U.S. patent application number 11/521032 was filed with the patent office on 2010-01-07 for dual higher harmonic control (hhc) for a counter-rotating, coaxial rotor system.
This patent application is currently assigned to Sikorsky Aircraft Corporation. Invention is credited to William A. Welsh.
Application Number | 20100003133 11/521032 |
Document ID | / |
Family ID | 39184531 |
Filed Date | 2010-01-07 |
United States Patent
Application |
20100003133 |
Kind Code |
A1 |
Welsh; William A. |
January 7, 2010 |
DUAL HIGHER HARMONIC CONTROL (HHC) FOR A COUNTER-ROTATING, COAXIAL
ROTOR SYSTEM
Abstract
A dual, counter-rotating, coaxial rotor system provides
individual control of an upper rotor system and a lower rotor
system. The lower rotor control system and the upper rotor control
system provide six controls or "knobs" to minimize or theoretically
eliminate airframe vibration. In a dual, counter-rotating, coaxial
rotor system, application of a HHC system to the two rotor systems
individually but located on the common axis, will yield essentially
complete vibration reduction because the 6 controls will suppress
the 6 loads.
Inventors: |
Welsh; William A.; (North
Haven, CT) |
Correspondence
Address: |
CARLSON, GASKEY & OLDS, P.C.
400 WEST MAPLE ROAD, SUITE 350
BIRMINGHAM
MI
48009
US
|
Assignee: |
Sikorsky Aircraft
Corporation
|
Family ID: |
39184531 |
Appl. No.: |
11/521032 |
Filed: |
September 14, 2006 |
Current U.S.
Class: |
416/1 ;
416/31 |
Current CPC
Class: |
Y02T 50/34 20130101;
B64C 27/001 20130101; B64C 2027/7238 20130101; Y02T 50/30 20130101;
B64C 27/10 20130101; B64C 2027/7233 20130101 |
Class at
Publication: |
416/1 ;
416/31 |
International
Class: |
B64C 27/605 20060101
B64C027/605 |
Claims
1-4. (canceled)
5. A rotary-wing aircraft comprising: a dual, counter-rotating,
coaxial rotor system having an upper rotor system and a lower rotor
system rotatable about a common axis of rotation; an upper
swashplate linked to said upper rotor system: a lower swashplate
linked to said lower rotor system; a sensor system within an
airframe; an upper HHC actuator system which includes at least one
individually controllable actuator in each axis to control said
upper swashplate and said upper rotor system in said X-Y-Z axis; a
lower HHC actuator system which includes at least one individually
controllable actuator in each axis to control said lower swashplate
and said lower rotor system in said X-Y-Z axis; and a HHC
controller in communication with said sensor system, said upper HHC
actuator system, and said lower HHC system to individually control
said upper rotor system and said lower rotor system to reduce
vibration.
6. The aircraft as recited in claim 5, wherein said HHC controller
provides closed loop control of said upper HHC actuator system and
said lower HHC actuator system.
7. The aircraft as recited in claim 5, wherein said upper HHC
actuator system includes a fore-aft cyclic actuator, a left-light
cyclic actuator and a collective actuator to independently control
said upper swashplate and said lower HHC actuator system
respectively include a to independently control said lower
swashplate.
8. A method of reducing vibration in a rotary wing aircraft
airframe having a dual, counter-rotating, coaxial rotor system
having an upper rotor system and a lower rotor system rotatable
about a common axis of rotation comprising: individually
controlling at least one individually controllable actuator in each
axis to control an upper swashplate and control an upper rotor
system in said X-Y-Z axis with an upper HHC actuator system to
reduce vibration within an airframe of the aircraft; and
individually controlling a lower HHC actuator system which includes
at least one individually controllable actuator in each axis to
control a lower swashplate and control a lower rotor system in said
X-Y-Z axis: with a lower HHC actuator system to reduce vibration
within the airframe of the aircraft.
