Rotating head structure for reducing resistance by differential pressure and friction of aircraft

Abstract

The invention discloses a rotating head structure for reducing the pressure difference and friction resistance of an aircraft. The rotating head structure comprises a turbine disc body and blades. The turbine disc body comprises a hub head tip and a hub which are integrally formed; the hub head tip is located at the front end of the hub, and the rear end of the hub is provided with a wheel disc matching face. The multiple blades are installed on the outer wall of the hub in the circumferential direction at equal intervals. The blade comprises a blade front edge surface, a blade rear edge surface, a blade pressure surface and a blade suction surface; the blade front edge surface is close to the tip of the hub head, and the blade rear edge surface is close to the wheel disc matching surface; wherein the hub is of a rotary body structure formed by rotation of a concave arc; the blades are spirally and gradually expanded along the outer wall of the hub, and the heights of the blades and the distance between every two adjacent blades are sequentially increased in the spiral direction. According to the invention, the differential pressure resistance and frictional resistance borne by the aircraft can be obviously reduced, the thermal load is correspondingly reduced, and the aerodynamic performance of the aircraft is effectively improved.

Description

Technical Field

[0001] The present invention relates to the technical field of navigation equipment, in particular to a rotating head configuration for reducing pressure difference and friction drag of an aircraft. Background Art

[0002] When many navigation equipment such as aerial vehicles (such as airplanes, missiles, rockets), underwater vehicles (such as submarines), rail trains, and vehicles are in motion, there is fluid resistance that hinders navigation. The existence of resistance makes powerful propulsion devices necessary, and the greater the resistance, the lower the upper speed limit. The existence of resistance can be roughly divided into two forms: pressure difference resistance and friction resistance. After the existing aircraft has been optimized through the fluid shape design, the only available drag reduction technologies are surface treatment (grooves, bionic skin), installation of vortex generators to delay separation, etc., which can only reduce friction resistance to a limited extent.

[0003] For high-speed aircraft, such as hypersonic aircraft, there is also the problem of severe aerodynamic heating at the head. Hypersonic aircraft often use blunt bodies to reduce drag and heat, but blunt bodies will produce bow shock waves at the head due to the high compression of the oncoming airflow at high Mach numbers, which will in turn produce very high pressure at the front, resulting in high aerodynamic drag during atmospheric flight. Even if the blunt body can reduce aerodynamic heating to a certain extent, the strong heat load at the head of the aircraft is still very large, which will cause ablation of the aircraft surface material and even failure or damage of internal electronic equipment.

[0004] At present, the main methods to reduce the resistance of the aircraft head in this situation are: installing needle-shaped objects at the front of the aircraft, reverse jet and energy deposition. Installing needle-shaped objects is passive flow control, and the local heat flux density is extremely high, which is very easy to ablate; reverse jet belongs to active flow control, and the applicable angle of attack range is small, which may erode the aircraft surface and the engine inlet; in energy deposition, the laser plasma has little interference with the overall flow field, but the power requirement of the laser transmitter is high. In summary, there is still a lot of room for development in designing new heat reduction and drag reduction devices for the head of hypersonic aircraft.

[0005] Therefore, a rotating head configuration for reducing pressure difference and friction drag of the vehicle is proposed. Summary of the invention

[0006] The object of the present invention is to provide a rotating head configuration for aircraft pressure difference and friction drag reduction, aiming to solve or improve at least one of the above-mentioned technical problems.

