Delta Wing Configuration Research Articles

Low-speed aerodynamics of apex fences on a tailless delta configuration

Apex fences are highly swept, upper-surface spoiler-like surfaces hinged at the leading edges of a delta wing. The fences are deployed vertically to generate a strong pair of counter-rotating vortices at low angles of attack when the main-wing vortex system is weak. The intense suction induced by the fence vortices on the wing apex area creates a nose-up moment that, in conjunction with trailing-edge flaps, will enhance the trimmed-lift capability, thus reducing landing speed. Conversely, at high angles of attack, when the basic delta wing apex is highly loaded, the fences reduce the apex lift, resulting in pitch-down acceleration for a rapid return to normal flight attitude. These hypothesized characteristics of the apex fence concept were validated through low-speed tunnel experiments on a 60-deg delta wing configuration using flow visualization techniques, upper-surface pressure surveys, and balance measurements. This paper presents selected results of the investigation indicating the quantitative benefits that can be realized by the application of apex fences to tailless delta aircraft.

Prediction of vortex lift on interacting delta wings in incompressible flow

Vortex flow modeling is used to calculate the steady inviscid incompressible flow past one, two and three delta wing configurations. The wings are modeled with vortex lattices and the leading and trailing-edge sheets are modeled by segmented straight vortex filaments. Aerodynamic characteristics are obtained for a range of geometry and angle of attack.

Nonlinear Supersonic Potential Flow over Delta Wings

A numerical procedure has been developed for the computation of the steady, inviscid supersonic flow over aircraft configurations. The technique accounts for major nonlinear effects (shock waves, blunt leading edges) at low to moderate supersonic speeds. A fully implicit marching technique for the full potential equation is utilized in a stereographically projected, conformally mapped, spherical coordinate frame. Cross-flow planes are efficiently solved by type-dependen t relaxation techniques. Results are presented for several delta wing configurations and bodies of revolution, and are compared with existing experimental data, Euler's equations solutions, and results from linearized theories (panel methods). I. Introduction T HE goal of the present investigation is the development of a highly efficient and general numerical procedure for the computation of the steady, inviscid supersonic flow over aircraft configurations. In particular, this paper will concentrate on the development and verification of the method through application to isolated delta wings and bodies of revolution. Emphasis will be given to the prediction of nonlinear effects at imbedded shock waves and blunt leading edges in the low to moderate supersonic Mach number range. Existing computational techniques in this regime may be broadly characterized as panel methods and finite-difference methods. Panel methods, which have been demonstrated to be widely versatile and quite efficient, have reached a very advanced stage of development and utility for subsonic flows. Efforts are continuing to improve their capabilities at supersonic speeds (i.e., Ehlers et al. *). These methods address the solution to the first-order linear small-disturbance equation with various approximations for the boundary conditions. Finite-difference techniques for wing planforms in the supersonic regime concentrate on solutions to Euler's equations with differing treatments of shock waves, shock capturing2 or shock fitting,3'5 and a variety of methods for grid generation (i.e., Moretti5) and geometry input. The major drawbacks of these methods are associated with their large computational times which are due primarily to the fine grids necessary to resolve thin wings and the accompanying strict stability requirements. Other problem areas for the Euler solutions in the Mach number range considered here are due to subsonic Mach numbers in the axial marching direction, initial conditions at the apex of the wing or body, and vortical singularities of the leeward surface. In a recent paper by Mason and daForno6 the opportunities for performance gains through nonlinear aerodynamic effects are discussed including the impact on nonlinear computational analysis and design. A convincing case is made for the necessity of the inclusion of nonlinear effects particularly for wing leading edges and imbedded shock waves for the aerodynamic design and analysis of wings.

A NEW SPECIMEN OF THE MOUNT DOOLING IRON METEORITE FROM MOUNT MANNING, WESTERN AUSTRALIA

A 701 kg iron meteorite has recently been discovered near the Mount Manning Range in Western Australia. The meteorite has a fan‐shaped or delta wing configuration, one side being smooth and slightly concave with a well‐defined fusion crust, whilst the other side is rough, convex and possesses numerous regmaglypts. It is probable that the meteorite pentrated much of the earth's atmosphere in an aerodynamically stable orientation, typical of the stalled attitude of delta wing aircraft. The meteorite is a member of Chemical Group 1C. A comparison of the chemical composition, surface features, microstructure and location of this meteorite with the Mount Dooling meteorite confirms that the find is a larger specimen of Mount Dooling. It is possible that other fragments of the Mount Dooling meteorite may be found in the Mount Manning Range region.

Three-Dimensional Flow Separations on Aircraft and Missiles

William J. Rainbird; received his B.E.(Mech) at Canterbury University, Christchurch, New Zealand in 1947, and was awarded a Government scholarship to the College of Aeronautices, Cranfield, England, where he left in 1950 with a Post Graduate Diploma, with distinction, specializing in aerodynamics. After a year in Stockholm, Sweden working on transonic and supersonic wind-tunnel design at F.F.A., he returned to New Zealand and lectured at Canterbury University until September 1954. Returning to the College of Aeronautics, Cranfield that year, he was Lecturer in General and Experimental Aerodynamics for 3 years. Mr. Rainbird worked at the National Aeronautical Establishment, a Division of the National Research Council, Ottawa, from January 1958 to July 1970 and was a Senior Research Officer and Head of the High Speed Aerodynamics Section. In July 1970, Mr. Rainbird was appointed Professor of Engineering and Chairman of the Division of Aerothermodynamics, at Carleton University, Ottawa. In 1967, Mr. Rainbird spent a year’s study and refresher leave at the University of California, San Diego, Department of Aerospace and Mechanical Engineering Sciences. He is a Fellow of the C ASI, Associate Fellow of the Royal Aeronautical Society, Member of the AIAA, and was a Canadian Member of the AGARD Fluid Dynamics Panel from 1964-1969. His main field of specialization and interest in aerodynamics is in threedimensional boundary layers and separation. David J. Peake is a graduate of University of Bristol, England, Department of Aeronautical Engineering [B.Sc., 1960 (First Class Honors); M.Sc., 1962], Concurrent with University training, Mr. Peake completed an Engineering Student Apprenticeship at National Gas Turbine Establishment, Pyestock, Farnborough, England, during 1956- 1960. He remained at NGTE until May 1962 as Scientific Officer, working with Dr. Brian Stratford on the control of supersonic turbulent boundary layers with tangential air injection, and supersonic inlet design. Mr. Peake moved to National Research Council, Ottawa, Canada in May 1962, joining the Council's National Aeronautical Establishment in January 1964, in the High Speed Aerodynamics Laboratory. He worked with W. J. Rainbird in the investigation of three-dimensional separations in both subsonic and supersonic flow; transonic jet-flapped aerofoils, and axisymmetric turbulent boundary layers, all at high Reynolds numbers. Mr. Peake is studying for Ph.D. (parttime) at Carleton University, Ottawa and has membership in AIAA, RAeS and CASI. He has recently been awarded a Ph.D. from the University of Bristol for submitted published work. Edward G. Atraghji is a graduate of University of London (B.Sc., 1962). He conducted research on slender delta wing configurations at Queen Mary College, London, between 1962 and 1966. He joined the National Research Council, NAE Division, High Speed Aerodynamics Laboratory, 1966. Topics of research have included the flows about various aircraft and missile configurations.