This paper describes a novel approach for the actuation of a Ni-Mn-Ga single crystal in the martensite phase using transverse acoustic waves. The acoustic waves, generated by piezoelectric stacks, produce shear stress levels high enough to induce twin boundary motion in the Ni-Mn-Ga crystal. By using repeated asymmetric pulses, the crystal can be made to transform reversibly from one of two variants to the other. The resulting engineering shear strain as the crystal transforms is theoretically as high as 12%. Two experimental apparatuses were developed that are able to produce actuation through this acoustic mechanism. In the first, two 33-mode piezoelectric stacks are mounted at a 45 deg angle to the base of a single Ni-Mn-Ga crystal. This geometry allows the longitudinal motion of each stack to be converted into transverse motion of the Ni-Mn-Ga crystal. In experiments using this apparatus, we were able to obtain 7.7% engineering shear strain over the length of the Ni-Mn-Ga crystal, and about 10.6% engineering shear strain over that portion of the crystal that exhibited twin boundary motion, close to the theoretical maximum. Some portions of the crystal appeared to be inactive, due to locking of incompatible twin systems. In the second apparatus, a single 15-mode piezoelectric stack was used to generate shear waves directly. In preliminary testing, we were able to achieve only about 1.9% engineering shear strain, probably due to incomplete conditioning of the Ni-Mn-Ga crystal to improve twin boundary motion.
Single crystal Ni-Mn-Ga ferromagnetic shape memory alloys (FSMAs) are active materials that produce strain when a magnetic field is applied. The large saturation strain (6%) of Ni-Mn-Ga and material energy density comparable to piezoelectric ceramics make tetragonal Ni-Mn-Ga an interesting active material. However, the usefulness of the material is limited by the need for electromagnets to produce a magnetic actuation field. In this paper, an actuation method for shape memory alloys in the martensitic phase is described, in which asymmetric acoustic pulses are used to drive twin boundary motion. Experimental actuators were developed using a combination of Ni-Mn-Ga FSMA single crystals and a piezoelectric stack actuator. In bidirectional actuation without load, strains of over 3% were achieved using repeated pulses (at 100 Hz) over a 30 s interval, while 1% strain was achieved in under 1 s. The maximum strains achieved are comparable to the strains achieved using bidirectional magnetic actuation, although the time required for actuation is longer. No-load actuation also showed a nearly linear relationship between the magnitude of the asymmetric stress pulse and the strain achieved during actuation, and a positive correlation between pulse repetition rate and output strain rate, up to a pulse repetition rate of at least 100 Hz. Acoustic actuation against a spring load showed a maximum output energy density for the actuator of about 1000 J/m3, with a peak-to-peak stress and strain of 100 kPa and 2%, respectively.
A new wavenumber domain sensing method has been developed and applied to feedback controller design for active structural acoustic control. The approach is to minimize the total acoustic power radiated form vibrating structures in the wavenumber domain. If the disturbance spectrum is given, the target wavenumbers in the supersonic domain (i.e., the radiating wavenumbers) can be determined. Then, a state-space model can be found to estimate the magnitude of the supersonic wavenumber components. Once we have a state-space model that can be used for active structural acoustic control, a modern control design paradigm can be applied to minimize the acoustic power radiated from vibrating structures. The new sensing method was numerically validated on a thick-walled cylindrical shell with 55 active composite panels mounted. It is found that the method enables us to systematically find a state-space model for wavenumber components in the supersonic region, and therefore makes it easy to design MIMO LQG controller.
