AIRFOIL OR ROTOR BLADE HAVING A CONTINUOUS TRAILING EDGE FLAP

Abstract

Embodiments of the present invention include various airfoils constructions, such as airfoils and rotor blades having a continuous trailing-edge flap. According to an embodiment, an airfoil is comprised of a load-bearing structure in the forward portion of the airfoil; a skin to maintain the airfoil shape; a tapered composite structure having one or more actuators extending from the load-bearing structure toward the trailing edge, and joining the skin in the vicinity of the trailing edge; and a core extending from the load-bearing structure connecting the tapered composite structure to the load-bearing structure. In this airfoil, the tapered composite structure tapers from being relatively thick near the core to being relatively thin near the trailing edge. And the one or more actuators are configured to deflect the trailing edge so as to deform the shape of the airfoil cross-section.

Description

CROSS REFERENCE TO RELATED APPLICATION(S)

This application claims the benefit of U.S. Provisional Patent Application No. 61/943,642 filed Feb. 24, 2014, herein incorporated by reference in its entirety for all purposes.

GOVERNMENT INTEREST

Governmental Interest—The invention described herein may be manufactured, used and licensed by or for the U.S. Government.

BACKGROUND OF THE INVENTION

i) Field of Invention

This application generally relates to airfoil constructions, and more particular to, airfoils and rotor blades having a continuous trailing-edge flap.

ii) Description of Related Art

Small flaps attached to the trailing edge of helicopter rotor blades have been proposed for use in improving rotor performance, reducing vibration and providing primary flight control.

For example, traditional active-flap rotors, such as the one shown in FIG. 1, have been developed over the last several decades. The primary drawback to these blades is their mechanical complexity; but the aerodynamic inefficiency associated with the flap is also a concern. This is because the rotors use discrete flaps which require hinges and other mechanical linkages in order to operate. A spinning rotor induces very large centrifugal loads within the blade and it is very difficult to design a mechanical flap that moves freely and reliably in the presence of the centrifugal acceleration. In addition, rotor blades bend appreciably during operation, and this bending tends to induce binding within the mechanical linkages. The gap between the flap and the blade, along with any projections of the linkages into the airflow, reduces the aerodynamic efficiency of the rotor blade and increases the power required during flight.

An alternative design to the active flap is the active-twist rotor, such as the one shown in FIG. 2. In this design, piezoelectric plies are embedded in the spar such that twist is induced in the spar when voltage is applied to them. The induced local twist is relatively small, but when the actuators are used over the majority of the blade, the net effect provides control authority similar to the active flap. This design eliminates the disadvantages of the active-flap rotor since there are no mechanical linkages or gaps in the rotor blade, but requires a very large numbers of piezoelectric actuators to be used.

Improvements in airfoil and rotor blade design would be useful.

BRIEF SUMMARY OF THE INVENTION

Embodiments of the present invention include various airfoils constructions, such as airfoils and rotor blades having a continuous trailing-edge flap. According to an embodiment, an airfoil is comprised of a load-bearing structure in the forward portion of the airfoil; a skin to maintain the airfoil shape; a tapered composite structure having one or more actuators extending from the load-bearing structure toward the trailing edge, and joining the skin in the vicinity of the trailing edge; and a core extending from the load-bearing structure connecting the tapered composite structure to the load-bearing structure. In this airfoil, the tapered composite structure tapers from being relatively thick near the core to being relatively thin near the trailing edge. And the one or more actuators are configured to deflect the trailing edge so as to deform the shape of the airfoil cross-section.

The various elements of the airfoil embodiments may be judiciously designed and configured for different implementations and/or applications. For example, the skin may be composed of one or more discrete flexible sections above and below the tapered composite structure. The load-bearing structure may be composed of composite plies or an integral structure. And the core can be formed in various ways. For instance, (i) the core may be comprised of composite plies of the load-bearing structure extending rearward, (ii) the core can be formed of separate composite plies connected to the load-bearing structure, (iii) the core may be comprised of a separate piece connected to the load-bearing structure, or (iv) the core might also be comprised an integral portion formed on the load-bearing structure. The airfoil may further include honeycomb or foam cores, above and below, the tapered composite structure.

