Patent Description:
Rotary wing aircraft, or rotorcraft, can generate significant vibratory responses during operation. The primary source of such vibration is that generated by the main rotor system rotating at the blade passing frequency and the periodic loads acting on the rotor blades. Forces and moments are transmitted through the gearbox into the airframe, resulting in airframe vibration.

Active vibration control (AVC) systems that are characterized by anti-vibration actuators mounted in the fuselage or on or very near the helicopter main transmission to suppress otherwise high levels of vibration are often heavier than desirable, resulting in a reduced payload of the aircraft. For systems intended to generally completely suppress the aircraft vibration, six or more anti-vibration actuators, typically the heaviest components in AVC systems, are required. Thus it is desirable to reduce the weight and number of AVC actuators.

For rotor-based anti-vibration systems, the corresponding rotor-based actuators are typically oscillated at frequencies of the fuselage or other vibrations to be suppressed. In general, rotor-based systems cannot totally suppress all of the vibratory loads originating from the main rotor(s) either because the number of distinct load directions is greater than the number of controls or because the power or amplitude needed to provide complete suppression of the vibration by the rotor-based system alone is onerous. Also, the rotor-based system might be tasked with improving rotor efficiency or some other attribute and not tasked with reducing vibration. As a result, the residual vibration may thus "leak" into the airframe and cause unwanted vibration.

<CIT> discloses a rotary wing aircraft including a dual counter-rotating, coaxial rotor system having an upper rotor system and a lower rotor system rotatable about a common axis. A plurality of blade assemblies is mounted to a portion of either the upper rotor system or the lower rotor system. A plurality of individually controllable actuators is coupled to each of the plurality of blade assemblies. Each of the plurality of actuators is configured to control movement of the coupled blade assembly about a pitch axis. The rotary-wing aircraft additionally includes a sensor system within an airframe. A higher harmonic control (HHC) controller is arranged in communication with the sensor system and the plurality of actuators to individually control the upper rotor system and the lower rotor system to reduce vibration. In embodiments where six controls cannot suppress the six vibratory hub loads at location L, force generators of an active vibration control (AVC) system positioned throughout the airframe may be used in combination with the HHC system to minimize the vibration in the airframe. Further, document <CIT> according to its abstract discloses at least one of an integrated actuator and an intermediate actuator associated with a first source of vibration, a sensor configured to sense vibration from the first source of vibration, a dedicated actuator configured for association with a fuselage, and a controller configured to receive information from the sensor and configured to control the dedicated actuator and the at least one of the integrated actuator and the intermediate actuator.

The invention relates to a rotary-wing aircraft according to appended claim <NUM> and to a method of reducing vibration in a rotary wing aircraft according to appended claim <NUM>. Preferable embodiments are disclosed in the dependent claims.

The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:.

<FIG> illustrates an exemplary vertical takeoff and landing (VTOL) rotary wing aircraft <NUM> having a dual, counter-rotating main rotor system <NUM>, which rotates about a rotating upper main rotor shaft 14U, and a counter-rotating lower main rotor shaft <NUM> (<FIG> and <FIG>), both about an axis of rotation A. The aircraft <NUM> includes an airframe F which supports the dual, counter-rotating, coaxial main rotor system <NUM> 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 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 aircrafts will also benefit from the present invention.

A main gearbox G (<FIG>) located above the aircraft cabin drives the rotor system <NUM>. The translational thrust system T may be driven by the same main gearbox G which drives the rotor system <NUM>. The main gearbox G is driven by one or more engines (illustrated schematically at E in <FIG>). As shown, the main gearbox G may be interposed between the gas turbine engines E, the rotor system <NUM>, and the translational thrust system T.

Referring now to <FIG>, the dual, counter-rotating, coaxial rotor system <NUM> includes an upper rotor system <NUM> and a lower rotor system <NUM>. Each rotor system <NUM>, <NUM> includes a plurality of rotor blade assemblies <NUM> mounted to a rotor hub assembly <NUM>, <NUM> for rotation about the rotor axis of rotation A. The rotor hub assembly <NUM> is mounted to the upper rotor shaft 14U which counter rotates within the lower rotor shaft <NUM>, which rotates with the lower hub assembly <NUM>.

