Showing papers on "Pitching moment published in 2014"

Aerodynamic Design Optimization Studies of a Blended-Wing-Body Aircraft

Abstract: The blended wing body is an aircraft configuration that has the potential to be more efficient than conventional large transport aircraft configurations with the same capability. However, the design of the blended wing is challenging due to the tight coupling between aerodynamic performance, trim, and stability. Other design challenges include the nature and number of the design variables involved, and the transonic flow conditions. To address these issues, a series of aerodynamic shape optimization studies using Reynolds-averaged Navier–Stokes computational fluid dynamics with a Spalart–Allmaras turbulence model is performed. A gradient-based optimization algorithm is used in conjunction with a discrete adjoint method that computes the derivatives of the aerodynamic forces. A total of 273 design variables—twist, airfoil shape, sweep, chord, and span—are considered. The drag coefficient at the cruise condition is minimized subject to lift, trim, static margin, and center plane bending moment constraints. ...

Experimental Investigation of the NASA Common Research Model

Abstract: An experimental aerodynamic investigation of the NASA Common Research Model has been conducted in the NASA NTF (National Transonic Facility). Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the WB and WBT0 configurations. Data have also been obtained at a chord Reynolds number of 5 million for the WBNP, WBT+2 and WBT-2 configurations. Force and moment, surface pressure and surface flow visualization data were obtained but only the force and moment data are presented herein. Model deformation measurements, aeroelastic, nacelle/pylon Reynolds number and tail effects have been assessed. The model deformation measurements showed more twist as you go out the wing span, with a break in the high q(sub infinity) data close to CL = 0.6 which is consistent with separation near the tip. Increases in dynamic pressure give an increase in pitching moment and drag and a decrease in lift for the WB and WBT0 configuration at Mach = 0.7, 0.85 and 0.87. The 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. Increases in chord Reynolds number have been found to follow the normal Reynolds number trends except at the 19.8 million low q(sub infinity) cases. The abnormality of the 19.8 million low q(sub infinity) cases is being investigated. The tail effects also follow the expected trends. All of the data shown fall within the 2-sigma limits for repeatability.

Leading-Edge Separation Control Using Alternating-Current and Nanosecond-Pulse Plasma Actuators

Abstract: Wind-tunnel experiments were conducted to quantify the effectiveness of ac and nanosecond-pulse single dielectric barrier discharge plasma actuators to suppress leading-edge stall on a NASA Energy Efficient Transport airfoil at Mach numbers up to 0.4 and chord Reynolds numbers up to 2.3×106. The airfoil model was designed to have a removable leading edge to accommodate two different leading-edge plasma-actuator designs, either with a thick ceramic or a thin Kapton dielectric layer. The exposed electrode for both plasma actuators was located at the leading edge of the airfoil. The covered electrode for both was on the suction side of the leading edge. The model was mounted on stages that measured the lift and drag forces and the pitching moment about the quarter-chord location. Both steady and unsteady ac plasma-actuator operation were examined. By its nature, the nanosecond-pulse plasma actuator only operates in unsteady operation. The optimal unsteady frequencies with regard to lift, lift to drag, and pi...

Summary of the Fourth AIAA Computational Fluid Dynamics Drag Prediction Workshop

Abstract: Results from the Fourth AIAA Drag Prediction Workshop are summarized. The workshop focused on the prediction of both absolute and differential drag levels for wing–body and wing–body/horizontal-tail configurations of the NASA Common Research Model, which is representative of transonic transport aircraft. Numerical calculations are performed using industry-relevant test cases that include lift-specific flight conditions, trimmed drag polars, downwash variations, drag rises, and Reynolds-number effects. Drag, lift, and pitching moment predictions from numerous Reynolds-averaged Navier–Stokes computational fluid dynamics methods are presented. Solutions are performed on structured, unstructured, and hybrid grid systems. The structured-grid sets include point-matched multiblock meshes and overset grid systems. The unstructured and hybrid grid sets comprise tetrahedral, pyramid, prismatic, and hexahedral elements. Effort is made to provide a high-quality and parametrically consistent family of grids for each g...

