Blade noise reduction method, blade tip assembly, rotor and rotorcraft

Abstract

The present invention discloses a blade noise reduction method, a blade tip assembly, a rotor and a rotorcraft, and relates to the technical field of rotorcraft noise reduction equipment. The blade noise reduction method comprises step S1, processing the rotorcraft blade that needs noise reduction into a basic blade; step S2, evaluating the type of aerodynamic noise with the greatest impact in the entire flight envelope according to the flight conditions of the rotorcraft, and determining the type of blade tip assembly that can reduce the aerodynamic noise; step S3, designing specific parameters of the blade tip assembly; step S4, installing the blade tip assembly on the basic blade to form a complete blade; the basic blade is the remaining part of the blade after the blade tip is removed; the blade tip assembly is used to be installed in the original blade tip area of ​​the blade to form a new blade tip of the basic blade; the present invention manufactures the blade tip assembly through the above method, which can reduce the aerodynamic noise with the greatest impact in the flight envelope of the rotorcraft, without the need to design, process and install the entire blade, thus saving the cost of design, processing and installation.

Description

Technical Field

[0001] The present invention relates to the technical field of rotorcraft noise reduction equipment, and in particular to a blade noise reduction method, a blade tip assembly, a rotor and a rotorcraft. Background Art

[0002] High-speed impulse noise (HSN) is a type of noise primarily produced during the flight of rotorcraft. When HSN occurs, its sharp, explosive sound can cause significant acoustic pollution in the area where it propagates. Therefore, reducing blade noise has become a pressing technical challenge for those skilled in the art. Summary of the Invention

[0003] The object of the present invention is to provide a blade noise reduction method, a blade tip assembly, a rotor and a rotorcraft, so as to effectively reduce the noise generated by the blade.

[0004] To achieve the above object, the present invention provides the following solutions:

[0005] The present invention provides a method for reducing blade noise, comprising the following steps:

[0006] Step S1, processing the rotorcraft blades that need noise reduction into basic blades;

[0007] Step S2: evaluating the noise type with the greatest impact in the entire flight envelope based on the flight conditions of the rotorcraft, and determining a type of blade tip assembly capable of reducing the noise of that type;

[0008] Step S3, designing specific parameters of the blade tip assembly;

[0009] Step S4, installing the tip assembly on the base blade to form a complete blade;

[0010] The basic blade is the remaining portion of the blade after the blade tip is removed; the tip assembly is used to be installed in the original tip area of ​​the blade to form a new tip of the basic blade.

[0011] Preferably, the step S2 includes:

[0012] Step 1: Using the blade flow field solver module in the CLORNS series software, solve the flow field characteristics of the blade under calculation conditions to obtain the sound source information;

[0013] Step 2: Using the observation point information and the sound source information obtained in step 1 as input, the rotor noise characteristic calculation module in the CLORNS series software performs calculations to obtain the aeroacoustic characteristics of the blade under the corresponding calculation conditions.

[0014] Step ③, analyzing the aeroacoustic characteristics to obtain the components of aerodynamic noise in the entire flight envelope and the magnitude of each component, and determining the required blade tip assembly;

[0015] The calculation conditions refer to taking the flight conditions of the entire flight envelope as input to the blade flow field solution module in the CLORNS series software; and the observation point information refers to the location of the observed sound.

[0016] Preferably, step S3 includes:

[0017] In the first step, characteristic points on the blade tip assembly that can reflect the overall profile of the blade tip are selected as design variables;

[0018] The second step is to determine the mathematical function for constructing the lines between the feature points;

[0019] The third step is to impose size range constraints on the design variables;

[0020] The fourth step is to determine the optimization constraints and optimization goals;

[0021] Step 5: Input all the parameters obtained in Steps 1 to 4 into the CLORNS series software optimization module to obtain the design variables that meet the standards, so as to determine the processing parameters of the blade tip assembly;

[0022] The constraint condition is that the aerodynamic performance of the blade does not decrease compared to the original blade, and the optimization goal is to minimize the aerodynamic noise sound pressure level at the observation point.

[0023] Preferably, the blade includes a root cutting section, a basic blade section and a tip assembly section which are sequentially arranged in a direction away from the center of the hub, and the basic blade section constitutes the basic blade; the length ratio of the tip assembly section to the blade is 1:10.

[0024] Furthermore, the present invention provides a blade tip assembly for reducing noise having the greatest impact in the flight envelope of a rotorcraft, characterized in that the blade tip assembly is adapted to be detachably mounted on a distal end of a blade of the rotorcraft to form the blade tip, and the blade tip assembly has at least one of a forward-swept structure, a backward-swept structure, a tapered structure, a tapered structure, a downturned structure, and a twisted structure.

