Showing papers on "Lift-induced drag published in 1981"

Design and experimental results for a flapped natural-laminar-flow airfoil for general aviation applications

Abstract: A natural-laminar-flow airfoil for general aviation applications, the NLF(1)-0416, was designed and analyzed theoretically and verified experimentally in the Langley Low-Turbulence Pressure Tunnel. The basic objective of combining the high maximum lift of the NASA low-speed airfoils with the low cruise drag of the NACA 6-series airfoils was achieved. The safety requirement that the maximum lift coefficient not be significantly affected with transition fixed near the leading edge was also met. Comparisons of the theoretical and experimental results show excellent agreement. Comparisons with other airfoils, both laminar flow and turbulent flow, confirm the achievement of the basic objective.

An Aerodynamic Analysis of Bird Wings as Fixed Aerofoils

Abstract: The aerodynamic properties of bird wings were examined at Reynolds numbers of 1-5 × 10 4 and were correlated with morphological parameters such as apsect ratio, camber, nose radius and position of maximum thickness. The many qualitative differences between the aerodynamic properties of bird, insect and aeroplane wings are attributable mainly to their differing Reynolds numbers. Bird wings, which operate at lower Reynolds numbers than aerofoils, have high minimum drag coefficients (0·03-0·13), low maximum lift coefficients (0·8-1·2) and low maximum lift/drag ratios (3–17). Bird and insect wings have low aerofoil efficiency factors (0·2-0·8) compared to conventional aerofoils (0·9-0·95) because of their low Reynolds numbers and high profile drag, rather than because of a reduced mechanical efficiency of animal wings. For bird wings there is clearly a trade-off between lift and drag performance. Bird wings with low drag generally had low maximum lift coefficients whereas wings with high maximum lift coefficients had high drag coefficients. The pattern of air flow over bird wings, as indicated by pressure-distribution data, is consistent with aerodynamic theory for aeroplane wings at low Reynolds numbers, and with the observed lift and drag coefficients.

Oscillating flow over steep sand ripples

Abstract: Although the form and dimensions of steep vortex ripples are well studied in relation to the oscillating flow which generates them, nevertheless the accompanying fluid motion is not yet understood quantitatively. In this paper we present a method of calculation based on the assumption that the sand-water interface is fixed and that the effect of sand in suspension is, to a first approximation, negligible. The method employs a simple conformal transformation of the fluid flow onto the exterior of a polygon, and thence onto the interior of a unit circle. The initial, irrotational flow is represented by a logarithmic vortex at the centre of the circle. Other vortices within the fluid are each represented by a symmetric system of P vortices and their images in the unit circle, P being the number of sides of the original polygon. Typically P is equal to 5. However, P is not limited to integer values but may be any rational number greater than 2 (see § 15). To proceed with the calculation it is assumed that separation of the boundary layer takes place at the sharp crests of the ripples, and that the shed vorticity can be represented by discrete vortices, with strengths given by Prandtl's rule. (For a typical time sequence see figures 7 and 8.) After a complete cycle, a vortex pair is formed, which can escape upwards from the neighbourhood of the boundary. The total momentum per ripple wavelength and the horizontal force on the bottom are expressible very simply in terms of the shed vortices at any instant. The force consists of two parts: an added-mass term which dissipates no energy, and a ‘vortex drag’, which extracts energy from the oscillating flow. The calculation is at first carried out with point vortices, in a virtually inviscid theory. However, it is found appropriate to assume that each vortex has a solid core whose radius expands with time like [e( t − t n )] ½ , where t n denotes the time of birth, and e is a small parameter analogous to a viscosity. The expansion of the vortex tends to reduce the total energy (which otherwise would increase without limit) at a rate independent of e. If the cores of two neighbouring vortices overlap they are assumed to merge, by certain simple rules. Calculation of the effective vortex drag in an oscillating flow yields drag coefficients $\overline{C}_D$ of the order of 10 −1 , in good agreement with the measurements of Bagnold (1946) and of Carstens, Nielson & Altinbilek (1969). The tendency for the highest drag coefficients to occur when the ratio 2 a / L of the total horizontal excursion of the particles to the ripple length is about 1·5 is confirmed. When 2 a / L = 4, the drag falls to about half its value at ‘resonance’.

