KR100550093B1 - Composite beam for reducing vibration and blade structure with the beam - Google Patents

Abstract

The present invention relates to a blade of a helicopter, a windmill, an aircraft, etc., and at least one kind of lamina having one fiber angle of -θ, 0 and + θ uniformly formed along one length in one lamina; It is characterized by consisting of at least one heterogeneous lamina, which is laminated in combination with this kind of lamina, and two fiber angles of -θ and + θ are alternately formed at regular intervals along the length. Accordingly, the present invention can lower the dynamic hub action load of the composite beam mounted in the blades of helicopters, windmills, aircraft, etc., and ultimately significantly reduce the vibration transmitted from the blades through the hub to the fuselage of the device.

Composite beam, blade, vibration, hub action, load

Description

Low vibration composite beam and blade structure equipped with the composite beam {COMPOSITE BEAM FOR REDUCING VIBRATION AND BLADE STRUCTURE WITH THE BEAM}

1 is a schematic perspective view showing a blade generally applied to a helicopter or a windmill, and the like.

Figure 2 is a schematic partially enlarged cross-sectional perspective view showing a composite beam according to the prior art and a blade structure having the composite beam.

Figure 3 is a schematic partially enlarged cross-sectional perspective view showing a rectangular low vibration composite beam and a blade structure having the composite beam according to the present invention.

4 is a schematic plan view and a side view of the blade structure of FIG.

5 is a schematic side enlarged view of the blade structure of FIG.

Figure 6 is a schematic plan view and side view showing a D-shaped low vibration composite beam and a blade structure having the composite beam according to the present invention.

7 is a schematic side enlarged view of the blade structure of FIG.

8 is a schematic plan view and side view showing an elliptical low vibration composite beam and a blade structure having the composite beam according to the present invention.

9 is a schematic side enlarged view of the blade structure of FIG. 8;

10 is a schematic plan view and side view showing a streamlined low vibration composite beam and a blade structure having the composite beam according to the present invention.

11 is a schematic side enlarged view of the blade structure of FIG.

12 is a schematic plan view and side view showing an I-shaped low vibration composite beam and a blade structure having the composite beam according to the present invention.

Figure 13 is a schematic side enlarged view of the blade structure of Figure 12;

14 is a schematic view showing a lead-free composite beam according to the prior art.

15 is a schematic view showing a combined 1 composite beam according to the prior art.

16 is a schematic view showing a linked 2 composite beam according to the present invention.

17 is a schematic view showing the geometric dimensional relationship of the linkage 2 composite beam according to the present invention.

18 is a view showing the bending vibration secondary mode when rotating the linkage 2 composite beam according to the present invention.

19 is a graph showing the characteristics of the bending-free composite mode and twisted primary mode of the lead-free composite beam and linkage 1 composite beam according to the prior art, and the linkage 2 composite beam according to the present invention.

20 is a graph showing the characteristics of the lead-free composite beam and linkage 1 composite beam according to the prior art, and the bending secondary mode and torsional secondary mode of the linkage 2 composite beam according to the present invention.

21 is a graph showing the characteristics of the longitudinal, lateral and vertical shear force of the lead-free composite beam and linkage 1 composite beam according to the prior art, and the linkage 2 composite beam according to the present invention.

22 is a graph showing the characteristics of the pitching, rolling and yawing moment of the lead-free composite beam and linkage 1 composite beam according to the prior art, and the linkage 2 composite beam according to the present invention.

<Description of the symbols for the main parts of the drawings>

1: hub 2: pliable beam

4: envelope 5: blade

6: honeycomb core 6a: foam core

7: Counterweight 8: Abrasion Resistant Cap

9: beam 9a: square beam

9b: D-beam 9c: elliptical beam

9d: Streamlined beam 9e: I-beam

9ee: Supporting part 10: Lamina

11: heterogeneous lamina 12: filling material

The present invention relates to a blade of a helicopter, a windmill, an aircraft, and the like, in particular, by forming a fiber angle differently in a zigzag shape along a length in a predetermined lamina of the laminated lamina of the spar (spar) provided in the blade, so that the blade A low vibration composite beam that radically lowers the dynamic hub load generated when it is rotated in the shaft and ultimately significantly reduces the vibration of the fuselage caused by the dynamic hub load applied from the blades to the body of the equipment through the hub. It is about a blade structure equipped with a composite beam.

In general, a hub (1) such as a helicopter, a windmill, an aircraft is provided with a flexible beam (2), as shown in Figure 1, the free end of the flexible beam and the hollow beam (3) serving as a support In addition, a blade 5 made of a shell 4 which encloses the periphery of the beam in a streamlined form to minimize the resistance of the air is fixed through fasteners such as rivets.

