阅读说明：本技术 一种高超声速飞行器高效激波减阻系统 (High-efficiency shock wave resistance reduction system of hypersonic aircraft ) 是由 黄杰 高代阳 姚卫星 曹镜 于 2020-04-26 设计创作，主要内容包括：本发明公开了一种高超声速飞行器高效激波减阻系统,利用扰流管、反向喷流和环形喷流组合而成的系统实现对高超声速来流的压缩,减弱飞行器前方的激波强度,降低激波阻力。本发明的高超声速飞行器高效激波减阻系统能大幅降低飞行器的激波阻力,提高飞行器的飞行速度,减少飞行器所携带的燃料,推动高超声速技术的发展。此外,本发明还具有结构简单,易于制造、安装和更换的优点,且对高超声速飞行器具有通用性。(The invention discloses a high-efficiency shock wave resistance reduction system of a hypersonic aircraft, which is characterized in that a system formed by combining a turbulent flow pipe, a reverse jet flow and an annular jet flow is utilized to compress hypersonic incoming flow, the shock wave strength in front of the aircraft is weakened, and the shock wave resistance is reduced. The high-efficiency shock wave resistance reducing system of the hypersonic aircraft can greatly reduce the shock wave resistance of the aircraft, improve the flight speed of the aircraft, reduce fuel carried by the aircraft and promote the development of a hypersonic technology. In addition, the invention also has the advantages of simple structure, easy manufacture, installation and replacement and universality for the hypersonic flight vehicle.)

1. A high-efficiency shock wave resistance reducing system of a hypersonic aircraft is characterized by comprising a turbulence pipe, wherein the turbulence pipe is a hollow slender round pipe with a uniform cross section, one end of the turbulence pipe is fixedly connected with the head of the hypersonic aircraft, the axis of the turbulence pipe is superposed with the longitudinal axis of a coordinate system of the hypersonic aircraft, and the turbulence pipe is used for compressing hypersonic incoming flow and weakening the shock wave strength in front of the hypersonic aircraft so as to reduce the shock wave resistance of the aircraft;

the turbulent flow tube and the head of the hypersonic aircraft are adjacent, one end of the turbulent flow tube is connected with air flow led out from the interior of the hypersonic aircraft, the other end of the turbulent flow tube is provided with a reverse jet port for uniformly ejecting the connected air flow outwards along the turbulent flow tube, the jet direction is parallel to the axis of the turbulent flow tube, so that a front end stagnation point area of the turbulent flow tube is cooled, the pneumatic heating effect of the stagnation point area under the hypersonic incoming flow condition is weakened, the non-ablation of the turbulent flow tube is realized, and the high-efficiency shock wave resistance reduction system of the hypersonic aircraft can work safely and.

2. The hypersonic aircraft high-efficiency shock wave drag reduction system of claim 1, wherein the turbulent flow tube is provided with an annular jet flow port on the wall surface at one end provided with the reverse jet flow port, and the jet direction is perpendicular to the axis of the turbulent flow tube and faces outwards;

the annular jet flow port is used for jetting annular jet flow and is combined with the reverse jet flow to enhance the compression effect of the turbulent flow pipe on hypersonic incoming flow, weaken the shock wave strength in front of the aircraft and further reduce the shock wave resistance of the aircraft.

Technical Field

The invention relates to the technical field of aerospace, in particular to a high-efficiency shock wave drag reduction system of a hypersonic aircraft.

Background

In order to reduce the stagnation point heat flux density, the hypersonic aircraft is usually designed to be a blunt body, which is the most important difference in appearance from the supersonic aircraft. Due to the blunt design of the hypersonic aircraft, strong isolated shock waves can be formed in front of the hypersonic aircraft, and huge shock wave resistance is generated. The shock wave resistance increases fuel and launching cost carried by the hypersonic aircraft, reduces the range or range of the aircraft, reduces the flight Mach number of the aircraft and the like, and seriously influences the performance of the aircraft. Therefore, reducing shock resistance has become a key technology for designing hypersonic flight vehicles.

Several researchers have now devised and invented systems for reducing shock resistance in hypersonic aircraft, most typically head mounted pneumatic rods. The pneumatic rod system can reduce the shock wave resistance of the aircraft to a certain extent, and the value and the test prove that the pneumatic rod system can reduce the shock wave resistance of the aircraft. However, this system has two significant design drawbacks:

1. the drag reduction efficiency of the pneumatic rod is limited, and the requirement of efficient drag reduction of a hypersonic aircraft in the future cannot be met;

2. because the diameter of the pneumatic rod is very small, the front end stagnation point area of the pneumatic rod can bear extremely serious pneumatic heating effect, the temperature of the pneumatic rod is overhigh, the pneumatic rod is burnt, and the system can not work normally. If the pneumatic rod is not burnt, the flight speed of the hypersonic aircraft is required to be reduced, so that the system can only be used for shock wave drag reduction under a lower Mach number, and the application Mach number range is limited.