The actuators may be bonded and/or attached via adhesive, glue, and/or fasteners. For example, the actuators may be formed of multiple layers of the piezoelectric actuators connected together, such as being configured as a bimorph. While one actuator may be provided in a basic embodiment of the airfoil, in most embodiments, two or more actuators are provided, such as with at least one actuator provided at the top and the bottom of the tapered composite structure, respectively.

Moreover, the tapered composite structure may be attached to top and bottom portions of the core, respectively. Near the core, for instance, in one embodiment, the thickness of the tapered composite structure may be approximately 2% of the chord length of the airfoil, and near the trailing edge, the thickness of the tapered composite structure may be about 0.3% of the chord length. The taper of the tapered composite structure may be linearly tapering. And, the degree of tapering at the top and the bottom of the tapered composite structure may be different.

The innovative airfoil embodiments disclosed herein lend themselves to various applications. For instance, the airfoil may be configured for use as a helicopter or tiltrotor rotor blade, a wind turbine blade, an aerodynamic control surface, or the like.

BRIEF DESCRIPTION OF THE DRAWINGS

So that the manner in which the above recited features of the present invention can be understood in detail, a more particular description of the invention, briefly summarized above, may be had by reference to embodiments, some of which are illustrated in the appended drawings. It is to be noted, however, that the appended drawings illustrate only a few embodiments of this invention and are therefore not to be considered limiting of its scope, for the invention may admit to other equally effective embodiments. These embodiments are intended to be included within the following description and protected by the accompanying claims.

FIG. 1 shows an example of a conventional active flap rotor blade.

FIG. 2 shows a cross-sectional view of an example of a conventional active-twist rotor blade.

FIG. 3 shows a cross-sectional view of an airfoil construction according to an embodiment.

FIG. 4 is a detailed view of the tapered composite structure of an airfoil construction according to an embodiment.

FIG. 5 shows airfoils of various tapered composite structures formed of different numbers of actuators plies, optimized for a typical aerodynamic load.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 3 shows the construction of an exemplary airfoil 10 including a tapered composite structure according to an embodiment of the present invention. The innovative airfoil 10, as shown in the cross-sectional view here, is an aerodynamic control surface which is configured to create an aerodynamic force due to the relative interaction of the airfoil and a fluid (such as air or water). Thus, the term airfoil as used herein should not be construed as being limited to only air as fluid medium. For many applications, the airfoil 10 of chord length C may have a forward rounded leading edge LE, followed by a rear sloping trailing edge TE, with an asymmetric camber in between. The exterior shape, and thus the aerodynamic profile of the airfoil 10 can be the made to be the same or similar as traditional airfoils. Of course, other aerodynamic shapes are also possible in keeping with the scope of the invention.

In general, the airfoil 10 shown in FIG. 3 is constructed of a load-bearing structure 11, a skin 12, a tapered composite structure 13 having at least one actuator 14, and a core 15. As shown in FIG. 3, the load-bearing structure 11 is located typically in the forward portion of the airfoil near the leading edge LE. Vertical line L-L shows the rearward edge of the load-bearing structure 11. While it appears that the load-bearing structure 11 tapers to a point at line L-L, this configuration is not limiting, and other configurations are envisioned, such as the load-bearing structure 11 tapering to a relatively smaller curved rearward shape, or a straight rearward shape at this line. The load-bearing structure 11 may be formed of suitable material for aircraft or underwater applications which is sufficient to withstand aerodynamic forces exerted thereupon, including the centrifugal force and the bending and tensional forces experienced by the airfoil 10. Such material may include metals, wood, plastics, composites, fiberglass and/or the like. In some embodiments, the load-bearing structure 11 may be fabricated of composite plies (such as E-glass fabric) or an integral structure. To optimize strength-to-weight ratio, the load-bearing structure 11 may be configured as a spar, beam, truss, or box structure, as typically found in airfoil constructions. In some instances, load-bearing structure 11 may be the primary or only load-bearing structure of the airfoil 10. If the load-bearing structure 11 is (at least partially) hollow, it may be filled with a foam or core material, such as Rohacell foam.

