CN109630315B - Solid rocket scramjet engine, arc-shaped gas generator and central injection device - Google Patents
Solid rocket scramjet engine, arc-shaped gas generator and central injection device Download PDFInfo
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- CN109630315B CN109630315B CN201910136102.9A CN201910136102A CN109630315B CN 109630315 B CN109630315 B CN 109630315B CN 201910136102 A CN201910136102 A CN 201910136102A CN 109630315 B CN109630315 B CN 109630315B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
- F02K7/18—Composite ram-jet/rocket engines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/24—Charging rocket engines with solid propellants; Methods or apparatus specially adapted for working solid propellant charges
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/32—Constructional parts; Details not otherwise provided for
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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
Technical Field
The invention relates to the technical field of solid rocket scramjet engines, in particular to a solid rocket scramjet engine, an arc-shaped gas generator and a central injection device.
Background
The solid rocket scramjet is a power device of a hypersonic aircraft, and the structural design of a grain is one of key technologies of the solid rocket scramjet. In decades of engineering applications of solid rocket engines, a large number of related design techniques have been accumulated.
In the existing solid rocket sub-combustion ramjet charging mode, a solid propellant is mostly assembled at the front side position in a fuel gas generator, the solid propellant is combusted along the axial direction and jets out rich fuel gas, and then the solid propellant is mixed with incoming air in a combustion chamber for combustion. The working Mach number of the solid rocket sub-combustion ramjet is not high and generally not more than 4, and the incoming air is compressed by the air inlet channel and then becomes subsonic, so that the injection of the incoming air can be set as side air inlet. However, the operating mach number of the solid rocket scramjet engine is generally above 5, and in order to ensure the performance of the engine, the incoming air can only enter the afterburning chamber in a supersonic speed mode. Based on this, the air inlet channel is not suitable for being bent excessively, otherwise, a large total pressure loss is generated, and the flowing air entering the afterburning chamber is required to be ensured to flow along the axial direction of the engine. To ensure mixing, the injection of the solid propellant is at an angle to the direction of the incoming air flow. In addition, as the incoming air enters the afterburning chamber at the supersonic speed, the residence time is short, the mixing and burning time is short, generally in millisecond magnitude, so that the length of the afterburning chamber of the solid rocket scramjet engine needs to be increased on the basis of the length of the afterburning chamber of the solid rocket sublagration scramjet engine to ensure the combustion efficiency of the engine, and all solid propellants are assembled on the front side of the afterburning chamber, so that the length-diameter ratio of the whole engine is too large, and the structural layout and the control of an aircraft are not facilitated.
In addition, for the solid rocket scramjet engine, since the solid rocket scramjet engine does not carry liquid working medium and cannot carry out regenerative cooling, the thermal protection of the solid rocket scramjet engine in long-time work is a great challenge. For the existing side wall injection type solid rocket scramjet engine, the gas penetration capability of the engine is limited, and a high-temperature area is concentrated near the wall surface, so that the problem of thermal protection of the engine wall surface is increased.
Disclosure of Invention
The invention provides a solid rocket scramjet engine, an arc-shaped gas generator and a central injection device, which are used for overcoming the defects of unreasonable engine space application, low filling ratio, difficult thermal protection and the like in the prior art, realizing reasonable application of the engine space, improving the filling ratio, reducing the size of the solid rocket scramjet engine, realizing the long-range flight function of the solid rocket scramjet engine, improving the mixing combustion efficiency of the engine and reducing the thermal protection difficulty of the wall surface of the engine.
In order to achieve the purpose, the invention provides a solid rocket scramjet engine which 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; 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 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.
In order to achieve the purpose, the invention also provides an arc-shaped gas generator of the solid rocket scramjet engine, wherein the gas generator is arc-shaped and provided with a containing cavity capable of containing solid propellant, and the gas generator is fixedly assembled on the afterburning chamber or the air inlet channel or the tail nozzle in a whole or in sections 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.
In order to achieve the purpose, the invention also provides a central injection device of the solid rocket scramjet engine, wherein the injection device is arranged at the nozzle end of the draft tube and positioned between the afterburning chamber and the air inlet, and the central axis of the injection device is overlapped with the central axis 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.
