Showing papers in "Journal of Aircraft in 2007"

Experimental Evaluation of Sinusoidal Leading Edges

Abstract: P REVIOUS studies on increasing airfoil lift and improving stall characteristics have addressed various passive and active approaches to modifying the leading and trailing edge shapes. The passive approaches have covered such methods as rippling the trailing edge, applying serrated-edge Gurney flaps, or modifying the leading-edge (LE) profile [1,2]. Other efforts have effectively eliminated the dynamic stall of an NACA 0012 airfoil by perturbing the LE contour as little as 0.5–0.9%of the chord [3]. Levshin et al. [4] demonstrated that sinusoidal LE planforms on an NACA 63-021 airfoil section decreased maximum lift, but extended the stall angle by almost 9 deg. The larger amplitude sinusoids created “softer” stall characteristics by maintaining attached flow at the peaks despite separated flow in the troughs. These tests were performed to simulate the effects of LE tubercles on humpback whale (Megaptera novaeangliae) flippers. Prior work by the authors also reported wind tunnel measurements for idealized scale models of humpback whale flippers [5]. One model had a smooth leading edge and a secondmodel had sinusoidal bumps (tubercles) along the leading edge for the outer 2 3 of the span. It was found that the addition of tubercles to a 3-D idealized flipper increased the maximum lift coefficient while reducing the drag coefficient over a portion of the operational envelope. It is thought that the tubercles on the flipper leading-edge enhance the whale’s ability to maneuver to catch prey [6]. Though the work to date regarding sinusoidal or serrated leading-edge planforms is largely motivated by marine mammal locomotion, the effects of extending the stall point for lifting surfaces at similar Reynolds numbers (Re) may have application to small-UAV (unmanned aerial vehicle) design and the inevitable laminar stall problems [7]. However other relevant applications might benefit from the effects of simulated tubercles such as stall alleviation/separation control on sailboat centerboards or wind turbines, where an expanded operating envelope could improve the overall effectiveness of the blade [8,9]. In the present work, a better understanding is sought of the mechanism of the improvements measured in previous experiments, with a greater applicability in mind. The authors seek to determine whether the performance improvements resulted from enhancements to the sectional characteristics of wings with tubercles (i.e., essentially 2-D effects), or from Reynolds number effects on a tapered planform, or from other 3-D effects such as spanwise stall progression.

Development and testing of the mentor flapping-wing micro air vehicle

Abstract: In 1997 the Defense Advanced Research Projects Agency initiated a program to explore the possibility of micro air vehicles for the purpose of individually portable surveillance systems for close-range operations. The various contractors approached the problem in several ways, such as developing tiny fixed-wing airplanes, rotary-wing aircraft, and ornithopters mimicking animal flight This paper describes one such flapping-wing aircraft, which drew upon the clap-fling phenomenon that is exploited by many flying animals and insects for lift generation. Essentially this aircraft was a mechanical simulation of hummingbird flight, though with two sets of wings to eliminate the unbalanced side-to-side flapping forces. Two flying demonstration models were built, one with an internal-combustion engine and another with an electric motor. In both cases, these incorporated a drive train to reduce the high rpm rotary shaft motion to lower-frequency oscillation for flapping. Also required was a programmable logic board for stabilization. Successful hovering flight was achieved with both models, and initial studies of transition to horizontal flight were also explored.

Effect of Microvortex Generators On Seperated Normal Shock/ Boundary Layer Interactions

Abstract: Experiments have been performed in a blowdown supersonic wind tunnel to investigate the effect of subboundary layer vortex generators placed upstream of a normal shock/turbulent boundary layer interaction at a Mach number of 1.5 and a freestream Reynolds number of 28 x 10 6 . The Reynolds number based on the inflow boundary layer displacement thickness was 26,000. Two types of subboundary layer vortex generators were investigated: wedge-shaped and counter-rotating vanes. It was found that the vane-type subboundary layer vortex generators eliminated and the wedge-type subboundary layer vortex generators greatly reduced the shock-induced separation. When placed in the supersonic part of the flow, both types of subboundary layer vortex generators caused a wave pattern consisting of a shock, reexpansion, and shock. The reexpansion and double shocks are undesirable features because they equate to increased total pressure losses. Furthermore there are indications that the vortex intensity is reduced by the normal shock/boundary layer interaction. Overall, the vane-type subboundary layer vortex generators were the more effective devices as they eliminated the shock-induced separation and had the least detrimental effect on the shock structure.

