Showing papers on "Pitching moment published in 2012"

Aerodynamic performance of a hovering hawkmoth with flexible wings: a computational approach

Abstract: Insect wings are deformable structures that change shape passively and dynamically owing to inertial and aerodynamic forces during flight. It is still unclear how the three-dimensional and passive change of wing kinematics owing to inherent wing flexibility contributes to unsteady aerodynamics and energetics in insect flapping flight. Here, we perform a systematic fluid-structure interaction based analysis on the aerodynamic performance of a hovering hawkmoth, Manduca, with an integrated computational model of a hovering insect with rigid and flexible wings. Aerodynamic performance of flapping wings with passive deformation or prescribed deformation is evaluated in terms of aerodynamic force, power and efficiency. Our results reveal that wing flexibility can increase downwash in wake and hence aerodynamic force: first, a dynamic wing bending is observed, which delays the breakdown of leading edge vortex near the wing tip, responsible for augmenting the aerodynamic force-production; second, a combination of the dynamic change of wing bending and twist favourably modifies the wing kinematics in the distal area, which leads to the aerodynamic force enhancement immediately before stroke reversal. Moreover, an increase in hovering efficiency of the flexible wing is achieved as a result of the wing twist. An extensive study of wing stiffness effect on aerodynamic performance is further conducted through a tuning of Young's modulus and thickness, indicating that insect wing structures may be optimized not only in terms of aerodynamic performance but also dependent on many factors, such as the wing strength, the circulation capability of wing veins and the control of wing movements.

Further Investigation of the Support System Effects and Wing Twist on the NASA Common Research Model

Abstract: An experimental investigation of the NASA Common Research Model was conducted in the NASA Langley National Transonic Facility and NASA Ames 11-foot Transonic Wind Tunnel Facility for use in the Drag Prediction Workshop. As data from the experimental investigations was collected, a large difference in moment values was seen between the experiment and computational data from the 4th Drag Prediction Workshop. This difference led to a computational assessment to investigate model support system interference effects on the Common Research Model. The results from this investigation showed that the addition of the support system to the computational cases did increase the pitching moment so that it more closely matched the experimental results, but there was still a large discrepancy in pitching moment. This large discrepancy led to an investigation into the shape of the as-built model, which in turn led to a change in the computational grids and re-running of all the previous support system cases. The results of these cases are the focus of this paper.

Effects of Sideslip on the Aerodynamics of Low-Aspect-Ratio Low-Reynolds-Number Wings

Abstract: The growing interest in micro aerial vehicles has brought attention to the need for an improved understanding of the aerodynamics of low-aspect-ratio wings at lowReynolds numbers. In this study, flat plate wings with rectangular and tapered planforms were fabricated with aspect ratios of 0.75, 1, 1.5, and 3, and the aerodynamic loading was measured at Reynolds numbers between 5 10 and 1 10. Surface tuft visualization was used to observe the interactions between the tip vortices and the leading-edge vortex. The tests were initially conducted at a sideslip angle of 0 and were then repeated for 10, 20, and 35 with and without winglets. Measurements made with a sixcomponent force balance showed that a decrease in aspect ratio caused an increase in stall and CLmax due to the nonlinear lift induced by the interacting flow on the upper wing surface. In addition, the detachment of tip vortices after stall leads to a sudden decrease in drag coefficient as the magnitude of the induced drag drops significantly. At increasing sideslip angles, the effects of the crossflow still contribute to an increase in lift but significantly reduce the pitching moment about the quarter-chord, thus decreasing the wing’s ability to recover from angle-of-attack perturbations. These results show that, while the effects of tip vortices and the leading-edge vortex complicate the flowfield around a low-aspect-ratio wing, particularly at increased sideslip angles, their impact tends to improve the aerodynamic performance.

