Requirements of future military aircraft structures are constantly increasing with advancing technological progress. While performance is still the main focus, costs have become a major issue in military aircraft procurement.In order to efficiently support its technological base oriented on the future demands of the market Daimler Chrysler Aerospace/Military Aircraft Division has inaugurated the Advanced Aircraft Structures Program, a collaborative research effort together with the German Aerospace Center and Daimler Chrysler Research and Technology, the corporate research division of Daimler Benz. The two key technologies to be pursued within the framework of this program are cost- effective composite structures and smart materials. This paper will give an overview of the Advanced Aircraft Structures Program with particular emphasis on smart structures technology as applied to active vibration damping, vibration isolation of equipment and composite health monitoring.

Structural fin or wing vibrations are observed on high performance aircraft when flying at high angles of attach. The severity of the so-called buffeting vibration depends on the aircraft configuration and aerodynamic optimization of the configuration. The vibrations are caused by flow fluctuations resulting from flow separation at wings or from bursting of wing leading edge and front fuselage vortices. The resulting dynamic loads with maneuver loads lead to increased material fatigue and may require an augmented effort in aircraft maintenance.

One of the most innovative concepts for active fin-buffet alleviation in vertical tail aircraft is the use of piezoelectric patch actuators distributed across the tail surface to actively induce a counter-strain into the structure. This concept involves the development of a novel material compound structure consisting of a fiber-composite aircraft skin, a ceramic patch actuator and the bonding layer between both components. This actively controllable structure has to provide enough authority to dampen the fin- buffet vibrations. It also has to function reliably during long-term aircraft operation under severe mechanical and environmental load conditions.

Modern military aircraft are characterized by employment of optimized structural components. New demands on exploitation of lightweight construction technology will arise because even greater flexibility with increased maneuverability is desired. The structural integration of multifunctional, often called 'smart' elements, properly activated to e.g. reduce structural loading, offers great potential to necessary advances in military aircraft design. One major problem of modern military aircraft is the buffet loading on the fin structures. Flying the aircraft at high angles of attack allows vortices, evolving from the leading edge of the wing, to hit the fin and excite structural vibrations. This leads to structural attrition as well as a reduced aircraft maneuverability. With the aim to reduce these fin vibrations, an adaptive structure has been developed which is presented in this paper. A concept is discussed with which the vibrational loads are reduced by introduction of counteracting forces using an 'active interface'. This interface concept is characterized by the integration of active, piezoelectric elements directly into the bending support of the fin structure. To validate the stability of the interface FE calculations and extensive measurements on piezoceramic stack actuators have been performed. The manufactured interface was integrate in an existing test structure and realistically loaded. The result will be given in this presentation.

Introduction of new technologies to aerospace applications necessarily requires methods of non-destructive testing suitable to evaluate structural integrity. This important task also occurred when it was decided to develop and manufacture a large Fin-Box-Demonstrator equivalent to a fighter aircraft tail equipped with surface bonded piezoceramic actuators between DaimlerChrysler Aerospace - Military Aircraft Division and DaimlerChrysler Research and Technology. The objective of this project is to prove that structural vibrations of a fighter aircraft tail fin due to buffeting can be damped actively by means of surface bonded piezoceramic actuators.

The DLR Institute of Structural Mechanics is engaged in the construction and optimization of adaptive structures for aerospace and terrestrial applications. Due to the FFS- Project, one of the recent works of the Institute is the reduction of buffet induced vibration loads at a fin. The construction of modern aircrafts is influenced b the increasing use of fiber composites. They have more specific stiffness and strength properties than metals. On the other hand the layered structure leads to new kinds of damages like delaminations. In the fin interface there are actuators and sensors integrated. Therefore the fin is connected with a controller. For the extension of this adaptive system towards an on-line tool for health monitoring this controller can be used as an identifier of the structure's modal parameters. The most promising procedure is based on MX filters. These filters constitute the filter coefficients from which a fast transformation procedure extracts the modal parameters. The changes of these parameters are related to the location and extent of the damage. So when using the already integrate controller for system identification, one can have a low-cost on-line damage detection for dynamic adaptive structures. First off-line test at CFRP plates have shown the ability to detect delaminations.

Fiber optic sensors are being developed for health monitoring of future aircraft. Aircraft health monitoring involves the use of strain, temperature, vibration and chemical sensors. These sensors will measure load and vibration signatures that will be used to infer structural integrity. Since the aircraft morphing program assumes that future aircraft will be aerodynamically reconfigurable there is also a requirement for pressure, flow and shape sensors. In some cases a single fiber may be used for measuring several different parameters.

This experimental investigation focuses on the application of piezoelectric sensors/actuators for wing flutter and vertical tail buffet suppression. The test article consists of a foam airfoil shell enveloped around an aluminum plate support structure with bonded piezoelectric actuators and sensors. Wind-tunnel test results for the wind are presented for the open- and closed-loop systems. Piezoelectric actuators were effective in suppressing flutter and the wake-induced buffet vibration over the range of parameters investigated.