Vortex dynamics around pitching plates

Abstract: Vortex dynamics of wakes generated by rectangular aspect-ratio 2 and 4 and two-dimensional pitching flat plates in free stream are examined with direct numerical simulation and water tunnel experiments. Evolution of wake vortices comprised of tip, leading-edge, and trailing-edge vortices is compared with force history for a range of pitch rates. The plate pivots about its leading edge with reduced frequency from π/8 to π/48, which corresponds to pitching over 1 to 6 chord lengths of travel. Computations have reasonable agreement with experiments, despite large differences in Reynolds number. Computations show that the tip effects are confined initially near the wing tips, but begin to strongly affect the leading-edge vortex as the motion of the plate proceeds, with concomitant effects on lift and drag history. Scaling relations based on reduced frequency are shown to collapse aerodynamic force history for the various pitch rates.

RANS-based Aerodynamic Shape Optimization Investigations of the Common Research Model Wing

Abstract: The aerodynamic shape optimization of transonic wings requires Reynolds-averaged Navier–Stokes (RANS) modeling due to the strong nonlinear coupling between airfoil shape, wave drag, and viscous effects. While there has been some research dedicated to RANS-based aerodynamic shape optimization, there has not been an benchmark case for researchers to compare their results. In this investigations, a series of aerodynamic shape optimizations of the Common Research Model wing defined for the Aerodynamic Design Optimization Workshop are presented. The computational fluid dynamics solves Reynolds-averaged Navier–Stokes equations with a Spalart–Allmaras turbulence model. A gradient-based optimization algorithm is used in conjunction with a discrete adjoint method that computes the derivatives of the aerodynamic forces. The drag coefficient at the nominal flight condition is minimized subject to lift, pitching moment and geometric constraints. A multilevel acceleration technique is used to reduce the computational cost. A total of 768 shape design variables are considered, together with a grid with 28.8 million cells. The drag coefficient of the optimized wing is reduced by 8.5% relative to the baseline. The single-point design has a sharp leading edge that is prone to flow separation at off-design conditions. A more robust design is achieved through a multipoint optimization, which achieves more reliable performance when lift coefficient and Mach number are varied about the nominal flight condition. To test the design space for local minima, randomly generated initial geometries are optimized, and a flat design space with multiple local minima was observed.

An experimental and three-dimensional computational study on the aerodynamic contribution to the passive pitching motion of flapping wings in hovering flies.

Abstract: The relative importance of the wing's inertial and aerodynamic forces is the key to revealing how the kinematical characteristics of the passive pitching motion of insect flapping wings are generated, which is still unclear irrespective of its importance in the design of insect-like micro air vehicles. Therefore, we investigate three species of flies in order to reveal this, using a novel fluid-structure interaction analysis that consists of a dynamically scaled experiment and a three-dimensional finite element analysis. In the experiment, the dynamic similarity between the lumped torsional flexibility model as a first approximation of the dipteran wing and the actual insect is measured by the Reynolds number Re, the Strouhal number St, the mass ratio M, and the Cauchy number Ch. In the computation, the three-dimension is important in order to simulate the stable leading edge vortex and lift force in the present Re regime over 254. The drawback of the present experiment is the difficulty in satisfying the condition of M due to the limitation of available solid materials. The novelty of the present analysis is to complement this drawback using the computation. We analyze the following two cases: (a) The equilibrium between the wing's elastic and fluid forces is dynamically similar to that of the actual insect, while the wing's inertial force can be ignored. (b) All forces are dynamically similar to those of the actual insect. From the comparison between the results of cases (a) and (b), we evaluate the contributions of the equilibrium between the aerodynamic and the wing's elastic forces and the wing's inertial force to the passive pitching motion as 80–90% and 10–20%, respectively. It follows from these results that the dipteran passive pitching motion will be based on the equilibrium between the wing's elastic and aerodynamic forces, while it will be enhanced by the wing's inertial force.

