Note: Descriptions are shown in the official language in which they were submitted.
ROTOR ASSEMBLY WITH HIGH LOCK-NUMBER BLADES
TECHNICAL FIELD
This patent application relates to aircraft rotor assemblies.
BACKGROUND
The Lock number is a dimensionless parameter for aircraft rotor blades, and
the equation is
7= pacR4
where 7 = Lock number
p = air density
a= slope of the 2-D airfoil lift curve
c= chord length
R= rotor radius
lb = flapping moment of inertia.
The Lock number represents the ratio of aerodynamic forces, which act to lift
the blade, to inertial
forces, which act to maintain the blade in the plane of rotation. Typical
rotorcraft blades have a
Lock number of between 3 and 12.
The inertial forces are based on the mass of each blade, so a larger blade
tends to have a lower
Lock number. For example, a two-blade helicopter rotor typically has blades
with high inertia, and
this is due to the size of each blade required to achieve the desired amount
of lift. However, a rotor
can provide the same or more lift by using a larger number of smaller and
lighter (higher Lock
number) blades. This reduces the mass and total inertia of the rotor and
reduces the loads that must
be reacted by the rotor hub, allowing for a lighter hub. Another advantage to
reducing rotor mass
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CA 2995295 2018-02-14
and inertia is that the jump-takeoff load, which is used to design the roof
structure, increases with
rotor inertia. In addition, reducing the mass of the rotor reduces the load it
applies to the fuselage in
a crash. Therefore, reducing the mass of the rotor may allow for a lighter
fuselage design, with
fuselage mass perhaps being reduced by twice the amount removed from the
rotor.
Another engineering consideration is that the combined inertia of the blades
be high enough to
allow for autorotation after engine failure, so single-engine aircraft
typically have high-inertia
rotors, whereas multi-engine aircraft can use rotors with less inertia. One
way to achieve higher
inertia is to add tip weights to the blades, but another way is to add blades
to the rotor. As
described above, adding narrower, lighter blades with a higher Lock number can
allow for an
aircraft with reduced weight in both the rotor system and the fuselage.
Using an increased number of narrower blades has other advantages. One
advantage is that
reducing the chord width reduces material cost for each blade, which can
significantly reduce the
price of a shipset of blades. Also, the rotor is quieter during operation due
to the reduced blade
noise, which tends to vary with chord width, and to the increased number of
blade passages, which
coalesce into a higher frequency and less offensive sound.
SUMMARY
In one aspect, there is provided an aircraft rotor assembly, comprising: a
central hub; a plurality of
rotor blades coupled to the hub for rotation with the hub about an axis, each
blade having a Lock
number of approximately 5 or greater; and a lead-lag pivot for each blade
formed by a flexure
coupling the associated blade to the hub, each pivot being a radial distance
from the axis and
allowing for in-plane lead-lag motion of the associated blade relative to the
hub, each pivot
allowing for in-plane motion from a neutral position of at least 1 degree in
each of the lead and lag
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directions; wherein elastic deformation of the flexure produces a biasing
force for biasing the
associated blade toward the neutral position, wherein the biasing force is
selected to achieve a first
in-plane frequency of greater than 1/rev for each blade.
In another aspect, there is provided an aircraft rotor assembly, comprising: a
central hub; a plurality
of rotor blades coupled to the hub for rotation with the hub about an axis,
each blade having a Lock
number of approximately 5 or greater; a lead-lag pivot for each blade formed
by a flexure coupling
the associated blade to the hub, each pivot being a radial distance from the
axis and allowing for in-
plane lead-lag motion of the associated blade relative to the hub, each pivot
allowing for in-plane
motion from a neutral position of at least 1 degree in each of the lead and
lag directions; and a
discrete flap hinge for each blade; wherein elastic deformation of the flexure
produces a biasing
force for biasing the associated blade toward the neutral position, wherein
the biasing force is
selected to achieve a first in-plane frequency of greater than 1/rev for each
blade.
In a further aspect, there is provided an aircraft rotor assembly, comprising:
a central hub; a
plurality of rotor blades coupled to the hub for rotation with the hub about
an axis, each blade
having a Lock number of approximately 5 or greater; a lead-lag pivot for each
blade formed by a
flexure coupling the associated blade to the hub, each pivot being a radial
distance from the axis
and allowing for in-plane lead-lag motion of the associated blade relative to
the hub, each pivot
allowing for in-plane motion from a neutral position of at least 1 degree in
each of the lead and lag
directions; and a gimbal configured to allow pivoting of the hub relative to
the axis about flap axes
perpendicular to the axis; wherein elastic deformation of the flexure produces
a biasing force for
biasing the associated blade toward the neutral position, wherein the biasing
force is selected to
achieve a first in-plane frequency of greater than 1/rev for each blade.
