RU2743306C1 - Rotary-winged aircraft with stabilizer wing - Google Patents

Abstract

FIELD: aviation.SUBSTANCE: invention relates in particular to the structures of rotary-wing aircraft. The rotorcraft with a fuselage has a centerline 230, at least one main rotor that generates vortices during operation, and a stabilizer wing 200. The stabilizer wing 200 is shaped in plan that reduces unstable aerodynamic loads caused by the wake stream of at least one main rotor. In particular, the stabilizer wing 200 may be provided with a left wing end 260, a right wing end 260, a quarter-chord line 240 with non-zero curvature so that the interaction between the vortices generated by at least one rotor and the quarter-chord line 240 is timed, a leading edge 210 that is arcuate, and a trailing edge 220 that is arcuate.EFFECT: improvement of controllability and flight stability is provided.16 cl, 10 dwg, 3 tbl

Description

The present embodiments relate to a rotary-wing aircraft, and more particularly to a rotary-wing aircraft with at least one main rotor and a stabilizer wing, whereby the stabilizer wing is shaped in plan that reduces unstable aerodynamic loads caused by it with a wake of at least one main rotor.

The rotorcraft is usually equipped with one or more wings that help improve the stability and controllability of the rotorcraft during flight. Therefore, such wings are sometimes also called horizontal stabilizers, horizontal stabilizers, or stabilizer wings.

Traditionally, helicopters have one or more stabilizer wings located at the aft end of the tail assembly. On some models, stabilizer wings extend from the rear of the tail boom or from the keel. In other models, a single stabilizer wing is placed at the top of the keel. The latter configuration, in which a single stabilizer wing is placed at the top of the keel, is sometimes also referred to as the T-tail.

Due to the location of the stabilizer wings in the tail region of the rotorcraft, they are typically exposed to airflow patterns generated by at least one rotor. These airflow patterns, sometimes also referred to as wake or rotor wake, interact with the stabilizer wings and determine the aerodynamic load on the stabilizer wings.

The force of interaction between the wake of the main rotor and the stabilizer wing depends on the position of the stabilizer wing relative to the propeller and on the state of flight (for example, during takeoff and landing or during level flight). However, the formation of aerodynamic loads on the stabilizer wing, which are significant for determining the structural dimension of the stabilizer wing, can usually not be prevented.

The stabilizer wing generates some amount of static aerodynamic load intentionally (for example, to improve the stability and controllability of the rotorcraft during flight). However, undesirable unstable aerodynamic loads are also obtained as a result of the interaction of the rotor wake jet with the stabilizer wing.

These unstable aerodynamic loads can sometimes reach significant amplitudes, which can lead to fatigue failure of the stabilizer wing. Consequently, special care during the design of a stabilizer wing is often required to ensure a structural design that greatly reduces the risk of fatigue damage caused by unwanted unstable aerodynamic loads.

The problem is often compounded by the fact that one of the dynamic natural frequencies of the stabilizer wing and tail boom oscillations is close to the rotor wake. Such resonance conditions between the rotor wake and the stabilizer wing can be prevented by increasing either the structural rigidity or the stabilizer wing mass.

However, increasing the structural rigidity or mass of the stabilizer wing adds weight in the sensitive area of the rotorcraft as it moves the center of gravity further back. Especially when the stabilizer is placed on the keel, structural and dynamic problems are difficult to solve. As a result, stabilizer wings are rarely placed on the keel, which would otherwise be very beneficial to handling and performance.

Instead, most conventional solutions place the stabilizer wing in a position that is vertically below the plane of the main rotor (for example, on the tail boom). As a result, the interaction between the rotor wake and the stabilizer wing mainly occurs at lower flight speeds. Since air loads increase with flight speed, the interaction between the rotor wake and the stabilizer wing at lower flight speeds produces relatively moderate unstable aerodynamic loads, while the static air load on the stabilizer wing is relatively low. Thus, since the stabilizer wing structure is subject to the imposition of static air load and unstable aerodynamic loads, the stabilizer wing structure does not reach its fatigue limits at lower flight speeds.

At high flight speeds, the main rotor wake jet changes its trajectory and passes over a low stabilizer wing (i.e., a stabilizer wing, which is located on the tail boom, for example). Thus, even if the static air load on the stabilizer wing is greater at high flight speeds compared to low flight speeds, the relatively small amplitude of unstable aerodynamic loads of the low stabilizer wing has the effect of keeping the total air load good in the fatigue range. ...

However, a low stabilizer wing exhibits reduced handling and performance when compared to a high stabilizer wing (eg a stabilizer wing at the top of the keel). For example, during hovering and / or at low airspeed, the low stabilizer wing lies in the rotor jet and generates a force directed towards the ground, which is sometimes also called downforce. This downward force causes the rotorcraft to lift up the nose, sometimes also called nose-up.

Pitch-up can drastically reduce the pilot's line of sight to the ground. The cabling additionally creates a relatively high mechanical stress on the main rotor column. In addition, the downforce can significantly reduce the payload of the rotorcraft. By way of example, the downforce can reduce the payload of a helicopter in the three to four ton class by an amount that is equivalent to one passenger.

In addition, a low-slung fender should usually be mounted on the tail boom. This leads to a shorter lever arm and thus a reduced stabilizing effect compared to a high stabilizer wing.

Some helicopter manufacturers are trying to overcome the dilemma between handling and performance, on the one hand, and flight stability, on the other, by using a high-mounted stabilizer wing. However, as mentioned above, these high-positioned stabilizer wings face the problem of high unstable aerodynamic loads caused by the rotor wake interaction during cruise flight.

Because unstable aerodynamic loads pose serious design problems, the relatively few helicopters designed to date feature a high stabilizer wing. Often, helicopters with high wing stabilizers include additional means to reduce erratic aerodynamic loads. Among such means are the tapering of the stabilizer wings (ie, the stabilizer wing has a chord length that is longer at the root than the ends) or the inclusion of structural reinforcements. These reinforcements typically consist of a thickening of the stabilizer wing at its connection to the tail boom or keel, or the installation of struts that externally support the stabilizer wing.

For example, the improved AV-02 Hughes YAH-64 fire support helicopter has a T-shaped tail (i.e., a stabilizer wing on the keel) with a sweep (i.e., the stabilizer wing is angled backward from the bark to ends) and narrowing. The Agusta-Bell BA609 features a T-shaped tail unit with a taper, resulting in a stabilizer wing with straight leading and trailing edges. The RAH-66 Comanche has a T-shaped tail with a straight stabilizer wing (i.e., a stabilizer wing at the top of the keel with straight leading and trailing edges, but no sweep). Kamov Ka-60 has a T-shaped tail unit with a straight stabilizer wing and supports to the keel.

Both the narrowing of the stabilizer wing and the reinforcement are aerodynamically disadvantageous in that they increase drag, decrease the efficiency of the stabilizer wing, and reduce the range of flight states in which the stabilizer wing has the ability to stabilize the rotorcraft.

EP 2 899 118 A1 describes a rotorcraft with a fuselage and at least one rotor, said at least one rotor is actuated to control the associated pitch angle of said rotorcraft in operation, and said fuselage is equipped with at least one passive wing-type aerodynamic device, which is adapted to generate, independently of said associated pitch angle, a lift acting on said fuselage, said lift is oriented perpendicular to the air flow that is directed during operation of said rotorcraft against said passive wing-type aerodynamic device. The rotorcraft further comprises a keel with a rudder and a T-shaped tail wing provided on the fuselage tail boom. The tail wing can be tilt adjustable and act as an additional lifting surface.

