Showing papers on "Pitching moment published in 2011"

Development of the D8 Transport Configuration

Abstract: The D8 jet transport configuration is presented. Focus is on the aerodynamic characteristics of its distinguishing features: the wide “double-bubble” fuselage with beneficial pitching moment and carryover lift characteristics, the nearly-unswept wing and reduced cruise Mach number, and the boundary layer ingesting rear engine installation together with the twin “pi-tail” configuration. The merit of each feature is evaluated in terms of mission fuel burn, calculated via an MDO optimizer which accounts for the weight and propulsive efficiency of the design features. Evaluations via panel methods are also performed, and wind tunnel test data from a 20:1 low speed model is also presented. The calculations and tunnel data indicate that the advantages of the D8 configuration would lead to a very large 33% fuel burn reduction compared to an optimized conventional transport configuration with the same materials and engine technology.

Experimental Investigations of the NASA Common Research Model in the NASA Langley National Transonic Facility and NASA Ames 11-Ft Transonic Wind Tunnel (Invited)

Abstract: Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility and the NASA Ames 11-ft wind tunnel. Data have been obtained at chord Reynolds numbers of 5 million for five different configurations at both wind tunnels. Force and moment, surface pressure and surface flow visualization data were obtained in both facilities but only the force and moment data are presented herein. Nacelle/pylon, tail effects and tunnel to tunnel variations have been assessed. The data from both wind tunnels show that an addition of a nacelle/pylon gave an increase in drag, decrease in lift and a less nose down pitching moment around the design lift condition of 0.5 and that the tail effects also follow the expected trends. Also, all of the data shown fall within the 2-sigma limits for repeatability. The tunnel to tunnel differences are negligible for lift and pitching moment, while the drag shows a difference of less than ten counts for all of the configurations. These differences in drag may be due to the variation in the sting mounting systems at the two tunnels.

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.

Aerodynamic effects of corrugation in flapping insect wings in hovering flight

Abstract: We have examined the aerodynamic effects of corrugation in model insect wings that closely mimic the wing movements of hovering insects. Computational fluid dynamics were used with Reynolds numbers ranging from 35 to 3400, stroke amplitudes from 70 to 180 deg and mid-stroke angles of incidence from 15 to 60 deg. Various corrugated wing models were tested (care was taken to ensure that the corrugation introduced zero camber). The main results are as follows. At typical mid-stroke angles of incidence of hovering insects (35-50 deg), the time courses of the lift, drag, pitching moment and aerodynamic power coefficients of the corrugated wings are very close to those of the flat-plate wing, and compared with the flat-plate wing, the corrugation changes (decreases) the mean lift by less than 5% and has almost no effect on the mean drag, the location of the center of pressure and the aerodynamic power required. A possible reason for the small aerodynamic effects of wing corrugation is that the wing operates at a large angle of incidence and the flow is separated: the large angle of incidence dominates the corrugation in determining the flow around the wing, and for separated flow, the flow is much less sensitive to wing shape variation. The present results show that for hovering insects, using a flat-plate wing to model the corrugated wing is a good approximation.

Computational Drag Prediction of the DPW4 Configuration Using the Far-Field Approach

Abstract: This paper presents the computational studies done at ONERA in the context of the 4th AIAA CFD Drag PredictionWorkshop (DPW4). Furthermore, it gives a detailed description of the far-field methods developed in the Applied Aerodynamics Department. Concerning the DPW4 configuration, a grid convergence study and a downwash study are proposed. Then the effects due toMach andReynolds numbers variations are quantified. All the multiblock structured grids used in this work have been provided by The Boeing Company to the drag prediction community. All the Reynolds-averaged Navier–Stokes computations are performed by using the ONERA-elsA solver with the Spalart–Allmaras turbulence model, and the solutions are postprocessed with the ONERA-ffd72 far-field drag-extraction tool. Concerning drag predictions, a very good agreement has been observed between ONERA-elsA results and the near-field drag coefficients (pressure and friction) computed by other DPW4 participants such asBoeing orAirbus.Moreover, the far-field software ffd72 givesONERA the singular capability to determine the values of the different physical drag components (viscous, wave, and lift-induced productions). Concerning the pitching moment, ONERA results are very close to Boeing, Airbus, or DLR, German Aerospace Center predictions.

