CN114165357B

CN114165357B - Rocket-based combined cycle engine based on detonation and detonation principles and application method

Info

Publication number: CN114165357B
Application number: CN202111484265.XA
Authority: CN
Inventors: 滕宏辉; 刘帅; 杨鹏飞; 郗雪辰; 冯占林
Original assignee: Beijing Institute of Technology BIT
Current assignee: Beijing Institute of Technology BIT
Priority date: 2021-12-07
Filing date: 2021-12-07
Publication date: 2022-12-06
Anticipated expiration: 2041-12-07
Also published as: CN114165357A

Abstract

The invention discloses a rocket-based combined cycle engine based on deflagration and detonation principles and an application method thereof, wherein the rocket-based combined cycle engine based on deflagration and detonation principles comprises: a center cone, a housing, a support, and a fuel supply system. Through forming the structure of two combustion chambers, two spray pipes, a propellant supply system of incomplete closed circulation is proposed, organically combines a liquid rocket engine and a ram engine based on deflagration and detonation principles, can meet the working requirements of low-speed start, wide-speed range/airspace range flight and hypersonic cruise, and has the characteristics of large specific impulse, strong robustness, high circulation heat efficiency and compact layout.

Description

Rocket-based combined cycle engine based on detonation and detonation principles and application method

Technical Field

The invention relates to the technical field of aerospace engines, in particular to a rocket-based combined cycle engine based on deflagration and detonation principles and an application method thereof.

Background

Combustion modes in nature can be divided into two categories, one being detonation and the other being detonation. Deflagration combustion is the propagation of deflagration waves by transferring energy released by reactions into a front unburnt mixture through heat conduction, molecular diffusion and turbulent flow transport mechanisms. Deflagration waves travel at relatively slow velocities, typically a few meters to tens of meters per second, subsonically relative to the reactants in front of them. The detonation combustion is mainly based on the detonation wave to organize the combustion, the detonation wave can be described as a strong shock wave with chemical reaction and capable of self-sustaining propagation, the propagation speed of the strong shock wave is far higher than the propagation speed of the detonation wave, the strong shock wave is an ultrasonic combustion wave, and the propagation speed can reach the magnitude of thousands of meters per second. The detonation combustion has the advantages of high energy release rate, high cycle thermal efficiency and the like, and has very wide development prospect in the field of hypersonic propulsion

The realization of wide-speed-range/airspace air-breathing type flight in the atmosphere is a research focus in the aerospace field all the time, and the adoption of a rocket-based stamping combined engine is a popular power generation scheme at present. The ramjet part of the traditional combined engine mainly organizes combustion based on a deflagration mode, and although flame stabilizing devices such as concave cavities are adopted, with the further improvement of Mach number, the reduction of flame stability and combustion sufficiency makes the engine difficult to generate effective thrust. Currently, rocket-based ramjet combos have fallen into bottlenecks in further broadening the air-breathing flight speed range.

Because detonation combustion has the advantages of fast energy release rate, high cycle thermal efficiency, and the like, researchers try to organize combustion by adopting detonation waves in hypersonic flight. In recent decades, researchers have proposed three propulsion schemes, namely pulse detonation, rotational detonation and oblique detonation based on detonation combustion, and have conducted extensive research on related problems. The invention patent with application publication number CN112228246A discloses a rocket-based knocking-punching combined cycle engine and a using method and application thereof, the rocket-based knocking-punching combined cycle engine disclosed by the invention mainly provides thrust by a punching type rotating knocking engine during air-breathing flight, the principle of the engine is that a rotating knocking wave tissue which is circumferentially propagated is utilized for combustion, but the engine possibly burns in advance under a higher Mach number (Ma > 7) and is difficult to continue to utilize the rotating knocking wave tissue for combustion, and in practical application, a shock wave front transmission interference air inlet channel exists, the magnitude and direction of thrust eccentric moment and lateral component force can be periodically changed, namely the engine can possibly generate periodic high-frequency vibration and other problems. The oblique detonation ramjet engine can overcome the difficult problem of combustion wave stationing by propelling, does not need additional ignition, is an engine based on a novel combustion organization mode of oblique detonation waves, is expected to realize the technical breakthrough of air-breathing hypersonic propulsion with the Mach number of 9 and above, but has the defect of higher starting speed (Ma > 8) in practical application.

Therefore, in order to further widen the speed range of air-breathing type flight and make up for the defects of an oblique detonation engine, the invention provides a rocket-based combined cycle engine based on the principles of deflagration and detonation and an application method thereof.

Disclosure of Invention

In view of this, the invention provides a rocket-based combined cycle engine based on the principle of deflagration and detonation and an application method thereof.

In one aspect, the present invention provides a rocket-based combined cycle engine based on detonation and detonation principles, comprising:

the central cone comprises a front body, a middle body and a rear body which are sequentially arranged along a first direction, wherein the front body is connected with the middle body through a first telescopic structure, and the middle body is connected with the rear body through a second telescopic structure; the rear body is sunken towards the front body to form an inner spray pipe and an inner combustion chamber in sequence;

the shell extends along the first direction and is sleeved outside the central cone, the central line of the shell is superposed with the central line of the central cone, a cavity is formed between the interior of the shell and the outer wall of the central cone, and the cavity comprises a mixed pressure type air inlet channel, an outward expansion type mixing section, an outer combustion chamber and an outer spray pipe which are sequentially arranged along the first direction; the diameter of the part of the shell corresponding to the mixed pressure type air inlet channel is continuously increased along the direction that the front body points to the rear body;

the support is a hollow cylinder and corresponds to the position of the outward-expanding mixing section, the support comprises a first support and a second support, the first support extends in the direction perpendicular to the central line of the central cone, one end of the first support is connected with the shell, the other end of the first support is connected with the middle body, and the first support is communicated with an injector; the second bracket is positioned on one side of the first bracket, which is far away from the front body, one end of the second bracket is connected with the shell, and the other end of the second bracket is connected with the middle body; the second bracket at least comprises a first bracket A and a second bracket B;