Modeling and Flight Control of Large-Scale Morphing Aircraft

Abstract: As morphing is an emerging topic of interest in aircraft research, the following article provides a review of the subject, with specific focus on modeling and flight control of large-scale planform altering flight vehicles. Our discussion proceeds in a fundamental manner to demonstrate that, although design methods for rigid aircraft have become highly developed, the consideration of morphing necessitates further investigation into the typically disparate fields of dynamic modeling, aerodynamic theory, and flight control theory. To clarify these points, the equations of atmospheric flight are derived in a general form, methods of integrating the aerodynamic forces are examined, and we distinguish between various approaches and methods of flight control.

Plasma Actuators for Hingeless Aerodynamic Control of an Unmanned Air Vehicle

Abstract: The use of dielectric barrier discharge plasma actuators for hingeless flow control over a 47-deg 1303 unmanned combat air vehicle wing is described. Control was implemented at the wing leading edge to provide longitudinal control without the use of hinged control surfaces. Wind-tunnel tests were conducted at a chord Reynolds number of 4.12 x 105 and angles of attack ranging from 15 to 35 deg to evaluate the performance of leading-edge plasma actuators for hingeless flow control. Operated in an unsteady mode, the actuators were used to alter the flowfield over the lee-side wing to modify the aerodynamic lift and drag forces on the vehicle. Multiple configurations of the plasma actuator were tested on the lee side and wind side of the wing leading edge to affect the wing aerodynamics. Data acquisition included force-balance measurements, laser fluorescence, and surface flow visualizations. Flow visualization tests mainly focused on understanding the vortex phenomena over the baseline uncontrolled wing to aid in identifying optimal locations for plasma actuators for effective flow manipulation. Force-balance results show considerable changes in the lift and drag characteristics of the wing for the plasma-controlled cases compared with the baseline cases. When compared with the conventional traditional trailing-edge devices, the plasma actuators demonstrate a significant improvement in the control authority in the 15- to 35-deg angle-of-attack range, thereby extending the operational flight envelope of the wing. The study demonstrates the technical feasibility of a plasma wing concept for hingeless flight control of air vehicles, in particular, vehicles with highly swept wings and at high angles of attack flight conditions in which conventional flaps and ailerons are ineffective.

Numerical investigations of injection-slot-size effect on the performance of coflow jet airfoils

Abstract: Three co-flow jet (CFJ) airfoils with injection slot size differed by two times consecutively are calculated by using a RANS CFD solver with 1-equation Spalart-Allmaras model. At the same angle of attack(AoA), the twice larger injection slot size airfoil passes about twice larger jet mass flow rate with the momentum coefficients also nearly doubled. The CFJ airfoil with the largest slot size has the least stall angle of attack(AoA). When the injection slot size is reduced by half, the stall AoA and the maximum lift coefficient is increased. However, when the injection slot size is further reduced by half, the stall AoA is still increased, but the maximum lift coefficient is lower due to the smaller momentum coefficient. The trend of the stall AoA and maximum lift agree with the experiment. At low AoA, both the computed lift and drag agree fairly well with the experiment. At high AoA, both the lift and drag are underpredicted. The reason may be that the RANS model can not handle the turbulence mixing well at high AoA.

