CN109630315B - Solid rocket scramjet engine, arc-shaped gas generator and central injection device - Google Patents

Abstract

The invention discloses a solid rocket scramjet engine, an arc-shaped gas generator and a central injection device, wherein the engine comprises a forebody, an air inlet channel, a afterburning chamber, a tail nozzle, at least one gas generator and at least one flow guide pipe, wherein the gas generator is arc-shaped and provided with an accommodating cavity capable of accommodating a solid propellant, and the accommodating cavity is integrally or sectionally fixedly assembled on the afterburning chamber or the air inlet channel or the tail nozzle along the axial direction parallel to the afterburning chamber or is arranged in the forebody; and the injection device of the flow guide pipe is positioned between the afterburning chamber and the air inlet, the end of the injection device close to the air inlet is a flow guide cone, and the end of the injection device close to the afterburning chamber is a mixing enhancement structure. Compared with the existing solid rocket secondary combustion and scramjet engines, the solid rocket secondary combustion and scramjet engine provided by the invention has the advantages that the space application is reasonable, the filling ratio is improved, the size of the engine is reduced, and the long-range flight function can be realized; meanwhile, the thermal protection pressure of the engine is effectively reduced.

Description

Solid rocket scramjet, arc gas generator, central injection device

technical field

The invention relates to the technical field of a solid rocket scramjet, in particular to a solid rocket scramjet, an arc-shaped gas generator and a central injection device.

Background technique

The solid rocket scramjet is the power unit of the hypersonic vehicle, and the structure design of the grain is one of the key technologies of the solid rocket scramjet. In the engineering application of solid rocket motors for decades, a large number of related design technologies have been accumulated.

Most of the existing solid rocket sub-combustion ramjet charging methods assemble the solid propellant at the front position inside the gas generator, and the solid propellant burns along the axial direction and injects rich fuel gas, which is then mixed with the supplementary combustion chamber. Combustion with incoming air. Since the working Mach number of the solid rocket sub-combustion ramjet is not high, generally not more than 4, the incoming air will become subsonic after being compressed by the intake port, so the injection of the future air can be set to the side intake. However, the working Mach number of the solid rocket scramjet is generally above 5. In order to ensure the performance of the engine, the incoming air can only enter the supplementary combustion chamber at supersonic speed. Based on this, the intake port should not be bent excessively, otherwise a large total pressure loss will occur. Generally, it is necessary to ensure that the incoming air entering the supplementary combustion chamber flows along the axial direction of the engine. In order to ensure mixing, the injection of solid propellant needs to be at a certain angle to the direction of flow of the incoming air. In addition, since the incoming air enters the after-combustion chamber at supersonic speed, the residence time is short, and the mixing combustion time is short, usually in the order of milliseconds. Therefore, in order to ensure the combustion efficiency of the engine, the length of the after-combustion chamber of the solid rocket scramjet needs to be in the solid state. The basis of the length of the afterburning chamber of the rocket sub-combustion ramjet has been increased. At this time, all solid propellants are assembled on the front side of the afterburning chamber, which will make the aspect ratio of the entire engine too large, which is not conducive to the structural layout and control of the aircraft. .

In addition, for a solid rocket scramjet, since it does not carry a liquid working medium, it cannot perform regenerative cooling, and its thermal protection for long-term work is a big challenge. For the existing sidewall injection solid rocket scramjet, due to its limited gas penetration capability, the high temperature area is concentrated near the wall surface, which increases the problem of thermal protection of the engine wall surface.

SUMMARY OF THE INVENTION

The invention provides a solid rocket scramjet, an arc-shaped gas generator, and a central injection device, which are used to overcome the defects of the prior art such as unreasonable use of engine space, low filling ratio, difficult thermal protection, etc. Reasonable use and increasing loading ratio can reduce the size of the solid rocket scramjet engine, realize its long-range flight function, improve the mixing and combustion efficiency of the engine, and reduce the difficulty of thermal protection on the engine wall.