Analytical Prediction of Vortex Lift

Abstract: General integral expressions are derived for the nonlinear lift and pitching moment of arbitrary wing planforms in subsonic flow The analysis uses the suction analogy and an assumed pressure distribution based on classical linear theory results The potential flow lift constant and certain wing geometric parameters are the only unknowns in the integral expressions Results of the analysis are compared with experimental data and other numerical methods for several representative wings, including ogee and double-delta planforms The present method is shown to be as accurate as other numerical schemes for predicting total lift, induced drag, and pitching moment b c c CL CD Cm CT Cs cc, ccd E2 Nomenclature = aspect ratio = wing span =chord = reference length = lift coefficient = drag coefficient = pitching moment coefficient = thrust coefficient = suction coefficient = section lift coefficient = section induced drag coefficient = section suction coefficient = pressure loading coefficient = drag = proportionality constant, Eq (32) = proportionality constant, Eq (53) = chordwise function, Eq (44) ff(rj) = span wise f unction, Eq (28) K = potential constant L =lift loading constant, Eq (5) S = suction force SR = reference area s = suction force per unit length T = leading edge thrust, Eq (7) T' = leading edge thrust per unit length V = freestream speed Wj = downwash velocity component, Eq (11) a = angle of attack F = vorticity p = freestream density £ = nondimensional chordwise coordinate 77 = nondimensional spanwise coordinate A = leading edge sweep angle Subscripts P = potential flow E =edge / = induced VLE = leading edge vortex VSE = side edge vortex

Wind-tunnel results for a modified 17-percent-thick low-speed airfoil section

Abstract: Wind-tunnel tests were conducted in the Langley low-turbulence pressure tunnel to evaluate the effects on performance of modifying a 17-percent-thick low-speed airfoil. The airfoil contour was altered to reduce the pitching-moment coefficient by increasing the forward loading and to increase the climb lift-drag ratio by decreasing the aft upper surface pressure gradient. The tests were conducted over a Mach number range from 0.07 to 0.32, a chord Reynolds number range 1.0 x 10 to the 6th power to 12.0 x 10 to the 6th power, and an angle-of-attack range from about -10 deg to 20 deg.

Wind-Tunnel Measurements of Wing-Canard Interference and a Comparison with Various Theories

Abstract: Wind-tunnel tests and analyses of the aerodynamics of wing-canard combinations for low speed applications are presented. Systematic tests are conducted in a 7 x 10 wind tunnel to explore various combinations of wing-canard vertical and horizontal positioning. The goals of the tests are (1) to investigate potential improved stalling characteristics over conventional tail-aft configurations, (2) to investigate the existence of a lift coefficient advantage, and (3) to determine induced drag levels. The measurements obtained are compared with calculations made using the Prandtl-Munk theory, and with a vortex-lattice panel code. Results indicate that the panel code gives excellent results for lift and induced drag at moderate lift coefficient, whereas Prandtl-Munk theory gives conservative results for induced drag. The application is a light transport aircraft used for short-haul operations.

Investigation of Delta Wing Leading-Edge Devices

Abstract: Leading-edge flow manipulators for alleviating the subsonic lift-dependent drag of highly-swept wings were investigated experimentally. The potential of several devices—fences, chordwise slots, pylon vortex generators, and a vortex plate concept—was evaluated in wind tunnel tests on a 60 deg cropped delta wing research model. Simultaneous balance and pressure measurements at increasing angles of attack provided an insight into the spanwise leading-edge flow development and its modification by the devices. The results demonstrated significant drag reductions through partial recovery of leading-edge suction at elevated angles of attack. In most cases, improvement in longitudinal stability also was obtained.

Optimization and performance calculation of dual-rotation propellers

Abstract: An analysis is given which enables the design of dual-rotation propellers. It relies on the use of a new tip loss factor deduced from T. Theodorsen's measurements coupled with the general methodology of C. N. H. Lock. In addition, it includes the effect of drag in optimizing. Some values for the tip loss factor are calculated for one advance ratio.

Longitudinal grooves for bluff body drag reduction

Abstract: Wind tunnel test results are presented for four axisymmetric bluff body configurations in order to determine their effect on form and pressure drag. It was found that drag reductions on the order of 40% are obtainable with an afterbody incorporating four longitudinal 'V' grooves. Although this effect may be due to the functioning of the grooves as longitudinal, continuous vortex generators, it is concluded that further research is needed to elucidate the physical basis of the test results. Optimization of the effect will be useful in base drag reduction for such vehicles as automobiles and cargo aircraft with sharply upswept afterbodies.

