A high-speed vehicle with flow layout and its design optimization method

Abstract

The present invention provides a high-speed aircraft with a flow layout, including an aircraft body. The aircraft body is streamlined along the navigation direction, and the aircraft body is symmetrical with respect to a horizontal plane passing through the axis, or symmetrical with respect to a vertical plane passing through the axis. The aircraft body is divided into a flow-facing portion and a flow-removing portion with its widest position as the boundary. The flow-facing portion is a cone with spikes, and the horizontal or vertical section line of the flow-facing portion passing through the axis is two line segments symmetrical along the axis. Both section lines are described by an inverse tangent function, and the parameters of the function are adjusted to change the spanwise dimension of the aircraft body and the length of the spikes. The present invention can significantly reduce the flow resistance of the aircraft during navigation. Due to the significant reduction in flow resistance, the maximum speed of the aircraft can be significantly increased under the same power, significantly reducing energy consumption during navigation, which is of great significance for energy conservation and emission reduction.

Description

Technical Field

[0001] The present invention relates to a high-speed aircraft, in particular to a high-speed aircraft with a flow layout and a design optimization method thereof. Background Art

[0002] High-speed vehicles mainly refer to man-made equipment and devices that travel at high speed in the air or water, including airplanes, missiles or other aircraft flying at high speed in the air, surface ships sailing at high speed in the water (especially the parts below the waterline), and submarines, unmanned submarines and other submersibles sailing at high speed underwater; the so-called high speed means that the vehicle has a high sailing speed. For aircraft, it usually refers to a speed exceeding the speed of sound. For underwater vehicles, it usually refers to a speed exceeding 50km / h.

[0003] Flow layout mainly refers to the optimized design of the flow shape of the aircraft based on the principles of fluid dynamics, as well as the related components set up to improve flow efficiency, reduce flow resistance and improve the function of the aircraft.

[0004] Objects traveling through a fluid are subject to flow resistance. A vehicle's propulsion is primarily used to overcome this resistance during navigation. This resistance increases with speed, generally proportional to the square of the speed. Therefore, as speed increases, resistance rapidly increases, requiring high-speed vehicles to consume significant amounts of energy to overcome resistance. This rapidly increasing resistance is a major factor limiting a vehicle's maximum speed. Therefore, effectively reducing resistance during high-speed navigation presents a significant challenge. Summary of the Invention

[0005] The technical problem to be solved by the present invention is to address the deficiencies of the above-mentioned existing technologies and to provide a high-speed aircraft with a flow layout and a design optimization method thereof. The aircraft can significantly reduce the flow resistance of the aircraft during navigation. Since the flow resistance is significantly reduced, the maximum speed of the aircraft can be greatly improved under the same power, which significantly reduces the energy consumption during navigation, and is of great significance for energy conservation and emission reduction.

[0006] To solve the above technical problems, the present invention adopts a technical solution: a high-speed vehicle with a flow layout, characterized by including a vehicle body, the vehicle body being streamlined and symmetrical about a horizontal plane passing through an axis, or about a vertical plane passing through the axis, the axis being a line connecting the centers of the vehicle body's cross sections, the axis generally coinciding with the vehicle's heading, the horizontal length of the vehicle body being greater than its vertical width, the vehicle body being divided into a flow-incoming portion and a flow-outward portion, the boundary between the flow-incoming portion and the flow-outward portion being located at the position where the vehicle body's horizontal and vertical widths are greatest, the flow-incoming portion being a cone, the front end of the flow-incoming portion being provided with a spike, the horizontal or vertical section line of the flow-incoming portion passing through the axis being two symmetrical line segments about the axis, both sections being described by a function with an inverse tangent function as a key factor. Specifically, if the horizontal section line of the flow-incoming portion passing through the axis and the vertical section line passing through the axis have the same dimensional parameters, and if the dimensional parameters of any section line passing through the axis are the same, then the vehicle body is a body of revolution.

[0007] Preferably, the spikes are fixed or retractable. The spikes are usually longer, which will increase the axial size of the aircraft body. The spikes are arranged into multiple segmented retractable structures, and the extension and retraction are controlled by a hydraulic system. They can be retracted into the head of the aircraft body when moored or sailing at low speed, and extended when sailing at high speed.

[0008] Preferably, the aircraft body is provided with fins, and the number of fins is one or more. The main function is to guide the flow, so that the heading of the aircraft remains stable and prevents rolling, and the heading of the aircraft can also be changed according to control. When a vertical fin is provided at the rear of the aircraft body, the rear half of the fin is rotatable along the axis, and the rotating part can act as a rudder to control the rolling of the aircraft when turning. When two symmetrical fins with a large area are provided in the horizontal direction of the aircraft body, for the aircraft body flying at high speed in the air, the two fins can act as wings and generate lift. When the rear half of the fin is rotatable, the movable part acts as an elevator to control the pitch of the aircraft. The rudder and the elevator can be used in combination to control the heading of the aircraft.

