Efficient mute paddle suitable for rotor aircraft

Abstract

The invention provides an efficient mute paddle suitable for a rotor type aircraft, the efficient mute paddle comprises a paddle root area, a paddle inner side area and a paddle main area which are sequentially and continuously arranged, the chord length of the paddle is gradually increased from the head end of the paddle inner side area, and the chord length of the paddle is gradually reduced to a paddle tip area after entering the paddle main area to reach the maximum chord length; by setting the plane shape, the wing type distribution and the torsion distribution of the blades, the compression effect of the advancing blades is reduced to the minimum, and meanwhile, the high-speed impulse noise caused by delocalization can be reduced; due to the sweepforward-sweepback characteristic of the paddle main area, the sharpening characteristic of the tip area and the thin wing type of the tip area, the rotation noise of the rotor wing can be reduced, meanwhile, the paddle vortex interference mode is changed, and the paddle vortex interference noise is reduced; the dynamic droop front edge area periodically deflects when the blade rotates, the bending degree of the blade section airfoil profile can be changed, the dynamic stall of the backward blade is controlled, the aerodynamic performance is improved, and the vibration level is improved.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft design, in particular to a high-efficiency silent blade suitable for rotorcraft. Background Art

[0002] Rotorcraft, with their advantages of vertical takeoff and landing, hovering, and excellent low-altitude and low-speed performance, play an important role in both military and civilian applications. As the core component of a rotorcraft, the rotor system determines its flight performance, including speed, range, noise, vibration, and payload.

[0003] Helicopters are typical representatives of rotary-wing aircraft. When a helicopter is in a hovering state, as the collective pitch increases, the lift coefficient and torsion coefficient increase, the helicopter's hovering efficiency rises to a peak, and then decreases. When the helicopter is in forward flight, the non-uniform inflow formed by the cyclic pitch change, flapping, and distorted wake of the blades causes the blade profile airfoil angle of attack to change significantly with the azimuth angle. When the rotor disc load is high, the retreating blades operate at a larger angle of attack, making complex dynamic stalls more likely to occur. Although the dynamic stall of the rotor can increase the peak lift, it will also cause a sudden increase in drag and torque, and the aerodynamic center of the airfoil will no longer be stable, resulting in a large unstable pitch rod load, leading to rotor vibration problems and limiting the helicopter's flight envelope.

[0004] Rotational noise is generated when the blades rotate, and high-speed impulse noise is generated when the local flow around the leading blade approaches the speed of sound. During low-speed descents or maneuvers, the tip vortices of the rotor blades approach or even collide with other blades, causing severe vortex interference and leading to a series of problems such as aerodynamics, noise, and vibration. Furthermore, phenomena such as blade-wake interference and trailing-edge turbulent separation also contribute to significant broadband noise.

[0005] Therefore, it is necessary to provide a rotor blade for a rotorcraft to improve the aerodynamic performance of the rotor and reduce the noise level and vibration level of the rotor. Summary of the Invention

[0006] In order to solve the above problems in the prior art, the present invention provides a high-efficiency silent blade suitable for rotorcraft, which can improve the aerodynamic performance of the rotor and reduce the noise level and vibration level of the rotor.

[0007] To achieve the above-mentioned purpose, the present invention is implemented through the following technical solutions:

[0008] A high-efficiency silent blade suitable for a rotorcraft, comprising a blade root region, a blade inner region, and a blade main region arranged in sequence, wherein the blade main region includes a blade tip region, and the blade tip region is located at the end of the blade main region;

[0009] The chord length of the blade gradually increases from the beginning of the inner region of the blade, reaches a maximum chord length in the main region of the blade, and then gradually decreases to the end of the tip region of the blade; the blade radius defined from the rotation axis of the blade to the end of the tip region of the blade is R, and the maximum chord length is between 60%R and 90%R;

[0010] The chord length of the blade in the root area of the blade is the root chord length, and the ratio of the root chord length to the maximum chord length is between 0.55 and 0.85; the chord length of the blade in the tip area of the blade is the tip chord length, and the ratio of the tip chord length to the maximum chord length is between 0.1 and 0.75.