An improved method for wind-tunnel wall corrections deduced by iterating from measured wall static pressure

Abstract: The purpose of this research was to demonstrate the viability of a method, due to Professor W. R. Sears, for obtaining wind-tunnel wall-corrections from measurements of near-field flow parameters by an interative procedure. A case is made for the improved accuracy of this method over the standard method of images. The wall-correction method was applied to an actual wind-tunnel test of a slightly oversized wing model at low subsonic speeds (Mach number ≈ 0.1). The wind tunnel facility and experimental setup and method are described and discussed. The wall-correction method consists of iterating between the region of space exterior to the test section boundary and the one interior to it. The flow fields in both regions are defined in terms of plane singularity elements each with an unknown, constant strength distribution. The method for expressing these flow fields as a linear system and for obtaining the associated matrices is described. The boundary conditions for the inner flow are slightly different from those of the outer flow because of the presence of the wing. There are actually two different but consistent sets of boundary conditions at the wing which lead to two different but compatible calculations for the wall-correction. The near-field flow parameter measured during the wind-tunnel test is the wing perturbation velocity potential, obtained from the quantity p - pᵢ. Here, i represents any of the 46 static taps distributed over the test section walls. It was decided to use 140 singularity elements for the outer flow description; therefore, a method was devised for fitting a least-squares surface to the measured pᵢ's and integrating to obtain 140φᵢ's. The procedure for the iterations is described and the criterion for convergence to unconfined flow is presented. Test cases consisting of known, simple flows are used along the way to verify the computational methods. Finally, the wall correction to the lift curve of the wing model is presented as well as the correction at a typical tail position and the correction to the induced drag of the wing.

A review of flight-to-wind tunnel drag correlation

Abstract: Comparisons are made of wind-tunnel-model and flight drag data for various configurations representing aircraft from the mid-1940s to the 1970s. Discrepancies between model and flight data such as Reynolds number effects, wall interference, and aeroelastic problems are discussed. String support effects and the inability of models to simulate surface deflections for longitudinal trim are also studied. A wind tunnel-to-flight correlation of turbulent friction drag confirms the incompressible Karman-Schoenherr variation of turbulent skin friction with Reynolds number and the T' method for accounting compressibility effects. NASA tested 10 deg cone research indicates that model tests which are affected by tunnel noise may require the lower disturbance level environment available in flight, and it is concluded that new cryogenic facilities will improve the fidelity of model simulations of full-scale flight flow phenomena.

Numerical optimization techniques for bound circulation distribution for minimum induced drag of Nonplanar wings: Computer program documentation

Abstract: A two dimensional advanced panel far-field potential flow model of the undistorted, interacting wakes of multiple lifting surfaces was developed which allows the determination of the spanwise bound circulation distribution required for minimum induced drag. This model was implemented in a FORTRAN computer program, the use of which is documented in this report. The nonplanar wakes are broken up into variable sized, flat panels, as chosen by the user. The wake vortex sheet strength is assumed to vary linearly over each of these panels, resulting in a quadratic variation of bound circulation. Panels are infinite in the streamwise direction. The theory is briefly summarized herein; sample results are given for multiple, nonplanar, lifting surfaces, and the use of the computer program is detailed in the appendixes.

Full-scale Wind-Tunnel Study of the Effect of Nacelle Shape on Cooling Drag

Abstract: Tests were made in the Ames 40 by 80 ft Wind Tunnel of a semispan wing with a nacelle (no propeller) from a typical, general aviation twin-engine aircraft. Measurements were made of the effect on drag of the flow of cooling air through the nacelle. Internal and external nacelle pressures were measured. It was found that the cooling airflow accounts for about 13% of the total estimated airplane drag during both cruise and climb. The now of cooling air through the nacelle accounts for 30% of the airflow drag component during cruise and 42% during climb; the balance, in both cruise and climb, is attributed to [he external shape of the nacelle. It was suggested that improvements could possibly be made by relocating both the inlet and the outlet for the cooling air.