Investigation on high angle of attack characteristics of hypersonic space vehicle

Abstract: The high angle of attack characteristics play an important role in the aerodynamic performances of the hypersonic space vehicle. The three-dimensional Reynolds Averaged Navier-Stokes (RANS) equations and the two-equation RNG k-ɛ-turbulence model have been employed to investigate the influence of the high angle of attack on the lift-to-drag ratio and the flow field characteristics of the hypersonic space vehicle, and the contributions of each component to the aerodynamic forces of the vehicle have been discussed as well. At the same time, in order to validate the numerical method, the predicted results have been compared with the available experimental data of a hypersonic slender vehicle, and the grid independency has been analyzed. The obtained results show that the predicted lift-to-drag ratio and pitching moment coefficient show very good agreement with the experimental data in the open literature, and the grid system makes only a slight difference to the numerical results. There exists an optimal angle of attack for the aerodynamic performance of the hypersonic space vehicle, and its value is 20°. When the angle of attack is 20°, the high pressure does not leak from around the leading edge to the upper surface. With the further increasing of the angle of attack, the high pressure spreads from the wing tips to the central area of the vehicle, and overflows from the leading edge again. Further, the head plays an important role in the drag performance of the vehicle, and the lift percentage of the flaperon is larger than that of the rudderevator. This illustrates that the optimization of the flaperon configuration is a great work for the improvement of the aerodynamic performance of the hypersonic space vehicle, especially for a high lift-to-drag ratio.

Support System Effects on the NASA Common Research Model

Abstract: An experimental investigation of the NASA Common Research Model was conducted in the NASA Langley National Transonic Facility and NASA Ames 11-Foot Transonic Wind Tunnel Facility for use in the Drag Prediction Workshop. As data from the experimental investigations was collected, a large difference in moment values was seen between the experimental and the computational data from the 4th Drag Prediction Workshop. This difference led to the present work. In this study, a computational assessment has been undertaken to investigate model support system interference effects on the Common Research Model. The configurations computed during this investigation were the wing/body/tail=0deg without the support system and the wing/body/tail=0deg with the support system. The results from this investigation confirm that the addition of the support system to the computational cases does shift the pitching moment in the direction of the experimental results.

Low-speed aerodynamic characteristics improvement of ATR 42 airfoil using a morphing wing approach

Abstract: This paper presents a morphing wing technique for extending laminar flow over the airfoil of the ATR 42 regional transport airplane. The objective is to reduce aerodynamic drag by delaying the transition from laminar to turbulent flow over the airfoil upper surface. The optimization of the airfoil shape is performed for several incompressible subsonic regimes, using a genetic algorithm tool, coupled with the subsonic aerodynamic solver XFOIL. A reduction of the drag coefficient with up to 26.73% and a delay in the transition point of up to 24.81% has been achieved.

Kinematic control of aerodynamic forces on an inclined flapping wing with asymmetric strokes

Abstract: In the present study, we conduct an experiment using a one-paired dynamically scaled model of an insect wing, to investigate how asymmetric strokes with different wing kinematic parameters are used to control the aerodynamics of a dragonfly-like inclined flapping wing in still fluid. The kinematic parameters considered are the angles of attack during the mid-downstroke (α(md)) and mid-upstroke (α(mu)), and the duration (Δτ) and time of initiation (τ(p)) of the pitching rotation. The present dragonfly-like inclined flapping wing has the aerodynamic mechanism of unsteady force generation similar to those of other insect wings in a horizontal stroke plane, but the detailed effect of the wing kinematics on the force control is different due to the asymmetric use of the angle of attack during the up- and downstrokes. For example, high α(md) and low α(mu) produces larger vertical force with less aerodynamic power, and low α(md) and high α(mu) is recommended for horizontal force (thrust) production. The pitching rotation also affects the aerodynamics of a flapping wing, but its dynamic rotational effect is much weaker than the effect from the kinematic change in the angle of attack caused by the pitching rotation. Thus, the influences of the duration and timing of pitching rotation for the present inclined flapping wing are found to be very different from those for a horizontal flapping wing. That is, for the inclined flapping motion, the advanced and delayed rotations produce smaller vertical forces than the symmetric one and the effect of pitching duration is very small. On the other hand, for a specific range of pitching rotation timing, delayed rotation requires less aerodynamic power than the symmetric rotation. As for the horizontal force, delayed rotation with low α(md) and high α(mu) is recommended for long-duration flight owing to its high efficiency, and advanced rotation should be employed for hovering flight for nearly zero horizontal force. The present study suggests that manipulating the angle of attack during a flapping cycle is the most effective way to control the aerodynamic forces and corresponding power expenditure for a dragonfly-like inclined flapping wing.