The smooth exterior shape of the airfoil 10 is maintained by the outer skin 12 which can form at least a portion of the exterior surface of the airfoil 10. The skin 12 can be formed of a flexible material sufficient to maintain the airfoil shape, but enable deformation of the airfoil cross-section. The flexible skin 12 material can be formed of aerospace-grade nylon, for instance, although other flexible materials may be similarly used. Moreover, the material used for the skin 12 should have as low stiffness as possible to maximize deflection. For example, the skin 12 may be composed of one or more discrete flexible sections, located above and below, the tapered composite structure 13.

The region between the skin 12 and the tapered composite structure 13 can be filled with a core material 16 such as honeycomb or structural foam. One material that may be used for the core material 16 is Nomex honeycomb. This combination ensures that the airfoil will maintain its shape in the presence of aerodynamic loads, but will be flexible enough that the tapered composite structure 13 will be able to bend the cross-section when one or more of the actuators 14 are energized or otherwise actuated. The core material 16 may be configured so as to be relatively stiff in the direction transverse to the tapered composite structure 13, but relatively flexible in the chordwise direction. This enables the core material 16 to maintain the airfoil shape when aerodynamic pressure is applied to the skin 12 while allowing the tapered composite structure 13 to bend when the actuators 14 are energized or actuated. The skin material 12 should also have sufficient structural durability and stiffness to avoid skin dimpling over core material 16.

In order to be able to both produce sufficient deformation and resist aerodynamic pressure on the trailing edge TE, the tapered composite structure 13 uses a tapered geometry which is relatively thicker at its front near the core 15 where it joins or connects to the load-bearing structure 11 than at its rear, near the trailing edge TE. The increased thickness near the core 15 reduces the bending curvature the tapered composite structure 13 can produce, but significantly increases the bending stiffness. The structural moment induced by the aerodynamic forces is typically small at the trailing edge TE, but increases towards the front of the tapered composite structure 13. Thus, the increased bending stiffness at its front is necessary for the airfoil to be able to resist the aerodynamic pressure. Similarly, the thinner thickness in the vicinity of the trailing edge TE produces greater bending during actuation, and thus maximizes the aerodynamic lift and moment generated by the airfoil 10.

FIG. 4 is a detailed view of the tapered composite structure 13 of the airfoil 10 shown in FIG. 3 from line L-L to the trailing edge TE. It is noted that this figure is not to scale, and that the tapered composite structure 13 and thus the tailing edge TE are deflected somewhat. The tapered composite structure 13 generally extends rearward from the load-bearing structure 11 toward the trailing edge TE and joins the skin 12 in the vicinity of the trailing edge TE. And the tapered composite structure 13 tapers from being relatively thicker in the vicinity where it joins the core 15 compared to being relatively thinner in the vicinity of the trailing edge TE. The centerline of the airfoil 10 is indicated as line 17.

For example, in one embodiment, near the core 15, the total thickness D₁ of the tapered composite structure 13 (as measured from top to bottom) may be approximately 2% of the chord length C of the airfoil, whereas near the trailing edge TE, the total thickness D₂ of the tapered composite structure 13 may be only about 0.3% of the chord length C. The actual dimensions of the airfoil 10 will vary as desired for a particular application, and in general can be scaled. The tapered composite structure 13 may taper from near the core 15 to near the trailing edge TE in a linear or non-linear manner, such as in a curved or arched profile, although, a linear taper may be preferred for some embodiments. A linear tapered shape of the tapered composite structure 13 has been found to be optimum for some embodiments; in fact, this tapering has been found to generate on average 63% more deflection than those of a rectangular-shaped bimorph bender (having no taper) used instead of the tapered composite structure 13.