9. A method as recited in claim 8, wherein said individually
controlling further comprises: individually controlling a fore-aft
cyclic actuator, a left-right cyclic actuator and a collective
actuator of the upper rotor system; and individually controlling a
fore-aft cyclic actuator, a left-right cyclic actuator and a
collective actuator of the lower rotor system independent of the
respective fore-aft cyclic actuator, the left-right cyclic actuator
and the collective actuator of the upper rotor system.
10.-11. (canceled)
12. A method as recited in claim 8, wherein said individually
controlling further comprises: overlaying pilot inputs to the dual,
counter-rotating, coaxial rotor system the HHC actuator system and
the lower HHC actuator system inputs to the dual, counter-rotating,
coaxial rotor system.
13. A method as recited in claim 8, wherein the upper rotor system
and the lower rotor system each generate six unique vibratory loads
such that the counter rotating, coaxial rotor system generates
twelve vibratory hub loads, the twelve vibratory hub loads combine
in the counter rotating, coaxial rotor system to yield six net
vibratory loads applied to the airframe; and suppressing the six
net vibratory loads by individually controlling a fore-aft cyclic
actuator, a left-right cyclic actuator and a collective actuator of
the upper rotor system and individually controlling a fore-aft
cyclic actuator, a left-right cyclic actuator and a collective
actuator of the lower rotor system.
14. The aircraft as recited in claim 5, wherein said upper rotor
system and said lower rotor system each generate six unique
vibratory loads such that said counter rotating, coaxial rotor
system generates twelve vibratory hub loads, the twelve vibratory
hub loads combine in said counter rotating, coaxial rotor system to
yield six net vibratory loads applied to said airframe, said HHC
controller operable to individually control said at least one
individually controllable actuator in each axis to control said
upper swashplate and said upper rotor system in said X-Y-Z axis and
said at least one individually controllable actuator in each axis
to control said lower swashplate and said lower rotor system in
said X-Y-Z axis to suppress the six net vibratory loads.
15. The aircraft as recited in claim 14, wherein said HHC
controller utilizes matrix arithmetic with a square matrix that
quantitatively relates the influence of said at least one
individually controllable actuator in each axis of said upper rotor
system and said at least one individually controllable actuator in
each axis of said lower rotor system on the six unique vibratory
loads.
16. The aircraft as recited in claim 15, wherein said HHC
controller utilizes an inverse of said square matrix to determine a
control solution which controls said at least one individually
controllable actuator in each axis of said upper rotor system and
said at least one individually controllable actuator in each axis
of said lower rotor system.
Description
BACKGROUND OF THE INVENTION
[0001] The present invention relates to a rotary-wing aircraft, and
more particularly to a HHC vibration control system therefor.
[0002] The reduction of vibrations is a primary goal in rotary-wing
aircraft design. Such vibrations may contribute to crew fatigue,
increased maintenance, and operating costs. A major cause of such
vibrations is periodic aerodynamic loads on the rotor blades.
[0003] An effective method of reducing rotor-blade induced
vibrations is to control the harmonic airload at the source, i.e.
on the rotor blades. For an N-bladed rotor, harmonic loads at a
frequency rate of (N-1) per revolution, N per revolution and (N+1)
per revolution are transmitted to the rotor hub. All three harmonic
load frequencies in the rotating rotor frame of reference result in
fuselage vibration in the non-rotating frame of reference at the
frequency rate of N/revolution (hereinafter NP). The function of
HHC devices is to generate additional airloads on the rotor so as
to cancel the NP vibratory hub load in the non-rotating frame of
reference.
[0004] Various schemes for reducing helicopter vibrations by HHC
have been investigated. Some approaches are based on passive
vibration control concepts involving dynamically tuned mechanisms
which actuate either the swash plate or the tab surfaces on the
rotor blade. Other HHC concepts make use of high frequency active
control systems which, when coupled with vibration sensors, provide
vibration reduction by either manual control or closed loop
feedback control.