Three Unstructured Computational Fluid Dynamics Studies on Generic Unmanned Combat Aerial Vehicle

Abstract: Three independent studies from the United States (NASA), Sweden (Swedish Defense Research Agency), and Australia (Defense Science and Technology Organization) are analyzed to assess the state of current unstructured grid computational fluid dynamic tools and practices for predicting the complex static and dynamic aerodynamic and stability characteristics of a generic 53-deg swept, round-leading-edge uninhabited combat air vehicle configuration, calledSACCON (which stands for “stability and control configuration”). NASAexercised theUSM3D tetrahedral cell-centered flow solver, while the Swedish Defense Research Agency and the Defense Science and Technology Organization applied the Swedish Defense Research Agency/EDGE general-cell vertex-based solver. The authors primarily employ the Reynolds-averaged Navier–Stokes assumption, with a limited assessment of the EDGEdetached eddy simulation extension, to explore sensitivities to grids and turbulencemodels. Correlations with experimental data are provided for force and moments, surface pressure, and off-body flow measurements. The vortical flowfield over SACCON proved extremely difficult to model adequately. As a general rule, the prospect of obtaining reasonable correlations of SACCON pitching moment characteristics with the Reynolds-averaged Navier–Stokes formulation is not promising, even for static cases. However, dynamic pitch oscillation results seem to produce a promising characterization of shapes for the lift and pitching moment hysteresis curves. Future studies of this configuration should include more investigation with higher-fidelity turbulence models such as detached eddy simulation.

Force Measurements in Hypervelocity Flows with an Acceleration Compensated Strain Gage Balance

Abstract: Force and moment measurements are performed on a 6-inch Apollo shaped capsule in the LENS I and 48-Inch reflected shock tunnels at CUBRC in cold hypersonic nitrogen flows and for high enthalpy flows (10 MJ/kg) of air, nitrogen, pure oxygen and a 50%-50% oxygen-argon mixture. Measurements with an acceleration compensated flexured strain gage force balance are compared with non-equilibrium Navier-Stokes computations. For the low enthalpy nitrogen runs, axial and normal force coefficients are within 2% of the numerical predictions whereas the pitching moment is under-predicted by 10%. Excellent agreement within 0.5% for the lift-to-drag ratio is achieved. For the first time, the effect of thermodynamic and chemical non-equilibrium on the force and moment coefficients is measured on a reentry capsule in a reflected shock tunnel and compared with state of the art numerical simulations. Both computations and experiments show an increase in the axial force, normal force and pitching moment coefficients and a decrease in the lift-to-drag ratio.

Advances in Rotor Performance and Turbulent Wake Simulation Using DES and Adaptive Mesh Refinement

Abstract: Time-dependent Navier-Stokes simulations have been carried out for a rigid V22 rotor in hover, and a flexible UH-60A rotor in forward flight. Emphasis is placed on understanding and characterizing the effects of high-order spatial differencing, grid resolution, and Spalart-Allmaras (SA) detached eddy simulation (DES) in predicting the rotor figure of merit (FM) and resolving the turbulent rotor wake. The FM was accurately predicted within experimental error using SA-DES. Moreover, a new adaptive mesh refinement (AMR) procedure revealed a complex and more realistic turbulent rotor wake, including the formation of turbulent structures resembling vortical worms. Time-dependent flow visualization played a crucial role in understanding the physical mechanisms involved in these complex viscous flows. The predicted vortex core growth with wake age was in good agreement with experiment. High-resolution wakes for the UH-60A in forward flight exhibited complex turbulent interactions and turbulent worms, similar to the V22. The normal force and pitching moment coefficients were in good agreement with flight-test data.