In order to achieve the above purpose, the present invention proposes a solid rocket scramjet, including a precursor, an air inlet, a supplementary combustion chamber, a tail nozzle, at least one gas generator, and at least one guide tube; the gas generator It is arc-shaped and has an accommodating cavity capable of accommodating solid propellant. The gas generator is fixedly assembled in the after-burning chamber as a whole or in sections along the axial direction parallel to the after-burning chamber by using the space gap between the outer casing and the inner flow channel of the aircraft. Or on the intake or tailpipe, or inside the precursor;

The guide pipe includes an injection device arranged at the nozzle end; the injection device is located between the supplementary combustion chamber and the intake port, and the injection device and the supplementary combustion chamber have the same central axis ; The end near the inlet port of the injection device is a guide cone, and the end near the supplementary combustion chamber is a mixing enhancement structure.

In order to achieve the above purpose, the present invention also proposes an arc-shaped gas generator for a solid rocket scramjet. The space gap of the flow channel is used to fix and assemble the gas generator on the supplementary combustion chamber, the intake port or the tail nozzle as a whole or in sections along the axial direction parallel to the supplementary combustion chamber.

In order to achieve the above purpose, the present invention also proposes a central injection device for a solid rocket scramjet, the injection device is arranged at the nozzle end of the guide pipe, and is located between the supplementary combustion chamber and the air inlet. , and the central axis of the injection device overlaps with the central axis of the supplementary combustion chamber; the end of the injection device near the intake port is a guide cone, and the end near the supplementary combustion chamber is a mixing enhancement structure.

Compared with the prior art, the beneficial effects of the present invention are:

1. The solid rocket scramjet engine provided by the present invention includes a precursor, an air inlet, a supplementary combustion chamber, a tail nozzle, at least one gas generator, and at least one guide tube, and the gas generator is in an arc shape, It has an accommodating cavity capable of accommodating solid propellant, and uses the space gap between the outer casing of the aircraft and the inner flow channel to fix the gas generator to the supplementary combustion chamber or the air intake channel as a whole or in sections along the axial direction parallel to the supplementary combustion chamber. or on the tailpipe. In the solid rocket scramjet, the incoming air enters the supplementary combustion chamber at supersonic speed, and the residence time of the incoming air and the rich gas in the supplementary combustion chamber is short, so the mixing combustion time of the incoming air and the rich gas is short, so Compared with the solid rocket sub-combustion ramjet, the length of the supplementary combustion chamber of the solid rocket scramjet needs to be increased accordingly. The solid rocket scramjet solid propellant of the present invention is loaded in the accommodating cavity of the gas generator, and there is no need to set all the solid propellants in the axial direction of the gas generator like the existing solid rocket sub-combustion ramjet. This design reduces the overall length of the engine and avoids the problem of excessive engine aspect ratio. The number of gas generators provided by the present invention can be set in one or more according to the amount of solid propellant to be loaded, and the number and surrounding position of the gas generators can also be set according to the structure and layout of the engine, so as to maximize the utilization of the aircraft Internal space, improve volume utilization, easy to carry more propellant, and meet the requirements of long-distance flight.

2. In the central injection device of the solid rocket scramjet provided by the present invention, the primary rich fuel gas generated by the gas generator enters the central injection device through the guide pipe, and is ejected from the nozzle near the end of the supplementary combustion chamber of the injection device, The main combustion area is limited to the central area away from the engine wall, forming an internal flow field of the "wind-pack fire" engine, which effectively reduces the thermal protection pressure that the engine avoids. In order to achieve further mixing enhancement, a mixing enhancement structure is installed near the end of the supplementary combustion chamber of the injection device, which can effectively enhance the mixing of the primary rich combustion gas of the engine and the incoming air.

Description of drawings

In order to explain the embodiments of the present invention or the technical solutions in the prior art more clearly, the following briefly introduces the accompanying 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, and for those of ordinary skill in the art, other drawings can also be obtained according to the structures shown in these drawings without creative efforts.

1 is a schematic diagram of a solid rocket scramjet engine provided in Embodiment 1;

Figure 2a is a side view of the central injection device provided in the first embodiment;

Figure 2b is a front view of the center injection device provided in the first embodiment;

3 is a schematic diagram of a solid rocket scramjet engine provided in Embodiment 2;

Figure 4a is a side view of the central injection device provided in the second embodiment;

Fig. 4b is the front view of the center injection device provided by the second embodiment;

5 is a schematic diagram of a solid rocket scramjet engine provided in Embodiment 3;

6a is a side view of the center injection device provided in Embodiment 3;

Figure 6b is a front view of the center injection device provided in the third embodiment;

FIG. 7 is a schematic diagram of the solid rocket scramjet provided in the fourth embodiment.

