CN118188220A - A full flow channel for RBCC engine ejection mode - Google Patents

Abstract

The present invention discloses a full flow passage for the ejection mode of an RBCC engine, which consists of an air inlet, an isolating section, a combustion chamber and a nozzle which are connected in sequence from front to back; a fuel support plate is arranged at the isolating section near the air inlet; an ejection rocket is arranged at the connection between the isolating section and the combustion chamber; an external rocket is arranged at the connection between the combustion chamber and the nozzle; the present invention has a higher ejection ratio and secondary combustion efficiency while taking multi-mode into consideration, thereby improving the thrust and specific impulse of the engine.

Description

A full flow channel for RBCC engine ejection mode

Technical Field

The invention belongs to the field of rocket-based combined cycle power technology, and in particular relates to a full flow channel for an RBCC engine ejection mode.

Background technique

The rocket-based combined cycle (RBCC) engine is an organic combination of a rocket engine and a ramjet engine. The rocket engine has a high thrust-to-weight ratio, but carries a large amount of oxidizer, which makes the specific impulse low; the ramjet engine has a simple structure and can use oxygen in the atmosphere, and has a high specific impulse performance, but the engine performance is too low or even cannot work normally when flying at low speeds. Therefore, by making full use of the advantages of the two propulsion methods, the engine can work in the optimal way at different altitudes and Mach numbers, and has the characteristics of high thrust-to-weight ratio and high specific impulse. In addition, the RBCC engine can take into account multiple working modes and achieve a smooth transition between modes. It has the characteristics of autonomous horizontal take-off and landing and wide-range multi-tasking, and is one of the main driving forces for future aerospace transportation.

The stage from zero speed to inlet start of the RBCC engine is called the ejection mode. The ejection mode is the key link that determines whether the RBCC engine can start at zero speed. It mainly uses the main rocket jet to eject ambient air into the combustion chamber, and organizes secondary combustion in the combustion chamber to obtain thrust gain and improve engine performance. In view of the multi-mode characteristics of the RBCC engine, the ejection mode needs to match the subcombustion/scramsonic mode. However, the larger inlet contraction ratio and nozzle expansion ratio required by the subcombustion/scramsonic mode and the pure expansion type combustion chamber will reduce the performance of the ejection mode and limit the flow path design of the ejection mode. Therefore, the matching principle of the ramjet flow path still needs further research. In addition, the location and method of injecting secondary fuel will affect the amount of air ejected by the rocket, and then affect the thrust and specific impulse performance. Therefore, the appropriate secondary combustion organization strategy is an important issue in the ejection mode design work. In response to these two major problems, a lot of research has been carried out in the industry. Based on a large number of previous engine configurations, there are still many problems: the contact area between the single-stage rocket ejector and the air is small, and the ability to work on the air is weak, resulting in insufficient engine ejection and air suction capacity; while taking into account multiple modes, the inlet duct contraction ratio is not well selected, resulting in large aerodynamic resistance and small ejection ratio; the secondary combustion produces a higher combustion chamber pressure, resulting in pressure forward transmission and serious inlet overflow; the secondary fuel and incoming air are not mixed sufficiently, resulting in low secondary combustion efficiency; the combustion chamber channel area is too small, and the suction effect of the rocket jet is weakened, resulting in a reduction in the amount of ejected air; the fixed structure nozzle configuration cannot take into account the working requirements of each mode.

Summary of the invention

The purpose of the present invention is to provide a full flow channel for the ejection mode of an RBCC engine, which has a higher ejection ratio and secondary combustion efficiency while taking into account multiple modes, thereby improving the thrust and specific impulse of the engine.

The present invention adopts the following technical scheme: a full flow passage for the ejection mode of the RBCC engine, which is composed of an air inlet, an isolation section, a combustion chamber, and a nozzle which are sequentially connected from front to back;

A fuel support plate is provided in the isolation section near the air inlet;

A launching rocket is arranged at the connection point between the isolation section and the combustion chamber;

An external rocket is provided at the connection point between the combustion chamber and the nozzle;

When the incoming flow Mach number is between 0 and 1.5, both the ejector rocket and the external rocket are turned on to start the engine at zero speed. As the incoming flow Mach number increases from 1.5 to 2, the flow rate of the ejector rocket is continuously reduced, and the injection amount of the fuel support fuel is increased, so that the fuel and the incoming air are fully mixed in the isolation section and enter the combustion chamber to be ignited and burned by the ejector rocket.