[0011] Preferably, the maximum chord length is between 70%R and 85%R.

[0012] Preferably, the ratio of the root chord length to the maximum chord length is between 0.65 and 0.75.

[0013] Preferably, the ratio of the tip chord length to the maximum chord length is between 0.25 and 0.35.

[0014] Preferably, the blade root region extends from its starting position to 30%R.

[0015] Preferably, the main area of the blade extends from 60%R to 100%R.

[0016] Preferably, the chord length of the blade root region is constant.

[0017] Preferably, the chord length of the inner region of the blade is an increasing linear transition or a nonlinear transition.

[0018] Preferably, the chord length of the main area of the blade is a linear transition or a nonlinear transition.

[0019] Preferably, the starting position of the blade tip area is not lower than 80%R.

[0020] Preferably, the outer shape of the blade tip region has a tapered geometric feature and / or a forward-swept geometric feature and / or a backward-swept geometric feature and / or an upward-inverted geometric feature and / or a downward-inverted geometric feature.

[0021] Preferably, the thickness of the airfoil in the blade root region is greater than the thickness of the airfoil in the blade main region, and the thickness of the airfoil in the blade inner region is the transition from the thickness of the airfoil in the blade root region to the thickness of the airfoil in the blade main region.

[0022] Preferably, the airfoil in the blade tip region is a transonic airfoil.

[0023] Preferably, the torsion rate of the portion of the blade located within 30%R is zero, the torsion rate of the portion of the blade located between 30%R and 70%R is negative, the torsion rate of the portion of the blade located between 70%R and 90%R is positive, and the torsion rate of the portion of the blade located between 90%R and 100%R is negative.

[0024] Preferably, the torsion angle of the blade at 70%R is 0°.

[0025] Preferably, a dynamic drooping leading edge is provided in the main region of the blade, the dynamic drooping leading edge is connected to the main region of the blade via a hinge, and the dynamic drooping leading edge is used for periodically deflecting around the hinge rotation axis when the blade rotates.

[0026] Preferably, the dynamic droop leading edge is located between 60%R and 85%R.

[0027] Preferably, the dynamic droop leading edge extends from 70%R to 80%R.

[0028] Preferably, the droop amplitude of the dynamic drooping leading edge is 20° or 30°.

[0029] Preferably, the frequency of the periodic deflection of the dynamic drooping leading edge is 1 or 2 times the rotation frequency of the blade.

[0030] Preferably, on the blade cross section, the hinge is located on the airfoil chord when the dynamically drooping leading edge is not drooped, and the chordwise distance from the blade trailing edge is 75%C or 80%C, where C is the chord length.

[0031] Compared with the prior art, the beneficial effects of the present invention are:

[0032] The present invention minimizes the compression effect of the advancing blade by setting the plane shape, airfoil distribution and torsion distribution, and at the same time reduces the high-speed pulse noise caused by delocalization; the forward-swept-backward characteristics of the main area of the blade, the tapered characteristics of the tip area and the thin airfoil of the tip area can reduce the rotational noise of the rotor, and at the same time change the mode of vortex interference to reduce the vortex interference noise; the dynamic drooping leading edge area periodically deflects when the blade rotates, which can change the curvature of the blade cross-section airfoil, control the dynamic stall of the retreating blade, enhance the aerodynamic performance and improve the vibration level. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 is a schematic diagram of the overall structure of an exemplary vertical take-off and landing rotorcraft using a rotor system including the blade of the present invention;

[0034] Figure 2 1 is a schematic diagram of the planar shape of a high-efficiency silent blade suitable for a rotary-wing aircraft according to the present invention;

[0035] Figure 3 is a schematic diagram of a rear view structure of a tip region having an anhedral angle and / or an anhedral angle in the invention;

[0036] Figure 4 This is a schematic diagram of the airfoil distribution of the high-efficiency silent blade of the present invention applicable to a rotorcraft when the blade is not twisted;

[0037] Figure 5 The invention discloses a torsion angle distribution of a high-efficiency silent blade along the blade span direction for a rotorcraft;

[0038] Figure 6 The thickness distribution of the high-efficiency silent blades of the present invention is applicable to rotorcraft;

[0039] Figure 7 This is a rear view of the high-efficiency silent blade suitable for rotorcraft of the present invention;

[0040] Figure 8 It is an axonometric view of the high-efficiency silent blade suitable for rotorcraft of the present invention;

[0041] Figure 9 It is a cross-sectional schematic diagram of the dynamic drooping leading edge area in the present invention.