Nondimensional Modeling of Ducted-Fan Aerodynamics

Abstract: The unique aerodynamic nature of the ducted-fan configuration makes it difficult to represent in terms of traditional nondimensional coefficients. Analysis of wind-tunnel data for an axisymmetric generic ducted-fan configuration has led to a new nondimensional modeling scheme. Force coefficients are plotted versus advance ratio for a range of angles of attack, yielding several observations. Each angle of attack yields a linear trend versus advance ratio, with all lines converging through a single point. This fulcrum point shares the same thrust coefficient value as hover tests but at a nonzero advance ratio. This suggests that a hovering ducted fan self-induces a freestream flow, even though the vehicle is stationary. The complicated pitching moment behavior is captured succinctly through modeling the center of pressuremovement. The fan power required was also successfully modeled using the figure of merit, typically a hover performance metric, over the entire flight regime to relate power required to thrust generated. A new wind-tunnel velocity correction for ducted fans is also presented. The new statistical modeling technique attains very high correlation for present and legacy data. It is possibly the most concise representation of ducted-fan aerodynamics to date, using 12 nondimensional coefficients.

Hypersonic Ludwieg Tube Design and Future Usage at the US Air Force Academy

Abstract: Hypersonic flows are usually characterized by the presence of strong shocks and equilibrium or nonequilibrium gas chemistry. Accurate prediction of these effects is critical to the design of any vehicle that flies at hypersonic velocities. The pressures and skin friction acting on the surface of the vehicles are integrated over the complete configuration to define the aerodynamic forces (e.g., lift, drag, pitching moment, and control surface effectiveness). The peak heat-transfer rate and the heating load are mapped over the vehicle surface as part of the process to design the thermal protection system. All of these are challenging parameters to predict at hypersonic speeds, requiring specialized experimental facilities or advanced computational simulation tools. Toward improving our ability to understand these flows, the U.S. Air Force Academy Department of Aeronautics will soon place into service a Mach 6 Ludwieg tube. This facility will enable the Academy to be better prepared to support the hypersonic research goals of the Department of Defense and other organizations. A Ludwieg tube is a high speed wind tunnel which does not require a total pressure control device or large settling chamber which is common for conventional blow-down tunnels. This greatly reduces the size and cost of operating the tunnel, since large compressors, heaters, and pressure vessels are not required for operation. The operational costs for a Ludwieg tube have been further reduced by the use of a fast-acting valve instead of the traditional bursting diaphragm. The facility will be compatible with the Boeing/AFOSR Mach 6 Quiet Tunnel at Purdue University, and will be able to support work from that facility, such as hypersonic transition. In addition, the facility will be able to support one of the difficult flow regimes encountered by NASA, namely acquiring data on blunt vehicles typical of lower ballistic coefficient configurations necessary for space activities. Finally, the facility will be able to conduct research on shock wave/boundary layer interactions at hypersonic speeds, which will support research for hypersonic air-breathing propulsion for future projects.

Analysis of Thrust Characteristics of a Magnetic Sail in a Magnetized Solar Wind

Abstract: Themagnetic sail is an advanced space propulsion concept that uses an artificial magnetosphere for capturing the solar wind energy. In this study, the interaction of the solar wind with themagnetosphere of a magnetic sail has been simulated based on the resistive magnetohydrodynamics model in two-dimensional space and the plasmadynamic characteristics of magnetic sail were evaluated. When the solar wind is not magnetized by the interplanetary magnetic field, the attitude of the magnetic sail spacecraft is static stable when the magnetic moment vector is perpendicular to the solar wind flow direction. The interplanetary magnetic field not only enhances a drag force in the direction leaving the sun (i.e., thrust) but also acts on the pitching moment; the pitching moment due to the interplanetary magnetic field rotates the magnetic sail spacecraft so as to align the magnetic moment vector parallel to the interplanetarymagnetic field. Despite the weak interplanetarymagnetic field adopted in the simulation, which is 1 order of magnitude lower than the typical value, the pitching moment coefficient is significant. The attitude stability of the magnetic sail is hence strongly affected by the interplanetary magnetic field.