Description of reference numerals: 1: solid propellant; 2: gas generator; 3: afterburner; 4: intake port; 5: tail nozzle; 6: guide pipe; 7: precursor; 8: guide Cone; 9: Blending enhanced structure.

The realization, functional characteristics and advantages of the present invention will be further described with reference to the accompanying drawings in conjunction with the embodiments.

Detailed ways

The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relationship between various components under a certain posture (as shown in the accompanying drawings). The relative positional relationship, the movement situation, etc., if the specific posture changes, the directional indication also changes accordingly.

In addition, descriptions such as "first", "second", etc. in the present invention are only for descriptive purposes, and should not be construed as indicating or implying their relative importance or implicitly indicating the number of indicated technical features. Thus, a feature delimited with "first", "second" may expressly or implicitly include at least one of that feature. In the description of the present invention, "plurality" means at least two, such as two, three, etc., unless otherwise expressly and specifically defined.

In the present invention, unless otherwise expressly specified and limited, the terms "connected", "fixed" and the like should be understood in a broad sense, for example, "fixed" may be a fixed connection, a detachable connection, or an integrated; It can be a mechanical connection, an electrical connection, a physical connection or a wireless communication connection; it can be a direct connection or an indirect connection through an intermediate medium, and it can be the internal connection of two elements or the interaction between the two elements. unless otherwise expressly qualified. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific situations.

In addition, the technical solutions between the various embodiments of the present invention can be combined with each other, but must be based on the realization by those of ordinary skill in the art. When the combination of technical solutions is contradictory or cannot be realized, it should be considered that the combination of technical solutions does not exist and is not within the scope of protection claimed by the present invention.

The present invention proposes a solid rocket scramjet engine, which includes a precursor 7, an air inlet 4, a supplementary combustion chamber 3, a tail nozzle 5, at least one gas generator 2, and at least one guide tube 6 for central injection. ;

The gas generator 2 is arc-shaped and has an accommodating cavity capable of accommodating the solid propellant. The gas generator 2 is integrated or integrated along the axial direction parallel to the supplementary combustion chamber 3 by using the space gap between the outer casing and the inner flow channel of the aircraft. The segment is fixedly assembled on the supplementary combustion chamber 3 or the intake port 4 or the tail nozzle 5, or placed inside the precursor 7;

The guide pipe 6 includes an injection device arranged at the nozzle end; the injection device is located between the supplementary combustion chamber 3 and the intake port 4, and the injection device is connected to the supplementary combustion chamber. 3. Concentric axis; the end of the injection device near the intake port is the guide cone 8, and the end near the supplementary combustion chamber is the mixing enhancement structure 9.

The air intake 4 compresses the incoming supersonic air, and the compressed incoming air is still supersonic after entering the supplementary combustion chamber 3, but compared with the incoming uncompressed air, the compressed incoming air is still and quiet. The pressure is increased and the speed is decreased. The solid propellant 1 undergoes single-side end face combustion in the gas generator 2 to generate high-temperature rich combustion gas, and the rich combustion gas passes through the gas generator nozzle guide tube 6 and the central injection device (including the guide cone 8 and the mixing enhancement). Structure 9) It is injected into the supplementary combustion chamber 3 at subsonic speed or supersonic speed, and is mixed with the compressed supersonic incoming air for combustion.

Preferably, the radian of the gas generator 2 is greater than 0 and less than or equal to 2π.

Preferably, the gas generator 2 can be arranged in sections along the axial or circumferential direction; the total cavity volume of the accommodating cavity of the gas generator 2 matches the total volume of the solid propellant 1 .

The arc size, segment design and quantity of the gas generator 2 are all set according to the amount of solid propellant 1 to be loaded and the engine structure layout, maximizing the use of the interior space of the aircraft, and improving the volume utilization of the aircraft as much as possible. Rate.