[0042] Description of Reference Numerals

[0043] 1-rotorcraft, 2-rotor system, 3-fuselage, 4-tail rotor system, 5-rotor hub, 6-blade, 7-blade cuff, 8-blade root area, 9-blade inner area, 10-blade main area, 11-blade tip area, 11a-anhedral portion, 11b-anhedral portion, 12-end of tip area, 13-leading edge, 14-trailing edge, 15-blade root, 16-dynamic droop leading edge, 17-hinge. DETAILED DESCRIPTION

[0044] In order to make the purpose, technical solutions and advantages of the implementation of the present invention clearer, the technical solutions of the present invention will be described in detail below with reference to the accompanying drawings. In the accompanying drawings, the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. It should be understood here that the described embodiments are partial embodiments of the present invention, not all embodiments. The following embodiments are exemplary and are intended to be used to explain the present invention, and should not be understood as limitations on the present invention. Based on the embodiments of the present invention, other embodiments obtained by ordinary technicians in this field without making creative work are all within the scope of protection of the present invention.

[0045] Figure 1An exemplary vertical take-off and landing (VTOL) rotorcraft 1 is shown that utilizes a rotor system incorporating the blades of the present invention. Rotorcraft 1 includes a fuselage 3 supporting a rotor system 2 and a tail rotor system 4 for balancing anti-torque and controlling heading. While a specific aircraft configuration is shown in this embodiment, other aircraft types, such as coaxial rotor aircraft, tiltrotor aircraft, turboprop aircraft, and eVTOL aircraft, would also benefit from the present invention.

[0046] Rotor system 2 includes a plurality of blades 6 mounted to a rotor hub 5 for rotation about an axis of rotation A. Any number of blades 6 may be used with rotor system 2. Rotor system 2 is coupled to a speed reducer driven by one or more engines.

[0047] See also Figure 2 , a high-efficiency silent blade suitable for rotorcraft, comprising a blade root area 8, a blade inner area 9 and a blade main area 10 arranged in sequence, wherein the blade main area 10 includes a blade tip area 11, and the blade tip area 11 is located at the end of the blade main area 10; the chord length of the blade 6 gradually increases from the head end of the blade inner area 9, enters the blade main area 10 to reach the maximum chord length, and then gradually decreases to the end 12 of the blade tip area; the blade radius defined from the rotation axis A of the blade 6 to the end of the blade tip area 11 is R, and the maximum chord length is located at The chord length of the blade 6 at the blade root area 8 is between 60%R and 90%R, preferably between 65%R and 85%R, and more preferably between 70%R and 85%R; the chord length of the blade 6 at the blade root area 8 is the root chord length, and the ratio of the root chord length to the maximum chord length is between 0.55 and 0.85, preferably between 0.6 and 0.8, and more preferably between 0.65 and 0.75; the chord length of the blade 6 at the blade tip area 11 is the tip chord length, and the ratio of the tip chord length to the maximum chord length is between 0.1 and 0.75, preferably between 0.25 and 0.5, and more preferably between 0.25 and 0.35.

[0048] The chord length of the blade gradually increases from the head end of the inner area of the blade and reaches the maximum chord length when entering the main area of the blade. During the rotation of the blade, the larger chord length area can increase the contact area between the blade and the air, thereby generating greater lift at the same speed. The maximum chord length is between 60%R and 90%R (the better range is 70%R to 85%R), which is located in the middle and outer part of the blade radius. At this position, the linear velocity of the blade is relatively large, and combined with the larger chord length, it can more effectively utilize the high-speed airflow to generate lift, so that the overall lift efficiency of the blade is significantly improved.

