CN119412245B - Integrated flow path of wide-range rocket-based ramjet engine coupled with oblique detonation engine flow path - Google Patents
Integrated flow path of wide-range rocket-based ramjet engine coupled with oblique detonation engine flow pathInfo
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- CN119412245B CN119412245B CN202411627253.1A CN202411627253A CN119412245B CN 119412245 B CN119412245 B CN 119412245B CN 202411627253 A CN202411627253 A CN 202411627253A CN 119412245 B CN119412245 B CN 119412245B
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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
- 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
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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
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- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Testing Of Engines (AREA)
Abstract
The invention provides an integrated runner coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine runner, and relates to the technical field of aerospace transportation power systems. The flow channel comprises an air inlet channel, a combustion chamber isolation channel, a first-stage concave cavity, a transition channel, a second-stage concave cavity and a tail spraying channel which are sequentially communicated, wherein a mixing support plate is arranged in the combustion chamber isolation channel, spraying holes for spraying fuel into the combustion chamber are formed in the mixing support plate, a rocket support plate is arranged in the first-stage concave cavity, the axial section of the transition channel is rectangular when the Mach number is 0-8, and the axial section of the transition channel is expanded to be trapezoid when the Mach number is greater than 8. According to the invention, different expansion ratios are realized by adjusting the angle of the rear transition section of the first-stage concave cavity, and the transition from stamping to knocking mode is realized under high Mach incoming flow.
Description
Technical Field
The invention relates to the technical field of aerospace transport power systems, in particular to an integrated runner coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine runner.
Background
The aerospace transport aircraft has the characteristics of high flying speed and the like, and the economic value of the aerospace transport aircraft is highly valued by world space researchers. The reusable aerospace transport system is an important way for human beings to realize aerospace flight and is a key for aerospace fusion development. This requires that the aerospace vehicle have the ability to start at zero speed on the ground and to operate efficiently and stably in a wide speed range. This clearly places high demands on the engines of the aerospace vehicle.
The main power scheme of the current wide-range air-to-air transport aircraft is a ramjet engine and a scramjet engine. Because the high-speed incoming flow is required to be decelerated and pressurized to meet the working conditions when the punching and scramjet engine works, the punching and scramjet engine needs to reach a certain flying speed to work normally, and how to realize the zero-speed starting of the punching and scramjet engine becomes a big problem.
The combined power scheme is a relatively mature solution at present, wherein a rocket-based ramjet combined cycle engine organically combines a rocket engine with high thrust-weight ratio and low specific impulse with an air suction ramjet engine with low thrust-weight ratio and high specific impulse, fully exerts the advantages and characteristics of two propulsion modes, successfully solves the problem of zero-speed starting of the scramjet engine through conversion among a pure rocket ejection mode, a punching mode and a scramjet mode, widens the working speed range of the scramjet engine and shows good performance. However, when the working speed is further increased, generally, the mach number is greater than 8, the friction resistance between the inner wall surface of the combustion chamber and the incoming air is greatly increased, and the performance of the combustion chamber is obviously reduced, so that a new power scheme needs to be introduced.
The oblique knocking engine has the potential to widen the operating speed range of the ramjet engine to Ma 8-ma16+. The ODE is characterized in that fuel is injected into a precursor or an air inlet channel in advance, the precursor or the air inlet channel is premixed with air to a certain extent and then enters a combustion chamber, oblique detonation waves which can be resident in high-speed airflow are induced by oblique wedges and the like, mixed gas combustion heat release is completed within a short distance, and a high-temperature high-pressure combustion product is expanded and accelerated through a spray pipe to generate thrust, so that the engine based on supersonic detonation combustion with the thermal cycle efficiency far higher than that of conventional isobaric combustion is realized.
