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Application of the Kutta-Joukowski theorem to the relationship between airfoil lift and circulation is described. A number of formulas concerning the conduct of vortex systems derived from the theorem are presented. The application of this line of reasoning to several problems of airfoil theory and the observed relations are discussed.
The viscous transonic flow around a low-aspect-ratio wing has been computed using an implicit, three-dimensional, 'thin-layer' Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. The geometry chosen for this study is a low-aspect-ratio wing with large sweep, twist, taper, and camber. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. Using this grid, the flow around the wing was computed for a free-stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three-dimensional flow separation off the wing surface are observed. Particle path lines indicate that the three-dimensional flow separation on the wing surface is part of the roots of the tip-vortex formation. The lifting of the tip vortex before the wing trailing edge is clearly observed by following the trajectory of particles released around the wing tip.
Numerical simulations have been performed for a vertical tail configuration with deflected rudder. The suction surface of the main element of this configuration, just upstream of the hinge line, is embedded with an array of 32 fluidic actuators that produce oscillating sweeping jets. Such oscillating jets have been found to be very effective for flow control applications in the past. In the current paper, a high-fidelity computational fluid dynamics (CFD) code known as the PowerFLOW R code is used to simulate the entire flow field associated with this configuration, including the flow inside the actuators. A fully compressible version of the PowerFLOW R code valid for high speed flows is used for the present simulations to accurately represent the transonic flow regimes encountered in the flow field due to the actuators operating at higher mass flow (momentum) rates required to mitigate reverse flow regions on a highly-deflected rudder surface. The computed results for the surface pressure and integrated forces compare favorably with measured data. In addition, numerical solutions predict the correct trends in forces with active flow control compared to the no control case. The effect of varying the rudder deflection angle on integrated forces and surface pressures is also presented.
Deblurring capabilities would significantly improve the Flight Science Support Office's ability to monitor the effects of lift-off on the shuttle and landing on the orbiter. A deblurring program was written and implemented to extract information from blurred images containing a straight line or edge and to use that information to deblur the image. The program was successfully applied to an image blurred by improper focussing and two blurred by different amounts of blurring. In all cases, the reconstructed modulation transfer function not only had the same zero contours as the Fourier transform of the blurred image but the associated point spread function also had structure not easily described by simple parameterizations. The difficulties posed by the presence of noise in the blurred image necessitated special consideration. An amplitude modification technique was developed for the zero contours of the modulation transfer function at low to moderate frequencies and a smooth filter was used to suppress high frequency noise.
A fast multigrid solver for the steady, incompressible Navier-Stokes equations is presented. The multigrid solver is based upon a factorizable discrete scheme for the velocity-pressure form of the Navier-Stokes equations. This scheme correctly distinguishes between the advection-diffusion and elliptic parts of the operator, allowing efficient smoothers to be constructed. To evaluate the multigrid algorithm, solutions are computed for flow over a flat plate, parabola, and a Karman-Trefftz airfoil. Both nonlifting and lifting airfoil flows are considered, with a Reynolds number range of 200 to 800. Convergence and accuracy of the algorithm are discussed. Using Gauss-Seidel line relaxation in alternating directions, multigrid convergence behavior approaching that of O(N) methods is achieved. The computational efficiency of the numerical scheme is compared with that of Runge-Kutta and implicit upwind based multigrid methods.
As part of the subsonic transport high-lift program, flight experiments are being conducted using NASA Langley's B737-100 to measure the flow characteristics of the multi-element high-lift system at full-scale high-Reynolds-number conditions. The instrumentation consists of hot-film anemometers to measure boundary-layer states, an infra-red camera to detect transition from laminar to turbulent flow, Preston tubes to measure wall shear stress, boundary-layer rakes to measure off-surface velocity profiles, and pressure orifices to measure surface pressure distributions. The initial phase of this research project was recently concluded with two flights on July 14. This phase consisted of a total of twenty flights over a period of about ten weeks. In the coming months the data obtained in this initial set of flight experiments will be analyzed and the results will be used to finalize the instrumentation layout for the next set of flight experiments scheduled for Winter and Spring of 1995. The main goal of these upcoming flights will be: (1) to measure more detailed surface pressure distributions across the wing for a range of flight conditions and flap settings; (2) to visualize the surface flows across the multi-element wing at high-lift conditions using fluorescent mini tufts; and (3) to measure in more detail the changes in boundary-layer state on the various flap elements as a result of changes in flight condition and flap deflection. These flight measured results are being correlated with experimental data measured in ground-based facilities as well as with computational data calculated with methods based on the Navier-Stokes equations or a reduced set of these equations. Also these results provide insight into the extent of laminar flow that exists on actual multi-element lifting surfaces at full-scale high-life conditions. Preliminary results indicate that depending on the deflection angle, the slat and flap elements have significant regions of laminar flow over a wide range of angles of attack. Boundary-layer transition mechanisms that were observed include attachment-line contamination on the slat and inflectional instability on the slat and fore flap. Also, the results agree fairly well with the predictions reported in a paper presented at last year's AIAA Fluid Dynamics Conference. The fact that extended regions of laminar flow are shown to exist on the various elements of the high-lift system raises the question what the effect is of loss of laminar flow as a result of insect contamiantion, rain or ice accumulation on high-life performance.
