Showing papers in "AIAA Journal in 2011"

Review of Output-Based Error Estimation and Mesh Adaptation in Computational Fluid Dynamics

Abstract: Error estimation and control are critical ingredients for improving the reliability of computational simulations Adjoint-based techniques can be used to both estimate the error in chosen solution outputs and to provide local indicators for adaptive refinement This article reviews recent work on these techniques for computational fluid dynamics applications in aerospace engineering The definition of the adjoint as the sensitivity of an output to residual source perturbations is used to derive both the adjoint equation, in fully discrete and variational formulations, and the adjoint-weighted residual method for error estimation Assumptions and approximations made in the calculations are discussed Presentation of the discrete and variational formulations enables a side-by-side comparison of recent work in output-error estimation using the finite volume method and the finite element method Techniques for adapting meshes using output-error indicators are also reviewed Recent adaptive results from a variety of laminar and Reynolds-averaged Navier-Stokes applications show the power of output-based adaptive methods for improving the robustness of computational fluid dynamics computations However, challenges and areas of additional future research remain, including computable error bounds and robust mesh adaptation mechanics

Performance Variations of Leading-Edge Tubercles for Distinct Airfoil Profiles

Abstract: An experimental investigation has been undertaken to determine the influence of sinusoidal leading-edge protrusions on the performance of two NACA airfoils with different aerodynamic characteristics. Force measurements on full-span airfoils with various combinations of tubercle amplitude and wavelength reveal that when compared to the unmodified equivalent, tubercles are more beneficial for the NACA 65-021 airfoil than the NACA 0021 airfoil. It was also found that for both airfoil profiles, reducing the tubercle amplitude leads to a higher maximum lift coefficient and larger stall angle. In the poststall regime, however, the performance with largeramplitude tubercles is more favorable. Reducing the wavelength leads to improvements in all aspects of lift performance, including maximum lift coefficient, stall angle, and poststall characteristics. Nevertheless, there is a certain point at which further reduction in wavelength has a negative impact on performance. The results also suggest that tubercles act in a manner similar to conventional vortex generators.

Parameter-Free Simple Low-Dissipation AUSM-Family Scheme for All Speeds

Abstract: This paper presents a new, simple low-dissipation numerical flux function of the AUSM-family for all speeds, called the simple low-dissipation AUSM. In contrast with existing all-speed schemes, the simple low-dissipation AUSM features low dissipation without any tunable parameters in a low Mach number regime while it keeps the robustness of the AUSM-family fluxes against shock-induced anomalies at high Mach numbers (e.g., carbuncle phenomena). Furthermore, the simple low-dissipation AUSM has a simpler formulation than the other all-speed schemes. These advantages of the present scheme are demonstrated in numerical examples of a wide spectrum of Mach numbers.

Aeroelastic and Aerothermoelastic Analysis in Hypersonic Flow: Past, Present, and Future

Abstract: H YPERSONIC flight began in February 1949 when a WAC Corporal rocket was ignited from a U.S.-captured V-2 rocket [1]. In the six decades since this milestone, there have been significant investments in the development of hypersonic vehicle technologies. The NASA X-15 rocket plane in the early 1960s represents early research toward this goal [2,3]. After a lull in activity, the modern era of hypersonic research started in the mid-1980s with the National Aerospace Plane (NASP) program [4], aimed at developing a single-stage-to-orbit reusable launch vehicle (RLV) that used conventional runways. However, it was canceled due mainly to design requirements that exceeded the state of the art [1,5]. A more recent RLV project, the VentureStar program, failed during structural tests, again for lack of the required technology [5]. Despite these unsuccessful programs, the continued need for a low-cost RLV, as well as the desire of the U.S. Air Force (USAF) for unmanned hypersonic vehicles, has reinvigorated hypersonic flight research. An emergence of recent and current research programs [6] demonstrate this renewed interest. Consider, for example, the NASA Hyper-X experimental vehicle program [7], the University of Queensland HyShot program [8], the NASA Fundamental Aeronautics Hypersonics Project [9], the joint U.S. Defense Advanced Research Projects Administration (DARPA)/USAF Force Application andLaunch fromContinentalUnited States (FALCON) program [10], the X-51 Single Engine Demonstrator [11,12], the joint USAF Research Laboratory (AFRL)/Australian Defence Science and Technology Organisation Hypersonic International Flight Research Experimentation project [13], and ongoing basic hypersonic research at the AFRL (e.g., [14–20]). The conditions encountered in hypersonic flows, combined with the need to design hypersonic vehicles, have motivated research in the areas of hypersonic aeroelasticity and aerothermoelasticity. It is evident from Fig. 1 that hypersonic vehicle configurations will consist of long, slender lifting body designs. In general, the body, surface panels, and aerodynamic control surfaces are flexible due to minimum-weight restrictions. Furthermore, as shown in Fig. 2, these

