Aerodynamic Pitching Moment Research Articles

Multi-objective optimization of composite airfoil fibre orientation under bending–torsion coupling for improved aerodynamic efficiency of horizontal axis wind turbine blade

Modern multi-megawatt wind turbines have long, slender and flexible blades, which experience significant vibration due to aerodynamic loads. These blades are generally made of composite layups with cambered airfoil sections, which induce elastic and dynamic coupling between bending and torsional modes. In addition to these phenomena, the aerodynamic pitching moment also arises due to the offset of the blade pitch axis from the aerodynamic centre. With this in mind, the present work aims to develop a computationally efficient complete multi-body dynamic aero-servo-elastic model of an onshore horizontal axis wind turbine, including bending–torsion coupling and aerodynamic pitching moment. 3D wind fields are generated using TurbSim, and the responses of the coupled system are solved under different operational scenarios, which signify the impact of the aforementioned coupling. It is found to be particularly detrimental at higher wind speeds than the rated value. But, proper design and implementation of structural layup can enhance the aerodynamic performance of the blade. Finally, the optimal design of composite fibre orientation of a benchmark turbine blade is investigated to show its beneficial effects in load/response mitigation.

Parametric Calculations of the Aerodynamics of a Descent Vehicle

We present the methodology and results of parametric aerodynamic studies of vehicles descending into the planet’s atmosphere. The proposed computational approach might serve as the basis for solving a number of problems such as predicting and optimizing the descent trajectory of the vehicle, the search for a rational aerodynamic layout of the vehicle, i.e., tasks requiring massive parametric calculations. The systematization of such calculations is the first step towards the creation of a specialized database that includes sets of input and output data (flight speed, angles of attack, drag and lift coefficients, aerodynamic pitching moment, etc.) and the corresponding three-dimensional fields of gas-dynamic quantities together with computational meshes of various granularity and parameters of the computational model. Additional information to each element of the database might be a set of variables, parameterizing the geometry of the vehicle, experimental data, etc. The probability of forming the information content of such a data-base using modern supercomputer systems is shown. The capabilities of the domestic supercomputer aerodynamic code NOISEtte are demonstrated in the field of multiparametric three-dimensional calculations of descent vehicles based on the numerical solution of the Navier --- Stokes equations on three-dimensional unstructured meshes

Trajectory and attitude deviations for internal store separation due to unsteady and quasi-steady test method

The deviations of trajectory and attitude angle for internal store separation are evaluated by two wind tunnel test methods. One is the Freedrop Test (FDT), which is known as unsteady and time-dependent method of scaled model. The other is the Captive Trajectory System (CTS) test, which is usually regarded as a quasi-steady and time-averaged test technology. The result shows that there is a streamwise adverse pressure gradient on the cavity resulting in a nose-up pitching moment coefficient (>0) acting on the store model. When the initial pitch angular velocity is 0, the store exits the shear-layer with a nose-up pitch attitude causing the store to climb back towards and collide with the aircraft. However, the store passes through the shear-layer into the freestream with a nose-down pitch, which causes a successful separation event when the initial pitch angular velocity < 0. The pitch angle obtained by unsteady test method is different from that by quasi-steady test method. The time-dependent test includes the aerodynamic force induced by pitch angular velocity whereas the time-averaged method (CTS) cannot reflect the effect of unsteady aerodynamic force. The deviation of vertical displacement is not obvious for FDT and CTS test since the store has an initial vertical velocity, which is dominant for the vertical displacement. This means that the highly unsteady flow can create unpredictability in aerodynamic pitching moment of the store, which can lead to the deviation of pitch angle for internal store separation.

Improved sine-trim2 model for the simultaneous elimination of aerodynamic lift, roll, and pitch moment vibrations of rigid rotor helicopters

