CN111878253A - Wave-lobe rocket nozzle and rocket base combined circulating propulsion system - Google Patents

Abstract

The invention provides a lobe type rocket nozzle and a rocket base combined circulating propulsion system, which comprise an air flow channel and a rocket gas nozzle, wherein the air flow channel is communicated with an air inlet channel, the number of the rocket gas nozzles is more than 2, each rocket gas nozzle is distributed on the outer side of the air flow channel, and a partition wall between the air flow channel and the rocket gas nozzle is in a lobe structure. The lobe structure has a plurality of convex and concave lobes in series. The structural design correspondingly generates a mixing boundary layer with a plurality of continuous convex lobe and concave lobe shapes in application, the mixing area is increased compared with the traditional near-circular mixing boundary layer, and the mixing efficiency of two flows is enhanced. On the other hand, the continuous convex lobes and concave lobes can generate more longitudinal eddy currents, and the mixing efficiency is improved. And each rocket gas nozzle can be independently opened, closed and adjusted, so that the flow of the rocket can be adjusted, and the aim of adjusting the working condition of the rocket-based combined circulating propulsion system in real time can be achieved.

Description

A lobed rocket nozzle and a rocket-based combined cycle propulsion system

technical field

The invention relates to the technical field of rocket nozzles, in particular to a lobe type rocket nozzle.

Background technique

The rocket-based combined cycle propulsion system (RBCC) utilizes the coupling of the rocket and the ramjet. It mainly has four modules: ejector mode, sub-combustion ramjet mode, scramjet mode and pure rocket mode, among which the ejector mode It is necessary to use the gas generated by the rocket to capture the air in the ejection inlet. The mixing of the rocket gas and the captured air in the inlet is the determining factor of the ejection modal performance. How to improve the ejection mixing efficiency of the rocket gas and the captured air becomes the rocket A key factor in the design of a combined cycle propulsion system.

At present, the research on the arrangement of rocket nozzles is relatively conventional. The invention patent publication No. 109779784A discloses an inner flow channel of an RBCC engine with a rocket front center layout. The overall RBCC configuration adopts an axisymmetric configuration. The rocket is located at the center, and the nozzle The shape of the outlet is a conventional circular outlet with a simple structure but no structure to promote blending. The invention patent publication number 109882886A discloses a RBCC engine combustion chamber with a ramp rocket layout and a design method thereof. The overall RBCC configuration adopts a two-dimensional rectangular configuration, and the rocket nozzle is located on one side of the wall and injected in the form of a jet, The use of rocket jet to achieve combustion chamber ignition and flame stability, the equipment processing difficulty is small, and the flame stability is good, but there is a problem of low mixing efficiency of rocket gas and air.

At present, the research and optimization of rocket nozzle position and configuration for RBCC is still in its infancy. Whether it is an axisymmetric configuration of RBCC or a two-dimensional rectangular configuration, there are problems of low mixing rate and long mixing length, which greatly restrict the introduction of RBCC. The performance of the injection mode is improved.

SUMMARY OF THE INVENTION

In view of the defects existing in the prior art, the present invention proposes a lobe-type rocket nozzle and a rocket-based combined cycle propulsion system.

For realizing the above-mentioned technical purpose, the concrete technical scheme that the present invention adopts is as follows:

A lobe-type rocket nozzle, including an air flow channel communicating with an air intake channel and a rocket gas nozzle, the rocket gas nozzle is provided with more than two, each rocket gas nozzle is distributed on the outside of the air flow channel, and the air flow channel is connected to the rocket gas nozzle. The partitions between the gas nozzles are in a lobed structure.

As a preferred solution of the present invention, the partition wall between the air flow channel and the rocket gas nozzles has continuous inner lobes, the inner lobes include a plurality of convex lobes and concave lobes, and each rocket gas nozzle is distributed on the inner lobes The trailing edge of each concave lobe of the lobe.

As a preferred solution of the present invention, the air flow channel is located in the center, and more than two rocket gas nozzles are arranged on the outside of the air flow channel. Further, more than two rocket gas nozzles are evenly distributed on the outside of the air flow channel.

As a preferred solution of the present invention, the rocket gas nozzles are distributed on the same side of the air flow channel as the air flow channel.

As a preferred solution of the present invention, each rocket gas nozzle is independently controlled, and each can be opened and closed independently. In this way, the real-time adjustment of the gas flow can be realized by controlling the switch of each rocket gas nozzle.