This study, started in late 1997, evaluates the concept of tab-assisted control (TAC), and the use of shape memory alloy (SMA) actuator in that connection. Under the TAC concept, a small tab, typically 10 percent of the mean chord of the entire control surface structure, is appended to the trailing edge of the primary control surface, or flap. This small tab vastly enhances the versatility of the control surface system. Depending on the orientation of the tab with respect to the flap and the amount of tab deflection, this tab may be used to modify lift and torque, actuate the flap, or provide precision control; if the tab is aligned with the flap, TAC reverts itself to the conventional configuration. Despite its many benefits, TAC faces one practical challenge in implementation. Due to the particular TAC configuration, the actuating system for the tab must be compact enough to fit in the limited real estate available within the flap. This makes SMA actuator a promising contender for TAC implementation. This paper presents some of the experimental result relevant to the design of the SMA actuator and addresses implementation issues such as power usage, life cycle, frequency response, and reliability.

This paper describes recent test result obtained on a prototype SMA-actuated foil that serves as a key element in a vortex wake control scheme for lifting surfaces. Previous papers have described the theoretical basis and feasibility studies for this scheme - which is based on a novel wake control known as vortex leveraging - as well as prior work on device design, test planning, and fabrication. The critical item in the realization of this scheme is a Smart Vortex Leveraging Tab (SVLT), a device designed to provide perturbations in the vortex system downstream of lifting surfaces at frequencies and amplitudes carefully selected to accelerate overall wake breakup. The paper summarizes the background of the effort, but focuses on the detail design and fabrication techniques used in the construction of a prototype SVLT and the results of water tunnel tests of a near full-scale prototype device.

A 2D spatially distributed smart skin sensor for real-time seat occupant position sensing is presented. The sensor exploits principles of spatial aperture shading of distributed transducers such as piezo-electric polymers and resistors, which are used as the active sensing medium. An example application is presented in which the sensor is used to report passenger position to an automobile air bag control system. The real-time data is used to modulate airbag deployment energies, mitigating passenger injury.

This paper discusses the conceptual design of a family of specially-designed temperature surety sensors made with shape-memory alloys (SMA). These sensors are capable of detecting a one time temperature excursion or variance form a predetermined temperature range. The propose designs can also be used to detect a one-time temperature rise and persistence above a certain pre-selected critical temperature. In that respect, these sensors relate to a family of one-time thaw sensors detecting whether or not frozen food items or other frozen products or objects experience a thawing-refreezing process in their journey from point A to point B. The proposed sensor can also detect a one time temperature excursion into non-allowable temperatures for non-frozen food, as well as pharmaceutical or other medical products. The essential design of these smart sensor is a lever arm attached to an SMA wire whose temperature is initially below Austenite start temperature or well into the Martensite region. As a given product experiences an undesirable temperature range which pushes the SMA material into the Austenite region the wire contracts and moves the lever arm outside a display window area and exposes either a red working indicator or a graduated scale calibrated to the range of temperature excursion experienced by the product. The sensor is designed such that if the temperature returns to normal the excursion indication will not disappear, but will permanently shown the amount of excursion above the temperature surety region for that product. Several possible design variations are presented and discussed. The proposed embodiments include a rupture type thaw sensor made with short SMA springs or bellows, SMA foil roll-up type sensors, SMA wire-loaded shutter type thaw sensors, SMA torsion strut-loaded shutter type thaw sensors, multiple shutter SMA wire-loaded thaw sensors, multiple shutter, SMA torsion-rod-loaded thaw sensors, and rupture-Type SMA spring-loaded thaw sensors.

A feasibility study in the use of induced strain actuators for active seal isolation is described. The focus of the work is the isolation of lightweight automotive seats for hybrid-electric vehicles. The feasibility study is based on a numerical analysis of a three-degree-of-freedom vibration model of the seat. Mass and inertia properties are based on measurements from a powered seat that is found in current model year automobiles. Tradeoffs between vertical acceleration of the seat, actuator stroke requirements, and isolation frequency are determined through numerical analysis of the vibration model. Root mean square accelerations and actuator strokes are computed using power spectral densities that model broadband excitation and road excitation that is filtered by the vehicle suspension. Numerical results using the road excitation indicate that factors of two to three reduction in vertical acceleration are achieved when the active isolation frequency is reduced to approximately 1 Hz with damping factors on the order of 10 to 30 percent critical. More significant reductions are achieved in the case of broadband floor excitation. Root mean square actuator strokes for both case are int he range of 0.4 to 50 mm. Root mean square accelerations in the vertical direction are consistent with the levels found in standard comfort curves.

Fiber optic sensors have the potential to be used in the very hostile environments necessary for advanced aerospace platforms. This paper reviews some of the key issues associated with the implementation of distributed fiber optic sensors in harsh environments and outlines baseline system designs.