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BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is an oblique view of an aircraft comprising a rotor assembly
according to this disclosure.
Figure 2 is an oblique view of a portion of the aircraft of Figure 1 and
showing the rotor assembly.
Figure 3 is a top view of a portion of the aircraft of Figure 1 and showing
the rotor assembly.
Figure 4 is an oblique isolation view of the rotor assembly of Figure 1, some
components being
removed for ease of viewing.
Figure 5 is an oblique view of the rotor assembly of Figure 1, some components
being removed for
ease of viewing.
Figure 6 is an oblique view of a portion of the rotor assembly of Figure 1,
some components being
removed for ease of viewing.
Figure 7 is an oblique view of a portion of an aircraft having an alternative
embodiment of a rotor
assembly according to this disclosure.
Figure 8 is an oblique view of the portion of the aircraft of Figure 7 and
showing the alternative
embodiment of the rotor assembly.
Figure 9 is an oblique view of another alternative embodiment of a rotor
assembly according to this
disclosure, some components being removed for ease of viewing.
Figure 10 is an oblique enlarged view of the rotor assembly of Figure 9.
Figure 11 is an oblique view of another alternative embodiment of a rotor
assembly according to
this disclosure, some components being removed for ease of viewing.
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Figure 12 is an enlarged top view of the rotor assembly of Figure 11.
Figure 13 is an oblique view of a tiltrotor aircraft comprising rotor
assemblies according to this
disclosure.
Figure 14 is an oblique isolation view of a rotor assembly for use with the
aircraft of Figure 13,
some components being removed for ease of viewing.
Figure 15 is an oblique exploded view of the rotor assembly of Figure 14.
Figure 16 is an oblique isolation view of another embodiment of a rotor
assembly for use with the
aircraft of Figure 13, some components being removed for ease of viewing.
Figure 17 is an oblique exploded view of the rotor assembly of Figure 16.
DETAILED DESCRIPTION
In the specification, reference may be made to the spatial relationships
between various
components and to the spatial orientation of various aspects of components as
the devices are
depicted in the attached drawings. However, as will be recognized by those
skilled in the art after a
complete reading of this disclosure, the devices, members, apparatuses, etc.
described herein may
be positioned in any desired orientation. Thus, the use of terms such as
"above," "below," "upper,"
"lower," or other like terms to describe a spatial relationship between
various components or to
describe the spatial orientation of aspects of such components should be
understood to describe a
relative relationship between the components or a spatial orientation of
aspects of such
components, respectively, as the device described herein may be oriented in
any desired direction.
CA 2995295 2018-02-14
This disclosure divulges a new concept for rotor assemblies that are stiff-in-
plane and soft out-of-
plane while using high Lock-number blades and with a first in-plane frequency
over 1/rev and
requiring no dampers.
The Lock number of a blade and its first in-plane frequency correlate
directly, so that, all other
things being equal, a higher Lock number produces a higher in-plane frequency.
Two ways to
adjust the first in-plane frequency are to soften the yoke in the chord, or in-
plane, direction or add
tip weight to the blade. Adding tip weight is contrary to the goal of reducing
the mass of the
aircraft, so removing stiffness from the yoke is a preferred way to achieve
the desired frequency.
With stiff and/or light blades, the required yoke stiffness in the chordwise
direction falls to the
point that the system is still a stiff-in-plane rotor but allows the blade to
move relative to the yoke
in in-plane lead and lag directions enough to reduce loads in the rotor.
Because the fundamental loads on a rotor blade are a function of mass, a
reduction in mass results
in a reduction of load, and this new rotor configuration could not have
existed prior to the
development of very stiff and light materials for bade construction. Allowing
a blade to move
relative to the yoke reduces loads, and the loads are inversely proportional
to the amount of
movement. Additional mass savings are achieved by there being no need for lead-
lag dampers, as
the configuration produces an in-plane natural frequency above 1/rev. This
type of rotor design
leads to solutions with more blades than traditional helicopters, but the
loads, weight, noise,
vibration, and cost are reduced when the weight and chord width of the blades
are reduced, as with
a high Lock-number blade.