An improved sine-trim method called sine-trim2 is proposed from the original sine-trim method, based on the phasor superposition principle, to simultaneously eliminate the aerodynamic lift, roll, and pitch moment vibrations of rigid rotor helicopters. The advancing blade concept rotor aerodynamic analysis model is modified and applied on the basis of the free-wake method associated with the inner-trim and outer-trim modules. A hybrid individual blade control system with the blade pitch and linear twist is applied to simultaneously adjust the blade lift and lift arm at each azimuth angle. In the sine-trim2 method, the number of rotor blades should be greater than or equal to four to successfully apply the phasor superposition principle for all kinds of aerodynamic vibration phasors of each blade. Two kinds of lift arm distributions (constant lift arm and sinusoidal lift arm) in one revolution that can both meet the requirements of vibration elimination are considered. Calculation results show that although the sine-trim2 method can eliminate the aerodynamic lift, roll, and pitch moment vibrations of the rigid rotor with either lift arm distribution. A proper sinusoidal lift arm as the control target will have lower control difficulty and efficiency loss than the constant lift arm. The optimal individual blade pitch and twist control distributions, which are both comprehensive superpositions of many higher harmonic inputs, are obtained in each azimuth in one rotation revolution.

SIMULATION OF 2D FLOW AROUND OF AIRFOILS AT LOW-SPEED WIND TUNNEL WITH OPEN JET TEST-SECTION

At present, there is a great interest in the development of new airfoils for wind turbines and high-lift wings of unmanned aerial vehicles (UAV). The requirements for such airfoils differ from conventional aircraft airfoils, because of structural reasons and extreme operating conditions. So, wind turbine airfoils operate frequently under fully separated flow when stall is used for power regulation at high wind speeds. At the same time design of airfoils for wings UAV poses the problem of availability of high-lift at low Reynolds number. Modern airfoils are to a large extent developed from numerical methods. However, the complex flow conditions such as separation at high angles of attack, laminar separation bubbles and the transition from laminar to turbulent flow are difficult to predict accurately. Hence, testing of airfoils at a two-dimensional condition is an important phase in airfoil design. The development and validation of a 2D testing facility for investigation of single and multi-element airfoils in the wind tunnel Т-102 with open test section are considered in this article. T-102 is a continuous-operation, closed-layout wind tunnel with two reverse channels. The test section has an elliptical cross-section of 4 ×2,33 m and a length of 4 m. Two big flat panels of the L × H=3 ×3,9 m size installed upright on balance frame aligned with the free stream are used for simulating two-dimensional flow in the tunnel test section. The airfoil section in the layout of a rectangular wing is mounted horizontally between flat panels with minimum gaps to ensure 2D flow conditions. The aerodynamic forces and pitch moment acting on the model were measured by wind tunnel balance. To determine boundary corrections for a new test section of wind tunnel, the experimental investigation of three geometrically similar models has been executed. The use of boundary corrections has provided good correlation of the test data of airfoil NACA 6712 with the results obtained from the wind tunnel except for lift and drag coefficient values at high angles of attack.

Flight-Test Techniques for Quantifying Pitch Rate and Angle-of-Attack Rate Dependencies

Three different types of maneuvers were designed to separately quantify the pitch rate and angle-of-attack rate contributions to the nondimensional aerodynamic pitching moment coefficient. These maneuvers combined pilot inputs and automatic multisine excitations, and they were demonstrated with the subscale T-2 and Bat-4 airplanes using the NASA Airborne Subscale Transport Aircraft Research flight-test facility. Stability and control derivatives (in particular, and ) were accurately estimated from the flight-test data. These maneuvers can be performed with many types of aircraft, and the results can be used to improve physical insight into the flight dynamics, facilitate more accurate comparisons with wind-tunnel experiments or numerical investigations, and increase simulation prediction fidelity.

Surrogate modeling of a 3D scramjet-powered hypersonic vehicle based on screening method IFFD

The emphasis of this paper lies in the development of an efficient approach to reproduce the behaviors of the scramjet-powered hypersonic system with high fidelity. The modeling of the dual-mode scramjet powered hypersonic vehicle dynamics with shock interaction, Ram-to-Scram transition, and finite rate chemistry reaction is firstly introduced. The structure of surrogate model is identified with the implement of iterative fractional factorial design (IFFD). In order to declare the reliability of the surrogate models, ν-gap metric is applied to distinguish the difference among these surrogate models in terms of closed-loop performance. The results show that the influence of Mach number on the aerodynamics should not be overlooked, and the effect of propulsion system to the aerodynamic pitch moment is dramatic. Further, the partial Kriging model appears to have the closest plants throughout the flight envelope compared with the full Kriging model and polynomials model. Nevertheless, considering the briefness of analytical expression, the polynomials model may be an alternative approach for design-oriented modeling.

