Controllable variable twist rotor blade assembly

A controllable variable twist rotor blade assembly wherein the twist along the span of the blade is continuously variable and may be controlled either manually by the pilot or automatically according to data input from flight sensors monitoring flight parameter data. The continuously variable twist of the blade is controlled by an out-of-plane weight assembly having a weight member which is selectively movable between a first position and a second position relative to the elastic axis of the blade so as to optimize the twist of the blade during various flight conditions.

BACKGROUND OF THE INVENTION 
1. Field of Invention 
The present invention relates generally to blade assemblies for rotary wing 
aircraft, and more particularly, but not by way of limitation, to an 
improved blade assembly for rotary wing aircraft wherein the blade may be 
controllably twisted to various positions in flight to optimize the angle 
of attack along the span of the blade for various flight conditions. 
2. Description of Related Art 
The lift developed by a conventional aircraft wing depends primarily on two 
factors: the angle of attack of the wing and the velocity of the air in 
relation to the wing. In the case of a rotary wing aircraft such as a 
helicopter, the air velocity in relation to the wing is produced by a 
combination of aircraft motion and the rotation of rotor blades. 
The air velocity over a particular section of the blade, referred to herein 
as a blade element, is a function of the distance of the blade element 
from the rotational axis of the rotor. In other words, air velocity is 
greater at the outer end of the blade than it is near the rotational axis 
of the rotor since the speed of the outer end of the blade is greater than 
the speed of the blade near the rotational axis of the rotor. As a result, 
each blade of a rotor assembly may be formed so that it is twisted about 
its longitudinal axis in order to maintain a favorable angle of attack 
along the span of the blade. 
Helicopter and tilt-rotor blades are supported in bearings that allow the 
blades to be rotated about their longitudinal axis. This rotation allows 
the pilot to make a gross change in blade pitch, either collectively or 
cyclically, in order to control the total lift and horizontal flight 
direction of the aircraft. This rotation of the entire blade does not, 
however, allow each blade to be variably twisted so as to set each blade 
at an optimum pitch at each blade element along the span of each blade for 
a specific flight condition. 
The optimum performance of a rotor blade assembly occurs when the angle of 
attack of each blade element along the span of each blade is related to 
the angle of other blades in such a way as to provide a consistent 
inflow/outflow velocity across the rotational plane of the blades. This 
requires each blade element along the span of each blade of the rotary 
blade assembly to have the proper angle of attack in relation to the air 
velocity at that particular blade element. For a rigid blade, the twist 
distribution is constant regardless of the flight conditions. Thus, the 
twist can be optimized for only one flight condition. For example, if 
hover performance is most critical, blade twist is optimized for airflow 
conditions that prevail during hover, with the consequence of degraded 
performance in all other flight conditions. Similarly, optimization of 
blade twist for high speed cruise degrades performance in all other flight 
conditions, including hover. 
In addition, the ride quality of a helicopter with a fixed blade pitch 
distribution is degraded by the non-optimum angle of attack of the 
individual blade elements from root to tip, and by the inability of 
relatively more rigid blade systems to respond to gusts or other 
turbulence through which the helicopter may be flying. If the blade angles 
can respond to these varying flight conditions in such a way as to 
minimize short-term fluctuations in spanwise lift, the ride quality will 
be improved throughout the range of the flight conditions. 
As an alternative to a blade having a fixed or constant twist, a blade 
which can be variably twisted about the longitudinal axis of the blade can 
provide optimum blade pitch at each blade element along the span of the 
blade for different flight conditions. However, in order for the blades to 
have the potential for variable twist, they must be designed with a 
controlled amount of torsional flexibility. Unfortunately, torsional 
flexibility brings with it the disadvantage of making the blades more 
susceptible to torsional flutter, a dynamic interaction between 
aerodynamic, flexural and mass properties of the blades. Torsional flutter 
can be described as a self-sustaining torsional vibration of the blades in 
which the blade pitch varies above and below acceptable limits at 
relatively high frequency and in such a way as to decrease the aerodynamic 
performance of the blade and impose high structural loads on the blade. 