Moreover, the degree of tapering at top and bottom portions of the tapered composite structure 13 may each vary. For instance, the top tapering could be much more pronounced than the bottom tapering, or vice versa.

FIG. 5 shows airfoils of various tapered composite structures 13 formed of 2, 3, 4, and 5 of actuator plies, optimized for a typical aerodynamic load. The complete airfoil is first shown in proportion, and then the trailing-edge section is displayed with amplification to show the tapered composite structures 13 in detail. The initial configuration of the stack design is shown next, which is selected to resemble the final optimal design based on the geometry of a solid structure. The design variables are the length and shape of the core and the chordwise locations of middle plies. The top and bottom plies are fixed at the root and tip of the tapered composite structure respectively. The results show that all the stacked, tapered composite structures 13 converge to a shape similar to the optimal solid structure, being thicker at the root and thinner at the tip. This is attained by having the core thickness tapering off toward the trailing edge. The two-ply structure has the largest core while the five-ply bender has the smallest, again to have the actuator plies forming the optimal taper shape as the solid structure. The three-ply structure has the chordwise location of its middle ply at almost equal distance from the top and bottom plies. The four-ply and five-ply structures have the extra middle plies located close to the structure root.

The one or more actuators 14 are configured to deflect the trailing edge TE so as to deform the shape of the airfoil cross-section and change the aerodynamic lift and moment which it produces. In a basic embodiment of the airfoil, one actuator may be provided; although, it should be appreciated that a greater degree of deflection of the trailing edge TE and thus control of the airfoil performance can be realized when multiple actuators are provided. Thus, in most embodiments, two or more actuators 14 are provided, such as with at least one actuator may be provided at the top and the bottom of the tapered composite structure 13, respectively. This arrangement is shown in FIGS. 3 and 4. More particularly, the tapered composite structure 13 may be attached to top and bottom portions of the core 15 as more clearly shown in FIG. 4.

Deformation of the tapered composite structure 13 via the actuators 14 changes the aerodynamic lift and moment of the airfoil 10. The actuators 14 may be piezoelectric actuators, for example. A piezoelectric actuator is actuator device in which a thin layer of piezoelectric material is combined with electrodes such that the actuator expands or contracts when voltage is applied to it. Piezoelectric actuators have a very high bandwidth, so they can be actuated in a sinusoidal manner at frequencies many times higher than the typically rotational speed of the rotor blades of many types of rotorcraft (such as helicopters). In general, piezoelectric actuators are typically very thin (e.g., approximately 0.3 mm). Therefore, all but the smallest of rotor blades (e.g., typically less than 2 m in radius), likely will require multiple actuators to be stacked together on one or both sides of the tapered composite structure 13. One exemplary actuator which may be used in embodiments includes a Macro Fiber Composite (MFC), a low-profile actuator, produced by Smart Material Corp., for example.

In some embodiments, the actuators 14 may be formed of multiple layers of piezoelectric actuators connected together, such as being configured as a bimorph. A bimorph is a structure comprised of at least two layers of actuator, which can produce a displacement. The piezoelectric actuators may be bonded and/or attached via adhesive, glue, and/or fasteners (such as rivets). The bimorph could have two, three, four, five or potentially even more stacks. Although, a four-ply bimorph configuration of the tapered composite structure 13 may be an optimal choice among the stack layouts because of its large output of angle relative to the number of plies required. If a tapered bimorph is embedded with the structure of a rotor blade, such that one end is attached to the load-bearing structure 11 and the other extends to the trailing edge TE, then the bending of the bimorph will result in deformation of the airfoil with the trailing edge TE moving up or down. This deformation results in changes in aerodynamic lift and moment, similar the effect of a discrete flap.