Aerodynamics Analysis of Small Horizontal Axis Wind Turbine Blades by Using 2D and 3D CFD Modelling

Abstract: Please write a brief description of your work, or copy an abstract you have included in the Thesis Wind power is one of the most important sources of renewable energy. Wind-turbines extract kinetic energy from the wind. Currently much research has concentrated on improving the aerodynamic performance of wind turbine blade through wind tunnel testing and theoretical studies. These efforts are much time consuming and need expensive laboratory resources. However, wind turbine simulation through Computational Fluid Dynamics (CFD) software offers inexpensive solutions to aerodynamic blade analysis problem. In this study, two-dimensional aerofoil (i.e. DU-93 and NREL-S809) CFD models are presented using ANSYS-FLUENT software. Using the Spalart-Allmaras turbulent viscosity, the dimensionless lift, drag and pitching moment coefficients were calculated for wind-turbine blade at different angles of attack. These CFD model values we then validated using published calibrated lift and drag coefficients evident in the literature. Optimum values of these coefficients as well as a critical angle were found from polar curves of lift, drag and moment modelling data. These data were exploited in order to select the aerofoil with best aerodynamic performance for basis of a three-decisional model analogue. Thereafter a three-dimensional CFD model of small horizontal axis wind-turbine was produced. The numerical solution was carried out by simultaneously solving the three-dimensional continuity, momentum and the Naveir-Stokes equations in a rotating reference frame using a standard non-linear k-ω solver so that the rotational effect can be studied. These three-dimensional models were used for predicting the performance of a small horizontal axis wind turbine. Moreover, the analysis of wake effect and aerodynamic noise can be carried out when the rotational effect was simulated.

Numerical investigation of air jets for dynamic stall control on the OA209 airfoil

Abstract: The design and numerical investigation of constant blowing air jets as fluidic control devices for helicopter dynamic stall control is described. Prospective control devices were first investigated using 3D RANS computations to identify effective configurations and reject ineffective configurations. Following this, URANS investigations on the dynamically pitching OA209 airfoil verified that configurations had been selected which reduced the peaks in pitching moment and drag while preserving at least the mean lift and drag from the clean wing. Two configurations using jets at 10% chord on the airfoil top were identified, and one configuration using a tangential slot at 10% chord on the airfoil top, with each configuration evaluated for two jet total pressures. For the best configuration, a reduction in the pitching moment peak of 85% and in the drag peak of 78% were observed, together with a 42% reduction in the mean drag over the unsteady pitching cycle.

Nonlinear control of a spacecraft with multiple fuel slosh modes

Abstract: This paper studies the modeling and control problem for a spacecraft with fuel slosh dynamics. A multi-pendulum model is considered for the characterization of the most prominent sloshing modes. The control inputs are defined by the gimbal deflection angle of a non-throttable thrust engine and a pitching moment about the center of mass of the spacecraft. The control objective, as is typical in a thrust vector control design, is to control the translational velocity vector and the attitude of the spacecraft, while attenuating the sloshing modes. A nonlinear mathematical model that reflects all of these assumptions is first derived. Then, a Lyapunov-based nonlinear feedback controller is designed to achieve the control objective. Finally, a simulation example is included to demonstrate the effectiveness of the controller.

Low Boom Configuration Analysis with FUN3D Adjoint Simulation Framework

Abstract: Off-body pressure, forces, and moments for the Gulfstream Low Boom Model are computed with a Reynolds Averaged Navier Stokes solver coupled with the Spalart-Allmaras (SA) turbulence model. This is the first application of viscous output-based adaptation to reduce estimated discretization errors in off-body pressure for a wing body configuration. The output adaptation approach is compared to an a priori grid adaptation technique designed to resolve the signature on the centerline by stretching and aligning the grid to the freestream Mach angle. The output-based approach produced good predictions of centerline and off-centerline measurements. Eddy viscosity predicted by the SA turbulence model increased significantly with grid adaptation. Computed lift as a function of drag compares well with wind tunnel measurements for positive lift, but predicted lift, drag, and pitching moment as a function of angle of attack has significant differences from the measured data. The sensitivity of longitudinal forces and moment to grid refinement is much smaller than the differences between the computed and measured data.

