阅读说明：本技术 具有稳定翼的旋翼飞行器 (Rotorcraft with stabilizer blades ) 是由 马丁·恩巴赫 托比亚斯·里斯 克里斯蒂安·埃克特 托马斯·克奈施 于 2020-04-23 设计创作，主要内容包括：本发明涉及一种旋翼飞行器,特别是涉及一种其包括具有中心线(230)的机身、在运行期间产生涡流的至少一个主旋翼以及稳定翼(200)的旋翼飞行器,其中稳定翼(200)具有使至少一个主旋翼的尾流产生的非定常气动载荷减小的平面形状。特别地,稳定翼(200)可具有：左翼尖(260)；右翼尖(260)；具有非零曲率的四分之一翼弦线(240),使得至少一个主旋翼产生的涡流与该四分之一翼弦线(240)之间的相互作用随时间发展而扩散；弧形的前缘210；以及弧形的后缘220。(The invention relates to a rotorcraft, in particular a rotorcraft comprising a fuselage with a centre line (230), at least one main rotor that generates vortices during operation, and a stabilizer (200), wherein the stabilizer (200) has a plan shape that reduces unsteady aerodynamic loads generated by the wake of the at least one main rotor. In particular, the stabilizer (200) may have: a left wingtip (260); a right wing tip (260); a quarter chord line (240) having a non-zero curvature such that the interaction between the vortices generated by the at least one main rotor and the quarter chord line (240) spreads as time progresses; a curved leading edge 210; and an arcuate trailing edge 220.)

1. A rotary-wing aircraft (100), comprising:

a fuselage (104) having a centerline (130);

at least one main rotor (102) that generates vortices (460, 470, 480) during operation; and

a stabilizer wing (133) connected to the fuselage (104),

characterized in that the stabilizer blade (133, 400) has:

a left wing tip (160, 457);

a right wing tip (160, 457);

a quarter chord line (440) having a non-zero curvature, the non-zero curvature of the quarter chord line (440) causing a reduction in unsteady aerodynamic loads on the stabilizer (133, 400) generated by the wake of the at least one main rotor (102) such that the interaction between the vortex (460, 470, 480) generated by the at least one main rotor (102) and the quarter chord line (440) spreads over time;

A curved leading edge (410); and

an arcuate trailing edge (420).

2. The rotary wing aircraft (100) according to claim 1, wherein the fuselage (104) extends in a longitudinal direction (X) between a nose section (108) and a tail section (109), and wherein the quarter chord line of the stabilizer wing (133) is closer to the tail section (109) at least one of the left wing tip or the right wing tip (160) than at the centerline (130) of the fuselage (104).

3. The rotary wing aircraft (100) according to claim 1, wherein the fuselage (104) extends in a longitudinal direction (X) between a nose section (108) and a tail section (109), and wherein the quarter chord line of the stabilizing wing (133) is closer to the nose section (108) at least one of the left wing tip or the right wing tip (160) than at the centerline (130) of the fuselage (104).

4. The rotary wing aircraft (100) according to any preceding claim, wherein an absolute value of the non-zero curvature of the quarter chord line (240) is greater than 0.2, wherein the quarter chord line (240) is a function and the non-zero curvature is a second derivative of the function.

5. The rotary-wing aircraft (100) of claim 4, wherein the absolute value of the non-zero curvature of the quarter-chord line (240) is less than 4.

6. The rotary wing aircraft (100) according to any preceding claim, wherein the non-zero curvature of the quarter chord line (240) progresses from a first value at a first distance from the centerline (230) of the fuselage (104) to a second value at a second distance from the centerline (230) of the fuselage (104) such that the non-zero curvature of the quarter chord line (240) is not constant between the first and second distances from the centerline (230) of the fuselage (104).

7. The rotary wing aircraft (100) according to any one of the preceding claims, wherein the stabilizer wing (200) has a half-span (257, 258) between the center line (230) and the left or right wing tip (260), wherein a first point (272), a second point (274), a third point (276) and a fourth point (278) are located at a first distance (282), a second distance (284), a third distance (286) and a fourth distance (288), respectively, from the center line (230) of the fuselage on the quarter-chord line (240), wherein the first distance (282), the second distance (284), the third distance (286) and the fourth distance (288) are 10%, 30%, 50% and 70% of the half-span (257, 258), respectively, wherein the non-zero curvature of the quarter-chord line (240) is not zero between the first point (272) and the fourth point (278), and wherein a straight line (290) between the first point (272) and the fourth point (278) is a fifth distance (270) from a quarter chord line (240) between the second point (274) and the third point (276), the fifth distance being greater than 2% of the semi-span (257, 258).

8. The rotary wing aircraft (100) according to any preceding claim, wherein the quarter chord line (240) is symmetrical with respect to the centerline (230) of the fuselage.

9. The rotary wing aircraft (100) according to any preceding claim, wherein the stabilizer wings (133) are mounted to at least one of the fuselage (104), a tail boom (121) attached to the fuselage (104), or a vertical tail fin (131) attached to the tail boom (121).

10. The rotary wing aircraft (100) according to any preceding claim, wherein the stabilizer (400) has a constant chord length (450) between the left and right tips (457) of the stabilizer (400).

11. The rotary wing aircraft (100) according to any one of claims 1 to 9, wherein a chord length (250) of the stabilizer (200) decreases from the centerline (230) of the fuselage towards the left and right wingtips (260) of the stabilizer (200).

12. The rotary-wing aircraft (100) according to any of claims 1-7, wherein the stabilizer wing (500) further comprises:

a left wing half (504) having a first sweep; and

a right wing half (502) having a second sweep different from the first sweep.

13. The rotary wing aircraft (100) of claim 12, wherein the left wing half (504) has a quarter chord line (540) with a first curvature, wherein the right wing half (502) has a quarter chord line (540) with a second curvature, and wherein the second curvature is different than the first curvature.

14. The rotary wing aircraft (100) of claim 12, wherein the left wing half (504) has a first length and the right wing half (502) has a second length different from the first length.

15. The rotary wing aircraft (100) according to any preceding claim, wherein the stabilizer wing (133) further comprises a winglet (138).

16. The rotary-wing aircraft (100) according to any preceding claim, wherein the rotary-wing aircraft is embodied as a helicopter.

Technical Field

The present invention relates to a rotorcraft, and in particular to a rotorcraft having at least one main rotor and a stabilizer wing, wherein the stabilizer wing has a planform that reduces unsteady aerodynamic loads thereon generated by wakes of the at least one main rotor.

Background

Rotorcraft are typically equipped with one or more wings that help improve the stability and maneuverability of the rotorcraft during flight. Accordingly, such wings are sometimes also referred to as horizontal stabilizers, horizontal stabilizers or stabilizers.

Typically, helicopters have one or more stabilizer wings disposed at the aft end of the tail unit. In some models, the stabilizer projects from the rear of the tail boom or from the vertical tail plane. In other models, a single stabilizer is provided on top of the vertical tail. This configuration, in which a single stabilizer is provided on top of the vertical tail, is sometimes referred to as a T-tail.

Since the stabilizer wings are located in the tail section of the rotorcraft, they are generally affected by the structure of the air flow generated by at least one main rotor. These airflow structures, sometimes referred to as wakes or rotor wakes, interact with and determine the aerodynamic loads on the stabilizer.

The strength of the interaction between the rotor wake and the stabilizer depends on the positioning of the stabilizer relative to the rotor and the flight conditions (e.g., during take-off and landing or during horizontal flight). However, it is generally not avoidable to generate aerodynamic loads on the stabilizer wing which have an influence on the structural dimensions of the stabilizer wing.

The stabilizer wings intentionally create a certain amount of static aerodynamic loading (e.g., to improve stability and handling of the rotorcraft during flight). However, undesirable unsteady aerodynamic loads may also be generated by the interaction between the rotor wake and the stabilizer.

These unsteady aerodynamic loads sometimes reach large magnitudes, which may lead to fatigue failure of the stabilizer blade. Therefore, during the development of a stabilizer blade, special care is usually taken to ensure a structural design that significantly reduces the risk of fatigue failure due to undesired unsteady aerodynamic loads.