Autonomous Sensing and Control of Wing Stall Using a Smart Plasma Slat

Abstract: DOI: 10.2514/1.24057 The concept of a self-governing smart plasma slat for active sense and control of flow separation and incipient wing stall is presented. The smart plasma slat design involves the use of an aerodynamic plasma actuator on the leading edge of a two-dimensional NACA 0015 airfoil in a manner that mimics the effect of a movable leading-edge slat of a conventional high-lift system. The self-governing system uses a single high-bandwidth pressure sensor and a feedback controller to operate the actuator in an autonomous mode with a primary function to sense and control incipient flow separation at the wing leading edge and to delay incipient stall. Two feedback control techniques are investigated. Wind tunnel experiments demonstrate that the aerodynamic effects of a smart actuator are consistent with the previously tested open-loop actuator, in that stall hysteresis is eliminated, stall angle is delayed by 7 deg, and a significant improvement in the lift-to-drag ratio is achieved over a wide range of angles of attack. These feedback control approaches provide a means to further reduce power requirements for an unsteady plasma actuator for practical air vehicle applications. The smart plasma slat concept is well suited for the design of low-drag, quiet, highlift systems for fixed-wing aircraft and rotorcraft.

Morphing wing flight control via postbuckled precompressed piezoelectric actuators

Abstract: 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.

Gust Loads Alleviation Using Special Control Surfaces

Abstract: Control laws are designed for the alleviation of dynamic gust loads on a wind-tunnel model of a transport aircraft using wing-mounted control surfaces. Three different control surfaces are used: the symmetrically actuated main ailerons, special underwing forward-positioned control surfaces at about 0.8 of the wingspan, and special wing-tip forward-positioned control surfaces. The 5.3-m-span cable-mounted wind-tunnel model was constructed at the TsAGI laboratories and is being tested as there part of the 3AS 5th Framework European Commission research project. The length of one-minus-cosine vertical gust velocity profile is tuned to yield maximal wing-root bending moment All the control laws are based on simple low-pass filters for easy and robust application in the wind tunnel. Each is based on single input of a wing-tip accelerometer, which is shown to react sufficiently before the peak of the wing-root bending moment

Two-Dimensional Wind Tunnel and Computational Investigation of a Microtab Modified Airfoil

Abstract: *† ‡ A computational and wind tunnel investigation into the effectiveness of a microtab-based aerodynamic load control system is presented. The microtab-based load control concept consists of a small tab, with a deployment height on the order of 1% of chord, which emerges approximately perpendicular to a lifting surface in the vicinity of the trailing edge. Lift mitigation is achieved by deploying the tabs on the upper (suction) surface of a lifting surface. Similarly, lift enhancement can be attained by tab deployment on the lower (pressure) surface of a lifting surface. A sensitivity analysis using Reynolds-averaged NavierStokes methods was conducted to determine optimal sizing and positioning of the tabs for active load control at a chord Reynolds number of 1.0 million for the S809 baseline airfoil. These numerical simulations provide insight into the flow phenomena that govern this promising load control system and guided tab placement during the wind tunnel study of the S809 airfoil. The numerical and experimental results are largely in agreement and demonstrate that load control through microtabs is viable. Future efforts will include a study of the unsteady load variations that occur during tab deployment and retraction, and three-dimensional issues involving spanwise tab placement and tab gaps.

Design and Testing of a Pneumatic Telescopic Wing for Unmanned Aerial Vehicles

Abstract: This paper discusses the design, development, and testing of an unmanned aerial vehicle pneumatic telescopic wing that permits a change in the wingspan, while simultaneously supporting structural wing loads. The key element of the wing is a pressurized telescopic spar able to undergo large-scale spanwise changes while supporting wing loadings in excess of 15 lb/ft 2 . The wing cross section is maintained by NACA 0013 rib sections fixed at the end of each element of the telescopic spar. Hollow fiberglass shells are used to preserve the spanwise airfoil geometry and ensure compact storage and deployment of the telescopic wing. A full-scale telescopic wing assembly was built and tested in the Glenn L. Martin Wind Tunnel at the University of Maryland. These tests included aerodynamic measurements at a variety of Reynolds numbers. The telescopic wing was tested in three different configurations and experimental results are compared with finite wing theory and results obtained on a rigid fixed-wing counterpart Preliminary aerodynamic results were promising for the variable wingspan telescopic wing. As expected, the telescopic wing at maximum deployment incurred a slightly larger drag penalty and a reduced lift-to-drag ratio when compared to its solid fixed-wing counterpart. However, the penalty was minimal and thus the development of an unmanned aerial vehicle with a pneumatic variable span wing is feasible.