the fuel supply system is arranged in the central cone and comprises a first pipeline, one end of the first pipeline is communicated with the second bracket A, and the other end of the first pipeline is communicated with a fuel pump; the fuel pump is characterized by also comprising a second pipeline, wherein one end of the second pipeline is communicated with the fuel pump, and the other end of the second pipeline is communicated with a precombustion chamber; one end of the third pipeline is communicated with the second pipeline, and the other end of the third pipeline is communicated with the internal combustion chamber; one end of the fourth pipeline is communicated with the fuel pump, and the other end of the fourth pipeline is communicated with a propellant shunting chamber; one end of the fifth pipeline is communicated with the propellant shunting chamber, and the other end of the fifth pipeline is communicated with the first bracket; the fuel pump also comprises a first transmission shaft, one end of the first transmission shaft is communicated with the fuel pump, and the other end of the first transmission shaft is communicated with a turbine; one end of the sixth pipeline is communicated with the precombustion chamber, and the other end of the sixth pipeline is communicated with the turbine; the turbine is communicated with the hot jet regulation and control chamber; the heat jet flow regulating and controlling device further comprises an eighth pipeline, one end of the eighth pipeline is communicated with the heat jet flow regulating and controlling chamber, and the other end of the eighth pipeline is communicated with a heat jet flow shunting chamber; one end of the ninth pipeline is communicated with the hot jet flow shunting chamber, and the other end of the ninth pipeline is communicated with the rear body; the turbine also comprises a second transmission shaft, one end of the second transmission shaft is communicated with the turbine, and the other end of the second transmission shaft is communicated with an oxidant pump; the system also comprises a tenth pipeline, wherein one end of the tenth pipeline is communicated with the oxidizer pump, and the other end of the tenth pipeline is communicated with the propellant shunting chamber; one end of the eleventh pipeline is communicated with the oxidant pump, and the other end of the eleventh pipeline is communicated with the second bracket B; the device also comprises a twelfth pipeline, one end of the twelfth pipeline is communicated with the oxidant pump, and the other end of the twelfth pipeline is communicated with the precombustion chamber; one end of the thirteenth pipeline is communicated with the twelfth pipeline, and the other end of the thirteenth pipeline is communicated with the thermal jet regulation chamber; and one end of the fourteenth pipeline is communicated with the thirteenth pipeline, and the other end of the fourteenth pipeline is communicated with the internal combustion chamber.

Preferably, the number of the first brackets is multiple, and the first brackets are distributed around the central line array of the central cone;

the fifth pipelines correspond to the first supports one by one.

Preferably, the rear body includes a first portion and a second portion, the first portion is a cylinder, one end of the first portion is connected to the second telescopic structure, the other end of the first portion is connected to the second portion, the second portion includes a first circular table and a second circular table, the first circular table includes a first surface and a second surface which are oppositely arranged, the second circular table includes a third surface and a fourth surface, the second surface is connected to the third surface, and the diameter of the second surface is equal to the diameter of the third surface; the first face is connected with the first portion, and the diameter of the first face is equal to that of the first portion; along the first direction, the diameter of the first circular truncated cone is continuously increased, and the diameter of the second circular truncated cone is continuously reduced; the fourth surface is sunken towards the first surface to form the inner spray pipe and the internal combustion chamber in sequence, a section is taken through the central line of the central cone, and the inner spray pipe and the internal combustion chamber are hourglass-shaped on the section.

Preferably, the ninth pipeline comprises a ninth pipeline A and a ninth pipeline B, one end of the ninth pipeline A is communicated with the hot jet flow splitting chamber, and the other end of the ninth pipeline A is communicated with the side surface of the first part; and one end of the ninth pipeline B is communicated with the hot jet flow shunting chamber, and the other end of the ninth pipeline B is communicated with the side surface of the first round platform in the second part.

Preferably, the number of the ninth pipeline a and the ninth pipeline b is multiple, the multiple ninth pipelines a are distributed around the centerline array of the central cone, and the multiple ninth pipelines b are distributed around the centerline array of the central cone.

Preferably, the first telescopic structure comprises a first sleeve body and a second sleeve body, the first sleeve body is arranged in the second sleeve body, and the first sleeve body and the second sleeve body are connected in a sliding manner along the first direction; the first sleeve body is connected with the front body, and the second sleeve body is connected with the middle body;

the second telescopic structure comprises a third sleeve body and a fourth sleeve body, the third sleeve body is sleeved outside the fourth sleeve body, and the third sleeve body and the fourth sleeve body are connected in a sliding mode along the first direction; the third sleeve body is connected with the middle body, and the fourth sleeve body is connected with the rear body;

the middle body is provided with a power supply which is respectively electrically connected with the first telescopic structure and the second telescopic structure and used for controlling the first sleeve body and the fourth sleeve body to move along the first direction.

In another aspect, the present invention provides a method for using a rocket-based combined cycle engine based on detonation and detonation principles as described above, comprising:

the first telescopic structure is in a contraction state, so that the mixed pressure type air inlet channel is matched with the flight working condition of Mach number of 0-2.5;

combusting part of fuel and oxidant discharged from the fuel pump and the oxidant pump in the pre-combustion chamber to generate high-temperature and high-pressure combustion gas to drive the turbine to rotate so as to drive the fuel pump and the oxidant pump to work;

delivering fuel and an oxidant to the internal combustion chamber, wherein fuel gas generated by combustion of the fuel and the oxidant in the internal combustion chamber is accelerated to be sprayed out through the internal spray pipe to provide main thrust;

the rich combustion hot jet flowing out of the turbine is in the hot jet regulation and control chamber and an oxidant react and heat up, sequentially passes through the hot jet flow splitting chamber and the ninth pipeline B to be jetted into the outer combustion chamber, forms an acute angle with the incoming flow direction, introduces external air into the outer combustion chamber through the combined action of stamping and injection and suction, and conducts secondary combustion with rich combustion hot jet flow tissues, and fuel gas generated by the outer combustion chamber is sprayed out through the outer spray pipe in an accelerated mode to provide auxiliary thrust.

The invention provides another application method of the rocket-based combined cycle engine based on the principles of detonation and detonation, which comprises the following steps:

the first telescopic structure moves along the direction that the rear body points to the front body, so that the mixed pressure type air inlet channel is matched with the flight working condition of Mach number 2.5-9;

conveying part of fuel to the propellant shunting chamber, sequentially passing through the fifth pipeline and the first bracket, spraying the part of fuel into the outward-expanding mixing section by the injector, utilizing shock waves generated by the precursor to compress air to flow, and fully mixing the air with the fuel; the rich-combustion hot jet flowing out of the turbine reacts with an oxidant in the hot jet regulation chamber to raise the temperature, sequentially passes through the hot jet flow splitting chamber and the ninth pipeline A to be injected into the external combustion chamber, is at a right angle with the incoming flow direction, and is used as an ignition source and a flame stabilizing source of the premixed incoming flow, so that the premixed gas and the rich-combustion hot jet are fully combusted in the external combustion chamber in a deflagration mode, and the generated gas is accelerated to be ejected through the external nozzle pipe to provide main thrust;

and delivering the other part of the fuel and the oxidant to the internal combustion chamber, wherein the fuel gas generated by the combustion of the fuel and the oxidant in the internal combustion chamber is accelerated and sprayed out through the internal spray pipe to provide auxiliary thrust.