Compared with the prior art, the invention has the beneficial effects that:
1. the invention provides a solid rocket scramjet engine which comprises a front body, an air inlet channel, a afterburning chamber, a tail nozzle, at least one fuel gas generator and at least one flow guide pipe, wherein the fuel gas generator is arc-shaped and is provided with a containing cavity capable of containing solid propellant, and the fuel gas generator 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 by utilizing a space gap between an aircraft outer shell and an inner flow channel. In the solid rocket scramjet engine, incoming air enters the afterburning chamber at supersonic speed, the residence time of the incoming air and rich fuel gas in the afterburning chamber is short, so that the mixing and burning time of the incoming air and the rich fuel gas is short, and the length of the afterburning chamber of the solid rocket scramjet engine needs to be correspondingly increased compared with that of a solid rocket sublagration ramjet engine. The solid propellant of the solid rocket scramjet engine is loaded in the accommodating cavity of the fuel gas generator, and the solid propellant does not need to be arranged at the axial front end of the fuel gas generator like the conventional solid rocket sublagration ramjet engine, so that the total length of the engine is reduced, and the problem of overlarge length-diameter ratio of the engine is avoided. The number of the gas generators provided by the invention can be set to be single or multiple according to the amount of the solid propellant to be loaded, and meanwhile, the number and surrounding positions of the gas generators are also set according to the structural layout of the engine, so that the internal space of the aircraft is utilized to the maximum extent, the volume utilization rate is improved, more propellants are convenient to carry, and the requirement of long-range flight is met.
2. According to the central injection device of the solid rocket scramjet engine, the primary fuel-rich gas generated by the gas generator enters the central injection device through the guide pipe, is injected from the nozzle close to the afterburning chamber end of the injection device, and limits the main combustion area to the central area far away from the wall surface of the engine, so that an internal flow field of the engine with 'wind-wrapped fire' is formed, and the heat protection pressure of the engine is effectively reduced. In order to realize further mixing enhancement, the end of the injection device, which is close to the afterburning chamber, is provided with a mixing enhancement structure, so that the mixing of primary rich fuel gas and incoming air of the engine can be effectively enhanced.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the structures shown in the drawings without creative efforts.
FIG. 1 is a schematic view of a solid rocket scramjet engine according to one embodiment;
FIG. 2a is a side view of a central injector apparatus according to one embodiment;
FIG. 2b is a front view of a central injector apparatus according to one embodiment;
FIG. 3 is a schematic view of a solid rocket scramjet engine provided in the second embodiment;
FIG. 4a is a side view of a central injector apparatus according to the second embodiment;
FIG. 4b is a front view of the central injector apparatus provided in the second embodiment;
FIG. 5 is a schematic view of a solid rocket scramjet engine provided in the third embodiment;
FIG. 6a is a side view of a central injector provided in accordance with a third embodiment;
FIG. 6b is a front view of a central injector provided in the third embodiment;
FIG. 7 is a schematic view of a solid rocket scramjet engine according to the fourth embodiment.
The reference numbers illustrate: 1: a solid propellant; 2: a gas generator; 3: a afterburning chamber; 4: an air inlet channel; 5: a tail nozzle; 6: a flow guide pipe; 7: a precursor; 8: a flow guide cone; 9: and (4) blending a reinforcing structure.
The implementation, functional features and advantages of the objects of the present invention will be further explained with reference to the accompanying drawings.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that all the directional indicators (such as up, down, left, right, front, and rear … …) in the embodiment of the present invention are only used to explain the relative position relationship between the components, the movement situation, etc. in a specific posture (as shown in the drawing), and if the specific posture is changed, the directional indicator is changed accordingly.
In addition, the descriptions related to "first", "second", etc. in the present invention are only for descriptive purposes and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "a plurality" means at least two, e.g., two, three, etc., unless specifically limited otherwise.
In the present invention, unless otherwise expressly stated or limited, the terms "connected," "secured," and the like are to be construed broadly, and for example, "secured" may be a fixed connection, a removable connection, or an integral part; the connection can be mechanical connection, electrical connection, physical connection or wireless communication connection; they may be directly connected or indirectly connected through intervening media, or they may be connected internally or in any other suitable relationship, unless expressly stated otherwise. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In addition, the technical solutions in the embodiments of the present invention may be combined with each other, but it must be based on the realization of those skilled in the art, and when the technical solutions are contradictory or cannot be realized, such a combination of technical solutions should not be considered to exist, and is not within the protection scope of the present invention.