Typically, the expectation is that being stiff-in-plane means that a rotor is
rigid, with a first in-plane
frequency of >1/rev and lead-lag motion of less than 1 degree in each of the
lead and lag directions
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with no lead-lag hinge. In fact, stiff-in-plane and rigid rotors have 1st in-
plane frequencies that are
similar for blades with a Lock number of approximately 4 or less, but the
frequencies for these
types of rotors begin to diverge when blade Lock numbers are approximately 5
and above. By a
Lock number of 10, they have diverged greatly.
The rotor designs according to this disclosure fall under the definition of a
stiff-in-plane rotor but
allow the blades to move as much as a soft-in-plane rotor, with lead-lag
motion of greater than 1
degree in each direction. Because of the need to keep first in-plane
frequencies on either side of
1/rev, a soft-in-plane rotor requires dampers to keep the frequency below
1/rev, whereas the rotor
designs according to this disclosure require no dampers and have a frequency
above 1/rev. These
rotor designs will typically use blades having a Lock number of approximately
6 to approximately
11, but particular applications may use a higher Lock number. For example, a
large helicopter may
have a high number of skinny blades formed from a high-stiffness material,
such as graphite, and
having a Lock number of, for example, up to 14.
While the rotor designs according to this disclosure have advantages for
helicopter application,
tiltrotor applications can have greater gains when compared to current rotor
designs. There is no
requirement in a tiltrotor to have enough rotor mass for autorotation, so
rotors may be configured
to use blades having a Lock number of, for example, 12 or higher, and the
upper limit on the Lock
number used may only be from the limits of blade construction. Additionally,
tiltrotors have
stability issues in whirl mode, which is an aerodynamic term that includes
rotor mass, and a higher
rotor mass leads to an undesirable increase in whirl. Therefore, using reduced-
mass rotors, such as
those according to this disclosure, on a tiltrotor will result in less whirl.
Another advantage is that
the reduced mass of high Lock-number blades further increases the advantage
that aerodynamic
forces of the blades have over the inertial mass forces. A tiltrotor rotor
does not normally cone
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CA 2995295 2018-02-14
much during flight in airplane mode due to the reduced thrust force when
compared to helicopter
mode, but the reduced mass of the blades may allow for a combination of
sensors and improved
swashplate control to cause coning of the rotors during airplane mode to
increase the distance
between the blades and the wing.
In a specific tiltrotor example, the first in-plane frequency for the three-
blade rotors on a
Bell/Boeing V-22 tiltrotor is 1.23/rev, and it may be desirable to increase
the frequency to 1.5/rev.
A four-blade rotor design according this disclosure and used on a V-22 will
have the desired
increase in frequency but will also reduce mass of the rotors, and the
reduction may be as much as
10001bs per rotor. This leads to reduction in the mass of other aircraft
components, such as the
fuselage, landing gear, etc., and this multiplier effect leads to a
significantly reduced overall mass
of the aircraft.
Figure 1 illustrates an aircraft 11 comprising a main rotor assembly according
to this disclosure.
Aircraft 11 comprises a fuselage 13 and a rotor assembly 15 with a plurality
of blades 17. Each
blade 17 has a Lock number of approximately 5 or greater. Rotor assembly 15 is
driven in rotation
about mast axis 19 by torque provided by an engine housed within fuselage 13.
Though aircraft 11
is shown as a helicopter having a single main rotor, rotor assembly 15 can
alternatively be used on
other types of aircraft, such as, but not limited to, helicopters having more
than one main rotor or
on tiltrotor aircraft. Also, rotor assembly 15 is shown as a main rotor for
providing vertical lift and
requiring collective and cyclic control, though rotor assembly 15 may
alternatively be configured
to provide longitudinal or lateral thrust, such as in a helicopter tail rotor
or airplane propeller.
Figures 2 through 6 illustrate rotor assembly 15, various components being
removed for ease of
viewing. A central yoke 21 is coupled to a mast 23 (shown in Figure 4) for
rotation with mast 23
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about mast axis 19. Yoke 21 has a honeycomb configuration in the embodiment
shown, though in
other embodiments yoke 21 may have another configuration, such as a central
portion with radially
extending arms. Yoke 21 is preferably formed from a composite material, though
yoke 21 may be
formed from any appropriate material. In the embodiment shown, yoke is
configured for use with
five rotor blades, though yoke 21 may be configured for use with any
appropriate number of
blades.