DLR Results from the Third AIAA Computational Fluid Dynamics Drag Prediction Workshop

Abstract: A summary about the DLR, German Aerospace Center results from the fourth AIAA Computational Fluid Dynamics Drag Prediction Workshop is presented. Compared to the investigations in the previous three workshops, the latest workshop had a stronger focus on drag and trim drag predictions as well as pitching moment calculations. Therefore, the new Common Research Model developed by NASA’s Subsonic Fixed Wing Aerodynamics Technical Working Group and tested in NASA wind tunnels is used. It represents a state-of-the-art transonic transport aircraft configuration, and in contrast to the configurations previously taken, it includes an optional horizontal tailplane with three different tail settings. DLR has defined three objectives for its activities in the fourth drag prediction workshop. At first, investigations should identify solution accuracy and grid convergence behavior using prismatic element dominant grids for boundary-layer resolution in comparison to hexahedral element dominant grids. Second, the influen...

Numerical Study of Aerodynamics of a Wing-in-Ground-Effect Craft

Abstract: The flow around a wing-in-ground-effect craft flying at α=3 deg and α=9 deg over flat and wavy ground is simulated and investigated by using ANSYS FLUENT, employing the compressible Reynolds-averaged Navier–Stokes equations and the Spalart–Allmaras turbulence model. The sliding mesh technology is used to simulate the relative motion between the wing-in-ground-effect craft and wavy ground. The effects of the wavy ground, flight height, and angle of attack on the aerodynamic characteristics and flowfield are analyzed in detail. The aerodynamic forces are found to be periodic when the wing-in-ground-effect craft flies over wavy ground. The aerodynamic forces over both flat and wavy ground vary with flight height in the same pattern. As the flight height reduces, the lift, drag, and nose-down pitching moment all increase at both angles of attack (α=3 deg and α=9 deg); however, the lift-to-drag ratio increases for all flight heights at α=3 deg, while it first increases and then decreases at α=9 deg. Redu...

Application of Jetstream to a Suite of Aerodynamic Shape Optimization Problems

Abstract: This work demonstrates the performance of Jetstream, a high-fidelity aerodynamic shape optimization methodology for three-dimensional turbulent flows. The geometry parameterization and mesh movement is accomplished using B-spline volumes and linear elasticity mesh movement. The Euler or Reynolds-averaged Navier-Stokes (RANS) equations are solved at each iteration using a parallel Newton-Krylov-Schur method. The equations are discretized in space using summation-by-parts operators with simultaneous approximation terms to enforce boundary and block interface conditions. The gradients are evaluated using the discrete-adjoint method to allow for gradient-based optimization using a sequential quadratic programming algorithm. The goal of this work is to investigate the performance of Jetstream for three test problems. The first problem is the drag minimization of a two-dimensional symmetric airfoil in transonic inviscid flow, under a geometric constraint that the airfoil have a thickness greater than or equal to that of a NACA 0012 airfoil. Although the shock waves are not quite eliminated, they are substantially weakened, such that the drag coefficient is reduced by 86% compared to the NACA 0012 airfoil. The second problem is drag minimization through optimizing the twist distribution of a three-dimensional wing characterized by NACA 0012 sections in subsonic inviscid flow, subject to a lift constraint. A nearly elliptical spanwise lift distribution is achieved by the optimized twist distribution, leading to a span efficiency factor of 0.98. The third problem is drag minimization through optimizing the sections and twist distribution of the blunt-trailing-edge Common Research Model wing in transonic turbulent flow, subject to lift and pitching moment constraints. For this case the optimization is performed based on the solution of the RANS equations, with the Spalart-Allmaras turbulence model fully coupled and linearized. The drag coefficient is reduced by eleven counts, or 6%, when analyzed on a fairly fine mesh.