The invention provides a solid rocket scramjet engine which comprises a forebody 7, an air inlet 4, a afterburning chamber 3, a tail nozzle 5, at least one fuel gas generator 2 and at least one flow guide pipe 6, wherein central injection is carried out;
the gas generator 2 is arc-shaped and is provided with an accommodating cavity capable of accommodating a solid propellant, and the gas generator 2 is integrally or sectionally fixedly assembled on the afterburning chamber 3 or an air inlet channel 4 or a tail nozzle 5 or arranged in a precursor 7 along the axial direction parallel to the afterburning chamber 3 by utilizing the space gap between the outer shell of the aircraft and the inner flow channel;
the honeycomb duct 6 comprises an injection device arranged at an injection end; the injection device is positioned between the afterburning chamber 3 and the air inlet 4, and the injection device and the afterburning chamber 3 have the same central axis; the end of the injection device close to the air inlet channel is a guide cone 8, and the end close to the afterburning chamber is a mixing enhancement structure 9.
The air inlet channel 4 compresses the supersonic air inflow, the compressed inflow air is still supersonic after entering the afterburning chamber 3, but the static temperature and static pressure of the compressed inflow air are improved and the speed is reduced compared with the uncompressed inflow air. The solid propellant 1 is subjected to unilateral end face combustion in the fuel gas generator 2 to generate high-temperature fuel-rich gas, the fuel-rich gas is injected into the afterburning chamber 3 through a fuel gas generator nozzle flow guide pipe 6 and a central injection device (comprising a flow guide cone 8 and a mixing enhancement structure 9) at subsonic speed or supersonic speed, the fuel-rich gas is mixed and combusted with compressed supersonic-speed incoming flow air, and the high-temperature high-pressure gas formed by combustion is expanded through a tail nozzle 5 to do work to generate thrust.
Preferably, the arc of the gasifier 2 is greater than 0 and equal to or less than 2 pi.
Preferably, the gas generator 2 can be arranged in axial or circumferential sectors; the total cavity volume of the gas generator 2 accommodating cavity is matched with the total volume of the solid propellant 1.
The radian size, the sectional design and the number of the gas generators 2 are set according to the quantity of the solid propellant 1 to be loaded and the structural layout of the engine, so that the internal space of the aircraft is utilized to the maximum extent, and the volume utilization rate of the aircraft is improved as much as possible.
Preferably, the included angle between the generatrix of the guide cone 8 of the guide pipe injection device and the central axis is less than or equal to 45 degrees; the generatrix line of the guide cone 8 can be at least one of linear type or streamline type; the guide cone 8 is arranged for reducing the influence of the central injection device on the whole flow field and is beneficial to the inflow air to enter the afterburning chamber 3, and the generatrix of the guide cone 8 can be designed into a linear type or a streamline type according to the structure of the whole engine, so that the influence of the central injection device on the whole flow field is minimum.
The mixing and reinforcing structure 9 is of a lobe type, and at least one injection hole is formed in the wall surface of the lobe type; the injection hole is in any geometric shape, such as a circle, a square, a rhombus and the like; the number and the size of the injection holes depend on the required injection flow and the mixing effect, and the mixing combustion efficiency is controlled by changing the size, the number and the geometric shape of the injection holes.
Preferably, the bending angle of the nozzle end of the draft tube is more than 0 degree and less than or equal to 90 degrees. The fuel-rich gas can be injected into the afterburning chamber 3 at different angles by designing the bending angle of the nozzle end of the flow guide pipe 6, so that the fuel-rich gas and the incoming air can be better mixed.
The arrangement number of the flow guide pipe 6 is set according to the number and the size of the gas generator 2, so that rich fuel gas generated in the gas generator 2 is ensured to be injected into the afterburning chamber 3 through the flow guide pipe 6 as completely as possible.