Dynamic Response of Highly Flexible Flying Wings

Abstract: *† This paper presents a method to model the coupled nonlinear flight dynamics and aeroelasticity of highly flexible flying wings, as well as analyze their nonlinear characteristics. A low-order, nonlinear, strain-based finite element framework is used, which is capable of assessing the fundamental impact of structural nonlinear effects in a computationally effective formulation target for preliminary vehicle design and control synthesis. The crosssectional stiffness and inertia properties of the wings are calculated along the wing span, and then incorporated into the 1-D nonlinear beam model. A proposed model for the effects in the torsional stiffness of skin wrinkling due to large bending curvature of the wing is also presented. Finite-state unsteady subsonic aerodynamic loads are incorporated to complete the aeroelastic representation of a flying wing. In studying flying wing dynamic response, a spatially-distributed discrete gust model is introduced and its impact on the time-domain solutions is investigated.

Metamodeling Method Using Dynamic Kriging for Design Optimization

Abstract: Metamodeling has been widely used for design optimization by building surrogate models for computationally intensive engineering application problems. Among all the metamodeling methods, the kriging method has gained significant interest for its accuracy.However, in traditional krigingmethods, themean structure is constructed using a fixed set of polynomial basis functions, and the optimization methods used to obtain the optimal correlation parameter may not yield an accurate optimum. In this paper, a new method called the dynamic kriging method is proposed to fit the true model more accurately. In this dynamic kriging method, an optimal mean structure is obtainedusing thebasis functions that are selected bya genetic algorithm from the candidate basis functions based on a new accuracy criterion, and a generalized pattern search algorithm is used to find an accurate optimum for the correlation parameter. The dynamic kriging method generates a more accurate surrogate model than other metamodeling methods. In addition, the dynamic kriging method is applied to the simulation-based design optimization with multiple efficiency strategies. An engineering example shows that the optimal design obtained by using the surrogate models from the dynamic kriging method can achieve the same accuracy as the one obtained by using the sensitivity-based optimization method.

High-Speed Boundary-Layer Instability: Old Terminology and a New Framework

Abstract: The discrete spectrum of disturbances in high-speed boundary layers is discussed with emphasis on singularities caused by synchronization of the normal modes. Numerical examples illustrate different spectral structures and jumps from one structure to another with small variations of basic flow parameters. It is shown that this singular behavior is due to branching of the dispersion curves in the synchronization region. Depending on the locations of the branch points, the spectrum contains an unstable mode or two. In connection with this, the terminology used for instability of high-speed boundary layers is clarified. It is emphasized that the spectrum branching may cause difficulties in stability analyses based on traditional linear stability theory and parabolized stability equations methods. Multiple-mode considerations and direct numerical simulations are needed to clarify this issue.

Impact of Fluid-Thermal-Structural Coupling on Response Prediction of Hypersonic Skin Panels

Abstract: DOI: 10.2514/1.J050617 The goal of the United States Air Force to field durable platforms capable of sustained hypersonic flight and responsive access to space depends on the ability to predict the response and the life of structures under combined aerothermal andaeropressure loading. However,current predictive capabilities are limitedfor these conditions due in part to the inability to seamlessly address fluid-thermal-structural interactions. This study aims to quantify the significance of a frequently neglected interaction, namely: the mutual coupling of structural deformation and aerodynamic heating, on response prediction. The quasi-static response of a carbon–carbon skin panel is investigated. It is found that the significance of this coupling depends largely on the in-plane boundary conditions, since increasing resistance to thermal expansion results in buckling and increasing deflections into the flow. Including these deformations in aerodynamic heating results in O10% increase in peak temperature and O100% increase in surface ply failure index for deflections O1% of panel length. In these cases, the locations of peaktemperaturesandstressesaresignificantlyaltered.Finally,neglectingdeformationsintheaeroheatinganalysis results in the prediction of snap-through for a gradual heating trajectory, whereas, inclusion leads to a higher mode dominated, dynamically stable response.

