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It was a large plane for its time, a revolution wearing wings. The silver-gray Boeing 247 could carry ten passengers. Developed in absolute secrecy, it made its first flight at [...]

The replacement of the analysis portion of an optimization problem by its equivalent metamodel usually results in a lower computational cost. In this paper, a conventional nonapproximative approach is compared against three different metamodels: quadratic-interpolation-based response surfaces, Kriging, and artificial neural networks. The results obtained from the solution of four different case studies based on aircraft design problems reinforces the idea that quadratic interpolation is only well-suited to very simple problems. At higher dimensionality, the usage of the more complex Kriging and artificial neural networks models may result in considerable performance benefits. DOI: 10.2514/1.45790

Late on February 19, 1934, a brand new Douglas DC-1 named, 'City of Los Angeles,' took off from that city and headed east for Newark, New Jersey. The flight was [...]

A methodology for improving the quality of high-lift-system performance prediction within a multidisciplinary conceptual design process is presented. The high-lift-system geometry is explicitly modeled and a multiple-lifting-line method is used to compute its aerodynamic characteristics. Computation times are acceptable for use in a conceptual design process. The results for several test cases show good agreement with wind-tunnel and/or high-fidelity numerical data. In addition, the method allows for further enhancement by using nonlinear airfoil polars for interpolation, improving drag prediction, and introducing some degree of nonlinear aerodynamic behavior. DOI: 10.2514/1.42845

Numerical simulations of a turbulent flow over a backward-facing step are performed. The governing equations are solved by the finite volume code Cobalt. Unsteady three-dimensional detached eddy simulations are carried out with Menter's shear stress transport turbulence model acting as a subgrid-scale model. Mean flow quantities such as pressure, velocity, and skin-friction coefficients are accurately predicted. Velocity and pressure fluctuations are resolved by the three-dimensional computations. The dominant frequency is in good agreement with existing experimental data, and power spectral analysis of wall pressure fluctuation is consistent with empirical relations found in the literature for several operating conditions. DOI: 10.2514/1.43912

A method to concisely quantify, compare, and identify mode shapes for flutter analysis using orthogonalized basis vectors is developed. The procedure numerically describes wing structural mode shapes as scalar projections onto a reduced set of basis vectors. The basis vectors are baseline wing shapes derived from the mode shapes of a reference wing model, and mode shapes from other wing models are concisely related to the basis vectors using the scalar projections. The resulting quantifier describes the wing modes in terms of the physical wing shapes from the reference model. The technique is demonstrated on a simple wing model from which the basis vectors are derived. Mass elements are added to the wing in four different configurations to produce various mode shapes for comparison to the basis set. These mode shapes are quantified in terms of basis vector components that are expressed concisely in matrix form. Visual comparisons of the mode shapes to the basis vectors show that the numeric characterizations are consistent with the visual interpretations. Comparison of the four configurations to one another indicates that the numeric components distinguish subtle differences in wing deformation as well as differences in the overall shape of the deformed wing. DOI: 10.2514/1.42591

This paper presents the development and evaluation of a computational fluid dynamics based methodology to predict the aerodynamic forces produced by a flexible flapping wing. The computational fluid dynamics analysis code solves the compressible Reynolds-averaged form of the Navier-Stokes equations on structured curvilinear grids. A grid deformation algorithm is devised that deforms the body-conforming volume grid at each time step consistent with the measured wing motions. This algorithm is based on geometric considerations and is both computationally efficient and capable of handling very large deformations. This methodology is validated using experimental data obtained from a test on an ornithopter with flexible wings. Test data include measurements of the wing surface deformations as well as the generated forces in the horizontal and vertical directions. Correlation with test data shows good agreement with measured vertical force and satisfactory agreement with measured horizontal force at low flapping frequencies. However, the prediction accuracy degrades with an increase in flapping frequency. Evidence of resonance in the vehicle system was detected from the analysis of the experimental data. Unmodeled inertial effects from the vehicle body and support mounts may be one of the contributors to disagreement between the data and analysis. DOI: 10.2514/1.43187