The throat diameter of the flow guide pipe 6 is set to control the pressure inside the gas generator 2, so as to control rich fuel gas generated by combustion in the gas generator 2 to enter the afterburning chamber 3 at subsonic or supersonic speed.
The rich fuel gas entering the afterburning chamber 3 has the following functions: 1) the rich-combustion gas is mixed with incoming flow air for combustion, chemical energy is converted into heat energy, the heat energy is further expanded in the tail nozzle 5 for acting, the heat energy is converted into mechanical energy, and thrust is provided for an engine; 2) the rich fuel gas is used as a disturbance source, the mixing of gas phase and particle phase in the fuel gas and supersonic incoming flow air is enhanced, and the combustion is stable; 3) the rich fuel gas is used as a high-temperature ignition source to ignite the combustible gas and the particle phase.
Preferably, the air inlet 4, the afterburning chamber 3 and the tail nozzle 5 are connected in sequence; the draft tube 6 is disposed at an axial end of the gas generator 2.
The invention also provides an arc-shaped fuel gas generator of the solid rocket scramjet engine, wherein the fuel gas generator 2 is arc-shaped and is provided with a containing cavity capable of containing solid propellant, and the fuel gas generator 2 is integrally or sectionally fixedly assembled on the afterburning chamber 3 or the air inlet channel 4 or the tail nozzle 5 along the axial direction parallel to the afterburning chamber 3 by utilizing the space gap between the outer shell of the aircraft and the inner flow channel.
The invention also provides a central injection device of the solid rocket scramjet engine, wherein the injection device is arranged at the nozzle end of the flow guide pipe 6 and positioned between the afterburning chamber 3 and the air inlet 4, and the central axis of the injection device is overlapped with the central axis of the afterburning chamber 3; the end of the injection device close to the air inlet channel is a guide cone 8, and the end close to the afterburning chamber is a mixing enhancement structure 9.
The solid propellant 1 of the solid rocket scramjet engine is loaded in the accommodating cavity of the gas generator 2, and the solid propellant 1 is not required to be arranged at the axial front end of the gas generator 2 like the conventional solid rocket sublagration ramjet engine, so that the total length of the engine is reduced, and the problem of overlarge length-diameter ratio of the engine is avoided.
The number of the gas generators 2 provided by the invention can be set to be single or multiple according to the quantity of the solid propellant 1 to be loaded, and meanwhile, the number and surrounding positions of the gas generators 2 are set according to the structural layout of the engine, so that the internal space of the aircraft is utilized to the maximum extent, the volume utilization rate is improved, more propellants are convenient to carry, and the requirement of long-range flight is met.
In addition, the central injection device provided by the invention enhances the mixing of rich fuel gas and incoming air, and simultaneously limits high-temperature combustion gas in a central area far away from the wall surface, thereby obviously reducing the difficulty of thermal protection of the wall surface of the engine.
Example one
Referring to fig. 1, the present embodiment provides a solid rocket scramjet engine, an arc gas generator, and a central injection device,
the solid rocket scramjet engine comprises a forebody 7, an air inlet 4, a afterburning chamber 3, a tail nozzle 5, a fuel gas generator 2 with radian of 2 pi and fixedly wound on the afterburning chamber 3, and four guide pipes 6 which are arranged on the fuel gas generator 2 in a central symmetry manner and are close to the end of the air inlet 4 in the axial direction;
the gas generator 2 is cylindrical and has a containing cavity capable of containing solid propellant, and the containing cavity is sized to be completely loaded with the required amount of solid propellant 1;
the honeycomb duct 6 comprises an injection device arranged at a nozzle end, and the nozzle end is arranged perpendicular to the flowing direction of incoming flow air; the injection device is positioned between the afterburning chamber 3 and the air inlet 4, and the central axis of the injection device is overlapped with the central axis of the afterburning chamber 3; the end of the injection device close to the air inlet channel is a guide cone 8, and the end close to the afterburning chamber is a mixing enhancement structure 9.
An included angle between a linear bus of the guide cone 8 and the central axis is 30 degrees; the number of the injection holes is 4, and the injection holes are all circular, as shown in fig. 2a and 2 b.