Experimental Investigation of Unsteady Transition on a Pitching Rotor Blade Airfoil

Abstract: The unsteady flow around the pitching helicopter main rotor blade airfoil EDI-M109 was experimentally investigated at conditions similar to those existing on a retreating rotor blade in forward flight High speed pressure measurements and hot film anemometry were used to investigate the unsteady transition characteristics of the airfoil Results are presented for dynamic test points with attached flow, light dynamic stall and deep dynamic stall at M = 03 and Re = 18 x 10^6 The results include the discussion of the periodicity of the hot film signals for different flow states The transition process of the pitching airfoil is analysed and the significance of the intermittent region is described A time delay between the transition and the model motion is discussed and a linear relationship between the transition position and the time is observed The influences of the pitching amplitude on the transition characteristics are discussed and the flow separation initiating dynamic stall is analysed

Numerical Calculation of Effect of Elastic Deformation on Aerodynamic Characteristics of a Rocket

Abstract: The application and workflow of Computational Fluid Dynamics (CFD)/Computational Structure Dynamics (CSD) on solving the static aeroelastic problem of a slender rocket are introduced. To predict static aeroelastic behavior accurately, two-way coupling and inertia relief methods are used to calculate the static deformations and aerodynamic characteristics of the deformed rocket. The aerodynamic coefficients of rigid rocket are computed firstly and compared with the experimental data, which verified the accuracy of CFD output. The results of the analysis for elastic rocket in the nonspinning and spinning states are compared with the rigid ones. The results highlight that the rocket deformation aspects are decided by the normal force distribution along the rocket length. Rocket deformation becomes larger with increasing the flight angle of attack. Drag and lift force coefficients decrease and pitching moment coefficients increase due to rocket deformations, center of pressure location forwards, and stability of the rockets decreases. Accordingly, the flight trajectory may be affected by the change of these aerodynamic coefficients and stability.

Experimental investigation of air jets to control shock-induced dynamic stall

Abstract: This paper shows experimental results for dynamic stall control on a dynamically pitching OA209 airfoil at Mach 0.5, with Reynolds numbers 1.9 × 106 and 0.85 × 106 . The control was by supersonic constant blowing on the suction side of the airfoil. Dry compressed air was blown normal to the airfoil chord, from portholes at 10% chord, with diameter 1% chord. At both Reynolds numbers, the OA209 without blowing experienced shock-induced stall with a hysteresis in lift and pitching moment around the static values, rather than the overshoot in forces typically associated with a dynamic stall vortex. The forces and the stall control were primarily functions of the maximum angle of attack, with full control of stall possible for maximum angles of attack of 14◦ and less. The higher Reynolds number required relatively more blowing (higher Cq , Cμ ) to control the dynamic stall. Drag was reduced for separated flow, but the energy required in compressed air to achieve this was more than the savings in drag, and no cases were found in which flow control resulted in a reduction in total power used. Increasing the jet spacing resulted in equivalent flow control with less air use. Jets spaced at 20% chord and mass flux ratio Cq = 0.004 (momentum ratio Cμ = 0.016) resulted in a reduction of the pitching moment peak by 60%. The flow control with air jets was uncritical regarding the aerodynamic damping.

Free-flight measurement technique in the free-piston high-enthalpy shock tunnel

Abstract: A novel multi-component force-measurement technique has been developed and implemented at the impulse facility JAXA-HIEST, in which the test model is completely unrestrained during the test and thus experiences free-flight conditions for a period on the order of milliseconds. Advantages over conventional free-flight techniques include the complete absence of aerodynamic interference from a model support system and less variation in model position and attitude during the test itself. A miniature on-board data recorder, which was a key technology for this technique, was also developed in order to acquire and store the measured data. The technique was demonstrated in a HIEST wind-tunnel test campaign in which three-component aerodynamic force measurement was performed on a blunted cone of length 316 mm, total mass 19.75 kg, and moment of inertia 0.152 kgm2. During the test campaign, axial force, normal forces, and pitching moment coefficients were obtained at angles of attack from 14° to 32° under two conditions: H0 = 4 MJ/kg, P0 = 14 MPa; and H0 = 16 MJ/kg, P0 = 16 MPa. For the first, low-enthalpy condition, the test flow was considered a perfect gas; measurements were thus directly compared with those obtained in a conventional blow-down wind tunnel (JAXA-HWT2) to evaluate the accuracy of the technique. The second test condition was a high-enthalpy condition in which 85% of the oxygen molecules were expected to be dissociated; high-temperature real-gas effects were therefore evaluated by comparison with results obtained in perfect-gas conditions. The precision of the present measurements was evaluated through an uncertainty analysis, which showed the aerodynamic coefficients in the HIEST low enthalpy test agreeing well with those of JAXA-HWT2. The pitching-moment coefficient, however, showed significant differences between low- and high-enthalpy tests. These differences are thought to result from high-temperature real-gas effects.