System-of-Systems Inspired Aircraft Sizing and Airline Resource Allocation via Decomposition

Abstract: The phrase "system-of-systems" describes a large system of multiple systems, each capable of independent operation, which have been brought together to provide capabilities beyond those of each individual constituent system. Formulating and solving a system-of-systems design problem has become increasingly important in the aerospace and defense industries as customers have begun to ask contractors for broad capabilities and solutions rather than for specific individual systems. Part of a system-of-systems design problem is determining the appropriate mix of both existing and yet-to-be-designed systems. Whereas determining an appropriate mix of existing systems falls into the category of resource allocation, including features of a yet-to-be-designed system complicates the problem by requiring the allocation of a variable resource. In this paper, an airline wishing to investigate how a new, yet-to-be-designed aircraft will impact the fleet operating costs provides a simple example of this type of problem. The resulting statement is a mixed-integer, nonlinear programming problem. Both a traditional approach and a new decomposition approach were used to solve the problem. The traditional approach is a mixed-integer nonlinear branch-and-bound. The decomposition approach applied to the problem is analogous to those of multidisciplinary optimization, in which there is an allocation domain and an aircraft sizing domain. When the problem size increases to the point where the traditional mixed-integer, nonlinear programming approaches cannot obtain a solution, the decomposition approach can find solutions for these larger problems. The multidisciplinary design optimization-motivated decomposition approach appears to have promise for the allocation of variable resources challenge presented by many system-of-systems design problems.

Thin-Airfoil Theoretical Interpretation for Gurney Flap Lift Enhancement

Abstract: Thin-airfoil theory is applied to the lift problem of an airfoil with a Gurney flap. The lift and pitching moment coefficient increments are given as a square-root function of the relative Gurney flap height, and they are proportionally related. This model interprets the Gurney flap lift enhancement as a special camber effect. The theoretical relations are in good agreement with experimental and numerical data for several different wings. The theoretical method developed in this paper can be applied to similar trailing-edge devices for lift enhancement, and it is useful in the preliminary design of these flow control devices.

Piezoelectric transducer actuated leading edge de-icing with simultaneous shear and impulse forces

Abstract: Piezoelectric actuations for simultaneous generation of shear and impulse forces for effective and energy efficient de-icing applications are proposed. Aircraft leading edge structures are considered for analysis and lab experiments. Piezoelectric actuators are affixed on the inner surface of the leading edge at the locations where highest amount of ice accretion on the outer surface has occurred. Simultaneous shear and impulse force generation is achieved with actuators consisting of two sets of electrodes, one arranged in parallel to the poling direction, and the other perpendicular to it, to generate shear forces and normal forces, respectively. Finite element models of the leading edge structure with ice accretion layer are formulated. Simulations of the de-icing process are performed and the actuator locations, electric charge applied, and impulse duration are optimized to achieve effective ice removal

Assessment of Ionosphere Spatial Decorrelation for Global Positioning System-Based Aircraft Landing Systems

Abstract: Ground-based augmentations of the global positioning system demand guaranteed integrity to support aircraft precision approach and landing navigation. To quantitatively evaluate navigation integrity, an aircraft computes vertical and lateral protection levels as position-error bounds using the standard deviation of ionosphere spatial decorrelation. Thus, it is necessary to estimate typical ionospheric gradients for nominal days and to determine an appropriate upper bound to sufficiently cover the differential error due to the ionosphere spatial decorrelation. Both station-pair and time-step methods are used to assess the standard deviation of vertical (or zenith) ionospheric gradients (σ vig ). The station-pair method compares the simultaneous zenith delays from two different reference stations to a single satellite and observes the difference in delay across the known ionosphere pierce point separation. Because these ionosphere pierce point separations limit the observability of the station-pair method, the time-step method is also used to better understand ionospheric gradients at short distance scales (10-40 km). The time-step method compares the ionospheric delay of a single line of sight at one epoch with the delay for the same line of sight at another epoch a short time (a few to tens of minutes) later. This method has the advantage of removing interfrequency bias calibration errors on different satellites and receivers while possibly introducing an estimation error due to temporal ionospheric gradients. The results of this study demonstrate that typical values of σ vig are on the order of 1-3 mm/km for nonstormy ionospheric conditions. As a result, σ vig of 4 mm/km is conservative enough to bound ionosphere spatial decorrelation for nominal days and still leave enough margin for more active days and for non-Gaussian tail behavior.

