Background Of The Invention In addition to the main and tail rotor systems, a rotorcraft typically includes various lifting/moment-producing airfoil structures/surfaces for stabilizing the rotorcraft in various flight regimes. One such airfoil structure, to which the present invention is directed, is the horizontal stabilizer which is typically mounted to a rearward portion of the tail boom assembly. The principle function of the horizontal stabilizer is to produce a downwardly-directed force vector, i. e., negative lift, when the rotorcraft is in a forward flight operating mode so as to produce a pitching moment which counteracts or balances the pitching moment produced by the main rotor system. That is, the horizontal stabilizer counteracts the nose-down pitching moment produced by the main rotor system in forward flight so as to"level out"the fuselage and, consequently, reduce profile drag.

Prior art configurations include these and other aerodynamic and/or structural benefits. For example, the horizontal stabilizer may be fixed or pivotable with respect to the tail boom assembly. Generally, to avoid complexity, it is desirable to employ a fixed horizontal stabilizer which is optimally oriented and configured for a single, most frequently occupied, operating regime. Examples of fixed horizontal stabilizers are described and depicted in U. S. Patents 5,102,067 and Des. 361,053. Alternatively, it may be desirable to employ a pivotable horizontal stabilizer for the purpose of minimizing impingement of the rotor downwash on the horizontal stabilizer. For example, by controllably pivoting the stabilizer, it may assume a position parallel to the freestream downwash to minimize profile area, and consequently, profile drag. This capability is particularly valuable in a hover or low-speed operating mode wherein, if the stabilizer were fixed in a horizontal plane, the downwash would impinge on the entire planform

area of the stabilizer. Examples of pivotable horizontal stabilizers are described and depicted in U. S. Patents 4,103,848, and 4,247,061.

A prior art horizontal stabilizer typically defines a rectangular planform and a substantially constant angle of incidence along its length, i. e., the angle of incidence from one cross-section to another is equal. Furthermore, in cross-section, such horizontal stabilizers commonly comprise conventional symmetric or cambered airfoil shapes which may include a constant or tapering thickness dimension. To avoid the introduction of large bending moments at the structural interface of the horizontal stabilizer, it is generally desirable to attach the horizontal stabilizer in substantially symmetrical relation to is adjoining structure, e. g., a vertical fin. That is, it is desirable to mount the horizontal stabilizer about its center of mass to balance the bending moments acting on the structural interface.

Summary of the Invention It is an object of the present invention to provide a horizontal stabilizer for rotorcraft which reduces bending moment loads/stresses acting on various structural interfaces and on the tail boom assembly of the rotorcraft for reducing the structural weight requirements thereof.

This objective is achieved by a horizontal stabilizer defining first and second spanwise stations wherein the first spanwise station defines a first angle of incidence and the second spanwise station defines a second angle of incidence and wherein the angles of incidence from one station to the other are different, e. g., one is greater than the other.

Various embodiments of the horizontal stabilizer include the use of vertically extending tabs along the trailing edge of the horizontal stabilizer, a stepped-transition to abruptly change the angle of incidence from one station to another, and a distributed twist which gradually changes the angle of incidence.

Brief Description Of The Drawings A more complete understanding of the present invention and the attendant features and advantages thereof may be had by reference to the following detailed description of the invention when considered in conjunction with the following drawings wherein:

Fig. la depicts a side view of a rotorcraft employing a horizontal stabilizer according to the present invention; Fig. lb depicts a top view of the rotorcraft illustrated in Fig. la; Fig. 1 c depicts a rear view of the rotorcraft illustrated in Fig. 1 a; Fig. 2 depicts an isolated perspective view of the horizontal stabilizer according to the present invention including upwardly and downwardly extending tabs along the trailing edge of the horizontal stabilizer; Fig. 2a is a cross-sectional view taken substantially along line 2a-2a of Fig. 2; Fig. 2b is a cross-sectional view taken substantially along line 2b-2b of Fig. 2; Fig. 3 depicts an isolated perspective view of an alternate embodiment of the inventive horizontal stabilizer including a stepped transition along a midplane thereof ; Fig. 3a is a cross-sectional view taken substantially along line 3a-3a of Fig. 3; Fig. 3b is a cross-sectional view taken substantially along line 3b-3b of Fig. 3; Fig. 4 depicts an isolated perspective view of an alternate embodiment of the inventive horizontal stabilizer which defines multiple angles of incidence, i. e., a distributed twist, along the span of the horizontal stabilizer; Fig. 4a is a cross-sectional view taken substantially along line 4a-4a of Fig. 4; Fig. 4b is a cross-sectional view taken substantially along line 4b-4b of Fig. 4; Fig. 4c is a cross-sectional view taken substantially along line 4c-4c of Fig. 4; Fig. 4d is a cross-sectional view taken substantially along line 4d-4d of Fig. 4; and Fig. 5 is a graph comparing the bending moment loads/stresses produced by a prior art horizontal stabilizer and the horizontal stabilizer according to the present invention.