[0049] After reaching the maximum chord length, the chord length gradually decreases to the end of the blade tip area. The blade tip has the highest linear velocity, and if the chord length is too large, the air resistance at the blade tip will increase significantly. By reducing the tip chord length, the interaction area between the blade tip and the air can be reduced, thereby reducing air resistance. The ratio of the tip chord length to the maximum chord length is between 0.1 and 0.75 (the optimal range is between 0.25 and 0.35). This ratio setting allows the blade tip to maintain a certain structural strength while effectively reducing resistance and improving the overall aerodynamic efficiency of the blade. Reasonable setting of the blade tip chord length, that is, maintaining an appropriate ratio between the tip chord length and the maximum chord length, helps to reduce the intensity of the tip vortex. The tip vortex is a vortex formed at the blade tip due to the pressure difference between the upper and lower surfaces of the blade and is a significant source of noise. A smaller tip chord length can reduce the pressure difference at the blade tip, thereby weakening the formation of the tip vortex and reducing noise.

[0050] The gradual change in blade chord length from the inner side to the main area and then to the tip provides a more even distribution of airflow around the blade. Uneven airflow creates turbulence and shock waves, which are sources of noise. By properly varying the chord length, airflow flows more smoothly over the blade, reducing airflow disturbances and shock waves, thereby reducing aerodynamic noise.

[0051] The ratio of the root chord length to the maximum chord length is between 0.55 and 0.85 (the optimal range is between 0.65 and 0.75), and the root chord length is relatively small. The root area of the blade mainly serves to connect and transmit torque, and does not require an excessively large chord length to generate lift. A smaller root chord length can reduce the use of material at the root of the blade, thereby reducing the overall weight of the blade. For rotorcraft, reducing weight helps to improve their endurance and flight performance. Although the root chord length is relatively small, the structural strength of the blade root can still be guaranteed through reasonable design and material selection. At the same time, the setting of the maximum chord length can also provide the blade with sufficient load-bearing capacity, ensuring that the blade will not be deformed or damaged when rotating at high speed and bearing large loads, thereby ensuring the reliability and safety of the blade.

[0052] In some optional embodiments, the blade root region 8 extends from its starting position to 30%R.

[0053] In some optional embodiments, the blade main region 10 extends from 60%R to 100%R.

[0054] The blade root area 8 is usually partially or even completely enclosed in the blade sleeve 7 ( Figure 1, thus having relatively little aerodynamic significance. Blade cuff 7 mounts blade 6 to rotor hub 5 by any means known in the art, and therefore no further detailed discussion is required herein. Located between blade root region 8 and blade main region 10 is blade inboard region 9.

[0055] In some optional embodiments, the chord length of the blade root region 8 is constant.

[0056] In some optional embodiments, the chord length of the inner region 9 of the blade is an increasing linear transition or a nonlinear transition.

[0057] In some optional embodiments, the chord length of the blade main region 10 is a linear transition or a nonlinear transition.

[0058] The leading edge 13 and the trailing edge 14 typically have a piecewise straight or curved profile to define a varying blade chord length. The chord length of the blade 6 preferably increases smoothly along a continuous curve from the blade inboard region 9 to a maximum chord length and then tapers smoothly to the end of the blade tip region 11.

[0059] Although the geometry in the figures shows a smooth and continuous transition between the inboard area 9 and the main area 10 of the blade, in reality this transition does not need to be smooth or continuous, nor does it have to be defined by a strict mathematical equation (elliptical, polynomial, or any other equation). In other words, although the examples show a continuous chord length distribution function with a combined forward-sweep and backward-sweep characteristic, the chord length variation can be piecewise linear or arbitrary.

[0060] In some optional embodiments, the blade tip region 11 begins at no less than 80% R. In this embodiment, the blade tip region 11 is located between 95% R and 100% R. The blade tip region 11 can be defined by several design features that distinguish it from the main blade region 10. For example, the transition to a transonic airfoil, a change in twist, and a combination of other geometric features such as forward sweep, backward sweep, anhedral, and dihedral.

[0061] In some optional embodiments, the outer shape of the blade tip region 11 has a tapered geometric feature and / or a forward-swept geometric feature and / or a backward-swept geometric feature and / or an upward-inverted geometric feature and / or a downward-inverted geometric feature.