This paper considers a wireless approach to health and usage monitoring of advanced composite materials. It uses a modular approach supporting, in particular, high performance system fabricated with composite panels fastened to an underlying frame. The panels are fabricated so as to include integrated circuit (IC) sensors placed in the various prepreg layers during layup. Various aspects in the development and performance of these IC sensors are described in the paper. Since the embedded sensors do not include an internal power source, interrogator circuitry has also been developed. This unit supplies power to sensor within the composite panel and obtains transducer data from the sensors, both via wireless techniques. This unit and its interface to the embedded sensors are discussed. Finally, a description is given of expected overall system monitoring using the sensor and interrogator devices being presented.

This paper describes the development and testing of a 4KVpp, 750 ma piezo drive switching amplifier. This amplifier is used to drive Piezo Fiber Composite material imbedded in a 1/6 scale CH-47 blade. This amplifier will allow higher harmonic control of the blade thus reducing rotor craft vibration and nose. The amplifier recycles reactive energy required to drive piezo material allowing for an efficient amplifier design. A multi level topology is used allowing solid sates switching devices with voltage rating of half the output drive voltage. The amplifier modular design allows easy migration to the power levels required to drive a full size CH-47 blade. This work was done in conjunction with the Smart Structures for Rotor Craft Control in support of DARPA/ONR.

The application of smart structures to helicopter rotors has received widespread study in recent years. This is one of the major thrusts of the Shape Memory Alloy Consortium (SMAC) program. SMAC includes 3 companies and 4 Universities in a cost sharing consortium funded under DARPA Smart Materials and Structures program. This paper describes the objective of the SMAC effort, and its relationship to a previous DARPA smart structure rotorcraft program from which it originated. The SMAC program includes NiTinol fatigue/characterization studies, SMA actuator development, and ferromagnetic SMA material development. The paper summarizes the SMAC effort, and includes background and details on Boeing's development of a SMA torsional actuator for rotorcraft applications. SMA actuation is used to retwist the rotorcraft blade in flight, and result in a significant payload increase for either helicopters or tiltrotors. This paper is also augmented by several other papers in this conference with specific results from other SMAC consortium members.

Active Fiber Composites (AFCs) provide a novel method for large scale actuation and sensing in active structures. The composite comprises unidirectionally aligned piezoelectric fibers, a resin matrix system, and interdigital electrode. AFCs have demonstrated distinct advantages over current monolithic piezoceramic actuators, including: higher planar actuation strains, tailorable orthotropic actuation, robustness to damage, conformability to curved surfaces, and potential for large area distributed actuation/sensing systems. The Active Fiber Composite Consortium (AFCC) consists of six organizations: piezoelectric material supplier, composite manufacturers, device packages, technology development, and several end-user application groups. The objective of the AFCC is to develop and demonstrate AFC materials for insertion into both military and commercial applications. These objectives are being achieved through 1) a consortium program in low cost manufacturing using innovative manufacturing techniques and economy of scale, 2) targeting specific military and commercial applications which could benefit substantially from successful AFC integration. This manuscript highlights recent activity in the AFCC, describing approaches to manufacturing of the fiber and composite system, advancements in the materials and characterization for improved performance, and an overview of the applications under development.

Several analytical and experimental studies clearly demonstrate that piezoelectric materials can be used as actuators to actively control vibratory response, including aeroelastic response. However, two important issues in using piezoelectrics as actuators for active control are: 1) the potentially large amount of power required to operate the actuators, and 2) the complexities involved with active control. Active or passive damping augmentation using shunted piezoelectrics may provide a viable alternative. This approach requires only simple electrical circuitry and very little or no electrical power. The current study examines the feasibility of using shunted piezoelectrics to reduce aeroelastic response using a typical-section representation of a wing and piezoelectrics shunted with a parallel resistor and inductor. The aeroelastic analysis shows that shunted piezoelectrics can effectively reduce aeroelastic response below flutter and may provide a simple, low-power method of subcritical aeroelastic control.

The objective of this paper is to describe the creation, modeling and testing of an active diaphragm flexure. The flexure is a circular stainless steel plate with three sets of spatial concentric slits. Such a flexure is commonly used as a passive centering spring but may be made active with the addition of piezoceramics. By adding piezoceramics as bi-morph pairs, out of plane motion is induced in the flexure. This active flexure is modeled using finite elements to determine the flexure's stiffness and natural frequency. A simplified analytical model is also developed. From this simple model, design equations for stiffness and frequency can be determined. By changing flexure properties, such as thickness, ring width or plate materials; approximate values of stiffness and natural frequency can be determined. For this paper only variations in the thickness of the flexure are examined. A comparison between experimentally obtained results and both the finite element and analytical models is presented. Results show that both numerical and analytical models can be used as predictive tools to determine flexure properties.