CN 103 979 105 A describes a vertical takeoff and landing variable wing aircraft. An aircraft with a variable geometry wing with a vertical take-off and landing type comprises an aircraft body, a main wing, a left wing and a right wing, the main wing being located in the lower part of the aircraft body; the left wing and the right wing are respectively attached to the two ends of the main wing; a shaft sleeve for turning the left wing and a shaft sleeve for turning the right wing are installed in the main wing; one set of the left wing rotation shaft in the left wing rotation shaft sleeve is located in the left wing; one set of the right wing pivot shaft in the right wing pivot shaft sleeve is located on the right wing; the left wing steering mechanism and the right wing steering mechanism are also housed in the main wing; the left wing steering mechanism is used to drive the left wing steering shaft to adjust the direction of the left wing; the right wing steering mechanism is used to drive the right wing pivot shaft to adjust the direction of the right wing. Variable geometry wing aircraft with vertical take-off and landing type is adapted to realize vertical take-off and landing of model airplane and unmanned aerial vehicle, quickly realize various flight conditions, and efficient braking and backward motion in flight.

EP 2 666 719 B1 describes an aircraft that includes a fuselage at least one main rotor provided with a plurality of blades, at least a pusher rotor with a variable pitch propeller, at least one airfoil and at least one motor driving main rotor and each push rotor. The airfoil may comprise two half-wings located on both sides of the fuselage and at least one horizontal stabilizer located at one end of the aircraft and provided with at least one movable surface.

WO 1999/067130 A1 describes a horizontal stabilizer defining the first and second nodes located along the wingspan, the first node located along the wingspan defining the first pitch angle, and the second node located along the wingspan defining the second pitch angle, the angle of inclination differs from one node to another, for example, one more than the other. The horizontal stabilizer functions to favorably influence the distribution of lift over the wingspan to reduce bending moments near its mounting surface. Various embodiments of the horizontal stabilizer include the use of vertically extending ridges along the trailing edge of the horizontal stabilizer, a stepped transition to abruptly change the angle of inclination from one node to another, and distributed twisting that gradually changes the angle of inclination.

EP 0 254 605 A1 describes a directional and stabilizing device which comprises a tunnel tail rotor driven in rotation in a transverse seam formed in a fairing inclined with respect to the vertical by an angle between 0 ° and 45 °. It also includes an "inward V-shaped" empennage attached to the top of the cowl and two aerofoils extending asymmetrically about a vertical plane passing through the top of the fairing. Aerodynamic surfaces are inclined in angular ranges between 0 ° and 45 °.

US 3,464,650 A describes a propeller / wing aircraft with flaps for vertical and short take-off and landing (V / STOL). At the apex of the vertical tail, the aircraft has a horizontal stabilizer, which can be of a fully movable type or have separate movable fragments. These individual pieces will act as traditional elevators, or can be linked together to function as ailerons to control the roll in forward flight in a dual function lifting surfaces. Conventional aircraft controls are used to engage the horizontal stabilizer and the vertical tail rudder for forward flight control.

Other stabilizer wings are described by way of example in documents CN 106 516 082 A, US 2016/0031554 A1, WO 2005/005250 A2, US 5,738,301 A and US 3,902,688 A. These documents respectively describe stabilizer wings in a rotorcraft, which are at least similar to the stabilizer wings described above. EP2409917 and WO2009155584 have been cited.

In short, the interaction between rotor wake and horizontal stabilizer wings in rotary-wing aircraft depends on the flight mode (i.e. hovering, forward flight at low or high speed) and the placement of the stabilizer wing on the rotorcraft (i.e. ., low or high), which has an impact on controllability, performance and flight stability. High stabilizer wings provide better controllability, performance and flight stability compared to low stabilizer wings. However, as mentioned above, high-positioned stabilizer wings face the problem of high unstable aerodynamic loads caused by interaction with the rotor wake during cruise flight.

The aim is therefore to provide a rotorcraft with a fuselage of at least one rotor and a stabilizer wing. The stabilizer wing should reduce unstable aerodynamic loads caused by the interaction with the rotor wake of at least one main rotor while providing improved controllability, performance and flight stability compared to traditional stabilizer wings.

These goals are solved by a rotary-wing aircraft, which contains the distinctive features of paragraph 1 of the claims. More specifically, the rotorcraft may comprise a fuselage having a centerline of at least one main rotor that creates turbulence during operation and a stabilizer wing that connects to the fuselage. The stabilizer wing is provided with a left wing end, a right wing end, a chord quarter line with nonzero curvature, so that the interaction between the vortices created by at least one main rotor and the chord quarter line is distributed in time, the leading edge, which is arcuate, and the trailing edge. an edge that is arcuate.

Advantageously, in a rotorcraft with at least one main rotor, the above-described stabilizer wing with a quarter-chord line having a non-zero curvature leads to reduced unstable aerodynamic loads caused by the wake of at least one main rotor on the stabilizer wing. Such stabilizer wings can be used on any rotorcraft with at least one main rotor, including vertical takeoff and landing aircraft, multicopters, helicopters, drones, etc.

The stabilizer wing has a separate shape in plan view (i.e., a separate shape when viewed from above). The separate shape in plan of the stabilizer wing reduces unstable loads generated by interaction with the rotor wake of at least one main rotor. In contrast to the existing stabilizer wing planes, which are mainly rectangular or consist of two trapezoids, a plan view is proposed in which the leading and trailing edges differ in curvature.

Accordingly, the quarter-chord line of the stabilizer wing is curved. In other words, the line that connects all points that are in the quarter of the chord length, which is the distance from the leading edge to the trailing edge of the stabilizer wing along the chord, from the leading edge, is curved. The presence of a curved line of chord quarters predominantly changes the response of the stabilizer wing to wake oscillations of the main rotor as compared to the stabilizer wings with a straight line of chord quarters.

In fact, the rotor wake manifests a regular airflow pattern during cruise flight, and this pattern determines the temporal sequence of loading conditions on the stabilizer wing. The curved line of the quarter-chords changes the relationship between the stabilizer wing and the rotor wake flow pattern, reducing the amplitude of unstable loads. Thus, the stabilizer wing becomes less susceptible to transient loading conditions of the type that create high mechanical stress at the junction of the stabilizer wing with the rotorcraft (for example, on the tail boom or keel), where the stabilizer wing is structurally sensitive. In particular, the instantaneous oscillations of the stabilizer wing about the longitudinal axis of the rotorcraft benefit from the reduced interaction between the rotor wake and the stabilizer wing, since these instantaneous oscillations are the most critical for the junction of the stabilizer wing to the rotorcraft from a structural point of view. ...

Since the proposed new shape in terms of the stabilizer wing significantly reduces unstable aerodynamic loads at the attachment point of the stabilizer wing, the stabilizer wing mounting reinforcement can be designed more economical in weight, and more economical in space (i.e., with less potentially negative influence on the aerodynamic shape of the interconnection area between the stabilizer wing and the tail boom or keel) compared to the fastening reinforcement of traditional stabilizer wings.

Significant weight savings can be achieved due to the reduced unstable aerodynamic loads. As an example, a stabilizer wing may require less structural reinforcement. As another example, the strength and weight of structures to which the stabilizer wing is attached (eg, tail boom or keel), and to which the stabilizer wing transmits unstable aerodynamic loads, can also be structurally reduced.

The keel of a rotary-wing aircraft is often a structurally weak component. Consequently, a stabilizer wing with a curved quarter-chord line, which greatly reduces unstable loads caused by rotor wake, is particularly useful when mounted on the keel. In fact, the aerodynamic interaction between the rotor wake and the stabilizer wing occurs at high speed, and unstable loads on the stabilizer wing usually reach their highest amplitudes when the stabilizer wing is mounted on the keel.

In addition, the exposure time of the stabilizer wing to these high unstable loads caused by the rotor wake is a large proportion of the rotorcraft's operating time. The large distance between the stabilizer wing and the fuselage turns the keel attachment into a junction that reacts with large deflections to the stabilizer wing loads and exhibits a low natural vibration frequency, which is difficult to correct if necessary from the point of view of structural dynamics.