the first telescopic structure moves along the direction that the rear body points to the front body, so that the mixed pressure type air inlet channel is matched with the flight working condition of Mach number 9-15;

conveying part of fuel to the propellant shunting chamber, sequentially passing through the fifth pipeline and the first bracket, spraying the part of fuel into the outward-expanding mixing section by the injector, utilizing shock waves generated by the precursor to compress air to flow, and fully mixing the air with the fuel; the rich-combustion hot jet flowing out of the turbine reacts with an oxidant in the hot jet regulation chamber to increase the temperature, then sequentially passes through the hot jet flow splitting chamber and the ninth pipeline B to be jetted into the external combustion chamber, an acute angle is formed between the rich-combustion hot jet and the incoming flow direction, the oblique detonation wave tissue premixed gas generated on the side surface of the first circular truncated cone is rapidly combusted, the rich-combustion hot jet emitted after the oblique detonation wave is taken as a flame stabilizing source, the premixed gas and the rich-combustion hot jet are fully combusted in the external combustion chamber in a detonation combustion tissue mode, and the generated gas is accelerated and ejected through the external nozzle to provide main thrust;

the first telescopic structure is in a fully extended state, so that the mixed pressure type air inlet channel is matched with a flight working condition with the Mach number higher than 15;

delivering fuel and part of oxidant to the internal combustion chamber, wherein fuel gas generated by combustion of the fuel and the oxidant in the internal combustion chamber is accelerated to be sprayed out through the internal spray pipe to provide main thrust;

conveying the other part of the oxidant to the propellant shunting chamber, sequentially passing through the fifth pipeline and the first support, and spraying the oxidant into the outward-expanding mixing section by the injector; follow the rich combustion heat efflux that the turbine flows out passes through in proper order hot efflux diverging chamber ninth pipeline first jets into outer combustion chamber is the right angle with the incoming flow direction organize oxidant and rich combustion heat efflux postcombustion in the outer combustion chamber, the gas warp of production outer spray tube is spout with higher speed, provides auxiliary thrust.

Compared with the prior art, the rocket-based combined cycle engine based on the principles of deflagration and detonation and the application method thereof provided by the invention at least realize the following beneficial effects:

1. the invention provides a rocket-based combined cycle engine based on deflagration and detonation principles and an application method thereof.A rear body is sunken towards a front body to form an inner spray pipe and an inner combustion chamber in sequence; the liquid rocket engine and the ram engine based on detonation and detonation principles are organically combined, the working requirements of low-speed starting, wide-speed range/airspace range flight and ultra-supersonic cruise (Ma > 9) can be met, the characteristics of large specific impulse, strong robustness, high circulating heat efficiency and compact layout are achieved, the respective advantages of the liquid rocket engine, ram detonation and ram oblique propulsion detonation are fully exerted, and the aim is to meet the multi-mode combined propulsion requirements of future horizontal take-off, single-stage orbit entry or two-stage orbit entry, reusable aerospace vehicle high-efficiency air-breathing type flight in the atmosphere and cross-atmosphere rocket orbit entry flight.

2. According to the rocket-based combined cycle engine based on the principle of deflagration and detonation and the application method thereof, the front body is connected with the middle body through the first telescopic structure, the middle body is connected with the rear body through the second telescopic structure, and the first telescopic structure enables the front body to reciprocate along the first direction, so that the mixed pressure type air inlet channel is adjusted; the second telescopic structure enables the rear body to reciprocate along the first direction, and the position of the side face of the first circular truncated cone relative to the shell and the structure of the outer spray pipe are adjusted, so that the requirements of organizing combustion and accelerating airflow under each mode are met.

3. In the rocket-based combined cycle engine based on the principle of deflagration and detonation and the application method thereof, one end of a first transmission shaft is communicated with a fuel pump, and the other end of the first transmission shaft is communicated with a turbine; one end of the sixth pipeline is communicated with the precombustion chamber, and the other end of the sixth pipeline is communicated with the turbine; one end of the second transmission shaft is communicated with the turbine, and the other end of the second transmission shaft is communicated with the oxidant pump; the high-pressure fuel gas rich in combustion generated by the pre-combustion chamber provides power to drive the turbine to rotate, so that the turbine drives the fuel pump and the oxidant pump to deliver fuel and oxidant. The rich-combustion hot jet flowing out of the turbine is deeply regulated and controlled in the hot jet regulation and control chamber, so that the rich-combustion hot jet can play the roles of an ignition source and a flame stabilizing source in the outer combustion chamber, and the combustion sufficiency of the outer combustion chamber and the thrust stability of the outer spray pipe are enhanced.

Of course, it is not necessary for any product in which the present invention is practiced to achieve all of the above-described technical effects simultaneously.

Other features of the present invention and advantages thereof will become apparent from the following detailed description of exemplary embodiments thereof, which proceeds with reference to the accompanying drawings.

Drawings

The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description, serve to explain the principles of the invention.

FIG. 1 is a schematic diagram of a rocket-based combined cycle engine based on detonation and detonation principles provided by the present invention;

FIG. 2 isbase:Sub>A cross-sectional view taken along line A-A' of FIG. 1;

FIG. 3 is an enlarged view at D of FIG. 2;

FIG. 4 is another cross-sectional view taken along line A-A' of FIG. 1;

FIG. 5 isbase:Sub>A further sectional view taken along line A-A' of FIG. 1;

FIG. 6 is a schematic diagram of a fully closed mixed-pressure inlet;

FIG. 7 is a cross-sectional view taken along line B-B' of FIG. 1;

FIG. 8 is a cross-sectional view taken along line C-C' of FIG. 1;

FIG. 9 is a schematic flow diagram of a method of using a rocket-based combined cycle engine based on detonation and detonation principles provided in accordance with the present invention;

FIG. 10 is a schematic flow diagram of a method of using a rocket-based combined cycle engine based on detonation and detonation principles provided in accordance with the present invention;

FIG. 11 is another schematic flow diagram of a method of using a rocket-based combined cycle engine based on detonation and detonation principles provided in accordance with the present invention;

FIG. 12 is a schematic flow diagram of another method of application of the rocket-based combined cycle engine based on detonation and detonation principles provided by the present invention;

1-central cone, 2-front body, 3-middle body, 4-rear body, 5-first telescopic structure, 6-second telescopic structure, 7-inner nozzle, 8-inner combustion chamber, 9-shell, 10-mixing pressure type air inlet channel, 11-outer expanding type mixing section, 12-outer combustion chamber, 13-outer nozzle, 14-bracket, 15-first bracket, 16-second bracket, 17-injector, 18-second bracket A, 19-second bracket B, 20-fuel supply system, 21-first pipeline, 22-fuel pump, 23-second pipeline, 24-precombustion chamber, 25-third pipeline, 26-fourth pipeline, 27-propellant flow splitting chamber, 28-fifth pipeline, 29-first transmission shaft, 30-turbine, 31-sixth pipeline, 32-seventh pipeline, 33-thermal jet regulation chamber, 34-eighth pipeline, 35-thermal jet diversion chamber, 36-ninth pipeline, 37-second transmission shaft, 38-oxidant pump, 39-tenth pipeline, 40-eleventh pipeline, 41-twelfth pipeline, 42-thirteenth pipeline, 43-fourteenth pipeline, 44-first part, 45-second part, 46-first circular truncated cone, 47-second circular truncated cone, 48-ninth pipeline A, 49-ninth pipeline B, 50-first sleeve, 51-second sleeve, 52-third sleeve, 53-fourth sleeve, 54-power supply and X-first direction.