The objective of this research is to design and develop a working prototype of a linear traveling wave motor that utilizes Thin-layer composite-Unimorph piezoelectric Driver (THUNDER) technology. THUNDER technology is used to create curved actuators from piezoceramic wafers. These flexures are arranged alternatively bowed towards and away from a flat surface. The basis of motion rest on the fact that upon actuation the flexures will change both chord length and radius of curvature. This allows the flexures to either grip the driving rod or extend axially. Motion is achieved through sequential operation of individual flexures. This method yields a lightweight actuator with power off holding capability and a larger step size than motors that use stacked actuators. The simple design lends itself to inexpensive fabrication. A finite element model was used to predict the initial curvature and the height displacement of a single flexure. The model takes into account material properties, physical layout, fabrication techniques and driving voltage. These theoretical predictions were compared to experimental results. A prototype was developed but no movement has been realized in this configuration. However, it is shown that the flexures have the capability to achieve step sizes 1-2 orders of magnitude greater than other similar linear actuators.

A family of high authority actuators was developed at the Naval Research Laboratory. These actuators are based on displacement amplification within a compact, solid state, monolithic piezoelectric actuator, by using a telescoping tube design. In this design, concentric tubes are mechanically connected in series. This gives an effective actuator length that is equal to the sum of the lengths of the individual concentric elements. The high displacement output of this actuator permits efficient coupling of the actuator output into a load of similar impedance, and thereby much greater effective actuator output. Initial prototypes were made of commercially available PZT tubes of three different diameters and wall thickness. These tubes were pulsed through the thickness of the walls and the change in their lengths were used for actuation. Their actuation is therefore making use of the d₃₁ piezoelectric coefficient. Alternatively, electrodes can be applied to the ends of the individual concentric tubes and their lengthwise displacement will subsequently be proportional to the d₃₃ parameter of the material. The tubes were bonded at their ends to alumina plates using epoxy-based adhesive. The displacement obtained from the assembly is close to the sum of those of the three individual tubes at the same applied field. Other parameters such as blocking force and energy densities are also reported. These actuators have applications where high force and simultaneously large displacement are required and space is limited. Potential uses include high end aerospace as well as low tech commercial applications.

Cryogenic magnetostrictive materials, such as rare earth zinc crystals, offer high strains and high forces with minimally applied magnetic fields, making the materials ideally suited for deformable optics applications. For cryogenic temperature applications, such as Next Generation Space Telescope, the use of superconducting magnets offer the possibility of a persistent mode of operation, i.e., the magnetostrictive material will maintain a strain field without power. High temperature superconductors (HTS) are attractive options if the temperature of operation is higher than 10 degrees Kelvin (K) and below 77 K. However, HTS wires have constraints that limit the minimum radius of winding, and even if good wires can be produced, the technology for joining superconducting wires does not exist. In this paper, the design and capabilities of a rare earth zinc magnetostrictive actuator using bulk HTS is described. Bulk superconductors can be fabricated in the sizes required with excellent superconducting properties. Equivalent permanent magnets, made with this inexpensive material, are persistent, do not require a persistent switch as in HTS wires, and can be made very small. These devices are charged using a technique which is similar to the one used for charging permanent magnets, e.g., by driving them into saturation. A small normal conducting coil can be used for charging or discharging. Very fast charging and discharging of HTS tubes, as short as 100 microseconds, has been demonstrated. Because of the magnetic field capability of the superconductor material, a very small amount of superconducting magnet materials is needed to actuate the rare earth zinc. In this paper, several designs of actuators using YBCO and BSCCO 2212 superconducting materials are presented. Designs that include magnetic shielding to prevent interaction between adjacent actuators will also be described. Preliminary experimental results and comparison with theory for BSSCO 2212 with a magnetostrictive element will be discussed.

The concept of using piezoelectric actuators in devices that alter the way in which an airfoil interacts with its environment is not new. In fact, several notable research institutions, federal laboratories and industrial partners are actively pursuing this type of research. The main driving force for this activity is increased fuel economy, lighter aircraft and the elimination of hydraulically actuated control surfaces. Several years ago, researchers at NASA developed a process that uniformly prestressed the piezoelectric actuators resulting increased movement at low frequencies. The key to this increased motion was the ability to develop an evenly prestressed actuator that behaved like a leaf spring. In order to take full advantage of this piezoelectric wafer, the fixturing and drive electronics had to be developed. This is a critical issue for all piezoelectric systems. This paper describes the characteristics and performances of these high displacement actuators and the devices that incorporate these actuators to create the synthetic jets. It is envisioned that these devices will play a critical role in the future of aeronautics.

The DARPA/AFRL/NASA Smart Wing program, conducted by a team led by Northrop Grumman Corp. under the DARPA Smart Materials and Structures initiative, addresses the development of smart technologies and demonstration of relevant concepts to improve the aerodynamic performance of military aircraft. This paper present an overview of the smart wing program.