Multistage-Fusion Algorithm for Estimation of Aerodynamic Angles in Mini Aerial Vehicle

Abstract: Nature is the best teacher of many scientific research and development. One such nature inspired invention is the flight vehicle system. Among these, Mini Aerial Vehicles (MAVs) are a class of flight vehicles which have size close to birds. The current work is inspired by the fact that MAV with Low Aspect Ratio (LAR) wings (Aspect Ratio < 2) has nonlinear lift curves and it does not stall sharply as compared to high aspect ratio wings. This can be utilized in flying the MAVs at lower speeds and at higher angle of attacks. In LAR wings, the lift is constant or keeps increasing up to large Angle of Attack (AOA) which can help in flying the MAV at slow speed. Main focus of the present work is to estimate the AOA, which can be further used for development of control laws. This paper presents a data fusion algorithm for estimating the aerodynamic angles in a MAV. For simulation, true states of the aircraft motion are generated using a flight simulation program and a zero mean white noise is added to few of these states as required for the measurements. These noisy states are used as sensor measurements for estimating the AOA and Side Slip Angle (SSA). The proposed scheme is a multi-stage in which initially Euler angles are estimated and later stages are used for AOA and SSA estimation. In the first stage, the Euler angles are estimated using the accelerometer outputs and the rate gyro outputs and V in an Extended Kalman Filter (EKF) algorithm. As a first attempt for estimation of AOA and SSA, only acceleration, angular rates and airspeed were used as measurement in six states EKF. It was observed that there is a bias of more than 10° ° ° ° present in the estimated aerodynamic angles. Since, it is not possible to have any additional sensors on-board due to weight restriction, a new modification is proposed in which a pseudo estimation of AOA and SSA was used as measurement. This effectively reduces the estimation bias and the mean error in both AOA and SSA, and are below 1°. Nomenclature Fx = Total force along the body x-axis Fy = Total force along the body y-axis Fz = Total force along the body z-axis L = Total moment along the body x-axis M = Total moment along the body y-axis N = Total moment along the body z-axis p = Roll rate q = Pitch rate r = Yaw rate Cx = Force coefficient in the x direction Cy = Force coefficient in the y direction Cz = Force coefficient in the z direction Cl = Moment coefficient in the x direction Cm = Moment coefficient in the y direction

Numerical Prediction of Pitch Damping Derivatives for a Finned Projectile at Angles of Attack

Abstract: This study extends the time-accurate forced-oscillation planar pitching ReynoldsAveraged method to predict the pitch damping dynamic stability derivative at non-zero angles of attack for projectile munitions. The simulations were performed for two configurations of the Army-Navy Finner Missile at Mach 1.96, for two Reynolds number test conditions, and for angles of attack ranging from 0–85°. Comparisons with archival wind tunnel data indicated good to excellent numerical prediction for the pitch damping moment. For the static aerodynamic coefficients, the method predicts the pitching moment with excellent accuracy, and the pitching moment slope with fair accuracy.

Experimental Study of Free Stream Turbulent Effects on Dynamic Stall of Pitching Airfoil by Using Particle Image Velocimetry

Abstract: The unsteady flow fields above NACA 0015 airfoil pitching with/without upstream turbulence generator are investigated in a water tunnel by mean of particle image velocimetry (PIV). The turbulence was generated by a square bar mesh situated at the inlet of the test section. The airfoil pitching waveform is performed under the condition calculated from the angle of attack histogram of a vertical axis wind turbine (VAWT). By using PIV, the instantaneous vortex structures above the pitching airfoil can be revealed. It allows us to study the free stream turbulence effects on dynamic stall over an airfoil at pitching waveform the same as VAWT. It is found that the free stream turbulence intensity has significant impacts on the dynamic stall process. The dynamic stall process is delayed to higher incidence angles on increasing the turbulence intensity.