Yoke 21 has a bearing pocket 25 for each blade 17, each pocket 25 carrying a
spherical bearing 27.
Each bearing 27 is spaced a radial distance from axis 19 and transfers
centrifugal force from the
associated blade 17 to yoke 21. Each bearing 27 forms a lead-lag pivot to
allow for limited rotation
of the associated blade 17 relative to yoke 21 in in-plane lead-lag
directions, and bearing 27 also
allows for limited rotation in out-of-plane flapping directions and limited
rotation about a pitch
axis 29. While each blade 17 can lead-lag about the associated bearing 27,
during operation the
centrifugal force tends to force each blade 17 toward a centered, neutral
position, from which each
blade 17 can lead, by rotating forward (in the direction of rotation about
mast axis 19) in-plane
relative to yoke 21, or lag, by rotating rearward in-plane relative to yoke
21.
A blade grip 31 couples each blade 17 to the associated bearing 27, each grip
31 being shown as an
elongated U-shaped structure, comprising an upper plate 33, a lower plate 35,
and a curved inner
portion 37 connecting plates 33, 35. Each grip 31 is connected to an inner end
of a blade 17 with
fasteners 39, thereby allowing loads from each blade 17 to be transferred
through grip 31 and inner
bearing 27 to yoke 21. A pitch horn 41 is mounted to each grip 31 by an
integral plate 43, allowing
for actuation by a flight control system of a pitch link 45 coupled to pitch
horn 41 for causing
rotation of grip 31 and blade 17 together about pitch axis 29 for cyclic and
collective control of
blades 17.
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To provide for a stiff-in-plane configuration, each blade 17 and/or grip 31 is
coupled to an outer
bearing, such as outboard shear bearing 47, supported by an in-plane flexure
assembly. The
following describes the flexure assembly for one blade 17, but, in the
embodiment shown, each
blade 17 has its own flexure assembly.
Each end of an elongate outboard bearing support 49 is coupled to yoke 21 by
upper and lower
brackets 51 using spherical bearings 53, allowing for bearing support 49 to
rotate only in out-of-
plane directions (away from or toward fuselage 13). This provides a discrete
flap hinge, allowing
for a limited amount of flapping and coning motion of the associated blade 17.
Each bracket 51 is
U-shaped and comprises a base portion 55, which is fastened rigidly to an
outer portion of yoke 21,
and two radially extending arms 57, which extend beyond the periphery of yoke
21. As visible in
the figures, the U-shape of brackets 51 allows pitch link 45 to be located and
operate between arms
57. The inner portion of each bearing 53 is fastened within a clevis formed by
corresponding arms
57 of two brackets 51, and the outer portion of each bearing 53 is installed
in a bearing pocket 58
of bearing support 49, such that bearing support 49 can rotate as described
relative to brackets 51
and yoke 21 for flapping and coning motions of blade 17.
A lead-lag flexure 59 is rigidly fastened at the inner end to a central boss
61 of bearing support 49,
and shear bearing 47 is rigidly fastened by clevis 62 to the outer end of
flexure 59. Shear bearing
47 is coupled to grip 31 (coupling not shown) for providing shear support of
blade 17 and
cooperates with inner bearing 27 to define pitch axis 29. Flexure 59 is
preferably formed from a
composite material, though flexure 59 can alternatively be formed from another
appropriate
material or combination of materials. For example, flexure 59 may be formed
solely from
fiberglass or a similar composite, or flexure from a combination of materials,
such as with a
laminated construction.
CA 2995295 2018-02-14
In the embodiment shown, flexure 59 is formed as a beam having a generally
rectangular cross-
section, with flexure 59 oriented to have a bending stiffness greater in the
out-of-plane directions,
shown by arrows 63, 65, than a bending stiffness in the in-plane directions,
indicated by arrows 67,
69. This means that flexure 59 is stiff to out-of-plane flapping motions of
the associated blade 17,
and the flapping motion occurs with movement relative to yoke 21 of bearing
support 49 at
bearings 53. However, flexure 59 acts as a spring by bending through elastic
deformation to allow
for a selected amount of rotation of the associated blade 17 at least 1 degree
in each of the lead and
lag directions, the bending of flexure 59 producing a biasing force opposing
the lead-lag motions
and biasing the blade toward the neutral position. This allows for selection
of the in-plane stiffness
of flexure 59 to "tune" the first in-plane frequency to be above 1/rev, and no
dampers are required
to achieve the desired frequency. This configuration is a new class of rotor
assembly, which may
be termed a "compliant stiff-in-plane" rotor.