Aerodynamic Design Criteria for Supersonic Climb-Cruise Missiles

Abstract: An analytical technique is presented which establishes the relationship between aerodynamic characteristics and total mission performance for ramjet-powered missiles which are launched at low altitude and must climb to high altitude for a long-range cruise at supersonic speeds. The analytical expressions obtained can be used to determine desirable aerodynamic characteristics and their variation with velocity given missile propulsion and design parameters such as specific fuel consumption, thrust coefficient, planform area, nozzle exit area, and weight. The expressions are derived for Rutowski minimum fuel-to-climb profiles and optimum specific fuel consumption cruise conditions. The low-altitude segment of the climb is shown to be dominated by the drag coefficient CD . and its variation with velocity. At high altitudes during the climb, the induced drag factor K becomes important. Therefore, reducing Q>min at low supersonic speeds and K at high supersonic speeds reduces fuel consumption during the climb. The desirable lift coefficient at minimum drag CL() is shown to be small at low speeds and to increase with velocity. The ideal climb-cruise configuration will have a Q>mjn and K which are low at low speeds and decrease strongly with increasing speed. The CL() will also be low at low speed but increase greatly with increasing speed. These results and the analytical expressions derived can be used by the missile designer as aerodynamic design criteria during configuration development.

Full-scale wind-tunnel study of wing-fuselage interaction and comparison with paneling method

Abstract: Full-scale wind-tunnel tests, conducted to determine the effects of three different airfoil sections on the aft-fuselage drag of a low-wing aircraft, are described. The measurements indicate a maximum difference in aft fuselage drag between the three airfoils of about 0.002. Measured changes in the locations of the fuselage pressure contours with airfoil section correlated well with the changes predicted by a three dimensional paneling code. A criterion, based on the inviscid computer code, is then proposed as an indicator for possible adverse viscous interactions at the wing-fuselage juncture.

Computer program documentation for a subcritical wing design code using higher order far-field drag minimization

Abstract: A subsonic, linearized aerodynamic theory, wing design program for one or two planforms was developed which uses a vortex lattice near field model and a higher order panel method in the far field. The theoretical development of the wake model and its implementation in the vortex lattice design code are summarized and sample results are given. Detailed program usage instructions, sample input and output data, and a program listing are presented in the Appendixes. The far field wake model assumes a wake vortex sheet whose strength varies piecewise linearly in the spanwise direction. From this model analytical expressions for lift coefficient, induced drag coefficient, pitching moment coefficient, and bending moment coefficient were developed. From these relationships a direct optimization scheme is used to determine the optimum wake vorticity distribution for minimum induced drag, subject to constraints on lift, and pitching or bending moment. Integration spanwise yields the bound circulation, which is interpolated in the near field vortex lattice to obtain the design camber surface(s).

Effect of sweep angles on aerodynamic performance of double arrow wing - An analytical study

Abstract: Hybrid wing planforms are studied for adoption on supersonic transport and fighter aircraft. The free vortex sheet method is used to determine effects of the leading-edge sweep angles on the aerodynamic performance of a double arrow wing with a strake. Results show lift and drag increase with the increase of the inboard and outboard leading-edge sweep angles. However, the lift-to-drag ratio is little influenced by the changes in these sweep angles. Spanwise surface pressure distributions on the aft region are influenced by the inboard sweep angle while the outboard sweep angle has no effect on these pressures. Finally, the experimental data and predicted results are compared to show good agreement.

Optimal trajectories for maximum endurance gliding in a horizontal plane

Abstract: Nomenclature C0,C/,C2,Cj = constants of integration CD = drag coefficient CDO =zero lift drag coefficient CL = lift coefficient CL* = lift coefficient for maximum lift-to-drag ratio D = drag force E* = maximum lift-to-drag ratio h = altitude H =Hamiltonian function kj,k2,k3 =constants, =C//C0, C2/C0, C3/C0, respectively K = induced drag factor L =lift force m =mass of the vehicle Px>Py>Pu>Pt>Pe = adjoint variables associated with state variables r — radius of penetration 5 = reference area / =time u = dimensionless speed V — speed of vehicle W — weight of vehicle x,y = dimensionless coordinates X,Y = position coordinates of vehicle A =tan/>t 0 = dimensionless time X = normalized lift coefficient jn =bank angle p = density of atmosphere T = normalized time 0 = velocity yaw angle co = dimensionless wing loading

An investigation of drag reduction for a standard truck with various modifications