The development and testing of a new actuator for helicopter rotor control, the Double X-Frame, is described. The actuator is being developed for wind tunnel and flight testing on an MD-900 helicopter. The double X-frame actuator has a number of design innovations to improve its performance over the original X-frame design of Prechtl and Hall. First, the double X-frame design uses two X-frames operating in opposition, which allows the actuator stack preloads to be applied internally to the actuator, rather than through the actuation path. Second, the frames of the actuator have been modified to improve the actuator form factor, and increase the volume of active material in the actuator. Testing of the double X-frame piezoelectric actuator was conducted in order to determine its performance (stroke and stiffness) and robustness. In general, stiffness test data compared well with the analytical predictions. The actuator stroke was about 15% less than expected, probably due to the stack output being less in the actuator than as measured in single stack segment testing in the lab. The actuator was also tested dynamically, to determine its frequency response. Actuator robustness was evaluated by measuring its performance when subjected to the effects of blade bending, vibration, and centrifugal loading. Blade elastic bending and torsion deformations were simulated by shimming of the actuator mounts. To assess the impact of the blade vibrations, the actuator and bench test rig were mounted on a hydraulic shaker and subjected to flapwise or chordwise accelerations up to 30 g. To assess the impact of centrifugal force loading, the actuator and bench test rig were spun in the University of Maryland vacuum chamber, so that the actuator was subjected to realistic accelerations, up to 115% of nominal. Results showed that actuator output (force times stroke) was largely unaffected by dynamic and steady accelerations or elastic blade deformations.
We describe structural-acoustic control experiments on a model fuselage test-bed, using collocated pairs of piezoelectric sensors and actuators. The test-bed is a hybrid-scaled model fuselage designed to be representative of complex aircraft structures with rib and stringer construction, which results in a structure with high modal density and complex behavior. The sensor/actuator pairs consist of PVDF film and PZT ceramic sheets bonded to the surface of the model fuselage. Closed- loop control of the fuselage skin was carried out with 30 collocated sensor/actuator pairs, covering approximately 10% of the surface area of the test-bed. The disturbance source is a PZT patch bonded to an adjacent panel. Rate feedback was applied to each collocated pair simultaneously (independent loop closure). Accelerometers attached to the panels and microphones located inside the test-bed were used as performance sensors. The experimental results show a reduction of as much as 20 dB in structural acceleration and up to 10 dB of attenuation in the interior acoustic pressure levels at resonant peaks, over the frequency range of 100 - 2000 Hz.
KEYWORDS: Analog electronics, Linear filtering, Feedback control, Signal attenuation, Digital filtering, Acoustics, Control systems, Composites, Digital signal processing, Actuators
We consider the problem of reducing the noise radiation from a thick-walled cylindrical shell by actively controlling the motion of the shell's outer surface. Because the shell is very stiff, it is difficult to directly control the shell deflections. Instead, the proposed approach is to cover the shell's outer surface with curved active composite panels. Each panel contains several embedded accelerometers mounted to its outer and inner surfaces, which can sense both the motion of the panel base (i.e., the outer motion of the shell) and the outer surface of the panel (i.e., the radiating surface). The accelerometers are used in both feedback and feedforward architectures, in which the accelerometer signals are used to command the panel displacement, in order to reduce the motion of the panel outer surface, reducing the radiated noise. Experimental results show that, in the best case, 10 - 30 dB of surface vibration reduction can be achieved in the frequency range of interest, which is 250 - 2000 Hz.
The results of preliminary tests on an active helicopter rotor blade are presented. The blade, a Mach-scaled model of a CH-47D helicopter blade, has a discrete piezoelectric actuator embedded within the spar that controls a trailing edge flap via a pushrod. Ultimately, the blade will be tested on a helicopter rotor hover stand at MIT. In this paper, we describe the tests performed prior to hover testing. First, the actuator was tested on the bench to determine its control authority and frequency response. Second, the actuator was tested on a shake table to simulate the out-of-plane accelerations that would be encountered in a full-scale helicopter in forward flight. Third, the actuator was embedded in the model blade, and its response to low-frequency sinusoidal actuation was obtained and compared to the bench test results. Finally, the frequency response of the actuator in the blade was determined using swept sine excitation. All test results indicate that the actuator should produce the desired level of control authority in the model-scale rotor.
A design is presented of a 1/6 Mach scaled CH-47D rotor blade incorporating a X-Frame discrete actuator for control of a trailing edge servo-flap. The second generation design of the X-Frame actuator is described focusing on the design changes made from the actuator prototype. The function of the components that restrain the actuator to the rotor blade and connect it to the servo-flap are described. The major challenge in placing a discrete actuator into a rotor blade is in allowing the required functionality in the aggressive acceleration environment of the blade. In particular, a new centrifugal flexure is used to restrain the actuator in the spanwise direction and special fittings are incorporated into the blades to allow the required actuator degrees of freedom while reacting the out of plane vibrational accelerations of the blade. Concentric steel rods are used to transfer actuator motion to the servo-flap and to eliminate the compliant blade fairing from the actuation load path. A slotted flap design was used to reduce the required hinge moments. The aerodynamic implications of using such a flap design are described. Furthermore, retention of the flap and the pre-stress of the actuator were accomplished by a steel wire centered on the flap rotational axis. The design of this part and its influence on choosing an optimum flap length is discussed. The manufacture of the composite rotor blades is described. The diversion of composite unidirectional plies to allow access to the actuator bay within the blade spar is described.