A stabilizer wing with a curved chord quarter can achieve a critical amount of reduction in unstable load that is necessary to realize the otherwise difficult design of a keel-mounted stabilizer. Several important controllability and performance advantages can be realized if a curved quarter-chord stabilizer is mounted on the keel, as detailed in the disadvantages of the prior art.

In one aspect, the fuselage extends longitudinally between the nose region and the tail region, and the quarter-chord line of the stabilizer wing is closer to the tail region at at least one of the left and right ends of the wing than the center line of the fuselage.

In one aspect, the fuselage extends longitudinally between the nose and the aft region, and the quarter-chord line of the stabilizer wing is closer to the nose at at least one of the left and right ends of the wing than the centerline of the fuselage.

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

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

In one aspect, the non-zero curvature of the quarter-chord line changes from a first value at a first distance from the fuselage centerline to a second value at a second distance from the fuselage centerline, such that the non-zero curvature of the quarter-chord line is non-constant between the first and second distance from the fuselage centerline.

In one aspect, the stabilizer wing has a wing half extending between the centerline and the wing tip from the left and right ends of the wing, with the first, second, third, and fourth points located on a quarter-chord line at first, second, third and fourth distances from the centerline. fuselage lines, respectively, with the first, second, third and fourth distances being 10%, 30%, 50% and 70% of the wing span, respectively, while the nonzero curvature of the chord quarter line is nonzero between the first and fourth points, and at the same time the straight line between the first and fourth points has the fifth distance from the quarter-chord line between the second and third points, which is more than 2% of the wing span.

In one aspect, the quarter-chord line is symmetrical to the fuselage centerline.

In one aspect, a stabilizer wing is mounted on at least one of a fuselage, a tail boom that attaches to the fuselage, or a vertical keel that attaches to a tail boom.

In one aspect, the stabilizer wing has a constant chord length between the left and right ends of the stabilizer wing.

In one aspect, the stabilizer wing has a chord length that decreases from the centerline of the fuselage towards the left and right ends of the stabilizer wing.

In one aspect, the stabilizer wing further comprises a left wing half having a first sweep and a right wing half having a second sweep that is different from the first sweep.

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

In one aspect, the left wing has a first length and the right wing has a second length that is different from the first length.

In one aspect, the stabilizer wing further comprises aerofoil end surfaces.

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

The embodiments are outlined by way of example in the following description with reference to the attached drawings. In these accompanying drawings, identical or identically functioning components and elements may be designated by identical reference numbers and symbols and, therefore, are described only once in the following description.

- Fig. 1A is a schematic diagram of an illustrative rotorcraft with a stabilizer wing, in accordance with some embodiments,

- Fig. 1B is a simplified side view of an illustrative rotorcraft in accordance with some embodiments,

- Fig. 1C is a simplified top view of an illustrative rotorcraft in accordance with some embodiments,

- Fig. 2 is a diagram of an exemplary stabilizer wing having a non-zero curvature chord quarter line in accordance with some embodiments,

- Fig. 3A is a diagram of an illustrative interaction between a rotor blade tip vortex line reaching a non-zero curvature chord quarter line of the left wing half of a stabilizer wing, in accordance with some embodiments.

- Fig. 3B is a diagram of an exemplary interaction between a rotor blade tip vortex line passing through a non-zero curvature chord quarter line of the left wing half of the stabilizer wing, in accordance with some embodiments.

- Fig. 3C is a diagram of an exemplary interaction between a rotor blade end vortex centerline reaching the intersection of a non-zero curvature chord line of the stabilizer wing and a fuselage centerline, in accordance with some embodiments,

- Fig. 4 is a sequence diagram of a rotor blade tip vortex center line reaching an exemplary stabilizer wing having a constant chord length and a non-zero curvature chord quarter line in accordance with some embodiments.

- Fig. 5 is a sequence diagram of a rotor blade tip vortex center line reaching an exemplary stabilizer wing having an asymmetric bend between left and right wing lobes in accordance with some embodiments, and

- Fig. 6 is a diagram of an exemplary stabilizer wing having a tapered chord length, a non-zero curvature chord quarter line, and a wing centerline that is perpendicular to a series of rotor blade end vortex centerlines in accordance with some embodiments.

FIG. 1A, 1B and 1C show a rotorcraft 100 with a fuselage 104 and a rotor 102. FIG. 1A shows a three-dimensional representation of rotorcraft 100, while FIG. 1B illustrates a simplified side view of rotorcraft 100, and FIG. 1C is a simplified top view of rotorcraft 100.

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

It should be noted, however, that the present embodiments are not limited to helicopters and can similarly be applied to other rotary-wing aircraft that are equipped with pivoting wings, whether such pivoting wings define articulated, non-pivoting, or pivotless and bearingless multi-bladed propellers. It should be further noted that the present embodiments can also be applied in cases where more than one rotor is provided.

In fact, exemplary embodiments may be incorporated into any rotorcraft with at least one main rotor and stabilizer wing, in which unstable aerodynamic loads caused by the wake of at least one main rotor on the stabilizer wing are to be reduced. Examples of such vehicles may include rotary-wing aircraft such as vertical takeoff and landing aircraft, multicopters, helicopters, drones, etc.

Illustratively, the rotorcraft 100 may have a fuselage 104 that forms the airframe of the rotorcraft 100. The fuselage 104 may be coupled to a suitable landing gear and aft fuselage. By way of example, the landing gear may be a skid-type 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 coupled to the tail boom 121. The fuselage 104, for example, forms a cockpit 104a defining the nose 107 of the fuselage in the nose region 108 of the rotorcraft 100.

FIG. 1B and 1C show rotorcraft 100 in a Cartesian coordinate system with X, Y and Z axes. The X axis extends in the longitudinal direction of rotorcraft 100 through nose 108 and tail 109 and is sometimes also called longitudinal or length axis.

The Y-axis extends laterally to the rotorcraft 100 through the left and right wing ends 160 for the stabilizer wing 133 and is sometimes also referred to as the transverse axis, lateral axis, or breadth axis. The Z axis extends in the vertical direction of the rotorcraft 100 and is sometimes also referred to as the vertical or altitude axis.

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

By way of example, rotorcraft 100 may include at least one reluctance device configured to provide reluctance during operation, i.e., resist the torque generated by rotation of at least one multi-blade propeller 102. in order to balance the rotorcraft 100 in terms of yaw. If desired, the reaction torque device can be covered with a cover. At least one torque device is illustratively provided in the tail region 109 of the rotorcraft 100 and may have a tail rotor 122.

The tail region 109 of the rotorcraft 100 may have a keel 131 that attaches to the tail boom 121 if desired. In some embodiments, the keel 131 may be provided with a rudder 132. The rudder 132 can be adapted to provide improved direction control and for optimized directional balancing of the rotorcraft 100. If desired, the rudder 132 can be deflected to large angles to reduce the lateral drag of the keel 131 during lateral flight.

Illustratively, the keel 131 may be provided with a suitable horizontal stabilizer wing 133 in the form of a T-tail. In other words, as shown in FIG. 1A and 1B, the horizontal stabilizer wing 133 may be coupled to the fuselage 104 via the keel 131 and tail boom 121. As shown, the stabilizer wing 133 may be mounted on top of the keel 131.

If desired, the stabilizer wing 133 may be mounted in a different position on the rotorcraft 100. For example, the stabilizer wing 133 may be mounted directly on the fuselage 104, tail boom 121, and / or keel 131 at a position that is on the Z-axis of height below keel summits 131.

The stabilizer wing 133 may be provided with left and right wing ends 160, a leading edge 136 that is arcuate, and a trailing edge 137 that is arcuate.