Disclosure of Invention
In order to solve the defects in the background art, the invention mainly solves the problems that when Mach number is larger than 8, the friction resistance between the inner wall surface of the combustion chamber and incoming air can be increased sharply and the performance of the combustion chamber can be obviously reduced because the combustion chamber of the scramjet engine is longer in the prior art. The invention provides an integrated runner coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine runner, which is designed by coupling a rocket-based combined cycle engine (RBCC) and a diagonal detonation engine (ODE) runner, namely the wide-range rocket-based ramjet engine and the diagonal detonation engine runner integrated combined engine, and simultaneously absorbs the advantages of a rocket, a ramjet engine and the diagonal detonation engine, and realizes the ultra-wide speed range (Ma0-Ma8+) operation of the engine through an injection mode, a sub-combustion punching mode, a super-combustion punching mode and a diagonal detonation mode so as to meet the requirements of repeated use and air-sky round trip.
The first object of the invention is to provide an integrated runner for coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine runner, which comprises an air inlet channel, a combustion chamber isolation channel, a first-stage concave cavity, a transition channel, a second-stage concave cavity and a tail spray channel which are communicated in sequence;
the mixing support plate is provided with injection holes for injecting fuel into the combustion chamber;
Rocket support plates are arranged in the first-stage concave cavities;
when the Mach number is 0-8, the axial section of the transition channel is rectangular, and when the Mach number is greater than 8, the axial section of the transition channel expands into a trapezoid.
Preferably, the first-stage concave cavity is used as a flame stabilizer and is simultaneously geometrically coupled with the rocket support plate to provide a geometric condition for generating the oblique detonation wave.
Preferably, the combustion chamber isolation channel is a channel for isolating the combustion chamber from the air inlet channel, and comprises a hollow shell with an air inlet end and an air outlet end, wherein the air inlet end is connected with the air outlet end of the throat section of the device, so that the air inlet channel is prevented from being started due to conduction before back pressure of the combustion chamber.
Preferably, the secondary cavity is a combustion heat release area with Mach number of 0-8, or an afterburning area with fuel and air mixed when Mach number is greater than 8.
Preferably, the front section of the rocket support plate is in a wedge shape, and the tip adopts a round angle to reduce flow resistance;
The front section of the rocket support plate is provided with bulges at the wedge positions, and the bulges are used for promoting the generation of oblique blast waves.
Preferably, the tail injection channel provides space for the fuel gas that is not fully reacted to continue to react, and simultaneously enables the high-temperature and high-pressure fuel gas to fully expand under a certain geometric constraint.
Preferably, the injection holes are arranged at the left and right sides of the middle section of the mixing support plate for injecting fuel into the combustion chamber.
Preferably, when the Mach number is greater than 8, the axial section of the transition channel expands from a rectangular shape to a trapezoid shape, the whole engine runner is converted into a configuration of oblique detonation mode operation, and detonation is completed through oblique splitting at the front edge of the rocket support plate, so that oblique detonation combustion is performed.
Compared with the prior art, the invention has the beneficial effects that:
the invention provides an integrated runner coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine runner, according to the invention, different expansion ratios are realized by adjusting the angle of the rear transition section of the first-stage concave cavity, and the transition from stamping to knocking mode is realized under high Mach incoming flow.
The invention solves the problem of insufficient starting power of the high-altitude overhead transport aircraft at zero speed on the ground, and adopts the liquid rocket to provide strong power on the ground so as to enable the high-altitude overhead transport aircraft to smoothly enter a low Mach stamping mode. Meanwhile, the support plate is highly coupled with the liquid rocket, the inside of the engine is simple, the overall reliability is stronger, the production cost is low, and the economic benefit is high.
The mode overregulation scheme of the invention is simple, convenient and efficient, the transition from stamping to inclined knocking mode can be completed by simply regulating the expansion ratio of the transition section, and the reliability of the combustion chamber is greatly improved.
According to the invention, the rocket-based ramjet engine and the oblique detonation engine are combined together, so that the working speed range is remarkably widened, and the ultra-high speed range adaptability of Ma0-Ma8+ is achieved.