Detailed Description

Various exemplary embodiments of the present invention will now be described in detail with reference to the accompanying drawings. It should be noted that: the relative arrangement of the components and steps, the numerical expressions and numerical values set forth in these embodiments do not limit the scope of the present invention unless specifically stated otherwise.

The following description of at least one exemplary embodiment is merely illustrative in nature and is in no way intended to limit the invention, its application, or uses.

Techniques, methods, and apparatus known to those of ordinary skill in the relevant art may not be discussed in detail but are intended to be part of the specification where appropriate.

In all examples shown and discussed herein, any particular value should be construed as merely illustrative, and not limiting. Thus, other examples of the exemplary embodiments may have different values.

It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, further discussion thereof is not required in subsequent figures.

With reference to fig. 1 to 8, fig. 1 is a schematic structural diagram of a rocket-based combined cycle engine based on the principle of detonation and detonation according to the present invention; fig. 2 isbase:Sub>A cross-sectional view taken alongbase:Sub>A-base:Sub>A' of fig. 1, fig. 2 illustrating an ejector rocket mode with Ma = 0-2.5; FIG. 3 is an enlarged view at D of FIG. 2;

fig. 4 is another cross-sectional view taken along linebase:Sub>A-base:Sub>A' of fig. 1, fig. 4 illustratingbase:Sub>A ram detonation mode with Ma = 2.5-8; fig. 5 isbase:Sub>A further cross-sectional view taken along linebase:Sub>A-base:Sub>A' of fig. 1, fig. 5 illustratingbase:Sub>A ramjet oblique knock mode with Ma = 8-15; FIG. 6 is a schematic diagram of a fully closed mixed-compression air inlet, and FIG. 6 illustrates a pure rocket mode suitable for Ma >15 in-orbit or on-orbit flight; FIG. 7 is a cross-sectional view taken along line B-B' of FIG. 1; FIG. 8 is a cross-sectional view taken along line C-C' of FIG. 1, illustrating a particular embodiment of a rocket-based combined cycle engine based on detonation and detonation principles provided by the present invention, including: the central cone 1, the housing 9, the support 14 and the fuel supply system 20;

the central cone 1 comprises a front body 2, a middle body 3 and a rear body 4 which are sequentially arranged along a first direction X, the front body 2 is connected with the middle body 3 through a first telescopic structure 5, and the middle body 3 is connected with the rear body 4 through a second telescopic structure 6; the rear body 4 is recessed towards the front body 2 to form an inner spray pipe 7 and an inner combustion chamber 8 in sequence;

the internal combustion chamber 8 is a combustion chamber of LRE (Liquid Rocket Engine), and the internal nozzle 7 accelerates the ejection of high-temperature and high-pressure gas generated in the internal combustion chamber 8 to generate thrust.

The shell 9 extends in the first direction X and is sleeved outside the central cone 1, the central line of the shell 9 is overlapped with the central line of the central cone 1, a cavity is formed between the interior of the shell 9 and the outer wall of the central cone 1, and the cavity comprises a mixed pressure type air inlet 10, an outward expansion type mixing section 11, an outer combustion chamber 12 and an outer spray pipe 13 which are sequentially arranged in the first direction X; the diameter of the part of the shell 9 corresponding to the mixed pressure type air inlet 10 is continuously increased along the direction of the front body 2 pointing to the rear body 4;

wherein, the front end of the front body 2 is a wedge surface, and when the front body 2 moves forward relative to the shell 9, the mixed pressure type air inlet channel 10 can be completely closed. The mixed pressure type air inlet 10 mainly performs deceleration and diffusion on incoming flow; the outward-expanding mixing section 11 can avoid the formation of thermal congestion as much as possible; the external combustion chamber 12 is a combustion chamber for secondary combustion of rich combustion heat jet, ramjet combustion and oblique combustion; the outer nozzle 13 is mainly used for accelerating the ejection of high-temperature and high-pressure gas generated in the outer combustion chamber 12 to generate thrust. In addition, the outer nozzle 13 can also help the incompletely expanded fuel gas ejected from the inner nozzle 7 to be further expanded under the condition of low pressure so as to improve the thrust and specific impulse.

The liquid rocket engine, the stamping detonation and the stamping oblique detonation propulsion have the characteristics of large specific impulse, strong robustness, high cycle thermal efficiency and compact layout, fully play the respective advantages of the liquid rocket engine, the stamping detonation and the stamping oblique detonation propulsion, and aim to meet the multi-mode combined propulsion requirements of future horizontal take-off and landing, single-stage or two-stage orbit entry, reusable aerospace vehicle for efficient air-breathing type flight in the atmosphere and cross-atmosphere rocket orbit entry flight.

The support 14 is a hollow cylinder and corresponds to the position of the outward-expanding mixing section 11, the support 14 comprises a first support 15 and a second support 16, the first support 15 extends in the direction perpendicular to the central line of the central cone 1, one end of the first support 15 is connected with the shell 9, the other end of the first support is connected with the middle body 3, and the first support 15 is communicated with an injector 17; the second bracket 16 is positioned on one side of the first bracket 15 far away from the front body 2, one end of the second bracket 16 is connected with the shell 9, and the other end is connected with the middle body 3; the second bracket 16 at least comprises a first bracket A18 and a second bracket B19;

wherein the bracket 14 may connect the central cone 1 and the housing 9 as one piece. Because the support 14 is a hollow cylinder, the first support 15 provides a channel, so that the fuel, the oxidant and the like in the propellant shunting chamber 27 are sprayed into the outward-expanding mixing section 11; the second support 16 provides a passage for the fuel and the oxidizer of the external tank to enter the central body 3, the fuel being delivered by the first support 18 and the oxidizer by the second support B19; the oxidant may also be delivered through the first stent 18 and the fuel through the second stent 19. The number of the brackets 14 can be reasonably adjusted according to parameters such as the diameter of an engine, the flow rate of propellant and the like.