Near-space station-keeping performance of a large high-altitude notional airship

Abstract: A key technical challenge for high-altitude (near-space) concepts is autonomous station keeping, or the ability to remain fixed over a geo-location in the presence of winds. Although operational altitudes of 65,000 ft. altitude and above are above weather (i.e., storms, rain), winds still exist. And to station keep in the presence of these winds requires propulsive power. This paper focuses on the analysis of the station keeping performance of a notional solar-powered airship operating at an altitude around 65,000. The vehicle has a lifting-gas volume of 6.1 million ft 3 , is 450 ft in length, and is solar-electric powered with the entire upper surface covered by solar cells. Electric motors and propellers provide propulsion, additional on-board power is required for the payload and auxiliary equipment, and batteries are used for power storage. In this paper, the aerodynamic drag of the vehicle is estimated, the wind environment is characterized, and the solar-power available is determined for several geo-locations and time of year. This power available is then compared to the power required for station keeping in the presence of winds. It is shown that due to diurnal solar variations and season winds a large airship may be unable to meet station-keeping performance requirements in certain months of the year due to power limitations

Aerodynamic Modeling of Morphing Wings Using an Extended Lifting-Line Analysis

Abstract: This paper presents an extension of Weissinger’smethod and its use in analyzing morphing wings. Thismethod is shown to be ideal for preliminary analyses of these wings due to its speed and adaptability to many disparate wing geometries. It extends Prandtl’s lifting-line theory to planform wings of arbitrary curvature and chord distribution and nonideal airfoil cross sections. The problem formulation described herein leads to an integrodifferential equation for the unknown circulation distribution. It is solved using Gaussian quadrature and a sine-series representation of this distribution. In this paper, this technique is used to analyze the aerodynamics of a morphable gull-like wing. Specifically, this wing’s ability to manipulate lift-to-drag efficiency and center of pressure location is discussed.

Investigation of Membrane Actuation for Roll Control of a Micro Air Vehicle

Abstract: A class of micro air vehicles uses a flexible membrane wing for weight savings and passive shape adaptation. Such a wing is not amenable to conventional aileron mechanisms for roll control, due to a lack of internal wing structure. Therefore, morphing (in the form of asymmetric twisting) is implemented through the use of a torque-actuated wing structure with thousands of discrete design permutations. A static aeroelastic model of the micro air vehicle is developed and validated to optimize the performance of the torque-actuated wing structure. Objective functions include the steady-state roll rate and the lift-to-drag ratio incurred during such a maneuver. An optimized design is obtained through the use of a genetic algorithm presenting significant improvements in both performance metrics compared with the baseline design.

Active separation control on the flap of a two-dimensional generic high-lift configuration

Abstract: The paper describes experimental results of controlling flow separation by periodic excitation on the flap of a generic high-lift configuration. The single slotted flap of the two-dimensional test model is equipped with a robust and reliable actuator system that fits inside the flap. The flow is excited using a pulsed wall jet that emanates from the upper surface near the flap's leading edge through a small spanwise-oriented slot By preventing the flow from separating or by reattaching the separated flow, lift and drag are substantially improved, resulting in a lift-to-drag ratio enhancement of 20-25 %. Because of the actuator assembly with spanwise individually addressable segments, the separated flow can be forced to attach only to certain parts of the flap. Local spanwise excitation is thus used to generate a rolling moment without the need to deflect an aileron.