A wind-tunnel investigation was conducted to examine the low-speed static stability and control characteristics of a 0.10 scale model of a STOL supersonic cruise fighter concept. The concept, referred to as a twin boom fighter, was designed as a STOL aircraft capable of efficient long range supersonic cruise. The configuration name is derived from the long twin booms extending aft of the engine to the twin vertical tails which support a high center horizontal tail. The propulsion system features a two dimensional thrust vectoring exhaust nozzle which is located so that the nozzle hinge line is near the aircraft center of gravity. This arrangement is intended to allow large thrust vector angles to be used to obtain significant values of powered lift, while minimizing pitching moment trim changes. Low speed stability and control information was obtained over an angle of attack range including the stall. A study of jet induced power effects was included.
Recent interest in the induced drag characteristics of crescent-shaped wings has led to a closer look at the methods used for determination of induced drag from computational aerodynamic methods. Induced drag may be computed by integration of surface pressure, or by evaluation of a contour integral in the Trefftz plane. A high-order panel method was used to study the induced drag of crescent and elliptical wings using both techniques. Induced drag computations using surface-pressure integration were strongly affected by panel density and angle of attack. Accurate drag computations for the crescent wing were obtained only when the spanwise as well as chordwise panel density was extremely high. Trefftz-plane results for the two wing planforms do not show this sensitivity to panel density or angle of attack. Span efficiencies of 0.994 for the crescent wing and 0.987 for the elliptical wing were computed by the Trefftz-plane technique. Substitution of a force-free, rolled-up wake geometry on the crescent wing did not change the pressure-integrated drag significantly. The slightly higher span efficiency of the crescent wing is attributed to a more nearly elliptical spanwise lift distribution. The chord distribution of the elliptical wing was modified to produce an elliptical span-loading on a wing with an unswept quarter-chord line. This wing demonstrated a span efficiency equal to that of the crescent wing.
Advanced methods of liquid oxygen (LO2) propellant conditioning were studied as part of an effort for increasing reliability and operability while reducing cost of future heavy lift launch vehicles. The most promising conditioning concept evaluated was no-bleed (passive recirculation) followed by low-bleed, helium injection, and use of a recirculation line. Full-scale cryogenic testing was performed with a sloped feedline test article to validate models of behavior of LO2 in the feedline and to prove no-bleed feasibility. Test data are also intended to help generate design guidelines for the development of a main propulsion system feed duct. A design-of-experiments matrix of over 100 tests was developed to test all four propellant conditioning concepts and the impact of design parameters on the concepts. Liquid nitrogen was used as the test fluid. The work for this project was conducted from October 1992 through January 1994 at the hydrogen cold flow facility of the west test area of MSFC. Test data have shown that satisfactory temperatures are being obtained for the no-bleed conditioning concept.
An analysis of the flow state on a trapezoidal wing model from the NASA 3-D High Lift Flow Physics Experiment is presented. The objective of the experiment was to characterize the flow over a non-proprietary semi-span three-element high-lift configuration to aid in assessing the state of the art in the computation of three-dimensional high-lift flows. Surface pressures and hot-film sensors are used to determine the flow conditions on the slat, main, and flap. The locations of the attachments lines and the values of the attachment line Reynolds number are estimated based on the model surface pressures. Data from the hot-films are used to determine if the flow is laminar, transitional, or turbulent by examining the hot-film time histories, statistics, and frequency spectra.