The problem is often exacerbated by the fact that one of the dynamic natural frequencies of the stabilizer and tail boom is close to the excitation frequency of the rotor wake. Such a resonance condition between the rotor wake and the stabilizer can be avoided by increasing the structural stiffness or mass of the stabilizer.

However, increasing the structural stiffness or mass of the stabilizer increases the weight of the sensitive area of the rotorcraft as the center of gravity moves further aft. In particular, when the stabilizer is provided on the vertical stabilizer, the problems of structural and dynamic properties are difficult to solve. As a result, the stabilizer is rarely arranged on the vertical stabilizer, which would otherwise be very advantageous for the quality and performance of the manoeuvre.

Instead, most conventional solutions provide the stabilizer wings in a position vertically below the rotor plane (e.g. at the tail boom). As a result, the interaction between the rotor wake and the stabilizer mainly occurs at lower flight speeds. Since aerodynamic loads increase with increasing flight speed, the interaction between the rotor wake and the stabilizer at lower flight speeds will produce relatively moderate unsteady aerodynamic loads while the static aerodynamic loads on the stabilizer are relatively low. Thus, the structure of the stabilizer does not reach its fatigue limit at lower flight speeds, since it is influenced by the superposition of static aerodynamic loads and unsteady aerodynamic loads.

At higher flight speeds, the rotor wake changes its trajectory and passes over a low-lying stabilizer (i.e. a stabilizer located, for example, at the tail boom). Thus, although the static aerodynamic loads on the stabilizer are greater at higher flight speeds than at lower flight speeds, the relatively small magnitude of the unsteady aerodynamic loads of the low-lying stabilizer has the effect of keeping the total aerodynamic load well within the fatigue limit.

However, low-profile stabilizers exhibit reduced handling qualities and performance as compared to high-profile stabilizers (e.g., stabilizers at the top of the vertical tail). For example, during hover and/or at lower flight speeds, the low-profile stabilizer wings are positioned within the rotor downwash and generate a force directed toward the ground, sometimes referred to as downforce. This downforce causes the rotorcraft to pitch up the nose, which is sometimes referred to as nose pitch-up.

Pitching up the head severely reduces the driver's view of the ground. Nose pitch up also produces relatively high stresses in the rotor mast. Furthermore, downforce can significantly reduce the payload of the rotorcraft. For example, downforce may reduce the payload of a three to four ton class helicopter by an amount equivalent to one passenger.

Furthermore, it is often necessary to install low-level stabilizing wings at the tail boom. This results in a shorter lever arm compared to the high-position stabilizer, which reduces the stabilizing effect.

Some helicopter manufacturers attempt to overcome the dilemma between maneuverability and performance on the one hand and flight stability on the other by employing high-attitude stabilizers. However, as mentioned above, these high-attitude stabilizer wings face the problem of strong unsteady aerodynamic loads caused by interaction with the rotor wake in cruising flight.

Since unsteady aerodynamic loads present serious structural problems, only a relatively small number of helicopters have been designed to date with high-attitude stabilizer wings. Typically, helicopters with high-attitude stabilizers include additional methods for reducing unsteady aerodynamic loads. Among these methods are tapering the stabilizer (i.e. the chord length of the stabilizer is longer at the root than at the tip) or including structural reinforcement. These reinforcements usually consist of thickening the stabilizer at its connection to the tail boom or vertical tail plane, or of installing struts which support the stabilizer from the outside.

For example, the T-tail (i.e., the stabilizer on the vertical tail) of the AV-02 Husky YAH-64 advanced attack helicopter has a swept back (i.e., the stabilizer slopes backwards from the root to the tip) and is tapered. The atubel model BA609 has a tapered T-shaped tail with stabilizers having straight leading and trailing edges. The T-tail of the komanchi model RAH-66 has straight stabilizers (i.e. stabilizers with straight leading and trailing edges but no sweep at the top of the vertical tail). The T-shaped tail of the Kamof Ka-60 type has straight stabilizers and struts for the vertical tail.

Both the tapering and stiffening of the stabilizer are disadvantageous from an aerodynamic point of view, because they increase drag, reduce the efficiency of the stabilizer and reduce the range of flight conditions over which the stabilizer can stabilize the rotorcraft.

Document EP2899118a1 describes a rotorcraft having a fuselage and at least one main rotor that can be driven for controlling an associated pitch attitude of said rotorcraft in operation, and the fuselage being equipped with at least one passive airfoil aerodynamic device adapted to generate, independently of said associated pitch attitude, a lift force acting on the fuselage oriented perpendicularly to an airflow directed towards said passive airfoil aerodynamic device in operation of said rotorcraft. The rotorcraft further comprises a vertical tail with a rudder and a tail in the form of a T-tail arranged at the tail boom of the fuselage. The tail fin can be adjusted in inclination and serves as an additional lifting surface.

Document CN103979105A describes a variable wing aircraft of the vertical take-off and landing type. The vertical take-off and landing type variable wing aircraft comprises an aircraft body, a main wing, a left wing and a right wing, wherein the main wing is arranged at the lower part of the aircraft body; the left wing and the right wing are respectively connected to two ends of the main wing; a left wing rotating shaft sleeve and a right wing rotating shaft sleeve are arranged in the main wing; a group of left wing rotating shafts in the left wing rotating shaft sleeve is arranged in the left wing; a group of right wing rotating shafts in the right wing rotating shaft sleeve is arranged on the right wing; a left wing steering engine and a right wing steering engine are also arranged in the main wing; the left wing steering engine is used for driving the left wing rotating shaft to adjust the direction of the left wing; the right wing rudder machine is used for driving the right wing rotating shaft to adjust the direction of the right wing. The vertical take-off and landing type variable wing aircraft can realize vertical take-off and landing of model airplanes and unmanned aircrafts, quickly realize various flight states and effectively brake and move backwards in flight.

Document EP2666719B1 describes an aircraft comprising a fuselage, at least one main rotor having a plurality of blades, at least one variable-pitch propeller, at least one airfoil and at least one motor rotationally driving the main rotor and each propeller. The wing profile may comprise two wing halves on either side of the fuselage and at least one horizontal stabilizer at one end of the aircraft and having at least one moving surface.

Document WO1999/067130a1 describes a horizontal stabilizer defining a first spanwise standing position and a second spanwise standing position, wherein the first spanwise standing position defines a first erection angle and the second spanwise standing position defines a second erection angle, and wherein the erection angles are different from one standing position to another, e.g. one erection angle is larger than the other. The horizontal stabilizer is used to advantageously affect the spanwise lift distribution to reduce bending moments around its mounting interface. Various embodiments of the horizontal stabilizer include the use of vertically extending tabs along the trailing edge of the horizontal stabilizer, a stepped transition that abruptly changes the setting angle from one station to another, and a distributed twist that gradually changes the setting angle.

Document EP0254605a1 describes a orienting and stabilizing device comprising a ducted tail rotor driven in rotation in a transversal slit formed in a fairing inclined at an angle between 0 ° and 45 ° with respect to the vertical. It also includes a tail fin "chevrons" attached to the shroud top and two aerodynamic surfaces extending asymmetrically with respect to a vertical plane passing through the fairing apex. The aerodynamic surface is inclined in an angular range between 0 ° and 45 °.

Document US3464650A describes an aircraft with a winged rotor/wing for vertical and short take-off and landing (V/STOL) performance. The aircraft has a horizontal stabilizer at the top of the vertical tail, which may be of the fully movable type or may have independently movable parts. These separate sections act as conventional elevators or may be connected to act as ailerons for roll control in a dual function elevating surface when flying forward. Conventional aircraft controls are used to operate the horizontal stabilizer and rudder on the vertical tail to control when flying forward.

Other stabilizer wings are described by way of example in documents CN106516082A, US2016/0031554A1, WO2005/005250A2, US5738301A and US 3902688A. These documents each describe a stabilizer in a rotorcraft, which is at least similar to the above-mentioned stabilizer. Documents EP2409917 and WO2009155584 have been cited.

