Showing papers on "Vortex lattice method published in 1984"

Lateral aerodynamics of delta wings with leading-edge separation

Abstract: An unsteady vortex lattice method is presented for the calculation of the aerodynamic forces acting on lifting surfaces undergoing complex three dimensional motion. For the present case the nonsymmetric motion of a slender delta wing was considered and the resulting lateral characteristics were calculated. The flow separation line was specified along the wing leading edge and the emanating vortex sheet shape and rollup was then calculated. Numerical results are presented for the combined high angle of attack and side slip condition and for the wing constant roll and coning motions.

Optimization of Propeller Blade Twist by an Analytical Method

Abstract: Based on the vortex lattice method and nonlinear programming, the problem for predicting optimum propeller blade twist has been formulated and solved to maximize the propulsive efficiency under the constraint of constant power consumption. The propeller is represented by a curved lifting line and a number of control points. The optimum twist distribution can be determined for a specified geometry of the lifting line. The method can be applied to complex blade shapes (swept, bent, propletted, and biblades). To demonstrate the method, the geometry of the lifting line of a straight blade and a propletted blade has been employed. The twist distribution and the ideal efficiency for the optimized and unoptimized blades are compared. The predicted improvement in ideal efficiency is about 1-6% for the optimized blade with proplet over that of the original propeller.

A vortex-lattice method for calculating longitudinal dynamic stability derivatives of oscillating delta wings

Abstract: A nonsteady vortex-lattice method is introduced for predicting the dynamic stability derivatives of a delta wing undergoing an oscillatory motion. The analysis is applied to several types of small oscillations in pitch. The angle of attack varied between + or - 1 deg, with the mean held at 0 deg when the flow was assumed to be attached and between + or - 1 deg and the mean held at 15 deg when both leading-edge separation and wake roll-up were included. The computed results for damping in pitch are compared with several other methods and with experiments, and are found to be consistent and in good agreement.

Errata and Addenda to "Application of Oscillatory Aerodynamic Theory to Estimation of Dynamic Stability Derivatives"

Abstract: T pitching moment coefficients in Ref. 1 are referred to an axis through the wing apex with the pitch axis through the quarter-chord point of the mean aerodynamic chord (MAC). It was the authors' intention to present the coefficients using the same reference and pitch axes at the quarterchord point of the MAC. The example wing had an aspect ratio of 8.0, a taper ratio of 0.25, a quarter-chord sweep of 30 deg., and was flying at a Mach number of 0.8. Using a new lattice idealization of the wing (since the original idealization was not recorded) defined by 5 equal chordwise divisions and 15 variable spanwise (narrower toward the tip) divisions, the new lift and corrected moment coefficients for the quarterchord MAC reference axis and pitch axis are presented in Table 1 for the nine values of reduced frequency previously considered. The calculations were based on the doublet-lattice method (DLM) of Ref. 2 as integrated into NASTRAN® and were carried out using MSC/NASTRAN; the original calculations were based on the DLM of Ref. 4 which was later developed into Ref. 5. Slight differences in the lift coefficients between the present and earlier solutions can be attributed to a different lattice idealization. If we extend Ref. 1 to permit different reference chords for reduced frequency (k = ub/V) and pitching moment coefficient

Wind tunnel wall interference in closed, ventilated and adaptive test sections

Abstract: A wall interference correction method for closed rectangular test sections was developed which uses measured wall pressures. Measurements with circular discs for blockage and a rectangular wing as a lift generator in a square closed test section validate this method. These measurements are intended to be a basis of comparison for measurements in the same tunnel using ventilated (in these case, slotted) walls. Using the vortex lattice method and homogeneous boundary conditions, calculations were performed which show sufficiently high pressure levels at the walls for correction purposes in test sections with porous walls. In Gottingen, an adaptive test section (which is a deformable rubber tube of 800 mm diameter) was built and a computer program was developed which is able to find the necessary wall adaptation for interference-free measurements in a single step. To check the program prior to the first run, the vortex lattice method was used to calculate wall pressure distributions in the nonadapted test section as input data for the one-step method. Comparison of the pressure distribution in the adapted test section with free-flight data shows nearly perfect agreement. An extension of the computer program can be made to evaluate the remaining interference corrections.