Therefore, there is a need to design a novel shock wave drag reduction system for hypersonic aircraft, which has high drag reduction efficiency and can work safely and reliably even under the condition of extremely high flight Mach number.

Disclosure of Invention

The invention aims to solve the technical problem of providing a high-efficiency shock wave drag reduction system of a hypersonic aircraft aiming at the defects in the background technology so as to solve the problems of low drag reduction efficiency and limited application Mach number range of the traditional drag reduction system.

The invention adopts the following technical scheme for solving the technical problems:

a high-efficiency shock wave resistance reducing system of a hypersonic aircraft comprises a turbulence pipe, wherein the turbulence pipe is a hollow slender round pipe with a uniform cross section, one end of the turbulence pipe is fixedly connected with the head of the hypersonic aircraft, the axis of the turbulence pipe is overlapped with the longitudinal axis of a coordinate system of the hypersonic aircraft, and the turbulence pipe is used for compressing hypersonic incoming flow and weakening the shock wave strength in front of the hypersonic aircraft so as to reduce the shock wave resistance of the aircraft;

the turbulent flow tube and the head of the hypersonic aircraft are adjacent, one end of the turbulent flow tube is connected with air flow led out from the interior of the hypersonic aircraft, the other end of the turbulent flow tube is provided with a reverse jet port for uniformly ejecting the connected air flow outwards along the turbulent flow tube, the jet direction is parallel to the axis of the turbulent flow tube, so that a front end stagnation point area of the turbulent flow tube is cooled, the pneumatic heating effect of the stagnation point area under the hypersonic incoming flow condition is weakened, the non-ablation of the turbulent flow tube is realized, and the high-efficiency shock wave resistance reduction system of the hypersonic aircraft can work safely and.

As a further optimization scheme of the high-efficiency shock wave resistance reducing system of the hypersonic aircraft, an annular jet flow port is arranged on the wall surface of one end, provided with the reverse jet flow port, of the turbulent flow pipe, and the jet direction is perpendicular to the axis of the turbulent flow pipe and faces outwards;

the annular jet flow port is used for jetting annular jet flow and is combined with the reverse jet flow to enhance the compression effect of the turbulent flow pipe on hypersonic incoming flow, weaken the shock wave strength in front of the aircraft and further reduce the shock wave resistance of the aircraft.

Compared with the prior art, the invention adopting the technical scheme has the following technical effects:

1. the combined action of the reverse jet flow and the annular jet flow can enhance the compression effect of the turbulent flow pipe on hypersonic incoming flow, continuously weaken the shock wave strength in front of the aircraft and further reduce the shock wave resistance of the aircraft;

2. the reverse jet flow can cool the stagnation point area at the front end of the turbulent flow pipe, the pneumatic heating effect of the stagnation point area under the condition of hypersonic incoming flow is weakened, and the non-ablation of the turbulent flow pipe is realized;

3. the problem that the resistance reduction efficiency of a traditional hypersonic aircraft resistance reduction system is low can be solved;

4. the problem that the Mach number range of the traditional drag reduction system of the hypersonic aircraft is limited can be solved, and the drag reduction system can work safely and reliably under any Mach number;

5. the length of the turbulent flow pipe and the total pressure of the reverse jet flow and the annular jet flow can be adjusted according to requirements, and the resistance reduction efficiency is controlled;

6. the method has universality on shock wave drag reduction of the hypersonic aircraft;

7. simple structure, easy manufacture and installation.

Drawings

FIG. 1 is a perspective view of the present invention;

FIG. 2 is a front view of the present invention;

FIG. 3 is a left side view of the present invention;

FIG. 4 is a global dimensioning and direction of hypersonic incoming flow;

FIG. 5 is a partial dimensional label of the front end of the turbulator tube;

FIG. 6 is a computational grid;

FIG. 7 is a Mach and pressure cloud plot of a flow field for a drag-free system;

FIG. 8 is a Mach and pressure cloud plot of a flow field of a conventional drag reduction system;

FIG. 9 is a Mach cloud and pressure cloud of the flow field of the drag reduction system of the present invention;

FIG. 10 is a comparison of aircraft wall pressure distributions for a drag reduction system without a drag reduction system, a conventional drag reduction system, and a drag reduction system of the present invention.

Figure 11 is a comparison of heat flux density distributions at the front ends of the turbulators for a conventional drag reduction system aerodynamic rod and a drag reduction system of the present invention.