Preferably, the included angle between the generatrix of the diversion cone 8 of the diversion pipe injection device and the central axis is ≤45 degrees; the generatrix shape of the diversion cone 8 can be at least one of a straight line or a streamlined shape; The purpose of setting the guide cone 8 is to reduce the influence of the central injection device on the entire flow field, which is beneficial for the incoming air to enter the supplementary combustion chamber 3, and the busbar shape of the guide cone 8 can be designed according to the entire engine structure. The streamlined shape minimizes the influence of the central injection device on the entire flow field.

The mixing enhancement structure 9 is a lobe type, and at least one injection hole is provided on the wall surface of the lobe type; the injection hole is in any geometric shape, such as a circle, a square, a diamond, etc.; The number and size depend on the required injection flow rate and the blending effect, and the blending combustion efficiency is controlled by changing the size, number and geometry of the injection holes.

Preferably, the bending angle of the spout end of the guide tube is greater than 0 degrees and less than or equal to 90 degrees. By designing the bending angle of the nozzle end of the guide pipe 6, the rich fuel gas can be injected into the supplementary combustion chamber 3 at different angles, so as to facilitate better mixing of the rich fuel gas and the incoming air.

The number of the guide pipes 6 is set according to the number and size of the gas generators 2 to ensure that the rich combustion gas generated in the gas generator 2 is injected into the supplementary combustion chamber through the guide pipes 6 as much as possible. 3.

The size of the throat diameter of the guide pipe 6 can be set to control the pressure inside the gas generator 2, and then control the rich combustion gas generated by the combustion in the gas generator 2 to enter the supplementary combustion chamber 3 at subsonic speed or supersonic speed.

The functions of the rich combustion gas entering the supplementary combustion chamber 3 are as follows: 1) The rich combustion gas is mixed and burned with the incoming air, and the chemical energy is converted into thermal energy, which is further expanded in the tail nozzle 5 to do work, and the thermal energy is converted into mechanical energy. The engine provides thrust; 2) The rich combustion gas is used as a disturbance source to enhance the mixing of the gas phase and particle phase in the gas with the supersonic incoming air to stabilize the combustion; 3) The rich combustion gas is used as a high temperature ignition source to ignite the combustible gas and the particle phase.

Preferably, the intake port 4 , the supplementary combustion chamber 3 and the tail nozzle 5 are connected in sequence; the guide pipe 6 is placed at one end of the gas generator 2 in the axial direction.

The present invention also proposes an arc-shaped gas generator for a solid rocket scramjet. The gas generator 2 is arc-shaped and has an accommodating cavity capable of accommodating the solid propellant. The space between the outer casing and the inner flow channel of the aircraft is utilized. The gas generator 2 is fixedly assembled on the supplementary combustion chamber 3 or the intake port 4 or the tail nozzle 5 as a whole or in sections along the axial direction parallel to the supplementary combustion chamber 3 with the clearance.

The present invention also proposes a central injection device for a solid rocket scramjet, wherein the injection device is arranged at the nozzle end of the guide pipe 6, between the supplementary combustion chamber 3 and the air inlet 4, and The central axis of the injection device overlaps with the central axis of the supplementary combustion chamber 3;

The solid rocket scramjet solid propellant 1 of the present invention is loaded in the accommodating cavity of the gas generator 2, and there is no need to set all the solid propellants 1 in the gas generator like the existing solid rocket scramjet engine. 2 Axial front end, this design reduces the overall length of the engine and avoids the problem of excessive engine aspect ratio.

The number of the gas generators 2 provided by the present invention can be set in one or more according to the amount of the solid propellant 1 to be loaded. At the same time, the number and the surrounding position of the gas generators 2 can also be set according to the structure and layout of the engine, so as to maximize the It can effectively utilize the internal space of the aircraft, improve the volume utilization rate, and facilitate the carrying of more propellants to meet the requirements of long-range flight.

In addition, the central injection device provided by the present invention enhances the mixing of the rich combustion gas and the incoming air, and at the same time confines the high temperature combustion gas to the central area away from the wall surface, which significantly reduces the difficulty of thermal protection of the engine wall surface.