Drawings
FIG. 1 is a cross-sectional view of an integrated flow passage coupling a wide-area rocket-based ramjet engine with a diagonal detonation engine flow passage;
FIG. 2 is a simulation result of injection mode values;
FIG. 3 is a numerical simulation result of a sub-combustion stamping mode;
FIG. 4 is a numerical simulation result of the scramjet mode;
FIG. 5 is a graph of results of numerical simulation of a diagonal knock mode;
Fig. 6 flame factor (Fi) cloud.
Wherein, 1, a combustion chamber inlet, 2, a combustion chamber isolation section, 3a first-stage concave cavity, 4 a transition section, 5a second-stage concave cavity, 6a tail nozzle, and 7, a combustion chamber outlet, 8, a blending support plate, 9, a rocket support plate, 10, a bulge, 11, a fuel injection hole and 12, a rocket outlet.
Detailed Description
Several embodiments of the present invention will be described in detail below, but it should be understood that the scope of the present invention is not limited by the embodiments.
The invention aims to provide an integrated runner coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine runner, different expansion ratios are realized by adjusting the angle of a rear transition section of a first-stage concave cavity, and the transition from the ramjet mode to the detonation mode is realized under high Mach inflow.
In order to achieve the above purpose, the invention provides an integrated runner for coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine runner, which comprises an air inlet channel, a combustion chamber isolation channel, a first-stage concave cavity, a transition channel, a second-stage concave cavity and a tail spray channel which are communicated in sequence;
the mixing support plate is provided with injection holes for injecting fuel into the combustion chamber;
Rocket support plates are arranged in the first-stage concave cavities;
when the Mach number is 0-8, the axial section of the transition channel is rectangular, and when the Mach number is greater than 8, the axial section of the transition channel expands into a trapezoid.
The first-stage concave cavity is used as a flame stabilizer and is simultaneously geometrically coupled with the rocket support plate to provide geometric conditions for generating oblique detonation waves.
The combustion chamber isolation channel is a channel for isolating the combustion chamber from the air inlet channel, and comprises a hollow shell with an air inlet end and an air outlet end, wherein the air inlet end is connected with the air outlet end of the throat section of the equipment, so that the air inlet channel is prevented from being not started due to conduction before back pressure of the combustion chamber.
The secondary concave cavity is a combustion heat release area when the Mach number is 0-8, or is an afterburning area when the Mach number is greater than 8, and fuel and air are further mixed.
The front section of the rocket support plate is in a wedge shape, the tip end of the rocket support plate adopts a round angle to reduce flow resistance, a rocket nozzle is reserved at the rear end of the rocket support plate and used for starting an engine on the ground at 0 speed, and bulges are arranged at the wedge positions of the front section of the rocket support plate and are used for promoting the generation of oblique explosion waves.
The tail spray channel provides a space for the fuel gas which is not fully reacted to continue the reaction, and simultaneously enables the high-temperature and high-pressure fuel gas to fully expand under a certain geometric constraint.
The injection holes are arranged on the left side and the right side of the middle section of the mixing support plate and are used for injecting fuel into the combustion chamber.
When the Mach number is greater than 8, the axial section of the transition channel expands from a rectangular shape to a trapezoid shape, the whole engine runner is converted into a configuration of oblique detonation mode operation, and detonation is completed through oblique splitting at the front edge of the rocket support plate, so that oblique detonation combustion is performed.
The expansion mode in the transition channel is that the rear edge of the first-stage concave cavity is used as a fulcrum, and the transition section is expanded outwards by 15 degrees.
In order to further explain the integrated runner coupling the wide-range rocket-based ramjet engine and the oblique detonation engine runner, the invention is described with reference to the accompanying drawings.
The air inlet channel related in the following embodiment is a combustion chamber inlet, the combustion chamber isolation channel is a combustion chamber isolation section, the transition channel is a transition section, and the tail jet channel is a tail jet pipe.