This paper presents a practical methodology for static aeroelastic analysis and aeroelastic optimization via coupling of high-fidelity commercial codes. A finite-volume-based flow solver FLUENT is used to solve three-dimensional Euler equations, Gambit is used to generate mesh in the fluid domain, and CATIA is used to model parametric solid geometry. Abaqus, a structural finite element method solver, is used to compute the structural response of the aeroelastic system. The mesh-based parallel-code coupling interface MpCCI is used to exchange the pressure and displacement information between FLUENT and Abaqus to perform a loosely coupled aeroelastic analysis by a staggered algorithm, and modeFRONTIER software is used as the optimization driver for scheduling a nondominated sorting genetic algorithm initiated with design of experiments. First, an AGARD 445.6 wing configuration is optimized with objectives of maximum lift/drag ratio and minimum weight. Optimization variables are chosen as sweep angle at the quarter-chord and the taper ratio of the wing. Second, a more realistic wing model, ARW-2, is optimized for thickness values of the inner ribs and spars. Aeroelastic analysis produce consistent results with experimental data, and the applied optimization methodology results in Pareto-optimal solutions. DOI: 10.2514/1.41491

The axisymmetric Busemann biplane was investigated numerically. This configuration is composed of two cylindrical wings for which the cross section was similar to a regular two-dimensional Busemann biplane. The influence of various flow conditions on the drag coefficient was investigated. The axisymmetric configuration was found to be less effective than a two-dimensional Busemann biplane. However, drag was significantly smaller than that of a single-wing shape. DOI: 10.2514/1.41472

Electrical-discharge wire cutting is a promising technique that provides flexibility and lightness for a flapping micro aerial vehicle. Electrical-discharge wire cutting is used to fabricate the high-aspect-ratio structure of the four-bar linkage gear transmission module made of aluminum-alloy 7075. Aluminum-alloy 7075 has excellent specific strength (yield strength/density) good for durability of the transmission module in a micro aerial vehicle's tuff operation. A new flapping micro aerial vehicle of 21.6 cm wing span consequently has a minimum body mass of 5.9 g after installing the transmission module and a flexible wing frame made of carbon fibers and polyethylene terephthalate film. This micro aerial vehicle can endure a flight time of 6 min 7 s with the wingbeat frequency of 10-20 Hz. The lift and thrust coefficients of the micro aerial vehicle have been investigated through wind-tunnel testing. The proposed flapping micro aerial vehicle also exhibits the improved characteristic in the scaling law with respect to wingbeat frequency versus body mass. DOI: 10.2514/1.38862

NASA and industry recently ended the High Speed Civil Transport program. The objective of the High Speed Civil Transport program was to develop critical technologies to support the potential development of viable supersonic commercial transport aircraft. The aerodynamic design development activities benefited greatly from the use of the prior design, analysis, and prediction methods as well as the understanding of the fundamental physics inherent in an efficient supersonic aircraft design. It was recognized that the critical strengths of the aerodynamic processes included the blending of the computational power offered by computational fluid dynamics methods with the fundamental knowledge and rapid design development and assessment capabilities inherent in the existing linear aerodynamic theory methods. Nonlinear design optimization studies are typically initiated with an initial optimized linear theory baseline configuration design. In this paper, a new supersonic linear theory wave-drag optimization methodology using far-field wave-drag methodology is introduced. The method is developed using the class-function/ shape-function transformation concept of an analytic scalar wing definition. The methodology is applied to an arrowwing planform to illustrate its versatility as well as to demonstrate the usefulness of the class-function/shape-function transformation analytic wing concept for aerodynamic design optimization.

