Showing papers on "Wing root published in 1997"

Effect of Canards on Delta Wing Vortex Breakdown During Dynamic Pitching

Abstract: The effect of a canard on delta wing vortices was investigated in the 2 3 3 ft water tunnel at Wichita State University. It is well known that the leading-edge vortices generated by a delta-shaped wing greatly enhance a vehicle’ s performance at high angles of attack. In this experiment, different canards were placed in front of a 70-deg swept main delta wing. Dye e ow visualization was used to observe the vortex breakdown location during dynamic pitch-up and pitch-down motion with varying pitch rates. Compared to the no-canard cone guration, results showed that there was a delay in vortex breakdown because of the presence of the canard and the dynamic pitch motion. The most favorable delay was obtained when the canard was located closest to the main delta wing and the model was pitched up at a fast rate or pitched down at a slow rate. Complete vortex breakdown on the main delta wing (i.e., full stall ) occurred at 53 deg for the static case without canard. In comparison, complete vortex breakdown occurred past 90 deg when a canard cone gured delta wing was pitched up at the fastest rate tested (i.e., k = 0.2).

Diamond, Cropped, Delta, and Double-Delta Wing Vortex Breakdown During Dynamic Pitching

Abstract: A series of experiments were conducted on the effect of different delta wing shapes on vortex breakdown under dynamic pitching conditions.

Loop-shaped transverse drive body or split-wing loop for aircraft

Abstract: The drive body is based on principles derived from the vortex flow of the wings of gliding birds. At its origin, or in wingspan direction from the wing root to the wing tip, the transverse drive creating structure is connected to a long base part (base wing) to provide overall an almost continuous distribution of the transverse drive at the transition point to two narrower transverse drive bodies which merge with each other without sharp bends.

Chimera Donor Cell Search Algorithm Suitable for Solving the Full Potential Equation

Abstract: An approximate iterative search algorithm for finding donor cells associated with the chimera zonal grid approach is presented. This new algorithm is both fast and simple. It is used in conjunction with a chimera-based full potential solver for computing transonic flow solutions about wing and wing/fuselage configurations. Within each grid zone a fully implicit approximate factorization scheme is used to advance the solution one iteration. This is followed by the explicit advance of all common intergrid boundaries using a trilinear interpolation of the velocity potential. The presentation is highlighted with numerical result comparisons, a grid refinement study, and parametric variation of pertinent algorithm parameters. The new search algorithm produces donor cells for the two-zone wing problem at a rate in excess of 60,000 cells/s (single processor Cray C90). The approximate nature of the search algorithm, which causes some of the donor cells to be approximated by nearest neighbor cells, does not cause any impact on solution accuracy. Overall the results indicate that the present chimera zonal grid approach is a viable technique for solving the full potential equation for aerodynamic applications

Using flow control devices in small wind tunnels

Abstract: Complex configurations such as transonic transport aircrafts are well suited for testing modern tools of aircraft design. These tools are computer based systems as well as wind tunnel test devices. This paper reports about combining theoretical, numerical and experimental techniques to analyze and influence specific aerodynamic phenomena such as the flow quality in the wing root area. Large size wind tunnel models are needed to investigate the flow around such specific parts of the model. The aim of controlling transonic flow quality at the wing-body junction in a small wind tunnel has motivated the design for an incomplete and modular configuration. Specific devices are needed to simulate a lift distribution resulting from the complete model: Circulation Control Splitter Blades (CCSB) are replacing the wing tips. The investigated high wing configuration is a generic model for a new military transport aircraft. The model, called the DLR-F9, was tested at the transonic wind tunnel (TWG) at the DLR Gottingen. Results from numerical simulation and a first set of experiments using CCSB's are reported.

Hybrid blade for gas turbine

Abstract: PROBLEM TO BE SOLVED: To enhance the resistance of an aircraft fan blade against impact rupture caused by collision with birds by letting the hybrid blade be provided with a first and a second segment, forming the first segment out of metallic material, and also forming the second segment out of composite material. SOLUTION: A gas turbine blade 10 is formed out of a shank part 12, and of a solid wing foil shaped part 22 composed of a first and a second segment 14 and 16. The first segment 14 is made out of metallic material, and has all of a suction side 30, a leading edge 24 and a trailing edge 26 contained in an area from a wing root 32 to a wing tip 34, and the second segment 16 is made out of composite material, and is partially provided for a pressure side in an area from a space close to the leading edge 24 to a space close to the trailing edge 26 over an area from the wing root 32 to a space close to the wing tip 34. And the resistance is thereby enhanced based on a fact that the traces of impact remain firstly in the area of the first segment 14 close to the leading edge 24 in the pressure side 28, and secondly in the area of the second segment 16 adjacent to the pressure side 28 at the time of collision with birds.