In summary, the interaction between the rotor wake and the horizontal stabilizer in a rotorcraft depends on the flight mode (i.e., hovering, low or high forward flight) and the position of the stabilizer on the rotorcraft (i.e., low or high), which affects maneuverability, performance, and flight stability. High stability wings provide better maneuverability, performance and flight stability than low stability wings. However, as mentioned above, high-attitude stabilizer wings face the problem of strong unsteady aerodynamic loads generated by interaction with the rotor wake in cruising flight.

Disclosure of Invention

It is therefore an object to provide a rotorcraft having a fuselage, at least one main rotor, and a stabilizer wing. The stabilizer should reduce unsteady aerodynamic loads generated by rotor wake interaction with at least one main rotor while providing improved handling, performance, and flight stability as compared to conventional stabilizers.

The above object is achieved by a rotorcraft comprising the features of claim 1. More specifically, a rotary wing aircraft may include a fuselage having a centerline, at least one main rotor that generates vortices during operation, and a stabilizer connected to the fuselage. The stabilizer has: a left wingtip; a right wingtip; a quarter chord line having a non-zero curvature such that the interaction between the vortices generated by the at least one main rotor and the quarter chord line spreads as time progresses; a curved leading edge; and an arcuate trailing edge.

Advantageously, in this rotorcraft with at least one main rotor, the above-mentioned stabilizer wing with a quarter chord line of non-zero curvature reduces the unsteady aerodynamic loads on the stabilizer created by the wake of the at least one main rotor. Such a stabilizer may be used with any rotorcraft having at least one main rotor, including a vertical take-off and landing aircraft, a multi-rotor helicopter, a drone, and the like.

The stabilizer has a specific planar shape (i.e., a specific shape when viewed from above). The particular planform of the stabilizer reduces unsteady loads generated by rotor wake interaction with the at least one main rotor. In contrast to the prior art shape of the stabilizer airfoil, which is mostly rectangular or consists of two trapezoids, a shape is proposed in which the leading and trailing edges have a curvature.

Thus, the quarter chord line of the stabilizer is curved. In other words, the line connecting all points located at a quarter of the chord length, which is the distance from the leading edge to the trailing edge of the stabilizing wing along the chord, is curved. Having curved quarter chord lines advantageously changes the response of the stabilizer to rotor wake fluctuations as compared to a stabilizer having straight quarter chord lines.

In fact, the rotor wake exhibits a regular airflow pattern in cruising flight, and this pattern determines the time sequence of the load conditions at the stabilizer wings. The curved quarter chord line alters the relationship between the stabilizer and rotor wake flow patterns by reducing the magnitude of unsteady loads. As a result, the stabilizer becomes less susceptible to transient load conditions of the type that generate high stresses at the stabilizer's connection to the rotorcraft (e.g., at the tail boom or vertical tail) where the stabilizer is sensitive to structural changes. In particular, the oscillating moments of the stabilizer about the longitudinal axis of the rotorcraft benefit from a reduction in the interaction between the rotor wake and the stabilizer, since these oscillating moments are most critical to the connection of the stabilizer to the rotorcraft from a structural point of view.

Since the proposed new planar shape of the stabilizer substantially reduces the unsteady aerodynamic loads originating from the attachment location of the stabilizer, the attachment connection of the stabilizer can be designed more weight and space saving (i.e. with less potential adverse effect on the aerodynamic shape of the interconnection area between the stabilizer and the tail boom or the vertical tail) compared to the attachment connection of conventional stabilizers.

Significant weight savings are possible due to the reduction of unsteady aerodynamic loads. For example, the stabilizer blade may require less structural reinforcement. As another example, the strength and weight of the structure to which the stabilizer is attached (e.g., tail boom or vertical tail fin) and the structure to which the stabilizer transmits unsteady aerodynamic loads may also be structurally reduced.

The vertical tail of a rotorcraft is typically a structurally weak component. A stabilizer wing with a curved quarter chord line that significantly reduces unsteady loads generated by rotor wakes when mounted to a vertical tail wing is therefore particularly advantageous. In fact, the aerodynamic interaction between the rotor wake and the stabilizer occurs at high speeds and the unsteady loads at the stabilizer usually reach maximum amplitudes when the stabilizer is mounted on the vertical tail.

Furthermore, the time that the stabilizer is exposed to these high unsteady loads generated by the rotor wake is relatively large during rotorcraft runtime. The large distance between the stabilizer and the fuselage makes the vertical tail attachment behave as a kind of joint which responds to stabilizer loads in case of large deflections and exhibits a low natural frequency, which is difficult to correct when needed from a structural dynamics point of view.

A stabilizer with a curved quarter chord line can achieve a decisive unsteady load reduction, which is necessary for achieving an otherwise challenging design of a stabilizer mounted on a vertical tail. As described in the disadvantages of the prior art, if a stabilizer having a curved quarter chord line is mounted to a vertical tail wing, several important benefits in handling quality and performance can be realized.

According to one aspect, the fuselage extends in a longitudinal direction between a nose section and a tail section, and the quarter chord line of the stabilizing wing is closer to the tail section at least one of the left wing tip or the right wing tip than at the centerline of the fuselage.

According to one aspect, the fuselage extends in a longitudinal direction between a nose section and a tail section, and the quarter chord line of the stabilizing wing is closer to the nose section at least one of the left wing tip or the right wing tip than at a centerline of the fuselage.

According to one aspect, the absolute value of the non-zero curvature of the quarter chord line is greater than 0.2.

According to one aspect, the absolute value of the non-zero curvature of the quarter chord line is less than 4.

According to one aspect, the non-zero curvature of the quarter-chord line progresses from a first value at a first distance from the centerline of the fuselage to a second value at a second distance from the centerline of the fuselage such that the non-zero curvature of the quarter-chord line is not constant between the first distance and the second distance from the centerline of the fuselage.

According to one aspect, the stabilizer wing has a semi-span between the center line and the left or right wing tip, wherein the first, second, third and fourth points are located at first, second, third and fourth distances, respectively, from the center line of the fuselage on a quarter-chord line, wherein the first, second, third and fourth distances are 10%, 30%, 50% and 70% of the semi-span, respectively, wherein the non-zero curvature of the quarter-chord line is non-zero between the first and fourth points, and wherein a straight line between the first and fourth points is a fifth distance from the quarter-chord line between the second and third points that is greater than 2% of the semi-span.

According to one aspect, the quarter chord line is symmetrical with respect to a centerline of the fuselage.

According to one aspect, the stabilizer is mounted to at least one of the fuselage, a tail boom attached to the fuselage, or a vertical tail fin attached to the tail boom.

According to one aspect, the stabilizer has a constant chord length between a left tip and a right tip of the stabilizer.

According to one aspect, the chord length of the stabilizer decreases from the centerline of the fuselage towards the left and right wingtips of the stabilizer.

According to one aspect, the stabilizer wing further includes a left half wing having a first swept portion and a right half wing having a second swept portion different from the first swept portion.

According to one aspect, the left wing half has a quarter chord line with a first curvature, the right wing half has a quarter chord line with a second curvature, and the second curvature is different from the first curvature.

According to one aspect, the left wing half has a first length and the right wing half has a second length different from the first length.

According to one aspect, the stabilizer further comprises a winglet.

According to one aspect, the rotorcraft may be implemented as a helicopter.

Drawings

Embodiments are summarized by way of example in the following description with reference to the drawings. In these figures, identical or functionally identical parts and elements are denoted by the same reference numerals and characters and are therefore described only once in the following description.