Application of the Vortex Lattice Method to the Three-Dimensional Theory of a Cavitating Propeller

Abstract: A new theory on a cavitating propeller is proposed, treating the cavitation threedimensionally, and using a lifting surface theory. A numerical method has also been developed based on this theory, appying the vortex lattice method and replacing the cavitation and the blade thickness by source elements, for a cavitating propeller operating in a uniform flow.To evaluate the present method, the calculations are carried out for a model propeller and compared with other prediction methods and the model test results.It is confirmed that the present method gives considerable improvements, in comparison with the conventional methods, on the evaluation of the cavity thickness due to the three-dimensional flow, and on the elimination of the instability of the solution caused by the two-dimensional theory.

Aerodynamic Design Optimization Trim Analysis of Canard Conventional Configurations

Abstract: A design study has been conducted to optimize trim cruise flight of high performance general aviation canard aircraft which achieve minimum drag. In order to investigate the advantages and disadvantages of canard configured aircraft, corresponding conventional tail-aft "baseline" aircraft were designed and used for comparison. Two-dimensional coupled lift and drag coefficient predictions of the canard and wing airfoil shapes were obtained by coupling inviscid results from a vortex panel multielement program to a momentum integral boundary layer analysis. By using the results of the two-dimensional vortex panel analysis, a vortex lattice method was employed to predict the finite wing results. The analysis utilized a turbulent airfoil and a natural laminar airfoil, which are two NASA state-of-the-art airfoil sections. The canard aircraft designs give quantitative results of wing and canard loadings, wing-to-canard moment arm ratios, and aspect ratio effects for trim cruise flight for a wide range of wing-to-canard area ratios. Both canard and baseline aircraft achieved a 25-30% improvement in performance over typical current technology aircraft, but high canard loading necessary for trim resulted in slightly poorer cruise performance of the canard aircraft for equal fuel as compared to the baseline designs. However, when takeoff gross weight was held the same by reducing the fuel payload, the canard designs achieved longer ranges than the baselines. The required positive decalage angle between the canard and the wing guarantees that the canard will stall first, thereby preventing the wing from stalling, and thus having a stall- and, hence, spin-proof configuration.

The lateral-directional characteristics of a 74-degree Delta wing employing gothic planform vortex flaps

Abstract: The low speed lateral/directional characteristics of a generic 74 degree delta wing body configuration employing the latest generation, gothic planform vortex flaps was determined Longitudinal effects are also presented The data are compared with theoretical estimates from VORSTAB, an extension of the Quasi vortex lattice Method of Lan which empirically accounts for vortex breakdown effects in the calculation of longitudinal and lateral/directional aerodynamic characteristics It is indicated that leading edge deflections of 30 and 40 degrees reduce the magnitude of the wing effective dihedral relative to the baseline for a specified angle of attack or lift coefficient For angles of attack greater than 15 degrees, these flap deflections reduce the configuration directional stability despite improved vertical tail effectiveness It is shown that asymmetric leading edge deflections are inferior to conventional ailerons in generating rolling moments VORSTAB calculations provide coarse lateral/directional estimates at low to moderate angles of attack The theory does not account for vortex flow induced, vertical tail effects

An analytical design procedure for the determination of effective leading edge extensions on thick delta wings

Abstract: An analytical design procedure for leading edge extensions (LEE) was developed for thick delta wings. This LEE device is designed to be mounted to a wing along the pseudo-stagnation stream surface associated with the attached flow design lift coefficient of greater than zero. The intended purpose of this device is to improve the aerodynamic performance of high subsonic and low supersonic aircraft at incidences above that of attached flow design lift coefficient, by using a vortex system emanating along the leading edges of the device. The low pressure associated with these vortices would act on the LEE upper surface and the forward facing area at the wing leading edges, providing an additional lift and effective leading edge thrust recovery. The first application of this technique was to a thick, round edged, twisted and cambered wing of approximately triangular planform having a sweep of 58 deg and aspect ratio of 2.30. The panel aerodynamics and vortex lattice method with suction analogy computer codes were employed to determine the pseudo-stagnation stream surface and an optimized LEE planform shape.

Modeling of the aerodynamic response to arbitrary aircraft maneuvres - A numerical validation

Abstract: The regime of validity of an aerodynamic mathematical model, applicable to describe the nonlinear aerodynamic reactions to a delta wing maneuvering at high angles of attack is investigated. An unsteady vortex-lattice method is used to compute the unsteady flowfields, and thus to evaluate the aerodynamic data required by the model, in terms of specified characteristic motions. Time-histories of the aerodynamic responses to complex motions are generated by means of the model and the evaluated aerodynamic data, and are compared with baseline aerodynamic responses obtained from direct vortex-lattice computations. The validity of the mathematical modeling approach for the maneuvering delta wing is demonstrated by the close agreement of the force and moment responses obtained from the two approaches.