Numerical Analysis of an Oscillating-Wing Wind and Hydropower Generator

Abstract: The extraction of energy from wind or water streams is generally accomplished by means of rotary systems. However, it is recognized and it has been demonstrated that oscillating wings can also be used for this purpose. A newly developed oscillating-wing wind and hydropower generator is described. Its potential for the generation of electric power from tidal flows and high-altitude jet streams is studied using two-dimensional Navier―Stokes simulations at Re = 20, 000. Results for a single NACA 0014 wing power generator undergoing nonsinusoidal pitch― plunge motion indicate around 17 % increase in power generated and around 15 % increase in efficiency over that for sinusoidal motion. Two airfoils operating in tandem, undergoing both sinusoidal and nonsinusoidal motions, are also studied. It is found that for sinusoidal motion both averaged power output and efficiency per foil are reduced by around 20% for tandem configurations compared with a single foil in sinusoidal motion, and similar performance reductions are experienced for nonsinusoidal motions.

Well-Posed Boundary Condition for Acoustic Liners in Straight Ducts with Flow

Abstract: DOI: 10.2514/1.J050723 The Myers boundary condition for acoustics within flow over an acoustic lining has been shown to be ill-posed, leading to numerical stability issues in the time domain and mathematical problems with stability analyses. This paper gives a modification (for flat or cylindrical straight ducts) to make the Myers boundary condition well posed, and indeed more accurate, by accounting for a thin inviscid boundary layer over the lining and correctly deriving the boundary condition to first order in the boundary-layer thickness. The modification involves two integraltermsovertheboundarylayer.The firstmaybewrittenintermsofthemass,momentum,andkinetic-energy thicknessesoftheboundarylayer,whichareshownto physicallycorrespondtoamodifiedboundarymass,modified grazingvelocity,andatensionalongtheboundary. Thesecondintegral termisrelatedtothecriticallayerwithinthe boundary layer. A time domain version of the new boundary condition is proposed, although not implemented. The modified boundary condition is validated against high-fidelity numerical solutions of the Pridmore-Brown equation for sheared inviscid flow in a cylinder. Absolute instability boundaries are given for certain examples, though convective instabilities appear to always be present at certain frequencies for any boundary-layer thickness.

Introducing Loading Uncertainty in Topology Optimization

Abstract: Uncertainty is an important consideration in structural design and optimization to produce robust and reliable solutions. This paper introduces an efficient and accurate approach to robust structural topology optimization. The objective is to minimize expected compliance with uncertainty in loading magnitude and applied direction, where uncertainties are assumed normally distributed and statistically independent. This new approach is analogous to a multiple load case problem where load cases and weights are derived analytically to accurately and efficiently compute expected compliance and sensitivities. Illustrative examples using a level-set-based topology optimization method are then used to demonstrate the proposed approach.

Modeling and Simulation of Nonlinear Dynamics of Flapping Wing Micro Air Vehicles

Abstract: included. Simulations are compared with previous modeling efforts, which neglected the wings’ mass and the associated inertial coupling effects on the body. Simulations show a qualitative consistency for the nonlinear model with wing effects when different aerodynamic models are chosen as inputs. Simulation results show a significant differenceinthemodelbehaviorwhenthemassofthewings,initiallysetat5.7%ofthebodymass,isincludedversus whenthemassisneglected.Asthemassofthewingsisdecreased,thesimulationresultsofthemodelwithwingeffects approachtheresultswhenthestandardaircraftmodelisused.Simulationsleadtotheconclusionthatthemasseffects of the wings are important for dynamics, stability, and control analyses.