As an example, the stabilizer wing 133 may have a constant chord length. In other words, the chord of the stabilizer wing 133, which is an imaginary straight line that connects the leading edge 136 and the trailing edge 137 of the stabilizer wing 133, may have a constant length for each point on the leading edge 136. If desired, the stabilizer wing 133 may have tapering chord length. In other words, the chord length of the stabilizer wing 133 may be greater at the centerline 130 of the fuselage 104 as compared to the left and right ends 160 of the wing.

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

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

If desired, the quarter-chord line of the stabilizer wing 133 may be curved forward. In other words, when a connecting line is drawn between two arbitrary points on the quarter chord line of the stabilizer wing 133, the quarter chord line segment between these two arbitrary points is closer to the tail region 109 than to the connecting line.

If desired, the stabilizer wing 133 can have vertical or nearly vertical extensions at the left and right ends 160 of the wing. Such vertical or nearly vertical wing tip extensions are sometimes also referred to as aerofoil end surfaces 138. As shown in FIG. 1A, stabilizer wing 133 has downwardly angled aerofoils 138. However, stabilizer wing 133 may have upwardly angled aerofoils 138 or aerofoil ends 138 that extend up and down as desired.

The rotor 102, and thus the plurality of rotor blades 102a, 102b, 102c, 102d, 102e, is propeller driven, i.e., controllable, to influence the associated pitch angle of the rotorcraft 100 in operation. Illustratively, the main rotor 102 defines the center 103 of the main rotor and includes a main rotor column. The main rotor column has a rotor column axis (eg, the main rotor column axis 103a in FIG. 1B) that defines the rotational axis of the main rotor 102.

During operation, the rotor 102 can form airflow patterns to provide lift and forward or backward thrust. Such airflow patterns are sometimes also referred to as rotor wake or wake. The main rotor wake for the main rotor 102 can interact with the stabilizer wing 133. In particular, the main rotor wake can impose an aerodynamic load on the stabilizer wing 133.

The aerodynamic load on the stabilizer wing 133 can be divided into a static aerodynamic load and an unstable aerodynamic load. Static aerodynamic loading is desirable because it improves flight stability and performance. The unstable aerodynamic load is undesirable since it creates structural stress on the stabilizer wing 133 and the fragments of the rotorcraft 100 to which the stabilizer wing 133 is attached.

Unstable aerodynamic loads on stabilizer wing 133 are caused by changes in air flow rate and / or airflow direction and can cause lift changes in stabilizer wing 133. These lift changes on stabilizer wing 133 occur in response to changes in flow on the wing chord line - stabilizer 133.

Important changes in air velocity and direction occur in the rotor wake in the form of vortices. Vortexes are volumes of air that rotate around an elongated, usually curved axis that is the center of the vortex. In a rotorcraft 100, vortices come from the ends of the propeller blades 102a, 102b, 102c, 102d, 102e.

Due to the rotation of the rotor blades and the simultaneous forward movement in the direction of flight, the shape of the end vortex axis of the rotor is similar to that of a spiral when viewed from above. The helix, while being created at the ends of the rotor blades, moves backward relative to the rotorcraft 100 during forward flight at a speed almost identical to the flight speed. Some of the turbulence at the tips of the rotor blades will move through the stabilizer wing 133 and cause unstable aerodynamic loads.

Since the changes in lift on the stabilizer wing 133 arise in response to changes in flux on the quarter chord line of the stabilizer wing 133, unstable aerodynamic loads on the stabilizer wing 133 can be modeled approximately by the interaction between the vortices at the tips of the rotor blades and the quarter chord line of the stabilizer wing. 133.

As mentioned above, the chord quarter line of the stabilizer wing 133 has a non-zero curvature. Thus, the interaction between the vortices generated by the at least one rotor 102 and the quarter chord line can be timed compared to straight quarter chord stabilizer wings.

In fact, due to the helical structure, the vortices at the ends of the rotor blades periodically pass through the line of the chord quarters of the stabilizer wing 133. For a propeller with N propeller blades, N helical end vortex systems are created, and, therefore, N vortices at the ends of the rotor blades pass through the wing stabilizer 133 for each turn of the propeller. Accordingly, unstable aerodynamic loads on the stabilizer wing 133 occur periodically with a fundamental frequency N for each rotor revolution.

Due to the curved shape of the quarter-chord line, the shape of periodic unstable aerodynamic loads at the attachment of the stabilizer wing 133 on the keel 131 or tail boom 121 can have smoother peaks and reduced amplitudes compared to a stabilizer wing with a straight quarter-chord line. This is further illustrated in FIG. 3A-3C.

FIG. 2 shows an illustrative stabilizer wing having a non-zero curvature quarter-chord line, in accordance with some embodiments. As shown in FIG. 2, the stabilizer wing 200 may have an arcuate leading edge 210, an arcuate trailing edge 220, and a quarter-chord line 240 with non-zero curvature.

The expression "arcuate", when applied to leading edge 210, means that the leading edge 210 is curved at least partially between the wing tips 260. Preferably, the leading edge 210 is curved at least in the vicinity of the center line 230. If desired, the arcuate leading edge 210 may be partially straight.

Likewise, the expression "arcuate" when applied to trailing edge 220 means that trailing edge 220 is curved at least partially between the wing tips 260. Preferably, the trailing edge 220 is curved. If desired, the arcuate trailing edge 220 may be partially straight.

By way of example only, and not limiting the present embodiments accordingly, arcuate objects include at least partially C-shaped objects, objects having a shape like a segment of a parabola, objects having a shape like a segment of an ellipse, semi-oval objects, objects that are are rounded in some sections and straight in others, etc.

As shown in FIG. 2, the stabilizer wing 200 may have a wing span 255, which may be the distance between the left and right wing ends 260. Wingspan 255 can be the sum of left wing span 257 and right wing span 258, whereby left wing span 257 may be the distance between left wing 260 and centerline 230, and right wing span 258 may be the distance between right wing 260 and center line 230.

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

Consider a scenario in which a coordinate system is introduced as shown in FIG. 2 with the Y-axis as the lateral or transverse axis and the X-axis as the length or longitudinal axis. Consider additionally that the Y axis is scaled such that the left and right ends 260 of the wing coincide with coordinates Y = -1 and Y = 1, respectively, and that the origin is at center line 230. Consider additionally that the X axis should be scaled like this that the distance from the origin to X = 1 is equal to the span of the wing.

If desired, in the case of a stabilizer wing 200 with unequal left and right wing span, a separate coordinate system for each wing can be introduced. However, as shown in FIG. 2, the stabilizer wing 200 has equal wingspan of the left and right half-wings 257, 258 and a quarter-chord line 240 that is symmetrical to the fuselage centerline 230.

In this scenario, the quarter-chord line 240 can be described as a function of X = f (Y), whereby the curvature of the quarter-chord line 240 is the second derivative of d ² / dY ² (f (Y)). Since the X-axis is positive towards the leading edge 210, rearward bend (i.e., the bend opening is oriented toward the tail region of the rotorcraft (e.g., tail region 109 of rotorcraft 100 in FIG. 1)) corresponds to negative values (i.e. i.e., d ² / dY ² (f (Y)) <0) and forward bend (i.e., the bend opening is oriented toward the nose region of the rotorcraft (e.g., nose region 108 of rotorcraft 100 in FIG. 1 )) - positive values (i.e., d ² / dY ² (f (Y)) <0), while the straight line of quarter chords has zero curvature (i.e., d ² / dY ² (f (Y)) = 0).

As an 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 / dY ² (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 (ie, | d ² / dY ² (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 * | Y | (i.e., | d ² / dY ² (f (Y)) | <4-2 * | Y |).

Illustratively, the quarter-chord line 240 may have non-zero curvature (i.e., d ² / dY ² (f (Y)) ≠ 0) over the entire wing span 255. If desired, the quarter-chord line 240 may have non-zero curvature (ie, d ² / dY ² (f (Y)) ≠ 0) over the ranges Y =] 0.1; 0.9] and Y =] -0.1; -0.9] or parts of these ranges. As an example, the quarter-chord line 240 may have non-zero curvature in at least the ranges Y =] 0.1; 0.7] and / or Y =] -0.1; -0.7].