Figures 7 through 12 illustrate alternative embodiments of a rotor assembly
according to this
disclosure. Like rotor assembly 15, as described above, these additional
embodiments, and
variations thereof, use blades with a high Lock number of approximately 5 or
greater to achieve a
reduced-mass rotor. The configurations each have lead-lag pivots radially
spaced from the mast
axis and allowing for in-plane lead-lag motion of the blades of at least 1
degree in each direction
from a neutral position, components for producing a biasing force through
elastic deformation that
opposes lead-lag motion of the blades and that bias the blades toward the
neutral position, and a
first in-plane frequency above 1/rev without the need for dampers. As with
rotor assembly 15,
shown and described above, these alternative embodiments are shown as a single
main rotor
assembly for a helicopter, those these rotor assemblies can alternatively be
used on other types of
aircraft, such as, but not limited to, helicopters having more than one main
rotor or tiltrotors. Also,
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CA 2995295 2018-02-14
these rotor assemblies are shown as a main rotor for providing vertical lift
with collective and
cyclic control, though they may alternatively be configured to provide
longitudinal or lateral thrust,
such as in a helicopter tail rotor or airplane propeller.
Figures 7 and 8 illustrate an alternative embodiment of a rotor assembly
according to this
disclosure, the views having various components removed for ease of viewing.
Aircraft 71
comprises a fuselage 73 and a rotor assembly 75 with a plurality of blades 77.
Rotor assembly 75 is
driven in rotation about mast axis 79 by torque provided by an engine housed
within fuselage 73.
In rotor assembly 75, a central yoke 81 is coupled to a mast 83 for rotation
with mast 83 about
mast axis 79. Yoke 81 has a honeycomb configuration in the embodiment shown,
though in other
embodiments yoke 81 may have another configuration, such as a central portion
with radially
extending arms. Yoke 81 is preferably formed from a composite material, though
yoke 81 may be
formed from any appropriate material. In the embodiment shown, yoke is
configured for use with
five rotor blades, though yoke 81 may be configured for use with any
appropriate number of
blades. Yoke 81 has a bearing pocket 85 for each blade 77, each pocket 85
carrying a spherical
bearing 87, which transfers centrifugal force from the associated blade 77 to
yoke 81. Each bearing
87 allows for limited rotation of the associated blade 77 relative to yoke 81
in in-plane lead and lag
directions, as indicated by arrows 89, 91, respectively, and in out-of-plane
flapping directions, as
indicated by arrows 93, 95. Each bearing 87 also allows for limited rotation
of the associated blade
77 about a pitch axis 97 for changing the pitch of blade 77.
A blade grip 99 couples each blade 77 to the associated bearing 87, each grip
91 being shown as an
elongated U-shaped structure, comprising an upper plate 101, a lower plate
103, and a curved inner
portion 105 connecting plates 101, 103. Each grip 99 is connected to an inner
end of blade 77 with
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fasteners 107, thereby allowing loads from each blade 77 to be transferred
through grip 99 and
bearing 87 to yoke 81. Each bearing 87 forms a lead-lag pivot, allowing for
lead-lag motion of the
associated blade. A pitch horn 109 is installed on each grip 99, allowing for
actuation by a flight
control system of a pitch link 111 coupled to pitch horn 109 for causing
rotation of grip 99 and
blade 77 together about pitch axis 97 for cyclic and collective control of
blades 77. Though not
shown, a droop stop limits droop of each blade 77 and grip 99 assembly toward
fuselage 73 when
rotor is slowly rotating about mast axis 79 or at rest.
Each blade 77 is coupled to each adjacent blade 77 by a spring assembly 113,
each spring
assembly 113 providing a biasing force and cooperating with each adjacent
spring assembly 113 to
bias each associated blade 77 toward a neutral position in lead-lag directions
89, 91. Each spring
assembly 113 comprises a spring 115, a spring perch 117 at each end of spring
115, and a
telescoping stabilizing rod 119 extending between perches 117. Spring 115 may
be formed from
metal, as shown, or spring 115 may be formed from a composite material or a
low-damped
elastomer, and these may require a different configuration for spring assembly
113. A connector,
such as rod end bearing 121, is installed at each end of spring assembly 113.