[0005] Reducing rotor-blade induced vibrations in a dual,
counter-rotating, coaxial rotor system is further complicated as
control inputs to the upper rotor control system and lower rotor
control system are typically linked or slaved. As such, HHC systems
have heretofore been linked or slaved such that the HHC inputs to
the upper rotor system are a fixed multiple of the inputs to the
lower rotor system. Such linkage may be acceptable to minimize
vibrations to a certain extent but will not provide the more
significant vibration reduction levels demanded by current
rotary-wing aircraft operators.
[0006] 1980 AHS (American Helicopter Society) paper entitled
"Design of Higher Harmonic Control for the ABC", J. O'Leary and W.
Miao, publicly describe the originally HHC proposed system. The
system does control six "signal" actuators to provide HHC inputs to
the main servos i.e. three inputs to the signal actuators inputting
into main servo of the upper rotor and three inputs to the signal
actuators inputting into the main servos of the lower rotor. In
this case, however, the three inputs to the upper signal actuators
are fixed multiples of the three inputs to the lower signal
actuators such that the upper rotor signal inputs are "slaved" to
the lower rotor inputs. The pilot flight controls for the upper and
lower rotors were also slaved together. Thus, following this
slaving philosophy for the HHC inputs was a natural approach. It is
apparent in this AHS paper that the analyses projected imperfect
vibration control. This is because the slaving process only
produces a total of three unique controls whereas there are up to
six vibratory hub loads that require suppression for excellent
vibration reduction.
[0007] Accordingly, it is desirable to provide an HHC system which
essentially provides total suppression of vibration heretofore
unachieved in a dual, counter-rotating, coaxial rotor system.
SUMMARY OF THE INVENTION
[0008] The HHC system according to the present invention generally
includes an HHC controller, a sensor system in communication with
the HHC controller and an upper HHC actuator system and a lower HHC
actuator system which implements individual higher harmonic blade
pitch to the upper rotor system and the lower rotor system. The
upper HHC actuator system and the lower HHC actuator system each
include a fore-aft cyclic actuator, a left-right cyclic actuator,
and a collective actuator (x-y-z axes).
[0009] A dual, counter-rotating, coaxial rotor control system
provides individual control of the upper rotor system and the lower
rotor system. The lower rotor control system and the upper rotor
control system provide six controls or "knobs" to minimize or
theoretically eliminate airframe vibration. The current invention
is based upon the key realization that the two rotors enable a
total of six independent HHC controls which is equal to the number
of vibratory hub loads created by two co-axial rotors in forward
flight. For a vehicle equipped with co-axial rotors, the vibration
from the two rotors combine to produce only six vibratory hub loads
at the point where the rotor support shaft attaches to the
aircraft. Thus, applying HHC to the two rotors can yield six
independent controls to completely suppress six hub loads.
[0010] The present invention therefore provides a HHC system which
essentially provides total suppression of vibration heretofore
unachieved in a dual, counter-rotating, coaxial rotor system.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] 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:
[0012] FIG. 1 is a general perspective side view of an exemplary
rotary-wing aircraft embodiment for use with the present
invention;
[0013] FIG. 2 is an expanded partial phantom view of a dual,
counter-rotating, coaxial rotor system of the aircraft of FIG.
1;
[0014] FIG. 3 is a schematic longitudinal sectional view of a dual,
counter-rotating, coaxial rotor control system in a flight
position;
[0015] FIG. 4 is a block diagram of the forces moments and control
loads available for Higher Harmonic control of a dual,
counter-rotating, coaxial rotor system; and
[0016] FIG. 5 is a block diagram of a Higher Harmonic control
system for a dual, counter-rotating, coaxial rotor system; and
[0017] FIG. 6A is a matrix explanation for the Higher Harmonic
control system of a dual, counter-rotating, coaxial rotor system
having Forces and moments as represented in FIG. 6B.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
[0018] FIG. 1 illustrates an exemplary vertical takeoff and landing
(VTOL) rotary-wing aircraft 10 having a dual, counter-rotating,
coaxial rotor system 12 which rotates about a rotating main rotor
shaft 14U, and a counter-rotating main rotor shaft 14L (FIG. 2)
both about an axis of rotation A. The aircraft 10 includes an
airframe F which supports the dual, counter rotating, coaxial rotor
system 12 as well as an optional translational thrust system T
which provides translational thrust during high speed forward
flight generally parallel to an aircraft longitudinal axis L.