[0025] Among them, the line connecting the midpoints of the width lines of the forward-swept structure bends forward; the line connecting the midpoints of the width lines of the swept-back structure bends backward; the width line of the cross section of the tapered structure shrinks inward; the thickness line of the thinned structure shrinks inward; the line connecting the trailing edges of the airfoil of the cross section of the inverted structure bends downward; and the line connecting the trailing edges of the airfoil of the cross section of the twisted structure is not within the tip thickness range.

[0026] Preferably, the angle at which the line connecting the midpoints of the width lines of the forward-swept structure bends forward is 10° to 20°; the angle at which the line connecting the midpoints of the width lines of the backward-swept structure bends backward is 10° to 20°.

[0027] Preferably, the cross-sectional ratio of the end of the tapered structure away from the center of the hub to the end of the tapered structure close to the center of the hub is 0.5 to 0.9; the cross-sectional ratio of the end of the thinned structure away from the center of the hub to the end of the thinned structure close to the center of the hub is 0.7 to 0.9.

[0028] Preferably, the angle of the downward bending of the line connecting the trailing edges of the airfoil in the cross section of the downward inverted structure is 0-20°; the angle of the twisting of the line connecting the trailing edges of the airfoil in the cross section of the twisted structure is 0-30°.

[0029] In addition, the present invention also provides a rotor, which includes a blade tip assembly according to any one of claims 5 to 8, and the blade tip assembly is used to be detachably mounted on the end of the blade of the rotor to form the blade tip.

[0030] In addition, the present invention also provides a rotorcraft, wherein the blade of the rotorcraft includes a blade tip assembly according to any one of claims 5 to 8, and the blade tip assembly is used to be detachably mounted on the end of the blade of the rotorcraft to form the blade tip.

[0031] Compared with the prior art, the present invention has achieved the following technical effects:

[0032] The present invention manufactures a blade tip assembly through the above steps in the blade noise reduction method, thereby reducing the aerodynamic noise that has the greatest impact in the flight envelope of the rotorcraft. There is no need to design, process, and install the entire blade, saving the cost of design, processing, and installation and reducing the difficulty of installation. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments. 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 any creative work.

[0034] Figure 1 This is a schematic diagram of the installation position of the propeller tip assembly;

[0035] Figure 2 Schematic diagram of the principle of delaying the generation of shock waves when the blade tip assembly in the present invention only includes the forward-swept structure;

[0036] Figure 3is a top view schematic diagram of the reference blade tip;

[0037] Figure 4 is a schematic side view of a reference blade tip;

[0038] Figure 5 This is a top view schematic diagram of the forward-swept structure;

[0039] Figure 6 It is a top view schematic diagram of the swept structure;

[0040] Figure 7 It is a top view schematic diagram of the tapered structure;

[0041] Figure 8 This is a side view schematic diagram of the inverted structure;

[0042] Figure 9 is a side view schematic diagram of the thinning structure;

[0043] Figure 10 It is a side view schematic diagram of the torsion structure;

[0044] Figure 11 It is a schematic diagram of the structure of forward sweep and then backward sweep;

[0045] Figure 12 This is a schematic diagram of the structure of previously swept and then sharpened;

[0046] Figure 13 This is a schematic diagram of the structure that is first swept and then sharpened;

[0047] Figure 14 Schematic diagram of the structure of first inverting and then thinning;

[0048] Figure 15 A schematic diagram of a rotorcraft blade that needs to reduce noise being converted into a basic blade;

[0049] Figure 16 A schematic diagram of a rotor with a pointed blade tip assembly is provided;

[0050] Figure 17 A schematic diagram of a rotor with a swept-back configuration for installing a blade tip assembly;

[0051] Figure 18 Schematic diagram of the structure of different types of blade tip assemblies;

[0052] Figure 19 Schematic diagram of the noise reduction effect at the observation point obtained by simulating the CLORNS noise calculation code when the blade tip assembly only includes a swept structure;

[0053] Figure 20 is the calculation flow chart in step S2;

[0054] Figure 21Schematic diagram of the observation point location;

[0055] Figure 22 FIG1 is a parametric diagram of the planar shape of the blade tip assembly when the blade tip assembly is a combination of a forward-swept structure, a backward-swept structure, and a tapered structure;

[0056] Among them, 1. Reference blade; 2. Basic blade. DETAILED DESCRIPTION

[0057] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. 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 making creative efforts are within the scope of protection of the present invention.

[0058] 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.