The plurality of main rotor blade assemblies <NUM> project substantially radially outward from the hub assemblies <NUM>, <NUM>. Any number of main rotor blade assemblies <NUM> may be used with the rotor system <NUM>. Each rotor blade assembly <NUM> of the rotor system <NUM> generally includes a rotor blade <NUM> (illustrated somewhat schematically), a rotor blade spindle <NUM>, and a rotor blade bearing <NUM>, which supports the rotor blade spindle <NUM> within a bearing housing <NUM> to permit the rotor blade <NUM> to pitch about a pitching axis P. It should be understood that various blade attachments may be utilized with the present invention.

Referring to <FIG>, a lower rotor control system <NUM> may include a rotor blade pitch control horn <NUM> mounted for rotation with the rotor blade spindle <NUM> of each lower rotor blade <NUM>. Each lower rotor blade pitch control horn <NUM> is linked to a lower rotor swashplate <NUM> through pitch control rods <NUM>. The lower rotor swashplate has two halves. One half rotates with the pitch control rods <NUM> while the other half is non-rotating and is linked to a lower rotor servo mechanism <NUM> to impart the desired pitch control thereto. Similarly, an upper rotor control system <NUM> includes rotor blade pitch control horn <NUM> mounted for rotation with the rotor blade spindle <NUM> of each upper rotor blade <NUM>. The upper rotor blade pitch control horn <NUM> is linked to an upper rotor swashplate <NUM> through a pitch control rod <NUM>. The upper rotor swashplate has two halves. One half rotates with the upper rotor pitch control rods <NUM> while the other half is non-rotating and is linked to upper rotor servo mechanism <NUM> to impart the desired pitch control thereto.

In such embodiments, each rotor control system <NUM>, <NUM> is independently controlled through the separate swashplate assemblies <NUM>, <NUM> which selectively articulate each rotor system <NUM>, <NUM>. Generally, motion of the swashplate assemblies <NUM>, <NUM> along the rotor axis A will cause the rotor blades <NUM> of the respective rotor system <NUM>, <NUM> to vary pitch collectively and tilting of the swash plate assemblies <NUM>, <NUM> with respect to the axis A will cause the rotor blades <NUM> to vary pitch cyclically and tilt the rotor disk. The swashplate assemblies <NUM>, <NUM> translate and/or tilt by separate servo mechanisms <NUM>, <NUM>. The rotor pushrods are in the rotor rotating reference system of the respective rotor while the servos are in the non-rotating reference system which selectively articulates each rotor system <NUM>, <NUM> independently in both cyclic and collective in response to a rotor control system <NUM>, <NUM>. The rotor control systems <NUM>, <NUM> preferably communicate with a flight control system which receives pilot inputs from controls such as a collective stick, cyclic stick, foot pedals and the like.

It should be understood that the pitch control rods and servo mechanisms <NUM>, <NUM> for the upper rotor system <NUM> and a similarly for the lower rotor system <NUM> may be located internally or externally to the respective main rotor shaft 14U, <NUM> and that various pitch control rods, links and servo mechanisms at various locations for cyclic and collective pitch control of the rotor system <NUM> are contemplated herein.

Alternatively, with reference to <FIG>, the lower rotor control system <NUM> may be individual blade control (IBC) system including an electrical actuator <NUM> directly or indirectly coupled to the rotor blade spindle <NUM> of each lower rotor blade <NUM>. Similarly, an upper rotor IBC system <NUM> includes an electrical actuator <NUM> coupled to the rotor blade spindle <NUM> of each upper rotor blade <NUM> (<FIG>). The actuators <NUM>, <NUM> are configured to impart a desired pitch control to the rotor blades <NUM>. In the illustrated embodiment, each actuator <NUM>, <NUM> is mounted to a respective rotor hub <NUM>, <NUM> adjacent one of the plurality of rotor blade spindles <NUM> such that the actuators <NUM>, <NUM> rotate about an axis R parallel and/or arranged approximately within the same plane as the axis P of an adjacent rotor blade <NUM>. Rotation of each actuator <NUM>, <NUM> is transferred to a corresponding rotor blade <NUM> through a linkage <NUM>, such as a connector or gear chain for example. Although a particular IBC system configuration is illustrated in the disclosed embodiment, other IBC systems configurations, such as having actuators <NUM>, <NUM> mounted concentrically with the rotor blade spindles <NUM> are within the scope of the invention. Other IBC systems such as having the actuators <NUM> mounted radially are within the scope of the invention.