As described with reference to fig. 1, a configuration of a combustion chamber integrating a wide-range rocket-based ramjet engine and a runner of a detonation engine mainly comprises a combustion chamber inlet 1, a combustion chamber isolation section 2, a first-stage concave cavity 3, a transition section 4, a second-stage concave cavity 5, a tail nozzle 6 and a combustion chamber outlet 7 which are sequentially communicated;
The combustion chamber inlet, the incoming air enters the cross section of the combustion chamber.
The combustion chamber isolation section isolates the combustion chamber from the air inlet channel, is a hollow shell comprising an air inlet end and an air outlet end, and the air inlet end is connected with the air outlet end of the equipment throat section to prevent the air inlet channel from being not started due to conduction before back pressure of the combustion chamber.
The first-stage concave cavity is used as a flame stabilizer and is geometrically coupled with the rocket support plate to provide a geometrical condition for generating the oblique detonation wave.
In the transition section, as the oblique knocking mode can generate severe combustion in a smaller space, the pressure needs to be quickly released after the combustion, and therefore, a larger expansion ratio is needed, and the power requirement of the mode is met by adjusting the angle of the transition section.
The secondary concave cavity is mainly used for burning the heat release area under the low Mach working condition, and the fuel and the air are further mixed with the post-combustion area under the high Mach working condition.
The tail jet pipe provides a space for the fuel gas which is not fully reacted to continue to react, and simultaneously enables the high-temperature and high-pressure fuel gas to fully expand under a certain geometric constraint.
And the high-temperature fuel gas is discharged out of the section of the combustion chamber at the outlet of the combustion chamber.
The mixing support plate is positioned in the isolation section, generates a precombustion shock wave string and injects fuel to mix the fuel with the main stream.
The front section of the rocket support plate is designed to be in a wedge shape, the tip end of the rocket support plate adopts a round angle to reduce flow resistance, the rear end of the rocket support plate is provided with a rocket nozzle for 0-speed starting of the engine ground, and the rocket support plate integrally penetrates through the whole combustion chamber.
And the fuel support plate injection holes are arranged at the left side and the right side of the middle section of the blending support plate and are used for injecting fuel into the combustion chamber.
The concave cavity is a flame stabilizer which is proved to be efficient and reliable by a large number of researches, and the working principle of the flame stabilizer is that the combustion efficiency and the flame stabilizing capability are improved by increasing the backflow area in the combustion chamber. The support plate has a stabilizing effect while changing the flow field in the combustion chamber, has a similar function with the concave cavity, and can effectively improve the combustion efficiency by combining the support plate with the concave cavity. The combination of the two has the advantages of promoting fuel mixing and ignition, enhancing fuel convection, enabling combustion to be more sufficient, being beneficial to shortening the length of the combustion chamber and the like. Therefore, the invention adopts the scheme of combining three flame holders, namely the blending support plate, the rocket support plate and the concave cavity.
The rocket support plate can effectively provide zero-speed starting power for the aircraft through rocket injection, and the rocket support plate is designed at a position in front of the combustion chamber, so that rocket fuel gas can be fully injected and expanded in the injection process, and the efficient utilization of energy is realized. The front section of the rocket support plate is provided with a bulge 10 at the wedge position, and the bulge is used for promoting the generation of oblique blast waves.
Considering that the flow velocity of a flow field in a combustion chamber is high under the working condition of a sub-combustion/super-combustion stamping mode, a two-stage concave cavity is designed, and a backflow area is formed in the combustion chamber by utilizing the entrainment action of the concave cavity on a main flow so as to prolong the residence time of fuel/air mixed gas in the combustion chamber. The expansion ratio required by the combustion chamber of different working modes is different, and the expansion ratio required by the isovolumetric combustion of the inclined knocking mode is larger than that of the sub-combustion/scramjet mode, so that the transition section after the first-stage concave cavity is designed can be subjected to structural transformation in the equal straight section and the trapezoid expansion section so as to meet the requirements of combustion performance under different working modes.