The fuel supply system 20 is arranged inside the central cone 1 and comprises a first pipeline 21, one end of the first pipeline 21 is communicated with the second bracket A18, and the other end of the first pipeline 21 is communicated with a fuel pump 22; a second pipeline 23 is further included, one end of the second pipeline 23 is communicated with the fuel pump 22, and the other end of the second pipeline 23 is communicated with a precombustion chamber 24; the device also comprises a third pipeline 25, one end of the third pipeline 25 is communicated with the second pipeline 23, and the other end is communicated with the internal combustion chamber 8; a fourth pipeline 26, wherein one end of the fourth pipeline 26 is communicated with the fuel pump 22, and the other end of the fourth pipeline 26 is communicated with a propellant shunting chamber 27; the device also comprises a fifth pipeline 28, one end of the fifth pipeline 28 is communicated with the propellant shunting chamber 27, and the other end of the fifth pipeline is communicated with the first bracket 15; the device also comprises a first transmission shaft 29, one end of the first transmission shaft 29 is communicated with the fuel pump 22, and the other end is communicated with a turbine 30; a sixth pipeline 31 is further included, one end of the sixth pipeline 31 is communicated with the precombustion chamber 24, and the other end of the sixth pipeline 31 is communicated with the turbine 30; the device also comprises a seventh pipeline 32, one end of the seventh pipeline 32 is communicated with the turbine 30, and the other end of the seventh pipeline is communicated with a thermal jet regulation and control chamber 33; the device also comprises an eighth pipeline 34, one end of the eighth pipeline 34 is communicated with the hot jet regulation and control chamber 33, and the other end is communicated with a hot jet shunting chamber 35; the device also comprises a ninth pipeline 36, one end of the ninth pipeline 36 is communicated with the hot jet flow shunting chamber 35, and the other end is communicated with the rear body 4; the device also comprises a second transmission shaft 37, one end of the second transmission shaft 37 is communicated with the turbine 30, and the other end of the second transmission shaft 37 is communicated with an oxidant pump 38; a tenth pipeline 39 is further included, one end of the tenth pipeline 39 is communicated with the oxidizer pump 38, and the other end of the tenth pipeline 39 is communicated with the propellant shunting chamber 27; the device also comprises an eleventh pipeline 40, wherein one end of the eleventh pipeline 40 is communicated with the oxidant pump 38, and the other end of the eleventh pipeline 40 is communicated with the second bracket B19; a twelfth pipeline 41 is further included, one end of the twelfth pipeline 41 is communicated with the oxidant pump 38, and the other end of the twelfth pipeline 41 is communicated with the precombustion chamber 24; the device also comprises a thirteenth pipeline 42, one end of the thirteenth pipeline 42 is communicated with the twelfth pipeline 41, and the other end of the thirteenth pipeline 42 is communicated with the thermal jet regulation and control chamber 33; and the device also comprises a fourteenth pipeline 43, one end of the fourteenth pipeline 43 is communicated with the thirteenth pipeline 42, and the other end of the fourteenth pipeline 43 is communicated with the internal combustion chamber 8. And a fuel supply system 20 is arranged for adjustment, so that the conversion of an ejection rocket mode, a stamping detonation mode, a stamping oblique detonation mode and a pure rocket mode is realized.

In some alternative embodiments, the number of the first brackets 15 is plural, and the plural first brackets 15 are distributed around the central line array of the central cone 1; the fifth pipes 28 correspond one-to-one to the first racks 15.

The plurality of first brackets 15 are distributed in an array around the center line of the central cone 1, and fig. 7 only shows that the number of the first brackets 15 is 4, specifically, the number of the first brackets 15 and the angle between two adjacent first brackets 15 may be set according to actual conditions, and of course, the angle between two adjacent first brackets 15 in the plurality of first brackets 15 may be set to be different from the angle between two other adjacent first brackets 15.

In some optional embodiments, the rear body 4 includes a first portion 44 and a second portion 45, the first portion 44 is a cylinder, one end of the first portion is connected to the second telescopic structure 6, the other end of the first portion is connected to the second portion 45, the second portion 45 includes a first circular truncated cone 46 and a second circular truncated cone 47, the first circular truncated cone 46 includes a first surface and a second surface which are oppositely arranged, the second circular truncated cone 47 includes a third surface and a fourth surface, the second surface is connected to the third surface, and the diameter of the second surface is equal to the diameter of the third surface; the first face is connected to the first portion 44, the first face being of the same diameter as the first portion 44; in the first direction X, the diameter of the first truncated cone 46 is continuously increased, and the diameter of the second truncated cone 47 is continuously decreased; the fourth surface is sunken towards the first surface to form an inner spray pipe 7 and an inner combustion chamber 8 in sequence, a section is taken through the central line of the central cone 1, and the inner spray pipe 7 and the inner combustion chamber 8 are hourglass-shaped in section.

It can be understood that the second face is connected to the third face, the diameter of the second face is equal to the diameter of the third face, which can ensure the fluency of the rear body 4, and the side of the first circular truncated cone 46 is an axisymmetric wedge surface, which helps to induce oblique detonation waves in the outer combustion chamber 12 to organize combustion.

In some optional embodiments, the ninth conduit 36 includes a ninth conduit a 48 and a ninth conduit b 49, the ninth conduit a 48 communicates with the thermal jet diversion chamber 35 at one end and communicates with the side of the first portion 44 at the other end; and one end of the ninth pipeline B49 is communicated with the hot jet flow splitting chamber 35, and the other end of the ninth pipeline B is communicated with the side surface of the first round platform 46 in the second part 45.

It will be appreciated that the provision of ninth conduits a 48 and b 49 connecting different positions of the rear body 4 helps to achieve different modal adjustments.

In some optional embodiments, the number of the ninth conduits a 48 and the ninth conduits b 49 is plural, the plural ninth conduits a 48 are distributed around the centerline array of the central cone 1, and the plural ninth conduits b 49 are distributed around the centerline array of the central cone 1. The ninth pipelines a 48 and the ninth pipelines b 49 are distributed around the central line of the central cone 1 in an array manner, and can be distributed uniformly, and of course, the number and arrangement of the ninth pipelines a 48 and the ninth pipelines b 49 can be set according to actual requirements.

In some alternative embodiments, the first telescopic structure 5 comprises a first sheath 50 and a second sheath 51, the first sheath 50 is disposed inside the second sheath 51, and the first sheath 50 and the second sheath 51 are slidably connected along the first direction X; the first jacket body 50 is connected with the precursor body 2, and the second jacket body 51 is connected with the middle body 3;

the second telescopic structure 6 comprises a third sleeve 52 and a fourth sleeve 53, the third sleeve 52 is sleeved outside the fourth sleeve 53, and the third sleeve 52 and the fourth sleeve 53 are connected in a sliding manner along the first direction X; the third sleeve body 52 is connected with the middle body 3, and the fourth sleeve body 53 is connected with the rear body 4;

the middle body 3 is provided with a power supply 54, and the power supply 54 is electrically connected with the first telescopic structure 5 and the second telescopic structure 6 respectively and is used for controlling the first sleeve 50 and the fourth sleeve 53 to move along the first direction X.