Figure 1A is a view of an exemplary rotorcraft with stabilizer wings according to some embodiments,

figure 1B is a simplified side view of an exemplary rotorcraft according to some embodiments,

figure 1C is a simplified top view of an exemplary rotorcraft according to some embodiments,

FIG. 2 is a view of an exemplary stabilizer wing with a quarter chord line of non-zero curvature according to some embodiments,

FIG. 3A is a view of an exemplary interaction between the rotor tip vortex centerline at a quarter chord line with non-zero curvature to the left half of a stabilizer wing according to some embodiments,

FIG. 3B is a view of an exemplary interaction between a rotor tip vortex centerline with a quarter chord line of non-zero curvature across the left half of a stabilizer wing according to some embodiments,

FIG. 3C is a view of an exemplary interaction between the rotor tip vortex centerline at the intersection between the quarter chord line with non-zero curvature and the centerline of the fuselage to a stabilizer wing according to some embodiments,

FIG. 4 is a view of a series of rotor tip vortex centerlines to an exemplary stabilizer wing with constant chord length and quarter chord line of non-zero curvature according to some embodiments,

FIG. 5 is a view of a series of rotor tip vortex centerlines to an exemplary stabilizer wing with asymmetric sweep between left and right wing halves according to some embodiments, and

FIG. 6 is a view of an exemplary stabilizer wing with a tapered chord length, quarter chord line with non-zero curvature, and wing centerline perpendicular to a series of rotor tip vortex centerlines according to some embodiments.

Detailed Description

Fig. 1A, 1B, and 1C show a rotorcraft 100 having a fuselage 104 and a main rotor 102. Fig. 1A shows a three-dimensional representation of rotorcraft 100, while fig. 1B shows a simplified side view of rotorcraft 100 and fig. 1C shows a simplified top view of rotorcraft 100.

The rotorcraft 100 is illustratively embodied as a helicopter, and the main rotor 102 is illustratively embodied as a hingeless or hingeless and bearingless multi-bladed rotor having a plurality of rotor blades 102a, 102b, 102c, 102d, 102 e.

It should be noted, however, that the proposed embodiments are not limited to helicopters and can equally be applied to other rotorcraft equipped with rotary wings, irrespective of whether these rotary wings define articulated, hingeless or hingeless and bearingless multi-bladed rotors. It should also be noted that the proposed embodiment can also be applied in case more than one main rotor is provided.

In fact, exemplary embodiments may be included in any rotorcraft having at least one main rotor and a stabilizer wing that should reduce unsteady aerodynamic loads on the stabilizer wing that are generated by the wake of the at least one main rotor. Examples of such vehicles may include rotorcraft, such as vertical take-off and landing aircraft, multi-rotor helicopters, drones, and the like.

Illustratively, rotorcraft 100 may have a fuselage 104 that forms the backbone of rotorcraft 100. The fuselage 104 may be connected to suitable landing gear and rear fuselage. For example, the landing gear may be a skid landing gear as shown in FIG. 1A. As another example, the landing gear may have wheels as shown in FIG. 1B. The rear fuselage may be connected to the tail boom 121. Fuselage 104 illustratively forms a nacelle 104a that defines a fuselage head 107 at a nose portion 108 of rotorcraft 100.

Fig. 1B and 1C show rotorcraft 100 in a cartesian coordinate system having an X-axis, a Y-axis, and a Z-axis. The X-axis extends in the longitudinal direction of rotorcraft 100 through nose section 108 and tail section 109, and is sometimes also referred to as the longitudinal or length axis.

The Y-axis extends in the lateral direction of rotorcraft 100 through left and right wingtips 160, 160 of stabilizer wings 133, and is sometimes also referred to as the lateral, side, or width axis. The Z-axis extends in the vertical direction of rotorcraft 100 and is sometimes also referred to as the vertical or altitude axis.

The fuselage 104 may have a centerline 130. The centerline 130 may be parallel to or coincident with the X-axis passing through the main rotor center 103. In other words, a plane defined by the centerline 130 and the Z-axis may define an axis of symmetry of the fuselage 104 in the lateral direction.

For example, rotorcraft 100 may include at least one counter-moment device configured to provide a counter-moment during operation, i.e., a moment generated by rotation of at least one multi-bladed rotor 102 in order to balance rotorcraft 100 in yaw. The counter-torque device may be covered if desired. At least one counter-torque device is illustratively provided at the tail section 109 of the rotorcraft 100 and may have a tail rotor 122.

If desired, tail section 109 of rotorcraft 100 may include a vertical tail fin 131 attached to tail boom 121. In some embodiments, the vertical tail 131 may be provided with a rudder 132. The rudder 132 may be adapted to provide enhanced directional control and optimized yaw trim of the rotorcraft 100. The rudder 132 can be deflected to a greater angle to reduce the lateral drag on the vertical tail 131 during lateral flight, if desired.

Illustratively, vertical tail 131 may be provided with a suitable horizontal stabilizer 133 in the form of a T-shaped tail. In other words, as shown in FIGS. 1A and 1B, horizontal stabilizer 133 may be connected to fuselage 104 by vertical stabilizer 131 and tail boom 121. As shown, stabilizer 133 may be mounted to the top of vertical stabilizer 131.

Stabilizer 133 may be installed at another location on rotorcraft 100 if desired. For example, stabilizer 133 may be mounted directly to fuselage 104, tail boom 121, and/or vertical tail wing 131 at a location below the top of vertical tail wing 131 in height axis Z.

Stabilizer 133 may have left and right wingtips 160, curved leading edges 136, and curved trailing edges 137.

For example, the stabilizer wing 133 may have a constant chord length. In other words, the chord of stabilizer 133, which is an imaginary straight line connecting leading edge 136 and trailing edge 137 of stabilizer 133, may have a constant length for each point on leading edge 136. Stabilizing wing 133 may have a tapered chord length, if desired. In other words, the chord length of the stabilizing wing 133 may be greater at the centerline 130 of the fuselage 104 than at the left and right wingtips 160, 160.

Illustratively, the quarter chord line of the stabilizing wing 133, which is an imaginary line connecting all points on different chords that are one-fourth of the respective chord length from the leading edge 136, may have a non-zero curvature.

As shown in FIGS. 1A and 1C, the quarter chord line of stabilizer wing 133 is curved aft. In other words, when a connecting line is drawn between any two points on the quarter chord line of the stabilizing wing 133, the quarter chord line segment between the any two points is closer to the nose section 108 than the connecting line.

The quarter chord line of the stabilizer wing 133 may be curved forward if desired. In other words, when a connecting line is drawn between any two points on the quarter chord line of the stabilizer wing 133, the quarter chord line segment between the any two points is closer to the aft section 109 than the connecting line.

The stabilizing wing 133 may have a vertical or near vertical extension at the left wing tip 160 and the right wing tip 160, if desired. Sometimes, such a vertical or near vertical extension at the tip is also referred to as a winglet 138. As shown in fig. 1A, the stabilizer 133 has a downwardly angled winglet 138. However, the stabilizing wing 133 may have an upwardly angled winglet 138 or may have an upwardly and downwardly extending winglet 138 if desired.

The main rotor 102, and thus the plurality of rotor blades 102a, 102b, 102c, 102d, 102e, is drivable, i.e., controllable, to affect the associated pitch attitude of the rotorcraft 100 in operation. Illustratively, main rotor 102 defines a main rotor center 103 and includes a rotor mast. The mast has a mast axis (e.g., mast axis 103a of fig. 1B) that defines an axis of rotation of main rotor 102.

During operation, the main rotor 102 may generate airflow structures for providing lift and forward or aft thrust. This airflow configuration is sometimes referred to as rotor wake or wake. The rotor wake of the main rotor 102 may interact with the stabilizer wings 133. In particular, the rotor wake may apply aerodynamic loads to stabilizer wings 133.

Aerodynamic loads on stabilizer 133 may be divided into static aerodynamic loads and unsteady aerodynamic loads. Static aerodynamic loading is desirable because it improves flight stability and performance. Unsteady aerodynamic loading is undesirable because it creates structural stresses in stabilizer wing 133 and the portion of rotorcraft 100 to which stabilizer wing 133 is attached.

Unsteady aerodynamic loads on stabilizer wings 133 are caused by changes in airflow velocity and/or airflow direction and may result in changes in lift at stabilizer wings 133. These changes in lift on stabilizer wing 133 occur in response to changes in airflow at the quarter chord line of stabilizer wing 133.

