Patent Grant
9027873
The present application relates to vibration control in rotorcraft.
Aircraft, such as rotorcraft, experience vibration concerns due in part to one or more large rotating rotors. Rotors are a primary source of vibration upon the fuselage of rotorcraft. Significant fuselage vibrations can contribute to flight control problems, material fatigue, maintenance costs, and pilot fatigue, to name a few. Significant time and expense has been spent in the rotorcraft industry in attempts to reduce and cancel rotorcraft vibrations. Traditionally, rotorcraft vibrations have been addressed with passive devices, such as vibration isolation systems and dynamic absorbers. These passive devices, which are tuned relative to the operating frequency of the aircraft rotor, have proven to be very effective for legacy rotorcraft. However, variable rotor speed (RPM) rotorcraft configurations are becoming more attractive in part due to initiatives to improve performance and reduce noise pollution. As a result, the passive vibration solutions previously effective on single rotor speed rotorcraft fall short when implemented on a variable rotor speed rotorcraft.
Hence, there is a need for an improved system and method for controlling vibration in a rotorcraft.
The novel features believed characteristic of the system of the present application are set forth in the appended claims. However, the system itself, as well as a preferred mode of use, and further objectives and advantages thereof, will best be understood by reference to the following detailed description when read in conjunction with the accompanying drawings, wherein:
While the system of the present application is susceptible to various modifications and alternative forms, specific embodiments thereof have been shown by way of example in the drawings and are herein described in detail. It should be understood, however, that the description herein of specific embodiments is not intended to limit the application to the particular forms disclosed, but on the contrary, the intention is to cover all modifications, equivalents, and alternatives falling within the spirit and scope of the system of the present application as defined by the appended claims.
Illustrative embodiments of system of the present application are described below. In the interest of clarity, not all features of an actual implementation are described in this specification. It will of course be appreciated that in the development of any such actual embodiment, numerous implementation-specific decisions must be made to achieve the developer's specific goals, such as compliance with system-related and business-related constraints, which will vary from one implementation to another. Moreover, it will be appreciated that such a development effort might be complex and time-consuming but would nevertheless be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure.
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 the present application, 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.
The disclosed system and method employs an Active Vibration Control (AVC) system in order to reduce vibration in a rotorcraft. Embodiments of the present system and method can be used in rotorcraft, fixed-wing aircraft, tiltrotor, as well as hybrid aircraft being part rotorcraft and part road vehicle, to name a few. The present AVC system reduces fuselage vibration by utilizing actuators to create controllable actions that minimize vibration produced by the main rotor system. The actuators can function within the rotor system or function in the fuselage. It is preferred that the actuators used in the AVC system reside within the rotorcraft fuselage; however, alternative embodiments may employ effectors in a variable of locations, including outside the fuselage. Actuators can be of several forms such as, electro-mechanical force generating devices attached to the fuselage structure, devices attached at the fuselage/rotor interface, devices attached to each rotor blade, or devices that change the shape of the rotor blade in a controllable manner.
The AVC system of the present application is particularly well suited for variable rotor speed (RPM) rotorcraft configurations; however, even constant RPM rotorcraft can effectively employ the AVC system of the present application. For example, in any rotorcraft, the dynamic response of the rotorcraft can change significantly as the rotorcraft's gross weight varies as a result of fuel consumption and cargo changes. Rotorcraft airframe vibrations characteristics can also vary due to time, use, aging of the airframe, and installation of non-standard equipment. Further, vibration characteristics of individual rotorcraft vary due to normal tolerance variation in manufacturing and assembly processes, thereby making each rotorcraft unique to a certain degree. The AVC system of the present application is robust, adaptive, and has the ability to suppress vibrations for expected variations in airframe vibration characteristics. In addition, the adaptive features of the AVC system are transparent to the passengers of the aircraft, so that adjustments to the AVC system are not perceived by occupants of the rotorcraft. Further, the adaptive mechanisms of the AVC system are automated thereby not requiring pilot or crewmember inputs.