It can be understood that the first telescopic structure 5 makes the precursor 2 move back and forth along the first direction X, so as to realize the adjustment of the mixed pressure type air inlet 10; the second telescopic structure 6 makes the afterbody 4 reciprocate along the first direction X, and realizes the adjustment of the structure of the external combustion chamber 12, including the position of the wedge surface of the first circular truncated cone 46 relative to the shell and the switching of the types of the external spray pipes 13, wherein the types of the external spray pipes 13 include a contraction and expansion type and an expansion type.

In some alternative embodiments, the first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth, thirteenth, and

fourteenth conduits

21, 23, 25, 26, 28, 31, 39, 40, 41, 42, 43 are each provided with a flow regulator to facilitate detection and regulation of the fuel supply system 20.

With continuing reference to fig. 1, 2, 3, 4 and 9, fig. 9 is a flow chart of a method for applying a rocket-based combined cycle engine based on detonation and detonation principles provided by the present invention, illustrating a method for applying a rocket-based combined cycle engine based on detonation and detonation principles provided by the present invention as described in any of the above embodiments, comprising: a rocket ejection mode when Ma = 0-2.5;

s101: the first telescopic structure 5 is in a contracted state, so that the mixed pressure type air inlet 10 is matched with the flight working condition of Mach number 0-2.5 to keep a high-efficiency air compression state;

s102: combusting part of the fuel and oxidant discharged from the fuel pump 22 and the oxidant pump 38 in the pre-combustion chamber 24 to generate high-temperature and high-pressure gas to drive the turbine 30 to rotate so as to drive the fuel pump 22 and the oxidant pump 38 to work;

s103: fuel and oxidant are conveyed to an inner combustion chamber 8, and fuel gas generated by combustion of the fuel and the oxidant in the inner combustion chamber 8 is accelerated to be sprayed out through an inner spray pipe 7 to provide main thrust;

s104: the rich combustion hot jet flowing out of the turbine 30 reacts with an oxidant in the hot jet regulation chamber 33 to increase the temperature, sequentially passes through the hot jet flow splitting chamber 35 and the ninth pipeline B49 to be jetted into the outer combustion chamber 12, forms an acute angle with the incoming flow direction, introduces outside air into the outer combustion chamber 12 through the combined action of stamping and injection suction, and performs secondary combustion with the rich combustion hot jet flow tissue, and gas generated by the outer combustion chamber 12 is accelerated to be jetted out through the outer jet pipe 13 to provide auxiliary thrust.

With continuing reference to fig. 1, 2, 3, 4 and 10, fig. 10 is a schematic flow chart of a method for applying a rocket-based combined cycle engine based on detonation and detonation principles provided by the present invention, illustrating a method for applying a rocket-based combined cycle engine based on detonation and detonation principles provided by the present invention as described in any of the above embodiments, comprising: a punch knock mode when Ma =2.5 to 8;

s201: the first telescopic structure 5 moves along the direction that the rear body 4 points to the front body 2, so that the mixed pressure type air inlet 10 is matched with the flight working condition of Mach number 2.5-9, and the high-efficiency air compression state is kept;

s202: combusting part of the fuel and oxidant discharged from the fuel pump 22 and the oxidant pump 38 in the pre-combustion chamber 24 to generate high-temperature and high-pressure gas to drive the turbine 30 to rotate so as to drive the fuel pump 22 and the oxidant pump 38 to work;

s203: part of the fuel is conveyed to a propellant shunting chamber 27, passes through a fifth pipeline 28 and a first bracket 15 in sequence, is sprayed into the outward-expanding mixing section 11 by an injector 17, flows by utilizing shock wave compressed air generated by the precursor 2, and is fully mixed with the fuel; the rich-combustion hot jet flowing out of the turbine 30 reacts with an oxidant in the hot jet regulation chamber 33 to increase the temperature, sequentially passes through the hot jet flow splitting chamber 35 and the ninth pipeline A48 to be jetted into the external combustion chamber 12, is at a right angle with the incoming flow direction, and is used as an ignition source and a flame stabilizing source of the premixed incoming flow, so that the premixed gas and the rich-combustion hot jet are fully combusted in the external combustion chamber 12 in a deflagration mode, and the generated gas is accelerated to be jetted out through the external jet pipe 13 to provide main thrust;

s204: and delivering the other part of the fuel and the oxidant to the internal combustion chamber 8, accelerating the ejection of fuel gas generated by the combustion of the fuel and the oxidant in the internal combustion chamber 8 through the internal spray pipe 7 to provide auxiliary thrust, and adjusting the thrust by the LRE internal combustion chamber 8 to stabilize the total thrust of the combined engine.

With continuing reference to figures 1, 2, 3, 4, 5 and 11, figure 11 is another flow diagram of a method of using a rocket-based combined cycle engine based on detonation and detonation principles provided by the present invention, illustrating a method of using a rocket-based combined cycle engine based on detonation and detonation principles as described in any one of the above embodiments provided by the present invention, comprising: a stamping oblique knock mode when Ma = 8-15;

s301: the first telescopic structure 5 moves along the direction that the rear body 4 points to the front body 2, so that the mixed pressure type air inlet channel is matched with the flight working condition of Mach number 9-15, and a high-efficiency air compression state is kept;

s302: combusting part of the fuel and oxidant discharged from the fuel pump 22 and the oxidant pump 38 in the pre-combustion chamber 24 to generate high-temperature and high-pressure gas to drive the turbine 30 to rotate so as to drive the fuel pump 22 and the oxidant pump 38 to work;

s303: part of the fuel is conveyed to a propellant shunting chamber 27, passes through a fifth pipeline 28 and a first bracket 15 in sequence, is sprayed into the outward-expanding mixing section 11 by an injector 17, flows by utilizing shock wave compressed air generated by the precursor 2, and is fully mixed with the fuel; the rich-combustion hot jet flowing out of the turbine 30 reacts with an oxidant in the hot jet regulation chamber 33 to increase the temperature, the rich-combustion hot jet sequentially passes through the hot jet flow splitting chamber 35 and the ninth pipeline B49 to be injected into the external combustion chamber 12 and forms an acute angle with the incoming flow direction, the mixed gas of the oblique detonation wave tissue generated on the side surface of the first circular truncated cone 46 is rapidly combusted, the rich-combustion hot jet emitted after the oblique detonation wave is used as a flame stabilizing source, the mixed gas and the rich-combustion hot jet are sufficiently combusted in the external combustion chamber 12 in the organization mode of detonation combustion, and the generated gas is accelerated and ejected through the external nozzle 13 to provide main thrust;

s304: and delivering the other part of the fuel and the oxidant to the internal combustion chamber 8, and accelerating the ejection of the fuel gas generated by the combustion of the fuel and the oxidant in the internal combustion chamber 8 through the internal nozzle 7 to provide auxiliary thrust.