Significant changes in airflow velocity and direction occur in the rotor wake in the form of vortices. A vortex is a quantity of air that rotates about an elongated, generally curved axis (the center of the vortex). In the rotorcraft 100, the vortices shed from the tips of the rotor blades 102a, 102b, 102c, 102d, 102 e.

Due to the rotation of the rotor blades and the simultaneous forward movement in the flight direction, the axial shape of the rotor tip vortex, as seen from above, resembles a helix. The screw is moved backwards relative to rotorcraft 100 during forward flight at almost the same speed as the flight speed while being generated at the tip of the rotor blade. A particular portion of the rotor blade tip vortex will pass through the stabilizing wing 133 and create an unsteady aerodynamic load.

Since the changes in lift on stabilizer wing 133 occur in response to changes in airflow at the quarter chord line of stabilizer wing 133, unsteady aerodynamic loads on stabilizer wing 133 can be approximated by the interaction between the rotor blade tip vortex and the quarter chord line of stabilizer wing 133.

As described above, the quarter chord line of stabilizer wing 133 has a non-zero curvature. Thus, the interaction between the vortex generated by the at least one main rotor 102 and the quarter chord line may spread over time as compared to a stabilizer wing with a straight quarter chord line.

In fact, due to the helical flow pattern, the rotor blade tip vortex periodically passes through the quarter chord line of the stabilizer 133. For a rotor with N rotor blades, N helical rotor tip vortices are generated, so that N rotor blade tip vortices pass the stabilizer 133 per rotor revolution. Thus, unsteady aerodynamic loads at stabilizer wings 133 are generated periodically at fundamental frequency N for each rotor revolution.

The shape of the cyclic unsteady aerodynamic loads of the stabilizer 133 at the attachment at the vertical tail fin 131 or the tail boom 121 may have a smoother peak shape and reduced amplitude due to the curved shape of the quarter chord line compared to a stabilizer having a straight quarter chord line. This is further illustrated in fig. 3A to 3C.

FIG. 2 illustrates an exemplary stabilizer wing having a quarter chord line of non-zero curvature according to some embodiments. As shown in FIG. 2, the stabilizer wing 200 may have a curved leading edge 210, a curved trailing edge 220, and a quarter chord line 240 having a non-zero curvature.

The term "curved" as used with respect to the leading edge 210 means that the leading edge 210 is at least partially curved between the tips 260. Preferably, the leading edge 210 is curved at least near the centerline 230. The curved leading edge 210 may be partially flat if desired.

Similarly, the term "curved" as applied to the trailing edge 220 means that the trailing edge 220 is at least partially curved between the tips 260. Preferably, the trailing edge 220 is curved. The curved trailing edge 220 can be partially straight if desired.

By way of example only and not to limit the presented embodiments accordingly, curved objects include objects that are at least partially C-shaped, objects that are shaped like parabolic segments, objects that are shaped like elliptical segments, semi-elliptical objects, objects that are circular in some portions and flat in other portions, and the like.

As shown in FIG. 2, the span 255 of the stabilizer 200 may be the distance between the left tip 260 and the right tip 260. The span 255 may be the sum of a left semi-span 257 and a right semi-span 258, where the left semi-span 257 may be the distance between the left wing tip 260 and the centerline 230 and the right semi-span 258 may be the distance between the right wing tip 260 and the centerline 230.

Illustratively, the stabilizing wing 200 may have a chord length 250 that decreases from the centerline 230 of the fuselage toward the left and right wingtips 260, 260. In other words, the stabilizer wings 200 may be tapered. If desired, the stabilizer 200 may have a constant chord length for at least a portion of the span 255.

Consider the case of introducing the coordinate system shown in fig. 2, where the Y-axis is the lateral or transverse axis and the X-axis is the length or longitudinal axis. It is also contemplated that the Y-axis is scaled so that the left and right wingtips 260 and 260 coincide with the coordinates Y-1 and Y-1, respectively, and the origin is located on the centerline 230. It is also contemplated that the X-axis should be scaled so that the distance from the origin to X-1 is equal to the span of the wing half.

If desired, a separate coordinate system may be introduced for each wing half in the case where the wing spans of the left and right wing halves of the stabilizer 200 are not equal. However, as shown in FIG. 2, the stabilizer wing 200 has left and right wing halves 257, 258 that are equal in span and a quarter chord line 240 that is symmetrical with respect to the centerline 230 of the fuselage.

In this case, the quarter chord line 240 may be described as the function X ═ f (y), so the curvature of the quarter chord line 240 is the second derivative d²/dY²(f (Y)). When the X-axis is positive toward the leading edge 210, the backward curvature (i.e., the opening of the curve is oriented toward the tail portion of the rotorcraft (e.g., tail portion 109 of rotorcraft 100 of fig. 1)) corresponds to a negative value (i.e., d)²/dY²(f (Y)) < 0, and the forward curvature (i.e., the opening of the bend is oriented toward the nose portion of the rotorcraft (e.g., nose portion 108 of rotorcraft 100 of FIG. 1)) corresponds to a positive value (i.e., d)²/dY²(f (Y)) 0) and the curvature of the straight quarter chord line is zero (i.e. d)²/dY²(f(Y))＝0)。

For example, the absolute value of the non-zero curvature of the quarter chord line 240 may be greater than 0.2 (i.e., | d)²/dY²(f (Y)) > 0.2. As another example, the absolute value of the non-zero curvature of the quarter chord line 240 may be less than 4 (i.e., | d)²/dY²(f(Y))|＜4)。

If desired, the absolute value of the non-zero curvature of the quarter chord line 240 may have an upper limit based on the distance from the centerline 230. For example, the absolute value of the non-zero curvature of the quarter chord line 240 may be less than 4-2 x Y (i.e., | d |)²/dY²(f(Y))|＜4－2*|Y|)。

Illustratively, the quarter chord line 240 may have a non-zero curvature (i.e., d) across the span 255 ²/dY²(f) (Y) ≠ 0). If desired, the quarter chord line 240 may be defined when Y ═ 0.1; 0.9]And Y [ -0.1; -0.9]Have a non-zero curvature (i.e., d) within or over a portion of these ranges²/dY²(f(Y))≠0)。For example, the quarter chord line 240 may be at least when Y ═ 0.1; 0.7]And/or Y [ -0.1; -0.7]Has a non-zero curvature within the range of (a).

In other words, considering only the right wing half (i.e., Y ═ 0.1; 0.7), the two points 272, 278 may be located at distances 282, 288, respectively, on the quarter-chord line 240 from the centerline 230 of the fuselage, such that the magnitudes of the distances 282, 288 are 10% and 70% of the half- spans 257, 258, respectively, and thus the non-zero curvature of the quarter-chord line 240 is non-zero at least between the two points 272, 278.

If desired, the two additional points 274, 276 may be at distances 284, 286 from the fuselage centerline 230 on the quarter-chord line 240, respectively, such that the distances 284, 286 are in the magnitude of 30% and 50% of the half- span 257, 258, and the distance 270 of the straight line 290 between the two points 272, 278 and the quarter-chord line 240 between the two additional points 274, 276 may be greater than 2% of the half- span 257, 258.

In some embodiments, the non-zero curvature of the quarter chord line 240 progresses from a first value at a first distance from the centerline 230 of the fuselage to a second value at a second distance from the centerline 230 of the fuselage such that the non-zero curvature of the quarter chord line 240 is non-constant between the first and second distances from the centerline 230 of the fuselage.

For example, the curvature of the quarter chord line 240 may be 2.5 at the centerline 230 (i.e., at Y-0, | d²/dY²(f (Y) | 2.5) and decreases to a curvature of 0.5 toward the left and right wingtips 260 and 260 (i.e., at Y ═ 1, | d²/dY²(f (y)) | 0.5. As another example, the curvature of the quarter chord line 240 may be 3.0 at the centerline 230 (i.e., id where Y is 0)²/dY²(f (Y) | 3.0), and decreases to a curvature of 0.25 toward the left and right wingtips 260 and 260 (i.e., at Y ═ 1, | d²/dY²(f(Y))|＝0.25)。

In some embodiments, the quarter chord line 240 may have a break (kink). For example, the turn of the quarter chord line 240 may be at the centerline 230 (i.e., at Y-0). As another example, the quarter chord line 240 may have a turn toward the left and right wingtips 260, 260 (e.g., at Y ═ 0.95).