[0059] The present invention discloses a method for reducing blade noise, comprising the following steps:

[0060] Step S1, processing the rotorcraft blades that need noise reduction into basic blades;

[0061] Step S2: evaluating the noise type with the greatest impact in the entire flight envelope based on the flight conditions of the rotorcraft, and determining a type of blade tip assembly that can reduce the noise of that type;

[0062] Step S3, designing specific parameters of the blade tip assembly;

[0063] Step S4, installing the tip assembly on the base blade to form a complete blade;

[0064] Among them, the basic blade is the remaining part of the blade after the tip is removed; the tip assembly is used to be installed in the original tip area of ​​the blade to form a new tip of the basic blade.

[0065] like Figure 15 As shown in the figure, this figure is a schematic diagram of processing the blade of the rotorcraft that needs noise reduction into the basic blade 2. Compared with the blade of the rotorcraft that needs noise reduction, that is, the blade to be modified (the blade to be modified can also be called the reference blade 1), the basic blade 2 is only based on the blade of the rotorcraft that needs noise reduction, with the original blade tip removed.

[0066] like Figure 16-17 As shown, these two figures are schematic diagrams of the structure of the basic blade installation pointed structure or swept structure respectively.

[0067] The passive suppression methods of aerodynamic noise in the existing technologies are all irreversible and inflexible. For example, the existing technologies can reduce the high-speed pulse noise generated during the flight of a rotorcraft by designing and installing blades with swept tips. However, if the flight conditions / flight missions of the rotorcraft change, resulting in a change in the type of aerodynamic noise with the greatest impact in the flight envelope of the rotorcraft, the blades with swept tips installed on the rotorcraft will not be able to effectively reduce the aerodynamic noise that is not high-speed pulse noise. In other words, the existing technologies can only reduce a single type of aerodynamic noise by replacing the blades, but when the type of aerodynamic noise with the greatest impact in the flight envelope of the rotorcraft changes, the rotorcraft will not be able to effectively reduce other types of aerodynamic noise.

[0068] "Passive suppression of aerodynamic noise" means achieving effects such as reducing the intensity of the sound source, isolating the propagation path or performing sound absorption treatment through the selection of structure or materials.

[0069] While this method controls aerodynamic noise through a passive method, namely, blade tip assembly shape design, this process can be made reversible by replacing blade tips of different shapes, and it has the flexibility to reduce different types of aerodynamic noise. "Making this process reversible by replacing blade tips of different shapes" means that because the blade tip assembly in the present invention is modular and detachable, even if the type of aerodynamic noise that has the greatest impact in the flight envelope of the rotorcraft changes, the rotorcraft can effectively reduce different types of aerodynamic noise by replacing blade tip assemblies of different types. In other words, the present invention solves the problem of adaptability to different flight states that exists in passive control methods.

[0070] The flight conditions of a rotorcraft refer to the environment and mission requirements it faces. The flight envelope is a closed geometric figure that represents the aircraft's flight range and operational limitations, using parameters such as speed, altitude, overload, and ambient temperature as coordinates. The aerodynamic noise generated by rotorcraft blades primarily includes thickness noise, load noise, high-speed impulse noise, and blade-vortex interference noise.

[0071] Furthermore, step S2 includes:

[0072] Step 1: Using the blade flow field solver module in the CLORNS series software, solve the flow field characteristics of the blade under calculation conditions to obtain the sound source information;

[0073] Step 2: Using the observation point information and the sound source information obtained in step 1 as input, the blade noise characteristic calculation module in the CLORNS series software performs calculations to obtain the aeroacoustic characteristics of the blade under a specific operating state (i.e., the operating state corresponding to the above calculation conditions);

[0074] Step 3: Analyze the aeroacoustic characteristics to determine the components and magnitudes of the aerodynamic noise throughout the flight envelope, and determine the required blade tip assembly.

[0075] The calculation conditions refer to taking the flight conditions of the entire flight envelope as the input of the blade flow field solution module in the CLORNS series software; the observation point information refers to the location of the observed sound.

[0076] like Figure 20 As shown in the figure, the acoustic field analysis process of the rotor is performed to obtain the type of aerodynamic noise with the greatest impact in the flight envelope of the rotorcraft, and to determine the type of blade tip assembly that can reduce this type of aerodynamic noise.

[0077] Specifically, the calculation conditions are used as inputs to the blade flow field solver module in the CLORNS series software. The calculation conditions include parameters such as rotor size, forward flight speed, rotor speed, and air density. The blade flow field solver module in the CLORNS series software (i.e. Figure 20 The CFD module in the CLORNS series software is used to solve the flow field characteristics of the blade under the calculation conditions and obtain the sound source information; then the observation point information and sound source information are used as the blade noise characteristic calculation module in the CLORNS series software (i.e. Figure 20 The blade noise characteristic calculation module in the CLORNS series software calculates the aeroacoustic characteristics of the blade under a specific working state (i.e., the working state corresponding to the calculation conditions) using the noise prediction module in the CLORNS series software as input; the aeroacoustic characteristics are then analyzed to obtain the components of the aerodynamic noise in the entire flight envelope and the size of each component, and to determine the type of blade tip assembly required.