Now referring to
If Transfer Function (G) were to remain constant, or behave linearly, then a frequency domain optimal control formulation could be used. Expression (5), as noted in Wayne Johnson (1982) “Self-Tuning Regulators for Multicycle Control of Helicopter Vibrations, Nasa Technical Paper 1996, March 1982, is a frequency domain optimal control formulation which assumes a linear Transfer Function (G) relationship between the resulting vibrations and the oscillatory load inputs from the AVC actuators. Expressions (1)-(4) depict a derivation of the optimal control formulation in expression (5). In the frequency domain, the vibration measured by each sensor, or accelerometer, at the blade passage frequency (NΩ) can be represented as:
Similarly, a frequency domain representation of the actuator command signals can be expressed as:
Mathematically, the vibration control problem can be represented by a global model as:
{z}={x}+{y}={x}+[G]{u} (3)
where the vector {z} is the harmonic coefficients of the controlled vibration response, {x} is the uncontrolled vibration response due to external excitation such as the main rotor, and {y} is the harmonic coefficients of the vibration generated by the AVC system. Matrix [G] is the complex transfer function between the AVC actuator inputs {u} and the fuselage vibration response. Effective vibration suppression is achieved when the norm of vector {z} approaches zero, or when the harmonic coefficients of vector {y} are of equal magnitude and opposite sign to the corresponding coefficients of vector {x}. The appropriate control action {u} is computed based on optimal control theory, which minimizes a quadratic cost function. The cost function J is a weighted sum of the resultant vibration and the control actions. Defined as:
J={z}T[Wz]{z}+{u}T[Wu]{u} (4)
where [Wz] and [Wu] are diagonal matrices of weighting factors. The weighting matrices [Wz] and [Wu] can be tailored to emphasize vibration reduction over control effort reduction or vice versa. Also, [Wz] and [Wu] can be tailored to put more emphasis on certain elements of {z} or {u} respectively. By minimizing the cost function J, the optimal control solution is simultaneously achieving the greatest vibration reduction with the least control effort. Minimization of the cost function yields the following optimal control effort formulation:
{u}=−([G]T[Wz][G]+[Wu])−1*([G]T[Wz])*{z}=[H]*{z} (5)
The above control expression (5) is optimal and effective provided the transfer function [G] is an accurate representation of the actual AVC system dynamics. One embodiment of a digital feedback control law that utilizes the optimal control solution is the following expression (6):
{u}k+1={u}k+α[H]{z}k (6)
where the subscript k represents a discrete time index, α represents a caution factor (0<α<1), [H] is the optimal gain matrix, {z} is the vector of harmonic coefficients at the current time step, and {u} is the actuator command signal vector of harmonic coefficients. It should be understood that alternate feedback control laws are acceptable to be incorporated into the algorithm.
However, in practice the Transfer Function (G) is not constant throughout the flight of the rotorcraft, nor is it constant from one rotorcraft to another of the same model. Typically, the Transfer Function (G) is frequency dependent, rotorcraft airframe dependent, and time-varying throughout the flight of rotorcraft. Transfer Function (G) changes can result from rotorcraft in flight RPM changes, fuel consumption, cargo changes, rotorcraft aging, or manufacturing variations. If a constant feedback gain matrix (H) was to be used, and the change in Transfer Function (G) large, then such an AVC system would become unstable and increase vibration rather than reduce vibration. Variations in Transfer Function (G) can be in the form of magnitude changes, phase changes, or combined magnitude and phase changes. Because of variations in Transfer Function (G), AVC System Controller 115 of AVC system 101 requires an adaptive controller algorithm that can accommodate variations in Transfer Function (G) in order to optimally control vibration.
Now referring also to
The Least-Squares routine 205 is structured to minimize an error 209. The error 209 is the difference between a measured vibration response from sensors 107 and a predicted vibration response from the mathematical model of Transfer Function (G). The predicted vibration response is computed from the mathematical model of Transfer Function (G) and commands to actuators 113. When the error 209 approaches zero, the mathematical model of Transfer Function (G) accurately represents the true system dynamics, and the predicted vibration matches the measured vibration from sensors 107. When mathematical model of Transfer Function (G) is an accurate representation, then feedback gain (H) derived from mathematical model of Transfer Function (G) is also accurate, thereby providing optimum vibration attenuation. It should be appreciated that this transfer function identification method, as represented in part as adaptive reference model 211, does not intentionally excite actuators 113 in order to determine transfer function, which can create disturbing vibrations felt by helicopter pilots and crew. It should also be appreciated that adaptive reference model 211 does not require knowledge from aircraft parameters 109, such as gross weight, fuel quantity, airspeed, altitude, and the like, although such information may be beneficial.