Abstract: A wind tunnel investigation was conducted to determine the influence of several physical variables on the aerodynamic drag of a standard truck model. The physical variables included: a cab mounted air deflector; a boattail on the rear of the cargo compartment; flow-vanes on the front of the cargo compartment; and a forebody fairing over the cab. Tests were conducted at yaw angles (relative wind angle) of 0, 5, 10, 20, and 30 degrees and Reynolds numbers of 3.4 x 100,000 to 6.1 x 100,000 based upon the equivalent diameter of the vehicles. The forebody fairing and the flow-vane with the closed bottom were very effective in improving the flow over the forward part of the cargo compartment. The forebody fairing provided a calculated fuel saving of 5.6 liters per hour (1.5 gallons per hour) over the baseline configuration for a ground speed of 88.6 km/hr (55 mph) in national average winds.

Forebody Drag Reduction

Abstract: It is shown that the mechanisms of forebody drag reduction by means of either a spike or a forward-facing jet are similar, with the maximum achievable drag reduction being of the same order. Because the jet may be a relatively cool gas, however, the forward facing jet has the additional capability of reducing the aerodynamic heating that is so severe at high Mach numbers. By means of the correlation presented, jet ejection parameters may be chosen to achieve maximum permissible forebody drag reduction. The correlation method uses a momentum coefficient that characterizes jet efflux and freestream conditions.

A vortex-filament and core model for wings with edge vortex separation

Abstract: A method for predicting aerodynamic characteristics of slender wings with edge vortex separation was developed. Semiempirical but simple methods were used to determine the initial positions of the free sheet and vortex core. Comparison with available data indicates that: the present method is generally accurate in predicting the lift and induced drag coefficients but the predicted pitching moment is too positive; the spanwise lifting pressure distributions estimated by the one vortex core solution of the present method are significantly better than the results of Mehrotra's method relative to the pressure peak values for the flat delta; the two vortex core system applied to the double delta and strake wing produce overall aerodynamic characteristics which have good agreement with data except for the pitching moment; and the computer time for the present method is about two thirds of that of Mehrotra's method.

Reduction of Aerodynamic Drag

Abstract: : A study is being made of the effect on aerodynamic drag of boundary layer irradiation by radioactive sources. A blow-down wind tunnel and a skin friction drag balance have been designed and constructed. The balance is of the null-position type, and is operated using an automatic control system to maintain the null position of the drag plate. High accuracy and stability are observed at flow velocities up to 200 m/s and resolutions of up to 0.1% of changes in drag have been achieved. In separate experiments, a study of the effect of radioactive emission on gas viscosity is being made with a specially designed torsion disc viscometer. Measurements to date have been at atmospheric pressure with both pieces of apparatus. No significant changes in drag have been found except at low flow speeds. However, future work will be extended to lower pressures where there are reports of a decrease in viscosity with radiation. An objective of new research will be to find the optimum conditions for any drag reduction.

Drag measureiments on airfoils

Abstract: The drag of airfoils is evaluated from the loss of total head in the wake. This is achieved either by an integrated rake positioned vertical to the trailing edge at some distance from the airfoil trailing edge, or the wake is traversed by a single Pitot-tube point by point and the readings are integrated subsequently. In both cases the drag is evaluated only for a particular plane in spanwise direction. If the drag was measured by a balance, the mean value over the whole wind tunnel model could be attained. But, because of the undefinable influences of the tunnel walls, this procedure cannot be used satisfactorily. In the laminar wind tunnel at the Institut for Aerodynamics, University of Stuttgart, drag is measured by an integrating rake.

Measurements in the Near-Wake of a Transport Model, to Determine the Lift, Drag and Drag Components Using Maskell's Analysis.

Abstract: : Extensive measurements have been made on a plane normal to the wake of a swept-wing/fuselage model of airbus type, at C sub L = 0.49 in incompressible flow, using a computer-controlled wake-traverse system incorporating a null-reading five-hole yawmeter probe. The results define the detailed distributions of total pressure defect, flow velocity, flow angles and streamwise vorticity in the viscous wake at that plane. The results have been analyzed within the basic theoretical framework set out by Maskell (1973), allowing the calculation of lift and drag from traverse data, and the resolution of the drag into provisionally-defined components C sub D sub I and C sub D sub II, respectively relating to the profile drag and vortex drag. (Author)