To this end, a need has long existed for an improved rotor blade assembly 
which allows a torsionally flexible blade to be controlled in flight to 
optimize flight performance while also diminishing the likelihood of 
flutter. It is to such an improved rotor blade assembly that the present 
invention is directed.

DETAILED DESCRIPTION 
FIG. 1 shows a rotor assembly constructed in accordance with the present 
invention and generally designated by the reference numeral 10. The rotor 
assembly 10 is utilized on a rotary wing aircraft, such as a helicopter or 
tilt-rotor aircraft, and includes a hub 12 affixed to a drive shaft 14 
which is rotated by a prime mover (not shown). A pair of rotor blade 
assemblies 16 and 18 are connected to the hub 12 via a pair of articulated 
hinges 20 and 22, respectively. The articulated hinges 20 and 22 function 
in the manner of universal joints thereby allowing for leading, lagging 
and flapping of the blade assemblies 16 and 18 in a manner well known in 
the art. The blade assemblies 16 and 18 are rotated about the hub 12 in 
the direction indicated by the arrow 24. 
It will be appreciated that a helicopter rotor assembly employs a plurality 
of rotor blade assemblies consisting of two, three or more rotor blade 
assemblies, and that each of the rotor blade assemblies 16 and 18 is 
identical in construction. Therefore, although a pair of blade assemblies 
16 and 18 are shown, only one blade assembly 16 will be referred to 
hereinafter in describing and explaining the features and functions of 
both blade assemblies 16 and 18. 
The blade assembly 16 includes a torsional flexible blade 25 having a 
leading edge 26, a trailing edge 28, a radially inner fixed end 30, and a 
radially outer free end 32. An elastic axis 39 of the blade 25 extends 
through the torsionally flexible blade 25. The elastic axis 39 of the 
blade 25 is the radial axis about which the blade 25 will twist when a 
purely torsional force is applied to the blade 25. 
During rotation of the rotor assembly 10, the pitch of the blade assembly 
16 is cyclically controlled in order to control the vertical lift and to 
achieve horizontal flight. This cyclical control of the pitch of the blade 
assembly 16, as well as the use of the articulated hinge 20 as briefly 
mentioned above, is well known in the art. Thus, no further description of 
their various components or their operation is believed necessary in order 
to enable one skilled in the art to understand the rotor assembly 10 of 
the present invention and the relationship of the rotor assembly 10 to a 
rotary wing aircraft. 
Referring now to FIGS. 1-5, the blade assembly 16 is provided with an 
out-of-plane weight assembly 38 positioned ahead of the leading edge 26 
for controlling the torsional flex of the blade 25 in a manner which will 
be discussed in detail hereinafter. The out-of-plane weight assembly 38 
includes a rigid body member 40 having a first end 41 and a second end 42. 
The first end 41 is pivotally mounted to the blade 25 along the leading 
edge 26 thereof and a weight member 43 is connected to the second end 42 
of the body member 40 so as to be positioned ahead of the leading edge 26. 
It is desirable that the out-of-plane weight assembly 38 be located in 
close proximity to the radially outer free end 32 of the blade 25 so that 
the angle of attack of the blade 25 is varied along the entire length of 
the torsionally flexible blade 25. The out-of-plane weight assembly 38 is 
selectively movable between a first position where the weight member 43 is 
positioned above a plane of rotation 44 of the blade 25 (FIG. 2) and a 
second position where the weight member 43 is positioned below the plane 
of rotation 44 of the blade 25 (FIG. 4). It will be understood that the 
out-of-plane weight assembly 38 is selectively movable between the first 
and second positions and to positions therebetween. 
As will be described in greater detail hereinafter, the out-of-plane weight 
assembly 38 allows a twisting moment to be applied to the radially outer 
free end 32 of the blade 25 when the blade 25 is in the rotating 
condition. This twisting moment causes a spanwise twist in the blade 25. 