Furthermore, the injection port of the fuel support plate is located on the isolation section, and its inlet is connected to the external fuel system.

Furthermore, the expansion ratio between the outlet and the inlet of the isolation section is 1.4.

Furthermore, the expansion ratio between the combustion chamber outlet and the inlet is 1.98.

Furthermore, the expansion ratio between the position where the launch rocket is installed on the combustion chamber and the combustion chamber inlet is 1.14.

Furthermore, the air inlet duct consists of a first air inlet duct and a second air inlet duct which are connected in sequence. The first air inlet duct is inclined upward and is arranged close to the aircraft air inlet inlet, and the contraction ratio between the outlet and the inlet is 0.38. The inlet of the second air inlet duct is connected to the outlet of the first air inlet duct, and the contraction ratio between the outlet and the inlet is 0.54.

Furthermore, the nozzle is composed of a fixed shell located on the upper side and a movable sealing plate located on the lower side. The fixed shell is "n"-shaped, and the movable sealing plate is located at the lower opening of the fixed shell and is movably connected to the combustion chamber. The movable sealing plate can move up and down along its movable connection point, thereby changing the outlet area of the nozzle so as to adapt to different incoming flow Mach numbers.

Furthermore, the contraction ratio of the nozzle outlet to the inlet is 0.75 in the ejection mode, and the expansion ratio of the nozzle outlet to the inlet is 1.17 in the subsonic/scramfic mode.

A full flow passage for an RBCC engine in an ejection mode, wherein the engine structure of the full flow passage comprises a connecting plate, an air intake casing, an isolation casing, a combustion casing and a nozzle which are sequentially connected from front to back;

The connecting plate is used to extend into the front body of the aircraft and cooperate with the air intake groove of the aircraft to form an air intake channel, and the rear end of the air intake channel is the air intake duct;

A fuel support plate is installed on the isolation shell, the outlet of the fuel support plate is located in the inner cavity of the isolation shell, and the inlet of the fuel support plate is arranged close to the air inlet;

An ejection rocket is installed on the combustion shell, and the outlet of the ejection rocket is located at the inlet of the combustion shell; an external rocket is installed on the nozzle, and the outlet of the external rocket is located at the inlet of the nozzle.

Furthermore, two fuel support plates are provided and are parallel to each other.

The beneficial effects of the present invention are:

The air intake duct of the present invention is composed of two sections, a first air intake duct and a second air intake duct. The contraction ratio of the first air intake duct outlet to the inlet is 0.38, which can better capture the incoming air and increase the amount of air introduced; the contraction ratio of the second air intake duct is 0.54, which decelerates and pressurizes the incoming air.

The first air inlet of the present invention is arranged obliquely upward, and the upper inlet is longer than the lower inlet, so that more air can be captured and the amount of air introduced can be increased.

The expansion ratio of the isolation section of the present invention is 1.40, which is mainly to prevent the pre-combustion shock wave train from moving forward to damage the intake duct, and effectively prevent the back pressure of the combustion chamber from being transmitted forward to cause intake duct overflow; in addition, the incoming air is further decelerated and pressurized;

The fuel support plate of the present invention disperses and sprays the secondary fuel into the flow channel. Since the engine is in the ejection mode, the total temperature of the incoming air is low and insufficient to ignite the secondary fuel. Therefore, spraying the secondary fuel in the isolation section can make the secondary fuel and the incoming air mix in advance, thereby improving the secondary combustion efficiency and shortening the engine length.