In other words, consider only the right wing wing (i.e., Y =] 0.1; 0.7], two points 272, 278 can be located on the quarter-chord line 240 at a distance of 282, 288 from the fuselage centerline 230, respectively so that the distances 282, 288 are 10% and 70% of the half-wing span 257, 258, respectively, with the result that the non-zero curvature of the quarter-chord line 240 is non-zero between at least two points 272, 278.

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

In some embodiments, the non-zero curvature of the quarter-chord line 240 varies from a first value at a first distance from fuselage centerline 230 to a second value at a second distance from fuselage centerline 230 such that the non-zero curvature of quarter-chord line 240 is not constant between the first and second distance from center line 230 fuselage.

As an example, a quarter-chord line 240 may have a curvature of 2.5 on centerline 230 (i.e., | d ² / dY ² (f (Y)) | = 2.5 at Y = 0) and decrease towards the left and the right ends 260 of the wing to a curvature of 0.5 (i.e., | d ² / dY ² (f (Y)) | = 0.5 at Y = ± 1). As another example, the quarter-chord line 240 may have a curvature of 3.0 at centerline 230 (i.e., | d ² / dY ² (f (Y)) | = 3.0 at Y = 0) and decrease towards left and right ends 260 of the wing to a curvature of 0.25 (i.e., | d ² / dY ² (f (Y)) | = 0.25 at Y = ± 1).

In some embodiments, the quarter-chord line 240 may be kinked. As an example, a kink at quarter-chord line 240 may be at centerline 230 (i.e., at Y = 0). As another example, the quarter-chord line 240 may have a kink towards the left and right wing ends 260 (eg, at Y = ± 0.95).

If desired, a family of embodiments can be conceived of the average curvature over a particular Y-band. Thus, the average curvature over the Y-value range (ie, in the range [Y1; Y2] is defined as the absolute value of the difference between the first derivative at the beginning of the range (ie, d / dY (f (Y); Y = Y1)) and the first derivative at the end of the range (i.e., d / dY (f (Y); Y = Y2) divided by (Y2-Y1). The mean curvature is therefore | (d / dy (f (Y); Y = Y1) - d / dY (f (Y); Y = Y2)) / (Y2-Y1) |.

Examples for such families of embodiments are shown in Tables 1, 2 and 3. Table 1 shows families of embodiments for Y1 = 0.1, Table 2 shows families of embodiments for Y1 = 0.2, and Table 3 shows families of embodiments for Y1. = 0.3. Additional families of embodiments are possible for different values of Y1 and / or Y2.

Table 1

| Y1 | = 0,1, | Y2 | = 0.9 | Y1 | = 0,1, | Y2 | = 0.8 | Y1 | = 0,1, | Y2 | = 0.7 | Y1 | = 0,1, | Y2 | = 0.5 The value of the average curvature ϵ] 0.2; 0.5] Family 1a Family 1b Family 1c Family 1d The value of the average curvature ϵ] 0.5; 1.0] Family 2a Family 2b Family 2c Family 2d The value of the average curvature ϵ] 1.0; 2.0] Family 3a Family 3b Family 3c Family 3d The value of the average curvature ϵ] 2.0; 4.0] Family 4a Family 4b Family 4c Family 4d

table 2

| y1 | = 0.2, | y2 | = 0.9 | y1 | = 0.2, | y2 | = 0.8 | y1 | = 0.2, | y2 | = 0.7 | y1 | = 0.2, | y2 | = 0.5 The value of the average curvature ϵ] 0.2; 0.5] Family 1e Family 1f Family 1g Family 1h The value of the average curvature ϵ] 0.5; 1.0] Family 2e Family 2f Family 2g Family 2h The value of the average curvature ϵ] 1.0; 2.0] Family 3e Family 3f Family 3g Family 3h The value of the average curvature ϵ] 2.0; 4.0] Family 4e Family 4f 4g family Family 4h

Table 3

| y1 | = 0.3, | y2 | = 0.9 | y1 | = 0.3, | y2 | = 0.8 | y1 | = 0.3, | y2 | = 0.7 | y1 | = 0.3, | y2 | = 0.5 The value of the average curvature ϵ] 0.2; 0.5] Family 1i Family 1k Family 1l Family 1m The value of the average curvature ϵ] 0.5; 1.0] Family 2i Family 2k Family 2l Family 2m The value of the average curvature ϵ] 1.0; 2.0] Family 3i Family 3k Family 3l Family 3m The value of the average curvature ϵ] 2.0; 4.0] Family 4i Family 4k Family 4l Family 4m

Variability in the Y1 coordinate may be important for stabilizer wings 200 that protrude from the fuselage, keel, or tail boom (eg, fuselage 104, keel 131, or tail boom 121 of rotorcraft 100 in FIG. 1). For example, a stabilizer wing that protrudes from the tail boom may not have Y-values that would be inside the tail boom. Therefore, an embodiment can be selected from a family that differs by a value for Y1 that lies outside the tail boom.

As mentioned above, a propeller with N rotor blades creates N helical end vortex systems of the rotor, and therefore N vortexes at the ends of the rotor blades pass through the stabilizer wing 200 with each rotation of the rotor. Accordingly, unstable aerodynamic loads on the stabilizer wing 200 occur periodically with a fundamental frequency N for each rotor revolution.

Due to the curved shape of the quarter-chord line, the shape of periodic unstable aerodynamic loads at the attachment point of the stabilizer wing 200 on the keel or tail boom can have smoother peaks and reduced amplitudes compared to a stabilizer wing with a straight quarter-chord line.

FIG. 3A, 3B, and 3C are diagrams of an exemplary time sequence of interactions between the centerline of an end vortex of a propeller and a quarter chord line with non-zero curvature. As shown in FIG. 3A, 3B, and 3C, the stabilizer wing 300 may have an arcuate leading edge 310, an arcuate trailing edge 320, and a quarter-chord line 340. The fuselage centerline 330 to which the stabilizer wing 300 is symmetrical is shown illustratively. Therefore, the quarter-chord line 340 is symmetrical to the center line 330.

Assuming the rotorcraft is on its landing gear, the propeller can rotate counterclockwise when viewed from above, or can rotate clockwise when viewed from above.

FIG. 3A shows the center line 360 of the end vortex of the propeller reaching the quarter-chord line 340. In this initial situation (i.e., T = T1), the propeller tip vortex center line 360 is tangent to the quarter chord line 340 and the interaction between the propeller tip vortex center line 360 and the quarter chord line 340 begins.

FIG. 3B shows the rotor tip vortex centerline 370 passing through the quarter-chord line 340 a short time after interaction has begun (i.e., at T = T2 with T2> T1).

FIG. 3C shows the propeller tip vortex centerline 380 when interaction with quarter chord line 340 ends and propeller tip vortex centerline 380 reaches the intersection of quarter chord line 340 and center line 330 (i.e., at T = T3 with T3> T2) ...

The center line 360, 370, 380 of the propeller tip vortex moves for a longer time (i.e., from T1 to T3) and a longer distance across the quarter-chord line of the stabilizer wing 300 compared to a comparable straight line of quarter-chords of a comparable wing stabilizer.

Thus, the center line 360, 370, 380 of the end vortex of the rotor interacts for a longer duration of time with the curved line 340 of the quarter-chords than it will interact with the straight line of the quarter-chords. The longer duration of interaction between the center line 360, 370, 380 of the propeller tip vortex and the quarter-chord line 340 implies, at one particular point in time, a relatively locally more limited influence on the quarter-chord line 340, and thus less aerodynamic load on stabilizer wing 300.