The magnitude of the twisting moment and, therefore, the degree of twist 
imparted to the blade 25 is dependent on the length of the body member 40 
of the out-of-plane weight assembly 38, the weight of the weight member 43 
of the out-of-plane weight assembly 38, the angle at which the weight 
member 43 of the out-of-plane weight assembly 38 is disposed from the 
blade 25, and the centrifigal force acting on the out-of-plane weight 
assembly 38 because of the blade rotation about the drive shaft 14. 
Before explaining the theory and manner of operation of the blade assembly 
16, additional reference must be made to FIG. 1 to define certain terms 
and relationships necessary to understand the orientation of the rotor 
assembly 10 and the direction and magnitude of certain forces acting upon 
the blade assemblies 16 and 18 thereof. 
The rotation of the blade 25 produces a wind velocity vector 45 and air 
passing through the plane of rotation 44 produces an inflow velocity 
vector 46. These vectors are added to produce a resultant relative wind 
vector 48. The relative wind vector 48, as well as an inflow angle 50 
thereof, varies along the length of the blade 25; that is, the relative 
wind vector 48 increases from the radially inner fixed end 30 to the 
radially outer free end 32 and the inflow angle 50 decreases from the 
radially inner fixed end 30 to the radially outer free end 32. As stated 
above, it is this spanwise difference in the relative wind vector 48 that 
creates the need for a blade 25 having a variable twist. By employing the 
system of the present invention, the relative wind vector 48 is more 
uniformly maintained across the span of the blade 25. By controlling the 
twist of the blade 25, the blade 25 can be optimized for varying flight 
conditions. 
The term "continuously varied" as used herein is to be understood to mean 
that the variation in pitch differs along the span of the blade 25 from a 
first variation adjacent the out-of-plane weight assembly 38 to a second 
variation at the radially inner fixed end 30, the first variation being 
greater than the second variation. In other words, with the entire blade 
25 being torsionally flexible, the pitch is continuously varied along the 
entire span of the blade 25 between a zero pitch variation at the radially 
inner fixed end 30 and a maximum pitch variation adjacent the out-of-plane 
weight assembly 38. 
Referring more specifically to FIG. 2, the blade 25, the body member 40 and 
the weight member 43 of the out-of-plane weight assembly 38 are depicted 
with the blade 25 in a non-rotating condition with the out-of-plane weight 
assembly 38 positioned above the plane of rotation 44 of the blade 25. The 
out-of-plane weight assembly 38 is attached to the leading edge 25 of the 
blade 25 so as to be located forward of the elastic axis 39 of the blade 
25 to reduce the likelihood of flutter. However, it should be understood 
that the out-of-plane weight assembly 38 can be attached to the trailing 
edge 28 and thus located rearward of the elastic axis 39 of the blade 25 
without departing from the scope and spirit of the present invention. 
FIG. 3 is a cross sectional view of the blade 25 showing the twisting 
effect on the blade 25 when the blade 25 is in the rotating condition and 
when the weight member 43 of the out-of-plane weight assembly 38 is 
initially positioned above the plane of rotation 44 of the blade 25. A 
component, indicated by arrow 58, of the centrifugal force acting on the 
weight member 43 due to the rotation of the blade 25 tends to move the 
weight member 43 toward the plane of rotation 44 of the blade 25. As the 
weight member 43 of the out-of-plane weight assembly 38 approaches the 
plane of rotation 44 of the blade 25, the torsionally flexible blade 25 is 
twisted about the elastic axis 39 thereby resulting in a negative angle of 
twist 60 about the elastic axis 39. The blade 25 is therefore twisted to 
the position indicated in FIG. 3, wherein the pitch is continuously varied 
between a zero pitch variation at the radially inner fixed end 30 and a 
maximum pitch variation adjacent the out-of-plane weight assembly 38. 
FIG. 4 shows a cross-section of the blade 25 wherein the blade 25, the body 
member 40 and weight member 43 of the out-of-plane weight assembly 38 are 
depicted as if the blade 25 is in a non-rotating condition with the 
out-of-plane weight assembly 38 positioned below the plane of rotation 44. 