Flowfield of a Rotating-Wing Micro Air Vehicle

Abstract: Experiments were conducted to measure the performance of a rotor, typical of that used on a rotating-wing micro air vehicle, which was shown to develop relatively low hovering efficiency. This inefficiency can be traced to high profile drag losses on the blades and also to the relatively large turbulent and rotational aerodynamic losses that are associated with the structure of the rotor wake. High-resolution flow visualization images have divulged several interesting flow features that appear unique to rotors operating at low Reynolds numbers. The wake sheets trailing from the rotor blades were found to be much thicker and also more turbulent than their higher chord Reynolds number counterparts. Similarly, the viscous core sizes of the tip vortices were relatively large as a fraction of blade chord. However, the flows in the tip vortices themselves were found to be similar, with laminar flow near their core axis and an outer turbulent region. Particle image velocimetry measurements were made to quantify the structure and strength of the wake flow, including the tip vortices. An analysis of the vortex aging process was conducted. A new nondimensional equivalent time scaling parameter is proposed to normalize the rate of growth of the vortex cores that are generated at substantially different vortex Reynolds numbers.

Semi-Active Control of Vertical Stroking Helicopter Crew Seat for Enhanced Crashworthiness

Abstract: S HOCK load-induced injury minimization has become an important issue in helicopter seat design. Harsh vertical landings or crash landings of these aircraft tend to result in pilot or occupant spinal and pelvic injuries. The severity of injury, however, can be reduced if the vehicles are outfitted with crashworthy seat designs. Utilization of a seat suspension system to attenuate the vertical shock loads that are transmitted from the base frame of the aircraft of the vehicle and imparted into the human body is a prime factor in determining survivability [1]. Within the cockpit, energy-absorbing crew seats have greatly enhanced helicopter crash survivability. Energy absorbers (EAs) are a key component in these energy-absorbing seats. The first examples of crashworthy crew seat designs employed fixed-load energy absorbers (FLEAs) to limit an occupant’s spinal load. These FLEAs are not adjustable (i.e., passive) so that their stroke and load profile are fixed at a factory-established, constant load throughout their entire operating range. Variable load energy absorbers (VLEAs) were developed subsequently to permit the occupant to manually adjust the constant stroking load by setting a dial for occupant weight. The stroking load of the VLEA is selected a priori that is proportional to the occupant weight, so that each occupant will experience similar acceleration (typically 14.5 G) and use similar stroking space during a high sink rate event. VLEAs exploit the fact that the strength of an occupant’s spine is nearly proportional to occupant weight, so that the VLEA will deliver the same low injury risk regardless of occupant weight. This technology was applied in programs to retrofit new seats into platforms such as the U.S. Navy’s CH-53 Sea Stallion and SH-3 Sea King aircraft [2]. Both fixed and variable load energy absorbers, however, are passive, in that they cannot automatically adapt their energy absorption or stroking profiles as a function of occupant weight, or as a function of real-time environmental measurements such as vibration level, shock level, sink rate, etc. This motivates the development of a seat suspension that uses an electronically adjustable adaptive energy absorber that can respond to such changing environmental stimuli via commands from a real-time feedback control system. Magnetorheological energy absorbers (MREAs) offer an innovative way to achieve what is effectively a continuously adjustable profile EA [3]. Using feedback control, the MREA can smoothly adjust the load profile as the seat strokes during a crash. Thus, MREAs are expected to provide the optimum combination of short stroking distance and minimum spinal load, while automatically adjusting for the occupant weight and load level. Of the three potential seat suspension approaches (passive, semiactive, and active), the semi-active approach is very attractive. A major drawback of a passive seat suspension based on viscoelastic or hydraulic energy absorbers is that performance is limited because neither damping nor stiffness are controllable. Furthermore, compared to active approaches, semi-active systems tend to require less power and have no stability issues (because semi-active force is always dissipative). Many researchers have been inspired to develop novel seat suspensions showing improved shock and vibration attenuation performance by permitting stiffness or damping to be adaptable and controllable. Wu and Griffin [4] examined several semi-active control algorithms for reduction in the severity of seat suspension end-stop impacts. Choi et al. [5,6] evaluated the attenuation of seat vibration using skyhook and sliding mode control algorithms on both electrorheological (ER) andMR seat suspensions for commercial vehicles. Park and Jeon [7] developed a Lyapunovbased robust control algorithm which compensates for energy absorber time delay and evaluated vibration control performance of an MR seat suspension. McManus et al. [8] investigated the use of MR seat suspensions to reduce the incidence and severity of end-stop impacts, showing impressive end-stop impact attenuation performance and reduced vibration exposure levels. Recently, Choi and Wereley [1] analyzed the biodynamic response of the human body protected by a controlled MR rotorcraft seat suspension to both sinusoidal vibration and shock loads, and showed that the MR suspension had better performance than a passive hydraulic seat suspension. In the present study, a control algorithm is presented through which a magnetorheological energy absorber may be used in a crew seat suspension to automatically accommodate occupants of varying weight.