Effect of Fuel Sloshing in the External Tank on the Flutter of Subsonic Wings

Abstract: This study considers the aeroelastic instability of a subsonic wing with an external fuel tank store containing sloshing fuel. The wing is modeled using Euler-Bernoulli beam and the fuel sloshing is simulated by a mass-spring-damper system as an equivalent mechanical model. An extension of the Peters finite-state aerodynamic model along with the strip theory is utilized to obtain the external work due to the unsteady lift and pitching moment. Also, the aerodynamic loads on the store are calculated using the slender body theory. Using Lagrange equations, the governing equations of motion are derived and solved by the Rayleigh-Ritz method to determine the flutter speed of the wing/store. The effect of fuel sloshing on the aeroelastic instability of the wing is investigated by several numerical studies and some conclusions are outlined. The obtained results show that when the fuel-sloshing frequency in the external tank is between the frequencies of the fundamental bending and torsional modes of the wing, the sloshing/aeroelasticity coupling occurs. Numerical studies illustrate that the fundamental frequency and damping of sloshing in the tank are the main parameters determining the increase or decrease of the flutter speed and a proper tuning of these parameters can considerably extend the flutter-safe speed range.

On the effect of sweep on separation control

Abstract: The performance of a flapped wing based on a NACA 0012 airfoil section and equipped with a linear array of fluidic oscillators was investigated experimentally to assess the significance of wing sweep and aspect ratio on the efficiency of the actuation. The semi-span wing that was suspended from the wind tunnel ceiling through a six-component balance could be withdrawn partially from the test section and rotated in a plane parallel to the flow thus its sweep could vary from 0° to ±45° and its aspect ratio could change from 2.4 to 7.5. The wing incidence, its flap deflection, and the level and distribution of the actuation were the additional independent parameters investigated. The experiments were carried out at Reynolds numbers varying between 300,000 and 500,000. The boundary layer was tripped in order to fix the location at which transition to turbulence occurs. To overcome separation at high flap deflections in the absence of wing sweep, a minimum momentum coefficient of the order of 1% was required. However, on a swept-back wing a substantially lower input level could improve the lift generated by the wing by some 20% and alter the pitching moment provided the aggregate number of the actuators was small. Under these conditions, the actuators acted as fluidic boundary layer fences that can be switched ON or OFF on demand and change the aerodynamic characteristics of the wing for takeoff and landing purposes. An attempt was made to explain the mechanism that makes the fluidic oscillators so effective.

Experimental investigation of high-pressure pulsed blowing for dynamic stall control

Abstract: Dynamic stall control using pulsed blowing is compared with control by constant blowing for an OA209 airfoil. Flow control was by blowing from 42 portholes, flush with the airfoil surface, of diameter 1 % chord positioned at 10 % chord and with separation 6.7 % chord. Light stall at Mach 0.3 could be fully suppressed by constant blowing, and for deep stall a pitching moment peak reduction of 65% was seen. For the jet configuration and test cases investigated in this paper, pulsed blowing at 100–500 Hz was found to be at best as effective as constant blowing with the same mass flux for the control of dynamic stall.

Fifth Drag Prediction Workshop: ONERA Investigations with Experimental Wing Twist and Laminarity

Abstract: This paper is aimed at presenting the computational studies that have been performed at ONERA–The French Aerospace Lab in 2013 in the framework of the 5th AIAA Drag Prediction Workshop. As data concerning the Common Research Model configuration were collected from the NASA Langley Research Center and Ames Research Center wind tunnels, a significant discrepancy appeared between the experimental and computational pitching moment evaluations. Investigations carried out by NASA in 2012 showed that the experimental model and the numerical geometry were slightly different. Indeed, at the design point, the wing twist of the experiments was stronger than the one used for the computations. Therefore, new computational-fluid-dynamics studies using the Common Research Model wing–body configuration with the corrected twist have been completed at ONERA. The common multiblock grids provided by the Drag Prediction Workshop Committee have been adequately modified before being computed with the ONERA elsA Reynolds-average...