[0062] exist Figure 2 In the embodiment of the present invention, the blade tip region 11 is swept back. It is worth noting that the blade tip region 11 can taper at both the leading edge 13 and the trailing edge 14, but can also taper independently. Alternatively, the blade tip region 11 can be swept forward or backward. Furthermore, the blade tip region 11 can include an anhedral portion 11a or an anhedral portion 11b. Figure 3 The blade tip region 11 is shown as having an anhedral portion 11a, an anhedral portion 11b, or a combination thereof, as defined when viewed from the trailing edge. It should be understood that various combinations of these tip features may be utilized, such as a linear tip with anhedral, a forward-swept or aft-swept tip with a combination of anhedral and anhedral, and other combinations. Furthermore, the present invention may utilize a continuously varying or piecewise linear variation in the chord length distribution of the blade tip region 11, including forward sweep, aft sweep, anhedral, and anhedral.

[0063] In some optional embodiments, the thickness of the airfoil in the blade root area 8 is greater than the thickness of the airfoil in the blade main area 10, and the thickness of the airfoil in the blade inner area 9 is a transition from the thickness of the airfoil in the blade root area 8 to the thickness of the airfoil in the blade main area 10.

[0064] In some optional embodiments, the airfoil of the blade tip region 11 is a transonic airfoil.

[0065] Please refer to Figure 4 The airfoil profile along the blade's span is shown in sections AA through FF from the blade's root region 8 to the blade's tip region 11. The blade preferably incorporates an airfoil profile that varies from the blade's root region 8 to the blade's inboard region 9, from the blade's inboard region 9 to the blade's main region 10, and finally to the blade's tip region 11. The pitch axis P is the axis along which the blade's airfoil profile angle of attack is varied. Typical blade sections AA through FF taken along the blade's span along the pitch axis P illustrate exemplary airfoils at zero twist. Within the blade's root region 8, particularly section AA at the blade's root 15, the airfoil preferably has a short chord and a large thickness. Further outboard, typically within the blade's inboard region 9, the airfoil preferably has a longer chord and a smaller thickness (sections BB and CC), preferably located between 30%R and 60%R. The blade's main region 10 preferably utilizes an airfoil suitable for medium Mach number operation (sections DD and EE). The blade tip region 11 preferably utilizes a transonic airfoil (profile FF).

[0066] The airfoil's characteristic of short chord and thicker thickness at the blade's root provides excellent structural support for the blade. The blade's root must withstand significant loads such as centrifugal force and bending moment. The greater thickness increases the airfoil's structural strength and rigidity, ensuring stability during high-speed rotation and preventing deformation or damage at the root due to excessive forces. A shorter chord facilitates the design of the connection between the blade's root and components such as the hub. This design results in a more compact connection, facilitating installation and maintenance while also reducing the additional weight and stress concentration associated with complex connections. A longer chord and thinner thickness in the inner blade region relative to the root improves the lift coefficient in this area. The longer chord increases the area where the airflow interacts with the airfoil, while the thinner thickness reduces air resistance and improves the blade's overall aerodynamic efficiency. This airfoil modification helps optimize spanwise load distribution. By adjusting the chord and thickness, the forces in the inner blade region are more balanced, avoiding localized excessive loads and extending the blade's service life. The use of an airfoil suitable for medium Mach number conditions in the main blade region maintains excellent aerodynamic performance. At medium Mach numbers, this airfoil effectively reduces shock wave losses and flow separation, improving the blade's lift-to-drag ratio and ensuring efficient operation for the majority of the blade's operating time. This airfoil ensures excellent performance stability at medium Mach numbers. Even with variations in wind speed or other operating conditions, the blade maintains relatively stable aerodynamic performance, reducing power output fluctuations caused by operating condition fluctuations. The blade tip rotates at a higher speed, making it easier to reach transonic speeds. The use of a transonic airfoil effectively reduces transonic shock wave drag and improves aerodynamic efficiency at the blade tip. Transonic airfoils typically feature a specialized design that better adapts to the complex changes in airflow at transonic speeds, reducing the generation and intensity of shock waves. The use of a transonic airfoil also reduces noise generated at the blade tip. In the transonic state, the generation of shock waves will cause strong noise, while the optimized transonic airfoil can reduce the intensity of shock waves, thereby reducing noise pollution and improving the environmental friendliness of blade operation.