[0009] Preferably, a beam hood is provided around the outer periphery of the vehicle body. The beam hood is a thin-walled tubular component. The beam hood is wrapped around the outer periphery of the vehicle body and maintains a distance from the vehicle body. The beam hood is connected to the vehicle body via a fin. The space between the inner wall of the beam hood and the vehicle body constitutes a flow channel, through which the fluid around the vehicle body flows. The length of the beam hood is usually not greater than the length of the vehicle body. The function of the beam hood is to constrain the high-energy fluid around the vehicle body to flow within a certain spatial range, preventing the high-energy fluid from flowing into the surrounding environment, exchanging matter and energy with the surrounding environment, and thus causing energy leakage. During a vehicle's navigation, the fluid ahead of it is affected by the vehicle's incoming flow, causing changes in velocity direction, magnitude, and pressure, resulting in a spanwise velocity component, i.e., a velocity component across the width of the vehicle. For compressible fluids like air, the spanwise static pressure increases significantly away from the vehicle's centerline. For incompressible fluids like water, the spanwise velocity increases significantly away from the vehicle's centerline. Without a beam shield to constrain the spanwise periphery, both the compressible fluid with increased static pressure and the incompressible fluid with increased velocity will flow into the flow field surrounding the vehicle, exchanging matter and energy with the surrounding flow field, resulting in energy leakage. This energy leakage reduces the pressure on the outflow side of the vehicle, increasing the pressure differential resistance between the incoming and outflow sides. Therefore, the placement of a beam shield helps reduce fluid energy leakage and lower the vehicle's drag. While the beam shield increases the vehicle's spanwise dimensions, the spacing between the beam shield and the vehicle body should be within a reasonable range; a too small spacing can increase the vehicle's flow resistance. In practical applications, the beam shield can be positioned as needed.

[0010] Preferably, a propeller is provided at the tail end of the vehicle body. Using the fluid surrounding the vehicle body as a working medium, the propeller can provide thrust for the vehicle body's forward motion, overcoming the vehicle's flow resistance. While the propeller can be positioned elsewhere, placement at the tail end of the vehicle body generally provides higher propulsion efficiency.

[0011] A design optimization method for a high-speed vehicle with a flow layout, characterized by comprising the following steps:

[0012] S1. Use a function to describe the first section line of the horizontal section of the flow-impinging portion through the axis: the function describing the first section line of the flow-impinging portion is y1 = arctan(x);

[0013] S2. Introducing the width coefficient: In order to control the width of the spacecraft, the width coefficient w is introduced into the function y1, and y1 = w*arctan(x / w) is obtained. The larger w is, the larger the width of the spacecraft is, that is, the longer the dividing line is;

[0014] S3. To control the length of the spike, the spike length coefficient c is introduced into the function y1, and the function y1 = w*(arctan(x / w)-c) is obtained, where 0 < c < π / 2. The larger the value of c, the longer the spike length.

[0015] S4. Describe the second section line of the horizontal section of the flow section through the axis: Since the two section lines are symmetrical about the axis of the vehicle body, the describing function describing the second section line is the function of the first section line described by flipping it along the symmetry axis x, that is, y2 = -w*(arctan(x / w)-c);

[0016] S5. Let y1 = y2, and solve for the x coordinate value x1, and we get x1 = w*tan(c). x1 is the x coordinate value of the apex of the spike, and x2 represents the x coordinate of the maximum width of the vehicle body. If x2 < 0, the width of the vehicle body is 2*w*(c-arctan(x2 / w)). Thus, we get two descriptive functions of the horizontal section line of the flow section through the axis, y1 = w*(arctan(x / w)-c) and y2 = -w*(arctan(x / w)-c). The range of x is, x2 <x<x1。

[0017] Preferably, in order to obtain relatively ideal flow efficiency, x2 is usually taken as -1 / 2*w*π, c>0.95*π / 2. Considering that the front end of the spike is long and very thin, occupies a large space and is easy to break, the front end of the spike is usually appropriately cut off. The actual x value of the front end of the spike is smaller than x1, and x<0.8*w*tan(c) is usually taken.

[0018] S6. Repeat S1-S5 to obtain the description function of the two profile lines that describe the vertical profile of the frontal part through the axis. The width coefficient w and the spike length coefficient c can be set to different values from the horizontal profile line description function. When the w and c values are the same as those of the horizontal profile description function, the aircraft body is a rotating body.

