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
This study compares a hybrid-electric aircraft featuring a propulsive empennage and overthe-wing, distributed-propulsion to a conventional regional turboprop. The impact of multiple design parameters, mission requirements, and technology assumptions on maximum takeoff mass and payload-range energy efficiency is evaluated, in order to illustrate the sensitivities of the design. A preliminary sizing method that incorporates aero-propulsive interaction effects is used to obtain rapid estimations. Results show that, for the baseline mission, the hybridelectric variant is 2.5% heavier and consumes 2.5% more energy than the reference aircraft. In this process, several key design guidelines and challenges for distributed-propulsion aircraft are identified. Firstly, when comparing a hybrid-electric configuration to a conventional one, each aircraft must be sized at its optimum cruise altitude for the same payload and range requirements. Secondly, the hypothetical advantages of distributed propulsion described in literature do not easily lead to a benefit at aircraft level, if the aero-propulsive interaction effects and associated dependencies are incorporated in the design process. Thirdly, the power-control parameters affect practically all characteristics of the aircraft, and the optimal control strategy is highly dependent on the aero-propulsive interaction. The results suggest that the proposed configuration can constitute a low-noise alternative for the regional transport market if the performance of the over-the-wing distributed-propulsion system is optimized.
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 • .
A Flying V aircraft is a tailless, V-shaped flying wing with two cylindrical pressurized cabins placed in the wing leading edge and two over-the-wing engines. Elevons provide longitudinal and lateral control while two tip-mounted vertical tails double as winglets. The goal of the presented study is to estimate the lift-to-drag ratio of this configuration at the cruise condition: M = 0.85, h = 13, 000m, and CL = 0.26. A vortex-lattice method is used to rapidly investigate the feasible design space, whereas an Euler solver on an unstructured grid is adopted for a more accurate wave and vortex-induced drag estimation. The profile drag is computed by an empirical method. The NASA Common Research Model is adopted as a benchmark with an estimated lift-to-drag ratio of 18.9. The three-dimensional geometry of the Flying V is generated according to a multi-level parametrization: the planform shape is parametrized with 10 variables, five wing sections are identified and described by a total of 43 parameters, while the winglet planform is defined by 3 additional variables. After a multi-fidelity design space exploration, two design approaches are investigated: a dual-step optimization, where planform and airfoil variables are subsequently varied, and a single-step optimization, where planform and airfoil variables are varied simultaneously. The highest lift-to-drag ratio is attained with the single-step optimized configuration and amounts to 23.7. It is therefore concluded that the Flying V Aircraft can have a 25% higher lift-to-drag ratio than the reference aircraft.
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 potential benefits of hybrid-electric propulsion (HEP) have led to an increased interest in this topic over the past decade. One promising advantage of HEP is the distribution of power along the airframe, which enables synergistic configurations with improved aerodynamic and propulsive efficiency. The purpose of this paper is to present a generic sizing method suitable for the first stages of the design process of hybrid-electric aircraft, taking into account the powertrain architecture and associated propulsion-airframe integration effects. To this end, the performance equations are modified to account for aero-propulsive interaction. A powerloading constraint-diagram is used for each component in the powertrain to provide a visual representation of the design space. The results of the power-loading diagrams are used in a HEP-compatible mission analysis and weight estimation to compute the wing area, powerplant size, and takeoff weight. The resulting method is applicable to a wide range of electric and hybrid-electric aircraft configurations and can be used to estimate the optimal power-control profiles. For demonstration purposes, the method is applied a HEP concept featuring leadingedge distributed-propulsion (DP). Three powertrain architectures are compared, showing how the aero-propulsive effects are inlcuded in the model. The results confirm the method is sensitive to top-level HEP and DP design parameters, and indicate an increase in wing loading and power loading enabled by DP.
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