Considering that the formation condition of the oblique knocking is severe, the oblique shock wave needs to be generated by utilizing the oblique wedge under high-speed incoming flow, the rocket support plate and the oblique knocking wedge are coupled, and bulges are designed at the oblique wedge of the support plate so as to promote the generation of the oblique shock wave. And how to inject fuel under high-speed incoming flow is a big problem, as the incoming flow speed above Ma8 is extremely high, the flame stabilizing effect of the concave cavity and the mixing efficiency of fuel/air are greatly reduced. The invention designs the mixing support plate in the isolation section, and finishes mixing of the fuel and the main flow by injecting kerosene fuel at two sides of the mixing support plate, and finishes detonation combustion at the oblique shock wave generated by the support plate rocket, thereby solving the problems of difficult generation of oblique shock wave and difficult mixing of fuel/air.
In order to realize efficient and stable work under ultra-wide-domain incoming flow working conditions, the invention adapts to stable flame and performance requirements of different incoming flow working conditions through multi-mode switching, and the specific implementation scheme is as follows:
under the condition of Ma0-3 incoming flow, the rocket is ejected in a mode. The engine provides power for meeting the ground zero-speed take-off requirement, and the support plate rocket opens the rocket injection at the moment and quickly climbs by utilizing the strong acceleration performance of the rocket.
And 2, ma3-5, a sub-combustion stamping mode. At this time, the rocket is completely closed and works in a pure stamping mode. The pressure of the combustion chamber is matched with the length of the isolation section, and the shock wave string of the isolation section is enough to reduce the incoming flow to subsonic speed, so that the scramjet can be realized.
And 3, ma5-8, a scramjet stamping mode. The scramjet stamping mode is continuously accelerated to Ma5 and is naturally switched to the scramjet stamping mode, the total temperature of the incoming flow is high in the scramjet stamping mode, the chemical reaction scale is small, the reaction rate is high, and therefore scramjet stamping can be achieved.
And the mode of oblique knocking is more than Ma 8. The aircraft accelerates to Ma8, the ramjet engine stops working, at the moment, the transition section expands, the whole engine runner is converted into a configuration suitable for the work of the oblique detonation mode, and detonation is completed through the oblique wedge at the front edge of the support plate rocket, so that oblique detonation combustion is realized.
In order to meet the stability of the engine in the ultra-wide Mach number working combustion, ma2, ma4, ma6 and Ma8 are selected as reference working conditions, the expansion scheme of the transition section is shown in figure 1, and the reference working conditions are shown in tables 1,3, 4 and 5.
The specific working modes of the flow channel provided by the invention in different working modes are further explained through calculation results under four reference working conditions of Ma2, ma4, ma6 and Ma 8.
The Ma2 rocket ejection mode, at this time, the working condition of the inlet of the isolation section is shown in table 1, the rocket outlet (12) behind the rocket support plate (9) is opened and ejects supersonic high-temperature rich gas, the rocket parameters are shown in table 2, it can be seen from fig. 2 that the supersonic high-temperature rich gas and surrounding subsonic velocity inflow air form a reaction mixing layer under the action of shearing force, the non-premixed flame gradually transits to premixing, and the pressure of the combustion chamber rises. The rocket jet flow continues to develop downstream, the rocket high-speed jet flow is subjected to shearing force, the speed is gradually reduced, the incoming flow speed is gradually increased, the speed gradient between the jet flow and the air is reduced, the reaction mixing layer is gradually thickened, the mixing of the jet flow and the air is more complete, and the rich combustion jet flow continuously releases heat. At this time, fuel is injected from the wall surface of the second-stage concave cavity, and kerosene liquid drops are crushed, atomized and evaporated in the flow channel to form gaseous kerosene, and the gaseous kerosene and air are fully mixed and then enter the main flow of the flow channel. The mixed gas directly enters a reaction mixing layer of the rocket jet flow, wherein the environment with high temperature and high pressure provides good reaction conditions for kerosene fuel, and the kerosene can fully release heat. The OH cloud can show that as the reaction mixture layer gradually thickens, the area where the chemical reaction occurs gradually increases, and the heat release in the flow channel is more severe. As can be seen from the temperature cloud chart, the range of the high temperature zone gradually expands along with the downstream flow, which indicates that the thickening of the reaction mixture layer and the addition of kerosene fuel jointly enhance the chemical reaction, and the heat release is more severe. Under the working condition, the flame in the combustion chamber can be stabilized, the engine works normally, and the design requirement is met.