In this study, the three-dimensional flow computations over a realistic aircraft high-lift configuration with a flow-through nacelle with a pylon mounted beneath the main wing are performed using an unstructured mesh method to investigate the influence of the boundary-layer transition on the aerodynamic forces, the capability of a transition prediction method, and the influence of brackets to support the high-lift devices. First, the influence of the boundary-layer transition on the aerodynamic forces is shown by comparison of the computational results with/without the boundary-layer transition. Then a transition prediction method based on the [e.sup.N] method and semi-empirical approaches is evaluated. The capability and areas to be improved are discussed. The influence of brackets to support the high-lift devices on the aerodynamic forces is also discussed. Interference of disturbed wakes by the slat supports to the flows on the main wing and flap is shown.

A study of the fatigue strength of full-size nitrided crankshafts made with 4340 vacuum arc remelting steel is presented. This investigation was undertaken in response to several in-flight failures of crankshafts in aircraft engines. The absence of fatigue endurance limit data for full-size, nitrided crankshafts made from 4340 vacuum arc remelting steel was specifically noted by the Federal Aviation Administration. Both nitriding and size effects are known to increase the fatigue strength typically reported for 4340 vacuum arc remelting steel using unnitrided, cylindrical rotating-beam test samples. Full-size nitrided crankshafts were tested in complete load reversal and were found to have a fatigue endurance limit of approximately 160 ksi. This value was found to be in excellent agreement with theoretical predictions using fracture-mechanics analyses and with prior fatigue testing of full-size nitrided crankshafts using air-melted 4340 steel.

The nonlinear resonant decay method identifies a nonlinear dynamic system using a model based in linear modal space comprising the underlying linear system and a small number of additional terms that represent the nonlinear behavior. In this work, the method is applied to an aircraftlike wing/store/pylon experimental structure that consists of a rectangular wing with two stores suspended beneath it by means of nonlinear pylons with a nominally hardening characteristic in the store rotation degree of freedom. The nonlinear resonant decay method is applied to the system using multishaker excitation. The resulting identified mathematical model features five modes, two of which are strongly nonlinear, one is mildly nonlinear, and two are completely linear. The restoring force surfaces obtained from the mathematical model are in close agreement with those measured from the system. This experimental application of the nonlinear resonant decay method indicates that the method could be suitable for the identification of nonlinear models of aircraft in ground vibration testing.

This paper develops a framework of dynamic simulation driven fatigue life analysis of a landing probe system for a typical 12-ton tricycle landing gear helicopter for embarked operations on the typical frigate. By integrating a novel dynamic helicopter/ship interface simulation with the rainflow cycle counting method, fatigue spectra, including all possible probe load cases under the wide range of operating and environmental conditions, have been developed with a confidence level of greater than 99.9 %: otherwise, they would be practically unobtainable, even by limited sea trial testing. Furthermore, the fatigue stresses of the probe assembly were obtained by the finite element method, and the cumulative fatigue damage analyses were conducted by monitoring the fatigue life of the critical locations on each component of the probe assembly using the Palmgren--Miner rule against the design life requirement. This new approach provides an innovative and efficient design tool, through virtual prototyping, that can speed up the design process and reduce cost.

Morphing wings are desired for their ability to reduce drag, to change flight characteristics, and perhaps to reduce weight by eliminating flap/aileron mechanisms. Development of two generations of a morphing wing project is documented. The work shows how a relatively low cost but realistic morphing wing test-bed can be fabricated. Wing skin, actuator, and actuator attachment development are discussed, as well as possible auxetic skin behavior. Aerodynamic characterization of the wing will be discussed in another paper. A very simple morphing wing was fabricated in generation one. The nose was able to elastically camber down approximately 25 deg and the tail 20 deg. Actuation was provided by three pneumatic 'rubber muscle actuators' that produce high contractive/ tensile forces. Upper and lower wing skins were fabricated from carbon fiber/ polyurethane elastomer laminates. Lower skin buckling, actuator air leaks, and actuator attachment problems were resolved in the second generation. A finite element model of the second wing was developed and is being used to refine the morphing wing test-bed. The second wing fabrication methodology enabled smooth elastic cambering with no buckling or waviness in the skins. The nose cambered down 14 deg and the tail cambered down to 13 deg, and is capable of larger deformations. Improved leak-free biomimetic actuators and attach points now include no metal parts and have higher actuation forces due to new braided sheaths and functionally gradient matrix properties. [DOI: 10.111511.3159043]