The actuators or mechanisms, for example <NUM> and <NUM> or <NUM> and <NUM>, of each rotor system <NUM>, <NUM> may be independently controlled. However, the plurality of actuators within each rotor control system <NUM>, <NUM> are commonly controlled together. In one embodiment, the actuators or mechanisms are used to similarly rotate the rotor blades <NUM> of each rotor system <NUM>, <NUM>, thereby varying the pitch at frequencies of multiples greater than one of the rotor rotational speed Ω with respect to axes P for vibration control. Additionally for primary control, the plurality of actuators or mechanisms within each rotor control system <NUM>, <NUM> may be used to vary the pitch of the rotor blades <NUM> collectively at a frequency of zero and cyclically at a frequency of Ω. The rotor control systems <NUM>, <NUM> are configured to communicate with a flight control system (not shown) which receives pilot inputs from inceptors such as a collective stick, a cyclic stick, foot pedals, and the like, and upon which one or more vibration reducing commands are superimposed.

Referring to <FIG>, the main rotor system <NUM> is mounted to the airframe F (<FIG>) at a location L and vibrations thereto are transferred at location L. Each of the upper and lower rotor system <NUM>, <NUM> generates six unique vibratory loads. The coaxial rotor system <NUM> thereby provides twelve vibratory hub loads. The twelve vibratory hub loads combine in the rotor system <NUM> to yield six loads applied to the airframe F at the location L. The two rotor systems <NUM>, <NUM> do not produce the same set of three six-force patterns because of the difference in position of the two rotor systems <NUM>, <NUM> i.e., they have different "leverage" with regard to location L and experience different aerodynamic environments. The six net vibratory hub loads at location L call for individual suppression to reduce airframe vibration.

The dual, counter-rotating, coaxial rotor system <NUM> enables individual control of the upper rotor system <NUM> and the lower rotor system <NUM>. The lower rotor control system <NUM> provides up to three independent controls and the upper rotor control system <NUM> provides up to three independent controls. Together the two rotor systems <NUM>, <NUM> can provide a total of six controls or "knobs" to reduce or theoretically eliminate air-frame vibration. In a dual, counter-rotating, coaxial rotor system <NUM>, application of higher harmonic control to the two rotor systems <NUM>, <NUM> which are located on the common axis A, may yield a substantial vibration reduction by suppressing the six loads.

In embodiments where the six controls cannot suppress the six vibratory hub loads at location L, an active vibration control (AVC) system <NUM> is used in combination with the HHC system <NUM> to minimize the vibration in the airframe F. With reference now to <FIG>, a schematic diagram of an AVC system and a higher harmonic control (HHC) system for reducing the vibrations experienced by the aircraft due to the rotating main rotor assembly <NUM> is illustrated.

As shown, the aircraft <NUM> includes a controller <NUM>, such as a real time self adaptive (RTSA) controller for example, and a sensor system <NUM> including a plurality of sensors <NUM> mounted about the aircraft <NUM>. It should be understood that various types of sensors <NUM> arranged at various locations are contemplated herein. In an embodiment, the sensors <NUM> are mounted in the cockpit or in areas where crew or other passengers are located. The sensors <NUM> may be accelerometers configured to generate signals representative of dynamic changes at selected locations of the aircraft <NUM> as the main rotor assembly <NUM> operates. The controller <NUM> is arranged in communication with the sensor system <NUM> to sense vibration within the airframe F.