[0007] To achieve the above-mentioned purpose, the present invention provides the following solution: The present invention provides a rotating head configuration for aircraft pressure difference and friction drag reduction, characterized in that it includes:

[0008] A turbine disc body, the turbine disc body comprising an integrally formed wheel hub head tip and a wheel hub; the wheel hub head tip is located at the front end of the wheel hub, and the rear end of the wheel hub is provided with a wheel disc mating surface;

[0009] Blades, wherein a plurality of blades are provided, and the plurality of blades are circumferentially and evenly spaced and mounted on the outer wall of the hub; the blades include a leading edge surface, a trailing edge surface, a pressure surface and a suction surface; the leading edge surface of the blade is arranged close to the tip of the hub head, and the trailing edge surface of the blade is arranged close to the mating surface of the wheel disc;

[0010] Among them, the hub is a rotating body structure formed by rotating an inwardly concave arc; a plurality of blades are arranged along the outer wall of the hub in a spiral and gradually expanding manner, and the height of the blades and the distance between two adjacent blades increase successively along the spiral direction.

[0011] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the tip of the hub head is configured as a rounded corner.

[0012] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the wheel disc mating surface is a plane, the edge of the wheel disc mating surface is integrally processed with the rear end edge of the hub, and the wheel disc mating surface is used to connect the shaft body.

[0013] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the center of the wheel disc mating surface is taken as the coordinate origin, and the x direction, y direction and z direction respectively form coordinate axes to establish a rectangular coordinate system;

[0014] The roulette mating surface is the plane where the y-coordinate axis and the z-coordinate axis are located, and the concave arc is located in the plane where the x-coordinate axis and the z-coordinate axis are located; the two end points of the concave arc are set as point A and point B respectively, the coordinate origin is set as point O, the distance from point O to point B is the radius r of the roulette mating surface, and the distance from point O to point A along the x-axis is d;

[0015] The calculation formula of the concave arc is z=r*tanh(k*x)+r, (-d <x<0);

[0016] Where: k is defined as the shape factor, which represents the inclination of the concave arc, 0.06 <k<0.17;0.75<r / d<1.3。

[0017] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the shape of the leading edge surface of the blade is rectangular, and the leading edge surface of the blade is arranged parallel to the plane where the x-coordinate axis and the z-coordinate axis are located;

[0018] The length of the leading edge of the blade is h 1 ,0.04 <h 1 / r<0.12.

[0019] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the shape of the blade trailing edge surface is rectangular, and the length of the blade trailing edge surface is h 2 ,0.15 <h 2 / r<0.4, the width of the trailing edge of the blade is d 2 ,d 2 The value range is 0.5~1.5mm.

[0020] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the angle between the trailing edge surface of the blade and the plane where the x-coordinate axis and the z-coordinate axis are located is α, 40°<α<60°.

[0021] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the number of the blades is not less than six.

[0022] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the blade suction surface is an inner concave surface, and the blade pressure surface is an outer convex surface.

[0023] According to the rotating head configuration for aircraft pressure difference and friction drag reduction provided by the present invention, the radius of the fillet is 4 mm to 10 mm.

[0024] The present invention discloses the following technical effects:

[0025] The present invention arranges a plurality of blades at equal intervals along the circumference of the hub surface of the hub, and the height of the blades and the interval between two adjacent blades increase successively along the spiral direction. At the airflow inlet, the blade height is relatively small, which can quickly guide the airflow into the turbine disk and reduce the deflection and separation of the airflow. Then, the blade height gradually increases, so that the airflow maintains a stable and uniform flow, reduces its obstruction and reverse flow, and meets the dynamic change requirements of high-speed airflow; at the same time, as the blade height increases, the increase in the flow channel width also effectively reduces the pressure drop in the flow channel, so that the airflow is more stable when passing through the blades, reduces the rapid compression and turbulence of the airflow, and reduces the energy loss caused by the pressure drop. By optimizing the fluid distribution in the flow channel and reducing the pressure drop, the drag reduction effect of the aircraft can be significantly improved, and the aerodynamic performance can be improved;