This paper presents the work done in designing a morphing wing concept for a small experimental unmanned aerial vehicle to improve the vehicle's performance over its intended speed range. The wing is designed with a multidisciplinary design optimization tool, in which an aerodynamic shape optimization code coupled with a structural morphing model is used to obtain a set of optimal wing shapes for minimum drag at different flight speeds. The optimization procedure is described as well as the structural model. The aerodynamic shape optimization code, that uses a viscous two-dimensional panel method formulation coupled with a nonlinear lifting-line algorithm and a sequential quadratic programming optimization algorithm, is suitable for preliminary wing design optimization tasks. The morphing concept, based on changes in wing-planform shape and wing-section shape achieved by extending spars and telescopic ribs, is explained in detail. Comparisons between optimized fixed wing performance, optimal morphing wing performance, and the performance of the wing obtained from the coupled aerodynamic-structural solution are presented. Estimates for the performance enhancements achieved by the unmanned aerial vehicles when fitted with this new morphing wing are also presented. Some conclusions on this concept are addressed with comments on the benefits and drawbacks of the morphing mechanism design. DOI: 10.2514/1.39016

Analytical expressions are developed that describe the artifacts encountered when translating, rotating, and vibrating point sources are imaged by compact radar ranges (CRRs) emulating airborne synthetic-aperture radar (SAR). The approach starts with coherent-aperture imaging basics and develops a general solution for imaging using the Born approximation. We show that moving-target artifacts on CRRs are similar to the artifacts encountered with SAR moving targets, suggesting that CRRs may be suitable for such emulations.

Expressions describing the correspondence between moving-target artifacts for synthetic-aperture radar (SAR) and compact radar range (CRR) emulations are developed. Experimental results are then presented to verify the accuracy of these equations and to characterize the rapacity of CRRs to emulate translating, vibrating, and rotating targets. We show that phase-compensation techniques can be applied to CRR data, providing a suitable test bed for SAR phase-compensation techniques.

This paper reviews concepts for flow control based on jet spoilers and presents a new device based on the jet principle for application to fixed-wing aircraft. The new concept relies on the naturally blown inverse spoiler. The jet is blown parallel to the surface and points against the freestream. The theoretical basis of the device is presented. Calculations are performed with a Navier--Stokes program to investigate the key design parameters. These parameters include the spoiler deflection (equal or differential), the position of the hinge points, the shape of the bypass channel, the effects of backward- and forward-facing steps in a partial deflection, and the nature and magnitude of the control forces. A comparison is done with conventional inlay spoilers. Finally, a multidimensional optimization has been carried out to study the combined effects of hinge position, deflection, and spoiler length. Optimal configurations with respect to jet momentum and drag coefficient have been analyzed. DOI: 10.2514/1.40321

A simplified method for analyzing and designing a vertical tail buffeting alleviation system is developed. The vertical tail model in this study is equipped with surface-bonded piezoelectric actuators to suppress the buffeting responses induced by aerodynamic forces. The structural dynamics of the vertical tail and the electrodynamics of piezoelectric actuators are modeled using a finite element method and realized by commercial software packages ANSYS[R]. Afterward, the finite element model for both the host structure and the piezoelectric patches are imported into MATLAB[R], in which the dynamic equations of the system in modal coordinates are obtained through modal truncation. The motion-induced aerodynamic forces are computed by using the doublet-lattice method. The buffet input excitations are simulated by the proper orthogonal decomposition-based method. The suboptimal controller is designed to compute suitable control voltages needed to drive piezoelectric actuators. Finally, the effectiveness of the piezoelectric actuators in reducing vibratory responses due to buffet loads on the vertical tail is investigated numerically. The results demonstrate that the proposed method is feasible and effective. DOI: 10.2514/1.39464