When energy (voltage) is applied to the actuators 14 the trailing edge of the airfoil 10 will deflect up and/or down. If actuators 14 receive opposing voltages, then one will expand and the other will contract. This induces a structural moment and thus the tapered composite structure 13 will bend. As the tapered composite structure 13 bends downward, the trailing edge TE of the airfoil 10 is deflected down, increasing the lift and generating a nose-down pitching moment on the airfoil 10. The voltage applied to the actuators 14 can be reversed, causing the tapered composite structure 13 to bend up and the airfoil 10 to produce less lift and a nose-up pitching moment.

The relatively thinner tapered composite structure 13 near the trailing edge TE produces more deformation there; thus, the thin trailing edge creates increased deformation that is needed to produce the changes in aerodynamic lift and moment. By actuating individual or groups of actuators 14 during flight, it is possible to reduce vibratory loads, reduce rotor noise, and improve aerodynamic performance of the airfoil 10.

The core 15 extends rearward from the load-bearing structure 11 connecting the tapered composite structure 13 to the load-bearing structure 11. A strong structural connection between the tapered composite structure 13 and the load-bearing structure 11 is important for the airfoil. In particular, the bending moment produced by the aerodynamic forces is relatively large at the front of the tapered composite structure 13, so simply bonding or fastening the tapered composite structure 13 to the load-bearing structure 11 would not be sufficiently strong enough. Where tapered composite structure 13 joins with the load-bearing structure 11, the front of the tapered composite structure 13 is thus made thicker than at the rear of the airfoil in the vicinity of the trailing edge TE. At this location, the piezoelectric actuators 14 can be positioned towards the upper and lower surfaces in order to be most effective.

Since there are a number of different designs for the load-bearing structure 11 of airfoils, there are a number of variations that the core 15 might take. For instance, (i) the core 15 may be comprised of composite plies of the load-bearing structure 11 extending rearward, (ii) the core 15 can be formed of separate composite plies connected to the load-bearing structure 11, (iii) the core 15 may be comprised of a separate piece connected to the load-bearing structure 11, or (iv) the core 15 might also be comprised an integral portion formed on the load-bearing structure 11.

The load-bearing structure 11 may, in some embodiments, be manufactured so as to have a separate piece (or tongue) that extends rearward forming the core 15 of the tapered composite structure 13. For a metal load-bearing structure 11, this tongue can be manufactured through simple machining or welding; but for composite spars, plies can be laid up so that the aft wall of the spar curves from the upper and lower walls into the aft tongue, as shown in FIG. 3. More particularly, if the load-bearing structure 11 is formed of a solid-machined piece, then the core might be integrally machined into this piece and bonded to the tapered composite structure 13. Or, if the load-bearing structure 11 is a layered composite structure, then the layers may be configured to curve back towards the tapered composite structure 13 and form the core 15. Alternatively, the core 15 may be fabricated during the construction of the tapered composite structure 13 and extend forward to provide sufficiently large surfaces that may be bonded or fastened to the load-bearing structure 11.

In designing the airfoil 10, the general design variables are geometry and location of the tapered composite structure 13, airfoil skin stiffness, core material properties, etc. Constraints are maximum stresses and strains, cross-sectional center of gravity, centroid, and shear center locations, total weight, etc. The tapered composite structure 13 length and its location are important variables in determining the aerodynamic characteristics of the airfoil, equivalent to the flap chord size in a conventional discrete trailing-edge flap (DTEF) airfoil. Previous study has determined that a bender located close to middle of the airfoil and extending to the trailing edge is optimal in providing maximum pitching moment with low drag (e.g., equivalent to a 20% DTEF flap chord.). Thus, the location of the tapered composite structure 13 can be fixed at approximately 50% of the chord length C, and extended to the trailing-edge (98% of the chord length C), in some instances. Some analyses and optimization of airfoil designs according to embodiments of the present invention are further discussed in the paper in Appendix A of the provisional application.