[0026] The present invention adopts the specific design and mutual cooperation of the rotating configuration of the turbine disk body and the blades, and the blades are arranged in a spiral and gradually expanding manner along the outer wall of the hub. The centrifugal force generated by the rotation of the turbine disk directly throws out the fluid that hits the aircraft head-on laterally, forming a negative pressure space at the head, directly reducing the pressure load of the head and the heat load of the high-speed aircraft. In high-speed aircraft such as hypersonic aircraft, this negative pressure area can effectively alleviate the aerodynamic heating problem and avoid the risks of material ablation and internal equipment damage; the ejected fluid forms a swirling vortex at the rear body of the aircraft, which makes the flow on the surface of the aircraft fuselage laminar to a certain extent, effectively reducing the friction resistance on the rear body surface, thereby significantly reducing the pressure difference resistance and friction resistance of the aircraft. For supersonic aircraft, the compressed high-temperature gas encountered by the head is radially ejected, which can also significantly reduce the heat load, achieve a significant drag reduction effect, and effectively improve the aerodynamic performance of the aircraft;

[0027] The rotating head configuration of the present invention can be added to the head of an existing aircraft or directly replace the original aircraft head, generating a large negative pressure area on the head to achieve considerable pressure difference drag reduction; a number of blades are arranged in a spiral expansion along the outer wall of the hub, so that the downstream swirling vortex is wrapped around the aircraft fuselage, isolating the fuselage from the high-speed incoming flow, and effectively reducing friction resistance; the swirling vortex has the potential to laminarize the flow on the surface of the fuselage, thereby causing laminar drag reduction, and can be widely used in aircraft in the air, underwater and on the ground. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0029] Figure 1 It is a structural schematic diagram of the present invention;

[0030] Figure 2 Schematic diagram of the concave arc and the rectangular coordinate system in the present invention;

[0031] Figure 3 It is a front view of the present invention;

[0032] Figure 4 It is a left side view of the present invention;

[0033] Figure 5 It is a simulation streamline diagram of the head rotation configuration of the present invention;

[0034] Figure 6 The head configuration of the present invention has a drag coefficient of Ma=0.3, and the positive rotation is a rotation mode diagram that does not require motor drive;

[0035] Figure 7 The head configuration of the present invention has a drag coefficient of Ma=0.7, and the positive rotation is a rotation mode diagram that does not require motor drive;

[0036] Figure 8 The simulated pressure cloud diagram of the head configuration of the present invention at a drag coefficient of Ma=1.2 and without blades;

[0037] Fig. 9 The simulated pressure cloud diagram of the head configuration of the present invention under the resistance coefficient of Ma=1.2 under rotation;

[0038] Fig.10 This is a simulated pressure cloud diagram without blades;

[0039] Fig.11 This is the simulated pressure cloud diagram under positive rotation;

[0040] Among them, 1. The tip of the hub head; 2. The hub; 3. The mating surface of the wheel disc; 4. The leading edge surface of the blade; 5. The trailing edge surface of the blade; 6. The pressure surface of the blade; 7. The suction surface of the blade. DETAILED DESCRIPTION

[0041] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0042] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0043] Reference Figure 1-Figure 11 The present invention provides a rotating head configuration for aircraft pressure difference and friction drag reduction, including:

[0044] The turbine disc body includes a hub head tip 1 and a hub 2 which are integrally formed; the hub head tip 1 is located at the front end of the hub 2, and the rear end of the hub 2 is provided with a wheel disc mating surface 3;

[0045] The blades are provided with a plurality of blades, which are circumferentially and evenly spaced and mounted on the outer wall of the hub 2; the blades include a leading edge surface 4, a trailing edge surface 5, a pressure surface 6 and a suction surface 7, and a three-dimensional shape of the blades is formed by spatial curve lofting; the leading edge surface 4 of the blades is arranged close to the tip 1 of the hub head, and the trailing edge surface 5 of the blades is arranged close to the mating surface 3 of the wheel disc; the blades adopt a lofting form combined with the spatial curve, so that the geometric shape of the blades on the windward surface can smoothly guide the airflow;

[0046] The hub 2 is a rotating body structure formed by rotating an inwardly concave arc; a plurality of blades are arranged along the outer wall of the hub 2 in a spiral and gradually expand, and the height of the blades and the distance between two adjacent blades are increased in sequence along the spiral direction, and the increase rate is flexible and adjustable. The specific form can be determined according to the requirements of fluid dynamics, covering linear, nonlinear or other increasing rules;

[0047] Compared with the conventional fixed head configuration, the present invention can throw the oncoming flow laterally and form a vortex on the rear body, effectively and significantly reducing the friction resistance and pressure difference resistance, and improving the overall aerodynamic performance of the aircraft.