[0078] Among them, the acoustic propagation characteristics determine that when observing (listening to) sounds at different distances and directions from the sound source, the observation results (the sounds heard) will be quite different. The position where the observer observes the sound is called the observation point. Figure 21 As shown, the observation points mainly refer to the extension direction of the rotorcraft's blade tip plane, the oblique front of the flight direction, and directly below the fuselage.

[0079] The calculated acoustic characteristics mainly include parameters such as sound pressure (hereinafter referred to as SP) and sound pressure level (hereinafter referred to as SPL) at a specific observation point. The calculation results can be analyzed to separate the components of aerodynamic noise and the size of each component.

[0080] Step S3 includes:

[0081] In the first step, characteristic points on the blade tip assembly that can reflect the overall profile of the blade tip are selected as design variables;

[0082] The second step is to determine the mathematical function that constructs the lines between the feature points;

[0083] The third step is to impose size range constraints on the design variables;

[0084] The fourth step is to determine the optimization constraints and optimization goals;

[0085] Step 5: Input all the parameters obtained in steps 1 to 4 into the CLORNS series software optimization module to obtain design variables that meet the standards and determine the processing parameters of the blade tip assembly;

[0086] Among them, the constraint condition is that the aerodynamic performance of the blade does not decrease compared with the original blade, and the optimization goal is to minimize the sound pressure level of the aerodynamic noise at the selected observation point.

[0087] In this case, the characteristic parameter optimization design method is used to design the parameters of the blade tip assembly required under different conditions. The specific content of step S3 is as follows:

[0088] In the first step, the characteristic points on the blade tip assembly that can reflect the overall contour of the blade tip are selected as design variables. After the type of blade tip assembly is selected in step S2, the key points that determine the shape are established as the characteristic points of this optimization process. Figure 22 As shown in the figure, P1 and P4 are starting points, P2 and P5 are turning points, and P3 and P6 are end points. All of these are characteristic points. If there are no turning points, only the starting and end points of the leading and trailing edges of the blade can be selected as characteristic points. The unknown coordinates of the characteristic points are the design variables.

[0089] The second step is to establish the mathematical function that constructs the shape of the line formed between the feature points. How to connect the two points is determined by the designer based on the acceptable processing difficulty and the required aerodynamic performance. The designer has a lot of freedom to choose. The following formulas can be used to connect adjacent feature points using the first, second, and third functions, or a combination of the above functions can be used to connect adjacent feature points. In the following formula, (x s ,y s ) is the starting point of this segment of the line, (x, y) is the coordinate of the point on this segment of the line, (x e ,y e ) is the end point of this line. The units of all x and y above are meters (m). R is specified as the rotor radius and c is the reference chord, that is, the length of the characteristic line A before the change, and the units are all meters (m).

[0090] Linear function: y = -k(xx e )+y e

[0091] Where,

[0092] Quadratic function: y = -k1(xx s ) 2 +y s

[0093] Where:

[0094] Cubic function: y=a1(xx s ) 3 +b1(xx s ) 2 +y s

[0095] Where:

[0096] Use Figure 22 The advantage of the cubic function and parabola shown is that the connections are tangent, which can achieve a smooth transition effect. Using straight lines to connect will reduce the processing difficulty.

[0097] The third step is to impose size range constraints on the design variables; for example, the horizontal coordinate of P2 is between P1 and P3.

[0098] The fourth step is to establish the optimization constraints and optimization goals.

[0099] Step 5: Input all the parameters obtained in steps 1 to 4 into the CLORNS series software optimization module to obtain design variables that meet the standards and determine the processing parameters of the blade tip assembly;

[0100] Among them, the constraint condition is that the aerodynamic performance of the blade does not decrease compared with the original blade, and the optimization goal is to minimize the aerodynamic noise sound pressure level at the observation point.

[0101] The following text will use the AH-1 / OLS model rotor blade as an example to explain the design process in detail. The AH-1 / OLS is selected as the initial blade, and the other sample blades are modified with the blade tip shape based on the initial blade.