As will be discussed hereinafter, positioning the weight member 43 in this 
manner will cause the blade 25 (when the blade 25 is in the rotating 
condition) to twist in the opposite direction than the direction the blade 
25 twisted when the weight member 43 was positioned above the plane of 
rotation 44 of the blade 25. 
FIG. 5 is a cross sectional view of the blade 25 showing the twisting 
effect on the blade 25 when the blade 25 is in the rotating condition and 
when the weight member 43 of the out-of-plane weight assembly 38 is 
disposed below the plane of rotation 44 of the blade 25. A component, 
indicated by arrow 59, of the centrifugal force acting on the weight 
member 43 due to the rotation of the blade 25 tends to move the weight 
member 43 toward the plane of rotation 44 of the blade 25. Since the 
weight member 43 is disposed below the plane of rotation 44 of the blade 
25, this component 59 of centrifugal force acts in the opposite direction 
as the component 58 of centrifugal force shown in FIG. 4. As the weight 
member 43 of the out-of-plane weight assembly 38 approaches the plane of 
rotation 44 of the blade 25, the torsionally flexible blade 25 is twisted 
about the elastic axis 39 of the blade 25 thereby resulting in a positive 
angle of twist 62 about the elastic axis 39. The blade 25 is therefore 
twisted to the position indicated in FIG. 5 when the blade 25 is placed in 
the rotating condition, wherein the pitch is continuously varied between a 
zero pitch variation at the radially inner fix end 30 and a maximum pitch 
variation adjacent the out-of-plane weight assembly 38. 
The theory of operation of the pitch adjusting feature of the out-of-plane 
weight assembly 38 is explained in greater detail in U.S. Pat. No. 
4,291,235, issued to Karl H. Bergey on Sep. 22, 1981, the disclosure of 
which is hereby incorporated by reference. Also, while the blade assembly 
16 has been described as having a single out-of-plane weight assembly, it 
will be understood that a plurality of out-of-plane weight assemblies can 
be positioned along the span of the blade angle thereby tailoring the 
blade distribution along the blade span in ways that are less linear. Such 
arrangements of plural out-of-plane weight assemblies may provide 
improvements in both aerodynamic performance and ride quality over a wider 
variety of flight conditions. 
Referring now to FIG. 6, a schematic of a control mechanism 64 for 
selectively controlling the position of the weight member 43 of the 
out-of-plane weight assembly 38 is illustrated. A powered servo actuator 
65 is operably connected to a control unit 66 whereby the pilot may 
control the placement of the weight member 43 of the out-of-plane weight 
assembly 38 from the cockpit. The servo actuator 65 and control unit 66 
are elements well known in the art, therefore no further description of 
their features will be made herein. The servo actuator 65 is connected to 
a shaft 68 which is connected to the first end 41 of the body member 40 of 
the out-of-plane weight assembly 38. Upon receipt of appropriate signals 
from the control unit 66, the servo actuator 65 rotates the shaft 68 which 
moves the out-of-plane weight assembly 38 to the desired position. 
FIG. 7 shows another embodiment of a control mechanism 70. A control unit 
72, to which a servo actuator 74 is operably connected, comprises a flight 
control system which receives signals from a multiplicity of sensors that 
continuously monitor flight parameters, such as rotor RPM, vehicle 
velocity through the air, angle of attack, blade loads, and blade 
dynamics. The servo actuator 74 adjusts the position of the out-of-plane 
weight assembly 38 (and thereby the twist of the blade 25) in accordance 
with rotary wing aircraft control laws to optimize the twist of the blade 
25 for the existing flight conditions. In this type of system, the servo 
actuator 74 is programmed to move the weight member 43 of the out-of-plane 
weight assembly 38 according to a predetermined schedule which is 
dependent upon the status of the monitored flight conditions. 
Changes may be made in the embodiment of the invention described herein, or 
in parts or elements of the embodiment described herein without departing 
from the spirit and/or scope of the invention as defined in the following 
claims.