[0048] The present invention arranges a plurality of blades at equal intervals along the circumferential direction of the hub surface of the hub 2. The height of the blades and the interval between two adjacent blades increase successively along the spiral direction. At the airflow inlet, the blade height is relatively small, which can quickly guide the airflow into the turbine disk and reduce the deflection and separation of the airflow. Then, the blade height gradually increases, so that the airflow maintains a stable and uniform flow, reduces its obstruction and reverse flow, and meets the dynamic change requirements of high-speed airflow. At the same time, as the blade height increases, the increase in the flow channel width also effectively reduces the pressure drop in the flow channel, so that the airflow is more stable when passing through the blades, reduces the rapid compression and turbulence of the airflow, and reduces the energy loss caused by the pressure drop. By optimizing the fluid distribution in the flow channel and reducing the pressure drop, the drag reduction effect of the aircraft can be significantly improved, and the aerodynamic performance can be improved.

[0049] The present invention adopts the specific design and mutual cooperation of the rotating configuration turbine disc body and blades, and the blades are arranged in a spiral and gradually expanding manner along the outer wall of the hub 2. The centrifugal force generated by the rotation of the turbine disc directly throws out the fluid that hits the aircraft head-on laterally, forming a negative pressure space at the head, directly reducing the pressure load of the head and the heat load of the high-speed aircraft. In high-speed aircraft such as hypersonic aircraft, this negative pressure area can effectively alleviate the aerodynamic heating problem and avoid the risks of material ablation and internal equipment damage; the ejected fluid forms a swirling vortex at the rear body of the aircraft, which makes the flow on the surface of the aircraft fuselage laminar to a certain extent, effectively reducing the friction resistance on the rear body surface, thereby significantly reducing the pressure difference resistance and friction resistance of the aircraft. For supersonic aircraft, the compressed high-temperature gas encountered by the head is radially ejected, which can also significantly reduce the heat load, achieve a significant drag reduction effect, and effectively improve the aerodynamic performance of the aircraft;

[0050] The rotating head configuration of the present invention can be added to the head of an existing aircraft or directly replace the original aircraft head, generating a large negative pressure area on the head to achieve considerable pressure difference drag reduction; a number of blades are arranged in a spiral expansion along the outer wall of the hub, so that the downstream swirling vortex is wrapped around the aircraft fuselage, isolating the fuselage from the high-speed incoming flow, and effectively reducing friction resistance; the swirling vortex has the potential to laminarize the flow on the surface of the fuselage, thereby causing laminar drag reduction, and can be widely used in aircraft in the air, underwater and on the ground.

[0051] As a further optimization solution, the tip 1 of the hub head is set to be rounded, so that the airflow can smoothly transition when flowing through the tip 1 of the hub head, reducing turbulence and local high pressure caused by sharp corners, alleviating aerodynamic heating (high-speed aircraft), and effectively reducing the chance of pressure difference to a certain extent.

[0052] Further optimizing the scheme, the wheel disc mating surface 3 is a plane, the edge of the wheel disc mating surface 3 and the rear end edge of the hub 2 are processed in one piece, and the wheel disc mating surface 3 is used to connect the shaft body;

[0053] The wheel disc mating surface 3 is the key surface for connecting the wheel hub 2 with the rear body. The wheel disc mating surface 3 is in close contact with the edge of the hub surface of the wheel hub 2 and is processed as a whole, thereby ensuring the rotational stability of the turbine disc.