To verify the predicted benefits of the smart wing concept, two 16 percent scale wind tunnel models, one conventional and the other incorporating smart wing design features, were designed, fabricated and twice tested at NASA Langley's 16 ft Transonic Dynamic Tunnel, in two series of tests, conducted in May 1996 and June 1998, respectively. A key objective of the Smart Wing Phase 1 program was not only to construct wind tunnel models that could be used to validate the predicted benefits of using smart materials, but also to identify and reduce the risks involved in eventually integrating smart materials into an actual flight vehicle. Among the challenges encountered in developing the wind tunnel model were the attachment of the shape memory alloy (SMA) control surfaces to the wing box, integration of the SMA torque tube in the wing structure, installation of the instrumentation, and development of fail safe control mechanisms to protect the model and the tunnel in the event of failure of the smart systems. In this paper, design and fabrication details of the two Smart Wing Phase 1 wind tunnel models are presented. Among the topics covered are 1) model design requirements, model design and static testing; 2) manufacturing techniques with particular emphasis on the improvement in the design and fabrication of the SMA control surfaces from the first to the second test; 3) system integration; and 4) post-test analysis and planned improvement. Lessons learned from the Phase 1 effort are discussed along with plans for the Smart Wing Phase 2 program.

To quantify the benefits of smart materials and structures adaptive wing technology. Northrop Grumman Corp. built and tested two 16 percent scale wind tunnel models of a fighter/attach aircraft under the DARPA/AFRL/NASA Smart Materials and Structures Development - Smart Wing Phase 1. Performance gains quantified included increased pitching moment, increased rolling moment and improved pressure distribution. The benefits were obtained for hingeless, contoured trailing edge control surfaces with embedded shape memory alloy wires and spanwise wing twist effected by SMA torque tube mechanism, compared to convention hinged control surfaces. This paper presents an overview of the results from the second wind tunnel test performed at the NASA Langley Research Center's 16 ft Transonic Dynamic Tunnel in June 1998. Successful results obtained were: 1) 5 degrees of spanwise twist and 8-12 percent increase in rolling moment utilizing a single SMA torque tube, 2) 12 degrees of deflection, and 10 percent increase in rolling moment due to hingeless, contoured aileron, and 3) demonstration of optical techniques for measuring spanwise twist and deflected shape.

In recent Smart Wing wind tunnel tests at NASA Langley, we demonstrated over 5 percent of span-wise twist at M equals 0.205. This was a considerable improvement over the 1.25 degree of twist demonstrated during the initial tunnel test. Key to the improvements were two developments. First a different torque loading path in the structure, which resulted in torque being directly reacted from root to wing tip. Secondly, a new SMA actuator was developed, with a measured blocking torque of 3500 +/- 100 in-lb. The second round of tunnel test not only demonstrated increased wing twist; we also were able to command a variety of twist angles and were able to show that the wing could maintain a predetermined twist for over an hour with a stability of 0.05 degrees. Power consumption was recorded, with maximum power of 200W during twisting, and a power demand of 20W for maintaining wing twist.

Fiber reinforced composites offer excellent specific stiffness and strength and are therefore interesting for rotating machinery applications. The main disadvantage of high performance composites is the manufacturing process which is labor intensive and thus slow and expensive. The Thermoplastic Fiber Placement process overcomes these difficulties due to its high degree of automation. During the process, an impregnated tape is heated up and then consolidated in-situ under pressure. The process which is used at ABB consists of a six axis robot, a heat source and a pressure device for consolidation. Today mechanical roller element are used to apply the forces normal to the surface to the composite part. These forces are necessary for proper consolidation. The roller action prevents damage due to shearing of the tape during lay down. To improve the processing sped, and to expand the use of the Thermoplastic Fiber Placement process for more complex structures, two severe drawbacks of the solid roller approach need to be overcome; the small pressure contact area which limits the speed of the process and the poor conformability which prevents the process from being applied to highly 3D surfaces. Smart materials such as piezoelectrics, electrostrictives and magnetostrictives can produce high forces at high operating frequencies and enable a large, conformable actuated surface to be realized. A pressure device made with a magnetostrictive actuator has been tested. The main design goal is to apply the consolidation pressure correctly, without introducing shear forces on the tape, in order to produce parts with optimal mechanical properties.

Many manufacturing situations such as lithography for microelectronics require rapid high precision positioning of machinery. This paper will describe an existing stage system and its evolution with an emphasis upon reducing residual vibrations associated with stage stepping. Results of an experimental study are presented where the adaptive input shaping strategy and several input shaping strategy technique for the literature are used to generate motion commands. Stage performance is evaluated with an emphasis upon the affects of the stages natural frequency and mode shapes.

Spacecraft carry instruments and sensors that gather information from distant points, for example, from the Earth's surface several hundred kilometers away. Small vibrations on the spacecraft can reduce instrument effectiveness significantly. Vibration isolation system are one means of minimizing the jitter of sensitive instruments. This paper describes one such system, the Satellite Ultraquiet Isolation Technology Experiment (SUITE). SUITE is a piezoelectric-based technology demonstration scheduled to fly in 2000 on PICOSat, a microsatellite fabricated by Surrey Satellite Technology, Ltd. Control from the ground station is planned for the first year after launch. SUITE draws on technology from previous research programs as well as a commercial piezoelectric vibration isolation system. The paper details the features of SUITE, with particular emphasis on the active hexapod assembly. A description of the PICOSat spacecraft and the other considerations preceding the development of the flight hardware begins the paper. Experimental goals are listed. The mechanical and electromechanical construction of the SUITE hexapod assembly is described, including the piezoelectric actuators, motion sensors, and electromagnetic actuators. The data control system is also described briefly, including the digital signal processor and spacecraft communication. The main features of the software used for real-time control and the supporting Matlab software used for control system development and data processing are summarized.