Low Boom Configuration Analysis with FUN3D Adjoint Simulation Framework (Invited)

Abstract: Off-body pressure, forces, and moments for the Gulfstream Low Boom Model are computed with a Reynolds Averaged Navier Stokes solver coupled with the Spalart-Allmaras (SA) turbulence model. This is the first application of viscous output-based adaptation to reduce estimated discretization errors in off-body pressure for a wing body configuration. The output adaptation approach is compared to an a priori grid adaptation technique designed to resolve the signature on the centerline by stretching and aligning the grid to the freestream Mach angle. The output-based approach produced good predictions of centerline and off-centerline measurements. Eddy viscosity predicted by the SA turbulence model increased significantly with grid adaptation. Computed lift as a function of drag compares well with wind tunnel measurements for positive lift, but predicted lift, drag, and pitching moment as a function of angle of attack has significant differences from the measured data. The sensitivity of longitudinal forces and moment to grid refinement is much smaller than the differences between the computed and measured data.

Effect of cavity location on combustion flow field of integrated hypersonic vehicle in near space

Abstract: The cavity has been widely employed as the flame holder to prolong the residence time of fuel in supersonic flows since it improves the combustion efficiency in the scramjet combustor, and also imposes additional drag on the engine. In this paper, the two-dimensional coupled implicit Reynolds Average Navier–Stokes equations, the RNG k–e turbulence model and the finite-rate/eddy-dissipation reaction model have been employed to numerically simulate the combustion flow field of an integrated hypersonic vehicle. The effect of cavity location on the combustion flow field of the vehicle has been investigated, and the fuel, namely hydrogen, was injected upstream of the cavity on the walls of the first stage combustor. The obtained results show that the viscous lift force, drag force and pitching moment of the vehicle are nearly unchanged by varying the cavity location over the location range and designs considered in this article, namely the configurations with single cavity, double cavities in tandem and double cavities in parallel. The variation of the fuel injection strategy affects the separation of the boundary layer, and the viscous effect on the drag force of the vehicle is remarkable, but the viscous effects on the lift force and the pitching moment are both small and they can be neglected in the design process of hypersonic vehicles. In addition to varying the location of the cavities, three fuel injection configurations were considered. It was found that one particular case can restrict the inlet unstart for the scramjet engine.

Design and implementation of an internal stress wave force balance in a shock tunnel

Abstract: A stress wave internal force balance for the High Enthalpy Shock Tunnel Gottingen (HEG) of the German Aerospace Center (DLR) to measure lift, pitching moment and drag was designed, calibrated and tested. The balance is designed to measure forces in ground based test facilities with test times in the order of milliseconds on models additionally instrumented with surface pressure and wall heat flux gauges from angles of attack of −40° to 20°. Experiments in HEG were performed on a 303 mm long, 10° half angle blunt cone at angles of attack from −20° to 0°. The tests were conducted utilizing two different operating conditions at total specific enthalpies of 3.0 and 3.5 MJ/kg and dynamic pressures of 30 and 72 kPa. The performance of the balance was assessed by comparing the measured force and moment coefficients with computational fluid dynamics (CFD) predictions.

Numerical Simulation of Gurney Flap on RAE-2822 Supercritical Airfoil

Abstract: A two-dimensional steady Reynolds-averaged Navier–Stokes equation was solved to investigate the effects of a Gurney flap on RAE-2822 (Royal Aeronautical Establishment) supercritical airfoil aerodynamic performance. The heights ofGurneyflaps range from0.25 to 3%airfoil chord lengths. The incompressible/compressibleNavier–Stokes equations were used to simulate the flow structure around the airfoils in subsonic/transonic flows, respectively, with the Spalart–Allmaras turbulence model. In comparison with the clean airfoil, the Gurney flap can significantly increase the prestall lift and lift-to-drag ratio of an RAE-2822 airfoil at a small angle of attack. Nosedown pitching moment also increased with the Gurney flap height. At both takeoff-and-landing status and cruise phase, the aerodynamic performance of the airfoil was significantly improved byGurney flaps with the height below 1%airfoil chord length. In addition, the surface pressure distribution, wake flow velocity profile, and trailing-edge flow structure of the airfoil were illustrated, which helps to understand the mechanisms of the Gurney flap to improve RAE-2822 airfoil aerodynamic performance.