The Least-Squares routine 205 utilizes measured vibration from sensors 107, and also utilizes command forces to actuators 113. In order to make adjustments to the mathematical model of Transfer Function (G), a finite set of data must be collected and processed before an adjustment is made. Therefore, the Least-Squares routine 205 may not respond well to rapid transfer function changes. In order to overcome this limitation, adaptive reference model 211 is augmented with gain scheduling, via a gain schedule database 203, for known rapid transfer function changes. In addition, adaptive reference model 211 needs an initial starting guess to adapt upon. As such, an initial reference model 213 may be used as an initial starting guess. It should be appreciated that the gain scheduling functionality allows mathematical model of Transfer Function (G) in adaptive reference model 211 to change when subjected to a known sudden and quick change of the AVC Transfer Function (G). When the Transfer Function (G) changes slowly over time, then enough data can be processed for proper adjustments to mathematical model of Transfer Function (G) and the Least-Squares routine 205 alone responds well.
The Least-Squares routine 205 for reference model 211 is derived with discrete time variables. Error 209, being the difference between the actual measured vibration response and the predicted vibration response from mathematical model of Transfer Function (G), is represented by expression (7):
{e}k=({z}k−{z}k−1)−({x}k−{x}k−1)−[Ĝ]({u}k−{u}k−1){e}k={Δz}k−{Δx}k−[Ĝ]{Δu}k (7)
Where {e} is the error vector of harmonic coefficients for each control sensor, {z} is the measured vibration response, {x} is the uncontrolled vibration response, {u} is the AVC command signal. The matrix [Ĝ] is the mathematical reference model of the AVC Transfer Function (G). Each of the response vectors are written as differences between values at two discrete times. Two successive time indices have been used for illustrative purposes, but is not a requirement. It is assumed that between the selected time indices the command signal has changed ({Δu}>0), and the change in measured vibration {Δz} is mostly due to the AVC command signal change {Δu}. Thus, it is assumed {Δx} is small compared to {Δy} and the error formulation simplifies to expression (8):
{e}k≅{Δz}k−└Ĝ┘{Δu}k or {Δz}k≅└Ĝ┘{Δu}k (8)
The reference model matrix [Ĝ] is a (n×m) matrix and the command signal vector is (m×1). The reference model matrix and command vector can be reorganized such that:
[Φ]kT{{circumflex over (θ)}}=[Ĝ]{Δu}k≅{Δz}k{e}k≅{Δz}k−[Φ]kT{{circumflex over (θ)}} (9)
Vector {{circumflex over (θ)}} is (nm×1) and composed by stacking the rows of [Ĝ] into a column vector. Matrix [Φ]kT is (n×nm) and composed from the command signal vector {Δu} as repeated rows to form a block diagonal matrix. The error function can be minimized in a least-squares sense, and the solution for {{circumflex over (θ)}} by batch mode processing of several observations is:
{{circumflex over (θ)}}=([Φ]kT[Φ]k)−1[Φ]kT{Δz}k (10)
The solution vector {{circumflex over (θ)}} contains the reference model matrix [Ĝ] parameters that minimize the error function in a least-squares sense. Thus, based on the data that was collected and analyzed, the reference model is now an accurate representation of the actual AVC Transfer Functions (G). The solution can also be obtained by a recursive least-squares formulation of the form:
{{circumflex over (θ)}}k={{circumflex over (θ)}}k−1+([Φ]kT[Φ]k)−1[Φ]k({Δz}k−[Φ]kT{{circumflex over (θ)}}k−1) (11)
Another simplified recursive solution that avoids matrix inversion is a Least Mean Square (LMS) solution of the form:
{{circumflex over (θ)}}k={{circumflex over (θ)}}k−1+γ[Φ]k({Δz}k−[Φ]kT{{circumflex over (θ)}}k−1)0<γ<2 (12)
It should be appreciated that other variations on these least-squares solutions may exist, but the underlying solution technique is similar. How the solution is obtained is not critical, just that a solution is obtained to minimize error function.