Here, a honeycomb core 6 is filled in an inner space between one side of the beam 3 and the rear edge of the shell 4, and the other side of the beam 3 and the shell 4. An internal space between the front edges of the counterweight 7 is provided, and the outer surface of the front curved portion of the shell 4 is attached with a wear-resistant cap 8 to protect the surface from the impact of air fluid (Fig. 2).

On the other hand, the beam (3) consists of a number of lamina (3a) made of at least one composite material selected from graphite, fiberglass, epoxy, aramid fibers, etc. As a result, the beam (3) is a metal material It exhibits high specific stiffness (stiffness per weight) and specific strength characteristics.

However, in the related art, since each lamina of a beam consisting of a plurality of lamina has a shape having a constant fiber angle pattern, a large dynamic hub load is generated every time the blade is rotated, and the hub of the flexible beam is removed from the blade. There was a problem that the fuselage of the corresponding device is severely shaken as it is delivered to the fuselage of the corresponding device.

Accordingly, the present invention has been made to solve the above-described problems, the object of which is to alternately form the fiber angle in a zigzag form differently along the length to a predetermined lamina of the beam lamina mounted in the blade of the helicopter, windmill, aircraft, etc. It is to provide a low vibration composite beam that can significantly reduce the dynamic hub action load of the composite beam.

Another object of the present invention is to provide a blade structure that is equipped with a low vibration composite beam in the shell of the blade to ultimately significantly reduce the vibration transmitted from the blade through the hub to the body of the device.

In order to achieve the above object, the low vibration composite beam according to the present invention is at least one kind of lamina in which one fiber angle of -θ, 0 and + θ is constantly formed along the length of one lamina; It is characterized by consisting of at least one heterogeneous lamina, which is laminated in combination with this kind of lamina, and two fiber angles of -θ and + θ are alternately formed at regular intervals along the length.

In addition, the blade structure provided with a low vibration composite beam according to the present invention is a combination of at least one kind of lamina having one fiber angle of -θ, 0 and + θ in one lamina uniformly along the length, and the kind of lamina A beam consisting of at least one heterogeneous lamina, which is stacked and alternately formed in one lamina with two fiber angles of -θ and + θ at regular intervals along its length; At least one envelope covering the upper and lower surfaces of the beam at least and having an airfoil (cross section) in a streamlined shape; It is attached to the front edge of the envelope is characterized in that consisting of a wear-resistant cap to protect the surface from the strong impact of the fluid.

Hereinafter, an embodiment of the present invention will be described with reference to FIGS. 3 to 22.

3 is a schematic partially enlarged cross-sectional perspective view showing a rectangular low vibration composite beam and a blade structure having the composite beam according to the present invention, Figure 4 is a schematic plan view and side view of the blade structure of FIG.

As shown in FIGS. 3 to 13, the low vibration composite beam 9 according to the present invention has at least one type in which one fiber angle of -θ, 0 and + θ is uniformly formed along a length in one lamina. Lamina 10;

It is composed of a structure consisting of at least one heterogeneous lamina 11 alternately formed by laminating in combination with this kind of lamina and alternately forming two fiber angles of -θ and + θ at a predetermined interval along one length.

Here, the longitudinal cross-sectional shape of the beam 9 may be formed of any one of a rectangle 9a, a D-shape 9b, an ellipse 9c, a streamlined 9d, and an I-shaped 9e, and the I-shaped. The beam 9e is preferably formed such that the upper and lower support portions 9ee have a predetermined curvature inwardly (see FIG. 13). In addition, the boundary point between −θ and + θ of the heterogeneous lamina 11 is preferably formed at the inflection point and the node point of the bending secondary mode (see FIGS. 16 to 18).

The dissimilar lamina 11 is the boundary surface of each section is integrally fixed by heat bonding or adhesive or the like between two adjacent lamina (10 or 11) or adjacent lamina (10 or 11) and the shell (4) And interlocked together with their adjacent lamina or shell by heat and pressure in an autoclave oven (not shown) (see FIGS. 5, 7, 9, 11 and 13).

As a result, it is possible to remove in advance the non-uniformity that may occur at the angle change point of the dissimilar lamina 11, even if the dissimilar lamina is located on the outermost of the lamina laminate or the innermost of the lamina laminate. Even if such a non-uniformity can be eliminated because it is securely surrounded by the shell 4 of the blade according to the embodiment below.