Bilge voltex energy recovery device for ship

Abstract: PROBLEM TO BE SOLVED: To convert the energy of a bilge voltex flow into a propelling force and to reduce ship resistance by attaching fins, one each, to the hull surface of left and right gunwales in the front side of a propeller providing the wing root part of the fin on the hull surface and positioning this substantially in the center of the bilge voltex SOLUTION: Fins 8 are attached to the surface of a hull 1, one for each, of left and right gunwales in the front side of a propeller 3 and the wing root part 8c of the fin 8 is attached such that a fin wing end part 8b is positioned in a snail center 6 position in which the bilge voltex 7 is passed through the fin 8 position The camver 8a of the fin 8 is attached facing downward, a flow is generated so as to flow to the upper surface 8d and the lower surface 8e of the fin 8m, that is, negative pressure surface sides and a wing end voltex 11 rotated in a direction reverse to the direction of the bilge voltex 7 flowing out from the fin wing end part 8b to a downstream side is formed Thus, the energy of a rotational flow caused by the bilge voltex 7 of a stern is converted into a propelling force and thus ship resistance is reduced

Design-Oriented Leading-Edge Thrust Force Prediction for Supersonic Lifting Surfaces

Abstract: A design-oriented method for evaluating the leading-edge (LE) thrust force on supersonic lifting surfaces with subsonic LE is presented. The method is developed to overcome numerical noise and nonsmooth drag predictions observed when currently attainable LE-thrust techniques are used for wing planform shape optimization. The method is an extension of a design-oriented unsteady supersonic lifting surface capability developed previously for aeroservoelastic shape optimization of wings. It is a panel/lattice method where assumed pressure-weighting functions, taking the LE singularity into account, are prescribed to the LE panels while constant pressure panels are retained elsewhere. Explicit expressions for aerodynamic influence coefficients are retained over most of the wing, except for cases involving the LE panels, where numerical integration must be used. Planform shape sensitivities of the LE thrust and wing pressures are obtained using a combination of analytic and semianalytic techniques.

An experimental study of the breakdown of leading-edge vortices on diamond, cropped, delta, and double delta wings during dynamic pitching

Abstract: The effect of delta wing shape on leadingedge vortex breakdown was investigated in the 2 ft x 3 ft water tunnel at Wichita State University. It is well known that a vehicle's performance at high angles of attack is greatly influenced by the development of leading-edge vortices on a delta-shaped wing. In this experiment, the aft Vs of a 76° swept delta wing was modified to obtain diamond, cropped, standard delta, and double delta shapes. The vortex breakdown location during dynamic pitch-up and pitch-down motion was observed using dye flow visualization. Amongst the four shapes tested, the cropped delta wing had the longest unbursted leading-edge vortex during dynamic pitching while the double delta wing had the earliest vortex breakdown. NOMENCLATURE AR Wing Aspect Ratio, b /S b Wing Span cr Wing Root Chord q Dynamic Pressure, V&pUa, Re Reynolds Number, Uwcr Iv S Wing Area Uo,, Freestream Velocity a Angle of Attack a' Pitch Rate K Non-Dimensional Pitch Rate, a'cr /2UC A Sweep-back Angle v Freestream Flow Kinematic Viscosity p Freestream Flow Density INTRODUCTION Recent interest in high angle of attack aerodynamics has refocussed attention on delta shaped wings. Vortices are formed at non-zero angles of attack as flow separates along the leading edges of a delta shaped wing. Very low pressure is associated with these leading-edge vortices, and they can account for up to 30% of the total lift at moderate angles of attack. For example, lift continues to increase until about 40° angle of attack on a 76° swept delta wing. In comparison, symmetric two dimensional airfoils typically stall out around 10-15° angle of attack. Unfortunately, there are limits to the benefits produced by these delta wing vortices. As the angle of attack is increased, there is a sudden breakdown in vortex structure. This phenomena, also known as vortex "bursting" results in a sudden stagnation in core axial flow and an expansion in radial size. Once this occurs, lift is no longer enhanced aft of the burst point. Thus, the development and subsequent breakdown of leading-edge vortices is crucial to the performance of delta wing aircraft. There have been a number of attempts to control delta wing vortices including the use of blowing', suction', flaps", and canards'. The *Assistant Professor, Senior Member AIAA. Student, present address: 26A Lengkong Satu, Singapore 417500, Singapore. *Student, Student Member AIAA, present address: Department of Aerospace Engineering, University of Southern California, Los Angeles, CA 90089-1191. Associate Professor, Member AIAA. Copyright ® 1997 by Roy Y. Myose, Boon-Kiat Lee, Shigeo Hayashibara, and L. Scott Miller. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc. reader is referred to Lee and Ho for a more complete review on delta wing vortices. As the angle of attack is increased on delta wings, the unburst part of the leading-edge vortices become shorter. Under dynamic conditions, there is a hysteresis or phase lag in the vortex burst location. For example, the vortex burst location is further aft compared to the static case (at a given a) under pitch-up motion and further forward under pitch-down motion. This phase lag is larger as the pitch rate is increased.' Thus, fast pitch-up and slow pitch-down is desired in order to delay vortex breakdown. It is well known that the sweep-back angle on a delta wing affects the development and breakdown of the leading-edge vortices. For example, full stall angle of attack (under static conditions) occurs at 27° on a 55° swept delta wing while it is 38°on a 65° swept delta wing and 54° on a 75° swept delta wing. Thus, high sweep-back angles provide enhanced lift until high angles of attack. This principle is employed on modern day fighter aircraft using strakes in front of the main wing to form a double delta shape. In this case, the strakes provide enhanced lift in addition to the main whig. A number of unique delta whig shapes were also investigated by Gatlin and McGrath. However, all of these studies on the effect of delta wing shape were conducted under static conditions. Under dynamic conditions, investigations on only the basic shapes such as the delta'' and double delta wings have been conducted. Modern day military aircraft often employ novel wing shapes hi order to incorporate stealth technology. Furthermore, enhanced performance at high angles of attack and under unsteady conditions may be required of these military aircraft. Thus, a series of experiments were conducted on the effect of different delta wing shapes on vortex breakdown under dynamic pitching conditions. EXPERIMENTAL METHOD The experiment was conducted in the 2 ft x 3 ft water tunnel located at Wichita State University, National Institute for Aviation a) Diamond b) Crop c) Delta d) Double delta Fig. 1. Test model shapes. Research (NIAR). The facility is a closed-loop water tunnel containing a total of 3500 gallons of water. The flow velocity is adjustable up to 1.0 ft/s using an impeller pump driven by a 5 hp variable speed motor. The facility has excellent optical access providing two side views, a bottom view, and an end view. Fig. 1 shows a sketch of the four different whig shapes which were tested. All four shapes have a sweep-back angle of 76° at the wing apex and a root chord length of 9". The sweep-back angle of the aft Vs of each wing is different, corresponding to diamond, cropped, standard delta, and double delta shapes. Each whig is made out of 0.063" thick aluminum alloy, and both starboard and port sides are symmetrically beveled at a 45° angle. The wings are painted white to enhance visual contrast, but black reference grid lines perpendicular to the whig centerline are marked in 5% chord intervals to determine the vortex breakdown locations. Table 1 lists the specifications for each whig shape. A dynamic test mount consisting of a rotating turntable was used to obtain the dynamic pitch motion (see fig. 2). The position and rotational speed of this belt-driven turntable are controlled by a variable speed DC motor. The turntable is capable of 360° of rotation at rates up to 30 deg/s. The pitching motion produced by this dynamic test mount is a constant angular rate of change, i.e., a "ramp-type" pitching motion. Additional details about the dynamic test mount are presented by Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc. Table 1: Test model specifications.