Near-Wall Formulation of the Partially Averaged Navier-Stokes Turbulence Model

Abstract: The variable-resolution partially averaged Navier-Stokes bridging strategy is applied to the four-equation k-epsilon-zeta-f turbulence model. In this approach, the popular two-equation model is enhanced with an additional transport equation for the velocity scale ratio zeta and an equation for the elliptic relaxation function f for the purpose of improved near-wall behavior. By using the elliptic relaxation technique to model the wall blocking effect, the new four-equation partially averaged Navier-Stokes model retains the simplicity of the previous two-equation partially averaged Navier-Stokes versions but significantly improves predictions in the near-wall region. The proposed partially averaged Navier-Stokes k-epsilon-zeta-f model is evaluated in a turbulent channel flow and flow around a three-dimensional circular cylinder mounted vertically on a flat plate. The results clearly show benefits of the improved near-wall modeling and extend partially averaged Navier-Stokes applicability to a broader range of smooth bluff-body separated flows.

Turbulent Wall Model for Immersed Boundary Methods

Abstract: The paper describes the development of a wall model to extend the applicability of Immersed Boundary methods to high Reynolds number flows. A two-layer approach, based on a decomposition of the near-wall region, is adopted. An outer region is governed by the compressible RANS equations which are solved numerically by using a classical finite volume method. In the proximity of the wall, an inner zone is established and modelled by a simplified version of the thin-boundary-layer equations. The simulation platform is based on Cartesian meshes and an immersed boundary technique. It is able to solve the steady Euler/RANS equations in two- and three-dimensional coordinates. The robustness and the accuracy of the methodology are discussed. At present this work represents the last advance of a research activity whose final goal is a fast pre-design tool for aeronautical/industrial applications.

Multifidelity surrogate modeling of experimental and computational aerodynamic data sets

Abstract: This study presents a multifidelity surrogate modeling approach, combining experimental and computational aerodynamic data sets. A multifidelity cokriging regression surrogate model is used. This study highlights how lowfidelity data from computations contribute to improving surrogate models built with limited high-fidelity data from experiments. Various types of sampling design for low fidelity data are also examined to study the impact of characteristics of the sampling design on the final surrogate models. Replication, blocking, and randomization techniques originally developed for design of experiments are used to minimize random and systematic errors. Surrogate models representing the performance of an inverted wing with counter-rotating vortex generators in ground effect are constructed, where design variables of the wing ride height and incidence and the response of sectional downforce are examined. A cokriging regression containing 12 experimental and 25 computational data points sampled with a Latin hypercube design shows the best performance here, capturing general characteristics of the target map well.

Lift enhancement by means of small-amplitude airfoil oscillations at low Reynolds numbers

Abstract: Force and particle image velocimetry measurements were conducted on a NACA 0012 airfoil undergoing small-amplitude sinusoidal plunge oscillations at a poststall angle of attack and Reynolds number of 10,000. With increasing frequency of oscillation, lift increases and drag decreases due to the leading-edge vortices shed and convected over the suction surface of the airfoil. Within this regime, the lift coefficient increases approximately linearly with the normalized plunge velocity. Local maxima occur in the lift coefficient due to the resonance with the most unstable wake frequency, its subharmonic and first harmonic, producing the most efficient conditions for high-lift generation. At higher frequencies, a second mode of flowfield occurs. The leading-edge vortex remains nearer the leading edge of the airfoil and loses its coherency through impingement with the upward-moving airfoil. To capture this impingement process, high-fidelity computational simulations were performed that showed the highly transitional nature of the flow and a strong interaction between the upper and lower-surface vortices. A sudden loss of lift may also occur at high frequencies for larger amplitudes in this mode.

Low-Reynolds-Number Aerodynamics of a Flapping Rigid Flat Plate

Abstract: Two- and three-dimensional low-aspect-ratio (AR = 4) hovering airfoil/wing aerodynamics at a low Reynolds number (Re = 100) are numerically investigated. Regarding fluid physics, in addition to the well-known leading-edge vortex and wake-capture mechanisms, a persistent jet, induced by the shed vortices in the wake during previous strokes, and tip vortices can significantly influence the lift and power performance. While in classical stationary wing theory the tip vortices are seen as wasted energy, here, they can interact with the leading-edge vortex to contribute to the lift generated without increasing the power requirements. Using surrogate modeling techniques, the two- and three-dimensional time-averaged aerodynamic forces were predicted well over a large range of kinematic motions when compared with the Navier-Stokes solutions. The combined effects of tip vortices, leading-edge vortex, and jet can be manipulated by the choice of kinematics to make a three-dimensional wing aerodynamically better or worse than an infinitely long wing. The environmental sensitivity during hovering for select kinematics is also examined. Different freestream strengths and orientations are imposed, with the impact on vortex generation and wake interaction investigated.