With continuing reference to figures 1, 2, 3, 4, 6 and 12, figure 12 is another flow diagram of a method of application of the present invention to a rocket-based combined cycle engine based on detonation and detonation principles, illustrating a method of application of the present invention to a rocket-based combined cycle engine based on detonation and detonation principles as described in any of the above embodiments, including pure rocket modes suitable for Ma >15 in-orbit or in-orbit flight;

s401: the first telescopic structure 5 is in a fully extended state, so that the mixed pressure type air inlet channel 10 is matched with a flight working condition with the Mach number higher than 15, the flight Mach number is larger than 15, and the mixed pressure type air inlet channel 10 is difficult to effectively compress incoming air and is in a closed state;

s402: combusting part of the fuel and oxidant discharged from the fuel pump 22 and the oxidant pump 38 in the pre-combustion chamber 24 to generate high-temperature and high-pressure gas to drive the turbine 30 to rotate so as to drive the fuel pump 22 and the oxidant pump 38 to work;

s403: the fuel and part of the oxidant are delivered to the internal combustion chamber 8, and fuel gas generated by combustion of the fuel and the oxidant in the internal combustion chamber 8 is accelerated to be sprayed out through the internal spray pipe 7 to provide main thrust.

S404: another part of the oxidant is conveyed to the propellant flow distribution chamber 27, passes through the fifth pipeline 28 and the first bracket 15 in sequence, and is sprayed into the outward expansion type mixing section 11 by the injector 17; the rich-burning hot jet flowing out of the turbine 30 sequentially passes through the hot jet flow splitting chamber 35 and the ninth pipeline shell 48 to enter the outer combustion chamber 12, a right angle is formed with the incoming flow direction, the secondary combustion of the oxidant and the rich-burning hot jet flow is organized in the outer combustion chamber 12, and the generated fuel gas is accelerated to be sprayed out through the outer spray pipe 13 to provide auxiliary thrust.

The conventional closed cycle is to inject the rich gas discharged from the turbine 30 into the internal combustion chamber 8 and to perform the secondary combustion of the fuel and oxidant. According to the invention, firstly, the rich gas discharged by the turbine 30 reacts with the trace oxidant in the thermal jet regulation chamber 33 according to the requirement, releases heat, controllably heats up, and then is injected into the external combustion chamber 12 to be secondarily combusted together with the incoming air flow. The conventional ignition source and flame-holding source of the combined engine are plumes generated by the LRE, while the present invention uses rich-burning high-temperature hot jets formed by controllably raising the temperature of the rich-burning gas discharged from the turbine 30. The scheme is characterized in that the fuel is injected into the outer combustion chamber 12 from different positions in different modes, so that the mode switching of the engine is favorably realized. And in the rocket ejecting mode, the flame is ejected into the outer combustion chamber 12 from the ninth pipeline B49 to serve as an ignition source and a flame stabilizing source, in the oblique detonation mode, the flame is ejected into the outer combustion chamber 12 from the ninth pipeline B49 to serve as a flame stabilizing source, and in the stamping detonation mode and the pure rocket mode, the flame is ejected into the outer combustion chamber 12 from the ninth pipeline A48 to serve as the ignition source and the flame stabilizing source. The invention adopts a structure of double combustion chambers and double spray pipes, and is different from the traditional design that the ejection rocket is arranged in the stamping combustion chamber, the internal combustion chamber 8 of the combined engine designed by the invention is relatively independent from the stamping engine, and the internal combustion chamber and the stamping engine are basically not interfered with each other during the organization combustion. The front body 2 can intelligently move back and forth on the central cone 1 according to the inflow conditions to realize the adjustment of the mixed pressure type air inlet channel 10, and the rear body 4 can also intelligently move back and forth on the central cone 1 according to the requirement of tissue combustion to realize the functions of adjusting the structure of the external combustion chamber 12, adjusting the position of the axisymmetric wedge surface and switching the types of the external spray pipes 13. Different from the traditional method for inducing oblique detonation waves based on two-dimensional wedge surfaces, the oblique detonation wave generating method disclosed by the invention has the advantages that the axial-symmetry conical oblique detonation waves are induced and generated by utilizing the outer side surface of the axial-symmetry circular truncated cone and a rich-combustion thermal jet combination mode in the stamping oblique detonation mode, and the bell-shaped spray pipe is adopted to generate thrust. The combined engine designed by the invention is based on the principle of deflagration and detonation combustion, and adopts different working modes at different flight speed sections respectively so as to improve the combustion and propulsion efficiency.

According to the embodiment, the rocket-based combined cycle engine based on the principles of detonation and the application method thereof, which are provided by the invention, at least the following beneficial effects are realized:

2. According to the rocket-based combined cycle engine based on the principle of deflagration and detonation and the application method thereof, the front body is connected with the middle body through the first telescopic structure, the middle body is connected with the rear body through the second telescopic structure, and the first telescopic structure enables the front body to reciprocate along the first direction, so that the mixed pressure type air inlet channel is adjusted; the second telescopic structure enables the rear body to reciprocate along the first direction, and the position of the side face of the first circular truncated cone relative to the shell and the structure of the outer spray pipe are adjusted, so that the requirements of organizing combustion and accelerating airflow under various modes are met.

Although some specific embodiments of the present invention have been described in detail by way of examples, it should be understood by those skilled in the art that the above examples are for illustrative purposes only and are not intended to limit the scope of the present invention. It will be appreciated by those skilled in the art that modifications may be made to the above embodiments without departing from the scope and spirit of the invention. The scope of the invention is defined by the appended claims.