Modern satellites require the ability to slew and settle quickly in order to acquire or transmit data efficiently. Solar arrays and communication antennas cause low frequency disturbances to the satellite bus during these maneuvers causing undesirable induced vibration of the payload. The ability to develop and experimentally demonstrate attitude control laws which compensate for these flexible body disturbances is of prime importance to modern day satellite manufacturers. Honeywell has designed and fabricated an actively controlled Appendage Simulator Unit (ASU) which can physically induce the modal characteristics of satellite appendages on to a ground based satellite test bed installed on an air bearing. The ASU consists of two orthogonal fulcrum beams weighting over 800 pounds each utilizing two electrodynamic shakers to induce active torques onto the bus. The ASU is programmed with the state space characteristics of the desired appendage and responds in real time to the bus motion to generate realistic disturbances back onto the satellite. Two LVDT's are used on each fulcrum beam to close the loop and insure the system responds in real time the same way a real solar array would on-orbit. Each axis is independently programmable in order to simulate various orientations or modal contributions from an appendage. The design process for the ASU involved the optimization of sensors, actuators, control authority, weight, power and functionality. The smart structure system design process and experimental results are described in detail.

This paper is concerned with the viscoelastic strain-energy hinge for solar array deployment. The original strain-energy hinge proposed by TRW for solar array deployment was made of strip measures. Due to its structural simplicity, the strain-energy hinge has been considered as an alternative to the torsional spring type deployment mechanism. However, theoretical modeling of the strain-energy hinge is extremely difficult because of its nonlinear pre- and post-buckling dynamic behavior. To investigate its dynamic characteristics, series of buckling and deployment tests on a single strain-energy hinge and a solar array structure equipped with strain-energy hinge have been conducted. The deployment test results show that there remain residual vibrations after deployment, which are resulted from the rapid deployment and the bending flexibility of the strain- energy hinge. We propose the use of viscoelastic material embedded between the layers of the strip measure to increase the passive damping. It results in less residual vibrations and smooth deployment. Experimental results show that viscoelastic strain-energy hinge ins superior to the ordinary strain-energy hinge in deployments. Based on the experiments on the single strain-energy hinge, an equivalent 1D torsional spring model is proposed. Simulation results based on the equivalent model are fairly in good agreement with experimental results.

An innovative new hybrid isolation system has been developed to significantly increase the performance over a passive whole-spacecraft isolation design. The design builds upon the passive design and incorporates active components in parallel to the passive design. This means that if the active system fails, the passive system would be able to handle the isolation requirements. Preliminary results show that significant attenuation can occur using the hybrid isolation system over the passive isolation system. Also, it has been determined that the performance gained by the hybrid isolation system will be dependent on the stiffness of the launch vehicle. As this stiffness now becomes an important design parameter when developing a whole- spacecraft launch isolation system.

As commercial and government space applications shift to smaller satellites there is an increased need to control the environmental loads the satellite must experience during launch. Recently, a passive vibration isolation system was successfully deployed. Seeing to enhance the passive isolator's effects research has shifted to active isolation techniques. This paper summarizes the results of recent research to develop a launch vehicle flight model and suitable environmental conditions to analyze the effectiveness of various passive and hybrid isolation techniques. Results include comparison of the isolator's performance.

This paper describes the fabrication, testing, and analysis of a single axis piezoceramic gimbal. The fabrication process consists of pre-stressing a piezoceramic wafer using a high-temperature thermoplastic polyimide and a metal foil. The differential thermal expansion between the ceramic and metal induces a curvature. The pre-stressed, curved piezoceramic is mounted on a support mechanism and a mirror is attached to the piezoceramic. A plot of gimbal angle versus applied voltage to the piezoceramic is presented. A finite element analysis of the piezoceramic gimbal is described. The predicted gimbal angle versus applied voltage is compared to experimental results.

The SAMPSON program will demonstrate the application of Smart Materials and Structures to large-scale aircraft and marine propulsion systems and show that smart materials can be used to significantly enhance vehicle performance, thereby enabling new missions and/or expanding current missions. Two demonstrations will be executed in relevant environments and at scales representations of actual vehicle components. The demonstrations will serve to directly address questions of scalability and technology readiness, thereby improving the opportunities and reducing the risk for transitioning the technology into applications. The aircraft application to be examined is the in-flight structural variation of a fighter engine inlet. Smart technologies will be utilized to actively deform the inlet into predetermined configurations to improve the performance of the inlet at all flight conditions. The inlet configurations to be investigated consists of capture area control, compression ramp generation, leading edge blunting, and porosity control. The operation and demonstration of this Smart Inlet is described in detail.