Example two
Referring to fig. 3, the present embodiment provides a solid rocket scramjet engine, an arc gas generator, and a central injection device,
the solid rocket scramjet engine comprises a forebody 7, an air inlet 4, a afterburning chamber 3, a tail nozzle 5 and gas generators 2, wherein three radians of the gas generators 2 are 2 pi and are respectively and fixedly wound on the afterburning chamber 3, the air inlet 4 and the tail nozzle 5, two flow guide pipes 6 are arranged at the end, close to the air inlet 4, of each gas generator 2 in the axial direction, and the two flow guide pipes 6 are distributed on the end face in a central symmetry manner;
the gas generator 2 is cylindrical and is provided with a containing cavity capable of containing solid propellant, and the total volume of the three containing cavities can be completely loaded with the required amount of the solid propellant 1;
the honeycomb duct 6 comprises a jetting device arranged at a nozzle end, the nozzle end of the honeycomb duct 6 fixedly wound on the fuel gas generator 2 on the afterburning chamber 3 is perpendicular to the flow direction of incoming air, the nozzle end of the honeycomb duct 6 fixedly wound on the fuel gas generator 2 on the air inlet passage 4 is arranged to form an included angle of 45 degrees with the flow direction of the incoming air, and the nozzle end of the honeycomb duct 6 fixedly wound on the fuel gas generator 2 on the tail nozzle 5 is arranged to form an included angle of 60 degrees with the flow direction of the incoming air.
The injection device is positioned between the afterburning chamber 3 and the air inlet 4, and the central axis of the injection device is overlapped with the central axis of the afterburning chamber 3; the end of the injection device close to the air inlet channel is a guide cone 8, and the end close to the afterburning chamber is a mixing enhancement structure 9.
An included angle between a linear bus of the guide cone 8 and the central axis is 30 degrees; the generatrix of the guide cone 8 is linear and streamline; the number of the injection holes is 2, and each injection hole is a diamond shape, as shown in fig. 4a and 4 b.
EXAMPLE III
Referring to fig. 5, the present embodiment provides a solid rocket scramjet engine, an arc gas generator, and a central injection device,
the solid rocket scramjet engine comprises a forebody 7, an air inlet 4, a afterburning chamber 3, a tail nozzle 5 and a fuel gas generator 2, wherein the two fuel gas generators are respectively fixedly arranged in the forebody 7 and fixedly wound on the afterburning chamber 3 in a fixed mode, and the radian of each fuel gas generator is 2 pi; the gas generator 2 in the front body 7 is only provided with one guide pipe 6, the guide pipe 6 is arranged at the end of the gas generator 2 close to the air inlet 4, and the nozzle end of the guide pipe is parallel to the flow direction of incoming air, so that rich fuel gas enters the afterburning chamber 3 and then is parallel to the incoming air; the gas generator 2 fixedly wound on the afterburning chamber 3 is provided with four flow guide pipes 6, the four flow guide pipes 6 are centrosymmetrically distributed at the end, close to the air inlet 4, of the gas generator 2, and the nozzle end of the flow guide pipes is vertical to the flow direction of incoming air, so that rich fuel gas is vertical to the flow direction of the incoming air entering the afterburning chamber 3;
the gas generator 2 is cylindrical and is provided with a containing cavity capable of containing solid propellant, and the total volume of the two containing cavities can be completely loaded with the required amount of the solid propellant 1;
the four guide pipes 6 are fixedly arranged on the gas generator 2 which is fixedly wound on the afterburning chamber 3 and comprise an injection device arranged at a nozzle end, and the nozzle end is arranged in a direction vertical to the flowing direction of incoming flow air; the injection device is positioned between the afterburning chamber 3 and the air inlet 4, and the central axis of the injection device is overlapped with the central axis of the afterburning chamber 3; the end of the injection device close to the air inlet channel is a guide cone 8, and the end close to the afterburning chamber is a mixing enhancement structure 9.
An included angle between a streamline bus of the diversion cone 8 and the central axis is 15 degrees; the number of the injection holes is 4, and the injection holes are all circular, as shown in fig. 6a and 6 b.