On the other hand, the blade structure provided with a low vibration composite beam according to the present invention is at least one kind lamina 10 is formed in one lamina uniform fiber length of one of -θ, 0 and + θ along the length, and A beam (9) composed of at least one heterogeneous lamina (11) laminated in combination with a kind of lamina and alternately formed in one lamina with two fiber angles of -θ and + θ at regular intervals along its length;

At least one envelope 4, the airfoil of which the airfoil is in close proximity to at least the upper and lower surfaces of the beam 9;

It consists of a wear-resistant cap 8 attached to the front edge of the shell 4 to protect the surface from the strong impact of the fluid (see FIGS. 3 to 13).

In this case, the longitudinal cross-sectional shape of the beam 9 may be formed of any one of a rectangle 9a, a D-shape 9b, an ellipse 9c, a streamlined 9d, and an I-shaped 9e. The beams 9a and the elliptical beams 9c are disposed in the shells so as to form a predetermined inner space on both sides of the shell 4, and the counterweight 7 in the front inner space of the shell 4 is provided. It is provided, and the opposite inner space has a structure filled with the filler 12 (see FIGS. 4 and 5 and 8 and 9).

The D-shaped beam 9b has a structure in which the rounded portion of the D-shaped beam is intimately opposed to the inner side of the front edge of the shell 4, and the opposite inner space is filled with the filler 12 (Fig. 6 and FIG. 7).

The streamlined beams 9d are arranged closely to each other in the inner space of the shell 4 and have a structure in which the filler 12 is filled in the streamlined beams 9d (see FIGS. 10 and 11).

The I-shaped beams 9e are disposed so as to face each other in two inner spaces of the shell 4, and the upper and lower support portions 9ee of each of the I-shaped beams 9e are in close contact with the inner surface of the shell 4. The inner space between the shell and the two I-shaped beams 9e has a structure filled with a filler 12 (see FIGS. 12 and 13).

Each of the fillers 12 preferably consists of any one of a honeycomb core 6 and a foam core 6a.

On the other hand, each of the lamina (10, 11) of the beam (9) as described above is at least one composite material selected from the group consisting of graphite, fiberglass, epoxy, aramid fibers, etc. having a high specific rigidity and specific strength characteristics It is preferable to make.

On the other hand, the heterogeneous lamina 11 is wrapped by the adjacent one or more lamina 10 or the outer shell 4 is fixed to each other so as not to be separated from any impact under a predetermined design condition unless the operator intended do.

EXAMPLE

The low vibration composite beam according to the present invention configured as described above and the composite beam according to the prior art are applied to the blades such as helicopters or windmills, respectively, to find out the difference in the vibration movement of each blade generated when the blade rotates. The same experiment was performed.

First, in connection with a lead-free composite beam according to the prior art (lamina laminate composed of either 0 ° or 90 ° in one lamina fiber) 1 composite beam (fiber angle in one lamina -θ and Laminated laminate composed of any one of + θ and, in addition, compared with the lead-free composite beam and the linkage 1 composite beam, the linkage 2 composite beam according to the present invention (two of -θ and + θ in one lamina Branch fiber angles were alternately formed lamina laminate formed at regular intervals along the length) (see Figs. 14 to 16).

Here, the predetermined heterogeneous lamina 11 constituting the linkage 2 composite beam is divided into section I, section II and section III, and the fiber angle of section II adjacent to section I and section III adjacent to section II. Were formed differently.

The length of the section I (r ₁ ) is 0.2R (length of the composite beam), the fiber angle -15 ° from the start of the composite beam; The length of section II (r ₂ ) is 0.8R, fiber angle + 15 ° from the start of the composite beam; The length of the section III was formed with the remainder of R, the fiber angle of -15 °, and the boundary point between -θ and + θ of the heterogeneous lamina 11 was formed at the inflection point and the node point part of the bending secondary mode. (See FIGS. 16-18).

The composite beam width w; 0.02R, height; 0.01R, thickness t; It was produced at a ratio of 0.01c (length of chord), where c is 0.08R (length of blade) (see FIG. 17).

Then, the lead-free composite beam and the linkage 1 composite beam were mounted in the corresponding shell, respectively, to construct a blade according to the prior art, and the linkage 2 composite beam was mounted in the shell, and the blade according to the present invention was configured. The remaining components of the blade were configured under the same conditions to increase the reliability of the experiment.

Then, as a result of rotating the respective blades in a vacuum state, the bending primary mode of the lead-free composite beam and the torsion primary mode characteristics of the linkage 1 composite beam and the linkage 2 composite beam are shown in the graph of FIG. 19. In addition, the bending secondary mode of the lead-free composite beam and the torsion secondary mode characteristics of the linkage 1 composite beam and the linkage 2 composite beam are shown in the graph of FIG. 20.