The model for aeroelastic validation research involving computation semispan wind-tunnel model, a transport wing--fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental limit cycle oscillation behavior data at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wing-tip configurations: clean, tip store, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6-0.9, and maps of limit cycle oscillation behavior were made in the range of M = 0.85-0.95. The effects of dynamic pressure and angle of attack were measured. Testing in both R134a heavy gas and air provided unique data on the Reynolds number, transition effects, and the effect of speed of sound on limit cycle oscillation behavior. This report gives an overview of the test results, including experimental flutter boundaries, and the conditions involving shock-induced transonic flow separation onset at low wing angles, including maps of limit cycle oscillation behavior. DOI: 10.2514/1.30079

Generic wraparound aerodynamic shape optimization technology is presented and applied to a modern commercial aircraft wing in transonic cruise. The wing geometry is parameterized by a novel domain-element method, which uses efficient global interpolation functions to deform both the surface geometry and corresponding computational fluid dynamics volume mesh. The technique also provides a method that allows geometries to be parameterized at various levels, ranging from global three-dimensional planform alterations to detailed local surface changes. Combining all levels of parameterization allows for free-form design control with very few design variables. The method provides an efficient combined shape parameterization and high-quality mesh deformation technique that is totally independent of mesh type (structured or unstructured). Optimization independence from the flow solver is achieved by obtaining sensitivity information for an advanced gradient-based optimizer by finite differences. The entire optimization suite has also been parallelized to allow optimization with highly flexible parameterization in practical times. Results are presented for highly constrained optimizations of the modern aircraft wing in transonic cruise, using three levels of parameterization (number of design variables) to assess the effect of parameterization level on the optimization. The highest-level optimization results in a totally-shock-free geometry with an associated substantial reduction in drag.

The focus of this paper is an optimal design of morphing aircraft wings employing a wing structure composed of an internal layout of cables and struts. Cables are used to provide actuation and stiffness, and struts provide stiffness without actuation. Topology optimization is used to place cables and struts in a bay or a section of the wing. Nonlinear finite element analysis is used to capture the large deformations of the structure, and the optimization is achieved using the Nondominated Sorting Genetic Algorithm II. The optimization procedure is illustrated using a morphing-wing example. The effect of the upper limit on actuation forces is studied, and solutions are found with good agreement between the desired and obtained deflections under actuation and aerodynamic loads. The implemented parallelized optimization algorithm is successful in solving a computationally intense, multi-objective, multiconstraint problem with a large number of discrete and continuous design variables in a reasonable amount of time.

Successful realization of a flapping wing micro-air vehicle (MAV) requires development of a light weight drive mechanism that can convert the continuous rotary motion of the motor into oscillatory flapping motion of the wings. The drive mechanism should have low weight to maximize the payload and battery capacity. It should also have high power transmission efficiency to maximize the operational range and to minimize weight of the motor. In order to make flapping wing MAVs attractive in search, rescue, and recovery efforts, they should be disposable from the cost point of view. Injection molded compliant drive mechanisms are an attractive design option because of manufacturing scalability and reduction in the number of parts. However, realizing compliant drive mechanism using injection molding requires use of multipiece multigate molds. Molding process constraints need to be considered during the design stage to successfully realize the drive mechanism. This paper describes an approach for determining the drive mechanism shape and size that meets both the design and molding requirements. The novel aspects of this work include (1) minimizing the number of mold pieces and (2) the use of sacrificial shape elements to reduce the impact of the weld-lines on the structural performance. The design generated by the approach described in this paper was utilized to realize an operational flapping wing MAV. [DOI: 10.1115/1.3116258]