Iterative Response Surface Based Optimization Scheme for Transonic Airfoil Design

Abstract: An improved response surface based optimization technique is presented for two-dimensional airfoil design at transonic speed. The method is based on an iterative scheme where least-square fitted quadratic polynomials of objective function and constraints are repeatedly corrected locally, about the current minimum, by adding the actual function value to the data set used to construct the polynomials. When no further cost function reduction is achieved, the design domain upon which the optimization is initially performed is changed, preserving its initial size, by updating the center point with the position of the last minimum found. The optimization is then conducted by using the same approximations built over the initial design space, which are again iteratively corrected until convergence to a given tolerance. To construct the response surfaces, the design space is explored by using a uniform Latin hypercube, aiming at reducing the bias error, in contrast with previous techniques based on D-optimality criterion. The geometry is modeled by using the PARSEC parameterization

CFD-Based Models of Aerodynamic Gust Response

Abstract: The paper presents a novel approach to modeling unsteady aerodynamic forces in response to traveling gust excitation. Time domain convolution, Auto-Regressive Moving Average (ARMA), and state-space models are presented, that are identified based on CFD input-output data. These models are compact and computationally efficient, and reproduce the CFD response highly accurately in the subsonic flow regime. The proposed models can be straightforwardly integrated into the time-domain formulation of the aeroelastic equation of motion in response to gust excitation. They can be used for both discrete and continuous gust response, as presented in the manuscript.

Riveting Process Induced Residual Stresses Around Solid Rivets in Mechanical Joints

Abstract: The interference fit provided by solid rivets introduces a residual stress field beneficial to the fatigue life of riveted joints. Evolution in riveting technology has led to force-controlled riveters which provide greater consistency over the rivet installation process and the resulting residual stress field. By reexamining the rivet installation process and its effects on the formation of residual stresses, the fatigue benefits of rivets could be further exploited. Using a 3-D finite element model, installation of universal and countersunk rivets in monolithic aluminum sheet has been studied. Aspects of accepted riveting practice, including the degree of rivet flushness and the rivet squeeze force were found to play significant roles in the formation of residual stresses. Residual stresses beneath the rivet head were also found to be influenced primarily by through-thickness compression of the joined sheets during riveting, challenging the traditional analogy of riveting to radial expansion processes.

Experimental Aerodynamic Study of Tandem Flapping Membrane Wings

Abstract: A systematic series of wind-tunnel tests was conducted on an ornithopter configuration consisting of two sets of symmetrically flapping wings, located one behind the other in tandem. It was discovered that the tandem arrangement can give thrust and efficiency increases over a single set of flapping wings for certain relative phase angles and longitudinal spacing between the wing sets. In particular, close spacing on the order of 1 chord length is generally best, and phase angles of approximately 0 ± 50 deg give the highest thrusts and propulsive efficiencies. Asymmetrical flapping was also studied, which consists of the two sets of wings rocking relative to one another 180 deg out of phase. It was found that the performance of such an arrangement is poor, relative to the best performing symmetrical tandem flapping.

Thermal Buckling and Nonlinear Flutter Behavior of Functionally Graded Material Panels

Abstract: The nonlinear flutter and thermal buckling of an functionally gradient material panel under the combined effect of elevated temperature conditions and aerodynamic loading is studied. A nonlinear finite element model based on the first-order shear deformable plate theory and von Karman strain-displacement relations is adopted. The governing nonlinear equations are obtained using the principal of virtual work, adopting an approach based on the thermal strain being a cumulative physical quantity to account for temperature-dependent material properties. The aerodynamic pressure is modeled using the quasi-steady first-order piston theory. This system of nonlinear equations is solved by the Newton-Raphson numerical technique. It is found that the temperature increase has an adverse effect on the functionally gradient material panel flutter characteristics through decreasing the critical dynamic pressure. Decreasing the volume fraction enhances flutter characteristics, but this is limited by structural integrity aspect. The presence of aerodynamic flow results in postponing the buckling temperature and in suppressing the postbuckling deflection, and the temperature increase gives way for higher limit-cycle amplitude.