A family of embodiments can be conceived by the magnitude of the mean curvature within a particular Y range, if desired. The average curvature over a range of Y values (i.e. in the interval [ Y1; Y2 ]) is thus defined as the difference between the first derivative at the beginning of the range (i.e. d/dY (f (Y); Y ═ Y1)) and the first derivative at the end of the range (d/dY (f (Y); Y ═ Y2)) divided by the absolute value of (Y2-Y1). Therefore, the average curvature is | (d/dY (f (Y); Y ═ Y1) -d/dY (f (Y); Y ═ Y2))/(Y2-Y1) |.

Examples of such embodiment families are shown in tables 1, 2 and 3. Table 1 shows the family of embodiments where Y1 is 0.1, table 2 shows the family of embodiments where Y1 is 0.2, and table 3 shows the family of embodiments where Y1 is 0.3. Additional families of embodiments are possible for different values of Y1 and/or Y2.

TABLE 1

TABLE 2

TABLE 3

The variability of coordinate Y1 may be important for a stabilizer 200 protruding from a fuselage, vertical tail, or tail boom (e.g., fuselage 104, vertical tail 131, or tail boom 121 of rotorcraft 100 of fig. 1). For example, a stabilizer wing protruding from the tail boom may have no curvature value at the Y value located inside the tail boom. Thus, an embodiment may be selected from a family having a value of Y1 located outside the tail boom.

As described above, a rotor with N rotor blades generates N helical rotor tip vortices, so that N rotor blade tip vortices pass the stabilizer 200 per rotor revolution. Thus, unsteady aerodynamic loads at the stabilizer 200 are generated periodically at the fundamental frequency N per rotor revolution.

Due to the curved shape of the quarter chord line, the shape of the periodic unsteady aerodynamic loads of the stabilizer 200 at the attachment at the vertical tail or tail boom may have smoother shape peaks and reduced amplitudes compared to a stabilizer having a straight quarter chord line.

Fig. 3A, 3B, and 3C are views of an exemplary time series of the interaction between the rotor tip vortex centerline and a quarter chord line with non-zero curvature. As shown in FIGS. 3A, 3B, and 3C, the stabilizing wing 300 may have a cambered leading edge 310, a cambered trailing edge 320, and a quarter chord line 340. A centerline 330 of the fuselage is schematically illustrated, with respect to which the stabilizer 300 is symmetrical. Thus, quarter chord line 340 is symmetrical with respect to centerline 330.

Considering the rotorcraft supported on its landing gear, the rotor may rotate counterclockwise when viewed from above, or the rotor may rotate clockwise when viewed from above.

FIG. 3A shows rotor tip vortex centerline 360 reaching quarter chord line 340. In this initial condition (i.e., T ═ T1), rotor tip vortex centerline 360 is tangent to quarter chord line 340 and the interaction between rotor tip vortex centerline 360 and quarter chord line 340 begins.

Figure 3B shows rotor tip vortex centerline 370 passing through quarter chord line 340 shortly after interaction begins (i.e., T-T2, where T2 > T1).

Figure 3C shows rotor tip vortex centerline 380 at the end of the interaction with quarter chord line 340 and when rotor tip vortex centerline 380 reaches the intersection of quarter chord line 340 and centerline 330 (i.e., T-T3, where T3 > T2).

The rotor tip vortex centerlines 360, 370, 380 travel a longer time (i.e., T1 to T3) and a longer distance across the curved quarter chord line 340 of the stabilizer 300 than the similar straight quarter chord line of a similar stabilizer.

Thus, rotor tip vortex centerline 360, 370, 380 interacts with curved quarter chord line 340 for a longer duration than it does with a straight quarter chord line. The longer duration of the interaction between the rotor tip vortex centerlines 360, 370, 380 and the quarter chord line 340 means that the more impact limits are placed on the relatively local area of the quarter chord line 340 at a particular time, and therefore the less aerodynamic load input on the stabilizing wing 300.

In contrast, if the duration of the interaction is short, or in the extreme case the rotor tip vortex centerline 360 and quarter chord line only nearly coincide at a certain moment (e.g., more likely to occur with a straight quarter chord line than with a properly curved quarter chord line), then a short and strong aerodynamic load peak will result from the aerodynamic load variation occurring nearly instantaneously at the same moment along the entire span of the stabilizer wing.

By creating a large difference between the curvature and orientation of quarter chord line 340 and the curvature and orientation of rotor tip vortex centerline 360, 370, 380, the proposed embodiment of stabilizer 300 with curved quarter chord line 340 allows for a longer interaction time between rotor tip vortex centerline 360, 370, 380 and stabilizer 300 (e.g., T1-T3).

Fig. 3A, 3B, and 3C illustrate the interaction between forward curving rotor tip vortex centerlines 360, 370, 380 and aft curving quarter chord line 340 at times T1, T2, and T3, respectively. As shown in fig. 3A, 3B, and 3C, a rotor tip vortex centerline with forward curvature is generally more relevant to a T-tail (i.e., a stabilizer attached to the top of a vertical tail) in cruise flight because the forward curved rotor tip vortex centerline is generated at the trailing edge of the rotor disk and the T-tail stabilizer is geometrically close to the rotor trailing edge.

Thus, a T-tail stabilizer with a quarter chord line that is curved aft may be less affected by the centerline of the rotor tip vortex generated at the trailing edge of the rotor disk than a stabilizer with a straight quarter chord line.

However, the rotor also generates a rotor tip vortex centerline with aft curvature. For example, a rotor may generate a rotor tip vortex centerline with a rearward curvature at the leading edge of the rotor disk.

A stabilizer wing attached such that it interacts with the vortices generated at the leading edge of the rotor disk under a variety of flight conditions may interact with these backward curved rotor tip vortex centerlines for a longer period of time with a forward curvature of the quarter chord line of the stabilizer wing. Thus, a stabilizer wing with forward curvature may experience lower aerodynamic load peaks when exposed to a rotor tip vortex centerline with aft curvature.

Consider the case where the rotorcraft is a helicopter having one main rotor. Consider further that the main rotor center of the main rotor coincides with the centerline of the fuselage (e.g., main rotor center 103 coincides with centerline 130 of fuselage 104 of fig. 1C), and that the rotor tip vortex centerline emanated by the main rotor is asymmetric with respect to the centerline of the fuselage due to the asymmetry of the main rotor about the mid-plane of the helicopter.

In this case, the curved shape of the quarter chord line 340, and thus the planar shape of the stabilizer wing 300, may be tailored to minimize aerodynamic loads. For example, based on the operation of the rotorcraft, the planform of the stabilizer 300 may be designed such that the total aerodynamic load (e.g., obtained by integration across the span of the stabilizer 300) at a predetermined location (e.g., the attachment location of the stabilizer 300) does not exceed a predetermined value.

As another example, based on the operation of the rotorcraft, the planform of the stabilizer 300 may be designed such that the time series of local interactions between the rotor tip vortex centerlines 360, 370, 380 and the quarter chord line 340 at each location of the stabilizer 300 are suitably staggered to achieve a near constant total aerodynamic load over time at the relevant location of the stabilizer 300.

As shown in fig. 2, 3A, 3B, and 3C, the stabilizing wings may be tapered. In other words, the chord length 250 may be longer at the centerline 230 than at the tip 260, or the ratio of the chord length 250 at the tip 260 divided by the chord length 250 at the centerline 230 may be less than 1. For example, the ratio of the chord length 250 at the tip 260 divided by the chord length 250 at the centerline 230 may be in the interval between 0.2 and 0.8.

However, even though the stabilizing wings 200, 300 shown in fig. 2, 3A, 3B and 3C are tapered, the curved quarter chord lines 240, 340 may be combined with any kind of spanwise chord length evolution.

