Multi-hedral rotary wing

Abstract

A rotary-wing includes a multi-hedral tip section which utilizes a distribution of anhedral and/or dihedral angles that cause a tip vortex to be axially displaced relative to a region on a following blade such that the tip vortex generally avoids influence upon the following blade.

Description

BACKGROUND OF THE INVENTION

The present invention relates to aerodynamic structures, and more particularly to multi-hedral rotor blade, which reduces tip vortex influence on a following blade.

Aerodynamic surfaces produce tip vortices as an artifact of flow. During typical rotorcraft flight operations, rotor blades of a main rotor assembly, due to the airfoil profile and angle of attack of the rotor blades, create a high velocity low pressure field over the upper aerodynamic surface of each rotor blade and a low velocity high pressure field over the lower aerodynamic surface of each rotor blade. At the tip of each rotor blade, this pressure differential effectively engenders airflow circulation from the high pressure field to the low pressure field to create a tip vortex.

During rotorcraft flight operations, the tip vortex is shed from a preceding rotor blade and at least partially interferes with a following rotor blade. Hover performance of helicopter rotors are especially affected by the strength and location of the tip vortex trailed from the rotor blades. The magnitude of the local induced velocity variation is a strong function of the axial distance of the passing tip vortex beneath the rotor blade. Various rotor blade geometric arrangements, along with blade tip displacement such as anhedral, are utilized to increase hover performance. Although anhedral relatively improves hover performance, anhedral may have negative performance tradeoffs in forward flight performance. Moreover, the degree of anhedral is generally limited by structural considerations associated with the increased out-of-plane mass distribution of the anhedral rotor blade tip. Such increased out-of-plane mass distribution increases the stress in the blade structure and may negatively affect overall rotor system longevity.

Accordingly, it is desirable to provide a rotor blade tip configuration that reduces the tip vortex influence on a following rotor blade.

SUMMARY OF THE INVENTION

The rotary-wing according to the present invention provides a multi-hedral tip section which utilizes a distribution of anhedral and/or dihedral angles that cause a tip vortex to be axially displaced relative to the region on the following blade strongly impacted by the tip vortex such that the tip vortex passes the following blade to improve hover performance while maintaining acceptable forward flight performance.

The present invention therefore provides a rotor blade tip configuration that reduces the tip vortex influence on a following rotor blade.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows:

FIG. 1 is a general perspective view an exemplary rotary wing aircraft embodiment for use with the present invention;

FIG. 2 is a plan view of a rotor blade for use with the present invention;

FIG. 3 is an expanded view of a multiple of rotor blades illustrating a tip vortex;

FIG. 4A is an expanded plan view of a rotor blade with a multi-hedral tip section;

FIG. 4B is an expanded rear view of the rotor blade multi-hedral tip section of FIG. 4A;

FIG. 5 is an expanded view of a propeller with a multi-hedral tip section of the present invention;

FIG. 6 is an expanded rear view of another multi-hedral tip section;

FIG. 7 is an expanded rear view of another multi-hedral tip section;

FIG. 8 is an expanded rear view of another multi-hedral tip section;

FIG. 9 is an expanded rear view of another multi-hedral tip section;

FIG. 10 is an expanded rear view of another multi-hedral tip section; and

FIG. 11 is an expanded rear view of another multi-hedral tip section.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 schematically illustrates a rotary-wing aircraft 10 having a main rotor assembly 12. The aircraft 10 includes an airframe 14 having an extending tail 16 which mounts an anti-torque rotor 18. Although a particular helicopter configuration is illustrated in the disclosed embodiment, other machines such as turbo-props, tilt-rotor and tilt-wing aircraft will also benefit from the present invention.

Referring to FIG. 2, a rotor blade 20 (only one illustrated) of the rotor assembly 12 includes an inboard section 22, an intermediate section 24, and an outboard section 26. The inboard, intermediate, and outboard sections 22, 24, 26 define the span of the main rotor blade 20. The rotor blade sections 22, 24, 26 define a blade radius R between the axis of rotation A and a distal end 30 of a blade tip section 28.

The blade root portion 22 is attached to the rotor assembly 12 for rotating the rotor blade 20 about the axis of rotation A and for pitching about a longitudinal feathering axis P. The rotor blade 20 defines a leading edge 22a and a trailing edge 22b, which are generally parallel to each other. The distance between the leading edge 22a and the trailing edge 22b defines a main element chord length Cm.

The outboard section 26 includes the blade tip section 28 which defines the distal end 30 of the rotor blade 20. The blade tip section 28 may include variations in chord, pitch, taper, sweep, and airfoil distributions. Although a rotor blade is disclosed in the illustrated embodiment, other aerodynamic members such as aircraft and marine propellers, fans, tilt-rotors, wind turbines, and other rotary-wing devices will benefit from the present invention.

Referring to FIG. 3, a preceding rotor blade 20p, operating in hover, generates a tip vortex V which passes near a section of a following rotor blade 20f. The tip vortex V spirals inward of an outer diameter of the following rotor blade 20f. That is, the tip vortex V has a helical quality such that the following blade 20f receives passage of the tip vortex V along the tip section longitudinally inboard of the rotor blade distal end 30f. The passage of the tip vortex V from the preceding blade 20p induces a local velocity variation along the span of the following rotor blade 20f to generate a general downwash Vd inboard and upwash Vu outboard of the radial passage location 32, superimposed with global wake induced velocity. This induced velocity field and rotational speed, combined with the blade geometric pitch distribution determines the angle-of-attack distribution that the blade experiences in hover.