Experimental Investigation of the Unstart Process of a Generic Hypersonic Inlet

Abstract: To provide information for the detection, prediction, and control of the inlet unstart, the entire process from a started status to an unstarted status of a generic two-dimensional hypersonic inlet is studied experimentally at Mach 5. The movement of a flow plug at the exit of the duct is used to gradually increase the throttling to simulate the unstart process caused by the excessive heat release in the combustor. Simultaneous high-speed schlieren imaging and dynamic surface pressure measurements are used to record the unsteady flow structures and surface pressures of the unstart process. According to the internal and external flows, the unstart process can be divided into four stages, namely, shock train in the combustor, shock train in the isolator, separation bubble in the throat, and unstart. The transient flow patterns of each stage are substantially different and the corresponding dynamics pressures also have prominent time-frequency features, which make the detection and prediction of the inlet unstart based on dynamic pressures possible. Since the sensor placed at the end of the top surface of the inlet contract part, marked by C1, obtains the most abundant time-frequency characteristics which can be used to discern the different stages of the inlet unstart process, the location where it stays is regarded as the first choice for sensor placement to construct a practical inlet-status-monitoring system.

Reduced-Order Aerothermoelastic Framework for Hypersonic Vehicle Control Simulation

Abstract: Hypersonic vehicle control system design and simulation require models that contain a low number of states. Modeling of hypersonic vehicles is complicated due to complex interactions between aerodynamic heating, heat transfer, structural dynamics, and aerodynamics. Although there exist techniques for analyzing the effects of each of the various disciplines, thesemethods often require solution of large systems of equations, which is infeasible within a control design and evaluation environment. This work presents an aerothermoelastic framework with reducedorder aerothermal, heat transfer, and structural dynamicmodels for time-domain simulation of hypersonic vehicles. Details of the reduced-order models are given, and a representative hypersonic vehicle control surface used for the study is described. Themethodology is applied to a representative structure to provide insight into the importance of aerothermoelastic effects on vehicle performance. The effect of aerothermoelasticity on total lift and drag is found to result in up to an 8% change in lift and a 21% change in drag with respect to a rigid control surface for the four trajectories considered. An iterative routine is used to determine the angle of attack needed to match the lift of the deformed control surface to that of a rigid one at successive time instants.Application of the routine todifferent cruise trajectories shows a maximum departure from the initial angle of attack of 8%.

Silent Owl Flight: Bird Flyover Noise Measurements

Abstract: Mostgeneraofowls(Strigiformes)havetheabilityto flysilently.Themechanismsofthesilent flightoftheowlhave been the subject of scientific interest for many decades. The results from studies in the past are discussed in detail in thispaperandtherationaleforthepresentresearchisgiven,whichincluded flyovernoisemeasurementsondifferent species of birds. Successful acoustic measurements were made on a Common Kestrel, a Harris Hawk, and a Barn Owl. Measurements on three other birds did not lead to reliable results. The setup and procedure used for the outdoor measurements are discussed. These include the estimation of the trajectory from dual video camera recordings and microphone-array measurements with a moving-focus beamforming technique. The main result from the 50 successful flyovers is that the owl flight produces aerodynamic noise that is indeed a few decibels below thatofotherbirds,evenif flyingatthesamespeed.Thisnoisereductionissignificantatfrequenciesabove1.6kHz.At frequencies above 6.3 kHz the noise from the owl remains too quiet to be measured.