In contrast, if the duration of interaction is short, or if in the extreme case the center line 360 of the end vortex of the propeller and the line of quarter chords almost coincide only for a moment in time (for example, as may be more likely to occur in the case of a straight line of quarter chords as compared to the case correctly curved line of quarter chords), the change in aerodynamic load is instant when it occurs over the entire span of the stabilizer wing at almost the same time, thereby causing short and strong peaks in aerodynamic load.

The present embodiments of the stabilizer wing 300 with a curved quarter-chord line 340 allow long interaction times (e.g., T1 to T3) between the rotor tip vortex centerline 360, 370, 380 and stabilizer wing 300, creating a large difference in curvature and orientation. lines of 340 quarter chords from the curvature and orientation of the center lines 360, 370, 380 of the screw end swirls.

FIG. 3A, 3B and 3C show the interaction between the forward curved centerline 360, 370, 380 of the rotor tip vortex at times T1, T2 and T3, respectively, and the backward curved quarter-chord line 340.

The center lines of the end swirl of the forward bend rotor as shown in FIG. 3A, 3B, and 3C are generally more appropriate for a T-tail (i.e., a stabilizer wing that attaches at the top of the keel) in cruise flight because the forward curved centerlines of the propeller tip vortex are formed at the trailing edge of the propeller disc. , and the wing-stabilizer of the T-shaped tail is geometrically close to the trailing edge of the propeller.

Consequently, a T-tail stabilizer wing with a backward curved chord quarter may be less affected by rotor tip vortex centerlines formed at the trailing edge of the propeller disk compared to a straight chord quarter wing.

However, the screw also forms the center lines of the backward bend of the screw. For example, the propeller may form the propeller end swirl centerlines with a rear bend at the leading edge of the propeller disk.

A stabilizer wing that attaches such that it interacts in many flight conditions with vortices formed at the leading edge of the propeller disk can interact for a longer time with these backward-curved centerlines of the propeller tip vortex if the quarter-chord line of the stabilizer wing has front bend. Consequently, a forward bend stabilizer wing may experience lower aerodynamic load peaks when subjected to the center lines of the rear bend rotor end vortex.

Consider a scenario in which the rotorcraft is a single rotor helicopter. Consider additionally that the main rotor center for the main rotor coincides with the center line of the fuselage (for example, the center line 103 of the main rotor coincides with the center line 130 of the fuselage 104 in FIG.1C), and that, due to the asymmetric nature of the main rotor relative to the center plane of the helicopter, the center lines the end vortexes of the rotor, propagated by the main rotor, are not symmetrical to the center line of the fuselage.

In this scenario, the shape of the curvature of the quarter-chord line 340, and thus the plan shape of the stabilizer wing 300, can be adjusted in the direction of minimizing aerodynamic loads. As an example, the plan view of the stabilizer wing 300 may be designed such that, based on the action of the rotorcraft, the total aerodynamic loads (e.g., obtained by integrating over the wingspan for the stabilizer wing 300) at a predetermined location (e.g., stabilizer wing attachment positions 300) do not exceed a predetermined value.

As another example, the planar shape of the stabilizer wing 300 may be designed such that, based on the action of the rotorcraft, a time sequence of local interactions between propeller tip vortex centerlines 360, 370, 380 and quarter chord line 340 at each wing location the stabilizer 300 is suitably biased to achieve a nearly constant total aerodynamic load at the appropriate location of the stabilizer wing 300 over time.

As shown in FIG. 2, 3A, 3B and 3C, the stabilizer wing may be tapered. In other words, chord length 250 may be longer at centerline 230 compared to wing ends 260, or the ratio of chord length 250 at wing ends 260 divided by chord length 250 at centerline 230 may be less than one. As an example, the ratio of the chord length 250 at the wing tips 260 divided by the chord length 250 at the center line 230 may be between 0.2 and 0.8.

However, even if the stabilizer wing 200, 300 shown in FIG. 2, 3A, 3B and 3C, tapers, the curved line 240, 340 quarter chords can be combined with any kind of wing span chord length calculation.

By way of example, FIG. 4 is a sequence diagram of propeller tip vortex centerlines 460, 470, 480 reaching an exemplary stabilizer wing 400 with a constant chord length 450 in accordance with some embodiments. In other words, the stabilizer wing 400 may have a constant chord length 450 between the left and right wing ends 457 for the stabilizer wing 400.

As shown in FIG. 4, the stabilizer wing 400 may have an arcuate leading edge 410, an arcuate trailing edge 420, and a quarter-chord line 440 with non-zero curvature. Illustratively shown is the fuselage centerline 430 to which the stabilizer wing 400 is symmetrical. Therefore, the quarter-chord line 440 is symmetrical to the center line 430.

As an example, chord length 450 may be constant at any Y node along the wingspan (i.e., anywhere between Y = -1 and Y = 1). In this example, since the stabilizer wing 400 has an arcuate leading edge 410 and a constant chord length 450 at any span Y node, the leading edge 410, trailing edge 420, and quarter-chord line 440 have the same arcuate shape and the same non-zero curvature.

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

By way of example, FIG. 5 is a sequence diagram of propeller tip vortex centerlines 560, 570, 580 reaching an exemplary stabilizer wing 500 having an asymmetric sweep between left and right half wings 504, 502, in accordance with some embodiments.

In fact, as shown in FIG. 5, the stabilizer wing 500 may have a left wing 504 that has a first sweep and a right wing 502 that has a second sweep that is different from the first sweep. By way of example, the left wing 504 may have a lower sweep than the right wing 502. If desired, the left wing 504 may have a greater sweep than the right wing 502.

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

As an example, stabilizer wing 500 may have a constant chord length of 550. If desired, the chord length 550 can be constant at any Y node along the wingspan (i.e., anywhere between Y = -1 and Y = 1). In this example, since the stabilizer wing 500 has an arcuate leading edge 510 and a constant chord length 550 at any span Y node, the leading edge 510, trailing edge 520, and quarter chord line 540 have the same arcuate shape and the same non-zero curvature.

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

For example, right wing 502 may have a stronger curvature of quarter chord line 540 than left wing 504. If desired, right wing 502 may have less curvature of quarter chord line 540 compared to left wing 504.

If desired, the left wing 504 may have a first length and the right wing 502 may have a second length that is different from the first length. In other words, one of the left wing 504 and the right wing 502 can provide more than 50% of the wing span 555. As an example, the left wing 504 may be shorter than the right wing 502. As another example, the right wing 502 may be shorter than the left wing 504.

The lengths of the right and left half-wings 502, 504 can be chosen to balance the proportion of aerodynamic loads on the right and left half-wings 502, 504 at the moment of roll near the attachment of the stabilizer wing 500 (for example, on the keel, tail boom or fuselage of a rotorcraft). The reduction of either the static and / or the dynamic component of the aerodynamic load at the attachment point of the stabilizer wing can be determined based on design needs.

If desired, the stabilizer wing 500 may include end airfoils (eg, end airfoils 138 in FIG. 1). The aerodynamic end surfaces can increase the aerodynamic lift of the stabilizer wing 500 without increasing the wing span 555.

Since the wingspan 555, and more specifically the lengths of the right and left half-wings 502, 504, respectively, determines the effective lever arm for the roll moments at the attachment of the stabilizer wing 500, and since the orientation of the end airfoil is perpendicular to the center lines 560, 570, 580 of the propeller end vortex, which limits the interaction between the propeller end vortex centerlines 560, 570, 580 and the aerodynamic end surfaces, the use of aerofoil end surfaces has the potential to reduce heeling moment fluctuations at the attachment of the stabilizer wing 500 for a given aerodynamic lift goal.

Asymmetry in curvature of quarter-chord line 540 and / or bend between right wing 502 and left wing 504 may be used to further reduce unstable aerodynamic loads on stabilizer wing 500.

For example, the selection of a combination of a predetermined curvature of the quarter-chord line 540 and a predetermined sweep for the left and right half-wings 504, 502 may provide the ability to independently adjust the time sequence of interactions between the centerlines 560, 570, 580 of the propeller tip vortexes along the Y nodes along the wingspan for the left and the right half-wings 504, 502 of the stabilizer wing 500.