[0102] like Figure 22As shown in the figure, this figure is a schematic diagram of the planar shape parameterization of the blade tip assembly when the blade tip assembly is a combination of a forward-swept structure, a swept-back structure and a tapered structure. R represents the rotor radius, c represents the reference chord length, x represents the spanwise position, y represents the chordwise position, and the dotted line in the figure is a quarter chord line. P1, P2 and P3 are three points on the leading edge of the blade plane shape, and P4, P5 and P6 are three points on the trailing edge, where P1 and P4 correspond to the leading edge point and trailing edge point of the section at 0.9 times the radius in the spanwise direction, respectively. Var1, var2, var3, var4, var5 and var6 are the six design variables in the optimization process, replacing the unknown (x, y) values ​​of the original position. Their units are meters (m). The horizontal axis below is the coefficient after non-dimensionalization using R, and the vertical axis is the coefficient after non-dimensionalization using c. The cross-sectional chord length from the root of the blade to 0.9R in the spanwise direction is the reference chord length. In order to ensure a smooth transition between the leading and trailing edges of the blade tip, the shape between the feature points is constructed by mathematical functions. First, the leading edge shape between points P1 and P2 is parameterized, and the spanwise position coordinate of point P2 is defined as var1, and the chordwise position coordinate is defined as var2. The curve between points P1 and P2 is a cubic function curve; point P3 is located at the leading edge blade tip, and the chordwise position of P3 is defined as var3. The leading edge shape curve between P2 and P3 is a parabola; the trailing edge shape of the blade is similar to the leading edge shape, and the change of the trailing edge shape of the blade starts at point P4. The spanwise position coordinate of point P5 is var4, and the chordwise position coordinate is var5. The curve between points P4 and P5 is also a cubic function curve; point P6 is located at the trailing edge blade tip, and the chordwise position coordinate is var6. A straight line transition is used between points P5 and P6.

[0103] The propeller tip shape is parameterized as follows:

[0104] (1) The leading edge when 0.9R<x≤var1,

[0105] y=a1(x-0.9R) 3 +b1(x-0.9R) 2 +c

[0106] Where:

[0107] (2) The leading edge when var1<x≤R,

[0108] y=-k1(x-var1) 2 +var2

[0109] Where:

[0110] (3) The trailing edge when 0.9R<x≤var4,

[0111] y=a²(x-0.9R)3 +b2(x-0.9R) 2 +c

[0112] Where:

[0113] (4) The trailing edge when var4<x≤R,

[0114] y=-k2(x-var4)+var5

[0115] Where:

[0116] In order to effectively realize the forward-swept-backward-tapered blade tip shape construction in the variable space, it is necessary to impose certain numerical constraints on each variable:

[0117]

[0118] In order to obtain good noise characteristics, the rotor must also have good aerodynamic performance. Therefore, for the optimization of the blade tip shape of the low-noise rotor, it is also necessary to introduce the equivalent lift-to-drag ratio of the rotor in forward flight as a constraint into the optimization process. The rotor equivalent lift-to-drag ratio L / D is defined as: L / D = C L / (C D +C Q / μ). Where: C L represents the lift coefficient, C D represents the drag coefficient, C Q represents the torque coefficient, and μ represents the forward ratio. All of the above parameters can be obtained through the flow field solution module in step 3. L represents the lift of the rotor, and D is the drag of the rotor. The lift-to-drag ratio of the optimized blade rotor should not be lower than that of the original rotor. Therefore, the variable parameters and target response function of the optimization work of this patent are as follows:

[0119]

[0120] In the above formula, SPL represents sound pressure level. Considering the large number of variables involved in this optimization, the optimization module in the CLORNS code was used to optimize the low-noise blade tip shape. The final result is the values ​​of X1 to X6. Once these values ​​are determined, the optimized shape parameters are obtained. The coordinates of the key points in the diagram are obtained, and the blade tip shape parameters are selected.

[0121] The blade tip assembly is then processed according to the obtained dimensional parameters; and the processed blade tip assembly is installed on the reference rotor to perform the flight mission.

[0122] like Figures 1-19As shown, the present invention discloses a blade tip assembly, which is used to be detachably mounted on the end of a rotorcraft blade to form the blade tip. The blade tip assembly has at least one structure selected from the group consisting of a forward-swept structure, a backward-swept structure, a tapered structure, a thinned structure, a reversed structure, and a twisted structure.

[0123] Among them, such as Figure 3-Figure 4 As shown, the characteristic line A is the width line of the blade in different cross sections; the characteristic line B is the line connecting the midpoints of the blade width line; the characteristic line C is the blade thickness line; and the characteristic line D is the line connecting the trailing edges of the airfoil of each cross section of the blade. Figure 5 As shown in FIG, the line connecting the midpoints of the width lines of the blades of the forward-swept structure bends forward; Figure 6 As shown in FIG, the line connecting the midpoints of the width lines of the swept-back blade is bent backward; Figure 7 As shown, the width line of the cross section of the tapered structure shrinks inwards; Figure 9 As shown in FIG, the thickness line of the thinned blade shrinks inward; Figure 8 As shown in FIG, the line connecting the trailing edges of the airfoil of the cross section of the blade with the downward-inverted structure is bent downward; Figure 10 As shown, the line connecting the trailing edges of the airfoil of the twisted cross section is not within the blade tip thickness range.