[0067] In some optional embodiments, the torsion rate of the portion of the blade 6 located within 30%R is zero, the torsion rate of the portion of the blade 6 located between 30%R and 70%R is negative, the torsion rate of the portion of the blade 6 located between 70%R and 90%R is positive, and the torsion rate of the portion of the blade 6 located between 90%R and 100%R is negative.

[0068] In some optional embodiments, the torsion angle of the blade 6 at 70% R is 0°

[0069] Reference Figure 5Another feature of blade 6 is the torsion distribution of the blade. The blade preferably combines an unconventional combination of positive and negative torsion rates. That is, from the root of the blade to the area of 30%R, it is preferably present with a constant torsion angle and zero torsion rate. The torsion extending from 30%R to 70%R in the main area of the blade has a negative torsion rate, which may be constant, varied in stages or varied continuously. The torsion extending from 70%R in the main area of the blade to the inner side of 90%R in the tip area of the blade has a positive torsion rate, which may be constant, varied in stages or varied continuously. It transitions to a negative torsion rate at 90%R. The torsion on the tip area of the blade may be positive, negative, constant or a combination.

[0070] The combination of positive and negative twist overcomes the limitations of traditional single negative twist, enabling lift to be distributed on demand within the 30%R-90%R range, improving adaptability across operating conditions (such as low-speed start-up and high-speed cruise). Mid-section negative twist reduces induced drag, tip negative twist suppresses tip vortices, and the positive twist region compensates for outboard lift, enhancing overall aerodynamic efficiency. The positive twist region (70%R-90%R) maintains an effective angle of attack at high speeds or fluctuating incoming airflow, delaying stall and improving helicopter maneuverability. Zero twist at the root reduces structural stress, while the combination of mid-section negative twist and outboard positive twist adjusts aerodynamic load distribution, reducing blade root bending moment and fatigue loads, and extending blade life. A stepwise / continuously variable twist rate design (rather than a constant twist) allows for more precise matching of aerodynamic loads and avoids localized overloads.

[0071] Reference Figure 6 , shows the thickness distribution of blade 6. The ratio of airfoil thickness to chord length is Figure 6 , which generally corresponds to a rear view of the blade 6 ( Figure 7 ). The thickness variation is directly related to the airfoil distribution discussed previously and serves to further illustrate the distribution of the blade structure and aerodynamic characteristics. It is noteworthy that the blade root region 8 generally has a constant thickness.

[0072] refer to Figure 8 and Figure 9 In some optional embodiments, a dynamic drooping leading edge 16 is provided in the main region 10 of the blade, and the dynamic drooping leading edge 16 is connected to the main region 10 of the blade via a hinge 17. The dynamic drooping leading edge 16 is used to periodically deflect around the rotation axis of the hinge 17 when the blade 6 rotates.

[0073] In some optional embodiments, the dynamic droop leading edge 16 is located between 60%R and 85%R, and preferably extends from 70%R to 80%R.

[0074] In some optional embodiments, the drooping amplitude of the dynamic drooping leading edge 16 is preferably 30°, more preferably .

[0075] In some optional embodiments, the frequency of the periodic deflection of the dynamic drooping leading edge 16 is preferably twice the rotation frequency of the blade 6 , and more preferably 1 times.

[0076] In some optional embodiments, on the cross section of the blade 6, the hinge 17 is located on the airfoil chord when the dynamically drooping leading edge 16 is not drooping, and the chordwise distance from the blade trailing edge 14 is 75%C or 80%C, where C is the chord length.