Best Mode For Carrying Out The Invention Referring now to the drawings wherein like reference characters identify corresponding or similar elements throughout the several views, Figs. la-lb show a rotorcraft 10 having a main rotor 12, a tail boom assembly 14 employing a FANTAILTM anti-torque system 16, a vertical fin 18, and a horizontal stabilizer 20 according to the present invention. The horizontal stabilizer 20 is fixedly mounted to the upper end of the vertical fin 18 and defines a"mean angle of attack"of about seven (7) degrees relative to

the horizontal. As will become apparent in light of the subsequent discussion, it is useful to describe the"angle of attack"in terms of a mean value since it is a principle teaching of the invention that the horizontal stabilizer 20 defines at least two (2) angles of incidence at two spanwise stations, and. in at least one embodiment of the invention, defines multiple angles of incidence to form a distributed twist along is span.

Before discussing the structural and functional features of the inventive horizontal stabilizer 20, it is useful to understand certain aerodynamic characteristics which heretofore were not fully appreciated/understood. More specifically, referring to Figs. lb and Ic, during flight testing it was discovered that the helicopter 10 experienced unacceptably high bending moment loads/stresses at the structural interfaces 22,24 (Fig. I c) between the horizontal stabilizer 20 and the vertical fin 18, and between the vertical fin 18 and the tail boom assembly 14. Conventionally, structural augmentation of these interfaces 22,24 would be required to react the imposed bending moments, thereby incurring weight penalties. In contradistinction to the conventional solution, the inventor sought to correct the problem via analyzing the freestream airflow acting on the horizontal stabilizer 20.

The inventor discovered that for a main rotor 12 having a counter-clockwise rotation, the downwash 26 skews to the right as illustrated in Fig. Ib. Furthermore, due to the normal coning of the main rotor 12. the downwash 26 produces strong vortices which impinge on the stabilizer in an asymmetric manner. More specifically, and referring to Fig. lc, the vortex pattern 28 illustrates that the high energy vortices 30 are distally spaced relative to the left side 20L of the stabilizer 20 and are in close proximity to the stabilizer 20 on its right side 20R.

The various aerodynamic effects on the stabilizer 20 can be summarized as follows. When the downwash 26 skews to one side of the stabilizer 20 as shown in Figs. lb, a region A thereof operates at a different angle of attack than in a region B, which is operating in substantially less-affected or"clean"airflow. The affected airflow in region A tends to increase the angle of attack of the stabilizer producing an excess force V1 on the right side 20R relative to the left side 20L. In contrast, the clean airflow in region B, causes this portion of the stabilizer 20 to function as intended and produces a relatively small downwardly directed force vector V2 (Fig. 1 c) when the aircraft is in

level unaccelerated flight. Accordingly, it will be appreciated that an imbalance is produced resulting in a bending moment M about the structural interfaces 22,24.

In Figs. 2,3, and 4, exemplary embodiments of the inventive stabilizer 20 are shown in isolated perspective. In all described embodiments, the stabilizer 20 has a substantially rectangular planform and comprises a one piece construction, i. e., incorporates a common spar (not shown) along it length or span. Furthermore, the stabilizer 20 comprises a conventional airfoil shape, i. e., symmetric or cambered about its chord, though the angle of incidence from one spanwise station to another varies. While the term"angle of attack"could also be used, this is a term which is relative to the free- stream airflow and not to the horizontal stabilizer 20 in isolation. Hence, the inventor uses the term"angle of incidence"for defining the characteristics of various spanwise stations of the same stabilizer 20. Furthermore, the"angle of incidence"is defined as the angle defined by geometric chord of the airfoil section and the horizontal.

In the broadest sense of the invention, the horizontal stabilizer 20 defines at least two spanwise stations wherein the angle of incidence is different, i. e., one is greater than the other. In Figs. 2,2a, and 2b, a first embodiment of the horizontal stabilizer 20 is shown wherein the left side 20L of the stabilizer 20 comprises an upwardly extending flap or tab 36 along its trailing edge 20TE while the right side 20R comprises a downwardly extending flap or tab 38. Consequently, the angle of incidence X (Figs. 2a and 2b) on the left side 20L is different, e. g., less than, the angle of incidence Q2 on the right side 2OR. In operation, this feature changes the effective camber of the airfoil to compensate for differences in local angle of attack whereby the magnitude of the downwardly directed force vectors V1 and V2 are decreased and increased, respectively. Consequently, in the previously described flow environment illustrated in Figs. lb and Ic, bending moment loads/stresses are mitigated or balanced so as to reduce the bending moment stresses at the structural interfaces.