Discrete trailing edge servo-flap actuator designs for use in rotor control applications are considered. A survey was conducted comparing the pros and cons of a number of feasible actuator designs. The major conclusions from this survey indicate that any successful actuator design will utilize a high bandwidth active material, produce large amplification of the active material stroke, and incorporate a simple compressive pre-stress mechanism while remaining efficient in a mass normalized sense. The mass efficiency, defined as the ratio of the specific work performed to the specific energy available, was used as a metric to rate the actuators considered in this survey. Thus, unnecessarily heavy actuators are penalized, which is appropriate when designing components operating under high centrifugal forces. The most feasible discrete actuators are those where the active material reacts against an inert support frame housing. An upper bound on the mass efficiency of this type of actuator is shown to be a function of the ratio of active material to frame specific moduli. A new high efficiency discrete actuator called the x- frame actuator, developed at MIT, and designed in accordance with the lessons learned from the actuator survey, is described. A prototype of this actuator, 150% of model scale, was built and tested on the bench top to confirm the predicted performance. The prototype demonstrates an output energy density of 14.6 ft-lb/slug. It also has a bandwidth of 543 Hz upon driving a nearly impedance matched load. This performance is shown to correspond to a mass efficiency between 18 and 31%.
KEYWORDS: Acoustics, Chemical elements, Finite element methods, Skin, Aluminum, Aircraft structures, Systems modeling, Control systems, Spherical lenses, Computer aided design
A cylindrical test-bed has been designed and modeled to aid in the study and control of interior acoustics in aircraft. The test-bed accounts for local as well as global structural- acoustic dynamics encountered in typical aircraft. The design is based on several existing aircraft and models used to study aircraft dynamics. The test-bed incorporates the basic geometry and materials common to a majority of aircraft, including an aluminum skin shrouding a framed structure composed of ribs and stringers. The design is approximately a one-third scale representation with a cylindrical geometry measuring 0.91 m diameter and 1.98 m long. The test-bed has been modeled using finite element method and Rayleigh-Ritz assumed modes analysis. The models were used to refine the design of the test-bed as well as to model the coupled structural-acoustic dynamics. The models predict that the test-bed will have a modal structure commensurate with experimental identifications on existing aircraft.
A servoflap that uses a piezoelectric bender to deflect a trailing edge flap for use on a helicopter rotor blade was designed, built, and tested. This servoflap design is an improvement over a design developed previously at MIT. The design utilizes a new flexure mechanism to connect the piezoelectric bender to the control surface. The efficiency of the bender was improved by tapering its thickness with length. Also, the authority of the actuator was increased by implementing a nonlinear circuit to control the applied electric field, allowing a greater range of actuator voltages. Experiments were carried out on a bench test article to determine the frequency response of the actuator, as well as hinge moment and displacement capabilities. Flap detections of 11.5 deg or more were demonstrated while operating under no load conditions at frequencies up to 100 Hz. The data indicate that if properly scaled, the actuator will produce flap deflections greater than 5 deg at the 90% span location on a full- scale helicopter. In addition, the first mode of the actuator was at 7/rev frequency of the target model rotor. Proper inertial scaling of this actuator could raise this modal frequency to greater than 10/rev on an operational helicopter, which is adequate for most rotor control purposes. A linear state space model of the actuator was derived. Comparisons of this model with the experimental data highlighted a number of mild nonlinearities in the actuator's response. However, the agreement between the experiment and analysis indicate that the model is a valid tool for predicting actuator performance.
The progress to date and essential features of approaches to optics and controlled structures technology (CST) are reviewed. The CST framework is suggested as a means of gaining new insight on deformable optical surface control. Within the CST approach, control bandwidths may extend into the frequency range of structural vibration modes, enabling possible reduction in area density and improvements in performance.
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