Showing papers on "Wing root published in 2002"

Experimental and Theoretical Study of Gust Response for High-Aspect-Ratio Wing

Abstract: A nonlinear response analysis of a high-aspect-ratio wing aeroelastic model excited by gust loads is presented along with a companion wind-tunnel test program. For the wind-tunnel tests, a high-aspect-ratio wing aeroelastic experimental model with a slender body at the tip has been constructed, and a rotating slotted cylinder gust generator has been used to generate a gust excitation field. A LabVIEW 5.1 measurement and analysis system is used to measure the gust response, flutter boundary, and limit-cycle oscillation behavior. Structural equations of motion based on a nonlinear beam theory are combined with the ONERA aerodynamic stall model to study the effects of geometric structural nonlinearity and steady angle of attack on nonlinear gust response of high-aspect-ratio wings. Also a dynamic perturbation analysis about a nonlinear static equilibrium is used to determine the small perturbation flutter boundary. The fair to good quantitative agreement between theory and experiment demonstrates that the present analysis method has reasonable accuracy.

Navier-Stokes Optimization of Supersonic Wings with Four Objectives Using Evolutionary Algorithm

Abstract: The design optimization of wings for supersonic transport by means of Multiobjective Evolutionary Algorithms is presented. Three objective functions are first considered to minimize the drag for transonic cruise, the drag for supersonic cruise and the bending moment at the wing root at the supersonic condition. The wing shape is defined by planform, thickness distributions and warp shapes in total of 66 design variables. A Navier-Stokes code is used to evaluate the aerodynamic performance at both cruise conditions. Based on the results, the optimization problem is further revised. The definition of the thickness distributions is given more precisely by adding control points. In total 72 design variables are used. The fourth objective function to minimize the pitching moment is added. The results of the revised optimization are compared with the three-objective optimization results as well as NAL’s design. Two Pareto solutions are found superior to NAL’s design for all four objective functions. The planform shapes of those solutions are “Arrow wing” type.

A numerical investigation of nonlinear aeroelastic effects on flexible high aspect ratio wings

Abstract: A nonlinear aeroelastic analysis is developed to analyze the aeroelastic characteristics of flexible high-aspect-ratio wings at transonic speeds. This is achieved by directly coupling a three-dimensional geometric nonlinear methodology, based on a 12 degree-of-freedom beam finite element, with an Euler/Navier-Stokes computational fluid dynamics analysis. Static aeroelastic results are presented for an unswept and swept high-aspect-ratio wing. Unswept wing results show a reversal in twist due to the nonlinear torsion-bending coupling effects. Specifically, the torsional moments due to drag become large enough to cause the wing twist rotations to washin the wing tips, whereas the linear results show a washout twist rotation. The nonlinear twist results are attributed to the large bending displacements coupled with the large drag experienced by this flexible high-aspect-ratio wing at the transonic flow conditions. Swept wing results show that nonlinear torsion-bending effects tend to reduce the amount of washout as compared to a linear structural aeroelastic analysis, making the wing more prone to tip stall.

Hybrid Inverse Airfoil Design Method for Complex Three-Dimensional Lifting Surfaces

Abstract: A method is presented for inverse design of airfoils for complex three-dimensional wings in incompressible e ow. The method allows for prescription of inviscid velocity distributions over different cross sections of the wing in a multipoint fashion. A hybrid approach is used to determine the shapes of the wing cross sections that satisfy the design specie cations. The airfoils forming the cross sections of the wing are generated using an inverse code for isolated airfoil design. A three-dimensional panel method is then used to obtain the velocity distributions over the resulting wing. The isolated airfoil velocity distributions are then used as design variables in a multidimensional Newton iteration method to achieve the design specie cations on the wing. The method is particularly useful for complex geometries such as junctures, where three-dimensional and interference effects have to be accounted for in the design process. Akey feature ofthedesign method is a schemeto avoid using thepanel method for sensitivity computations for the Newton iterations. This scheme not only results in signie cant reductions in computation time but also enables the integration of any readily available three-dimensional analysis code in executable form. Examples are shown to demonstrate the usefulness of the method. Nomenclature Cl = airfoil lift coefe cient, chord D1 c = airfoil/section chord F = vector containing the residuals J = Jacobian matrix n = number of design variables V = airfoil/section velocity nondimensionalized by the freestream velocity v = desired value for the velocity difference over a segment ® = angle of attack, deg 1V = velocity difference over a segment normalized by the freestream velocity ±x = vector containing the corrections to the design variables Subscripts i = index of design variable for the Newton iteration r = property associated with the wing root t = property associated with the wing tip w = property associated with the wing 2D = property associated with an airfoil in isolation 3D = property associated with a cross section of a three-dimensional lifting surface

3D RANS Simulations for High-Lift Analysis of Transport Aircraft Configurations

Abstract: This paper presents the numerical simulation and the analysis of viscous high-lift flows around a complex wing/body configuration (DLR ALVAST) with deployed high-lift devices. The solution of the Reynolds averaged Navier-Stokes equations (RANS) is carried out using the unstructured DLR TAU-code. The results of the computation are compared to measurements. The investigations aim at a better understanding of the aerodynamics at the wing root and the lift breakdown.