The Ma4 subclinical stamping mode, at which the inlet working conditions of the isolation section are shown in Table 3. In this mode of operation, the rocket (12) will be fully closed, relying on a purely ram mode of operation. At the moment, the pressure of the combustion chamber is matched with the length of the isolation section (2), and the shock wave string of the isolation section (2) is enough to reduce the incoming flow to subsonic speed, so that the scramjet can be realized. The engine heat release starting position is designed inside the second stage cavity (5) because of the risk of flooding which is easy to occur due to the small total pressure of the incoming flow of Ma 4. The numerical simulation result under the working condition is shown in fig. 3, and it can be seen from the temperature cloud chart that the high-temperature area is mainly concentrated behind the rocket support plate (9) and in the second-stage concave cavity (5), which indicates that a low-speed reflux area is generated behind the rocket support plate (9) through the deceleration action of the rocket support plate (9), the flow time scale of the chemical reaction is reduced, the flow time scale can be matched with the chemical reaction time scale, and the fuel can be sucked into the rocket support plate and is subjected to chemical reaction. The low-speed airflow continues to flow downstream of the engine, enters a backflow area of the secondary cavity (5), generates large-range low-speed vortex and releases heat severely. OH groups can represent the heat release position to a certain extent, the highest OH content in the rocket support plate (9) and the secondary cavity (5) can be seen, the heat release mainly occurs in the range, and the actual combustion process accords with the designed heat release interval. The total temperature of Ma4 is calculated to be slightly higher than that of Ma2, and stable combustion can be established only by initial ignition energy. The area of combustion and heat release is behind the rocket support plate and inside the concave cavity, and has more sufficient space for carrying out sufficient chemical reaction, so that the combustion efficiency is higher than that of Ma2. Overall, the design requirements are met.
The Ma6 scramjet mode, when the inlet working condition of the isolation section is shown in table 4. Under the condition that the flying incoming flow is Ma6, as shown in fig. 4, the strength of the shock wave string of the isolation section (2) is insufficient to decelerate the incoming flow to subsonic speed, so that the incoming flow speed in front of the rocket support plate (9) is still kept at supersonic speed, and the density cloud image can obviously show the improvement of the incoming flow density after the normal shock wave. The mixing support plate (8) is sprayed from the side direction, and the mixing support plate (8) and the high-speed incoming flow are mutually disturbed and form a normal shock wave, so that the high-speed incoming flow has great speed loss and can be fully mixed with kerosene liquid drops, and the fuel begins to release heat at the mixing support plate (8). As can be seen from the temperature cloud chart, the total incoming flow temperature of Ma6 is about 1700K, so that the static temperature is increased to the ignition temperature of kerosene after the forward shock is decelerated, and the kerosene can be spontaneously ignited without forced ignition. The OH group active group is an intermediate component of kerosene combustion and can generally represent the place where the reaction occurs. From the OH cloud, it can be seen that the chemical reaction of kerosene, combustion heat release, is mainly in the shear layer of the combustion stream and air inflow, indicating that the shear layer promotes the blending of fuel with air, and that the chemical reaction is severe and combustion heat release is concentrated in this region. Overall, the working mode of the Ma6 scramjet stamping is high in overall temperature, small in chemical reaction scale and high in chemical reaction speed, so that the combustion efficiency is high, and the design requirement can be met.