[0019] After determining the describing functions of the horizontal and vertical section lines of the oncoming part through the axis, the other horizontal and vertical section lines of the oncoming part can be determined by interpolation and fitting, thereby completing the three-dimensional modeling design of the oncoming surface.

[0020] Among the two descriptive functions y1 and y2, the inverse tangent function is the key factor, which determines the overall flow shape of the oncoming part and plays a key role in improving flow efficiency and reducing flow resistance. w and c are adjustment parameters. By adjusting these two parameters, the needs of the vehicle under various application conditions can be met.

[0021] For high-speed aircraft, flow resistance is generally considered to consist of three components: pressure differential drag, induced drag, and frictional drag. Induced drag is also caused by pressure differential and can be categorized as pressure differential drag. When the aircraft's speed reaches or exceeds the critical Mach number, shock waves are generated, accompanied by shock wave drag. Shock wave drag originates from the total pressure loss of the incoming flow caused by the shock wave and the shock wave-induced separation, resulting in a significant pressure differential between the aircraft's incoming and outgoing flows. Therefore, it also constitutes pressure differential drag. Shock wave drag increases rapidly with increasing Mach number. Therefore, for high-speed aircraft, shock wave drag is the primary factor in flow resistance. The key to reducing flow resistance lies in reducing shock wave drag. The present invention addresses this issue from two perspectives: first, minimizing the fluid pressure in the incoming flow while simultaneously reducing fluid pressure loss in the outgoing flow; and second, reducing the pressure component of the fluid pressure in the incoming flow in the heading (axial) direction, as it is the pressure component in the heading (axial) direction that actually creates resistance to the aircraft.

[0022] The oncoming part of an aircraft flying at high speed in the air is designed to be cone-shaped with spikes. Its cross-section line is a function waveform segment with the inverse tangent function as the key factor and a function waveform segment symmetrically flipped along the axis. The spikes can optimize the shock wave distribution around the spikes and avoid the occurrence of strong shock waves. Specifically, there are two major benefits: on the one hand, when the supersonic airflow passes through the spikes at the front of the cone, a Mach wave array composed of countless conical Mach waves (or weak shock waves very close to Mach waves) will be formed in the area around the spikes. These shock waves will not superimpose into a strong shock wave, but will be regularly distributed in a certain area around the spikes. When the supersonic airflow passes through the Mach wave array, it can be considered to be isentropically compressed. The airflow around the spikes has not yet been completely compressed, and the speed is greater than the speed of sound, and the static pressure is small. The airflow flows through the rear of the spikes and passes through a weak positive shock wave (also a Mach wave, at this time the Mach number M=1) before the airflow speed drops to the speed of sound. This design can reduce the pressure on the overall oncoming part of the aircraft. The backflow part of the aircraft is designed to be streamlined, which can reduce the flow loss of the fluid flowing through the backflow part, which is beneficial to reducing the pressure difference resistance of the aircraft. On the other hand, the direction of the fluid pressure on the spike part in the front of the incoming part is the normal direction of the incoming part. The direction of the fluid pressure is almost perpendicular to the axis of the aircraft, and the angle with the span direction of the aircraft is small. The fluid pressure is decomposed in the heading (axial) and span directions. The component of the fluid pressure in the span direction is large (for a rotating body, this force component will be offset by a symmetrical and equal reverse force), and the component in the heading (axial) direction is very small, which is also very beneficial to reducing navigation resistance. Combining the above two factors, the front of the incoming part of the aircraft body is designed as a spike, which can greatly reduce the pressure difference resistance between the incoming part and the backflow part of the aircraft.

[0023] The spikes at the front of the incoming airflow create a Mach wave array around them. As the supersonic airflow passes through this array, it is gradually compressed by countless Mach waves. The entire compression stroke is roughly equivalent to the length of the spikes, typically on the order of decimeters and meters. Compared to a strong shock wave, the compression stroke is increased by millions to tens of millions of times. Within the Mach wave array, the gas pressure and velocity exhibit a very regular, cascaded distribution. The compression process is equivalent to adding a buffer zone millions to tens of millions of times longer. The intensity of collisions between gas molecules is greatly reduced, and the high-frequency harmonic components in the collision energy spectrum are significantly reduced. The proportion of energy absorbed by gas molecules and converted into molecular internal energy is also significantly reduced. Most of the gas's kinetic energy is converted into potential energy between gas molecules. This is the microscopic reason for achieving efficient isentropic compression.