The expansion ratio of the combustion chamber of the present invention is 1.98, which is to match the sub-/scram combustion mode; the combustion chamber length is appropriately lengthened to increase the air flow area, so that the rocket can fully exert the ejection and suction effect;

The present invention arranges an ejector rocket at a combustion chamber expansion ratio of 1.14, the interior of the ejector rocket is connected to the fuel system, the ejector rocket outlet is connected to the combustion chamber inlet, and the high-temperature and high-pressure combustion gas is sprayed into the combustion chamber through the ejector rocket outlet, and shears and entrains with the air in the flow channel, thereby achieving the purpose of ejecting air;

The movable sealing plate of the nozzle of the present invention can swing up and down. In the ejection mode, the movable sealing plate is located in the upper position to form a contraction type nozzle with a contraction ratio of 0.75; in the subcombustion/scramson mode, the movable sealing plate is located in the lower position to form an expansion type nozzle with an expansion ratio of 1.17;

The external rocket outlet of the present invention is connected with the nozzle inlet, so that the high-temperature and high-pressure combustion gas generated by the external rocket enters the nozzle through the external rocket outlet for further expansion, which can generate a large auxiliary thrust, is beneficial to the zero-speed starting of the engine, and improves the ejection mode performance.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG1 is a front view of the entire flow channel of the present invention;

FIG. 2 is a schematic diagram of the engine structure of the full flow channel of the present invention.

Among them: 1. Air inlet; 2. Fuel support plate; 3. Isolation section; 4. Launch rocket; 5. Combustion chamber; 6. External rocket; 7. Nozzle; 8. Connecting plate; 9. Isolation shell; 10. Combustion shell; 11. Air inlet shell.

Detailed ways

The present invention is described in detail below with reference to the accompanying drawings and specific embodiments.

In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside" and the like indicate positions or positional relationships based on the positions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the present invention. In addition, the terms "first" and "second" are only used for descriptive purposes, and cannot be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Therefore, the features defined as "first" and "second" may explicitly or implicitly include one or more of the features. In the description of the present invention, unless otherwise specified, "multiple" means two or more. The "direction" in the present invention is described based on the direction of the present invention when it is in the state of Figure 1.

The present invention discloses a full flow passage for an RBCC engine ejection mode, as shown in FIG1 , which is composed of an air inlet 1, an isolation section 3, a combustion chamber 5 and a nozzle 7 which are sequentially connected from front to back.

A fuel support plate 2 is provided at the isolation section 3 near the air inlet 1; an ejection rocket 4 is provided at the connection point between the isolation section 3 and the combustion chamber 5; an external rocket 6 is provided at the connection point between the combustion chamber 5 and the nozzle 7; when the incoming flow Mach number is 0-1.5, the ejection rocket 4 and the external rocket 6 are both turned on to start the engine at zero speed, and in the process of the incoming flow Mach number continuously increasing from 1.5 to 2, the flow rate of the ejection rocket 4 is continuously reduced, and the injection amount of the fuel of the fuel support plate 2 is increased, so that the fuel and the incoming air are fully mixed in the isolation section 3 and enter the combustion chamber 5 to be ignited and burned by the ejection rocket 4.

The injection port of the fuel support plate 2 is located on the isolation section 3, and its inlet is connected to the external fuel system. The expansion ratio between the outlet and the inlet of the isolation section 3 is 1.4. The expansion ratio between the outlet and the inlet of the combustion chamber 5 is 1.98. The expansion ratio between the position where the launch rocket 4 is installed on the combustion chamber 5 and the inlet of the combustion chamber 5 is 1.14.

The air inlet 1 consists of a first air inlet and a second air inlet connected in sequence. The first air inlet is inclined upward and arranged close to the aircraft air inlet inlet, and the contraction ratio between the outlet and the inlet is 0.38. The inlet of the second air inlet is connected to the outlet of the first air inlet, and the contraction ratio between the outlet and the inlet is 0.54.

The nozzle 7 is composed of a fixed shell at the upper side and a movable sealing plate at the lower side. The fixed shell is in an "n" shape. The movable sealing plate is located at the lower opening of the fixed shell and is movably connected to the combustion chamber 5. The movable sealing plate can move up and down along its movable connection point, thereby changing the outlet area of the nozzle 7 to adapt to different incoming flow Mach numbers. In the ejection mode, the contraction ratio of the nozzle 7 outlet to the inlet is 0.75; in the subcombustion/scramjet mode, the expansion ratio of the nozzle 7 outlet to the inlet is 1.17.