Here, it can be seen that there is no change in the shape of the bending mode in the lead-free, linkage 1, and linkage 2 composite beams, but the torsional mode is significantly changed. The shape change of the torsional mode is commonly observed in the first vibration mode or the second vibration mode, and it can be confirmed that the linkage 2 composite beam can be effectively used for vibration control compared to the conventional leadless composite beam or the linkage 1 composite beam. .

Meanwhile, the longitudinal, lateral and vertical shear forces of each of the leadless composite beam, the linkage 1 composite beam, and the linkage 2 composite beam are shown in the graph of FIG. 21, and the each of the leadless composite beam and the linkage 1 composite beam. Pitching, rolling, and yawing moments of the beam and linkage 2 composite beams are shown in the graph of FIG. 22.

Here, except that the vertical shear force of the linkage 2 composite beam according to the present invention is slightly higher than that of the linkage 1 composite beam according to the prior art, the longitudinal and lateral shear forces of the linkage 2 composite beam are different from that of the lead-free composite beam. Pitching and rolling moment of the linkage 2 composite beam except that the yaw moment of the linkage 2 composite beam according to the invention is slightly higher than that of the linkage 1 composite beam according to the prior art. Was lower than that of lead-free composite beams.

That is, since the linkage 2 composite beam according to the present invention generally generates a lower shear force and moment than the lead-free composite beam and linkage 1 composite beam according to the prior art to apply the linkage 2 composite beam according to the invention to the blade of the device. This will significantly reduce the dynamic hub load applied from the blades through the hub to the fuselage of the device.

On the other hand, when the blade to which the linkage 2 composite beam according to the present invention is applied to rotate by mounting a helicopter or windmill, etc., if the fiber angle is formed differently in a zigzag shape along the length of the laminated lamina of the linkage 2 composite beam along the length of the blade The mathematical explanation of the change in dynamic hub load is as follows.

First, the governing equation of the blade can be obtained based on Hamilton's principle.

[Equation 1]

는 블레이드의 변형율에너지 변분, 운동에너지 변분 및 공기력에 의한 가상일 변분을 나타내며, 상기 변형율에너지 변분의 수식은 다음과 같이 표현된다.Where t_OneAnd t₂Is a random time, each of the above                                                  

Denotes the strain energy variation, the kinetic energy variation, and the virtual work variation by the aerodynamic force, and the equation of the strain energy variation is expressed as follows.

[Equation 2]

은 수직응력과 전단응력을 나타내며, 상기 각

은 상기 각 응력에 대응되는 변형율 성분을 나타낸다. 한편, 2차원 평면응력 상태를 가정했을 때 xy 및 xz 평면에 대한 복합보 라미나의 응력 및 변형율 관계식은 다음과 같이 표현된다.Here, R is the length of the blade, A is the cross-sectional area of the blade, and each of the x, y, z represents the coordinates of the blade (see Fig. 1),

Represents vertical stress and shear stress,

Represents a strain component corresponding to each stress. On the other hand, assuming a two-dimensional plane stress state, the relationship between the stress and strain of the composite lamina for the xy and xz plane is expressed as follows.

[Equation 3]

[Equation 4]

In this case, the stiffness coefficient C _ij components are defined as follows.

[Equation 5]

는 복합보 라미나의 환원강성계수를 나타내며, 그들의 각 성분은 다음과 같다.Where                                                  

 Denotes the reduction stiffness coefficient of the composite lamina, and their respective components are as follows.

[Equation 6]

Here, θ represents the fiber angle of the composite lamina, and Q _ij is the stiffness coefficient of the composite _boramina . On the other hand, the displacement-strain relationship obtained by considering the second order nonlinearity is as follows.

[Equation 7]

는 블레이드의 기하학적 비틀림각

을 나타낸다.Here, each of u, v, w and φ represents the length of the blade, lag and flap displacement, and torsional displacement, and the superscript 'integrates with x, λ_TTorsional warping, θ₀Is the initial torsion angle,                                                  

 The geometric torsion angle of the blade

Indicates.

Substituting Equations 3, 4 and 7 into the Strain Equation 2 provides a strain energy formula consisting of displacement components of the blade. On the other hand, the kinetic energy variation of the equation 1 and the variation of the virtual work is expressed as follows.