Force Measurements in Hypervelocity Flows with an Acceleration Compensated Piezoelectric Balance

Abstract: A new 6-component acceleration compensated piezoelectric Force balance was designed, built and used to perform force and moment measurements on a 6-inch diameter Apollo shaped capsule in the 48-Inch tunnel at CUBRC in cold hypersonic nitrogen flows. Results were compared with CFD computations and measurements with a 3-component acceleration compensated strain gage balance under the same nominal flow conditions on the exact same aerodynamic model. Excellent agreement was found between the two methods as L/D was within 1.5% and pitching moment within 8%. Agreement with perfect gas Navier-Stokes computations was within 5% for all the measured aerodynamic coefficients. Compared to a previous strain balance gage design, the new piezoelectric balance has both frequency response and sensitivity increased by a factor of approximately four. The current design is expected to allow force measurement in the new LENS-XX expansion tunnel for test times as low as 1 ms. Simple modifications to the existing design should allow measurements for test times as low as 0.3 ms and force levels as high as 3000 lbf.

Transitory Control of Dynamic Stall on a Pitching Airfoil

Abstract: Transitory attachment of the flow over a stalled, 2-D airfoil is investigated in wind tunnel experiments using pulsed actuation. Actuation is provided by a spanwise array of momentary, combustion-based actuator jets having a characteristic time scale O[1 ms] that is an order of magnitude shorter than the convective time scale of the flow. It is shown that a single actuation pulse results in transitory flow attachment that is manifested by rapid increase in the global circulation and aerodynamic forces and persists for about ten convective time scales before the flow becomes fully stalled again. Large-scale changes in vorticity accumulation that are associated with repetitive, burst-modulated actuation pulses are exploited for significant extension of the streamwise domain and duration of the attached flow with a corresponding increase in circulation. The effects of the transitory actuation are further amplified when the airfoil is mounted on a dynamic 2DOF (pitch and plunge) traversing mechanism and the actuation is tested with pitch oscillations beyond the stall limit. In this configuration, the actuation is nominally two-dimensional within a spanwise domain measuring 0.21S that is bounded by end fences. It is shown that pulse actuation significantly increases the lift not only at post-stall but also at angles of attack that are below stall (ostensibly by trapping vorticity over the entire oscillation cycle).

Effects of ice geometry on airfoil performance using neural networks prediction

Abstract: Purpose – The purpose of this paper is to describe a methodology for predicting the effects of glaze ice geometry on airfoil aerodynamic coefficients by using neural network (NN) prediction. Effects of icing on angle of attack stall are also discussed.Design/methodology/approach – The typical glaze ice geometry covers ice horn leading‐edge radius, ice height, and ice horn position on airfoil surface. By using artificial NN technique, several NNs are developed to study the correlations between ice geometry parameters and airfoil aerodynamic coefficients. Effects of ice geometry on airfoil hinge moment coefficient are also obtained to predict the angle of attack stall.Findings – NN prediction is feasible and effective to study the effects of ice geometry on airfoil performance. The ice horn location and height, which have a more evident and serious effect on airfoil performance than ice horn leading‐edge radius, are inversely proportional to the maximum lift coefficient. Ice accretions on the after‐location...

Aerodynamic Effects of Simulated Ice Accretion on a Generic Transport Model

Abstract: An experimental research effort was begun to develop a database of airplane aerodynamic characteristics with simulated ice accretion over a large range of incidence and sideslip angles. Wind-tunnel testing was performed at the NASA Langley 12-ft Low-Speed Wind Tunnel using a 3.5 percent scale model of the NASA Langley Generic Transport Model. Aerodynamic data were acquired from a six-component force and moment balance in static-model sweeps from alpha = -5deg to 85deg and beta = -45 deg to 45 deg at a Reynolds number of 0.24 x10(exp 6) and Mach number of 0.06. The 3.5 percent scale GTM was tested in both the clean configuration and with full-span artificial ice shapes attached to the leading edges of the wing, horizontal and vertical tail. Aerodynamic results for the clean airplane configuration compared favorably with similar experiments carried out on a 5.5 percent scale GTM. The addition of the large, glaze-horn type ice shapes did result in an increase in airplane drag coefficient but had little effect on the lift and pitching moment. The lateral-directional characteristics showed mixed results with a small effect of the ice shapes observed in some cases. The flow visualization images revealed the presence and evolution of a spanwise-running vortex on the wing that was the dominant feature of the flowfield for both clean and iced configurations. The lack of ice-induced performance and flowfield effects observed in this effort was likely due to Reynolds number effects for the clean configuration. Estimates of full-scale baseline performance were included in this analysis to illustrate the potential icing effects.