Although a particular aircraft configuration is illustrated in the
disclosed embodiment, other counter-rotating, coaxial rotor systems
will also benefit from the present invention. Although a particular
counter-rotating, coaxial rotor system aircraft configuration is
illustrated in the disclosed embodiment, other rotor systems and
other aircraft types such as tilt-wing and tilt-rotor aircraft will
also benefit from the present invention.
[0019] A main gearbox G which may be located above the aircraft
cabin drives the rotor system 12. The translational thrust system T
may be driven by the same main gearbox G which drives the rotor
system 12. The main gearbox G is driven by one or more engines
(illustrated schematically at E). As shown, the main gearbox G may
be interposed between the gas turbine engines E, the rotor system
12 and the translational thrust system T.
[0020] Referring to FIG. 2, the dual, counter-rotating, coaxial
rotor system 12 includes an upper rotor system 16 and a lower rotor
system 18. Each rotor system 16, 18 includes a plurality of rotor
blade assemblies 20 mounted to a rotor hub assembly 22, 24 for
rotation about the rotor axis of rotation A. The rotor hub assembly
22 is mounted to the upper rotor shaft 14U which counter rotates
within the lower rotor shaft 14L which rotates the lower hub
assembly 24.
[0021] The plurality of the main rotor blade assemblies 20 project
substantially radially outward from the hub assemblies 22, 24. Any
number of main rotor blade assemblies 20 may be used with the rotor
system 12. Each rotor blade assembly 20 of the rotor system 12
generally includes a rotor blade 28 (illustrated somewhat
schematically), a rotor blade spindle 30, and a rotor blade bearing
32 which supports the rotor blade spindle 30 within a bearing
housing 34 to permit the rotor blade 28 to pitch about a pitching
axis P. It should be understood that various blade attachments may
also be utilized with the present invention.
[0022] Referring to FIG. 3, a lower rotor control system 36
includes a rotor blade pitch control horn 38 mounted for rotation
with the rotor blade spindle 30 of each lower rotor blade 28. Each
rotor blade pitch control horn 38 is linked to a lower swashplate
40 through a pitch control rod and servo mechanism 42 to impart the
desired pitch control thereto. An upper rotor control system 44
includes a rotor blade pitch control horn 46 mounted for rotation
with the rotor blade spindle 30 of each upper rotor blade 28. Each
rotor blade pitch control horn 46 is linked to an upper swashplate
48 through a pitch control rod and servo mechanism 50 to impart the
desired pitch control thereto.
[0023] Each rotor system 36, 44 is independently controlled through
the separate swashplate assemblies 40, 48 which selectively
articulates each rotor system 36, 44. Generally, motion of the
swashplate assemblies 40, 48 along the rotor axis A will cause the
rotor blades 20 of the respective rotor system 36,44 to vary pitch
collectively and tilting of the swash plate assemblies 40, 48 with
respect to the axis A will cause the rotor blades 20 to vary pitch
cyclically and tilt the rotor disk. The swashplate assemblies 40,
48 translate and/or tilt by a separate servo mechanism 42, 50. The
upper rotor pushrods are in the rotating reference system while the
servos are in the non-rotating reference system which selectively
articulates each rotor system 36, 44 independently in both cyclic
and collective in response to a rotor control system (illustrated
schematically). The rotor control system preferably communicates
with a flight control system which receives pilot inputs from
controls such as a collective stick, cyclic stick, foot pedals and
the like.