[0124] The present invention effectively reduces the noise generated during the flight of the rotorcraft through the blade tip assembly, thereby reducing sound pollution; at the same time, the detachable design of the blade tip assembly makes it possible for the rotorcraft to reduce noise by simply installing the blade tip assembly that can reduce the corresponding type of noise at the end of the rotorcraft blade when the rotorcraft needs to reduce noise, without having to remove the entire rotor, thereby simplifying the steps of replacing the blade tip assembly of the rotorcraft, making noise reduction of the rotorcraft easier, shortening the maintenance time of the rotorcraft, and improving the availability of the aircraft.

[0125] When the blade tip assembly of the present invention is applied to a military rotorcraft, the survivability of the military rotorcraft can be greatly improved; moreover, by adopting the blade tip assembly of the present invention, the vibration and noise of the rotorcraft can be reduced, thereby improving the scope of application of the rotorcraft.

[0126] The principles of noise reduction of the forward-swept structure, the swept-back structure, the tapered structure, the thinned structure, the reversed structure and the twisted structure in the present invention are as follows: the forward-swept structure and the swept-back structure reduce the component of the flow velocity in the tangential velocity, thereby reducing the magnitude of the tangential flow velocity, causing the blade tip speed to reach the speed of sound more slowly, delaying the generation of shock waves and thus suppressing high-speed pulse noise. Shock waves are strong compression waves in supersonic airflow, and shock waves often appear when the airflow speed exceeds the speed of sound. After the shock wave appears, the rotor sound field will produce extremely large high-speed pulse noise, and the forward-swept and swept-back designs suppress high-speed pulse noise by delaying the generation of shock waves. As Figure 2 As shown, Figure 23 is a schematic diagram showing the principle of the forward-swept structure delaying the generation of shock waves when the blade tip assembly in the present invention only includes the forward-swept structure.

[0127] The tapered and thinned structures of the present invention suppress thickness noise by reducing the volume of the blade tip and the volume of air displaced by the blade tip. The volume of air displaced per unit time is positively correlated with the magnitude of the thickness noise.

[0128] The twisting structure and the downward-reversed structure in the present invention change the direction of the tip vortex dragged out, so that the tip vortex dragged out by the front blade of a working rotor will not hit the following blade and generate a large blade-vortex interference noise.

[0129] When the rotorcraft is in an oblique descent state, the use of the blade tip assembly of the present invention (the blade tip assembly only includes a downward-reversed structure or a torsional structure, or includes both a downward-reversed structure and a torsional structure) can reduce the noise sound pressure level in the key area by more than 4dB, and the sound propagation range is reduced by more than 30%; when the rotorcraft is in a high-speed forward flight state, the use of the blade tip assembly of the present invention (the blade tip assembly only includes a forward-swept structure or a swept-back structure, or includes both a forward-swept structure and a swept-back structure) can reduce the sound pressure level of the observation point in the propagation direction of high-speed pulse noise by more than 5dB, and the sound propagation range is reduced by more than 30%.

[0130] Based on the above discussion, the blade tips in the rotor can effectively reduce the noise generated during the operation of the rotor by adopting different types of blade tip assemblies in the present invention.

[0131] Figure 18 The left column in the middle is a schematic diagram of the propeller blade from a top-down perspective, and the right column is a schematic diagram of the propeller blade from a side-on perspective.

[0132] Among them, "the blade tip assembly has at least one structure selected from the group consisting of a forward-swept structure, a backward-swept structure, a tapered structure, a thinned structure, a downward-inverted structure, and a twisted structure" means that the blade tip assembly in the present invention includes one structure selected from the group consisting of a forward-swept structure, a backward-swept structure, a tapered structure, a thinned structure, a downward-inverted structure, and a twisted structure; or includes any two structures, any three structures, any four structures, or any five structures selected from the group consisting of a forward-swept structure, a backward-swept structure, a tapered structure, a thinned structure, a downward-inverted structure, and a twisted structure; or the blade tip assembly includes a forward-swept structure, a backward-swept structure, a tapered structure, a thinned structure, a downward-inverted structure, and a twisted structure. Figure 11-14As shown in the figures, these are schematic diagrams of the structure when the blade tip assembly includes two structures. When the blade tip assembly includes multiple structures, the multiple structures can appear in stages or in sequence. For example, when the blade tip assembly includes a forward-swept structure and a swept-back structure, the blade tip assembly is sequentially provided with the forward-swept structure and the swept-back structure in the direction away from the center of the hub. In this case, it appears in stages. When the blade tip assembly includes a swept-back structure and a reversed structure, the blade tip assembly is reversed and then swept in the region of 0.9R to 1.0R, or swept first and then reversed. In this case, it appears in sequence.