To explain the action, the previously symmetrical stabilizer wing can be considered rotated by some angle Ψ relative to the center line of the fuselage, resulting in an asymmetric sweep.

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

As an example, the stabilizer wing 600 may have a tapered chord length 650. If desired, the chord length 650 can decrease from the wing centerline 635 towards the wing tips (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 a different arcuate shape and different non-zero curvature.

As shown in FIG. 6, the stabilizer wing 600 may be symmetrical with respect to the wing centerline 635. Since the center line 635 of the wing is rotated by an angle Ψ from the center line 630 of the fuselage, the stabilizer wing 600 can be asymmetrical with respect to the center line 630 of the fuselage.

Since the wing centerline 635 is perpendicular to the propeller end vortex centerline sequence 660, 670, 680, each propeller end vortex centerline 660, 670, 680 reaches both wing hemi of stabilizer wing 600 at the same time. Thus, bending moments about the wing centerline 635 for the stabilizer wing 600 occur synchronously in time.

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

As an example, an angle ψ between 0 ° and 30 ° may be selected for a single rotor rotorcraft that has a rotor center on the fuselage centerline 630 in front of the stabilizer wing 600 and rotates counterclockwise when viewed from above. As another example, an angle ψ between 0 ° and -30 ° may be selected for a rotorcraft having a single rotor that has a rotor center on the centerline 630 of the fuselage in front of the stabilizer wing 600 and rotates clockwise when viewed from above.

The resilience of the stabilizer wing 200, 300, 400, 500, or 600 shown in FIG. 2-6 can affect aerodynamic loading because the curvature of the quarter-chord line 240, 340, 440, 540, 640 can lead to coupled reaction in bending and twisting. Thus, the vertical bending of the stabilizer wing 200, 300, 400, 500, 600 due to aerodynamic loads may be accompanied by elastic twisting of the stabilizer wing about the Y axis. Providing a stabilizer wing that allows elastic twisting and / or that has a twisted structure around the Y-axis at the point of attachment of the stabilizer wing, can reduce aerodynamic loads, since twisting can change the angles of attack and, therefore, the aerodynamic loads on the stabilizer wing.

If desired, the stabilizer wing 200, 300, 400, 500 or 600 in FIG. 2-6 can have twist, which is the change in angles between the chord lines and the horizontal XY plane for different Y nodes in the wingspan. In other words, the stabilizer wing 200, 300, 400, 500 or 600 may have a variable pitch angle at least in a portion of the wing span.

If desired, the stabilizer wings 200, 300, 400, 500, or 600 in FIG. 2-6 may have twists in addition to a line of 240, 340, 440, 540, or 640 quarter-chords with non-zero curvature. Additional twisting of the stabilizer wings 200, 300, 400, 500 or 600 can compensate for the situation of asymmetric inflow flow on the left and right half wings of the corresponding stabilizer wing 200, 300, 400, 500 or 600.

Consider a scenario in which the rotor wake is not symmetrical about the centerline of the fuselage. For example, as shown in FIG. 5, the rotor wake that forms the rotor end vortex centerlines 560, 570, 580 is not symmetrical with respect to the plane defined by the fuselage centerline 530 and the z-axis.

In this scenario, propeller tip vortex centerlines 560, 570, 580 may lead to asymmetric mean aerodynamic loads between left wing 504 and right wing 502 in the case of a non-twisted horizontal stabilizer wing.

As an example, the stabilizer wing 500 may equalize the time average for lift and moment at the root of the left and right wings 504, 502 at the attachment of the stabilizer wing 500 for most appropriate flight conditions (e.g., forward level flight) by means of a curved the quarter-chord lines 540 and the variable pitch angle along the wingspan 555, thereby reducing the combined static roll moment that appears to be the left and right half-wings 504, 502 at the attachment point of the stabilizer wing 500.

If desired, the stabilizer wing 200, 300, 400, 500, 600 in FIG. 2-6 may include wing-span distribution of pitch angles. Span distribution of pitch angles can reduce the risk of stall at the root of the stabilizer wing. The risk of stalling increases with the magnitude of the negative pressure on the stabilizer wing.

Illustratively, the angle of attack may be decreased in the direction away from the wing tips for the stabilizer wing 200, 300, 400, 500, or 600 in FIG. 2-6 to the root of the stabilizer wing in order to reduce the amount of negative pressure created on the lower surface of the root of the stabilizer wing.

The portion of the stabilizer wing surface that is close to the root of the stabilizer wing is particularly prone to stalling when the stabilizer wing is mounted on the keel of the rotorcraft. In fact, the low-pressure region of the keel extends to the portion of the stabilizer wing surface close to the root of the stabilizer wing, in addition to the low pressure caused by the stabilizer wing itself.

The keel of the rotorcraft can provide counter torque against the propeller torque. For example, the keel can generate a force to the right (i.e., in the positive Y-direction) in a rotorcraft with a propeller that rotates counterclockwise when viewed from above. In fact, the force generated by the keel is achieved by negative pressure on the right side of the keel.

In this example, a zero degree pitch angle between the chord of the stabilizer wing and the XY plane could be on the right side of the keel (i.e., at a positive Y wing span node). If desired, the pitch angle between the chord of the stabilizer wing and the XY plane may increase continuously from the wing span node, where the pitch angle is zero degrees, to the left wing end (i.e., in the negative Y-direction) and steadily decreases from span, where the pitch is zero degrees, toward the right wing end (i.e., in the positive Y-direction).

Alternatively, for a clockwise rotating propeller, when viewed from above, the left surface of the keel will be subject to negative pressure. Accordingly, a zero degree pitch angle between the chord of the stabilizer wing and the XY plane may be on the left side of the keel (i.e., at a negative Y wing span node).

If desired, the pitch angle between the chord of the stabilizer wing and the XY-plane may increase continuously from the wing span node, where the pitch angle is zero degrees, to the right wing end (i.e., in the positive Y-direction) and decreases steadily from wing span, where the pitch is zero degrees, to the left end of the wing (i.e., in the negative Y-direction).

The pitch distributions over the entire wingspan of the stabilizer can be shifted generally along the ordinate (i.e., shifted along the Y-axis to uniformly increase or decrease the pitch angle over the entire wing span) without altering the oscillation along the abscissa (i.e. ., twisting along the X axis), but only the general tilt and, thus, the lift of the entire stabilizer wing.

As an example, the pitch angle may have a greater change in the direction of the root of the stabilizer wing as compared to the change in the direction of the ends of the stabilizer wing. In other words, for a given pitch angle θ and a transverse Y-axis from Y = -1 at the left wing end to Y = 1 at the right wing end for the stabilizer wing, the change in pitch angle dθ / dY is large compared to the mean twist (t ie, θ (Y = -1) - θ (Y = 1)) / 2) in the area along the wingspan Y = [-0.5 ... 0.5].

Illustratively, the wing span pitch distribution of the stabilizer can have predefined upper and lower bounds. If desired, an actual pitch angle distribution over the wingspan of the stabilizer can be selected that lies between predetermined upper and lower bounds. The selection of the actual pitch angle distribution along the stabilizer wing can be made based on the prevailing flight mode, for example.

It should be noted that the above-described embodiments are merely described to illustrate possible implementations of the present invention, and not to limit the present invention to them. Instead, many modifications and variations of the present embodiments are possible and should therefore also be considered part of the invention.

As an example, the asymmetric stabilizer wing 500 in FIG. 5 is shown with a constant chord length of 550. However, the asymmetrical stabilizer wing 550 can be tapered if desired. Similarly, the stabilizer wing 600 of FIG. 6 is shown tapering. However, the stabilizer wing 600 may have a constant chord length if desired.