[0077] Figure 9 A cross-section of a dynamic drooping leading edge 16 is shown. The shape of the dynamic drooping leading edge 16 when not drooping is indicated by a dashed line, and the shape of the dynamic drooping leading edge 16 when drooped is indicated by a solid line. Hinge 17 is located on the chord of the airfoil when the dynamic drooping leading edge 16 is not drooping, preferably at a chordwise distance of 80%C from the trailing edge 14, and more preferably at a chordwise distance of 75%C from the trailing edge, where C is the chord length. As blade 6 rotates, dynamic drooping leading edge 16 periodically deflects about hinge 17, changing the camber of the blade's cross-sectional airfoil. The deflection amplitude is preferably 30°, more preferably 20°, and the deflection frequency is preferably twice the blade's rotation frequency, and more preferably equal to the blade's rotation frequency. The mechanical structure of the dynamic drooping leading edge 16 and hinge 17 can be implemented by any means known in the art, and therefore, no further detailed structural discussion is required herein.

[0078] Setting the dynamic droop leading edge can make the blade actively change the curvature and adjust the lift in real time. Figure 9 The leading edge is deflected downward to increase the camber of the airfoil, which is equivalent to increasing the effective angle of attack and improving the local lift coefficient, which is suitable for the working condition of the retreating blade in the stall zone. Figure 9The dynamic droop leading edge can delay separation by changing the leading edge curvature when the airflow is about to separate (such as in the high angle of attack region), thereby widening the stall margin. In addition, when the rotor is rotating at high speed, dynamic camber adjustment is performed in the transonic region to reduce the interaction between the shock wave and the boundary layer, thereby reducing wave drag. As the blade rotates, the aerodynamic force varies with the azimuth angle. The dynamic droop leading edge deflects at the rotation frequency or its multiples to actively offset load fluctuations: when the blade enters the forward region from the trailing region, the leading edge periodically resets, suppressing the flapping load caused by the sudden increase in lift and reducing root fatigue stress. The passive torsion or flapping deformation of traditional blades has a response lag. The dynamic droop leading edge, as an active control surface, can provide damping by anti-phase deflection before aeroelastic vibrations (such as flutter and flapping) occur, thereby increasing the critical flutter speed and improving stability. The dynamic droop leading edge enables the blade to operate efficiently under multiple operating conditions. In low-speed and low-velocity scenarios, the leading edge is drooped throughout the entire cycle to maximize lift / thrust and improve hovering efficiency. In high-speed scenarios (such as forward flight and strong winds), the leading edge is drooped only at specific azimuth angles to avoid local overload while maintaining overall efficiency.

[0079] The high-efficiency silent blades of this embodiment, which are suitable for rotorcraft, minimize the compression effect of the forward blades by selecting the plane shape, airfoil distribution and torsion distribution, and at the same time reduce the high-speed pulse noise caused by delocalization; the forward-swept-backward characteristics of the main area of the blades, the tapered characteristics of the tip area and the thin airfoil of the tip area can reduce the rotational noise of the rotor, while changing the mode of vortex interference and reducing vortex interference noise; the dynamic droop leading edge area deflects periodically when the blades rotate, changing the curvature of the blade cross-section airfoil, controlling the dynamic stall of the backward blades, improving aerodynamic performance and vibration levels. Increasing the solidity of the rotor maximizes the performance efficiency of the design point while maintaining sufficient control margin. The definition of design parameters is to ensure that the aeroelastic requirements of the blades are met. These requirements are determined through independent structural dynamics and aeroelastic analysis and consideration of factors such as processing and manufacturing.

[0080] Those skilled in the art will appreciate that the various blade characteristics disclosed herein may be used individually or in any combination, depending on the specific design requirements of the aircraft. Furthermore, while the present invention is generally described in conjunction with single-rotor helicopters, the present invention is equally applicable to any aircraft, including but not limited to coaxial rotor aircraft, tiltrotor aircraft, turboprop aircraft, and eVTOL aircraft.

[0081] It should be understood that relative position terms such as "forward", "backward", "upper", "lower", "above", "below", "inside", "outside", etc. refer to the normal operating posture of the aircraft and should not be regarded as having other limitations.

[0082] Although a particular sequence of steps is described above, it should be understood that the steps may be performed in any order, separated or combined unless otherwise indicated, and still benefit from the present invention.

[0083] The above is only one specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any simple replacement or modification within the scope of the technical concept disclosed by the present invention and the technical solution of the present invention shall be within the scope of protection of the present invention. Therefore, the scope of protection of the present invention shall be based on the scope of protection of the claims.