Yet another embodiment of the horizontal stabilizer 20 is depicted Figs. 3,3a, and 3b, wherein the stabilizer 20 comprises a stepped-transition at a predefined spanwise station. In the described embodiment, the stepped-transition occurs at the midplane MP of the horizontal stabilizer 20. In this embodiment, the left side 20L of the stabilizer defines a nose down angle of incidence 3 (Fig. 3a) relative to the angle of incidence 4 on the right side 20R (Fig. 3b). The functional effects of the stepped-transition are the

same as those previously described for the upwardly and downwardly extending tabs 36,38.

In Figs. 4,4a-4d, the same functional effect may be produced by a uniform or non-uniform twist distribution at a multiplicity of spanwise stations of the stabilizer 20.

More specifically, and referring to Figs. 4a-4d, the horizontal stabilizer 20 defines a multiplicity of incidence angles P5-+8 so as to define a substantially nose down angle of incidence +5 on the left side 20L of the stabilizer 20 (Fig. 4a), relative to the angle of incidence fus on the right side 20R (Fig. 4d). In the described embodiment, the stabilizer 20 defines a substantially uniform twist which effects a gradual change in angle of incidence when viewing the cross-sections in sequence, e. g., from Fig. 4a through Fig. 4d.

Fig. 5 is a graph which shows the balancing or normalizing effects of the embodiment depicted in Fig. 2. The airspeed of the rotorcraft is plotted along the abscissa in KNOTS, and the bending moment (along the structural interface between the stabilizer and the vertical fin) is plotted along the ordinate axis in LBS/IN. The triangular data points 50 are indicative of a prior art stabilizer having a constant angle of incidence along its span while the square data points 60 are indicative of the horizontal stabilizer 20 according to the present invention. It will be appreciated by comparing the data trends of each that the stabilizer of the present invention produces relatively low steady bending moment stresses of about 4,000 lbs/in as the rotorcraft approaches 140 knots while the prior art imposes more than twice this bending moment stresses, i. e.. about 10,000 lbs./in.

In summary, the present invention varies the angle of incidence, and, consequently, the effective angle of attack, to reduce or cancel the bending moment loads/stresses acting on the structural interfaces 22,24 and, additionally, on the entire tail boom assembly 14. Accordingly, the invention provides a structurally efficient solution for reducing the overall weight of the rotorcraft.

Although the invention has been shown and described with respect to exemplary embodiments thereof, it should be understood by those skilled in the art that other changes, omissions and additions may be made therein and thereto, without departing from the spirit and scope of the present invention. For example, while the invention has been described in the context of a unitary structure, it will be appreciated that the horizontal stabilizer may comprise multiple sections. Furthermore, while the inventive

horizontal stabilizer is shown mounted atop a vertical fin of the tail boom assembly, it will be appreciated that the stabilizer may be mounted to any portion thereof.

Furthermore, the teachings of the invention are equally applicable to fixed, moveable or cantilevered stabilizer configurations. For example, it may be desirable to produce axisymmetric lift characteristics on a fixed cantilevered stabilizer such as that shown in U. S. Patent Des. 361,053.

While the detailed description depicts three embodiments of the invention for varying the angle of incidence, the invention contemplates any means known in the art of aerodynamics for producing such effect. For example, controllably blowing air over the surface of the stabilizer may be used at various spanwise stations to alter the stagnation points on the airfoil thereby varying its lift characteristics. Furthermore, while in one of the embodiments described herein, vertically extending tabs are employed on both sides of the horizontal stabilizer to varying its lift characteristics, it will be appreciated that a single tab disposed on one side of the stabilizer may be used. Furthermore, while the tabs are shown to be substantially vertically aligned and of constant height, it should be understood that the tabs may incorporate a horizontal component, i. e., not entirely vertical, and may vary in height along the span of the stabilizer. While in another embodiment, a single stepped-transition occurs at a midplane of the horizontal stabilizer, it will be appreciated that the invention may employ multiple steps disposed at any spanwise station. While in yet another embodiment described, a uniform distributed twist is employed, the invention contemplates a non-uniform twist distribution, i. e., wherein the change of angle of incidence varies from, for example, rapid to shallow, along the span. Moreover, while the invention has been described in the context of a rotorcraft having a rotor system driven in a counter-clockwise direction, it will be appreciated that the teachings are equally applicable to a rotor system having a clock- wise rotation.

What is claimed is: