Showing papers on "Lift-induced drag published in 1985"

Navier-Stokes computations of projectile base flow with and without mass injection

Abstract: A computational capability has been developed for predicting the flowfield about projectiles, including the recirculatory base flow at transonic speeds. In addition, the developed code allows mass injection at the projectile base and hence is used to show the effects of base bleed on base drag. Computations have been made for a secant-ogive-cylinder projectile for a series of Mach numbers in the transonic flow regime. Computed results show the qualitative and quantitative nature of base flow with and without base bleed. Base drag is computed and compared with the experimental data and semiempirical predictions. The reduction in base drag with base bleed is clearly predicted for various mass injection rates. Results are also presented that show the variation of total aerodynamic drag both with and without mass injection for Mach numbers of 0.9 < M< 1.2. The results obtained indicate that, with further development, this computational technique may provide useful design guidance for projectiles. MAJOR area of concern in shell design is the accurate prediction of the total aerodynamic drag. Both the range and terminal velocity of a projectile (two critical factors in shell design) are directly related to the total aerodynamic drag. The total drag for projectiles can be divided into three components: 1) pressure drag (excluding the base region), 2) viscous (skin friction) drag, and 3) base drag. At transonic speeds, base drag constitutes a major portion of the total drag. For a typical shell at M = 0.90, the relative magnitudes of the aerodynamic drag components are: 20% pressure drag, 30% viscous drag, and 50% base drag. The critical aerodynamic behavior of projectiles, indicated by rapid changes in the aerodynamic coefficients, occurs in the transonic speed regime and can be attributed in part to the complex shock structure existing on projectiles at transonic speeds. Therefore, in order to predict the total drag for projectiles, computation of the full flowfield (including the base flow) must be made. There are few reliable semiempirical procedures that can be used to predict shell drag; however, these procedures cannot predict the effects of mass injection. The objective of this research effort was to develop a numerical capability, using the Navier-Stokes computational technique, to compute the flowfield in the base region of projectiles at transonic speeds and thus to be able to compute the total aerodynamic drag with and without mass injection. The pressure and viscous components of drag generally cannot be reduced significantly without adversely affecting the stability of the shell. Therefore, recent attempts to reduce the total drag have been directed toward reducing the base drag. A number of studies have been made to examine the total drag reduction due to the addition of a boattail.1 Although this is very effective in reducing the total drag, it has a negative impact on the aerodynamic stability, especially at transonic

Length adjustable strut link with low aerodynamic drag

Abstract: A low aerodynamic drag structural link suitable for use within the housing of a turbofan jet engine; the link includes length adjustment capability, pivotal end mounting provision, maintained airstream orientation capability, low mass and jam nut length and orientation locking. Several variations in link construction including a single ball and socket arrangement, varying link cross-section along its longitudinal length and the use of fairing nose and tail inserts are disclosed.

Supercritical airfoil drag reduction by passive shock wave/boundary layer control in the Mach number range .75 to .90

Abstract: Airfoils operating in the transonic region are subject to large increases in drag due to shock wave/boundary layer interactions. The concept of passive shock wave/boundary layer control seeks drag reduction by placing a thin cavity with a porous top surface at the airfoil chordwise position where a shock wave would normally occur. The higher pressure behind the shock wave circulates flow through the cavity to the lower pressure ahead of the shock wave. The effects from this circulation prevent boundary layer separation and reduce entropy increases through the shock wave. In this investigation this concept is studied at a freestream Mach number range of .75 and .90. The Mach number distributions over the model, the wake impact pressure surveys used to determine profile drag and schlieren photographs for 2.8 percent porosity and solid airfoil cases are presented and compared. Results indicate that the profile drag coefficient can be reduced by as much as 40 percent through the use of this passive drag control system.

Aerodynamic tradeoff study of conventional, canard, and trisurface aircraft systems

Abstract: Conventional, canard, and three-surface aircraft configurations are investigated analytically to determine each configuration's induced and viscous drag under trimmed conditions. A three-surface vortex lattice method is used to trim the aircraft, as well as to predict the induced drag of each configuration. A vortex panel method in conjunction with the momentum integral boundary-layer method is used to predict inviscid and viscous characteristics. Parameters varied including wing to stabilator surface area ratio, static margin, canard to tail loading ratio, and CL^ . For all of the parameters considered, the conventional configuration had the highest ^tnm^' ^ e " stabilator aspect ratios, the CL . /CD of the conventional aircraft was the highest, whereas for the highest stabilator aspect ratio considered the canard configuration had the highest CL /Cp. The trisurface was superior to the canard at the lower aspect ratio with the canard becoming superior at the higher values.

The Theoretical Minimum Induced Drag of Three-Surface Airplanes in Trim

Abstract: Using the theories and theorems of Prandtl and Munk it is shown that "ideal" minimum induced drag can be achieved with a modern "three-surface" airplane in trim if equal and opposite vertical loads are applied by the forward and aft trimming surfaces. The minimum induced-drag trim condition can be attained at any center-ofgravity (e.g.) location because the two individual trim surface loads can be any size and sign as long as their sum is zero. This result is shown to be entirely consistent with the less favorable results for conventional and canard "two-surface" airplanes derived by others from the same theoretical basis. These types require zero trim surface load for minimum induced drag. This can be attained at only one aft e.g. location on a tail-aft design, and cannot be attained on a canard which requires a further forward e.g. location for inherent positive static longitudinal stability. The nonoptimum surface loadings required for trim and stability over a finite e.g. range on two-surface airplanes are reviewed. It is shown that each type incurs an induced-drag penalty that is related to the trim surface load. This penalty is significantly higher on the canard two-surface airplane due to the much larger surface loading required for trim.

A method for flight-test determination of propulsive efficiency and drag

Abstract: A flight-test method is described from which propulsive efficiency as well as parasite and induced drag coefficients can be directly determined using relatively simple instrumentation and analysis techniques. The method uses information contained in the transient response in airspeed for a small power change in level flight in addition to the usual measurement of power required for level flight. Measurements of pitch angle and longitudinal and normal acceleration are eliminated. The theoretical basis for the method, the analytical techniques used, and the results of application of the method to flight-test data are presented. Flight-test data showed performance parameters measured with a standard deviation of about 0.8% for propulsive efficiency, 0.3% for parasite drag coefficient, and 8% for the airplane efficiency factor, e.

Impact of Airfoil Profile on the Supersonic Aerodynamics of Delta Wings

Abstract: A nonlinear flow model was employed to predict the flowfield, pressure and force data for delta wings at supersonic speeds. The goal of the study was to investigate the influence the airfoil profiles has on the wing aerodynamics. The analysis covers wing aspect ratios from 0.5-3.0 with leading edge sweep of 0.5-4.0 on diamond, circular arc and NACA modified 4-digit airfoils. Nonlinear aerodynamics are approximated with nonlinear zero-lift wave drag curves, yielding results significantly different from those obtained from linear calculations. The analytical technique, useful in preliminary design studies, indicates in all cases that 90 percent of wave drag is generated at the wing apex and trailing edge.

Wing span loads of complex high-lift systems from wake measurements

Abstract: H IGH-lift systems of transport aircraft are very complex, usually consisting of a wing with leading-edge devices and multiple slotted trailing-edge flaps. Power plant installations and associated flap cutouts add to the geometric complexity. It is difficult to make accurate wind tunnel measurements of the loads on such a configuration since model flaps are often too small to allow installation of a sufficient number of surface pressure taps. The purpose of this Note is to show that accurate wing span loads can be obtained from wake surveys without the use of pressure taps. Wake surveys are usually conducted to measure profile drag and vortex drag (e.g., Refs. 1-3) or to gain a qualitative understanding of the flowfield. Such experiments are timeconsuming and expensive since a large number of data points must be acquired to get accurate drag data. Performing detailed wake surveys, however, becomes more attractive if the measurements will yield wing span loads in addition to drag. This Note presents wake data of a twin-engine transport in a takeoff configuration and outlines a method of calculating wing loads based on well-known concepts that relate loading to the strength of trailing vortices. The method has some features in common with that of ElRamly and Rainbird, who measured wing span loads on simple wing geometries. It differs in principle from the method of Orloff, who enriches the measured wake data using lifting line theory in order to find the circulation of a wing section. The method described below calculates wing circulation directly from the measured data without recourse to theoretical enrichment.

Full-Scale, Ort-Road Study of the Effect of Automobile Shape on its Aerodynamic Characteristics, and Comparison with Small-Scale Wind Tunnel Results

Abstract: The design of passenger vehicles for improved aerodynamic characteristics will result in reduced fuel consumption and better road handling during high-speed driving. In this research, techniques were developed to measure the aerodynamic drag and lift forces acting on a full-scale vehicle under road conditions and then were compared with results obtained on reduced-scale models in a wind tunnel. A number of configurations which characterize common vehicle forms were investigated for their effect on aerodynamic efficiency and fuel consumption. Experimental speeds were between 70 and 110 km/h, these being representative of highway driving conditions. A typical passenger vehicle of the three-box type was selected for the experiments, and its exterior form was altered by means of attaching various configurations to its front, rear, and underbody portions. These additions transformed the original vehicle into a "fastback" and "station wagon," and were used in combination with underbody alterations, such as front spoiler, side skirts, and smooth underbody. During road experiments, drag force was measured by means of a telemetric system receiving data on drive-shaft strains, whereas lift forces were measured by relative vertical displacements in the front and rear suspensions. Statistical analyses showed that the different configurations had a significant effect on the aerodynamic forces. The change in configurations brought about a maximum reduction in drag coefficient of 51%, relative to the original vehicle. As a result, fuel consumption was reduced by 13% (at 110 km/h). Lift forces dropped by as much as 47%. The most effective components were a smooth underbody and a "fastback" form for drag, and a smooth underbody and front spoiler for low lift. Results of the road experiments showed a reasonable correlation with those obtained using reduced-scale models in a wind tunnel.

Arrays for minimum wave drag of bodies of revolution

Abstract: The wave drag of two identical Sears-Haack bodies at transonic and supersonic speeds has been determined by using the supersonic area rule. The solution is found for these bodies displaced parallel to each other, both laterally and longitudinally. The results show that the drag of a pair of bodies can be either doubled, or nearly halved, depending upon the lateral and longitudinal spacings of the bodies. The magnitude of this drag is determined by the degree of mutual interference between the bodies. It is shown how reductions in wave drag can be obtained by proper spacing of external bodies. The regions of favorable mutual interference are delineated. It is also shown how to apply the two-body results to many-body arrays. Some remarks are made on applying the results to store-airframe interference and on further aspects of the store-airframe drag problem.

Hypersonic characteristics of an advanced aerospace plane

Abstract: A series of hypersonic wind-tunnel tests have been conducted in the NASA Langley Hypersonic Facilities Complex to obtain the static longitudinal and lateral-directional aerodynamic characteristics of an advanced aerospace plane. Data were obtained at 0 to 20 deg angles of attack and -3 to 3 deg angles of sideslip at Mach numbers of 6 and 10 in air and 20 in helium. Results show that stable trim capability exists at angles of attack near maximum lift-drag ratio (L/D). Both performance and stability exhibited some Mach number dependency. The vehicle was longitudinally unstable at low angles of attack but stable at angles of attack near and above maximum L/D. It was directionally unstable with positive dihedral effect. The rudder showed an inability to provide lateral-directional control, and removing the vertical tail resulted in increased directional instability. Analytical predictions of the static longitudinal aerodynamic coefficients gave relatively good comparisons with the experimental data.

Experimental and theoretical study of the longitudinal aerodynamic characteristics of delta and double-delta wings at Mach numbers of 1.60, 1.90, and 2.16

Abstract: An experimental and theoretical study was conducted to investigate the supersonic aerodynamic characteristics of delta and double-delta wings. Testing was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. The double-delta wings exhibited lower zero-lift drag values than the delta wings having the same aspect ratio, whereas delta wings provided the lower drag due to lift. Deflections of the trailing-edge flaps for pitch control revealed that the induced aerodynamic forces were only a function of the flap planform and were independent of wing planform. The supporting theoretical analysis showed that the supersonic design and analysis system (SDAS) did not consistently predict all the longitudinal aerodynamic characteristics of the low-sweep, low-fineness-ratio wing-body configurations under investigation.

An evaluation of three helicopter rotor sections

Abstract: Three helicopter rotor sections were tested in the NASA Ames Research Center 2- by 2-Foot Transonic Wind Tunnel over a Mach range from 0.2 to 0.88. The sections tested had maximum thickness/chord ratios of 0.078, 0.09, and 0.10. The thickest section was of early technology and had been tested previously in other wind tunnels. This section was included in the investigation to establish a basis for comparing the two thinner sections, which were of recent design. The results of the investigation showed that the pitching-moment characteristics for the three airfoil sections were acceptable. The drag divergence Mach numbers for the three sections were 0.80, 0.825, and 0.845 in order of decreasing thickness.

Study of winglets applied to biplanes

Abstract: The possibility of improving the aerodynamic characteristics of a biplane configuration by adding winglets is examined both theoretically and experimentally. Theoretical calculations show good agreement with experiment in predicting inviscid drag due to lift. Theoretical and experimental results indicate that the addition of winglets to an optimized biplane configuration can increase the ideal efficiency factor by up to 13 percent, as well as increasing the lift curve slope and maximum lift coefficient.

Interdependence of parameters important to the design of subsonic canard-configured aircraft

Abstract: An analysis is made of the interrelationship of the longitudinal parameters important to the aerodynamic design of an efficient canard or tandem wing configuration. It is shown that theoretical configuration span efficiencies substantially greater than one are feasible with the proper choice of parameters. This improvement can translate into significantly increased lift/drag ratios assuming fixed spans. The Prandtl-Munk relationship for induced drag is used as a convenient qualitative guide, with stability and trim criteria superimposed. An 'aspect-ratio ratio' parameter is introduced to aid in optimizing a configuration longitudinally. It is shown that a canard/wing 'aspect-ratio ratio' of approximately 3/2 to 2 is necessary to achieve peak span efficiency for a given span ratio and gap, assuming representative parameters.

Extension to the Panel Code NALSOP0509 to Include Designof Wing-Body Configurations for Minimum Induced Drag at given lift and pitching moment

Abstract: The code NALSOP0509, available in the SOFFTS Library of NAL has been extended to include a design of wing-body combinations to, minimize induced drag for specified lift and pitching moments at sub- or super--sonic flow conditions, Method of Lagrangean multiple is used to obtain the optimum camber and twist distributions for the wing in the presence of a body. SOFFTS Library of NAL has now been updated with this version of the code.

First order analysis of the effect of pitching on the drag coefficient

Abstract: A first order analysis was made for the drag coefficient of a pitching NACA 0015 airfoil below stall. The inviscid velocity ditribution for a translating NACA 0015 airfoil was superimposed with the additional circulation velocity for a pitching ellipse. The resulting velocity distribution was used to numerically integrate a momentum/boundary layer formulation to obtain the drag coefficient. For both laminar and turbulent boundary layers it was found that the effect of pitching on the drag coefficient can be approximated by a shift in angle of attack. The shift angle was found to be a linear function of the pitching velocity and to be less than the induced angle of attack caused by the pitching.

Arrays of bodies of revolution for minimum wave drag

Abstract: The wave drag of two identical Sears-Haack bodies at transonic and supersonic speeds has been determined by using the supersonic area rule. The solution is found for these bodies displaced parallel to each other, both laterally and longitudinally. The results show that the drag of a pair of bodies can be either doubled, or nearly halved, depending upon the lateral and longitudinal spacings of the bodies. The magnitude of this drag is determined by the degree of mutual interference between the bodies. It is shown how reductions in wave drag can be obtained by proper spacing of external bodies. The regions of favorable mutual interference are delineated. It is also shown how to apply the two-body results to many-body arrays. Some remarks are made on applying the results to store-airfram e interference and on further aspects of the store-airframe drag problem. a b CD CDl ACD D