Ma8 knocking mode, where the isolation zone inlet conditions are shown in Table 5. The operating mode simulation data is shown in fig. 5. And analyzing the flow field details of the rocket support plate (9) and the first-stage concave cavity (3). By observing the cloud chart, the concentration of OH groups at the position 1 is suddenly increased, and two strong oblique shock waves are accompanied, so that the successful initiation of the oblique shock waves generated by the oblique splitting of the front edge of the rocket support plate (9) is illustrated. The high-temperature reaction area exists at the position 2, the OH group concentration at the position is extremely high through observing an OH group cloud chart, and the analysis density chart can show that two strong combustion-supporting shock waves formed by coupling of diffusion shock waves and reflection shock waves exist at the position, the fuel which is not completely reacted further completes combustion at the position, and a large quantity of OH groups are enriched at the position due to the shrinkage of a runner. This shows that under the condition of Ma8, the oblique detonation mode of the combustion chamber can not only successfully detonate, but also enable the heat release position to provide sufficient space for combustion expansion of fuel gas, so that the fuel waste caused by that fuel should be blown out of the combustion chamber after incomplete reaction is reduced. To further analyze the intrinsic mechanism of the knocking modal combustion, a Flame factor (Flame Index) analysis method is combined, which is defined as the inner product of the fuel mass fraction gradient and the oxidant mass fraction gradient, when the Flame factor is positive, the Flame factor is premixed combustion, and when the Flame factor is negative, the Flame factor is diffusion combustion, so as to obtain a Flame factor cloud chart shown in fig. 6. According to analysis, the fuel is subjected to a precombustion shock wave string to cause certain diffusion combustion after being sprayed out of the blending support plate (8), the combustion is not strong according to a temperature cloud picture, more premixed gas formed by air and fuel is preheated, when the premixed gas of air and fuel flows through the oblique of the rocket support plate (9), combustion reaction caused by oblique shock waves rapidly occurs, severe premixed combustion can be found at the oblique shock wave generated by the rocket support plate (9) from observing the flame factor cloud picture, and no diffusion combustion phenomenon exists after the flame factor cloud picture, so that the combustion completion degree of the fuel and the air is quite high at the place, the characteristics of conforming to an oblique detonation combustion mode, and isovolumetric combustion can be completed in a very small shock wave space, and therefore, the integral coupling scheme of the rocket-based ramjet engine and the oblique detonation engine runner can be judged to realize a reliable oblique detonation combustion mode.
In summary, the integrated runner coupling the wide-range rocket-based ramjet engine and the oblique detonation engine runner creatively combines the rocket and the oblique detonation together, realizes the integrated runner design, and ensures that the variable runner structure meets the high-efficiency combustion structure under each mode, thereby supporting the wide-range Ma0-8 < + > flight.
Table 1 Ma2 isolated block inlet reference operating mode table
TABLE 2 rocket exit parameters for Ma2 rocket supports
Table 3 Ma4 isolated block inlet reference condition table
Table 4 Ma6 isolated segment inlet reference condition table
Table 5 Ma8 isolated block inlet reference condition table
Claims (7)
1. The integrated runner of the coupling of the wide-range rocket-based ramjet engine and the oblique detonation engine runner is characterized by comprising an air inlet channel, a combustion chamber isolation channel, a first-stage concave cavity, a transition channel, a second-stage concave cavity and a tail spray channel which are communicated in sequence;
the mixing support plate is provided with injection holes for injecting fuel into the combustion chamber;
Rocket support plates are arranged in the first-stage concave cavities;
When the Mach number is 0-8, the axial section of the transition channel is rectangular, and when the Mach number is greater than 8, the axial section of the transition channel expands into a trapezoid;
the transition channel expands in such a way that the rear edge of the first-stage concave cavity is used as a fulcrum, and the transition section expands outwards by 15 degrees;
the stable flame and performance requirements of different incoming flow working conditions are met through multi-mode switching, and the specific scheme is as follows:
Under the condition of Ma0-3 inflow, the rocket is ejected in a mode, and at the moment, the rocket of the support plate opens the rocket ejection, so that the rocket can quickly climb by utilizing the strong acceleration performance of the rocket;
Ma3-5, a sub-combustion stamping mode, wherein the rocket is completely closed and works in a pure stamping mode;
Ma5-8, scramjet mode;
and (8) in the oblique knocking mode, the ramjet engine stops working, at the moment, the transition section expands, the whole engine runner is converted into a configuration suitable for the oblique knocking mode, and detonation is completed through oblique splitting of the front edge of the support plate rocket, so that oblique knocking combustion is realized.