For example, fig. 4 is a view of a series of rotor tip vortex centerlines 460, 470, 480 to an exemplary stabilizer 400 having a constant chord length 450 according to some embodiments. In other words, stabilizer 400 may have a constant chord length 450 between left tip 457 and right tip 457 of stabilizer 400.

As shown in FIG. 4, the stabilizer 400 may have a cambered leading edge 410, a cambered trailing edge 420, and a quarter chord line 440 with a non-zero curvature. A centerline 430 of the fuselage is schematically illustrated, with respect to which the stabilizer 400 is symmetrical. Thus, quarter chord line 440 is symmetrical with respect to centerline 430.

For example, the chord length 450 may be constant at any span station position Y (i.e., at any location between Y-1 and Y-1). In this example, since the stabilizer 400 has a cambered leading edge 410 and a constant chord length 450 at any span station Y, the leading edge 410, the trailing edge 420, and the quarter chord line 440 have the same cambered shape and the same non-zero curvature.

As shown in fig. 2, 3A, 3B, 3C, and 4, the stabilizer wings 200, 300, 400 may be symmetrical with respect to the centerline 230, 330, 430. However, the stabilizing wings may be asymmetrical with respect to the centerline, if desired.

For example, fig. 5 is a view of a series of rotor tip vortex centerlines 560, 570, 580 to an exemplary stabilizer wing 500 having an asymmetric sweep between left wing half 504 and right wing half 502 according to some embodiments.

In fact, as shown in FIG. 5, left wing half 504 of stabilizer 500 may have a first sweep and right wing half 502 may have a second sweep different from the first sweep. For example, left wing half 504 may have a smaller sweep than right wing half 502. The left wing half 504 may have a greater sweep than the right wing half 502 if desired.

As shown in FIG. 5, the stabilizer wing 500 may have a cambered leading edge 510, a cambered trailing edge 520, and a quarter chord line 540 with a non-zero curvature.

For example, the stabilizer 500 may have a constant chord length 550. The chord length 550 may be constant at any span station Y (i.e., anywhere between Y-1 and Y-1), if desired. In this example, since the stabilizer 500 has a cambered leading edge 510 and a constant chord length 550 at any span station Y, the leading edge 510, the trailing edge 520, and the quarter chord line 540 have the same cambered shape and the same non-zero curvature.

However, since stabilizer wing 500 is asymmetric with respect to centerline 530, quarter-chord line 540 of left wing half 504 has a first curvature and quarter-chord line 540 of right wing half 502 has a second curvature, wherein the second curvature is different from the first curvature.

For example, right wing half 502 may have a greater curvature of quarter chord line 540 than left wing half 504. Right wing half 502 may have a smaller curvature of quarter chord line 540 than left wing half 504, if desired.

Left wing half 504 may have a first length and right wing half 502 may have a second length different than the first length, if desired. In other words, one of the left wing half 504 and the right wing half 502 may constitute more than 50% of the span 555. For example, left wing half 504 may be shorter than right wing half 502. As another example, right wing half 502 may be shorter than left wing half 504.

The lengths of right wing half 502 and left wing half 504 may be selected to equalize the contribution of aerodynamic loads on right wing half 502 and left wing half 504 to the roll moment around the attachment of stabilizer 500 (e.g., at the vertical tail, tail boom, or fuselage of a rotary wing aircraft). The amount of reduction of the static and/or dynamic component of the aerodynamic load at the attachment of the stabilizer wing may be determined based on the structural requirements.

The stabilizer 500 may include a winglet (e.g., winglet 138 of FIG. 1), if desired. A winglet may increase the aerodynamic lift of the stabilizer 500 without increasing the span 555.

Because the span 555, and in particular the length of the right and left wing halves 502, 504, respectively, defines an effective lever arm for the roll moment at the attachment of the stabilizer 500, and because the tip winglet is oriented perpendicular to the rotor tip vortex centerline 560, 570, 580 (which limits the interaction between the rotor tip vortex centerline 560, 570, 580 and the tip winglet), the use of a tip winglet may reduce the oscillating roll moment at the attachment of the stabilizer 500 for a given aerodynamic lift target.

The curvature of quarter chord line 540 and/or the asymmetry of the sweep between right wing half 502 and left wing half 504 may be utilized to further reduce unsteady aerodynamic loads on stabilizer wing 500.

For example, selecting a combination of a predetermined curvature of quarter chord line 540 and a predetermined sweep of left and right halves 504, 502 may allow for independent adjustment of the time series of interactions between rotor tip vortex centerlines 560, 570, 580 along the span station Y of left and right halves 504, 502 of stabilizer wing 500.

To explain this effect, it can be considered that the previously symmetrical stabilizer wings are rotated by an angle ψ relative to the centre line of the fuselage, which results in an asymmetry of the sweep.

FIG. 6 is a view of an exemplary stabilizer wing 600 having a quarter chord line 640 with non-zero curvature and a wing centerline 635 perpendicular to a series of rotor tip vortex centerlines 660, 670, 680, according to some embodiments.

For example, the stabilizer 600 may have a tapered chord length 650. The chord length 650 may decrease from the wing centerline 635 toward the wing tip if desired (i.e., the chord length 650 decreases with increasing distance from the wing centerline 635). In this example, leading edge 610, trailing edge 620, and quarter chord line 640 may all have different camber shapes and different non-zero curvatures.

As shown in FIG. 6, stability wing 600 may be symmetric with respect to wing centerline 635. Because wing centerline 635 is rotated at angle ψ relative to the centerline 630 of the fuselage, stability wing 600 may be asymmetric relative to the centerline 630 of the fuselage.

Because wing centerline 635 is perpendicular to the series of rotor tip vortex centerlines 660, 670, 680, each of rotor tip vortex centerlines 660, 670, 680 reach both halves of stabilizer 600 simultaneously. Thus, the bending moments around the wing centerline 635 of the stabilizer wing 600 occur in phase in time.

As shown in FIG. 6, stabilizer 600 may be rotated relative to fuselage centerline 630 such that rotor tip vortex centerlines 660, 670, 680 are perpendicular to wing centerline 635. The actual angle of rotation ψ between wing centerline 635 and fuselage centerline 630 may depend on the actual configuration of the rotorcraft.

For example, for a rotorcraft having one main rotor with its center located at the centerline 630 of the fuselage forward of the stabilizer 600 and rotating counterclockwise when viewed from above, an angle ψ between 0 ° and 30 ° may be selected. As another example, an angle ψ of between 0 ° and-30 ° may be selected for a rotary wing aircraft having one main rotor with its center located at the centerline 630 of the fuselage forward of stabilizer 600 and rotating clockwise when viewed from above.

The elasticity of the stabilizer wings 200, 300, 400, 500 or 600 shown in fig. 2-6 may affect aerodynamic loads, as the curvature of the quarter chord lines 240, 340, 440, 540, 640 may cause a coupling reaction in bending and torsion. Thus, the vertical bending of the stabilizer wings 200, 300, 400, 500, 600 caused by the aerodynamic load may be accompanied by elastic torsion of the stabilizer wings about the Y-axis. Providing a stabilizer wing that allows for elastic torsion at its attachment and/or a torsion structure around the Y-axis may reduce aerodynamic loads, as torsion may change the angle of attack and thus the aerodynamic load on the stabilizer wing.

If desired, the stabilizer wings 200, 300, 400, 500 or 600 of fig. 2 to 6 may have a twist, which is the change in angle between the chord line and the horizontal XY plane for different span stations Y. In other words, the stabilizer 200, 300, 400, 500 or 600 may have a variable stagger angle over at least a portion of the span.

The stabilizer wings 200, 300, 400, 500 or 600 of fig. 2 to 6 may also have a twist, if desired, in addition to the quarter chord line 240, 340, 440, 540 or 640 having a non-zero curvature. The additional twist of the stabilizer 200, 300, 400, 500 or 600 may compensate for asymmetric inflow conditions on the left and right halves of the respective stabilizer 200, 300, 400, 500 or 600.