Supersonic maneuverable aircraft

Abstract: aeronautical engineering. SUBSTANCE: aircraft has fuselage, wing, root wing extensions, horizontal tail, vertical tail, cabin, engine nacelles located in tail section of fuselage, engines mounted in nacelles and engine air intakes. Aircraft is provided with swept-forward wing with flaps. Root expansions are mounted along starboard and port nose section of fuselage; each expansion has pointed curvilinear leading edge at positive sweep angle. At point of conjugation of leading edges of extensions with leading edge of wing root, sweep angle ranges from 85 to 90 deg.; length of each extension along respective side of fuselage is equal to 0.55-0.75 of length of fuselage nose section. EFFECT: improved aerodynamic and maneuvering properties of aircraft. 24 cl, 1 dwg

Wind tunnel test model

Abstract: PROBLEM TO BE SOLVED: To manufacture a wind tunnel test model, in which many small holes opened on the surface of a wing and passages communicating the small holes with a wing root are arranged so as not to interfere with each other, at a low cost. SOLUTION: This wind tunnel test model is provided with many small holes Ha-Hd distributed on the surface of a main wing W, and the small holes Ha-Hd are communicated with many passages Pa-Pd formed in the span direction in the main wing W and extended to the wing root. The main wing W is molded by a photo-molding method, and the passages Pa-Pd are integrally formed in the main wing W when the main wind W is molded. The passage Pa communicated with the small hole Ha on a wing tip side is arranged on the center side of a wing cross section, and the passage Pd communicated with the small hole Hd on the wing root side is arranged on the surface side of the wing cross section. The passage Pa communicated with the small hole Ha on the wing tip side and the small hole Hd on the wing root side are reliably prevented from interfering with each other in the wind, and the degree of freedom on the number and layout of the small holes Ha-Hd is largely increased.

Protective cover assembly for tail wing root of helicopter

Abstract: The utility model provides a protective cover assembly for tail wing root of helicopter, which comprises a support beam and a carrying handle arranged above the support beam, and is characterized in that two ends of a screw rod are arranged inside a left bearing seat and a right bearing seat. Two sides of the screw rod are provided with symmetrical positive thread and negative thread. A semi-circular support plate is arranged above the vertical screw rod, and a guide key which is positioned at a support block that is arranged under the screw rod is moved along with the support plate. The plane of the support plate is provided with a support hole. A support rod is horizontally fixed inside the hole by a bearing. A rocking handle is fixed by a flat key positioned inside the right bearing seat. The utility model has the effects that the assembly can be designed based on the width of the tail wing of the helicopter, the protective cover can be rapidly assembled by that the protective cover extending to the required size and the assembly placed on the wing root in the assembling of the protective cover, which is labor-saving and time-saving.

Lift enhancement of a thick delta wing using rotational leading edges

Abstract: The subsonic flow for a 45deg-swept delta wing is numerically simulated to see the effects of rotational leading edges on aerodynamic forces. In the present study, part of the cylindrical leading edge can be rotated which extends from x = 0.25C to C. The computation has been performed at an angle of attack of 20deg and a Reynolds number of 2.0 x 104 based on the wing root chord. Under these conditions, the flow is fully separated from the wing surface in the case without leading edge rotation, so that no suction peaks due to leading edge vortices appear over the upper surface of the wing. Consequently, the delta wing will not have an effective vortical lift. However, by rotating the leading edges, leading edge vortices can be successfully produced, which leads to an increase in lift. Furthermore, the effect of asymmetric rotation on the rolling moment is presented.

Papers : Vortex Flow and Aerodynamic Load Characteristics of the Delta Wing / LEX Configuration in Sideslip

Abstract: The vortex flow and aerodynamic load characteristics of a sweep delta wing with the leading edge extension in sideslip condition is investigated experimentally. The freestream velocity is 40 m/sec, which corresponds to a Reynolds number per meter of 1.76X based on the wing root chord. The angles of attack range from to 28{^{\circ}}, and the sideslip angles treated are 0{^{\circ}}, -10{^{\circ}}, 20{^{\circ}}. The LEX vortex of the leeward side. The LEX and wing vortics coalesce to to become a concentrated strong vortex or to break down at down at downstream position. Due to the interation of the LEX and wing vortices, a high suction pressure is maintained on the windward wing surface, and a low suction pressure is formed on the leeward wing surface.ﾖ⨀

Visualization Study of High-Incidence Vortical Flow over the LEX/Delta Wing Configuration with Sideslip

Abstract: An off-surface flow visualization experiments have been performed to investigate the flow field over a delta wing with the leading edge extension(LEX). The model is a flat wing with sweepback angle. The free stream velocity is 6.2 m/s, which corresponds to Reynolds number of based on the wing root chord. The angle of attack and sideslip angle range from and , respectively. The visualization technique of using the micro water-droplet and the laser beam sheet enabled to observe the vortical flow structures, which can not be obtained by 5-hole probe measurements. With sideslip angle, the interaction and breakdown of the LEX and wing vortices was promoted in the windward side, whereas, it was suppressed in the leeward side.