Example 1

Please refer to FIG. 1 , the present embodiment provides a solid rocket scramjet, an arc-shaped gas generator, and a central injection device,

The solid rocket scramjet includes a precursor 7, an air inlet 4, a supplementary combustion chamber 3, a tail nozzle 5, a gas generator 2 with an arc of 2π and is fixed on the supplementary combustion chamber 3, four The guide pipe 6 is arranged in the axial direction of the gas generator 2 near the end of the air inlet 4 in a center-symmetrical manner;

The gas generator 2 is cylindrical, and has a accommodating cavity capable of accommodating the solid propellant, and the size of the accommodating cavity can fully load the required amount of the solid propellant 1;

The guide pipe 6 includes an injection device arranged at the nozzle end, and the nozzle end is arranged perpendicular to the flow direction of the incoming air; the injection device is located between the supplementary combustion chamber 3 and the air inlet 4, And the central axis of the injection device overlaps with the central axis of the supplementary combustion chamber 3 ; the end of the injection device near the intake port is the guide cone 8 , and the end near the supplementary combustion chamber is the mixing enhancement structure 9 .

The angle between the linear generatrix of the guide cone 8 and the central axis is 30 degrees; the injection holes are provided with 4, all of which are circular, as shown in Figures 2a and 2b.

Embodiment 2

Please refer to FIG. 3, the present embodiment provides a solid rocket scramjet, an arc-shaped gas generator, and a central injection device,

The solid rocket scramjet includes a precursor 7, an air inlet 4, a supplementary combustion chamber 3, a tail nozzle 5, and three arcs of 2π and are respectively fixed around the supplementary combustion chamber 3, the air inlet 4, and the tail. The gas generator 2 on the nozzle 5, each gas generator 2 is provided with two guide pipes 6 near the end of the air inlet 4 in the axial direction, and the two guide pipes 6 are centrally symmetrically distributed on the end face;

The gas generator 2 is cylindrical and has an accommodation cavity that can accommodate the solid propellant, and the total volume of the three accommodation cavities can fully load the required amount of the solid propellant 1;

The guide pipe 6 includes an injection device arranged at the nozzle end, and the nozzle end of the guide pipe 6 on the gas generator 2 fixed on the supplementary combustion chamber 3 is arranged perpendicular to the flow direction of the incoming air, and is fixed on the inlet side. The nozzle end of the guide pipe 6 on the gas generator 2 on the air passage 4 is arranged at an angle of 45 degrees with the flow direction of the incoming air, and the guide pipe 6 on the gas generator 2 is fixed around the tail nozzle 5. The nozzle end and the flow direction of the incoming air are arranged at an angle of 60 degrees.

The injection device is located between the supplementary combustion chamber 3 and the intake port 4, and the central axis of the injection device overlaps with the central axis of the supplementary combustion chamber 3; The end of the intake port is a guide cone 8 , and the end near the supplementary combustion chamber is a mixing enhancement structure 9 .

The angle between the linear busbar and the central axis of the diversion cone 8 is 30 degrees; the busbar type of the diversion cone 8 is straight + streamlined; the injection holes are provided with 2, all of which are diamond-shaped, as shown in the figure 4a and 4b.

Embodiment 3

Please refer to FIG. 5, the present embodiment provides a solid rocket scramjet, an arc-shaped gas generator, and a central injection device,

The solid rocket scramjet includes a precursor 7, an air inlet 4, a supplementary combustion chamber 3, a tail nozzle 5, two radians of 2π and are respectively fixedly installed inside the precursor 7 and fixedly surrounded by the supplementary combustion chamber. The gas generator 2 on the 3; the gas generator 2 inside the precursor 7 is only provided with a guide pipe 6, and the guide pipe 6 is arranged at the end of the gas generator 2 near the air inlet 4, and its spout end is parallel to the incoming air The flow direction is such that the rich fuel gas enters the supplementary combustion chamber 3 and is parallel to the incoming air; the gas generator 2 fixed on the supplementary combustion chamber 3 is provided with four guide pipes 6, and the four guide pipes 6 are centrally symmetrical. Distributed at the end of the gas generator 2 near the air inlet 4, the nozzle end is perpendicular to the flow direction of the incoming air, so that the rich-burning gas is perpendicular to the flow direction of the incoming air entering the supplementary combustion chamber 3;

The gas generator 2 is cylindrical and has a accommodating cavity capable of accommodating the solid propellant, and the total volume of the two accommodating cavities can completely load the required amount of the solid propellant 1;

The four guide pipes 6 provided on the gas generator 2 fixed around the supplementary combustion chamber 3 include an injection device arranged at the nozzle end, and the nozzle end is arranged perpendicular to the flow direction of the incoming air; the injection device It is located between the supplementary combustion chamber 3 and the intake port 4, and the central axis of the injection device overlaps with the central axis of the supplementary combustion chamber 3; the end of the injection device near the intake port is the guide. The flow cone 8, near the end of the supplementary combustion chamber, is the mixing enhancement structure 9.