Improving Low-Frequency Characteristics of Recycling/Rescaling Inflow Turbulence Generation

Abstract: For the study of turbulent flows with low-frequency dynamics (e.g. shock-wave/boundary-layer interactions), it is desirable that the inflow turbulence does not contaminate the solution with spurious spatiotemporal correlations introduced by the mechanism of inflow turbulence generation. To investigate the creation and mitigation of these adverse low-frequency effects, large-eddy simulation of a Mach 2.28 boundary layer over an adiabatic flat plate is carried out using a typical recycling/rescating procedure. Spurious temporal autocorrelations and energy spectral peaks are observed associated with the recycling frequency and its harmonics. Comparisons are made with common "synthetic" turbulence-generation techniques, and improvements to the standard recycling/rescaling procedure are suggested to substantially reduce or eliminate the inherent low-frequency contamination. It is found that by applying a nonconstant reflection or translation operation to the recycled turbulence plane at randomly-distributed time intervals, one is able to maintain realistic turbulence without low-frequency contamination.

Flutter and Stall Flutter of a Rectangular Wing in a Wind Tunnel

Abstract: The aeroelastic behavior of a rectangular wing with pitch and plunge degrees of freedom was observed experimentally using pressure, acceleration, and particle image velocimetry measurements. The wing was set at different static angles of attack and wind-tunnel airspeeds. The wing’s dynamic behavior was governed by a twoparameter bifurcation from steady to limit cycle oscillations, with the two parameters being the airspeed and the static angle of attack. At the lowest static angle, the wing underwent a classical flutter phenomenon that was transformed into a supercriticalHopf bifurcation at higher angles. The latterwas combinedwith a fold bifurcation at intermediate angles of attack. All limit cycle oscillations observed were either low-amplitude oscillations with timevarying amplitude or high-amplitude oscillations with nearly steady amplitude. They were caused by two different types of dynamic stall phenomena. During low-amplitude limit cycle oscillations the periodically stalled flow covered only the rear part of the wing. During high-amplitude limit cycle oscillations, trailing-edge and leading-edge separation occurred. Trailing-edge separation was characterized by a significant amount of unsteadiness, varying visibly from cycle to cycle. The occurrence of leading-edge separation was muchmore regular and had the tendency to stabilize the amplitude of the limit cycle oscillation motion.

Numerical Investigation of Plasma-Based Control for Low-Reynolds-Number Airfoil Flows

Abstract: Large-eddy simulations are carried out to investigate the use of plasma-based actuation for the control of flows over a finite span wing at low Reynolds numbers. The wing section corresponds to the SD7003 airfoil, which is representative of those employed for micro air vehicle applications. Dielectric-barrier-discharge plasma actuators are used to modify the transitional flow and improve aerodynamic performance. Solutions are obtained to the Navier-Stokes equations, which were augmented by source terms used to represent plasma-induced body forces imparted by the actuators on the fluid. Simple phenomenological models provided the body forces generated by the electric field of the plasma surrounding the actuators. The numerical method is based upon a high-fidelity time-implicit scheme, an implicit large-eddy-simulation approach, and domain decomposition in order to perform calculations on a parallel computing platform. Flow at a chord-based Reynolds number of 40,000 is considered in the investigation, which is characterized by laminar separation on the suction surface of the wing at low angles of attack. This separation then promotes transition to a more complex state, which can be modified by the use of plasma actuation. Several aspects of control are examined, including different actuator configurations, alternative plasma-force models, both continuous and pulsed modes of operation, and the magnitude of plasma force required for control.

Structural health monitoring of composite structures with piezoelectric wafer active sensors

Abstract: Piezoelectric wafer active sensors (PWAS) are lightweight and inexpensive enablers for a large class of structural health monitoring (SHM) applications such as: (a) embedded guided-wave ultrasonics, i.e., pitch-catch, pulseecho, phased arrays; (b) high-frequency modal sensing, i.e., the electro-mechanical (E/M) impedance method; (c) passive detection (acoustic emission and impact detection). The focus of this paper will be on the challenges and opportunities posed by the composite structures as different from the metallic structures on which this methodology was initially developed. After a brief introduction, the paper discusses damage modes in composites. Then, it reviews the PWAS-based SHM principles. It follows with a discussion of guided wave propagation in composites and PWAS tuning effects. Finally, the paper presents some experimental results with damage detection in composite specimens. The paper ends with conclusions and suggestions for further work

Hybrid Source Model for Predicting High-Speed Jet Noise

Abstract: This paper introduces a novel hybrid source model into an existing acoustic analogy approach to obtain improved predictions of the turbulent mixing noise from cold, round, subsonic, and supersonic jets. The model incorporates new features of the Reynolds stress autocovariance tensor components found in recent experiments. The model parameters are determined from a Reynolds-averaged Navier―Stokes flow solution and experimental data. It is shown that this model significantly improves the predictions relative to previous results, particularly at observer polar angles between 90 degrees to the jet axis and the peak noise direction, indicating the importance of properly modeling relatively subtle characteristics of the autocovariance functions. The results are used to infer the relative importance of individual terms that make up the formula for the acoustic spectrum as a function of jet Mach number, frequency, and observer location.

