignition and detonation. At the end of the isolation section B, before the combustion chamber, the fuel is fully mixed, and the speed is much greater than the forward speed of the oblique detonation combustion wave. Two hot jets with small apertures that can control the intensity flow from The injection from the upper and lower walls of the combustion chamber, that is, from the first hot jet nozzle 5 and the second hot jet nozzle 4, induces the supersonic incoming flow to form two bow-shaped shock waves, and the bow-shaped shock wave quickly ignites the incoming flow of combustible gas in an oblique explosion Shock waves, two oblique detonation waves intersect in the middle of the expansion nozzle behind the combustion chamber; at the wave surface of the oblique detonation wave, the fuel burns rapidly and becomes combustion products and releases a large amount of heat. Expansion occurs at the tube and the velocity increases, allowing the propulsion system to generate thrust.

Among them, the integrated combustion chamber expansion nozzle C is another feature of this propulsion system. The oblique detonation shock wave burns rapidly, so the design of the combustion chamber is very compact, and there is no obvious boundary with the tail expansion nozzle. The upper and lower walls of the tail expansion nozzle are expanded. The shape ensures that the combustion products after the two oblique detonation waves fully expand to generate thrust.

The propulsion system uses two controllable hot jets to detonate the oblique detonation wave, avoiding the effect of a single oblique detonation wave and the boundary layer, avoiding the influence of the boundary layer on the oblique detonation wave, and ensuring full combustion of fuel.

In addition to the above-mentioned propulsion system, the present invention also provides a supersonic detonation engine including the above-mentioned propulsion system. For the structure of other parts of the supersonic detonation engine, please refer to the prior art, which will not be repeated here.

Each embodiment in this specification is described in a progressive manner, each embodiment focuses on the difference from other embodiments, and the same and similar parts of each embodiment can be referred to each other.

The propulsion system provided by the present invention has been introduced in detail above. In this paper, specific examples are used to illustrate the principle and implementation of the present invention, and the descriptions of the above embodiments are only used to help understand the method and core idea of the present invention. It should be pointed out that for those skilled in the art, without departing from the principle of the present invention, some improvements and modifications can be made to the present invention, and these improvements and modifications also fall within the protection scope of the claims of the present invention.

Claims (8)

1. A propulsion system, characterized in that it comprises an air inlet (A), a combustion chamber and an expansion nozzle connected to the combustion chamber, the upper wall of the combustion chamber is provided with a first The hot jet nozzle (5), the lower wall is provided with the second hot jet nozzle (4) for the hot jet ejection, the first hot jet nozzle (5) and the second hot jet nozzle (4) can eject After the hot jet induces supersonic incoming flow to form two oblique detonation waves, and the two oblique detonation waves can intersect in the middle of the expansion nozzle; It also includes a controller connected to the first heat jet nozzle (5) and the second heat jet nozzle (4), and the controller is used for: Obtain the flow velocity of the incoming flow; Calculating the optimum injection intensity of the first heat jet and the optimum injection intensity of the second heat jet of the first heat jet nozzle (5) and the second heat jet nozzle (4) respectively according to the flow velocity of the incoming flow; Respectively adjust the intensity of the hot jets ejected from the first hot jet nozzle (5) and the second hot jet nozzle (4) to the optimum jet intensity of the first hot jet and the optimum jet intensity of the second hot jet Jet intensity. 2. The propulsion system according to claim 1, characterized in that, the opening of the first hot jet nozzle (5) is downward and inclined away from the air inlet (A); the second The opening of the hot jet nozzle (4) is upward. 3. The propulsion system according to claim 1, characterized in that, the first hot jet nozzle (5) and the second hot jet nozzle (4) are both arranged in the combustion chamber close to the expansion nozzle one end. 4. The propulsion system according to claim 1, characterized in that, the first hot jet nozzle (5) and the second hot jet nozzle (4) are respectively connected with a first hot jet delivery pipeline and a second Heat jet delivery pipeline. 5. The propulsion system according to any one of claims 1 to 4, characterized in that the combustion chamber and the expansion nozzle form an integrated combustion chamber expansion nozzle (C). 6. The propulsion system according to any one of claims 1 to 4, characterized in that, the air inlet (A) communicates with the combustion chamber through an isolation section (B). 7. The propulsion system according to any one of claims 1 to 4, characterized in that, the front end of the air inlet (A) is provided with a precursor fuel injection port (1), and the precursor fuel injection The port (1) is used for injecting fuel into the intake passage (A) to form premixed fuel (2). 8. A supersonic detonation engine, comprising a propulsion system, characterized in that the propulsion system is the propulsion system according to any one of claims 1-7.