Again referring to
Still referring to
The method continues with the Least-Squares routine 205 processing data from reference model [Ĝ], change in measured vibration response {γZ}, change in computed command signal {ΔU}, and error function vector {e}. If the magnitude of the change in computed command signal |{ΔU}| is greater than a specified threshold, then a parameter estimate {{circumflex over (θ)}}k of the reference model [Ĝ] is updated, otherwise the parameter estimate {{circumflex over (θ)}}k is left unchanged. This step is enforced to ensure there is persistent excitation for the Least-Squares routine 205. As such, reference model [Ĝ] is used to initialize the parameter estimate {{circumflex over (θ)}}k. If |{ΔU}|>magUthreshold, then {Δu}k is used to construct [Φ]kT and the parameter estimate is updated via {{circumflex over (θ)}}k={{circumflex over (θ)}}k−1+γ[Φ]k({Δz}k−[Φ]kT{{circumflex over (θ)}}k−1). If |{ΔU}|<magUthreshold, then {{circumflex over (θ)}}k={{circumflex over (θ)}}k−1. Based upon the previous criteria, the parameter estimate {{circumflex over (θ)}}k is updated repeatedly. After a specified number of updates, the parameter estimate {{circumflex over (θ)}}k is used to construct a new reference model [Ĝ]. The new reference model [Ĝ] is used to create a new feedback gain matrix [H], (e.g. [H]=−([Ĝ]T[Wz][Ĝ]+[Wu])−1*([Ĝ]T[Wz])). The new mathematical reference model [Ĝ] and the new feedback gain matrix [H] are output. The process continues with updating the gain schedule library 203 with the new reference model [Ĝ] where appropriate based on the other aircraft parameter data 109 and rotor speed RPM data 111. It should be noted that actuator command signals are being continuously computed via a parallel feedback path. Harmonic analysis routine 215 is utilized to output harmonic coefficients of vibration {z}, and combined with the current feed back gain matrix [H] and computes the next set of command signals with control law {u}k+1={u}k+α[H]{z}k.
Referring now to
It is apparent that an application with significant advantages has been described and illustrated. Although the present application is shown in a limited number of forms, it is not limited to just these forms, but is amenable to various changes and modifications without departing from the spirit thereof.
| Filing Document | Filing Date | Country | Kind | 371c Date |
|---|---|---|---|---|
| PCT/US2010/025720 | 3/1/2010 | WO | 00 | 8/22/2011 |
| Publishing Document | Publishing Date | Country | Kind |
|---|---|---|---|
| WO2010/099521 | 9/2/2010 | WO | A |
| Number | Name | Date | Kind |
|---|---|---|---|
| 4819182 | King et al. | Apr 1989 | A |
| 5126641 | Putnam et al. | Jun 1992 | A |
| 5219143 | Staple et al. | Jun 1993 | A |
| 5713438 | Rossetti et al. | Feb 1998 | A |
| 5954169 | Jensen | Sep 1999 | A |
| 6002778 | Rossetti et al. | Dec 1999 | A |
| 6229898 | Goodman | May 2001 | B1 |
| 6402089 | Kiss et al. | Jun 2002 | B1 |
| 6460803 | Kiss et al. | Oct 2002 | B1 |
| 6493689 | Kotoulas et al. | Dec 2002 | B2 |
| 6751602 | Kotoulas et al. | Jun 2004 | B2 |
| 6772074 | Millott et al. | Aug 2004 | B2 |
| 6856920 | Millott et al. | Feb 2005 | B2 |
| 7003380 | MacMartin et al. | Feb 2006 | B2 |
| 7107127 | Goodman | Sep 2006 | B2 |
| 7197147 | Millott et al. | Mar 2007 | B2 |
| 7216018 | Zuo et al. | May 2007 | B2 |
| 7224807 | Welsh et al. | May 2007 | B2 |
| 20060054738 | Badre-Alam | Mar 2006 | A1 |
| Entry |
|---|
| Examination report in related Chinese patent application No. 201080009801.4, mailed Aug. 30, 2013, 7 pages. |
| “Self-Turning Regulators for Multicyclic Control of Helicopter Vibration”, Wayne Johnson, NASA Technical Paper, 1996. |
| International Search Report mailed by ISA/USA, U.S. Patent and Trademark Office on Apr. 22, 2010 for International Patent Application No. PCT/US10/25720. |
| International Preliminary Examination Report mailed by IPEA/US, U.S. Patent and Trademark Office on Apr. 20, 2011 for International Patent Application No. PCT/US10/25720. |
| Corrected International Preliminary Examination Report mailed by IPEA/US, U.S. Patent and Trademark Office on Jun. 20, 2011 for International Patent Application No. PCT/US10/25720. |
| Johnson, Wayne, Self-tuning regulators for multicyclic control of helicopter vibration, NASA Technical Paper 1996, Mar. 1982. |
| Examination report in related Canadian patent application No. 2,752,701, mailed Jul. 17, 2013, 2 pages. |
| Office Action in related Canadian patent application No. 2,752,701, dated Jun. 12, 2014. |
| Extended European Search Report dated Jul. 9, 2014 from counterpart EP App. No. 10746964.5. |
| Number | Date | Country | |
|---|---|---|---|
| 20110303784 A1 | Dec 2011 | US |
| Number | Date | Country | |
|---|---|---|---|
| 61156141 | Feb 2009 | US |