[Equation 8]

[Equation 9]

는 허브에 대한 블레이드의 임의지점의 상대속도이다. 또한, 상기 수학식 9에서 상기 각 L_u, L_v, L_w , M_ψ는 인장, 래그, 플랩, 그리고 피칭 모멘트 하중을 나타낸다. 상기 수학식 2, 수학식 8 및 수학식 9를 상기 수학식 1에 대입하여 정리하면, 블레이드의 변위성분들에 의한 지배 운동방정식을 얻을 수 있게 된다. 그 지배 운동방정식에 유한요소법을 적용하여 이산화된 블레이드 지배 운동방정식을 구성하면 다음과 같은 2계 상미분 운동방 정식이 얻어진다.In Equation 8, ρ is the density of the blade,

Is the relative speed of any point of the blade with respect to the hub. In addition, in Equation 9, each of L _u , L _v , L _w , and M _ψ represents tensile, lag, flap, and pitching moment loads. By substituting Equation 2, Equation 8 and Equation 9 into Equation 1, a governing equation of motion by displacement components of the blade can be obtained. When the discretized blade governing equations of motion are constructed by applying the finite element method to the governing equations of motion, the following two-phase ordinary differential equations of motion are obtained.

[Equation 10]

Here, each of M, C, K, and F represents mass, damping, stiffness matrix, and external force vector, and q is a generalized coordinate of the blade.

The hub of the helicopter or windmill is subjected to periodic loads by the aerodynamic and inertial forces generated by the movement of the blade, and this hub acting load is obtained by summing the rotating loads acting on the N _b blades. This method is called the Force Summation Method, and using the force summation method, the hub acting load is expressed as the loads on the rotational coordinate system as follows.

[Equation 11]

Here, β _p is the initial cone angle (precone angle), ψ is the azimuth angle of the blade rotation surface, the angle F and M are the shear force and the moment component. In Equation 11, the superscript m means the m th blade, and the superscript H means a hub which is a non-rotating coordinate system. In Equation 11, the rotational load components for the m th blade are obtained as follows.

[Equation 12]

Here, each of L _u , L _v , L _w and M _u , M _v , M _w represents the load and moment components along the direction of motion u, v, w of the blade, and they apply the two-dimensional quasi-normal aerodynamic theory. To obtain.

In general, for a rotor consisting of N _b blades, the N _b / rev load is transmitted in a direction perpendicular to the hub. Separating these N _b / rev hub acting loads into higher order harmonic terms using the Fourier series gives N _b / rev hub acting loads delivered to the fuselage of the equipment. The final equation for the hub acting load is expressed as

[Equation 13]

Here, f ₀ represents a steady state component of the N _b / rev hub acting load, and each of the f _nc and f _ns represents cosine and sine components of the N _b / rev dynamic hub acting load.

In conclusion, when the fiber angle of a predetermined lamina among the laminated lamina of the composite beam constituting the blade such as a helicopter or a windmill is formed differently along the length of the lamina, the magnitude of N _b / rev load transmitted to the hub is different. In the design of the blade, such characteristics can be used to produce the optimum blade with minimal vibration.

As described above, the present invention is a composite beam by alternately stacking the fiber angle in a section of -θ and + θ along the length to a predetermined lamina of the laminated lamina of the beam mounted in the blade of the helicopter, windmill, aircraft, etc. It is possible to significantly lower the dynamic hub load of the.

In addition, the present invention is equipped with a low-vibration composite beam in the shell of the blade can ultimately significantly reduce the vibration transmitted from the blade through the hub to the body of the device.

Claims (6)

At least one kind of lamina in which one fiber angle of -θ, 0 and + θ is constantly formed along the length of one lamina; At least one heterogeneous lamina laminated in combination with this kind of lamina and alternately formed in one lamina with two fiber angles of -θ and + θ at regular intervals along the length Low vibration composite beam, characterized in that consisting of. The low vibration composite beam according to claim 1, wherein the longitudinal cross-sectional shape of the beam is any one of a rectangle, a D-shape, an ellipse, a streamline, and an I-shape. The low vibration composite beam according to claim 2, wherein the upper and lower support portions of the I-shaped beam form a predetermined curvature toward the inside. The low vibration composite beam of claim 1, wherein a boundary point between −θ and + θ of the heterogeneous lamina is formed at an inflection point and a node point of the bending secondary mode. At least one kind of lamina having a uniform angle along one of the lengths of -θ, 0 and + θ in one lamina, and laminated in combination with this kind of lamina, and having -θ and + θ A beam consisting of at least one heterogeneous lamina having two fiber angles alternately formed at regular intervals along its length; At least one envelope covering the upper and lower surfaces of the beam at least and the airfoils being streamlined; Abrasion resistant cap attached to the front edge of the shell to protect the surface from the strong impact of the fluid Blade structure provided with a low vibration composite beam, characterized in that consisting of. The blade structure according to claim 5, wherein the longitudinal cross-sectional shape of the beam is any one of a rectangle, a D-shape, an ellipse, a streamline, and an I-shape.

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