2. The integrated flow path of claim 1, wherein the primary cavity acts as a flame stabilizer and simultaneously creates geometric coupling with the rocket mount to provide geometric conditions for the creation of a detonation wave.
3. The integrated flow path coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine flow path of claim 1, wherein the combustion chamber isolation path is a path isolating the combustion chamber from the inlet path and comprises a hollow housing having an inlet end and an outlet end, the inlet end being connected to the outlet end of the throat section of the device to prevent the inlet path from being deactivated by conduction prior to back pressure of the combustion chamber.
4. The integrated flow path coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine flow path of claim 1, wherein the secondary cavity is a combustion heat release zone at a mach number of 0-8 or a fuel and air further blending afterburner zone at a mach number greater than 8.
5. The integrated runner of the wide-range rocket-based ramjet engine and the oblique detonation engine runner coupling according to claim 1, wherein the front section of the rocket support plate is in a wedge shape, and the tip adopts a round angle to reduce flow resistance;
The front section of the rocket support plate is provided with bulges at the wedge positions, and the bulges are used for promoting the generation of oblique blast waves.
6. The integrated flow path coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine flow path of claim 1, wherein the tail injection passage provides space for the incompletely reacted fuel gas to continue to react while allowing for sufficient expansion of high temperature, high pressure fuel gas under certain geometric constraints.
7. The integrated flow path coupling a wide-range rocket-based ramjet engine and a diagonal detonation engine flow path of claim 1, wherein the injection holes are disposed on the left and right sides of the middle section of the blending bracket for injecting fuel into the combustion chamber.
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| CN106150757A (en) * | 2016-08-10 | 2016-11-23 | 西北工业大学 | A kind of dual pathways becomes geometry rocket based combined cycle electromotor |
| CN112682219A (en) * | 2020-12-24 | 2021-04-20 | 中国人民解放军国防科技大学 | Wide-speed-range engine based on tail confluence rocket of annular supercharging central body |
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| FR2637017B1 (en) * | 1988-09-28 | 1990-11-30 | Snecma | NOZZLE STRUCTURE FOR TURBO-STATO-FUSEE COMBINED PROPELLER |
| US6857261B2 (en) * | 2003-01-07 | 2005-02-22 | Board Of Regents, The University Of Texas System | Multi-mode pulsed detonation propulsion system |
| FR2921119B1 (en) * | 2007-09-19 | 2010-10-01 | Novac Modoran | STATOREACTOR FUSE |
| KR101616647B1 (en) * | 2014-12-12 | 2016-04-28 | 한국항공우주연구원 | Combined cycle engine for hypersonic having a rectangle section |
| KR101954034B1 (en) * | 2018-10-11 | 2019-03-04 | 국방과학연구소 | Supersonic fuel injection apparatus |
| CN115807716B (en) * | 2022-12-05 | 2024-07-09 | 西安交通大学 | Oblique knocking engine |
| CN115899767B (en) * | 2022-12-08 | 2024-11-05 | 西北工业大学 | A mixing support plate suitable for a turboramjet combination engine |
| CN118189214A (en) * | 2024-02-29 | 2024-06-14 | 西北工业大学 | A flame stabilizing structure of a support plate rocket coupled with a cavity |
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| CN106150757A (en) * | 2016-08-10 | 2016-11-23 | 西北工业大学 | A kind of dual pathways becomes geometry rocket based combined cycle electromotor |
| CN112682219A (en) * | 2020-12-24 | 2021-04-20 | 中国人民解放军国防科技大学 | Wide-speed-range engine based on tail confluence rocket of annular supercharging central body |
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