The AVC system <NUM> includes one or more force generators <NUM> mounted about the fuselage F or at any suitable location of the aircraft <NUM>. The plurality of force generators <NUM> are coupled to a power source (not shown) such as an electric motor, air motor, hydraulic motor, or turbine for example, and is arranged in communication by means of electrical or wireless communication links <NUM> with the controller <NUM>. In response to the sensor data, the controller <NUM> is configured to output signals to the force generators <NUM> by means of the electrical or wireless communications links <NUM> to control phase and/or magnitude characteristics of the force generators <NUM>. Operation of the force generators <NUM> is continuously varied by the controller <NUM> to correspond to changing dynamic characteristics of the aircraft to reduce or eliminate the vibratory forces experienced by the aircraft <NUM>.

The HHC system <NUM> generally includes a HHC-controller <NUM> arranged in communication with the AVC- controller <NUM>. In some embodiments, the controller <NUM> includes a plurality of computers, for example three computers that provide triplex redundancy for flight critical components, such as the rotor control systems <NUM>, <NUM> for example. In the illustrated, non-limiting embodiment, the one or more computers within <NUM> are operably coupled to the controller <NUM> via a communication bus <NUM>. It should be understood that in the event that the communication between the controller <NUM> and the computers within <NUM> is lost, the controller <NUM> is capable of operating the AVC system <NUM> to maximize a reduction in the vibration of the aircraft <NUM> using only the force generators <NUM>. In normal operation, the AVC-controller <NUM> sends commands to both the AVC actuators <NUM> and the HHC-actuators <NUM>, <NUM> to minimize vibration.

Each of the one or more computers <NUM> are additionally arranged in communication with an upper HHC actuator system <NUM> and a lower HHC actuator system <NUM> which are operable to implement a higher harmonic blade pitch to the upper and lower rotor systems <NUM>, <NUM>. The computers within <NUM> provide closed loop control of the upper actuator system <NUM> and of the lower actuator system <NUM> individually to minimize vibration of the rotor system <NUM> in accordance with an HHC algorithm executed by the controller <NUM>.

Claim 1:
A rotary wing aircraft (<NUM>) comprising:
a rotor system (<NUM>) rotatable about an axis relative to an airframe (F) and which comprises a dual, counter-rotating, coaxial rotor system having an upper rotor system (<NUM>) and a lower rotor system (<NUM>) rotatable about a common axis relative to the airframe (F);
a plurality of blade assemblies (<NUM>) mounted to the rotor system (<NUM>);
a higher harmonic control system (<NUM>) including at least one computer (<NUM>) operably coupled to at least one actuator (<NUM>, <NUM>), the at least one computer (<NUM>) being operable to receive a higher harmonic control signal from a controller (<NUM>) and operate the at least one actuator (<NUM>, <NUM>) in response to the higher harmonic control signal;
an active vibration control system (<NUM>) including at least one force generator (<NUM>) mounted about the airframe (F) to further reduce vibration within the airframe of the aircraft, the active vibration control system (<NUM>) being operable to generate vibration forces about the aircraft (<NUM>) according to an active vibration control signal and including an electrical or wireless link (<NUM>) coupled to the at least one force generator (<NUM>), the link (<NUM>) being operable to control at least one of phase and magnitude characteristics of the at least one force generator (<NUM>) in response to the active vibration control signal from the controller (<NUM>); and
the controller (<NUM>) operably coupled to the at least one force generator (<NUM>) and operably coupled to the at least one computer (<NUM>) via a communication bus (<NUM>), the controller (<NUM>) being operable to selectively communicate the higher harmonic control signal to the at least one computer (<NUM>) and communicate the active vibration control signal to the at least one force generator (<NUM>) to coordinate operation of the higher harmonic control system (<NUM>) and the active vibration control system (<NUM>) to reduce vibration within the airframe (F),
wherein the at least one computer (<NUM>) includes a plurality of computers for providing redundancy and provides closed loop control of the at least one actuator (<NUM>, <NUM>).