Correlating CFD Simulation With Wind Tunnel Test for the Full-Scale UH-60A Airloads Rotor

Abstract: Data from the recent UH-60A Airloads Test in the National Full-Scale Aerodynamics Complex 40- by 80- Foot Wind Tunnel at NASA Ames Research Center are presented and compared to predictions computed by a loosely coupled Computational Fluid Dynamics (CFD)/Comprehensive analysis. Primary calculations model the rotor in free-air, but initial calculations are presented including a model of the tunnel test section. The conditions studied include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed into deep stall. Predictions show reasonable agreement with measurement for integrated performance indicators such as power and propulsive but occasionally deviate significantly. Detailed analysis of sectional airloads reveals good correlation in overall trends for normal force and pitching moment but pitching moment mean often differs. Chord force is frequently plagued by mean shifts and an overprediction of drag on the advancing side. Locations of significant aerodynamic phenomena are predicted accurately although the magnitude of individual events is often missed.

Using formal methods to verify safe deep stall landing of a MAV

Abstract: The various fields of application for miniature air vehicles often do not provide distinct landing areas or even require additional equipment like nets or parachutes to land the aircraft without damaging it This work introduces the deep stall landing (DSL) as a maneuver that uses the extraordinary aerodynamic characteristics of a delta wing MAV that come into effect after the angle of attack passes the stall angle This landing maneuver is modeled based on a longitudinal aerodynamic model that takes lift, drag, thrust, weight, and pitching moment into account By determining the operational modes that the aircraft has to perform in order to either complete the landing maneuver or abort it in case of a missed approach a hybrid system is identified This system contains both continuous and discrete state dynamics that model the aircraft in each landing phase Based on this hybrid system reachability analyses are performed which utilize level set methods to calculate backwards reachable sets These sets are used to identify transitions within the modeled system that bring the aircraft form one operational mode to another without leaving the safe flight envelope The final result is a discrete event system that covers all possible transitions within the refined model Based on this discrete model an autonomous system can be implemented that is able to determine whether the initiation of the landing maneuver is safe in terms of keeping the aircraft within the safe flight envelope during the whole maneuver Furthermore the results of the reachability analysis determine for which states of the aircraft it would be safe to initiate a recovery maneuver in case of a missed landing approach

Wind-tunnel tests of a tilt-rotor aircraft

Abstract: A wide aerodynamic test campaign has been carried out on the tiltrotor aircraft ERICA at the Large Wind Tunnel of Politecnico di Milano by means of a modular 1:8 scale model in order to produce a dataset necessary to better understand the aerodynamic behaviour of the aircraft and to state its definitive design. The target of the tests was the measurement of the aerodynamic forces and moments in several different configurations and different attitudes. The test program included some conditions at very high incidence and sideslip angles that typically belong to the helicopter-mode flight envelope and measurements of forces on the tail and on the tilting wings. A large amount of data has been collected that will be very useful to refine the aircraft design. In general the aircraft aerodynamics do not present any critical problems, but further optimisation is still possible. From the viewpoint of drag in the cruise configuration, the sponsons of the landing gear seem to be worth some further design refinement since they are responsible for a 20% drag increase with respect to the pure fuselage configuration. On the contrary, the wing fairing has proved to work well when the aircraft longitudinal axis is aligned with the wind, providing just a slight drag increase. Two other interesting aspects are the quite nonlinear behaviour of the side force for the intermediate sideslip angles as well as the noticeable hysteresis in the moment coefficient at very high incidence angles.

