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The design, modeling, and testing of a morphing wing for flight control of an uninhabited aerial vehicle is detailed. The design employed a new type of piezoelectric flight control mechanism which relied on axial precompression to magnify control deflections and forces simultaneously. This postbuckled precompressed bending actuator was oriented in the plane of the 12% thick wing and mounted between the end of a tapered D-spar at the 40% chord and a trailing-edge stiffener at the 98% chord. Axial precompression was generated in the piezoelectric elements by an elastic skin which covered the outside of the wing and served as the aerodynamic surface over the aft 70% of the wing chord. A two-dimensional semi-analytical model based on the Rayleigh-Ritz method of assumed modes was used to predict the static and dynamic trailing-edge deflections as a function of the applied voltage and aerodynamic loading. It was shown that static trailing-edge deflections of 3:1 deg could be attained statically and dynamically through 34 Hz, with excellent correlation between theory and experiment. Wind tunnel and flight tests showed that the postbuckled precompressed morphing wing increased roll control authority on a 1.4 meter span uninhabited aerial vehicle while reducing weight, slop, part-count, and power consumption. Nomenclature A = extensional stiffness matrix or aspect ratio B = coupled laminate stiffness matrix b = span C L , C l = three-dimensional, section lift coefficient c = chord D = bending laminate stiffness E = total energy F a = aerodynamic force F 0 = precompression force f = frequency K = structural stiffness K = stiffness matrix k = spring stiffness L = actuator length M = applied moment vector M = mass matrix m = mass N = applied force vector n = number of shape functions P = lift force p = pressure q = amplitude T = kinetic energy t = thickness or time U = internal energy or velocity u = horizontal displacement V = potential energy or voltage w = vertical displacement = angle of attack = trailing-edge deflection = normal strain = trailing-edge end rotation = curvature = unloaded actuator strain = potential energy = density = normal stress = velocity potential = disturbed velocity potential = shape function Subscripts a = actuator b = bonding layer c = circulatory ex = external h = hinge point l = laminate m = morphing part nc = noncirculatory sp = negative spring rate t = thermal

Current, highly active classes of adaptive materials have been considered for use in many different aerospace applications. From adaptive flight control surfaces to wing surfaces, shape-memory alloy (SMA), piezoelectric and electrorheological fluids are making their way into wings, stabilizers and rotor blades. Despite the benefits which can be seen in many classes of aircraft, some profound challenges are ever present, including low power and energy density, high power consumption, high development and installation costs and outright programmatic blockages due to a lack of a materials certification database on FAR 23/25 and 27/29 certified aircraft. Three years ago, a class of adaptive structure was developed to skirt these daunting challenges. This pressure-adaptive honeycomb (PAH) is capable of extremely high performance and is FAA/EASA certifiable because it employs well characterized materials arranged in ways that lend a high level of adaptivity to the structure. This study is centered on laying out the mechanics, analytical models and experimental test data describing this new form of adaptive material. A directionally biased PAH system using an external (spring) force acting on the PAH bending structure was examined. The paper discusses the mechanics of pressure adaptive honeycomb and describes a simple reduced order model that can be used to simplify the geometric model in a finite element environment. The model assumes that a variable stiffness honeycomb results in an overall deformation of the honeycomb. Strains in excess of 50% can be generated through this mechanism without encountering local material (yield) limits. It was also shown that the energy density of pressure-adaptive honeycomb is akin to that of shape-memory alloy, while exhibiting strains that are an order of magnitude greater with an energy efficiency close to 100%. Excellent correlation between theory and experiment is demonstrated in a number of tests. A proof-of-concept wing section test was conducted on a 12% thick wing section representative of a modern commercial aircraft winglet or flight control surface with a 35% PAH trailing edge. It was shown that camber variations in excess of 5% can be generated by a pressure differential of 40 kPa. Results of subsequent wind tunnel test show an increase in lift coefficient of 0.3 at 23 m s −1 through an angle of attack from −6 • to +20 • .

This paper describes how post-buckled precompressed (PBP) piezoelectric bender actuators are employed in a deformable wing structure to manipulate its camber distribution and thereby induce roll control on a subscale UAV. By applying axial compression to piezoelectric bimorph bender actuators, significantly higher deflections can be achieved than for conventional piezoelectric bender actuators. Classical laminated plate theory is shown to capture the behavior of the unloaded elements. A Newtonian deflection model employing nonlinear structural relations is demonstrated to predict the behavior of the PBP elements accurately. A proof of concept 100 mm (3.94 ) span wing employing two outboard PBP actuator sets and a highly compliant latex skin was fabricated. Bench tests showed that, with a wing chord of 145 mm (5.8 ) and an axial compression of 70.7 gmf mm −1 , deflection levels increased by more than a factor of 2 to 15.25 • peak-to-peak, with a corner frequency of 34 Hz (an order of magnitude higher than conventional subscale servoactuators). A 1.4 m span subscale UAV was equipped with two PBP morphing panels at the outboard stations, each measuring 230 mm (9.1 ) in span. Flight testing was carried out, showing a 38% increase in roll control authority and 3.7 times greater control derivatives compared to conventional ailerons. The solid state PBP actuator in the morphing wing reduced the part count from 56 down to only 6, with respect to a conventional servoactuated aileron wing. Furthermore, power was reduced from 24 W to 100 mW, current draw was cut from 5 A to 1.4 mA, and the actuator weight increment dropped dramatically from 59 g down to 3 g.

The deflection characteristics of Structures using directionally attached piezoelectric (OAP) and enhanced OAP (EDAP) elements are explored. Tests demonstrate that piezoceramic elements, which are isotropic, exhibit orthotropic behavior when directionally attached using any of three methods: (i) partial attachment, (ii) transverse shear lag, and (iii) differential stiffness bonding. Test results demonstrate that directional enhancement through transverse stiffening can increase OAP element strain from 5 to 25%. Closed form expressions 01 OAP/ EDAP strains based on classical laminated plate theory are presented. The models demonstrate that OAPIEDAP elements generate any in-plane strain (extensions and shear) or out-of plane curvature (bending in either direction and twist) independent of other Strains or curvatures. Test results show that fiberglass and aluminium oAP/EDAP beams produce torsional and bending deflections in excess of 30°m-' with theory and experiment in close agreement. The deflections of OAP/ EDAP and conventional piezoelectric active structures are compared. Tests show that oAp/EoAp elements can produce up to 16 times more twist than conventionally attached piezoceramic elements. Two wings were constructed with OAP and EDAP elements. EDAP elements were laminated into the skin of a graphite/epoxy supersonic wing that had a 9% thick diamond airfoil section and an aspect ratio of 3. OAP elements were also laminated to a torsion beam of a subsonic wing that had an NACA 0012 profile and an aspect ratio of 1.4. The supersonic wing demonstrated static twist deflections in excess of 2 ' . The subsonic wing demonstrated static pitch deflections of go. The lifting capability of the o A P / E w wings are compared to piezo-ailerons. The OAP~EOAP wings are shown to produce much larger changes in lift coefficient and greater deflection stability with increasing airspeed than the piezo-aileron configuration.