Unsteady Aerodynamics of Deformable Thin Airfoils

Abstract: The paper presents a theory for the unsteady aerodynamics of deformable thin airfoils. It extends the theory developed by Theodorsen and Garrick, which is restricted to rigid body motion. Frequency-domain lift, pitching moment, and thrust expressions are derived for an airfoil undergoing harmonic oscillations and deformation in the form of the Chebychev polynomials. The first two polynomials give the rigid body motion, whereas the rest represent the deformation. The results are verified with the time-domain unsteady aerodynamic theory of Peters. Numerical results are presented for several combinations of airfoil motion, which identify various possibilities for thrust generation using a deformable airfoil.

A novel strategy for designing and manufacturing a fixed wing mav for the purpose of increasing maneuverability and stability in longitudinal axis

Abstract: In this study, a novel simple strategy is proposed to choose and accommodate an airfoil based on the effects of airfoil type and plan-form shape on the flight performance of a micro air vehicle. In this strategy, after defining flight mission, the weight of the micro air vehicle is estimated and then, aerodynamic parameters and thrust force are calculated. In the next step, some different plan-forms and airfoils are investigated to be selected for decreasing the stall region in high attack anglesby open source software named XFLR5. Having calculated the aerodynamic center, the pitching moment needed to stabilize the micro air vehicle is computed. Due to the static margin, the airfoil camber line is changed to stabilize the micro air vehicle and then, its thickness is improved to reach to a high aerodynamic characteristic. To evaluate the software results, some flight tests are performed which then compared to the software results that show a good agreement. Finally, some adjustments and improvements are made on the micro air vehicle and then, its performance is obtained by the flight tests. The flight test results show it has an excellent aerodynamic performance, stability and maneuverability.

Design and Performance of a Quad-Shrouded Rotor Micro Air Vehicle

Abstract: This paper describes the experimental investigation of a quad-shrouded rotor micro air vehicle and focuses on the hover performance improvements over a conventional unshrouded micro quad rotor. The effects of number of rotor blades, blade root pitch angle, and shroud diffuser length on aerodynamic performance were studied. The rotor diameter was 6.6 cm, it had a tip Re of 20,000, and the gross weight of the vehicle was 90 g, with a shroud weight fraction of 13%. With the optimized design, the power loading of the quad-shrouded rotor was about 15% greater than the unshrouded configuration. To completely evaluate the configuration, the performance of the vehicle in edgewise flow was investigated. The drag and pitching moment for the shrouded rotor was about 2.5 times greater than the unshrouded vehicle. However, control-moment measurements suggested that the edgewise gust tolerance was at least 4 m/s. The vehicle prototype was also successfully flight tested in hover using onboard feedback regulation. Base...

Design and aerodynamic investigation of a parallel vehicle on a wide-speed range

Abstract: In order to design a hypersonic vehicle for a wide-ranged Mach number, a novel parallel vehicle for a wide-speed range has been proposed. In this paper, we employ a numerical method to investigate a parallel vehicle’s aerodynamic performance and flow field characteristics. The obtained results show that the aerodynamic performance of the novel parallel vehicle is better than that of the waverider designed with a single Mach number for the wide-speed range. With the increase in Mach number, the lift-to-drag ratio of the novel parallel vehicle first increases and then decreases. When the Mach number is 7 and the angle of attack is 3°, the lift-to-drag ratio is the largest, and its value is 3.968. When the angle of attack is 3°, the lift-to-drag ratio is not lower than 3.786 in the range considered in the current study, and the novel parallel vehicle’s aerodynamic performance is good. The wing changes the drag performance of the parallel vehicle remarkably, and results in the decrease of the lift-to-drag ratio. Meanwhile, the wing can enhance the pitching moment performance.