In the figure, 1-hypersonic aircraft, 2-turbulence pipe, 3-reverse jet, 4-annular jet, 5-original isolated shock wave without resistance-reducing system, 6-isolated shock wave in front of turbulence pipe, 7-reattachment shock wave of aircraft, and 8-reverse jet formed Mach disk.

Detailed Description

The technical scheme of the invention is further explained in detail by combining the attached drawings:

the present invention may be embodied in many different forms and should not be construed as limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the invention to those skilled in the art. In the drawings, components are exaggerated for clarity.

Fig. 1, fig. 2 and fig. 3 respectively show a perspective view, a front view and a left view of a high-efficiency shock wave drag reduction system of a hypersonic aircraft, the high-efficiency shock wave drag reduction system comprises a turbulent flow pipe, the turbulent flow pipe is a hollow slender round pipe with a uniform cross section, one end of the turbulent flow pipe is fixedly connected with the head of the hypersonic aircraft, the axis of the turbulent flow pipe is superposed with the longitudinal axis of a coordinate system of the hypersonic aircraft, and the turbulent flow pipe is used for compressing hypersonic incoming flow and weakening the shock wave intensity in front of the hypersonic aircraft, so that the shock wave resistance of the aircraft is reduced;

the turbulent flow tube and the head of the hypersonic aircraft are adjacent, one end of the turbulent flow tube is connected with air flow led out from the interior of the hypersonic aircraft, the other end of the turbulent flow tube is provided with a reverse jet port for uniformly ejecting the connected air flow outwards along the turbulent flow tube, the jet direction is parallel to the axis of the turbulent flow tube, so that a front end stagnation point area of the turbulent flow tube is cooled, the pneumatic heating effect of the stagnation point area under the hypersonic incoming flow condition is weakened, the non-ablation of the turbulent flow tube is realized, and the high-efficiency shock wave resistance reduction system of the hypersonic aircraft can work safely and.

The turbulent flow pipe is also provided with an annular jet port on the wall surface at one end provided with the reverse jet port, and the jet direction is vertical to the axis of the turbulent flow pipe and faces outwards;

the annular jet flow port is used for jetting annular jet flow and is combined with the reverse jet flow to enhance the compression effect of the turbulent flow pipe on hypersonic incoming flow, weaken the shock wave strength in front of the aircraft and further reduce the shock wave resistance of the aircraft.

The hypersonic aerocraft is simulated by a composite body formed by a hemisphere and a cylinder, and the turbulent flow tube and the hypersonic aerocraft have the same axis. The reverse jet flow nozzle is a reverse jet flow opening at one end of the head of the hypersonic aircraft far away from the turbulence pipe, the annular jet flow nozzle is an annular jet flow opening on the side wall of the turbulence pipe, and the two nozzles can uniformly spray air outwards.

FIGS. 4 and 5 are respectively a geometric model of the whole and a size label of the front end of the turbulence pipe, wherein D is the diameter of the hemisphere at the front part of the aircraft, h is the height of the cylinder at the rear part of the aircraft, L is the length of the turbulence pipe, D is the diameter of the turbulence pipe, and L is the diameter of the turbulence pipe₁Is the distance between the annular jet orifice and the reverse jet orifice, h₁Is the height of the reverse jet opening so as toThe values of the upper geometry are listed in table 1. Table 2 shows Mach number Ma of hypersonic incoming flow_∞Static pressure P_∞Static temperature T_∞And angle of attack α assuming that the reverse jet and the annular jet have the same jet parameters, Table 3 lists the Mach number Ma of the two jets_jTotal pressure P_0jAnd total temperature T_0j。

TABLE 1

TABLE 2

TABLE 3

The numerical model is as follows:

FIG. 6 is a calculation grid and boundary conditions of the numerical model, the outer wall surfaces of the aircraft and the turbulence tubes are defined as isothermal wall surfaces of 300K, the reverse jet flow and the annular jet flow are defined as pressure inlets, and other boundary conditions comprise a symmetry axis and a pressure far-field boundary, besides, the total number of units of the calculation grid is 392555, the height of the first layer of grid on the wall surface is 1 × 10^-3mm。