To provide for coupling of spring assemblies 113 to grips 99, a spring block
123 is rigidly coupled
to each grip 99 with fasteners 124, and each spring block 123 comprises a pair
of shafts 125 sized
for receiving rod end bearings 121. When assembled, each spring assembly 113
can be rotated a
limited amount relative to each spring block 123, allowing for the assemblies
of grips 99 and
blades 77 to move in lead and lag directions relative to each other and to
yoke 81. Also, the biasing
force of each spring assembly 113 is transferred to each grip through a spring
perch 117 and
associated bearing 121 and into spring block 123 for biasing grips 99 to
control relative motion
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between grips 99 and their associated blades 77. Selection of the biasing
force of springs 115
allows for tuning of the in-plane frequencies without the need for dampers.
The configuration of rotor assembly 75 allows blades 77 to "pinwheel" relative
to yoke 81, in
which all blades 77 rotate in the same lead or lag direction relative to yoke
81, and this may
especially occur in lag direction 91 during initial rotation about mast axis
79 of rotor assembly 75
from rest. As the centrifugal force on blades 77 builds with their increased
angular velocity, blades
77 will rotate forward in the lead direction 89 to their angular neutral
position relative to yoke 81.
Figures 9 and 10 illustrate another alternative embodiment of an aircraft
rotor assembly 127
according to this disclosure, the views having various components removed for
ease of viewing.
Rotor assembly 127 has a plurality of blades 129 and is driven in rotation by
mast 131 about mast
axis 133 by torque provided by an engine of an aircraft (both not shown).
Rotor assembly 127 is similar in configuration to rotor assembly 15, described
above, in that
assembly 127 has a lead-lag flexure 135 for each blade 129, flexures 135
providing a biasing force
for biasing the associated blade 129 toward a neutral lead-lag position
relative to a central yoke
137. Flexures 135 are similar in construction to flexures 59, described above,
flexures 135
preferably being formed from a composite material or any appropriate material
or combination of
materials.
Unlike yoke 21 of rotor assembly 75, yoke 137 comprises radially extending
arms 139 extending
from a central section 141. In the embodiment shown, each arm 139 has a
flexible portion 143 that
acts as a flexural flap hinge, allowing for out-of-plane flapping motion of an
outer portion of the
associated arm 139 and blade 129 together relative to central section 141 in
directions indicated by
arrows 145, 147. In addition, flexible portion 143 allows for limited rotation
of the outer portion of
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each arm 139 and blade 129 together about pitch axis 149 through force applied
to a pitch horn 151
by a pitch link 153.
A flexure mount 155 mounts each flexure 135 to the outer end of an associated
arm 139, each
mount 155 having an arm clevis 157 configured for attachment to the outer end
of arm 139 and a
flexure clevis 159 configured for attachment to the inner end of flexure 135.
For each mount 155,
flexure clevis 159 is clocked 90 degrees from the orientation of arm clevis
157, thereby orienting
the attached flexure 135 to provide a bending stiffness in flapping directions
145, 147 greater than
a bending stiffness in the lead and lag directions, indicated by arrows 161,
163, respectively. A
blade mount 165 couples each blade 129 to the associated flexure 135, each
mount 165 having a
flexure mount 167 and being coupled rigidly coupled to an inner end of blade
129 by fasteners
169. A grip 171 extends inward from the inner end of each blade 129 to a lead-
lag bearing 173,
which forms a lead-lag pivot for grip 171 and associated blade 129 to move
together relative to
yoke 137 in lead-lag directions 161, 163. Each grip 171 is coupled to blade
mount 165 by fasteners
174. As a blade 129 rotates about the associated lead-lag bearing 173 from the
neutral position, the
associated flexure 135 acts as a spring to provide a biasing force to bias the
blade toward the
neutral position.
Figures 11 and 12 illustrate another alternative embodiment of an aircraft
rotor assembly 175
according to this disclosure, the views having various components removed for
ease of viewing.
Rotor assembly 175 has a plurality of blades 177 and is driven in rotation by
mast 179 about mast
axis 181 by torque provided by an engine of an aircraft (both not shown).