Navier-Stokes Predictions of Aerodynamic Coefficients and Dynamic Derivatives of a 0.50-cal Projectile

Abstract: m m C C q  = pitch damping moment coefficient sum  m C = pitching moment coefficient derivative  N C = normal force coefficient derivative Cn = side moment coefficient  p n C = Magnus moment coefficient derivative  n C = slope of side moment coefficient with coning rate

Theoretical solutions for unsteady compressible subsonic flows past oscillating rigid and flexible airfoils

Abstract: This paper presents simple and efficient analytical solutions for unsteady compressible subsonic flows past flexible airfoils executing low frequency oscillations. These analytical solutions are obtained with a method using velocity singularities related to the airfoil leading edge and ridges (defined by the changes in the boundary conditions). Efficient analytical solutions in closed form are presented for the unsteady lift and pitching moment coefficients and for the chordwise distribution of the unsteady pressure difference coefficient in the general case of rigid or flexible airfoils executing oscillatory rotations and normal-to-chord translations and flexural oscillations. The method has been validated for the pitching and plunging oscillations of the rigid airfoils by comparison with results based on Jordan's data for compressible flows and by comparison with the solutions obtained by Theodorsen, Postel and Leppert and Mateescu and Abdo for the limit case of incompressible flows. A detailed analysis of the variations of the aerodynamic coefficients with the Mach number and with the reduced frequency of oscillations is also presented for the cases of oscillatory pitching rotations and normal-to-chord translations of rigid airfoils and for the flexural oscillations of flexible airfoils. Efficient analytical solutions are presented for the flexural oscillations of airfoils in compressible flows, which can be efficiently used in solving aeroelastic interaction problems.

CFD Computations for a Generic High-Lift Configuration Using TetrUSS

Abstract: Assessment of the accuracy of computational results for a generic high-lift trapezoidal wing with a single slotted flap and slat is presented. The paper is closely aligned with the focus of the 1st AIAA CFD High Lift Prediction Workshop (HiLiftPW-1) which was to assess the accuracy of CFD methods for multi-element high-lift configurations. The unstructured grid Reynolds-Averaged Navier-Stokes solver TetrUSS/USM3D is used for the computational results. USM3D results are obtained assuming fully turbulent flow using the Spalart-Allmaras (SA) and Shear Stress Transport (SST) turbulence models. Computed solutions have been obtained at seven different angles-of-attack ranging from 6 -37 . Three grids providing progressively higher grid resolution are used to quantify the effect of grid resolution on the lift, drag, pitching moment, surface pressure and stall angle. SA results, as compared to SST results, exhibit better agreement with the measured data. However, both turbulence models under-predict upper surface pressures near the wing tip region.

An accelerometer balance for the measurement of roll, lift and drag on a lifting model in a shock tunnel

Abstract: A force balance to measure roll, lift and drag on a lifting aerodynamic body in an ultrashort-duration hypersonic test facility, such as a shock tunnel, has been developed and tested on a flapped, blunt-nosed, triangular lifting body at a freestream Mach number of 8. The flow total enthalpy and the freestream unit Reynolds number were 0.83 MJ kg−1 and 0.98 million, respectively. The balance structure has a soft suspension that allows the model to have a free flight during the short-duration aerodynamic test. The balance was mounted inside the hollow model and was equipped with accelerometers to sense the aerodynamic moment and forces on the model. The measurements were carried out at different angles of incidence of the model and the acquired signals of the accelerometers were reduced to the aerodynamic moment and the force coefficients based on the theories of applied mechanics and aerodynamics. Also, the moment and force coefficients were theoretically calculated based on the Newtonian theory, which is an accepted analytical approach for hypersonic bodies. Good agreement has been observed between the experimental and the analytical results. The method of measurement of roll and lift, and the data on the rolling moment of a lifting body presented in this note are novel.

Unsteady Subsonic Aerodynamic Characteristics of Wing in Fold Motion

Abstract: Aerodynamic characteristics of a wing during fold motion were investigated in order to understand how variations or changes in such characteristics increase aircraft performance. Numerical simulations were conducted, and the results were obtained using the unsteady vortex lattice method to estimate the lift, drag and the moment coefficient in subsonic flow during fold motion. Parameters such as the fold angle and the fold angular velocity were summarized in detail. Generally, the lift and pitching moment coefficients decreased as the angle increased. In contrast, the coefficients increased as the angular velocity increased.