Blade for fan and manufacture thereof

Abstract: PROBLEM TO BE SOLVED: To provide a blade with a favorable aerodynamic characteristic and the less generation of noise by constituting the blade out of the extruded section having a wing shape section and its stem part strongly and lightly. SOLUTION: A flat and thick torque box part 5 is formed integrally on the extruded section for constituting a blade 1 and a stem part 20 is formed by expanding and deforming this torque box part 5 cylindrically at the wing root part of the blade 1.

Shock-wave/boundary-layer interaction in transonic flow around a forward-swept wing

Abstract: The shock-wave/boundary-layer interaction in transonic flow around a forward-swept wing is investigated experimentally and numerically. For low angles of attack, good agreement is obtained between the numerical and experimental position and intensity of the shock wave, except than close to the wing root. The explanation of the latter discrepancy is the presence in the experiments of a splitter plate, which affects the measurements, especially in the region close to the wing root. For low angles of attack, the numerical simulation has then been used to investigate the effects of Mach and Reynolds numbers on shockwave/boundary-layer interaction. For high angles of attack, in which large zones of separated flow are present, significant discrepancies are observed between experimental and numerical results. Therefore, the effects of shock-wave/boundary-layer interactions on the stall behavior of the forward-swept wing are derived only from the analysis of the experimental data

Delta wing vortex dynamics in high rate oscillations in pitch

Abstract: The increasing performance demands on advanced aircraft, including maneuvers at high angles of attack, has led to a need for the prediction of vehicle aerodynamics that are dominated by unsteady separated flow effects. For aircraft with highly swept wing leading edges the challenge is to fully understand the unsteady flow physics behind the observed dramatic effects of vortex breakdown. In the case of moderate leading-edge sweep the flow dynamics are even more complicated because of the presence of a partial-span leading-edge vortex. In the present paper an attempt is made to pinpoint the physical flow processes causing the experimentally observed unusual dynamic characteristics of a 45 deg delta wing describing high-rate pitch oscillations. Nomenclature c wing root chord f oscillation frequency k reduced frequency, = coc^U^ p static pressure; coefficient Cp = (p-P=o)/(PcoUro/2) s local semispan U horizontal velocity x axial body-fixed coordinate (Fig. 4) y spanwise coordinate (Fig. 4) z out-of-plane deflection (Fig. 4) * Engineering Consultant. Fellow AIAA. Copyright