Claims

1. A high-efficiency silent blade suitable for rotorcraft, characterized in that: The blade root region (8), the blade inner region (9) and the blade main region (10) are arranged in sequence, wherein the blade main region (10) includes a blade tip region (11), and the blade tip region (11) is located at the end of the blade main region (10); the blade radius defined from the rotation axis of the blade (6) to the end (12) of the blade tip region is R, the blade root region (8) extends from its starting position to 30%R, the blade main region (10) extends from 60%R to 100%R, and the starting position of the blade tip region (11) is not less than 80%R; The chord length of the blade (6) gradually increases from the head end of the blade inner region (9), reaches the maximum chord length in the blade main region (10), and then gradually decreases to the end of the blade tip region (11), wherein the maximum chord length is between 60%R and 90%R; The chord length of the blade (6) located in the blade root region (8) is the root chord length, and the ratio of the root chord length to the maximum chord length is between 0.55 and 0.85; the chord length of the blade (6) located in the blade tip region (11) is the tip chord length, and the ratio of the tip chord length to the maximum chord length is between 0.1 and 0.

75.

2. The high-efficiency silent blade suitable for a rotorcraft according to claim 1, characterized in that: The maximum chord length is between 70%R and 85%R, the ratio of the root chord length to the maximum chord length is between 0.65 and 0.75, and the ratio of the tip chord length to the maximum chord length is between 0.25 and 0.

35.

3. The high-efficiency silent blade suitable for a rotorcraft according to claim 1, characterized in that: The chord length of the inner region (9) of the blade is an increasing linear transition or a nonlinear transition; The chord length of the blade main region (10) is a linear transition or a nonlinear transition.

4. The high-efficiency silent propeller blade suitable for a rotorcraft according to claim 1, characterized in that: The outer shape of the blade tip region (11) has a tapered geometric feature and / or a forward-swept geometric feature and / or a backward-swept geometric feature and / or an upward-inverted geometric feature and / or a downward-inverted geometric feature.

5. The high-efficiency silent propeller blade suitable for rotorcraft according to claim 1, characterized in that: The thickness of the airfoil in the blade root region (8) is greater than the thickness of the airfoil in the blade main region (10), the thickness of the airfoil in the blade inner region (9) is a transition from the thickness of the airfoil in the blade root region (8) to the thickness of the airfoil in the blade main region (10), and the airfoil in the blade tip region (11) is a transonic airfoil.

6. The high-efficiency silent propeller blade suitable for a rotorcraft according to claim 1, characterized in that: The torsion rate of the portion of the blade (6) located within 30%R is zero, the torsion rate of the portion of the blade (6) located between 30%R and 70%R is negative, the torsion rate of the portion of the blade (6) located between 70%R and 90%R is positive, the torsion rate of the portion of the blade (6) located between 90%R and 100%R is negative, and the torsion angle of the blade (6) located at 70%R is 0°.

7. The high-efficiency silent propeller blade suitable for a rotorcraft according to claim 1, characterized in that: A dynamic drooping leading edge (16) is provided in the main region (10) of the blade. The dynamic drooping leading edge (16) is located between 60%R and 85%R. The dynamic drooping leading edge (16) is connected to the main region (10) of the blade via a hinge (17). The dynamic drooping leading edge (16) is used to periodically deflect around the rotation axis of the hinge (17) when the blade (6) rotates.

8. The high-efficiency silent propeller blade suitable for a rotorcraft according to claim 7, characterized in that: The dynamic droop leading edge (16) extends from 70%R to 80%R.

9. The high-efficiency silent propeller blade suitable for a rotorcraft according to claim 7, characterized in that: The drooping amplitude of the dynamic drooping leading edge (16) is 20° or 30°; The frequency of the periodic deflection of the dynamic drooping leading edge (16) is 1 or 2 times the rotation frequency of the blade.

10. The high-efficiency silent blade suitable for rotorcraft according to claim 7, characterized in that: On the cross section of the blade (6), the hinge (17) is located on the airfoil chord when the dynamically drooping leading edge (16) is not drooping, and the chordwise distance from the blade trailing edge (14) is 75%C or 80%C, where C is the chord length.

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