[0024] When the vehicle is submerged in water, the spikes effectively reduce the turning angle of the incoming flow section. This is particularly noticeable at the front end, where the turning angle is very small. Consequently, even at high speeds, the fluid pressure remains low. The fluid pressure on the incoming flow section acts in the normal direction, with a large angle between this pressure direction and the vehicle's axis. At the very front of the spikes, the angle is typically close to π / 2 (radians). This decomposition of the fluid pressure into the vehicle's heading (axial) and spanwise directions results in a larger component in the spanwise direction (for symmetrical rotating bodies, this force can be offset), while a smaller component in the heading (axial) direction helps reduce flow resistance.

[0025] Compared with the prior art, the present invention has the following advantages:

[0026] 1. The present invention can achieve a certain degree of energy conservation and emission reduction. Since the flow resistance of the aircraft is reduced, the required driving force power is greatly reduced at the same speed, which is particularly beneficial when sailing at high speed.

[0027] 2. The present invention enables aircraft to reach higher speeds. The maximum speed of an aircraft is typically limited by power and navigational economics. The present invention can significantly reduce the flow resistance of an aircraft during high-speed navigation. Therefore, the same power can significantly increase the maximum speed of the aircraft, while also improving navigational economics.

[0028] 3. When applied in the aviation field, the present invention can be applied to various aircraft, including supersonic aircraft, supersonic drones, and supersonic missiles, thereby achieving higher flight speeds and superior overall performance. The aircraft body is designed as a rotating body, with fins and a cylindrical beam shield installed on the exterior. During high-speed flight, it functions as a ramjet compressor. The flow within the compressor is theoretically isentropic, i.e., isentropic compression, with a theoretical compression efficiency of 100%. This ramjet compressor can be further integrated into a ramjet engine or a combined engine.

[0029] In the field of navigation, the part below the waterline of a ship can be designed and modified according to the flow layout of the present invention, which can greatly improve the ship's navigation speed and cruising range, reduce energy consumption, and have good economy; the submarine can be designed and modified according to the flow layout of the present invention, which can greatly improve the navigation speed and cruising range, and the beam cover can also play a certain role in sound insulation and noise reduction. If it is equipped with absorbing materials, the noise of the submarine is expected to be greatly reduced, greatly improving the stealth and survivability of the submarine.

[0030] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Figure 1 Schematic diagram of the flow layout of the ram compressor in Example 1 of the present invention.

[0032] Figure 2 This is a schematic diagram of the shock wave array of the ramjet compressor head in Example 1 of the present invention.

[0033] Figure 3 Schematic diagram of the flow layout of the submarine in Example 2 of the present invention.

[0034] Figure 4 This is a schematic diagram of the pressure decomposition of a certain micro unit of a submarine in Example 2 of the present invention.

[0035] Figure 5 It is a schematic diagram of the axial pressure distribution of the flow-facing part of Example 2 of the present invention.

[0036] Description of reference numerals:

[0037] 1—ramjet compressor body; 2—dividing line; 3—spike;

[0038] 4—compressor shroud; 5—streamflow section line; 6—streamflow section line;

[0039] 7—Fin; 8—Supersonic flow; 9—Mach wave;

[0040] 10—thruster; 11—rudder; 12—submarine vertical fin;

[0041] 13—Submarine body; 14—Submarine horizontal fin; 15—Fluid pressure;

[0042] 16—Axial component of pressure; 17—Spanwise component of pressure; 18—Submarine head after spike retraction;

[0043] 19—Axial pressure of the submarine's upstream part 20—Submarine after the spike is retracted 21—Axial pressure coordinate axis;

[0044] Force distribution curve; Axial pressure distribution curve of the flow-facing part;

[0045] 22—Submarine Spike; DETAILED DESCRIPTION

[0046] Example 1

[0047] like Figure 1 As shown, this embodiment provides a high-speed aircraft used in the air, namely a ram compressor, including a ram compressor body 1, a fin 7, and a compressor dome 4. The ram compressor body 1 is an axisymmetric body of revolution, divided into an incoming flow portion and an outgoing flow portion. The dividing line 2 is located at the maximum spanwise dimension of the ram compressor body 1. The portion facing the incoming flow is the incoming flow portion, which is a cone with a spike 3 at the front. The cross-section line 5 of the incoming flow portion is described by the following two functions:

[0048] y1=w*(arctan(x / w)-c) x2 <x<x1

[0049] y2=-w*(arctan(x / w)-c) x2 <x<x1

[0050] In this embodiment, the back flow portion is streamlined, and the back flow portion section line 6 is as shown in FIG. Figure 1 As shown in the figure; three fins 7 are usually provided, which fasten the compressor shroud 4 and the compressor body 1 and play a guiding role; the compressor shroud 4 is a thin-walled cylinder, covering the periphery of the compressor body 1, and the space between the compressor shroud 4 and the compressor body 1 constitutes an air flow channel, and the compressed air flows in the flow channel.