Consider the case where the rotor wake is not symmetric about the centerline of the fuselage. For example, as shown in FIG. 5, the rotor wake that generates rotor tip vortex centerlines 560, 570, 580 is asymmetric with respect to the plane defined by fuselage centerline 530 and the Z-axis.

In this case, rotor tip vortex centerlines 560, 570, 580 may create asymmetric average aerodynamic loads between left wing half 504 and right wing half 502 in the case of non-twisted horizontal stabilizer wings.

For example, the stabilizer wing 500 may equalize the time averages of the lift and wing root moments of the left and right wing halves 504, 502 at the attachment of the stabilizer wing 500 for the most relevant flight conditions (e.g., horizontal forward flight) by the curved quarter chord line 540 and the variable angle of installation across the span 555, thereby reducing the combined static roll moment exerted by the left and right wing halves 504, 502 at the attachment of the stabilizer wing 500.

The stabilizer wings 200, 300, 400, 500, 600 of fig. 2-6 may include a spanwise distribution of stagger angles, if desired. The spanwise distribution of the stagger angles may reduce the risk of flow separation at the root of the stabilizer. The risk of flow separation increases with increasing underpressure on the stabilizer blade.

For example, the attack angle may be reduced in a direction from the tip of the stabilizer 200, 300, 400, 500, or 600 of fig. 2 to 6 to the root of the stabilizer, thereby reducing the negative pressure value generated at the lower surface of the root of the stabilizer.

The portion of the stabilizer surface near the root of the stabilizer is particularly susceptible to flow separation if the stabilizer is mounted to the vertical tail of a rotary wing aircraft. In fact, in addition to the low pressure caused by the stabilizer itself, the low pressure field of the vertical tail also extends to the surface portion of the stabilizer near the root of the stabilizer.

The vertical tail of a rotorcraft may provide counter-torque against the torque of the rotor. For example, a vertical tail may generate a rightward (i.e., in the positive Y-direction) force in a rotorcraft having a rotor that rotates counterclockwise when viewed from above. In fact, the force generated by the vertical tail is achieved by the negative pressure on the right surface of the vertical tail.

In this example, the zero degree stagger angle between the chord of the stabilizer and the XY-plane may be to the right of the vertical tail (i.e., at a wing span station with a positive Y). If desired, the stagger angle between the chord of the stabilizer and the XY-plane may increase continuously from a wingspan station at a stagger angle of zero degrees towards the left wing tip (i.e., in the negative Y-direction) and decrease continuously from a wingspan station at a stagger angle of zero degrees towards the right wing tip (i.e., in the positive Y-direction).

Alternatively, for a rotor that rotates clockwise when viewed from above, the left vertical tail surface will be subjected to negative pressure. Thus, the zero degree stagger angle between the chord of the stabilizer and the XY-plane may be to the left of the vertical tail (i.e., at a wing span station with a negative Y).

If desired, the stagger angle between the chord of the stabilizer and the XY-plane may increase continuously from a wingspan station at a stagger angle of zero degrees towards the right wing tip (i.e., in the positive Y-direction) and decrease continuously from the wingspan station at a stagger angle of zero degrees towards the left wing tip (i.e., in the negative Y-direction).

The distribution of the stagger angle across the span of the stabilizer can be moved globally along the ordinate (i.e. along the Y-axis to increase or decrease the stagger angle evenly across the entire span) without changing the variation along the abscissa (i.e. the twist along the X-axis), but only the global stagger angle and hence the lift of the entire stabilizer.

For example, the variation of the stagger angle may be greater towards the root of the stabilizer than towards the tip of the stabilizer. In other words, for a given mounting angle θ and a transverse axis Y from Y ═ 1 at the left wing tip to Y ═ 1 at the right wing tip of the stabilizing wing, the change in mounting angle d θ/dY is greater than the average twist in the spread area Y [ -0.5 … 0.5] (i.e., θ (Y ═ 1) to θ (Y ═ 1)/2)

For example, the mounting angle distribution across the span of the stabilizer may have predetermined upper and lower limits. The actual mounting angle distribution across the span of the stabilizer may be selected between predetermined upper and lower limits, if desired. For example, the option of actual mounting angle distribution across the stabilizing wings may be selected based on the primary flight pattern.

It should be noted that the above-mentioned embodiments are only described for illustrating possible implementations of the present invention, and are not intended to limit the present invention thereto. On the contrary, many modifications and variations of the described embodiments are possible and should therefore also be considered as part of the present invention.

For example, the asymmetric stabilizer 500 of FIG. 5 is shown with a constant chord length 550. However, if desired, the asymmetric stabilizer 550 may be tapered. Similarly, stabilizer 600 of FIG. 6 is shown as being tapered. However, the stabilizer 600 may have a constant chord length, if desired.

As another example, the stabilizer wings 200, 300, 400, 500, 600 of fig. 2-6 are shown as having a predetermined sweep. The stabilizer wings 200, 300, 400, 500, 600 may have different sweepbacks, if desired. For example, the sweep of the stabilizer wings 200, 300, 400, 500, 600 may be greater or less than the sweep shown in fig. 2-6.

Further, the lower limit of the absolute value of the non-zero curvature of the quarter chord line 240 of the stabilizer blade 200 in the drawing is designated as 0.2 (i.e., | d)²/dY²(f (Y)) > 0.2, and the upper limit of the absolute value of the non-zero curvature of the quarter chord line 240 is specified as 4 (i.e., | d)²/dY²(f(Y))|＜4)。

However, the lower limit of the absolute value of the non-zero curvature of the quarter chord line 240 may be selected to be greater than or less than 0.2. Similarly, the upper limit of the absolute value of the non-zero curvature of the quarter chord line 240 may be selected to be greater than or less than 4. The lower and upper limits of the absolute value of the non-zero curvature of the quarter chord line 240 may be selected based on the shape of the rotor tip vortex centerline, if desired.

Similarly, the absolute value of the non-zero curvature of the quarter chord line 240 may have an upper limit based on the distance from the centerline 230, and the factor by which the upper limit is reduced may be a factor a, which may be calculated from the interval [ 0; b is]Where B is the upper limit of the absolute value of the non-zero curvature of the quarter chord line 240. For example, the absolute value of the non-zero curvature of the quarter chord line 240 may be less than B- (B-C) | Y |, where B e [ 1; 4]And C ∈ [ 0; 1](i.e. | d)²/dY²(f(Y))|＜B－(B－C)*|Y|)。

List of reference numerals

100 rotor craft

102 main rotor

102a, 102b, 102c, 102d, 102e rotor blade

103 main rotor center

103a rotor shaft axis

104 fuselage

104a nacelle

107 fuselage nose

108 head part

109 tail part

121 tail beam

122 tail rotor

130 center line

131 vertical tail

132 rudder

133 tail wing, stabilizer wing

136 leading edge

137 trailing edge

138 wingtip winglet

160 wing tip

200 stabilizing wing

210 leading edge

220 trailing edge

230 center line

240 quarter chord line

Length of 250 chord

255 span

257 half wingspan (left half wing)

258 semi-wingspan (Right semi-wing)

260 wingtip

270 distance

272. 274, 276, 278 quarter chord line

282 is a distance of 10% of the span

284 is a distance of 30% of the span

286 is a distance of 50% of the span

288 is a distance of 70% of the span

290 straight line

300 stabilizer

310 leading edge

320 trailing edge

330 center line

340 quarter chord line

360. 370, 380 rotor wing tip vortex centerline

400 stabilizer

410 leading edge

420 trailing edge

430 center line

440 quarter chord line

450 chord length

457 wing tip

460. 470, 480 rotor wing tip vortex center line

500 stabilizer

502 right wing half

504 left half wing

510 leading edge

520 trailing edge

530 center line

540 quarter chord line

550 chord length

555 wingspan

560. 570, 580 rotor wing tip vortex center line

600 stabilizer

610 leading edge

620 trailing edge

630 fuselage centerline

635 wing centerline;

640 quarter chord line

Length of 650 chord

660. 670, 680 rotor wing tip vortex centerline;

x longitudinal axis and length axis

Y horizontal axis, side axis and width axis

Z vertical axis and height axis