Planar Wing Research Articles

Aerodynamic Characteristics of Spanwise Cambered Delta Wings

A wind-tunnel investigation was undertaken to determine the effect of various spanwise camber distributions on the aerodynamic characteristics of a 75-deg delta wing. Data presented include force balance, seven-hole probe survey, as well as surface e ow visualization and vortex burst trajectories. The experimental results indicate that some of the tested dihedral and anhedral geometric variations can augment lift signie cantly, requiring only minor modie cations to the planar wing geometry to implement. It was found that anhedral generally promotes lift augmentation compared to the basic wing, whereas dihedral has the opposite effect. Anhedral and dihedral have the greatest effect when applied near the wing tips, with anhedral enhancing lift through an increase in strength of both the primary and secondary wing leading-edge vortices. Nonplanarity had a small effect on the vortex burst characteristics.

Effects of Anhedral and Dihedral on a 75-deg Sweep Delta Wing

An investigation into the effects of spanwise camber in the form of anhedral and dihedral on a 75-deg sweep delta wing is detailed. Data are presented encompassing force balance, surface pressure measurement, seven-hole probe surveys, and vortex burst trajectories. The results show that the net effect of this form of nonplanarity is an increase in lift for anhedral and a decrease in lift for dihedral compared to the planar wing. Small anhedral angles aremosteffectivein augmenting lift. Anhedral doesnot appearto greatly augmentthestrength of theleading-edge vortex. The major benee t from anhedral would appear to be due to its displacing effect on the vortex trajectory: drawing the vortex closer to the wing surface and inboard compared to the planar wing. As the vortex is drawn inboard, its induced surface loading acts on a greater area of the wing. Dihedral also draws the vortex closer to the wing surface (to a greater extent then anhedral ) while moving the vortex toward the wing leading edge. In addition, anhedral doesnotappearto introduceany detrimental effectson longitudinal stability and doesnot incur any penalties in terms of vortex burst characteristics.

Axis Expansion Method for Nearly Two Dimensional Objects - Abstract

An approximate theory of scattering by nearly two-dimensional (2D) objects (plane wings, leg bones,...) is presented and implemented numerically in the case of perfectly conducting diffraction gratings slightly modulated in the direction of the grooves. Derived from the Waterman formalism, this theory leads to the solution of several problems of scattering by 2D objects, illuminated by a sum of plane waves. It is shown from numerical results that this theory is able to deal with nearly two-dimensional objects in condition where, as far as we know, the only numerical tools are based on rigorous theories of three-dimensional scattering problems.

Aerodynamic characteristics of multi-surface aircraft configurations

This paper describes the wind tunnel testing of a specially designed aircraft model allowing systematic variation of geometric parameters related to overall aircraft configurations. The experiment work was carried out in the 1.8 m low-speed wind tunnel at VZLU, Aeronautical Research and Test Institute in Prague. The resultant data created an aerodynamic database for numerical modelling and verification. In addition, numerical validation of CFD package FLUENT was performed in the computerised fluid dynamic laboratory at the Department of Aerospace Engineering, University of Glasgow as a part of the ongoing research collaboration between both institutions. The wind-tunnel test program had two aims•to provide basic aerodynamic data effects of multi-surface aircraft configurations with a view to assessing the degree to which specific design features such as a combination of canard, wing and tail plane are beneficial to aircraft aerodynamic performance•to provide an aerodynamic database for numerical validation. The aerodynamic data for different geometrical aircraft configurations with the same wing, fuselage, horizontal tail and canard surfaces were compared. The results show the magnitude of lift, drag and pitching moment depending on the angle of attack and various positions of canard, wing and horizontal tail with respect to one another. They have been compared with studied references. The experimental results are in accordance with theoretical predictions. to provide basic aerodynamic data effects of multi-surface aircraft configurations with a view to assessing the degree to which specific design features such as a combination of canard, wing and tail plane are beneficial to aircraft aerodynamic performance to provide an aerodynamic database for numerical validation.

Cost-Effective Multipoint Design of a Blended High-Speed Civil Transport

The multipoint design of a high-speed civil transport (HSCT) is considered. Low computational costs are maintained by parameterizing the HSCT using a novel method of surface generation known as the partial differential equation method. This method enables a reduction in the computational costs because of its ability to describe complex surfaces using a small number of parameters. The two design points that are considered are the subsonic and supersonic (Mach 2) flight regimes with design criteria of maximized lift and minimized wave drag, respectively. Subsonic lift is estimated using Panel Method Ames Research Center, whereas supersonic wave drag is estimated using the Harris wave-drag code (which is based upon linearized potential flow theory and the supersonic area rule). An efficient quasi-Newton method of optimization is used to optimize a weighted combination of subsonic lift and supersonic wave drag subject to the constraints of a fixed wing plan area and fixed fuselage and wing volumes. A variety of test problems are considered in which the weightings of the two design criteria are varied

Gurney Flap Experiments on Airfoils, Wings, and Reflection Plane Model

The effect of Gurney e aps on two-dimensional airfoils, three-dimensional wings, and a ree ection plane model were investigated. There have been a number of studies on Gurney e aps in recent years, but these studies have been limited to two-dimensional airfoil sections. A comprehensive investigation on the effect of Gurney e aps for a wide range of cone gurations and test conditions was conducted at Wichita State University. A symmetric NACA 0011 and a cambered GA (W)-2 airfoil were used during the single-element airfoil part of this investigation. The GA (W)-2 airfoil was also used during the two-element airfoil study with a 25% chord slotted e ap dee ected at 10, 20, and 30 deg. Straight and tapered ree ection plane wings with natural laminar e ow (NLF) airfoil sections were tested for the three-dimensional wing part of this investigation. A fuselage and engine were attached to the tapered NLF wing for the ree ection plane model investigation. In all cases the Gurney e ap improved the maximum lift coefe cient compared to the baseline clean cone guration. However, there was a drag penalty associated with this lift increase.

Hysteresis of a separated variable-velocity flow about a straight-wing model

STRAIGHT-WING MODEL B. Yu. Zanin UDC 533.69.011 It is known that if the angle of attack of a wing increases and decreases sequentially, its aerodynamic characteristics, such as lift and drag [1], change, which leads to an ambiguous flow pattern within a certain range of the angles of attack [2, 3]. The author [4] and Lushin [5] have revealed that the hysteresis phenomena can be observed in the flow about a wing mounted at a constant angle of attack if the flow velocity first increases and then decreases. These papers dealt with a global separation (leading-edge stall) on a straight-wing model. The separation was shown in [4] to be eliminated by an increase in velocity, whereas it occurs again at a different (lower) velocity, i.e., the hysteresis is observed within a certain range of flow velocities. In this case, the angle of attack was not changed. It was also found [4] that the irreversible flow can reattach to the wing surface under the acoustic action on the flow in the hysteresis range of velocities, i.e., the separation is not restored when the sound is switched off. For flow velocities lower than the hysteresis ones, acoustic forcing leads to reversible reattachment -- the separation is restored after the sound effect is eliminated. With a global flow separation (stall) from the leading edge, a vast separation region is formed above the wing. This region extends over the entire wing surface from the separation line to the trailing edge. An important feature of global separation, which has been noted by many authors, is flow three-dimensionality in the separation region, which manifests itself in the emergence of a pair of large-scale vortices rotating in the wing plane (see, for example, [3, 5-8]). A typical flow pattern on the upper wing surface for a global stall is shown in Fig. 1. The contribution of the three-dimensional vortex structure of the separated flow to the appearance of hysteresis has not yet been studied. The goal of the present paper is to obtain flow patterns on the wing surface in the hysteresis range of flow velocities and to determine which changes in the three-dimensional structure of the separated flow are caused by an increase and a sequential decrease of the flow velocity and how these changes are related to the hysteresis. In addition, the effect of acoustic forcing on the flow pattern has been studied. The experiments were performed in a low-turbulence wind tunnel at the Institute of Theoretical and Applied Mechanics, Siberian Division, Russian Academy of Sciences. Use was made of the model with a rectangular wing in plan and with a symmetric profile whose relative thickness was 10%. The model span was 945 mm, and the chord length was 196 mm (aspect ratio 4.82). Top plates were installed at the model edges to avoid flow spillage. A 8 ~ angle of attack was kept constant during the experiments. The results were obtained using the oil-film visualization technique. With an increase in the flow velocity from 0 to 22 m/sec, stalling from the leading edge (global separation) occurred, which was confirmed by a hot-wire anemometer. Upon reaching a velocity of

A problem of minimal induced resistance in unsteady motion

AbstractEnergy considerations are used to find the spanwise circulatory lift distribution yielding minimal induced resistance for a harmonically oscillating planar wing in a steady incompressible flow. Under the restriction that the time-averaged and oscillatory constituents of the wing's circulatory lift are known, the optimal distribution is such that its time-averaged constituent is elliptical, but the oscillatory one is generally not. The latter varies from elliptical to rectangular as the frequency of the oscillations increases. When the oscillatory part of the lift is not prescribed, present results infer that its reduction — as, for example, by an elastic twist — will typically reduce the drag, even if the twist yields a non-optimal (for that lift) lift distribution.

Hypersonic flow past conical V-shaped wings

Hypersonic flow past delta wings with a V-shaped cross-section has been investigated both theoretically and experimentally. Much attention is given to the examination and classification of possible conical flow patterns in the vicinity of the windward surface using the thin shock layer method. Solutions for shocks both attached to and detached from the leading edges them are obtained. It is shown that qualitatively new flow patterns can appear in the flow past V-shaped wings as compared with the case of a planar wing.

Study of Noise on a Small‐Scale Hovering Tilt Rotor

A study of the noise from a small-scale, semi-span test of a hovering tilt rotor is presented. The effect on noise of the tilt rotor fountain and of several tilt rotor configuration changes is examined. Measurments were made of an isolated rotor; rotor and wing; rotor and image plane; and rotor, wing and image plane. This last configuration models a tilt rotor. With a tilt rotor, the wing turns some of the rotor downwash inboard. When the flows meet, some of the flow is forced upward creating a fountain. Some of the fountain is ingested into the rotor disc. Rotor fountain interactions produce very unsteady impulsive noise to the rear of the aircraft. Configuration changes to investigate download reduction included increasing the rotor/wing separation, deflecting the flap, adding the nacelle, adding vortex trapping plates and blowing on the upper surface. These configuration changes produced little or no effect on the noise, although they changed the net download. Increasing the wing span, however, reduced the annoying content of this impulse noise about 2dB.

Supercritical regimes of hypersonic flow past a planar yawed delta wing

Hypersonic viscous perfect gas flow past a planar delta wing in the viscous-inviscid interaction regime is considered. The effect of the yaw angle on the parameters of the laminar boundary layer on the cold wing and the formation of subcritical and supercritical flow regions is studied.

Drag reduction by wing tip slots in a gliding Harris' hawk, Parabuteo unicinctus

The anterior-most primary feathers of many birds that soar over land bend upwards and separate vertically to form slotted wing tips during flight. The slots are thought to reduce aerodynamic drag, although drag reduction has never been demonstrated in living birds. Wing theory explains how the feathers that form the tip slots can reduce induced drag by spreading vorticity horizontally along the wing and by acting as winglets, which are used on aircraft to make wings non-planar and to spread vorticity vertically. This study uses the induced drag factor to measure the induced drag of a wing relative to that of a standard planar wing with the same span, lift and speed. An induced drag factor of less than 1 indicates that the wing is non-planar. The minimum drag of a Harris' hawk gliding freely in a wind tunnel was measured before and after removing the slots by clipping the tip feathers. The unclipped hawk had 70­90 % of the drag of the clipped hawk at speeds between 7.3 and 15.0 m s-1. At a wing span of 0.8 m, the unclipped hawk had a mean induced drag factor of 0.56, compared with the value of 1.10 assumed for the clipped hawk. A Monte Carlo simulation of error propagation and a sensitivity analysis to possible errors in measured and assumed values showed that the true mean value of the induced drag factor for the unclipped hawk was unlikely to be more than 0.93. These results for a living bird support the conclusions from a previous study of a feathered tip on a model wing in a wind tunnel: the feathers that form the slotted tips reduce induced drag by acting as winglets that make the wings non-planar and spread vorticity both horizontally and vertically.

What makes the difference between shock boundary-layer interaction on an airfoil and on a swept wing?

Swept wing flows are characterized by the curvature of the streamlines in the projection to the wing plan and by the skewing of the velocity profile in the boundary layer. The aerodynamic performance of supercritical wings at transonic speeds is trongly influenced by the interaction between a weak shock front and a turbulent boundary layer. The characteristic elements of this interaction are the precompression, the post-shock expansion, and the shock diffusion. The differences between the interactive flow over an airfoil and over a swept wing are elaborated by the comparison between the two-dimensional case and the flow with superposed tangential velocity.

Comment on 'Canard-wing interaction in unsteady supersonic flow'

A N application of various supersonic oscillatory aerodynamic computer programs was made in Ref. 1 to an idealization of the SAAB 37 Viggen canard airplane. It was suggested that the harmonic gradient method (HGM) of Chen and Liu only had accuracy comparable to the PG version of Hounjet's potential gradient method. However, the HGM has been extended from a velocity potential formulation to an acceleration potential formulation since the publication of Ref. 2. The formulation of this HGM/acceleration-potential method has been published in Ref. 4, and the associated computer program, known as the ZONA51C code, now has comparable accuracy to Hounjet's CP method. It seems appropriate to compare the newer harmonic gradient results to the NLR SPNLRI-CP results on the Viggen configuration. The calculations were carried out in the Flutter Solution Sequence of MSC/NASTRAN for the Viggen idealization of Ref. 1 shown in Fig. 1. (ZONA51D, an improved version of ZONA51C, has been incorporated into Version 67 of MSC/ NASTRAN in a new option as a variation on the subsonic Doublet Lattice Method since the aerodynamic modeling is identical for the two methods; the branching within MSC/ NASTRAN is simply determined by the Mach number.) The canard is slightly above the wing plane by a distance of 0.1 length units. The generalized force Q12 shown is the lift coefficient of the wing due to a unit rotation of the canard about its midchord. The lift coefficient is obtained by dividing the generalized force by the dynamic pressure and the square of the wing span (1.20 units). As in Ref. 1, the wing lift is calculated by considering two rigid-body modes of motion, the first being the wing alone undergoing unit plunging, and the second being unit canard rotation. (Example HA75H of Ref. 6 illustrates the application of MSC/NASTRAN to this task, including the necessary DMAP Alters.) The reduced frequency is based on a reference semichord of 1.00 unit. The results are shown in Fig. 1 and almost coincide with the NLR SPNLRI-CP results up to the high reduced frequency of k 5.0. To verify convergence, two aerodynamic element (box) idealizations were considered. In the first, the canard was divided into 8 equal width strips with 8 equal chordwise divisions, and the wing was divided into 12 equal width strips and 12 equal chordwise divisions, giving a total of 208 boxes. In the second idealization, the equal divisions were 10 x 10 on the canard and 20 x 20 on the wing for a total of 500 ReQ12

Heat pipe and surface mass transfer cooling of hypersonic vehicle structures

The problem of determining the feasibility of cooling hypersonic vehicle leading-edge structures exposed to severe aerodynamic surface heating using heat pipe and mass transfer cooling techniques is addressed. A description is presented of a numerical finite-difference-based hypersonic leading-edge cooling model incorporating poststartup liquid metal heat pipe cooling with surface transpiration and film cooling to predict the transient structural temperature distributions and maximum surface temperatures of hypersonic vehicle leading edge. An application of this model to the transient cooling of a typical aerospace plane wing leading-edge section. The results of this application indicated that liquid metal heat pipe cooling alone is insufficient to maintain surface temperatures below an assumed maximum level of 1800 K for about one-third of a typical aerospace plane ascent trajectory through the earth's atmosphere.

Two complementary approaches to estimate downwash lag effects from flight data

ALIDATION of the wind-tunnel and analytical estimates of the aerodynamic derivatives with estimates from flight test data is an important application of the system identification methodology. Reliable and accurate estimates of a large number of aerodynamic derivatives are obtained from flight data using the maximum likelihood method, although, routinely, explicit accounting for certain unsteady aerodynamic effects such as downwash and sidewash interferences pose difficulties. Recent advances in both parameter estimation methods and in flight test techniques have provided a new impetus for modeling and identification of such unsteady interference effects in aircraft dynamics.1'3 This Note addresses in some depth the problem of accounting for downwash lag effects in aircraft parameter estimation from flight data. One possibility of including such an effect is to proceed from the fundamental and to model the forces acting on the wing and tail plane separately. In such a case it becomes possible to estimate explicitly the downwash angle and also to include the lag effect. It leads, however, to a model that is nonlinear in parameters and necessarily requires an estimation program capable of handling nonlinear model postulates. The other approach is a simplification based on linearization. It leads to the first-order approximation of aerodynamic derivatives with respect to translational acceleration. The derivatives due to the body-fixed translational acceleration in the vertical direction are equivalent in the stability axis to those with respect to the rate of change of angle of attack. The second approach was first investigated in Ref. 1 to estimate from flight data the pitching moment derivatives with respect to pitch rate and rate of change of angle of attack of a highly maneuver able aircraft with large roll angle capabilities. A parameter estimation program capable of handling linear as well as nonlinear system models is used here to compare the two approaches of accounting for the downwash. Further, an attempt is made to investigate the possibility of using the linearized approach to estimate separately the two pitch damping derivatives from flight tests with a larger aircraft having limited roll angle capabilities.

Bird wings and matrices

A simple matrix method for modelling the planform of a bird's wing is introduced. Two important properties of a wing, namely aspect ratio and wing loading, are described and the matrix approach is used to make certain deductions about particular wing plans. Some examples are provided which show how the matrix transformation of one wing plan into another brings about predictable changes in wing properties. The technique is relevant to biological education because it presents a non-trivial yet concrete introduction to matrices and also gives students an appreciation of the role of algebra in shape representation and transformation. Most of the ideas reported in this brief article reflect work carried out by sixth-formers undertaking a cross-disciplinary research project on wing aerodynamics. However, the basic matrix transformation technique has also been used with third-formers; some sixth-form researchers have extended it considerably.

Control of Spider Mites and Second-Generation Corn Borers, 1988

Abstract This experiment was conducted in sprinkler-irrigated corn in Finney County, Kans. Treatments were applied with a Bull Thrush 1820 fixed-wing plane (wing span, 451/2 ft), flying 130 mph using Spraying Systems 3 hollow-cone nozzles at 22 psi and delivering a volume of 2 gal/acre. All plots were 3 plane swaths wide (except for 1 control plot, which was 2 plane swaths wide) (75 rows or 188 ft) and oriented east-west. The treatments were applied 30 Jul when the temperature was 82°F and the wind was 5 mph out of the west-southwest. Three treatments and a control were arranged in a randomized complete block design with 3 replicates. BGM, mostly large adult females, were counted weekly by searching the leaves of 5 flagged plants in the central part of each plot 28 Jul through 18 Aug. SWCB and ECB evaluations were made 8 and 9 Sep by dissecting 25 plants/plot (5 consecutive corn plants next to each flagged plant). Larvae were recorded by species for each plant.

Efficacy of Capture Against Spider Mites and Second-Generation Corn Borers, Aerial Test, 1987

Abstract This experiment was conducted in sprinkler-irrigated corn in Finney County, Kans. Treatments were applied with a Thrush 600 fixed-wing plane (wing span, 45.5 ft), flying 105 mph using Spraying Systems # 5 hollow-cone nozzles at 22 psi and delivering a volume of 2 gal/acre. Plots were 50 rows wide (125 ft or 2 plane swaths). The Furadan + Supracide treatment was applied the evening of 5 Aug. The Capture treatments were applied the morning of 6 Aug. Three treatments and a check were initially arranged in a randomized complete block design with 3 replications; however, errors and substitutions during application destroyed the blocking and 2 plots were misapplied. The data were analyzed assuming a completely randomized design with 2 missing values. Spider mites (mostly large adult females) were counted weekly by searching every other leaf (one-half plant) of 5 flagged plants in the central part of each plot 31 Jul through 31 Aug. Corn borer evaluations were made 31 Aug by dissecting 25 plants in each plot (5 consecutive corn plants next to each flagged plant). Corn borer larvae (by species) were recorded for each plant.

Efficacy of Comite Against Spider Mites Corn, Aerial Test, 1986

Abstract Experiments were conducted in 2 corn fields (south and north fields) less than 1 mile apart in Finney County, KS. The south field was under center pivot irrigation and the north field was furrow-irrigated. Two treatments and a check were arranged in a randomized complete block design with 3 rephcations in each field. Treatments were applied with a Thrush 600 fixed wing plane (wing span = 45.5 ft), flying 105 mph using Spraying Systems #5 hollow cone nozzles at 22 psi and delivering a volume of 2 gpa. Treated plots were 75-rows wide (188 ft or 3 plane-swaths). The untreated plots were 50-rows wide (125 ft). Treatments were applied on 21 Jul, ca. 1 week pre-tassel. During application, the sky was overcast, air temperature was 75°F and the wind was 3-8 mph from the SSE. Spider mites were counted weekly on every other leaf (one half plant) of 5 flagged plants in the central part of each plot between 15 Jul and 21 Aug. Mite predator populations were also recorded. Spider mite populations in one south field check plot collapsed by week 3, apparently due to predator mite activity. Therefore, this plot was treated as a missine data value after week 2. The north field was oversprayed with Furadan 4F 1.0 lb Al/acre after the 2-week counts were made.