The angle between the streamlined generatrix of the diversion cone 8 and the central axis is 15 degrees; the injection holes are provided with four, all of which are circular, as shown in Figures 6a and 6b.

Embodiment 4

Please refer to FIG. 7, the present embodiment provides a solid rocket scramjet, an arc-shaped gas generator, and a central injection device,

The solid rocket scramjet includes a precursor 7, an air inlet 4, a supplementary combustion chamber 3, a tail nozzle 5, a gas generator 2 with a radian of π and is fixed on the supplementary combustion chamber 3, a set of A guide pipe 6 near the end of the air inlet 4 in the axial direction of the gas generator 2;

The gas generator 2 is cylindrical, and has a accommodating cavity capable of accommodating the solid propellant, and the size of the accommodating cavity can fully load the required amount of the solid propellant 1;

The guide pipe 6 includes an injection device arranged at the nozzle end, and the nozzle end is arranged perpendicular to the flow direction of the incoming air; the injection device is located between the supplementary combustion chamber 3 and the air inlet 4, And the central axis of the injection device overlaps with the central axis of the supplementary combustion chamber 3 ; the end of the injection device near the intake port is the guide cone 8 , and the end near the supplementary combustion chamber is the mixing enhancement structure 9 .

The angle between the linear generatrix of the guide cone 8 and the central axis is 30 degrees; the injection holes are provided with four, all of which are circular.

The above descriptions are only the preferred embodiments of the present invention, and are not intended to limit the scope of the present invention. Under the inventive concept of the present invention, the equivalent structural transformations made by the contents of the description and drawings of the present invention, or the direct/indirect application Other related technical fields are included in the scope of patent protection of the present invention.

Claims (5)

1. A solid rocket scramjet engine comprises a forebody, an air inlet channel, a afterburning chamber, a tail nozzle, at least one fuel gas generator and at least one flow guide pipe, and is characterized in that,

the gas generator is arc-shaped and provided with an accommodating cavity capable of accommodating a solid propellant, and the gas generator is integrally or sectionally fixedly assembled on the afterburning chamber or an air inlet channel or a tail nozzle along the axial direction parallel to the afterburning chamber by utilizing the space gap between the outer shell of the aircraft and the inner flow channel, or is arranged in the front body;

the honeycomb duct comprises an injection device arranged at an injection end; the injection device is positioned between the afterburning chamber and the air inlet, and the injection device and the afterburning chamber have the same central axis; the diameter of the injection device is smaller than that of the afterburning chamber; the end of the injection device close to the air inlet channel is a guide cone, and the end close to the afterburning chamber is a mixing enhancement structure;

the radian of the gas generator is more than 0 and less than or equal to 2 pi;

the gas generator can be arranged along the axial direction or the circumferential direction in a segmented mode;

the afterburning chamber and the tail nozzle are of a structure with a wide opening along the main flow direction;

the solid rocket scramjet engine is applied to a hypersonic aircraft.

2. A solid-rocket scramjet engine as recited in claim 1, wherein the total volume of said gas-generator-housing chamber cavity matches the total volume of solid propellant.

3. A solid rocket scramjet engine as recited in claim 1, wherein an angle between a generatrix of a guide cone of said guide tube injection device and a central axis is not more than 45 degrees; the generatrix of the flow guide cone is at least one of linear type or streamline type;

the mixing and reinforcing structure is of a lobe type, and at least one injection hole is formed in the wall surface of the lobe type.

4. A solid-rocket scramjet engine as recited in claim 3, wherein said nozzle end of said nozzle has a bending angle of greater than 0 degrees and equal to or less than 90 degrees.

5. A solid rocket scramjet engine as recited in any one of claims 1 to 4, wherein said inlet duct, afterburning chamber and tail nozzle are connected in sequence; the guide pipe is arranged at one axial end of the gas generator.

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