Inverse Airfoil Design Method for Low-Speed Straight-Bladed Darrieus-Type VAWT Applications

Abstract: The paper demonstrates the application of inverse airfoil design method to improve performance of a low-speed straight-bladed Darrieus-type VAWT. The study shows that an appropriate tailoring of the airfoil surface using the inverse airfoil design technique can help improve performance by eliminating undesirable flow field characteristics at very low Re, such as early transition due to presence of separation bubbles. The increase aerodynamic efficiency then translates into an improved aerodynamic performance of VAWTs specifically at very low chord Reynolds numbers. The study employs an interactive inverse airfoil design method (PROFOIL) that allows specification of velocity and boundary-layer characteristics over different segments of the airfoil subject to constraints on the geometry (closure) and the flow field (far field boundary). Additional constraints to satisfy some desirable features, such as pitching moment coefficient, thickness, camber, etc., along with a merit of performance of the VAWT, such ...

Effects of Active Sting Damping on Common Research Model Data Quality

Abstract: Recent tests using the Common Research Model (CRM) at the Langley National Transonic Facility (NTF) and the Ames 11-foot Transonic Wind Tunnel (11' TWT) produced large sets of data that have been used to examine the effects of active damping on transonic tunnel aerodynamic data quality. In particular, large statistically significant sets of repeat data demonstrate that the active damping system had no apparent effect on drag, lift and pitching moment repeatability during warm testing conditions, while simultaneously enabling aerodynamic data to be obtained post stall. A small set of cryogenic (high Reynolds number) repeat data was obtained at the NTF and again showed a negligible effect on data repeatability. However, due to a degradation of control power in the active damping system cryogenically, the ability to obtain test data post-stall was not achieved during cryogenic testing. Additionally, comparisons of data repeatability between NTF and 11-ft TWT CRM data led to further (warm) testing at the NTF which demonstrated that for a modest increase in data sampling time, a 2-3 factor improvement in drag, and pitching moment repeatability was readily achieved not related with the active damping system.

Flight Control Design for a Tailless Aircraft Using Eigenstructure Assignment

Abstract: We apply eigenstructure assignment to the design of a flight control system for a wind tunnel model of a tailless aircraft The aircraft, known as the innovative control effectors (ICEs) aircraft, has unconventional control surfaces plus pitch and yaw thrust vectoring We linearize the aircraft in straight and level flight at an altitude of 15,000 feet and Mach number 04 Then, we separately design flight control systems for the longitudinal and lateral dynamics We use a control allocation scheme with weights so that the lateral pseudoinputs are yaw and roll moment, and the longitudinal pseudoinput is pitching moment In contrast to previous eigenstructure assignment designs for the ICE aircraft, we consider the phugoid mode, thrust vectoring, and stability margins We show how to simultaneously stabilize the phugoid mode, satisfy MIL-F-8785C mode specifications, and satisfy MIL-F-9490D phase and gain margin specifications We also use a cstar command system that is preferable to earlier pitch-rate command systems Finally, we present simulation results of the combined longitudinal/lateral flight control system using a full 6DOF nonlinear simulation with approximately 20,000 values for the aerodynamic coefficients Our simulation includes limiters on actuator deflections, deflection rates, and control system integrators

Dynamic stall control by passive disturbance generators

Abstract: Passive cylindrical disturbance generators mounted near the leading edge of an airfoil significantly improved its performance under dynamic stall conditions. Time-resolved particle image velocimetry and simultaneous pressure measurements were conducted at the midchord of a pitching airfoil equipped with passive disturbance generators. The disturbance generators were effective in reducing the strength of the dynamic stall vortex and therefore the negative pitching moment peak and hysteresis effects. When the disturbance generators were applied, the flow separation type was altered from leading- to trailing-edge stall. In contrast to the clean case, reattachment was initiated immediately after the separation reached the leading-edge region. In addition to the circular shape, also backward- and forward-wedge-shaped disturbance generators were investigated. Although the backward wedge also showed favorable results, the forward wedge was less successful. The shape of the disturbance generators appears to have a strong influence on the effectiveness of reducing the negative impact of dynamic stall, depending on the sense of rotation of a pair of weak trailing vortices.