[0054] Further optimizing the scheme, taking the center of the wheel mating surface 3 as the coordinate origin, the x direction, y direction and z direction form coordinate axes respectively, and establishing a rectangular coordinate system;

[0055] The roulette mating surface 3 is the plane where the y-coordinate axis and the z-coordinate axis are located, and the concave arc is located in the plane where the x-coordinate axis and the z-coordinate axis are located; the two end points of the concave arc are set as point A and point B respectively, the coordinate origin is set as point O, the distance from point O to point B is the radius r of the roulette mating surface, and the distance from point O to point A along the x-axis is d;

[0056] The calculation formula of the concave arc is z = r*tanhk*x+r, -d <x<0;

[0057] Where: k is defined as the shape factor, which represents the inclination of the concave arc, 0.06 <k<0.17;0.75<r / d<1.3;

[0058] The wheel hub 2 is a rotating body structure formed by rotating an inward concave arc, and the wheel hub surface of the wheel hub 2 is stretched by rotating a curve of a hyperbolic tangent function, but is not limited to such mathematical expressions, and all curves that satisfy the arc-shaped change (such as exponential curves, parabolas, ellipses, etc.) are within the protection range;

[0059] Thus, the hub surface of the hub 2 has a specific curvature distribution, which can effectively deflect the inflow direction of the fluid, and optimize the fluid pressure gradient by reasonably selecting the curve parameters and definition domain;

[0060] The present invention utilizes a hyperbolic tangent function curve to rotate and stretch to form the hub surface of the hub 2, and adopts two rectangular surfaces to form blades on the disk along a space curve, and combines the disk matching surface to realize the purpose of turbine disk rotation.

[0061] According to a further optimization scheme, the shape of the blade leading edge surface 4 is rectangular, and the blade leading edge surface 4 is arranged parallel to the plane where the x-coordinate axis and the z-coordinate axis are located; the starting point of the blade leading edge surface 4 is at a position d shifted backward (15%-50%) from the tip 1 of the hub head along the positive direction of the x-axis;

[0062] The leading edge surface 4 of the blade needs to bear the airflow load. Within a certain thickness range, the bending stiffness of the rectangular cross section is relatively balanced, which can effectively disperse the pressure and shear force generated by the airflow and avoid local overload or buckling.

[0063] The length of the blade leading edge surface 4 is h 1 ,0.04 <h 1 / r<0.12, the width of the blade leading edge surface 4 is d 1 , according to different aircraft application scenarios, such as missiles and launch vehicles, the relevant parameters d 1 The value range is between 0.5 and 1.5 mm.

[0064] Further optimization scheme, the shape of the blade trailing edge surface 5 is rectangular, and the length of the blade trailing edge surface 5 is h 2 ,0.15 <h 2 / r<0.4, the width of the blade trailing edge surface 5 is d 2 ,d 2 The value range of is 0.5~1.5mm; In this embodiment, the length h of the blade trailing edge surface 5 2 Preferably 55 mm, width d 2 Preferably 1 mm.

[0065] According to a further optimization scheme, the included angle between the blade trailing edge surface 5 and the plane where the x-coordinate axis and the z-coordinate axis are located is α, 40°<α<60°; in this embodiment, the included angle α is preferably 50°.

[0066] According to a further optimization scheme, the number of blades is not less than six; in this embodiment, the number of blades is preferably nine.

[0067] To further optimize the solution, the suction surface 7 of the blade is a smooth concave surface, and the pressure surface 6 of the blade is a smooth convex surface, so that the airflow can flow along the surface of the blade when passing through, forming a stable laminar airflow area, thereby reducing the friction between the blade surface and the airflow and reducing frictional resistance.

[0068] According to the further optimization scheme, the radius of the fillet is 4 mm to 10 mm.