Analytical two-dimensional model of a hybrid (bolted/bonded) single-lap joint

Abstract: The mechanical behavior of hybrid (bolted/bonded) joints applied to aeronautic structures is investigated. The joints under study are balanced single-lap joints, and an elastic behavior of the materials is assumed. A fully parametric analytical two-dimensional model, based on the finite element method, is presented. A special finite element, named "bonded beams" element, is computed to simulate the bonded adherends. The method and the equations to obtain the stiffness matrix of this new element are presented in this paper. The simulation of fasteners is examined through experimental and numerical approaches. The bolt load transfer is measured by the use of specially instrumented bolts, whereas a finite element 3-D model is developed to calibrate the analytical model. Good agreement is found between the analytical, experimental, and numerical results.

Aerodynamic Loading for an Airfoil with an Oscillating Gurney Flap

Abstract: A study of aerodynamic loadings on a NACA 0012 airfoil with a static and an oscillating trailing-edge Gurney flap was made. The focus is on the experimental measurement of the static and dynamic-pressure distributions on the airfoil surface. The experimental results are also correlated with theoretical results obtained using the Navier-Stokes code INS2D, developed by NASA. A Reynolds number of 348,000, a flow velocity of 20 m/s (65.6 ft/s), and a reduced frequency from 0 to 0.4 based upon half-chord b and freestream velocity U are used. The experimental results show that the effect of the static and oscillating strips located near the trailing edge of the airfoil is to enhance the maximum lift and pitching-moment coefficients for both unstalled and stalled angles of attack. An increase of the oscillating frequency also enhances the aerodynamic loading. Reasonably good agreement between the experiment and theory is obtained. The experimental results confirm the idea that an oscillating small strip located near the trailing edge can be a useful tool for active aerodynamic flow control for a wing.

Hybrid Design for Aircraft Wind-Tunnel Testing Using Response Surface Methodologies

Abstract: A response surface methodology approach to wind-tunnel testing of high-performance aircraft is being investigated at the Langley Full-Scale Tunnel. An exploratory study was completed using a newly developed response surface methodology design in an effort to better characterize an aircraft's aerodynamic behavior while simultaneously reducing test time. This new design called a "nested face-centered design" was developed when classic designs were found to have inadequate prediction qualities over a cuboidal design space with factors at five levels. A 19% scale modified X-31 aircraft model was chosen for evaluation of the new response surface methodology design based on its nonlinear aerodynamic behavior at high angle of attack that is representative of modern fighter aircraft and due to a substantial preexisting data base. A five-level nested fractional factorial design, augmented with center points and axial points, produced regression models including pure cubic terms for the characteristic aerodynamic forces and moments over a cuboidal design space as a function of model position and control surface deflections. Model adequacy and uncertainty levels were described using robust statistical methods inherent to response surface methodology practice. Comparisons to baseline data and sample lateral-directional and longitudinal aerodynamic characteristics are given as validation of the new design.

Flight Control Using Synthetic Jets on a Cessna 182 Model

Abstract: The application of active flow control via synthetic jet actuators for separation and roll control on a scaled Cessna 182 model was investigated experimentally in a low-speed wind tunnel. The model was instrumented with either ailerons or synthetic jets embedded within the outer portion of the wings' span, in lieu of ailerons. Force and moment measurements were performed for various aileron deflections and synthetic jet momentum coefficients (on either both wingtips or only on one). The sensor's rms output was monitored in real time by a computer. When the rms reached a predetermined threshold value, the computer automatically turned on the synthetic jet actuators. Using the appropriate threshold value resulted in complete avoidance of wingtip stall at the angle of attack where separation would have occurred. In addition, the shear stress sensor and wind tunnel force data were used to identify the system dynamics