[0051] When the ramjet compressor is in operation, a shock wave array will be formed in the area around the spike 3 at the front of the incoming flow part of the compressor body 1. The schematic diagram of the shock wave array is shown in FIG. Figure 2 As shown, the Mach number M of the supersonic incoming flow 8 is greater than 1. When it flows through the front face of the compressor body 1, a conical Mach wave 9 will first be generated at the front end of the spike 3. The angle between this Mach wave and the supersonic incoming flow is the Mach angle μ, μ=arcsin(1 / M), where M is the incoming flow Mach number. At the design Mach number of the ramjet compressor, the leading edge of the compressor beam shroud 4 usually intersects with the front-end conical Mach wave 9. Such a compressor beam shroud 4 can prevent the energy of the shock wave compressed air from leaking to the outside. Each differential annular flow surface behind the front end of the spike will generate a conical Mach wave, but after the supersonic incoming flow passes through the previous Mach wave, the Mach number decreases slightly, the Mach angle increases slightly, and the static pressure increases slightly, as shown in FIG. Figure 2As shown in , these countless conical Mach waves form a Mach wave array. As supersonic incoming flow 8 passes through the Mach wave array, it undergoes isentropic compression, gradually decreasing its Mach number and increasing its static pressure, until it finally passes through a weak normal shock wave (also a Mach wave, at which point the Mach number M = 1), reducing its velocity to subsonic speeds. The subsonic airflow continues to flow backward along the flow path between the compressor dome 4 and the compressor body 1, further decreasing its velocity and increasing its static pressure, completing the supersonic compression process.

[0052] In the process of the supersonic incoming flow 8 passing through the entire Mach wave array, it is theoretically isentropically compressed with a theoretical compression efficiency of 100%. The theoretical pressure ratio of the ramjet compressor can be directly found in the isentropic flow characteristic table according to the incoming flow Mach number. For example, when the incoming flow Mach number is 2, the theoretical pressure ratio is 7.824. When the incoming flow Mach number is 3, the theoretical pressure ratio is 36.73. After deducting the boundary layer and other flow losses, the actual compression efficiency and pressure ratio should be quite close to the theoretical values.

[0053] The forces acting on the upstream part of the ram compressor body 1 are analyzed. For any micro-element surface on the upstream surface, let its spanwise projection area (windward area) be dS, the air pressure acting on the micro-element surface be P, the pressure acting on the micro-element surface be dF, and the angle between the normal of the micro-element surface and the axial direction of the ram compressor body be θ. Then the area of the micro-element surface is dS / cos(θ), the pressure acting on the micro-element surface is dF=P*dS / cos(θ), the axial component of dF is dFx=P*cos(θ)*dS / cos(θ), eliminating cos(θ), dFx=PdS, and the spanwise component of dF is dFy=P*sin(θ)*dS / cos(θ). Since the ram compressor body 1 is axisymmetric, the spanwise component dFy will be offset.

[0054] If the spike 3 at the front of the ram compressor body 1 is removed, the head of the ram compressor body will change from a pointed head to a blunt head, and the supersonic airflow will form a strong shock wave at the head. The pressure distribution at the head will change. The pressure is recorded as Ps. The supersonic incoming flow 8 will form a normal shock wave near the apex of the blunt head. The pressure Ps0 at the apex will reach a maximum value. Around the apex, the normal shock wave gradually changes into a weaker oblique shock wave, and the value of Ps gradually decreases. Ps0 can be calculated using the Rayleigh-Pitot tube formula. For any micro-surface on the frontal surface of the ram compressor body 1, assume that its spanwise projected area (frontal area) is dS, the air pressure acting on the micro-surface is Ps, the pressure acting on the micro-surface is dF, and the angle between the normal of the micro-surface and the axial direction of the ram compressor body is θ. Then the area of the micro-surface is dS / cos(θ), the pressure acting on the micro-surface dF=Ps*dS / cos(θ), the axial component of dF dFx=Ps*cos(θ)*dS / cos(θ), eliminating cos(θ), dFx=PsdS, the spanwise component of dF dFy=Ps*sin(θ)*dS / cos(θ), and since the ram compressor body 1 is axisymmetric, the spanwise component dFy will also be offset. Before the spike is removed, dFx=PdS, where P is the pressure of the supersonic airflow passing through the Mach wave array. The airflow is gradually compressed, and the P value is small. At the top of the spike 3, the P value is close to the static pressure of the supersonic incoming flow 8; after the spike 3 is removed, dFx=PsdS, where Ps is the pressure of the supersonic incoming flow 8 after passing through the normal shock wave and the oblique shock wave. Due to the compression of the strong shock wave, the Ps value is larger, that is, Ps>P. At the front end, the difference between the two is more obvious; on the other hand, after the supersonic airflow passes through the normal shock wave and the oblique shock wave, the total pressure loss is large, so the pressure loss of the backflow part of the ramjet compressor body is also large, further increasing the pressure difference resistance. From the above analysis, it can be concluded that the axial resistance of the head with the spike 3 is significantly smaller.