Claims

1. A rocket-based combined cycle engine based on deflagration and detonation principles, comprising:

the shell extends along the first direction and is sleeved outside the central cone, the center line of the shell is superposed with the center line of the central cone, a cavity is formed inside the shell and the outer wall of the central cone, and the cavity comprises a mixed pressure type air inlet, an outward expansion type mixing section, an outer combustion chamber and an outer spray pipe which are sequentially arranged along the first direction; the diameter of the part of the shell corresponding to the mixed pressure type air inlet channel is continuously increased along the direction that the front body points to the rear body;

the fuel supply system is arranged in the central cone and comprises a first pipeline, one end of the first pipeline is communicated with the second bracket A, and the other end of the first pipeline is communicated with a fuel pump; the fuel pump is characterized by also comprising a second pipeline, wherein one end of the second pipeline is communicated with the fuel pump, and the other end of the second pipeline is communicated with a precombustion chamber; one end of the third pipeline is communicated with the second pipeline, and the other end of the third pipeline is communicated with the internal combustion chamber; one end of the fourth pipeline is communicated with the fuel pump, and the other end of the fourth pipeline is communicated with a propellant shunting chamber; one end of the fifth pipeline is communicated with the propellant shunting chamber, and the other end of the fifth pipeline is communicated with the first bracket; the fuel pump is characterized by also comprising a first transmission shaft, wherein one end of the first transmission shaft is communicated with the fuel pump, and the other end of the first transmission shaft is communicated with a turbine; the device also comprises a sixth pipeline, wherein one end of the sixth pipeline is communicated with the precombustion chamber, and the other end of the sixth pipeline is communicated with the turbine; the turbine is communicated with the hot jet regulation and control chamber; the heat jet flow regulating and controlling device further comprises an eighth pipeline, one end of the eighth pipeline is communicated with the heat jet flow regulating and controlling chamber, and the other end of the eighth pipeline is communicated with a heat jet flow shunting chamber; one end of the ninth pipeline is communicated with the hot jet flow shunting chamber, and the other end of the ninth pipeline is communicated with the rear body; the turbine also comprises a second transmission shaft, one end of the second transmission shaft is communicated with the turbine, and the other end of the second transmission shaft is communicated with an oxidant pump; the system also comprises a tenth pipeline, wherein one end of the tenth pipeline is communicated with the oxidizer pump, and the other end of the tenth pipeline is communicated with the propellant shunting chamber; one end of the eleventh pipeline is communicated with the oxidant pump, and the other end of the eleventh pipeline is communicated with the second bracket B; the device also comprises a twelfth pipeline, one end of the twelfth pipeline is communicated with the oxidant pump, and the other end of the twelfth pipeline is communicated with the precombustion chamber; one end of the thirteenth pipeline is communicated with the twelfth pipeline, and the other end of the thirteenth pipeline is communicated with the thermal jet regulation chamber; and one end of the fourteenth pipeline is communicated with the thirteenth pipeline, and the other end of the fourteenth pipeline is communicated with the internal combustion chamber.

2. A rocket-based combined cycle engine based on the principles of deflagration and detonation as claimed in claim 1, wherein said first brackets are plural in number, said plural first brackets being distributed in an array around the centerline of said central cone;

the fifth pipelines correspond to the first supports one by one.

3. A rocket-based combined cycle engine based on detonation and detonation principles according to claim 1, characterized in that said afterbody comprises a first part and a second part, said first part being a cylinder, one end being connected to said second telescopic structure and the other end being connected to said second part, said second part comprising a first circular truncated cone and a second circular truncated cone, said first circular truncated cone comprising a first face and a second face arranged opposite to each other, said second circular truncated cone comprising a third face and a fourth face, said second face being connected to said third face, the diameter of said second face being equal to the diameter of said third face; the first face is connected with the first portion, and the diameter of the first face is equal to that of the first portion; along the first direction, the diameter of the first circular truncated cone is continuously increased, and the diameter of the second circular truncated cone is continuously reduced; the fourth surface is sunken towards the first surface to form the inner spray pipe and the internal combustion chamber in sequence, a section is taken through the central line of the central cone, and the inner spray pipe and the internal combustion chamber are hourglass-shaped on the section.

4. A rocket-based combined cycle engine based on detonation and detonation principles according to claim 3, characterised in that said ninth conduits comprise a ninth conduit a and a ninth conduit b, said ninth conduit a having one end communicating with said hot jet flow splitting chamber and the other end communicating with the side of said first portion; one end of the ninth pipeline B is communicated with the hot jet flow shunting chamber, and the other end of the ninth pipeline B is communicated with the side surface of the first round table in the second part.

5. A rocket-based combined cycle engine according to claim 4 wherein said ninth conduits A and said ninth conduits B are each plural in number, a plurality of said ninth conduits A are distributed around the centerline array of said central cone, and a plurality of said ninth conduits B are distributed around the centerline array of said central cone.

6. A rocket-based combined cycle engine based on the principles of deflagration and detonation as recited in claim 1, wherein said first telescopic structure comprises a first sleeve body and a second sleeve body, said first sleeve body being disposed within said second sleeve body, and said first sleeve body and said second sleeve body being slidably connected along said first direction; the first sleeve body is connected with the front body, and the second sleeve body is connected with the middle body;

the middle body is provided with a power supply, and the power supply is electrically connected with the first telescopic structure and the second telescopic structure respectively and is used for controlling the first sleeve body and the fourth sleeve body to move along the first direction.

7. A method of using a rocket based combined cycle engine based on deflagration and detonation principles as defined in claim 4, characterized in that it comprises:

follow the rich combustion heat efflux that the turbine flows is in hot efflux regulation and control room and oxidant reaction intensification pass through in proper order hot efflux shunting room the ninth pipeline B jets into outer combustion chamber, is the acute angle with the incoming flow direction, through punching press and injection suction combined action, introduces external air outer combustion chamber organizes the postcombustion with rich combustion heat efflux, the gas warp that outer combustion chamber produced outer spray tube is spout with higher speed, provides auxiliary thrust.

8. A method of using a rocket based combined cycle engine based on detonation and detonation principles, as defined in claim 4, characterized in that it comprises:

9. A method of using a rocket based combined cycle engine based on detonation and detonation principles, as defined in claim 4, characterized in that it comprises:

conveying part of fuel to the propellant shunting chamber, sequentially passing through the fifth pipeline and the first bracket, spraying the part of fuel into the outward-expanding mixing section by the injector, utilizing shock waves generated by the precursor to compress air to flow, and fully mixing the air with the fuel; the rich-combustion hot jet flowing out of the turbine reacts with an oxidant in the hot jet regulation chamber to increase the temperature, passes through the hot jet flow splitting chamber and the ninth pipeline B in sequence, is injected into the external combustion chamber, forms an acute angle with the incoming flow direction, and is rapidly combusted by virtue of oblique detonation wave tissue premixed gas generated on the side surface of the first circular truncated cone;

10. A method of using a rocket based combined cycle engine based on detonation and detonation principles, as defined in claim 4, characterized in that it comprises:

conveying the other part of the oxidant to the propellant shunting chamber, sequentially passing through the fifth pipeline and the first bracket, and spraying the oxidant into the outward-expanding mixing section by the injector; follow the rich combustion heat efflux that the turbine flows out passes through in proper order hot efflux diverging chamber ninth pipeline first jets into outer combustion chamber is the right angle with the incoming flow direction organize oxidant and rich combustion heat efflux postcombustion in the outer combustion chamber, the gas warp of production outer spray tube is spout with higher speed, provides auxiliary thrust.