[0024] It should be understood that the pitch control rods and
servo mechanisms 42, 50 for the upper rotor system 16 and a lower
rotor system 18 may be located internally or externally to the
respective main rotor shaft 14U, 14L and that various pitch control
rods, links and servo mechanism at various locations for cyclic and
collective pitch control of the rotor system 12 may be utilized
with the present invention. Furthermore, it should be understood
that rotor control systems other than swashplates will likewise be
usable with the present invention.
[0025] Referring to FIG. 4, the rotor system 12 is mounted to the
airframe F at a location L and vibrations thereto are transferred
at location L. Each rotor system 16, 18 generates six unique
vibratory loads. The counter rotating, coaxial rotor system 12
thereby provides twelve vibratory hub loads. The twelve vibratory
hub loads combine in the rotor system 12 to yield six loads applied
to the airframe F at the location L. The two rotor systems 16, 18
do not produce the same set of three six-force patterns because of
the difference in position of the two rotor systems 16, 18, i.e.
they have different "leverage" with regard to location L. The six
net vibratory hub loads at location L require individual
suppression to reduce airframe vibration.
[0026] The dual, counter-rotating, coaxial rotor system 12 provides
individual control of the upper rotor system 16 and the lower rotor
system 18. The lower rotor control system 36 and the upper rotor
control system 44 provide a total of six controls or "knobs" to
minimize or theoretically eliminate airframe vibration. In a dual,
counter-rotating, coaxial rotor system 12, application of HHC to
the two rotor systems 16, 18 which are located on the common axis
A, will yield essentially complete vibration reduction because the
six controls can suppress the six loads.
[0027] Referring to FIG. 5, a HHC system 52 generally includes a
HHC controller 54, a sensor system 56 in communication with the HHC
controller 54, an upper HHC actuator system 58 and a lower HHC
actuator system 60 which implements the higher harmonic blade pitch
to the upper rotor system 16 and the lower rotor system 18. The HHC
controller 54 is in communication with the sensor system 56 to
sense vibration within the airframe F. It should be understood that
various sensors at various locations may be utilized with the
present invention. The HHC controller 54 is preferably an adaptive
controller to individually control the upper HHC actuator system 58
and the lower HHC actuator system 60. The HHC controller 54
preferably provides closed loop control of the upper HHC actuator
system 58 and the lower HHC actuator system 60 to minimize
vibration thereof in accordance with an HHC algorithm (FIGS. 6A and
6B).
[0028] The upper HHC actuator system 58 and the lower HHC actuator
system 60 preferably each includes a fore-aft cyclic actuator; a
left-right cyclic actuator and a collective actuator (x-y-z axes).
It should be understood that the upper rotor control system 44 and
the lower rotor control system 36 are preferably overlaid or
integrated with the pilot inputs with the HHC actuator system 58
and the lower HHC actuator system 60. It should be understood that
various actuator systems may be utilized with the present invention
so long as active control is provided individually in each axis of
both the upper and lower rotor system.
[0029] The matrix arithmetic shown in FIG. 6a is a math model that
represents how the six HHC controls, U, three from each swashplate,
influence the 6 net hub loads F. The square matrix T quantitatively
relates the influence of U on F. Well known mathematics indicates
that if the matrix T is square, and each column of T is
independent, then matrix T can be "inverted" and a control
solution, U can be found that will make all elements of F equal to
zero. This is a math illustration of why 6 controls or knobs (U)
are required to completely nullify 6 hubloads (F). If only three
unique values of U existed, it would be impossible to completely
nullify F.
[0030] 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
aircraft and should not be considered otherwise limiting.
[0031] It should be understood that although a particular component
arrangement is disclosed in the illustrated embodiment, other
arrangements will benefit from the instant invention.
[0032] 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.
[0033] 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.
* * * * *