The Smart Aircraft and Marine Propulsion System Demonstration (SAMPSON) Program will culminate in two separate demonstrations of the application of Smart Materials and Structures technology. One demonstration will be for an aircraft application and the other for marine vehicles. The aircraft portion of the program will examine the application of smart materials to aircraft engine inlets which will deform the inlet in-flight in order to regulate the airflow rate into the engine. Continuous Moldline Technology (CMT), a load-bearing reinforced elastomer, will enable the use of smart materials in this application. The capabilities of CMT to withstand high-pressure subsonic and supersonic flows were tested in a sub-scale mini wind- tunnel. The fixture, used as the wind-tunnel test section, was designed to withstand pressure up to 100 psi. The top and bottom walls were 1-inch thick aluminum and the side walls were 1-inch thick LEXAN. High-pressure flow was introduced from the Boeing St. Louis poly-sonic wind tunnel supply line. CMT walls, mounted conformal to the upper and lower surfaces, were deflected inward to obtain a converging-diverging nozzle. The CMT walls were instrumented for vibration and deflection response. Schlieren photography was used to establish shock wave motion. Static pressure taps, embedded within one of the LEXAN walls, monitored pressure variation in the mini-wind tunnel. High mass flow in the exit region. This test documented the response of CMT technology in the presence of high subsonic flow and provided data to be used in the design of the SAMPSON Smart Inlet.

An investigation into the interaction of a synthetic jet actuator array with a thick, turbulent boundary layer was conducted and the results are presented here. This paper documents an experimental effort that was directed towards improving the understanding of synthetic jet actuator arrays and the mechanism by which these actuators control flows at scales much larger than themselves. Of specific interest in the interaction of a synthetic jet flow and the boundary layer flow on a flat or curved wall where the axis of the synthetic jet is perpendicular to the direction of the oncoming flow. The investigation addressed, in broad terms, a typical control application involving internal flows. A synthesis of the results will provide guidance and direction in future actuator designs and implementation strategies, and feed design input to test plans for an extensive study of actuator control effects in a duct with variable pressure gradient and variable streamline curvature.

Documented herein is a review of progress for the recently completed 'Smart Skin Structure Technology Demonstration' (S³TD) contract number F33615-93-C-3200 performed by Northrop Grumman Corporation, Hawthorne, California and TRW/ASD, Rancho Bernardo, San Diego, California under the Air Force Research Laboratory, Flight Dynamics Directorate, Structures Division's direction and sponsorship. S³TD was conceived as the first serious attempt, to made a complex antenna become a bone fide aircraft structural panel, without loss of overall structural integrity or electrical performance. The program successfully demonstrated the design, fabrication, and structural validation of a load bearing multifunction antenna component panel subjected to realistic aircraft flight load conditions. The final demonstration article was a structurally effective 36 by 36 inch curved multifunction antenna component panel that withstood running loads of 4,000 pounds per inch, and principal strain levels of 4,700 microstrain. Testing the structural component to ultimate, the panel failed at the predicted limit of 148 kips equating to 150 percent design limit load, after successfully completing one lifetime of fatigue. The load conditions were representative of a mid-fuselage F-18 class fighter component panel installation. The panel was designed not to buckle at ultimate failure, and the dominant failure mode was face sheet pull off, as predicted. Structural test data correlated closely with analysis. Wide band electrical performance for the component antenna panel was validated using anechoic chamber measurements and near field probing techniques, covering avionics communication navigation and identification and electronic warfare functions in the 0.15 to 2.2 GHz frequency regimes.

Turbomachinery such as fans, compressors and turbines are usually designed and optimized for one operating point. The efficiency of such a machine significantly decreases if it operates at off-design conditions. Therefore the efficiency could be increased if the flow channel geometry can be adapted to the varying flow conditions such as flow sped and mass flow.

Fiber reinforced composites offer excellent specific stiffness and strength and are therefore interesting for rotating machinery applications. Drive shafts are used to transfer motor torques in aircraft, vehicles or production lines. With the use of composite materials these shafts could be made lighter, with tailored mechanical properties, lower inertia and a benign failure behavior.

Windmills are large composite structures, usually located at difficult-to-access sites, bearing strong dynamic loads in a harsh environment. In-service inspection involves dismounting the blades, a very costly process. Instead of inspecting the windmills regularly, they are usually overdesigned. Furthermore, the design of a windmill is based on estimated wind conditions, while the real wind is seldom measured. The main inconvenience for it are the characteristics of the conventional equipment needed to carry out the measurements. The substitution of traditional sensors for Bragg gratings and piezoelectronics to measure the strain field and vibrational frequencies adds the advantages of a smaller size, no drift, no EMI and the possibility of embedding them while manufacturing, so the windmill is fully equipped when installed. Long-term measurements are possible to check both the in-service conditions and the degradation of the structure. Two possibilities were tested: embedding of the sensors while manufacturing with low disturbance of the process and surface-bonding of the sensors. Windmill qualification test were carried out to check the survivability of the Bragg gratings and the piezoelectric sensors under extreme environment are presently running. Results are hopeful. The next step could be a permanent connection via modem to a remote controller that can use the acquired data to map the wind conditions and/or the structural health.