As another example, the stabilizer wings 200, 300, 400, 500, 600 in FIG. 2-6 are shown with a predefined sweep. If desired, the stabilizer wings 200, 300, 400, 500, 600 can have different sweep shapes. For example, stabilizer wings 200, 300, 400, 500, 600 may have a sweep greater or less sweep than that shown in FIG. 2-6.

In addition, the lower bound for the absolute value of the non-zero curvature of the quarter-chord line 240 of the stabilizer wing 200 in the drawing was set to 0.2 (i.e., | d ² / dY ² (f (Y)) |> 0.2) , and the upper bound for the absolute value of the non-zero curvature of the chord quarter line 240 was specified as 4 (i.e., | d ² / dY ² (f (Y)) | <4).

However, the lower bound for the absolute value of the non-zero curvature of the quarter-chord line 240 may be selected greater than or less than 0.2. Likewise, the upper bounds for the absolute value of the non-zero curvature of the quarter-chord line 240 can be selected greater than or less than 4. If desired, the lower and upper bounds for the absolute value of the non-zero curvature of the quarter-chord line 240 may be selected based on the shape of the center lines of the screw tip curvatures.

Likewise, 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 center line 230, and the factor by which the upper limit is decreased may be the factor A, which may be selected from] 0; B [, with B being the upper bound for the absolute value of the non-zero curvature of the 240 quarter-chord line. For example, the absolute value of the nonzero curvature of the line 240 of quarter-chords may be less than B - (BC) * | Y |, while B ϵ [1; 4] and C ϵ [0; 1] (that is, | d ² / dY ² (f (Y)) | <B - (BC) * | Y |).

List of reference positions

100 rotorcraft

102 main rotor

102a, 102b, 102c, 102d, 102e propeller blades

103 main rotor center

103a main rotor axis

104 fuselage

104a cabin

107 fuselage nose

108 nasal area

109 tail region

121 tail boom

122 tail rotor

130 center line

131 keel

132 rudder

133 tail wing, stabilizer wing

136 leading edge

137 trailing edge

138 end aerofoil

160 wing end

200 wing stabilizer

210 leading edge

220 trailing edge

230 center line

240 quarter chord line

250 chord length

255 wingspan

257 wing span (left wing)

258 wing span (right wing)

260 wing end

270 distance

272, 274, 276, 278 points on the chord quarter line

282 distance equal to 10% of the wingspan

284 distance equal to 30% of the wingspan

286 distance equal to 50% of the wingspan

288 distance equal to 70% of the wingspan

290 straight line

300 stabilizer wing

310 leading edge

320 trailing edge

330 center line

340 quarter chord line

360, 370, 380 center line of screw end swirls

400 wing stabilizer

410 leading edge

420 trailing edge

430 center line

440 quarter chord line

450 chord length

457 wing end

460, 470, 480 center line of screw end swirls

500 wing stabilizer

502 right wing

504 left wing

510 leading edge

520 trailing edge

530 center line

540 quarter chord line

550 chord length

555 wingspan

560, 570, 580 center line of screw end swirls

600 wing stabilizer

610 leading edge

620 trailing edge

630 fuselage centerline

635 centerline of the wing

640 chord quarter line

650 chord length

660, 670, 680 center line of screw end swirls

X longitudinal axis, length axis

Y transverse axis, side axis, width axis

Z vertical axis, height axis

Claims (28)

1. Rotary-wing aircraft (100), containing: fuselage (104) having a center line (130); at least one rotor (102) that generates vortices (460, 470, 480) during operation; and a stabilizer wing (133), which is connected to the fuselage (104); characterized in that: stabilizer wing (133, 400) equipped with: the left end (160, 457) of the wing, the right end (160, 457) of the wing, the line (440) of the quarter chords with nonzero curvature, and the nonzero curvature of the line (440) of the quarter chords leads to reduced unstable aerodynamic loads caused by the wake of at least one main rotor (102) on the stabilizer wing (133, 400), so that the interaction between the vortices (460, 470, 480), formed by at least one main rotor (102), and the line (440) of the quarter chords is distributed in time, the leading edge (410), which is arcuate, and trailing edge (420), which is arcuate. 2. The rotorcraft (100) according to claim 1, in which the fuselage (104) extends in the longitudinal direction (X) between the nose region (108) and the tail region (109), and the line of the quarter-chords of the stabilizer wing (133 ) is closer to the tail region (109) at least one of the left or right ends (160) of the wing than the center line (130) of the fuselage (104). 3. Rotary-wing aircraft (100) according to claim 1, in which the fuselage (104) extends in the longitudinal direction (X) between the nose region (108) and the tail region (109), and the line of the quarter-chords of the stabilizer wing (133 ) is closer to the nose area (108) at least one of the left or right ends (160) of the wing than the center line (130) of the fuselage (104). 4. The rotorcraft (100) according to any one of the preceding claims, wherein the absolute value of the non-zero curvature of the quarter chord line (240) is greater than 0.2, the quarter chord line (240) is a function and the curvature is the second derivative of the function. 5. The rotorcraft (100) according to claim 4, wherein the absolute value of the non-zero curvature of the chord quarter line (240) is less than 4. 6. The rotorcraft (100) according to any one of the preceding claims, wherein the non-zero curvature of the quarter-chord line (240) varies 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), so that the non-zero curvature of the quarter-chord line (240) is variable between the first and second distance from the center line (230) of the fuselage (104). 7. The rotorcraft (100) according to any one of the preceding claims, in which the stabilizer wing (200) has a wing span (257, 258) between the centerline (230) and the wing tip from the left or right wing ends (260), with the first, second, third and fourth points (272, 274, 276, 278) are located on the line (240) of the quarter chords at the first, second, third and fourth distances (282, 284, 286, 288) from the center line (230) fuselage, respectively, while the first, second, third and fourth distances (282, 284, 286, 288) are 10%, 30%, 50% and 70% of the wing span (257, 258), respectively, with a nonzero curvature of the line (240) of the quarter chords is nonzero between the first and fourth points (272, 278), and the straight line (290) between the first and fourth points (272, 278) has a fifth distance (270) from the line (240) of the quarter chords between the second and the third point (274, 276), which is more than 2% of the span (257, 258) of the half wing. 8. The rotorcraft (100) according to any one of the preceding claims, wherein the quarter-chord line (240) is a symmetrical fuselage centerline (230). 9. The rotorcraft (100) according to any one of the preceding claims, in which the stabilizer wing (133) is mounted on at least one of the fuselage (104), the tail boom (121), which is attached to the fuselage (104), or the vertical keel (131), which is attached to the tail boom (121). 10. The rotorcraft (100) according to any one of the preceding claims, wherein the stabilizer wing (400) has a constant chord length (450) between the left and right ends (457) of the stabilizer wing (400). 11. Rotary-wing aircraft (100) according to any one of paragraphs. 1-9, in which the stabilizer wing (200) has a chord length (250) that decreases from the center line (230) of the fuselage to the left and right ends (260) of the wing for the stabilizer wing (200). 12. Rotary-wing aircraft (100) according to any one of paragraphs. 1-7, in which the stabilizer wing (500) additionally contains: left wing half (504), having the first sweep; and the right wing half (502) having a second sweep, which differs from the first sweep. 13. Rotary-wing aircraft (100) according to claim 12, in which the left wing (504) has a line (540) of quarter chords with a first curvature, and the right wing (502) has a line (540) of quarter chords with a second curvature, and in this, the second curvature differs from the first curvature. 14. The rotorcraft (100) according to claim 12, wherein the left wing (504) has a first length and the right wing (502) has a second length that differs from the first length. 15. The rotorcraft (100) according to any one of the preceding claims, wherein the stabilizer wing (133) further comprises end aerodynamic surfaces (138). 16. The rotorcraft (100) according to any one of the preceding claims, wherein the rotorcraft is in the form of a helicopter.

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