[0055] In this embodiment, the flow layout of the ramjet compressor can be combined and applied in multiple ways: if the rear of the ramjet compressor is connected to the combustion chamber and tail nozzle, it will form a complete ramjet engine; if it is connected to the air intake of the subsequent stage aircraft engine, it can form a combined engine; if a rocket engine is equipped at the rear end of the compressor body, it can become a supersonic missile.

[0056] If the rear of the ramjet is not connected to other equipment or devices, the compressed air in the outer flow passages of the rearward portion of the main body will re-expand and accelerate into supersonic airflow. Since the entire flow process from the air inlet to the exhaust is theoretically isentropic, the total air pressure loss is minimal, and the overall flow resistance is quite low. If two large wings, engines, and rudders are installed on either side of the shroud, the aircraft will be transformed into a supersonic aircraft with very low flow resistance.

[0057] Example 2

[0058] like Figure 3 As shown, this embodiment provides a high-speed underwater navigation vehicle, namely a submarine, including a submarine body 13, a submarine vertical fin 12, a rudder 11, a submarine horizontal fin 14, and a propeller 10. The submarine body 13 is the main body of the submarine, and its flow shape and related technical features are the same as those of the previous embodiment, with only the parameter settings being different. The cross-section line of the incoming flow portion of the submarine body 13 is also described by the following two functions:

[0059] y1=w*(arctan(x / w)-c) x2 <x<x1

[0060] y2=-w*(arctan(x / w)-c) x2 <x<x1

[0061] In this embodiment, the submarine spike 22 at the front of the upstream part of the submarine body 13 is configured to be multi-stage hydraulically controlled and retractable. After the submarine spike 22 is retracted, the head of the submarine will be similar to the head of an ordinary submarine. When the submarine is anchored or sailing at low speed, the submarine spike 22 is retracted. When sailing at high speed, the submarine spike 22 is extended; a rotatable rudder 11 is provided at the rear of the submarine vertical fin 12. In addition, two submarine horizontal fins 14 are provided below the submarine vertical fin 12, and elevators are provided at the rear of them. The rudder and elevators are used together to control the heading of the submarine; the propeller 10 uses the water flowing through the surface of the downstream part of the submarine body 13 as the working medium to provide thrust for the submarine's navigation, and can be a propeller or a pump-propelled device.

[0062] In this embodiment, Figure 4 This is a schematic diagram of the pressure decomposition of a certain micro-element surface on the submarine's frontal surface. Let its spanwise projection area (frontal area) be dS, and the angle between the normal line of the micro-element surface and the axial direction of the submarine body 13 be θ. Then the area of the micro-element surface is dS / cos(θ). The value of the fluid pressure 15 acting on the micro-element surface is dF, and the direction is the direction of the normal line of the micro-element surface. The value dF of the fluid pressure 15 can be decomposed into the axial pressure component 16, whose value is dFx, and the spanwise pressure component 17, whose value is dFy. dFx will hinder the submarine's progress. The smaller this component is, the better. dFy can be offset by another opposite component of the axisymmetric micro-element surface. From Figure 4It can be seen that the angle between dF and the axial direction is θ, so dFx=P*cos(θ)*dS / cos(θ)=PdS. It can be seen that there are two main factors that determine the axial component dFx of the pressure on the headwind surface of the submarine body 13. One is the surface pressure P, and the other is the headwind area dS. For the same headwind area, if the pressure P can be reduced, a smaller dFx can be obtained, thereby reducing the resistance. Pressure P consists of two components: P = Ph + Pc. One component is the static pressure Ph, which depends on the water depth h. It's generally assumed that the upstream and downstream components of the submarine's main body 13 cancel each other out and are therefore negligible. The other component, Pc, arises from the change in the incoming flow direction caused by the upstream surface. The momentum theorem shows that Pc is determined by the product of the flow rate through the microsurface and the change in velocity. The flow rate is determined by the product of the ship's speed and the water density, while the change in velocity is primarily determined by two factors. One is the angle between the microsurface and the incoming flow direction, which can be considered a first-order factor and plays a primary role. The other is the angle between the current microsurface and the previous microsurface in the incoming flow direction, which can be considered a second-order factor and plays a secondary role. To facilitate understanding the mechanisms of first- and second-order factors, the concept of a near-surface layer is introduced. The near-surface layer refers to a layer of fluid outside the boundary layer. The velocity of this layer has already changed, essentially flowing in close contact with the incoming surface. The first-order factor mainly affects the pressure on the outside of the near-surface layer (the side away from the oncoming surface), and the second-order factor mainly affects the pressure on the inside of the near-surface layer (the side close to the attachment layer). The following is a further analysis of these two factors:

[0063] First, analyze the first-order factor that affects Pc. This factor is the main factor. Derivative the profile description function y = w*(arctan(x / w)-c) to obtain the derivative function y' = 1 / (1+(x / w) 2 ), its value represents the slope of the tangent line of the section line at the x-coordinate, and also reflects the magnitude of the turning angle between the infinitesimal surface at the x-coordinate of the section line and the incoming flow direction. The larger the slope, the larger the turning angle. Analysis of the derivative function shows that the derivative value is maximum at x = 0, at which point y' = 1. Usually, the absolute value of x1 is greater than the absolute value of x2. Therefore, at the vertex of the submarine spike 22, x1 = w*tan(c), the derivative value is minimum, at which point y' = 1 / (1+(tan(c))) 2 ), it can be seen that from the front end of the submarine spike 22 to the rear end of the oncoming surface, the turning angle first increases gradually, reaches the maximum at x=0, and then gradually decreases, which reflects the overall change of the pressure Pc on the oncoming surface.

[0064] Then analyze the second-order factors affecting Pc, and the derivative function of the profile description function y'=1 / (1+(x / w) 2 ) Take the derivative again and we get y'=-2x / (w+x) 2, take the derivative of y' again and get y''=4x / (w+x) 3 -2 / (w+x) 2 , let y''=0, get x=w, substitute x=w into the second-order derivative function y', get y'=-1 / (2w), this is an extreme value of y' when x>0, usually a value much smaller than the maximum value of the first-order derivative y' 1. It can be seen that the second-order factor has a limited impact on the Pc value and is a secondary factor. Generally speaking, as long as the profile line of the oncoming surface is a smooth curve, the value of the second-order factor y' is usually relatively small, and has little effect on P c The impact of P c The first-order and second-order factors of

[0065] It can be concluded that the submarine spike 22 optimizes the flow field pressure distribution at the head of the submarine body 13, which is very beneficial to reducing the flow resistance of the submarine.

[0066] Through the above analysis, we can get the following: Figure 5 The schematic diagram of axial pressure distribution on the submarine's frontal surface is shown in the figure. Figure 5 The left side of the figure shows the flow-facing part of the submarine body 13, and the right side shows the axial pressure component curve along the span direction of the flow-facing part. The axial pressure distribution curve 19 of the submarine's flow-facing surface shows the distribution of the axial pressure component along the span direction on the flow-facing surface. It can be seen that the axial pressure component at the top of the submarine spike 22 in the flow-facing part is relatively small, because the pressure there is only slightly greater than the static pressure Ph. The integral of this curve is as follows: Figure 5 The shaded area represents the total value of the axial pressure of the fluid on the front surface.

[0067] In this embodiment, when the submarine spike 22 is retracted into the head of the submarine body under hydraulic control, the head of the submarine becomes the head of an ordinary submarine. The submarine head 18 after the submarine spike 22 is retracted is as shown in FIG. Figure 5 The dotted line on the left of the figure shows this. By decomposing the pressure on the incoming surface, we can obtain the axial pressure distribution curve 20 on the incoming surface of the submarine after the submarine spike 22 is retracted. The maximum value of the curve is on the axial pressure coordinate axis 21, which corresponds to the front end of the submarine body 13 and is a stagnation point of the incoming flow. The fluid deflection angle here can be considered to be π / 2 (radians), and the pressure is equal to the total pressure P0 of the incoming flow. According to the Bernoulli equation, we can get:

[0068] P0=Ph+1 / 2*ρ*V 2

[0069] Where Ph is the static pressure of water, ρ is the density of water, and V is the speed of the submarine. As can be seen, P0 includes the square factor of the speed. The higher the speed, the faster the value of P0 increases. This is an important factor that restricts the maximum speed of underwater vehicles.