The calculation results were analyzed as follows:

to verify the high efficiency of the drag reduction system of the present invention, numerical calculations were also performed here for both the drag reduction-free system and the hypersonic aircraft using the conventional drag reduction system (pneumatic rod system). Fig. 7, fig. 8 and fig. 9 are respectively a mach cloud and a temperature cloud of a flow field calculated by a drag reduction system without drag reduction, a traditional drag reduction system and a drag reduction system of the present invention. When no drag reduction system is adopted, the isolated shock waves in front of the hypersonic aircraft form huge shock wave resistance, the gas in the shock wave layer has high pressure after being strongly compressed, and the high-pressure area is positioned in a stagnation area at the front end of the aircraft. When the traditional drag reduction system is adopted, the hypersonic incoming flow is pre-compressed at the most front end of the pneumatic rod, a weak isolated shock wave is formed in front of the pneumatic rod, and the pre-compressed gas continues flowing downstream and is subjected to secondary compression of an aircraft. Therefore, the original isolated shock wave in front of the aircraft is converted into the reattachment shock wave, the shock wave intensity and the gas pressure in the reattachment shock wave layer are both reduced, and the high-pressure zone in the reattachment shock wave layer is positioned in the middle of the hemisphere. When the drag reduction system of the present invention is employed, comparing fig. 8 and 9, it can be observed that the pre-compression effect of the system on hypersonic incoming flow is significantly stronger than that of the conventional drag reduction system, which greatly reduces the air velocity in front of the re-attached shock wave, so that both the re-attached shock wave intensity and shock layer air pressure in front of the aircraft are significantly lower than those of the conventional drag reduction system.

Fig. 10 shows the wall pressure distribution of the aircraft calculated by the above analysis model 3, and the wall pressure distribution determines the magnitude of the aircraft drag. The results show that at x>Within the range of 20mm, the difference of the aircraft wall pressure distribution of the three analysis models is small. Aiming at the traditional drag reduction system, in the range of x being 0-5mm, the pressure intensity of the wall surface of the aircraft is obviously lower than the calculation result of the drag reduction-free system; and in the range of x being 5-20mm, the pressure of the wall surface of the aircraft is higher than the calculated result of the drag-free system. Aiming at the resistance reducing system, when x is 0-20mm, the pressure intensity of the wall surface of the aircraft is obviously lower than the calculation results of the non-resistance reducing system and the traditional resistance reducing system. Table 4 lists the aircraft wall pressure peak P for the 3 analytical models_maxAnd the drag coefficient C of the entire model_dCoefficient of resistance C_dIs defined as:

in the formula: f_dIs the aerodynamic resistance of the entire model; rho_∞Is the incoming flow density; v_∞Is the incoming flow rate; s is a reference area defined as π D²/4. The result shows that the aircraft wall pressure peak value P of the drag reduction system of the invention_maxCompared with the non-resistance-reducing system and the traditional resistance-reducing system, the resistance coefficient C of the resistance-reducing system is respectively reduced by 76.525 percent and 64.217 percent_dThe reduction is 66.220% and 58.225% compared with the non-drag reduction system and the traditional drag reduction system respectively. Therefore, the efficiency of the drag reduction system is obviously higher than that of the traditional drag reduction system, and the drag reduction system is an efficient drag reduction system.

TABLE 4

Fig. 11 shows the heat flux density distribution at the front end of the spoiler tube of the drag reduction system of the present invention and the pneumatic rod of the conventional drag reduction system. The result shows that the heat flux density distribution at the front end of the turbulence pipe of the drag reduction system is obviously lower than that of a pneumatic rod of the traditional drag reduction system. The pneumatic rod of the traditional drag reduction system is subjected to extremely serious pneumatic heating, and the stagnation point (most front end) heat flow of the traditional drag reduction system reaches 414.717kW/m²This would burn the pneumatic rod and the conventional drag reduction system would not work properly at high mach numbers. The front end of the turbulence pipe of the resistance reducing system is cooled by the reverse jet flow, and because the static temperature of the jet flow is lower than the isothermal wall surface of 300K, the front end of the turbulence pipe generates negative heat flow, and the non-ablation of the turbulence pipe is realized, so the resistance reducing system can still work safely and reliably under the high Mach number.

The efficient shock wave drag reduction system has strong universality on a hypersonic aircraft, the specific implementation mode of the efficient shock wave drag reduction system is explained only through a typical numerical calculation example, the length of the turbulent flow pipe and the total pressure of the reverse jet flow and the annular jet flow can be adjusted according to actual requirements, the drag reduction efficiency is controlled, and the problems that the traditional drag reduction system is low in drag reduction efficiency and limited in application Mach number range are solved.

It will be understood by those skilled in the art that, unless otherwise defined, all terms (including technical and scientific terms) used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. It will be further understood that terms, such as those defined in commonly used dictionaries, should be interpreted as having a meaning that is consistent with their meaning in the context of the prior art and will not be interpreted in an idealized or overly formal sense unless expressly so defined herein.

The above-mentioned embodiments, objects, technical solutions and advantages of the present invention are further described in detail, it should be understood that the above-mentioned embodiments are only illustrative of the present invention and are not intended to limit the present invention, and any modifications, equivalents, improvements and the like made within the spirit and principle of the present invention should be included in the protection scope of the present invention.