The present invention also discloses a full flow channel for the RBCC engine ejection mode. As shown in FIG. 2 , the engine structure of the full flow channel includes a connecting plate 8, an air intake casing 11, an isolation casing 9, a combustion casing 10 and a nozzle 7 which are sequentially connected from front to back.

The connecting plate 8 is used to extend into the front body of the aircraft and cooperate with the air intake groove of the aircraft to form an air intake channel, and the rear end of the air intake channel is the air intake channel 1; a fuel support plate 2 is installed on the isolation shell 9, and the outlet of the fuel support plate 2 is located in the inner cavity of the isolation shell 9, and its inlet is arranged close to the air intake channel 1.

The combustion shell 10 is provided with an ejection rocket 4, the outlet of which is located at the entrance of the combustion shell 10; the nozzle 7 is provided with an external rocket 6, the outlet of which is located at the entrance of the nozzle 7. Two fuel support plates 2 are provided and are parallel to each other.

The air inlet 1 of the present invention compresses the incoming air to decelerate and increase its pressure. In the ejection mode, a low-pressure environment is formed at the outlet of the air inlet 1 to eject and draw the incoming air into the combustion chamber 5. The isolation section 3 connects the air inlet 1 and the combustion chamber 5, and is mainly responsible for matching the pressures of the two. The combustion chamber 5 is a key component in the engine for organizing secondary combustion and heat release. The nozzle 7 mainly expands and accelerates the high-temperature and high-pressure combustion gas generated in the combustion chamber 5 to obtain thrust, and different working modes have different surface requirements. The ejection rocket 4 in the flow channel is responsible for ejecting the incoming atmosphere to organize secondary combustion and heat release in the ejection mode, and improves the engine's ability to resist back pressure. It plays a role in ignition and flame stabilization in the sub-combustion mode, and is closed or works at a small flow rate in the scramjet mode to act as a duty flame.

In the ejection-suction stage, the incoming flow Mach number is between 0 and 1.5, the ejection rocket 4 and the external rocket 6 are both turned on, and the engine starts at zero speed. In this stage, the suction effect of the ejection rocket 4 plays a dominant role in the ejection air flow; in the ejection-ramjet stage, the ejection rocket 4 flow becomes smaller, and the external rocket 6 is turned on. In this process, the ramjet effect plays a dominant role in the air flow entering the combustion chamber 5; in the process of the incoming flow Mach number continuously increasing from 1.5 to 2, the flow of the ejection rocket 4 is continuously reduced, and the injection amount of the fuel of the fuel support plate 2 is increased. Since the air inlet 1 is in an unstarted state, the incoming flow pressure and Mach number are low, and the incoming flow temperature is insufficient to ignite the secondary fuel. In this way, the secondary fuel is injected at the inlet of the isolation section 3, and can be mixed with the incoming air in advance to make the mixing more complete, and then enter the combustion chamber 5, ignite the ejection rocket 4, and the fuel-rich gas and the mixed airflow are fully burned, the combustion efficiency is greatly improved, a higher combustion chamber 5 pressure is established, and a higher thrust gain is obtained.

Example 1

A ground zero-speed ejection test was carried out for the flow channel of the present invention. This embodiment uses a fuel-rich external rocket 6, i.e., a two-stage rocket, to verify the ejection and suction effect of the ejection rocket 4 jet and the over-expanded flow state of the nozzle 7 airflow under the ground zero-speed state.

After verification, the chamber pressure of the external rocket 6 deviates from the design value by 0.4%, and the pressure platform time meets the test requirements. The experiment shows that in the zero-speed takeoff stage, since the external rocket 6 is far away from the air inlet 1, the air injection and suction ability is weak. Therefore, in actual application, the external rocket 6 works at a large flow rate and mainly provides the thrust required for the flight of the aircraft.