[0070] The axial pressure distribution curves 20 and 21 on the submarine's frontal surface Figure 5 The area enclosed by the central vertical axis is the total axial pressure on the submarine's frontal surface after the submarine spike 22 is retracted. Clearly, this area is significantly larger than the area enclosed by the vertical axis and the axial pressure distribution curve 19 on the submarine's frontal surface, indicating that the submarine experiences significantly greater axial resistance after the submarine spike 22 is retracted. The difference between the two values increases with increasing speed. Therefore, it can be concluded that the submarine spike 22 can reduce flow resistance, achieve higher speeds, extend range, and improve energy efficiency. To increase the submarine's displacement while minimizing flow resistance, the length of the submarine's rearward portion can be significantly lengthened.

[0071] This embodiment can be applied to surface ships. By transforming the part below the waterline of the surface ship into one or more flow layouts similar to the above-mentioned submarine, the speed of the ship will be greatly improved, while energy conservation and emission reduction can be achieved. It is especially suitable for application on warships and speedboats.

[0072] In the ocean, the fastest swimming fish are all pointed-mouth fish, such as swordfish and marlin, which can swim at speeds of over 150 km / h. They all have long spikes on their heads, which are very similar to the flow-facing part of the vehicle body of the present invention. The cross-sectional line of their heads is very similar to the waveform of the inverse tangent function. This is the result of long-term evolution in nature and can be considered as an example of the present invention in nature. Therefore, the present invention is also very consistent with the principles of bionics.

[0073] The above description is only a preferred embodiment of the present invention and does not limit the present invention in any way. Any simple modification, change and equivalent variation made to the above embodiment based on the essence of the invention technology shall still fall within the scope of protection of the technical solution of the present invention.

Claims

1. A high-speed aircraft with a flow layout, characterized in that: The invention comprises an aircraft body, wherein the aircraft body is streamlined and symmetrical with respect to a horizontal plane passing through an axis, or symmetrical with respect to a vertical plane passing through the axis, the axis is a line connecting the centers of cross sections of the aircraft body, the axis coincides with the heading of the aircraft, the horizontal length of the aircraft body is greater than the vertical width, the aircraft body is divided into a flow-facing part and a flow-repelling part, the boundary point between the flow-facing part and the flow-repelling part is located at a position where the spanwise dimension of the aircraft body is the largest, the flow-facing part is a cone, a spike is provided at the front end of the flow-facing part, the horizontal or vertical section line of the flow-facing part passing through the axis is two line segments symmetrical with respect to the axis, and both section lines are described by a function with an inverse tangent function as a key factor.

2. A high-speed aircraft with a flow layout according to claim 1, characterized in that: The spikes are fixed or retractable.

3. The high-speed aircraft with a flow layout according to claim 1, characterized in that: The aircraft body is provided with one or more fins.

4. The high-speed aircraft with a flow layout according to claim 1, characterized in that: A beam cover is provided on the periphery of the aircraft body. The beam cover is a thin-walled tubular component. The beam cover is wrapped around the periphery of the aircraft body and keeps a distance from the aircraft body. The beam cover and the aircraft body are fixedly connected by one or more fins.

5. The high-speed aircraft with a flow layout according to claim 1, characterized in that: A propeller is provided at the tail end of the aircraft body.

6. A design optimization method for a high-speed vehicle with a flow layout according to any one of claims 1 to 5, characterized in that: The following steps are involved: S1. Use a function to describe the first section line of the horizontal section of the incoming flow part through the axis: The function describing the first section line of the incoming flow part is ; S2. Introducing width coefficient: In order to control the width of the spacecraft, the width coefficient w is introduced into the function y1, and the result is , the larger w is, the wider the spacecraft body is, the longer the dividing line is; S3. In order to control the length of the spike, the spike length coefficient c is introduced into the function y1 to obtain the function , 0<c<π / 2, the larger the value of c, the longer the spike length; S4. Describe the second section line of the horizontal section of the flow section through the axis: Since the two section lines are symmetrical about the axis of the aircraft body, the description function describing the second section line is the function of the first section line description function flipped along the symmetry axis x-axis, that is, ; S5. Order , solve for the coordinate value x1 of x, and get , x1 is the x-coordinate value of the spike vertex, x2 represents the x-coordinate of the maximum width of the spacecraft body, and if x2 < 0, the width of the spacecraft body is , thus, we get two description functions of the horizontal section line of the flow section passing through the axis, and , the value range of x is, x2 <x<x1; S6. Repeat S1-S5 to obtain the description function of the two profile lines that describe the vertical profile of the frontal part through the axis. The width coefficient w and the spike length coefficient c can be set to different values from the horizontal profile line description function. When the w and c values are the same as those of the horizontal profile description function, the aircraft body is a rotating body.

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