Referring to FIGS. 4A and 4B, the multi-hedral blade tip section 28 includes a first segment 34 which defines a first axis T1, a second segment 36 which defines a second axis T2 and a third segment 38 which defines a third axis T3. The first axis T1 is preferably defined parallel to the blade feathering axis P. It should be understood that the Figures are illustrated without a twist within the rotor blade 20 for sake of clarity and that the first axis T1 is generally parallel to the longitudinal direction of the rotor blade 20, the leading edge 22a and/or the trailing edge 22b. The second axis T2 is transverse the first axis T1 to define a dihedral between the first segment 34 and the second segment 36 at, preferably a 90 percent of radius blade station. The third axis T3 is transverse the second axis T2 to define an anhedral between the second segment 36 and the third segment 38 at, preferably a 94 percent of radius blade station. Although a dihedral of 10 degrees and an anhedral of 20 degrees are disclosed in the illustrated embodiment, other angles will likewise benefit herefrom.

The intersection of the third axis T3 and the second axis T2 is most preferably the radial passage location 32 for the tip vortex V (FIG. 3). That is, the radial passage location 32 is axial displaced raised relative the tip vortex V such that tip vortex V is passes below the intersection of the third axis T3 and the second axis T2 and has minimal effect upon the rotor blade tip section 28.

Generally, the multi-hedral tip section 28 utilizes a distribution of anhedral and/or dihedral angles that cause the tip vortex to be at a lower axial position, relative to the region on the following blade strongly impacted by the tip vortex V such that the tip vortex V passes beneath the following blade to improve hover performance while maintaining an acceptable forward flight performance.

It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. In particular, the usage of the term “below” as in the tip vortex V passes “below” the intersection of the third axis T3 and the second axis T2 is to be broadly construed as below relative the lift and/or thrust generated by the blade, e.g., for a propeller below would indicate that the tip vortex V passes behind the propeller blade 35 behind the radial passage location 32p, such that the tip vortex V is axially displaced opposite the thrust direction which is generated by the rotary member (FIG. 5), in in the direction of the induced flow.

Referring to FIGS. 6-11, several schematics of various multi-hedral tip sections 28a-28f are illustrated. It should be understood that other tip sections which include multi-hedral sections will also benefit from the present invention.

FIG. 6 illustrates a multi-hedral tip section 28a which includes a first segment 34a which defines a first axis T1a, a second segment 36a which defines a second axis T2a and a third segment 38a which defines a third axis T3a. The second axis T2a is transverse the first axis T1a to define an anhedral between the first segment 34a and the second segment 36a. The third axis T3a is transverse the second axis T2a to define a second anhedral between the second segment 36a and the third segment 38a.

FIG. 7 illustrates a multi-hedral tip section 28b as more fully described with regard to FIG. 4A and 4B.

FIG. 8 illustrates a multi-hedral tip section 28c, which essentially combines FIG. 6 and 7 to include four sections.

FIG. 9 illustrates a multi-hedral tip section 28d, which includes a segment 40d, which is generally parallel to the first segment 34d.

FIG. 10 illustrates a multi-hedral tip section 28e, includes five segments to form a segmented arc.

FIG. 11 illustrates a multi-hedral tip section 28f, which is smooth. That is, the multi-hedral tip section 28f includes an infinite number of segments to form the smooth arcuate tip section.

Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention.

The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The preferred embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.

Claims

1. A rotary aerodynamic member comprising: a first segment which defines a first axis; a second segment which defines a second axis transverse said first axis; and a third segment which defines a third axis transverse said second axis.

2. The rotary aerodynamic member as recited in claim 1, wherein said first axis is defined parallel to a leading edge of said first segment.

3. The rotary aerodynamic member as recited in claim 1, wherein said first axis is parallel to a trailing edge of said first segment.

4. The rotary aerodynamic member as recited in claim 1, wherein said first axis is parallel to a feathering axis of said first segment.

5. The rotary aerodynamic member as recited in claim 1, wherein said second axis extends at least partially above said first axis.

6. The rotary aerodynamic member as recited in claim 1, wherein said second axis extends toward a first side of said first axis and said third axis extends toward a second side of said first axis.

7. The rotary aerodynamic member as recited in claim 1, wherein said second axis extends toward a first side of said first axis and said third axis extends toward said first side of said first axis.

8. The rotary aerodynamic member as recited in claim 1, wherein said second segment is a dihedral segment and said third segment is an anhedral segment.

9. The rotary aerodynamic member as recited in claim 1, wherein said second segment and said third segment form a rotor blade tip section.

10. The rotary aerodynamic member as recited in claim 1, wherein said second segment and said third segment form a propeller tip section.

11. The rotary aerodynamic member as recited in claim 1, further comprising a fourth segment generally parallel to said first segment.

12. The rotary aerodynamic member as recited in claim 1, wherein said second segment and said third segment are two of an infinite number of sections.

13. A rotor blade assembly comprising: a first rotor blade segment which defines a first axis; a second segment which defines a dihedral relative said first axis; and a third segment which defines an anhedral relative said first axis.

14. The rotor blade assembly as recited in claim 13, wherein said first rotor blade segment comprises a rotor blade root section.

15. The rotor blade assembly as recited in claim 13, wherein said third rotor blade segment comprises a distal end of a rotor blade tip section.

16. The rotor blade assembly as recited in claim 13, further comprising a fourth rotor blade segment between said second rotor blade segment and said third rotor blade segment.

17. A method of minimizing an influence of a tip vortex formed by a preceding rotary-wing upon a following rotary-wing comprising the step of: (1) axially displacing a multiple of segments of the preceding and receding rotary-wings relative a first axis to axially displace a tip vortex formed by the preceding rotary-wing relative the following rotary-wing.

18. A method as recited in claim 17, wherein step (1) further comprises the step of: displacing at least one of the multiple of segments above the first axis and at least one of the multiple of segments below the first axis.