As a preferred solution of the present invention, the convex lobes and the concave lobes are connected in a streamlined shape, and the inner side of the partition wall between the air flow channel and the rocket gas nozzle forms a streamlined curved wall surface. In this way, the wall surface of the streamlined curve can be designed by the method of rotating characteristic line and bidirectional streamline tracing method. The streamlined curve wall has a continuous change of the whole wall, low curvature, no geometrically discontinuous part, less shock waves, and no obvious shock wave structure in the flow field. While controlling the heat load, the fluid energy loss is minimized, and both active cooling and passive thermal protection technologies can be well used.

The present invention provides a rocket-based combined cycle propulsion system having any of the above-mentioned lobe-type rocket nozzles.

The beneficial effects of the present invention are as follows:

The partition wall between the air flow channel and the rocket gas nozzle of the present invention is in a lobe structure, and the lobe structure has a plurality of continuous convex lobes and concave lobes. Such a structural design will correspondingly generate a mixed boundary layer with continuous multiple convex lobes and concave lobes in application. Compared with the traditional near-circular mixed boundary layer, the mixed area of the two fluids increases, and the flow of the two fluids increases. Enhanced mixing efficiency. On the other hand, the continuous multiple convex lobes and concave lobes will generate more longitudinal vortices and improve the mixing efficiency.

Further, each rocket gas nozzle is independently controlled, and each can be opened and closed independently, so that the rocket flow, chamber pressure and other working conditions can be adjusted according to actual conditions, so that the rocket and the ramjet can work better coupled.

Further, the inner side of the partition wall between the air flow channel and the rocket gas nozzle forms a streamlined curved wall surface, which has the characteristics of continuous change and low curvature, which can minimize the loss of fluid energy while controlling the heat load.

The structure of the present invention is not only suitable for circular RBCC inner flow channel, rectangular two-dimensional RBCC inner flow channel can also be realized by nearly elliptical lobe ring arrangement, which can also achieve the purpose of promoting fluid mixing and improving RBCC operation flexibility.

Description of drawings

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

FIG. 1 is a schematic structural diagram of Embodiment 1. FIG.

Fig. 2 is the layout schematic diagram of air flow channel and rocket gas nozzle in embodiment 1;

3 is a partial sectional view of Embodiment 1;

FIG. 4 is a schematic structural diagram of Embodiment 2. FIG.

FIG. 5 is a second structural schematic diagram of Embodiment 2. FIG.

Detailed ways

In order to make the technical solutions and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention, but not to limit the present invention.

Example 1:

1 , 2 and 3 , the lobed rocket nozzle includes an air flow channel 1 communicating with the intake channel and a rocket gas nozzle 2 , and there are four rocket gas nozzles 2 . The four rocket gas nozzles 2 are evenly distributed on the outside of the air flow channel 1, and the partition wall between the air flow channel 1 and the rocket gas nozzle 2 has a lobe structure. Specifically, the partition wall between the air flow channel 1 and the rocket gas nozzle 2 has a continuous inner lobe 3 , and the inner lobe 3 includes a continuous plurality of convex lobes 4 and concave lobes 5 . This embodiment has 4 convex lobes 4 and 4 concave lobes 5 . Each rocket gas nozzle 2 is distributed on the trailing edge of each concave lobe 5 of the inner lobe 3 . The lobed rocket nozzle in this embodiment has an axisymmetric structure as a whole.

Each rocket gas nozzle 2 is independently controlled, and each can be opened and closed independently. In this way, the real-time adjustment of the gas flow can be realized by controlling the switch of each rocket gas nozzle.

The convex lobe 4 and the concave lobe 5 are connected in a streamlined shape, and the inner side of the partition wall between the air flow channel 1 and the rocket gas nozzle 2 forms a streamlined curved wall surface. In this way, the wall surface of the streamlined curve can be designed by the method of rotating characteristic line and bidirectional streamline tracing method. The streamlined curve wall has a continuous change of the whole wall, low curvature, no geometrically discontinuous part, less shock waves, and no obvious shock wave structure in the flow field. While controlling the heat load, the fluid energy loss is minimized, and both active cooling and passive thermal protection technologies can be well used.

The partition wall between the air flow channel and the rocket gas nozzle is a lobe structure. The lobe structure has a continuous plurality of convex lobes and concave lobes, which can generate longitudinal vortices and widen the mixing area. The mixing structure promotes the mixing of the two fluids. On the other hand, each rocket gas nozzle can be independently controlled, which can realize the opening and closing of some nozzles, so as to achieve real-time regulation of rocket gas flow and control of working conditions.

Example 2:

Referring to FIGS. 4 and 5 , the lobed rocket nozzle includes an air flow channel 1 communicating with the air intake channel and a rocket gas nozzle 2 . There are three rocket gas nozzles 2 . The three rocket gas nozzles 2 are evenly arranged on the same side outside the air flow channel, and the section of the air flow channel where the rocket gas nozzles 2 are arranged is elliptical or similar to an ellipse. The partition wall between the air channel 1 and the rocket gas nozzle 2 has a lobe structure. Specifically, the partition wall between the air flow channel 1 and the rocket gas nozzle 2 has a continuous inner lobe 3 , and the inner lobe 3 includes a continuous plurality of convex lobes 4 and concave lobes 5 . This embodiment has 2 convex lobes 4 and 3 concave lobes 5 . Each rocket gas nozzle 2 is distributed on the trailing edge of each concave lobe 4 of the inner lobe 3 . Each rocket gas nozzle 2 is independently controlled, and each can be opened and closed independently. In this way, the real-time adjustment of the gas flow can be realized by controlling the switch of each rocket gas nozzle.

Example 3

A rocket-based combined cycle propulsion system has the lobed rocket nozzle provided in Embodiment 1 or Embodiment 2.

In the actual use process, the rocket gas that works first is ejected through the rocket gas nozzle, resulting in a low-pressure area that ejects the incoming air from the air flow channel, and the two fluids are mixed under the action of curved lobe-shaped folds (lobe structure). At the same time, the convex lobes and the concave lobes will generate longitudinal vortices in the two fluids, which will further promote the mixing of the fluids. The two fluids continue to mix in the mixing section and the combustion chamber to form one fluid.

According to the change of the flight height and speed of the rocket-based combined cycle propulsion system, the total pressure of the incoming air and the required thrust will also change in real time. Through the independent opening and closing of each rocket gas nozzle, the real-time adjustment of the rocket flow can be realized. Broaden the operational flexibility of rocket-based combined cycle propulsion systems. In the sub-combustion/super-combustion mode, it is also possible to control the opening of some rocket nozzles to achieve the effect of boosting the rocket.

In order to control the heat load to the maximum extent, the flow channel partition wall with convex lobes and concave lobes in the present invention adopts a curved wall surface, and adopts the spiral characteristic line method and bidirectional streamline tracing to design the flow channel profile. The whole wall surface changes continuously, with low curvature, and there is no geometrically discontinuous part, which causes less shock waves, and there is no obvious shock wave structure in the flow field. While controlling the heat load, the fluid energy loss is minimized, and both active cooling and passive thermal protection technologies can be well used.

In summary, although the present invention has been disclosed above with preferred embodiments, it is not intended to limit the present invention. Any person of ordinary skill in the art, without departing from the spirit and scope of the present invention, can make various modifications. Therefore, the protection scope of the present invention shall be subject to the scope defined by the claims.

Claims (9)

1. Lobe formula rocket nozzle, including air runner and rocket gas nozzle with the intake duct UNICOM, its characterized in that: the number of the rocket gas nozzles is more than 2, each rocket gas nozzle is distributed on the outer side of the air flow channel, and a partition wall between the air flow channel and the rocket gas nozzles is of a lobe structure.

2. The lobed rocket nozzle of claim 1 wherein the intermediate wall between the air flow passage and the rocket gas nozzle has a continuous inboard lobe comprising a plurality of convex lobes and concave lobes, each rocket gas nozzle being distributed at a trailing edge of each concave lobe of the inboard lobe.

3. The lobed rocket nozzle of claim 2 wherein the air flow passage is centrally located and more than 2 rocket gas nozzles are circumferentially disposed outside the air flow passage.

4. The lobe rocket nozzle of claim 3, wherein more than 2 rocket gas nozzles are uniformly circumferentially distributed outside the air flow passage.

5. The lobed rocket nozzle of claim 2 wherein the rocket gas nozzles are distributed in the air flow path on the same side of the air flow path.

6. The lobed rocket nozzle of any of claims 1-5, wherein each rocket gas nozzle is independently controlled to open and close independently.

7. A rocket nozzle as claimed in any one of claims 2 to 5, wherein the lobes and lobes are streamlined, and the air flow passages are formed with a streamlined curved wall surface inside the partition wall between the air flow passage and the rocket gas nozzle.

8. The lobed rocket nozzle of claim 7 wherein the streamlined curved wall surface is designed with a flow channel profile using a spiral eigen-line method or a bi-directional streamline tracing method.

9. A rocket-based combined cycle propulsion system having the lobed rocket nozzle of claim 1.

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