[0133] It should be noted that the references to the forward-swept structure, the swept-back structure, the tapered structure, the thinned structure, the reversed structure and the twisted structure in the present invention, such as "the line connecting the midpoints of the width lines of the blades bends forward / backward, the width line of the cross section shrinks inward, the thickness line of the blades shrinks inward, the line connecting the trailing edges of the cross section bends downward and the line connecting the trailing edges of the cross section is not within the blade tip thickness range", are all compared with the blade tip that has not been modified, such as Figure 3-Figure 4 As shown in FIG. 1 , the two figures are a top view and a side view of an unmodified blade tip, i.e., a reference blade tip. The rotorcraft or rotorcraft mentioned in the present invention all refer to aircraft using rotors.

[0134] like Figure 1 As shown, Figure 1 Schematic diagram of the installation position of the blade tip assembly on the blade in the present invention, wherein point A represents the hub center of the blade, section AB is the undercut section (the undercut section of the blade of the rotor of the rotorcraft), section BC is the basic blade section, and section CD is the installation position of the original blade tip, i.e., the blade tip assembly section. Figure 1 The original blade tip at the middle CD section has been removed. At this time, the CD section is the installation position of the blade tip assembly in the present invention, and the AD section represents the structure from the center of the hub to the blade tip, that is, the blade.

[0135] The root cut section is a mechanical structure used to connect the blade to the hub. Since the root cut section is located at the root of the blade, close to the rotation center of the rotor, its linear speed is small during the rotation of the rotor, and its shape is not fixed. Its load is often ignored in simulation analysis. Therefore, this patent adopts the general viewpoint of simulation analysis and does not draw the shape of this section. The basic blade section is the blade to be installed with the tip assembly in this application after the original blade is removed. And this application uses the hub center, i.e. Figure 1 Point A is the origin, the direction of the blade extending away from the hub center is the positive direction, and the coordinate of the farthest point of the blade from the hub center is regarded as 1.0R. Specifically, the length of the AD segment is set to 1.0R, the length of the AC segment is set to 0.9R, and the length of the CD segment is set to 0.1R. Among them, the blade is regarded as a rotating part with forward flight speed, such as Figure 1As shown in Figure 2, the blade tip, further from the center of rotation, has a greater net velocity. A greater net velocity means a greater relative velocity with the fluid, making the blade tip more critical to aerodynamic performance than the blade root. Because the distance from the center of rotation and the proximity to the blade tip have the greatest impact on blade aerodynamic performance, the present invention selects the 0.9R to 1.0R segment as the modular tip CD segment.

[0136] Specifically, in the present invention, the angle at which the line connecting the midpoints of the width lines of the forward-swept structure bends forward is 10° to 20°; the angle at which the line connecting the midpoints of the width lines of the swept-back structure bends backward is 10° to 20°; the cross-sectional ratio of the end of the tapered structure away from the center of the hub to the end of the tapered structure close to the center of the hub is 0.5 to 0.9; the cross-sectional ratio of the end of the thinned structure away from the center of the hub to the end of the thinned structure close to the center of the hub is 0.7 to 0.9; the angle at which the line connecting the trailing edges of the airfoils of the cross-section of the blade of the downward-inverted structure bends downward is 0 to 20°; the angle at which the line connecting the trailing edges of the airfoils of the cross-section of the twisted structure twists is 0 to 30°.

[0137] Among them, when the forward-swept structure, the backward-swept structure, the tapered structure, the thinned structure, the downward-inverted structure and the twisted structure are within the above range, the present invention has a better noise reduction effect on different types of noise.

[0138] Furthermore, the present invention further provides a rotor, comprising the blade tip assembly of any one of the above embodiments, wherein the blade tip assembly is used to be detachably mounted on the end of the blade of the rotor to form the blade tip.

[0139] Furthermore, the present invention also discloses a rotorcraft, wherein the blade of the rotorcraft includes the tip assembly of any one of the above embodiments, and the tip assembly is used to be detachably mounted on the end of the blade of the rotorcraft to form the tip of the blade.

[0140] The present invention uses specific examples to illustrate the principles and implementation methods of the present invention. The above examples are only intended to help understand the method and core concept of the present invention. At the same time, those skilled in the art will find that the specific implementation methods and application scopes may vary based on the concept of the present invention. In summary, the contents of this specification should not be construed as limiting the present invention.