EXAMPLE

In one exemplary embodiment, the airfoil 10 is constructed to have a 25 cm (9.84 in) chord C and a span of 12.2 inches. The layout is generally identical to the one shown in FIG. 4. Its tapered composite structure 13 is composed of 28 M8557-P1 MFC actuators and 8 M8528-P1 MFC actuators from Smart Materials Corp. arranged in 4 embed stacked piezoelectric MFC actuator plies 14, above and below, the centerline 17 in the aft half of the airfoil 10. The plies are arrangement symmetrical about the centerline 17. This arrangement will cause the trailing edge TE to bend up or down when the MFCs are actuated.

The MFCs may have an active section that is 85 mm long and 57 mm wide. But the start and end positions of the MFC (as well as their lengths) may not be the same. For instance, as generally illustrated in FIG. 4, the two topmost/bottommost MFC plies may start as far forward as possible near the line L-L, and end short of the trailing edge TE, with the next MFCs plies below/above those plies being staggered somewhat rearward by about 0.075C, ultimately with the innermost MFCs extending closest to the trailing edge TE. (Other arrangements are certainly possible). D1 is approximately 0.189 inches and D2 is 0.047 inches. The skin 12 is 127-μm-thick Nylon film, and the cores 16 filled with Nomex honeycomb between the tapered composite structure 13 and the skin 12. The airfoil surface finish is at least 32 microinches rms. The airfoil 10 may be fabricated using a clam-shell type mold suitable for autoclave insertion as may be used for composite material fabrication of rotor blade articles. The taper of the composite structure 13 for this exemplary embodiment may be characterized according to following formula:

t=6.206−13.56x++7.101x²

where x is the horizontal position of the airfoil (measured from the leading edge LE) and t is the distance (measured vertically, whether above or below) from the centerline 17 of the airfoil 10. The formula is configured for x and t values being in millimeter units. For this exemplary embodiment, at the line L-L, x is 0.5C, and at the trailing edge TE, x is 1.0 C, where C is 25 cm. This embodiment is relative stiff having a blade section twist less than ±0.2° and no noticeable waviness in contour.

The innovative airfoil embodiments disclosed herein lend themselves to various implementations and applications. For instance, the airfoil may be configured for use as a helicopter or tiltrotor rotor blade, a wind turbine blade, an aerodynamic control surface, or the like. In some embodiments, the airfoil can be incorporated into existing helicopter rotor blades (for use in military and/or commercial helicopters) and actuated during flight to reduce vibration and noise. Aerodynamic and finite-element structural analyses have been computed by the inventor for an exemplary helicopter rotor blade according to an embodiment. The designed rotor blade has a 0.35-m chord length which is approximately the size of a light helicopter, like the Bell OH-58D. When actuated at ±1000V, the trailing edge deflects up and down by about 4.89 mm. Even at an airspeed of about Mach 0.65, the trailing edge will deflect by 3.42 mm, which in turn produces a change in lift coefficient of 0.425 and a change in moment coefficient of 0.057. This is comparable to the aerodynamic control produced by conventional active flap design.

The airfoils may be specifically adapted to provide for a Continuous Trailing-Edge Flap (CTEF) airfoil design where the rear of the airfoil 10 near the trailing edge TE functions as a flap. The CTEF will have the ability to perform all of the functions of an active flap on a rotorcraft blade, and can be actively controlled in-flight to reduce vibratory loads, reduce rotor noise, and improve rotor performance. A CTEF can also replace the trailing-edge tabs currently used to keep blades in track. Such a system would reduce the significant maintenance time and cost associated with blade tacking

CTEF devices can also be used for primary flight control, thus eliminating the helicopter swash-plate and reducing the weight and aerodynamic drag of the rotor hub. The CTEF is not only applicable to helicopters, but all rotorcraft. The airfoils also can be used to maintain rotor blade track, thereby reducing maintenance time and cost. The use of active CTEFs to reduce vibration, and thus improve passenger comfort, is also of interest. Swashplateless rotors can be developed that use CTEFs on each blade for primary flight control rather than a traditional swashplate. This would reduce the aerodynamic drag of the rotorcraft, thereby improving its fuel efficiency and maximum speed.