Rotor assembly 175 is unlike configurations described above, in that each
blade 177 has an inner
flexible portion 183 that allows for out-of-plane flapping motion of the outer
portion of blade 177
CA 2995295 2018-02-14
in directions indicated by arrows 185, 187 and acts as a lead-lag pivot to
allow for in-plane lead-lag
motions indicated by arrows 189, 191 respectively. In addition, bending of
flexible portion 183
creates a biasing force to bias the associated blade 177 toward a neutral lead-
lag position relative to
a central yoke 193. Each blade 177 is coupled to an arm 195 of yoke 193 with a
blade mount 197,
each blade 177 rigidly coupled to blade mount 197 by fasteners 199. Bearings
within each arm 195
react centrifugal and shear forces of blade 177 and allow for rotation of each
blade about pitch axis
201 through force applied to a pitch horn 203 on the associated mount 197.
Figure 13 illustrates a tiltrotor aircraft 205. Aircraft 205 comprises a
fuselage 207, a wing 209, and
nacelles 211 mounted to each end of wing 209. Nacelles 211 are configured for
a limited amount
of rotation relative to wing 209, and each nacelle 211 houses an engine (not
shown) for providing
torque to rotate an associated proprotor 213 having multiple blades 215. As
described above,
blades 215 have a Lock number of preferably greater than 6.
Figures 14 and 15 illustrate another alternative embodiment of an aircraft
rotor assembly 217
according to this disclosure, the views having various components removed for
ease of viewing.
As shown, rotor assembly 217 is configured for use as proprotor 213 on
aircraft 205, though rotor
assembly 217 or components thereof may be used on other type of aircraft.
Rotor assembly 217 is
configured to have discrete flap hinges.
Like rotor assembly 15, described above, rotor assembly 217 has a lead-lag
flexure 219 for each
blade 215, flexures 219 providing a biasing force for biasing the associated
blade 215 toward a
neutral lead-lag position relative to a central hub 221. Flexures 219 are
preferably formed from a
composite material but may be formed from any appropriate material or
combination of materials.
Each flexure 219 has a inboard section 223, comprising a hinge aperture 225
and a bearing pocket
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227, and an outboard section 229 with a tang 231 at the outboard end. In the
embodiment shown,
flexure 219 is formed as a beam having a generally rectangular cross-section
of varying size,
outboard section 229 having a smaller chordwise thickness than inboard section
223 to allow for
bending of outboard section 229 during use, forming a lead-lag pivot. Flexure
219 is oriented to
have a bending stiffness greater in the out-of-plane directions, shown by
arrows 233, 235, than a
bending stiffness in the in-plane directions, indicated by arrows 237, 239.
Flexure 219 is stiff to out-of-plane flapping motions of the associated blade
215, and flapping and
coning motion occurs with rotation of flexure 219 relative to hub 221. Hub 221
is affixed to mast
241 for rotation therewith about mast axis 242 and comprises a pair of pillow
block bearings 243
for each blade 215, each pair of bearings 243 cooperating to form a flap-hinge
axis. A shaft (not
shown) extends through bearings 243 of each pair and hinge aperture 225 of the
associated flexure
219, allowing for flexure 219 to rotate a limited amount relative to hub 221
to allow for out-of-
plane motion of blades 215.
Each flexure 219 has a centrifugal-force bearing 245 located in bearing pocket
227 and a bearing
assembly 247 mounted to tang 231. Bearing 245 is preferably a spherical
bearing, and bearing
assembly 247 comprises a mounting clevis 249, a shear bearing 251, and a
folding bearings 253.
Bearing 245 and shear bearing 251 cooperate to form a pitch axis 255, allowing
for blade 215 to
rotate a limited amount relative to hub 221 about pitch axis 255. The pitch
angle of blades 215 is
controlled using swashplate 257, which is translatably carried on splined
section 259 of mast 241.
Pitch links 261 connect swashplate 257 to pitch horns 263, each pitch horn 263
being coupled to a
blade 215, and translation or tilting of swashplate 257 relative to mast 241
causes changes in the
pitch angle of blades 215 about pitch axis 255.
17
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Each blade 215 is coupled to the associated flexure 219 through a blade grip
265. An inboard
section 267 of grip 265 is coupled to bearing 245, and an outboard section 269
is coupled to
bearing assembly 247. A clevis 271 of blade 215 receives grip 265, and clevis
271 and grip 265 are
positioned to align folding apertures 273 and 275 for receiving folding
bearings 253. This allows
for blade 215 to be selectively rotated relative to grip 265 about folding
bearings 253, such as may
be desirable only when aircraft 205 is not in operation.