Claims

1. A blade noise reduction method, characterized in that: The following steps are involved: Step S1, processing the rotorcraft blades that need noise reduction into basic blades; Step S2: evaluating the type of aerodynamic noise that has the greatest impact in the entire flight envelope based on the flight conditions of the rotorcraft, and determining a type of blade tip assembly that can reduce the aerodynamic noise of that type; Step S3, designing specific parameters of the blade tip assembly; Step S4, installing the tip assembly on the base blade to form a complete blade; The basic blade is the remaining portion of the blade after removing the blade tip; the blade tip assembly is used to be installed in the original blade tip area of ​​the blade to form a new blade tip of the basic blade; The step S2 comprises: Step 1: Using the blade flow field solver module in the CLORNS series software, solve the flow field characteristics of the blade under calculation conditions to obtain the sound source information; Step 2: Using the observation point information and the sound source information obtained in step 1 as input, the blade noise characteristic calculation module in the CLORNS series software performs calculations to obtain the aeroacoustic characteristics of the blade under the corresponding calculation conditions. Step ③, analyzing the aeroacoustic characteristics to obtain the components of aerodynamic noise in the entire flight envelope and the magnitude of each component, and determining the required blade tip assembly; The calculation conditions refer to the flight conditions of the entire flight envelope as input to the blade flow field solution module in the CLORNS series software; the observation point information refers to the location where the sound is observed; The step S3 comprises: In the first step, characteristic points on the blade tip assembly that can reflect the overall profile of the blade tip are selected as design variables; The second step is to determine the mathematical function for constructing the lines between the feature points; The third step is to impose size range constraints on the design variables; The fourth step is to determine the optimization constraints and optimization goals; Step 5: Input all the parameters obtained in Steps 1 to 4 into the CLORNS series software optimization module to obtain the design variables that meet the standards, so as to determine the processing parameters of the blade tip assembly; The constraint condition is that the aerodynamic performance of the blade does not decrease compared to the original blade, and the optimization goal is to minimize the aerodynamic noise sound pressure level at the observation point.

2. The blade noise reduction method according to claim 1, characterized in that: The blade includes a root cut section, a basic blade section and a tip assembly section which are sequentially arranged in a direction away from the center of the hub, and the basic blade section constitutes the basic blade; the length ratio of the tip assembly section to the blade is 1:

10.

3. A blade tip assembly using the blade noise reduction method according to claim 1, used to reduce the noise that has the greatest impact in the flight envelope of a rotorcraft, characterized in that: The blade tip assembly is used to be detachably mounted on the end of the rotorcraft blade to form the blade tip of the blade, and the blade tip assembly has at least one structure selected from the group consisting of a forward-swept structure, a backward-swept structure, a tapered structure, a thinned structure, a reversed structure, and a twisted structure. Among them, the line connecting the midpoints of the width lines of the forward-swept structure bends forward; the line connecting the midpoints of the width lines of the swept-back structure bends backward; the width line of the cross section of the tapered structure shrinks inward; the thickness line of the thinned structure shrinks inward; the line connecting the trailing edges of the airfoil of the cross section of the inverted structure bends downward; and the line connecting the trailing edges of the airfoil of the cross section of the twisted structure is not within the tip thickness range.

4. The blade tip assembly according to claim 3, wherein: The line connecting the midpoints of the width lines of the forward-swept structure bends forward at an angle of 10° to 20°; the line connecting the midpoints of the width lines of the backward-swept structure bends backward at an angle of 10° to 20°.

5. The blade tip assembly according to claim 3, wherein: The cross-sectional ratio of the end of the tapered structure away from the center of the hub to the end of the tapered structure close to the center of the hub is 0.5~0.9; the cross-sectional ratio of the end of the thinned structure away from the center of the hub to the end of the thinned structure close to the center of the hub is 0.7~0.

9.

6. The blade tip assembly according to claim 3, wherein: The angle of the downward bending of the line connecting the trailing edges of the airfoil of the cross section with the downward inverted structure is 0-20 degrees; the angle of the twisting of the line connecting the trailing edges of the airfoil of the cross section with the twisted structure is 0-30 degrees.

7. A rotor, characterized in that: The rotor comprises the tip assembly according to any one of claims 3 to 6, and the tip assembly is used to be detachably mounted on the end of the blade of the rotor to form the tip of the blade.

8. A rotary-wing aircraft, characterized in that: The blade of the rotorcraft comprises the tip assembly according to any one of claims 3 to 6, and the tip assembly is used to be detachably mounted on the end of the blade of the rotorcraft to form the tip of the blade.

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