The above are only preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims (10)

1. The full flow passage for the ejection mode of the RBCC engine is characterized by comprising an air inlet passage (1), an isolation section (3), a combustion chamber (5) and a spray pipe (7) which are sequentially communicated from front to back;

a fuel support plate (2) is arranged at the position, close to the air inlet channel (1), of the isolation section (3);

A rocket ejector (4) is arranged at the communication part of the isolation section (3) and the combustion chamber (5);

an external rocket (6) is arranged at the communication part of the combustion chamber (5) and the spray pipe (7);

When the Mach number of the incoming flow is 0-1.5, the ejector rocket (4) and the external rocket (6) are opened to enable the engine to start at zero speed, the flow of the ejector rocket (4) is continuously reduced in the process that the Mach number of the incoming flow is continuously increased from 1.5 to 2, the fuel injection quantity of the fuel support plate (2) is increased, and the fuel and the incoming flow air are fully mixed in the isolation section (3) and enter the combustion chamber (5) to be ignited and combusted through the ejector rocket (4).

2. A full flow channel for RBCC engine injection modes according to claim 1, characterized in that the injection port of the fuel strip (2) is located on the separation section (3) with its inlet in communication with the external fuel system.

3. A full flow channel for RBCC engine injection modes according to claim 1, wherein the expansion ratio of the outlet to the inlet of the isolation section (3) is 1.4.

4. A full flow channel for RBCC engine injection modes according to claim 1, characterized in that the expansion ratio of the outlet to the inlet of the combustion chamber (5) is 1.98.

5. A full flow channel for the ejector mode of a RBCC engine according to claim 1, characterized in that the expansion ratio of the position of the combustion chamber (5) where the ejector rocket (4) is mounted to the inlet of the combustion chamber (5) is 1.14.

6. A full flow channel for RBCC engine injection modes according to claim 1, characterized in that the inlet channel (1) consists of a first inlet channel and a second inlet channel in sequential communication, the first inlet channel being arranged obliquely upwards and close to the aircraft inlet with an outlet to inlet constriction ratio of 0.38, the second inlet channel being in communication with the outlet of the first inlet channel with an outlet to inlet constriction ratio of 0.54.

7. The full flow passage for the ejection mode of the RBCC engine according to claim 1, wherein the nozzle (7) is composed of a fixed shell positioned at the upper side and a movable sealing plate positioned at the lower side, the fixed shell is n-shaped, the movable sealing plate is positioned at the lower side opening of the fixed shell and is movably connected with the combustion chamber (5), and the movable sealing plate can move up and down along the movable connecting point of the movable sealing plate so as to change the outlet area of the nozzle (7) and be suitable for different incoming flow mach numbers.

8. A full flow channel for RBCC engine ejection mode according to claim 1, wherein the nozzle (7) outlet to inlet constriction ratio is 0.75 in ejection mode; the expansion ratio of the outlet to the inlet of the lance (7) in the sub-combustion/super-combustion mode is 1.17.

9. The full flow passage for the ejection mode of the RBCC engine is characterized in that the engine structure of the full flow passage comprises a connecting plate (8), an air inlet shell (11), an isolation shell (9), a combustion shell (10) and a spray pipe (7) which are sequentially connected from front to back;

The connecting plate (8) is used for extending into a precursor of the aircraft and is matched with an air inlet groove of the aircraft to form an air inlet channel, and the rear end of the air inlet channel is an air inlet channel (1);

The fuel support plate (2) is arranged on the isolation shell (9), an outlet of the fuel support plate (2) is positioned in an inner cavity of the isolation shell (9), and an inlet of the fuel support plate is close to the air inlet channel (1);

The combustion shell (10) is provided with a rocket ejector (4), and an outlet of the rocket ejector (4) is positioned at an inlet of the combustion shell (10); an external rocket (6) is mounted on the spray pipe (7), and an outlet of the external rocket (6) is positioned at an inlet of the spray pipe (7).

10. A full flow channel for RBCC engine injection modes according to claim 9, characterized in that said fuel support plates (2) are provided in two and mutually parallel.