[0069] Figure 10-11 It can be seen that the simulated pressure cloud diagram of the head configuration of the present invention at Ma=1.2, in the forward mode, the drag reduction rate can reach 57%; specifically, Fig.10In the absence of blades, there is a significant shock wave near the head configuration, and the positive rotation of the blades can effectively weaken this shock wave, thereby significantly reducing the overall drag coefficient.

[0070] In the description of the present invention, it should be understood that the terms "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside" and "outside" etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the present invention.

[0071] Obviously, the above embodiments of the present invention are merely examples for clearly explaining the present invention, and are not intended to limit the implementation methods of the present invention. For ordinary technical users in the relevant field, other different forms of changes or modifications can be made on the basis of the above description. It is not necessary and impossible to list all the implementation methods here. Any modifications, equivalent substitutions and improvements made within the spirit and principles of the present invention should be included in the protection scope of the claims of the present invention.

Claims

1. A rotating head configuration for aircraft pressure difference and friction drag reduction, characterized in that: Comprising: A turbine disk body, the turbine disk body includes an integrally formed hub head tip (1) and a hub (2); the hub head tip (1) is located at the front end of the hub (2), and a disk mating surface (3) is provided at the rear end of the hub (2); Blades, a plurality of the blades are circumferentially and equally spaced and installed on the outer wall of the hub (2); the blades include a blade leading edge surface (4), a blade trailing edge surface (5), a blade pressure surface (6) and a blade suction surface (7); the blade leading edge surface (4) is arranged close to the hub head tip (1), and the blade trailing edge surface (5) is arranged close to the disk mating surface (3); Wherein, the hub (2) is a rotating body structure formed by rotating an inward concave arc; a plurality of the blades are arranged in a spiral and gradually expanding manner along the outer wall of the hub (2), and the height of the blades and the spacing between two adjacent blades both increase sequentially along the spiral direction.

2. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 1, characterized in that: The hub head tip (1) is provided with a rounded corner.

3. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 1, characterized in that: The disk mating surface (3) is a plane, the edge of the disk mating surface (3) is integrally machined with the rear end edge of the hub (2), and the disk mating surface (3) is used to connect a shaft body.

4. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 1, characterized in that: Taking the center of the disk mating surface (3) as the coordinate origin, the x-direction, y-direction and z-direction respectively form coordinate axes to establish a rectangular coordinate system; The disk mating surface (3) is the plane where the y-axis and z-axis are located, and the inward concave arc is located in the plane where the x-axis and z-axis are located; the two end points of the inward concave arc are respectively set as point A and point B, the coordinate origin is set as point O, the distance from point O to point B is the radius r of the disk mating surface, and the distance from point O to point A along the x-axis is d; The calculation formula of the inward concave arc is z = r*tanh(k*x)+r, (-d < x < 0); Wherein: k is defined as a shape factor, indicating the inclination degree of the inward concave arc, 0.06 < k < 0.17; 0.75 < r / d < 1.

3.

5. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 4, characterized in that: The shape of the blade leading edge surface (4) is rectangular, and the blade leading edge surface (4) is arranged parallel to the plane where the x-axis and z-axis are located; The length of the blade leading edge surface (4) is h1, 0.04 < h1 / r < 0.

12.

6. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 1, characterized in that: The shape of the blade trailing edge surface (5) is rectangular, the length of the blade trailing edge surface (5) is h2, 0.15 < h2 / r < 0.4, and the width of the blade trailing edge surface (5) is d2, and the value range of d2 is 0.5 to 1.5 mm.

7. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 4, characterized in that: The included angle between the blade trailing edge surface (5) and the plane where the x-axis and z-axis are located is α, 40° < α < 60°.

8. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 1, characterized in that: The number of the blades is not less than six.

9. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 1, characterized in that: The blade suction surface (7) is an inward concave surface, and the blade pressure surface (6) is an outward convex surface.

10. The rotating head configuration for aircraft pressure difference and friction drag reduction according to claim 2, characterized in that: The radius of the rounded corner is 4 mm to 10 mm.

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