Vortical flow prediction of a diamond wing with rounded leading edges

The objective of this work is to assess the potential and limitations of current practice in computational fluid dynamic modeling in predicting vortical flowfields over a generic 53-degree swept diamond wing with rounded leading edges. This wing was designed under STO AVT Task Group 183 and has a constant NACA 64A-006 airfoil section with a leading edge radius of 0.264 percent chord. CFD simulations were run for different angles of attack at a Mach number of 0.15 and a Reynolds number of 2.7×106 based on the mean aerodynamic chord to match experiments. The wind tunnel experiments of the diamond wing were carried out in the Institute of Aerodynamics and Fluid Mechanics of the Technische Universität München, Germany and include aerodynamic lift, drag, and pitch moment measurements as well as span-wise pressure distributions at different chord-wise locations. This data set is used to validate the CFD results. The results presented demonstrate that the CFD compare well with the experiments at small angles of attack; the pitch moments predicted by the SARC turbulence model provide a better match to experimental results than the SA model at moderate angles of attack; and at high angles of attack, CFD predictions are not as good. The flow visualization results show that a leading-edge vortex is formed above the upper surface of the wing at an angle of attack of about eight degrees. This vortex becomes larger and stronger when the angle of attack is increased. With increasing angle of attack, the vortex formation point moves upstream and the vortex core moves inboard towards the wing center. Finally, the computational results show that the flow over the diamond wing is relatively steady throughout the range of angles of attack.

Flight Dynamics and Simulation of Laser Propelled Lightcraft

A seven degree-of-freedom (7DOF) dynamic model was developed to provide insight into the flight behavior of Type 200 and other related lightcraft, and to serve as a research tool for developing future engine-vehicle configurations for laser launching of nanosatellites (1–10+kg). Accurate engine, beam, and aerodynamics models are included to improve the predictive capability of the 7DOF code. The aerodynamic forces of lift, drag, and aerodynamic pitching moment were derived from FLUENT® computational fluid dynamics predictions, and calibrated against limited existing wind tunnel data. To facilitate 7DOF model validation, simulation results are compared with video analysis of actual flights under comparable conditions. Despite current limitations of the 7DOF model, the results compared well with experimental flight trajectory data.

Optimal discus trajectories

A general 3-D dynamic model for men's and women's discus flight is presented including precession of spin angular momentum induced by aerodynamic pitching moment. Dependence of pitching moment coefficient on angle of attack α is estimated from experiment. Numerical integration of 11 equations of motion for nominal release speed v 0=25 m/s and axial spin p 0=42 rad/s also requires 3 other release conditions; initial discus flight path angle β 0, pitch attitude θ 0, and roll angle φ 0. Optimal values for these release conditions are calculated iteratively to maximize range and are similar for both men and women. The optimal men's trajectory and range R=69.39 m is produced by the strategy β 0=38.4°, θ 0=30.7°, and φ 0=54.4°. Initial angular velocities except spin are chosen to minimize wobble but an optimal initial spin rate p 0=25.2 rad/s exists that also maximizes range. Optimal 3-D range exceeds that predicted by 2-D models because, although angle of attack and lift are negative initially, 3-D motion allows advantageous orientation of lift later in flight, with tilt of the axis of symmetry from vertical becoming much smaller at landing. Optimal strategies are discontinuous with wind speed, resulting in slicing and kiting strategies in large head and tail winds, respectively. Sensitivity of optimal range is largest to initial β 0 and least to φ 0. Present calculations do not account for dependence of initial release angle or spin on release velocity or among other release conditions.

Effect of Blade Torsion on Modeling Results for the Small Wind Research Turbine (SWRT)

The small wind research turbine (SWRT) testing project was initiated with the goal of better characterizing both small wind turbine loads and dynamic behavior. The main purpose of the testing was to produce high quality data sets for model development and validation and to help the wind industry further their understanding of small wind turbine behavior, including furling. Previous papers have discussed modeling results for the SWRT using the FAST aeroelastic simulator. This paper summarizes modeling results from both the FAST and ADAMS aeroelastic simulators. The ADAMS model allows for blade torsional stiffness to be modeled, which is shown to significantly improve the agreement between model and test results for rotor speed and explain variances in previous comparisons between the SWRT FAST modeling results and the SWRT test data. The ADAMS model results also show that the aerodynamic pitching moment impacts the blade torsion more than the inertial forces.