The CTEF airfoil construction also has the advantages of the active-twist rotor, in that the use of embedded actuators creates a monolithic structure free from mechanical linkages and having a smooth continuous aerodynamic profile. Also, since the flap structure is typically small relative to the length of the blade, fewer actuators are required to build the airfoil. The flap is also more scalable than the active-twist concept, meaning that it can easily be applied to both very small flaps used just for blade tracking and large flaps used for primary flight control.

The foregoing description, for purpose of explanation, has been described with reference to specific embodiments. However, the illustrative discussions above are not intended to be exhaustive or to limit the invention to the precise forms disclosed. Many modifications and variations are possible in view of the above teachings. The embodiments were chosen and described in order to best explain the principles of the present disclosure and its practical applications, to thereby enable others skilled in the art to best utilize the invention and various embodiments with various modifications as may be suited to the particular use contemplated.

While the foregoing is directed to embodiments of the present invention, other and further embodiments of the invention may be devised without departing from the basic scope thereof, and the scope thereof is determined by the claims that follow. All references cited above are hereby incorporated by reference herein for all purposes.

Claims

1. An airfoil comprising:

a load-bearing structure in the forward portion of the airfoil;

a skin to maintain the airfoil shape;

a tapered composite structure having one or more actuators extending from the load-bearing structure toward the trailing edge, and joining the skin in the vicinity of the trailing edge, the one or more actuators being configured to deflect the trailing edge so as to deform the shape of the airfoil cross-section; and

a core extending from the load-bearing structure connecting the tapered composite structure to the load-bearing structure, wherein the tapered composite structure tapers from being relatively thick near the core to being relatively thin near the trailing edge.

2. The airfoil of claim 1, wherein the skin is composed of one or more discrete flexible sections above and below the tapered composite structure.

3. The airfoil of claim 1, wherein the load-bearing structure is composed of composite plies or an integral structure.

4. The airfoil of claim 1, wherein the core comprises composite plies of the load-bearing structure extending rearward.

5. The airfoil of claim 1, wherein the core comprises separate composite plies connected to the load-bearing structure.

6. The airfoil of claim 1, wherein the core comprises a separate piece connected to the load-bearing structure.

7. The airfoil of claim 1, wherein the core comprises an integral portion formed on the load-bearing structure.

8. The airfoil of claim 1, further comprising honeycomb or foam cores, above and below, the tapered composite structure.

9. The airfoil of claim 1, wherein the one or more actuators are bonded and/or attached via adhesive, glue, and/or fasteners.

10. The airfoil of claim 1, wherein the one or more actuators are formed of multiple layers of the piezoelectric actuators connected together.

11. The airfoil of claim 10, wherein the piezoelectric actuators are configured as a bimorph.

12. The airfoil of claim 1, wherein two or more actuators are provided.

13. The airfoil of claim 12, at least one actuator is provided at the top and the bottom of the tapered composite structure, respectively.

14. The airfoil of claim 1, wherein the tapered composite structure is attached to top and bottom portions of the core, respectively.

15. The airfoil of claim 1, wherein, near the core, the thickness of the tapered composite structure is approximately 2% of the chord length of the airfoil, and near the trailing edge, the thickness of the tapered composite structure is about 0.3% of the chord length.

16. The airfoil of claim 1, wherein the taper of the tapered composite structure is linearly tapering.

17. The airfoil of claim 1, wherein the degree of tapering at the top and bottom of the tapered composite structure is different.

18. A helicopter or tiltrotor rotor blade comprising the airfoil of claim 1.

19. A wind turbine blade comprising the airfoil of claim 1.

20. An aerodynamic control surface comprising the airfoil of claim 1.