Example four
Referring to fig. 7, the present embodiment provides a solid rocket scramjet engine, an arc gas generator, and a central injection device,
the solid rocket scramjet engine comprises a forebody 7, an air inlet 4, a afterburning chamber 3, a tail nozzle 5, a fuel gas generator 2 with radian pi and fixedly wound on the afterburning chamber 3, and a flow guide pipe 6 arranged at the end, close to the air inlet 4, of the fuel gas generator 2 in the axial direction;
the gas generator 2 is cylindrical and has a containing cavity capable of containing solid propellant, and the containing cavity is sized to be completely loaded with the required amount of solid propellant 1;
the honeycomb duct 6 comprises an injection device arranged at a nozzle end, and the nozzle end is arranged perpendicular to the flowing direction of incoming flow air; the injection device is positioned between the afterburning chamber 3 and the air inlet 4, and the central axis of the injection device is overlapped with the central axis of the afterburning chamber 3; the end of the injection device close to the air inlet channel is a guide cone 8, and the end close to the afterburning chamber is a mixing enhancement structure 9.
An included angle between a linear bus of the guide cone 8 and the central axis is 30 degrees; the number of the injection holes is 4, and the injection holes are all circular.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention, and all modifications and equivalents of the present invention, which are made by the contents of the present specification and the accompanying drawings, or directly/indirectly applied to other related technical fields, are included in the scope 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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| CN110700963B (en) * | 2019-08-15 | 2021-03-02 | 西北工业大学 | Compact layout type solid rocket gas scramjet engine based on axial symmetry |
| CN110793062A (en) * | 2019-10-30 | 2020-02-14 | 北京空天技术研究所 | Scramjet engine and runner structure adopting central combustion |
| CN112431692B (en) * | 2020-11-17 | 2021-08-03 | 中国人民解放军战略支援部队航天工程大学 | Cooperation air-breathing liquid rocket engine propellant supply system |
| CN112664350A (en) * | 2020-12-15 | 2021-04-16 | 北京动力机械研究所 | Multi-stage charging solid fuel ramjet engine |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3754511A (en) * | 1954-12-30 | 1973-08-28 | Us Navy | Fuel and fuel igniter for ram jet and rocket |
| RU2195566C2 (en) * | 2000-02-21 | 2002-12-27 | Иркутский военный авиационный инженерный институт | Rocket ramjet engine |
| CN106870203A (en) * | 2017-03-30 | 2017-06-20 | 内蒙动力机械研究所 | The scramjet engine of fluidized powder propellant |
| CN107503862A (en) * | 2017-10-10 | 2017-12-22 | 北京航空航天大学 | A kind of hybrid rocket combination circulation propulsion system and its control method |
| CN109098891A (en) * | 2018-10-11 | 2018-12-28 | 中国人民解放军国防科技大学 | Cross-medium ramjet based on solid propulsion |
| CN109139297A (en) * | 2018-07-10 | 2019-01-04 | 西北工业大学 | A kind of device combining enhancing blending for solid-rocket scramjet engine |
| CN109322763A (en) * | 2018-09-19 | 2019-02-12 | 中国人民解放军国防科技大学 | Solid rocket powder scramjet engine |
-
2019
- 2019-02-25 CN CN201910136102.9A patent/CN109630315B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3754511A (en) * | 1954-12-30 | 1973-08-28 | Us Navy | Fuel and fuel igniter for ram jet and rocket |
| RU2195566C2 (en) * | 2000-02-21 | 2002-12-27 | Иркутский военный авиационный инженерный институт | Rocket ramjet engine |
| CN106870203A (en) * | 2017-03-30 | 2017-06-20 | 内蒙动力机械研究所 | The scramjet engine of fluidized powder propellant |
| CN107503862A (en) * | 2017-10-10 | 2017-12-22 | 北京航空航天大学 | A kind of hybrid rocket combination circulation propulsion system and its control method |
| CN109139297A (en) * | 2018-07-10 | 2019-01-04 | 西北工业大学 | A kind of device combining enhancing blending for solid-rocket scramjet engine |
| CN109322763A (en) * | 2018-09-19 | 2019-02-12 | 中国人民解放军国防科技大学 | Solid rocket powder scramjet engine |
| CN109098891A (en) * | 2018-10-11 | 2018-12-28 | 中国人民解放军国防科技大学 | Cross-medium ramjet based on solid propulsion |
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