As with flexures described above, flexure 219 acts as a spring by bending
through elastic
deformation to allow for a selected amount of rotation relative to hub 221 of
the associated blade
215 at least 1 degree in each of the lead direction 237 and lag direction 239,
the bending of flexure
219 producing a biasing force opposing the lead-lag motions and biasing the
blade toward the
neutral position. This allows for selection of the in-plane stiffness of
flexure 219 to "tune" the first
in-plane frequency to be above 1/rev, and no dampers are required to achieve
the desired
frequency.
Figures 16 and 17 illustrate another alternative embodiment of an aircraft
rotor assembly 277
according to this disclosure, the views having various components removed for
ease of viewing.
As shown, rotor assembly 277 is configured for use as proprotor 213 on
aircraft 205, though rotor
assembly 277 or components thereof may be used on other type of aircraft.
Rotor assembly 277 is
configured to gimbal relative to mast 241 and have discrete coning hinges.
Rotor assembly 277 is constructed and operated similarly to rotor assembly
217, as described
above, though components used to couple flexures 219 to mast 241 are
different. To allow for
gimbaling, a hub 279 is coupled to a gimbal 281. Hub 279 comprises a pair of
pillow block
bearings 283 for each blade 215, each pair of bearings 283 cooperating to form
a coning-hinge
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axis. A shaft (not shown) extends through bearings 283 of each pair and hinge
aperture 225 of the
associated flexure 219, allowing for flexure 219 to rotate a limited amount
relative to hub 279 to
allow for out-of-plane coning motion of blades 215. Hub 279 is affixed to
plate 285 of gimbal 281.
Gimbal 281 is configured to allow plate 285 and hub 279 to pivot together
relative to mast 241
about axes perpendicular to axis 242, and hub springs 287 bias plate 285 and
hub 279 toward a rest
position. Gimbal 281 allows for flapping motion of blades 215 to occur at
gimbal 281, whereas
coning motion of blades 215 occurs at bearings 283.
It should be understood that in one or more of the embodiments shown, the
flexural element
providing the biasing force may be formed as an integral component of the yoke
or hub.
At least one embodiment is disclosed and variations, combinations, and/or
modifications of the
embodiment(s) and/or features of the embodiment(s) made by a person having
ordinary skill in the
art are within the scope of the disclosure. Alternative embodiments that
result from combining,
integrating, and/or omitting features of the embodiment(s) are also within the
scope of the
disclosure. Where numerical ranges or limitations are expressly stated, such
express ranges or
limitations should be understood to include iterative ranges or limitations of
like magnitude falling
within the expressly stated ranges or limitations (e.g., from about 1 to about
10 includes, 2, 3, 4,
etc.; greater than 0.10 includes 0.11, 0.12, 0.13, etc.). For example,
whenever a numerical range
with a lower limit, RI, and an upper limit, Ru, is disclosed, any number
falling within the range is
specifically disclosed. In particular, the following numbers within the range
are specifically
disclosed: R=Ri +k * (R-R1), wherein k is a variable ranging from 1 percent to
100 percent with a
1 percent increment, i.e., k is 1 percent, 2 percent, 3 percent, 4 percent, 5
percent,...50 percent, 51
percent, 52 percent,..., 95 percent, 96 percent, 95 percent, 98 percent, 99
percent, or 100 percent.
Moreover, any numerical range defined by two R numbers as defined in the above
is also
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specifically disclosed. Use of the term "optionally" with respect to any
element of a claim means
that the element is required, or alternatively, the element is not required,
both alternatives being
within the scope of the claim. Use of broader terms such as comprises,
includes, and having should
be understood to provide support for narrower terms such as consisting of,
consisting essentially
of, and comprised substantially of. Accordingly, the scope of protection is
not limited by the
description set out above but is defined by the claims that follow, that scope
including all
equivalents of the subject matter of the claims. Each and every claim is
incorporated as further
disclosure into the specification and the claims are embodiment(s) of the
present invention. Also,
the phrases "at least one of A, B, and C" and "A and/or B and/or C" should
each be interpreted to
include only A, only B, only C, or any combination of A, B, and C.
CA 2995295 2018-02-14