Linearized Motion of a Fin-Stabilized Projectile Subjected to a Lateral Impulse

An existing analytical theory for modeling the free-e ight motion of nonspinning, statically stable projectiles is extended to include the effect of a simple lateral impulse applied during e ight. The extended theory is based on the incorporation of generalized lateral translational and angular disturbances into the familiar equations of projectile free-e ight motion. The applied disturbances are then modeled using specie ed mathematical forms, and the modie ed equations are solved to obtain the angular and translational motion of the projectile over the trajectory. The various components of the translational motion of the projectile are extracted and characterized. An idealized application is presented for a large-caliber e nned projectile, representative of the class of 120-mmlong rod e nned projectiles e red from current tracked vehicle weapon systems, subjected to a single lateral control impulse in e ight. The closed-form analytical solutions are compared against results obtained using a numerical trajectory simulation code that incorporates generalized guidance and control commands. Nomenclature A = reference area, od 2 =4, m 2 CL® = derivative of aerodynamic lift force coefe cient with respect to angle of attack CM® = derivative of aerodynamic pitching moment coefe cient with respect to angle of attack

Robust Longitudinal Control Design Using Dynamic Inversion and Quantitative Feedback Theory

The objective is to design low-bandwidth control laws that yield robust performance over a wide e ight envelope in the presence of 20% model uncertainties. This is achieved by using a hybrid controller in which a secondorder dynamic inversion transforms the nonlinear angle-of-attack dynamics to linear time-invariant form, while an outer loop, designed using quantitative feedback theory (QFT), ensures robust performance. The dynamic inversion reduces the plant variations due to nonlinearity, which are traditionally treated as uncertainty in QFT. Thereduced uncertainty allows lowercontrollerbandwidth fora givenlevel of performance. Thehybrid controller design is applied to the unstable, nonlinear, short-period dynamics of the F-16 aircraft and compared with a baseline QFT controller designed without dynamic inversion. Both designs provide satisfactory performance, but the baseline QFT controller requiresa bandwidth of around 4 ‐8 rad/s, whereasthe hybrid design has a bandwidth of only about 1 rad/s, leading to reduced demands on the control hardware. The success of the dynamic inversion depends strongly on the desired linear dynamics, and so recommendations are given on their selection. Nomenclature Cmtotal = nondimensional pitch moment coefe cient Cm®;Cmq;Cm±e = nondimensional partial derivative of Cmtotal with respect to ®;q, and ±e CXtotal;CZtotal = nondimensional coefe cient of x=z force N c = mean aerodynamic chord Iyy = moment of inertia about pitch axis M = aerodynamic pitch moment, M D N qSN cCmtotal m

NON-LINEAR DYNAMICS OF AN AIR FOIL FORCED TO OSCILLATE IN DYNAMIC STALL

The aeroelastic response of a NACA 0012 airfoil with freedom to move in pitch only, and forced to oscillate through dynamic stall in subsonic flow is simulated numerically using a finite difference method. The unsteady aerodynamic pitching moment is obtained via a semi-empirical dynamic stall model; comparison between the predicted pitching moment and experimental values is reasonable. It is demonstrated that the aerodynamic non-linearities associated with dynamic stall are sufficient to cause a chaotic response. Furthermore, it is suggested that there are at least two different routes to chaos depending on the ratio of the forcing frequency is considerably higher than the structural natural frequency, the suggested route to chaos is via a period doubling cascade. For the second, where the frequency of the applied torque is approximately equal to the structural frequency, it is suggested that the response is intermittently chaotic.

Experimental determination of the aerodynamic characteristics of ski-jumpers

SummaryThe aerodynamic pitching moment, lift and drag forces acting on a ski-jumper during the free flight and landing phases were measured using scaled models in a wind tunnel. Comparisons are made between those meaurements and observations of actual ski-jumps.It is shown that the high angles of incidence (typically α varies from 35° to 55°) adopted by ski-jumpers during free flight is determined by the need to prevent tumbling. Over this range of incidence angles, the opposing moments due to weight distribution and aerodynamic pressure distribution are approximately in equilibrium, and small out-of-balance moments can be accommodated by arm movements. The maximum lift/drag ratio on a ski-jumper in free flight occurs at α ≃ 25°. If it were possible to avoid tumbling at low incidence angles, maximum jump lengths would be achieved at α ≃ 8°.