Closed-Loop Control of Lift for Longitudinal Gust Suppression at Low Reynolds Numbers

Abstract: Experiments are conducted to investigate the ability of variable-pressure pulsed-blowing actuation to maintain a constant lift force on a low-aspect-ratio semicircular wing in a longitudinally gusting flow. Dynamic models of the lift response to actuation and the lift response to longitudinal gusting are obtained through modern system identification methods. Robust closed-loop controllers are synthesized using a mixed-sensitivity loop-shaping approach. An additional feedforward disturbance compensator is designed based on a model of the unsteady aerodynamics. The controllers show suppression of lift fluctuations at low gust frequencies, f < 0.8 Hz(reduced frequency, k < 0.09). At higher frequencies, the control performance degrades due to limitations related to the time for a disturbance, created by the actuators, to convect over the wing and establish the flowfield that leads to enhanced lift on the wing.

Investigations of Lift-Based Pitch-Plunge Equivalence for Airfoils at Low Reynolds Numbers

Abstract: The limits of linear superposition in two-dimensional high-rate low-Reynolds-number aerodynamics are examined by comparing the lift-coefficient history and flowfield evolution for airfoils undergoing harmonic motions in pure pitch, pure plunge, and pitch―plunge combinations. Using quasi-steady airfoil theory and Theodorsen's formula as predictive tools, pitching motions are sought that produce lift histories identical to those of prescribed plunging motions. It follows that a suitable phasing of pitch and plunge in a combined motion should identically produce zero lift, canceling either the circulatory contribution (with quasi-steady theory) or the combination of circulatory and noncirculatory contributions (with Theodorsen's formula). Lift history is measured experimentally in a water tunnel using a force balance and is compared with two-dimensional Reynolds-averaged Navier―Stokes computations and Theodorsen's theory; computed vorticity contours are compared with dye injection in the water tunnel. Theodorsen's method evinces considerable, and perhaps surprising, resilience in finding pitch-to-plunge equivalence of lift-coefficient―time history, despite its present application to cases in which its mathematical assumptions are demonstrably violated. A combination of pitch and plunge motions can be found such that net lift coefficient is nearly identically zero for arbitrarily high reduced frequency, provided that amplitude is small. Conversely, cancellation is possible at large motion amplitude, provided that reduced frequency is moderate. The product of Strouhal number and nondimensional amplitude is therefore suggested as the upper bound for when superposition and linear predictions remain valid in massively unsteady two-dimensional problems.

Aerodynamic Characteristics at Low Reynolds Numbers for Wings of Various Planforms

Abstract: The aerodynamic characteristics of various wing planforms at low Reynolds numbers (about 1 x 10 4 ) were studied by conducting wind-tunnel tests. These low Reynolds numbers correspond to the flights of small creatures, such as insects. Elliptical, rectangular, and triangular planforms with various aspect ratios were used in this study, as well as a swept rectangular (parallelogram) wing with an aspect ratio of four. The wing sections of all models were thin rectangular airfoils. The aerodynamic forces (lift and drag) and the pitching moment acting on the wing were measured for a wide range of angles of attack (including the maximum of 90 deg). Nonlinear characteristics of the lift coefficient were obtained, even at low angles of attack for high-aspect-ratio wings, whereas a small lift slope and a large maximum lift coefficient were obtained for low-aspect-ratio wings. The drag and the pitching moment coefficients also exhibited nonlinear characteristics. The effect of the Reynolds number based on the wing chord was comparatively small, but a distinctive phenomenon in low-Reynolds-number flow was observed in flow visualization using oil-film and smoke-wire methods.