The blades of helicopter rotors and turbine blades can be modeled as pretwisted beams/plates or shells. The angle of pretwist can affect the performance of these turbo-machinery systems. Being able to actively change the angle of pretwist even moderately can enhance or optimize these systems for various modes of operation. To actively change the angle of pretwist, a torque has to be introduced to the pretwist member that will cause it to either increase or decrease the existing pretwist angle of the blade. By building piezoelectric layers into the blade/plate and applying a controlled voltage the shape and or position of the cross- section can be changed. Due to the coupling of both bending modes and the extensional and torsional modes in pretwisted members a piezoelectric materials that can induce extension in the plate can also be used to twist or bend the shape of the plate. In this study a cantilevered pretwisted plates bonded to two piezoelectric layers on the outside is modeled using 3D linear elastic finite element approach with the pretwist built into the formulation. The static response of the pretwisted plate to a uniform voltage applied to the piezoelectric layers is investigated. Twisting and bending of the plate is accomplished through coupling of the bending modes and extensional-torsional coupling. The piezoelectric layer itself is not isotropic and so introduces additional coupling into the system.

Fatigue cracking and plastic deformation occur in both new and aging aircraft engines, requiring periodic manual inspection with ultrasonic and eddy current probes. Often cracking occurs in inaccessible areas that require engine disassembly just to perform the inspection. In many cases, the cost of assembling and reassembling the engine far exceeds the cost of inspection and can also induce new damage.

Smart materials based on 1-3 piezocomposite transducers, capable of both sensing and actuation, are being developed for active control applications. Large area, low profile SmartPanels, consisting of 1-3 piezocomposite actuators and pressure sensors and net-shape-molded PZT accelerometers, have been fabricated and evaluated for surface mounted boundary control applications. Single layer and two-layer 100 X 100 mm and 250 X 250 mm SmartPanels have been tested for actuator authority, surface displacement uniformity, sensor-actuator coupling, and surface vibration reduction. Single layer SmartPanels have shown broad band 20 dB surface vibration reduction.

A method for estimating the actuation efficiency of a structurally integrated active material is presented. A background literature search revealed many different expressions for efficiency depending upon the application and discipline of interest. Following the review of the literature, an efficiency expression was developed for a piezoelectric actuator in the frequency domain. The actuation efficiency of the piezoceramic actuator was define as the ratio of total mechanical energy imparted to the structure to total electrical energy drawn by the piezoceramic from an electrical power source. The efficiency expression is a function of the piezo electromechanical coupling coefficient, the mechanical impedance ratio of the structural to the piezoceramic material, and the frequency of operation. The developed expression was then used to analytically predict the efficiency of a single degree-of- freedom system actuated by a piezoceramic actuator. Static and quasi-static efficiency results agree with analytical results found in the literature. Dynamic analysis of the efficiency expression, however, produced unexpected and interesting results. For the given definition of efficiency, there exists a combination of material parameters and drive frequency that yield efficiencies greater than one.Further analysis provided evidence that frequency domain formulation provides that while there are instances when relatively large quantities of mechanical energy in the system exist relative to the quantity of electrical energy being drawn by the actuator, total energy is always conserved. The important result of this work was the knowledge gained in the fundamental understanding of power, energy, and efficiency as it relates to dynamic actuation of an electromechanical system. The results shown here support the concept of actuation at or near a system resonance to increase efficiency. This work is on-going; the ultimate goal of which is to develop a tool for aiding active material actuation feasibility studies by using actuation efficiency as a performance metric.

As complex intelligent manufacturing systems appear increasingly in the industry, the necessity of more user- friendly man-machine interfaces is becoming progressively crucial for their utilization, and consequently for their market success. This paper introduces some of the recent technological advances in the intelligent manufacturing systems that influence the design and development of man- machine interfaces. At the same time, some of the research works in the Research Lab. of Composite Materials at Harbin Institute of Technology are presented.

Active Fiber Composite (AFC) actuators comprise piezoelectric fibers, polymer matrix, and interdigital electrodes. They are conformable, robust and have higher strain actuation than standard in-plane PZT actuators due to the 33 mode of actuation offered by interdigital electrodes. Taking advantage of the 33 actuation necessitates the alignment of the poling axis and load path axis. Piezoelectric material are more vulnerable to stress induced depolarization when the loading and poling axes are aligned than when they are transverse. The poled PZT fibers will remain poled as long as they are under sufficiently low mechanical, thermal, and electric fields. Extreme values of any of these environments will cause depolarization, resulting in substantially reduced actuation. The compressive stress levels along the poling axis required to depole bulk PZT-5A piezoceramic samples under low field excitation are well documented. An AFC is a more complicated system, however, and the compressive stress levels which cause depolarization are not necessarily the same as for bulk piezoceramics. This paper describes a set of experiments designed to determine the compressive stress induced depolarization limits of AFC actuators with PZT-5A fibers. Results indicate that the AFCs are much more robust to compressive stress induced depolarization than is suggested by data published on bulk piezoceramics.