Numerical Investigation on Aerodynamic Characteristics of a Compound Wing-in-Ground Effect

Abstract: The performance of a wing-in-ground effect craft depends mostly on its wing configuration. The aerodynamic characteristics of compound wings were numerically investigated during the ground effect. The compound wing was divided into three parts with one rectangular wing in the middle and two reverse taper wings with anhedral angle at the sides. The NACA 6409 airfoil was employed as a section of wings. Three-dimensional computational fluid dynamics was applied as a computational scheme. The k-"turbulent model was used for the turbulent ?ow over wing surface. The computational results of a rectangular wing with aspect ratio 1, angle of attack 2 deg, and different ground clearance were compared with the experimental data of other published work. Next, the principal aerodynamic characters of compound wings and a rectangular wing were computed for various ground clearance. The aerodynamic characters of compound wings were compared with the rectangular wing, which has an acceptable increase in lift and decrease in drag. Consequently, the lift-to-drag ratio has a considerable improvement especially in small ground clearance, although their nose-down pitching moment has a little reduction. The performance of the wing improves noticeably for a certain total span of compound wing when the span of the middle part becomes smaller.

Trim analysis of a moving-mass actuated airplane

Abstract: This paper investigates the feasibility of trimming an airplane using internal mass motion as moment generation mechanisms. Two internally moving masses are considered: (i) mass moving laterally to generate rolling moment and (ii) mass moving longitudinally to generate pitching moment. These are to replace aileron and elevator, respectively. Various straightlevel ight conditions are considered to determine the trim conditions, particularly the masses and positions of the moving masses. For a small unmanned aerial vehicle, feasible level ight trim conditions are found using only the longitudinal moving mass, without the conventional aerodynamic control surface, elevator. Furthermore, endurance and range of for straight-level ight conditions are determined and compared with the aircraft with elevator.

Dynamic Load Control on a Finite Span Wind Turbine Blade Using Synthetic Jets

Abstract: The feasibility of active flow control to mitigate hysteresis loop due to a dynamically pitching finite span s809 blade was investigated experimentally at a Reynolds number of 220,000. Under normal operating conditions, hysteresis loop and tip vibrations exist, which with extended exposure would cause blade fatigue and eventually translate to a reduced lifetime of wind turbines. In this regard, active flow control via arrays of synthetic jet actuators was explored as a means to control flow separation over the finite span blade, which can lead to mitigation of these undesired unsteady loads. In the present work, a six-component load cell was used to measure the aerodynamic loading of lift, drag and pitching moment. Stereoscopic Particle Image Velocimetry (SPIV) measurements were also performed to understand the effects of synthetic jets on flow separation during dynamic pitch, and to correlate these effects to the forces and moment measurements. It was shown that active flow control could delay or minimize dynamic stall through the reduction of the hysteresis loop of the aerodynamic loads. This implies less unsteady aerodynamic loadings on the blade, which can potentially lead to prolonged life of wind turbines.

Thrust Vector Control of an Upper-Stage Rocket with Multiple Propellant Slosh Modes

Abstract: The thrust vector control problem for an upper-stage rocket with propellant slosh dynamics is considered. The control inputs are defined by the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. The rocket acceleration due to the main engine thrust is assumed to be large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. A multi-mass-spring model of the sloshing fuel is introduced to represent the prominent sloshing modes. A nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. The effectiveness of the controller is illustrated through a simulation example.

Control of a spacecraft with time-varying propellant slosh parameters

Abstract: This paper studies the thrust vector control problem for an upper-stage rocket with fuel slosh dynamics. The control inputs are the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. It is assumed that the rocket acceleration due to the main engine thrust is large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. The prominent sloshing modes are represented by a multi-mass-spring model with time-varying parameters. A time-varying nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. A simulation example is included to illustrate the effectiveness of the controller.