Analysis and numerical results are presented for the elastic shear stiffness of a corrugated shear web with a certain type of discrete attachments at the ends of the trough lines of the corrugations, namely point attachments to a rigid flange which interferes with the deformations of the end cross sections by preventing downward movement but permitting upward (lifting off) movement. The analysis is based on certain assumed modes of deformation of the cross sections in conjunction with the method of minimum total potential energy and the calculus of variations in order to obtain equations for the manner in which the assumed modes of deformation vary along the length of the corrugation. The numerical results are restricted to the case of equal-width crests and troughs but otherwise apply to a wide variety of geometries. They are in the form of graphs which give the overall shear stiffness as a fraction of the overall shear stiffness that could be obtained by having continuous attachment at the ends of the corrugations.
The directionality and strength of blade-vortex interactions (BVI) is explained through the radiation cone concept. BVI acoustic radiation is primarily the result of two sound mechanisms: the tip effect, and the radiation cone effect. The radiation cone effect is a highly directional mechanism which results when a lift distribution moves supersonically with respect to the fluid. After a physical explanation of the BVI mechanisms, sample cases using translating and rotating blades interacting with a straight line vortex are shown. The radiation cone concept is then applied to specific rotorcraft cases where it helps to explain zones of intense sound pressure level found in experimental results for the XV-15 tiltrotor and for a BO-105 helicopter scale model.
Recently proposed flight research at NASA Dryden Flight Research Center (DFRC) has prompted study into the aerodynamic effects of modifications made to the surfaces of laminar airfoils. The research is focused on the high-aspect ratio, laminar-flow type wings commonly found on UAVs and other aircraft with a high endurance requirement. A broad range of instrumentation possibilities, such as structural, pressure, and temperature sensing devices may require the alteration of the airfoil outer mold line as part of the installation process. This study attempts to characterize the effect of installing this additiona1 instrumentation on key airfoil performance factors, such as transition location, lift and drag curves, and stall point. In particular, the general case of an airfoil that is channeled in the spanwise direction is considered, and the impact on key performance characteristics is assessed. Particular attention is focused on exploring the limits of channel depth and low-Reynolds number on performance and stall characteristics. To quantify the effect of increased skin friction due to premature transition caused by protruding or recessed instrumentation, two simplified, conservative scenarios are used to consider two potential sources of diaturbance: A) that leading edge alterations would cause linearly expanding areas (triangles) of turbulent flow on both surfaces of the wing upstream of the natural transition point, and B) that a channel or bump on the upper surface would trip turbulent flow across the whole upper surface upstream of the natural transition point. A potentially more important consideration than the skin friction drag increment is the change in overall airfoil performance due to the installation of instrumentation along most of the wingspan. To quantify this effect, 2D CFD simulations of the flow over a representative mid-span airfoil section were conducted in order to assess the change in lift and drag curves for the airfoil in the presence of disturbed flow due to the installed instrumentation. A discussion as to the impact on high-altitude and low-speed operation of this and similar aircraft is provided.
This paper investigates the formation of nitric oxide (NO) and carbon monoxide (CO) emissions as a function of nitrogen (N 2 ) or carbon dioxide (CO 2 ) diluent content, added to a premixed reacting jet in crossflow. Here, reaction characteristics of a rich methane-air jet injected into a lean vitiated crossflow were analyzed at an elevated pressure of 5 atm. The jet was pre-heated and enriched with 0%, 15%, 30%, and 50% mass fraction to quantify the effect on pollutant emissions. Simulated results of the full chemistry Star-CCM+CFD model were verified with data taken in the experimental high-pressure combustion facility, which provides pressure, temperature and velocity profiles, as well as line-of-sight chemiluminescence and exit emission measurements. The analysis revealed the significant influence of the diluent to delay axial combustion; this mixing delay increased the axial flame lift-off and controlled any thermal ignition limitations. Increasing the diluent content increased the timescale for the flame to stabilize, which allowed for greater entrainment of crossflow oxidizer into the axial jet stream, and led to decreased pollutant emissions. Hence, crossflow entrainment is a critical driving force at high diluent content. Local nitric oxide (NO) emission formation in the axial stage was predicted numerically, showing the correlation between diluent addition, axial heat release, and the formation of nitric oxide pollution. The high diluent levels leaned out the jet mixture and delayed axial combustion, while reducing prompt and thermal NOx by mitigating flame hotspots and minimizing the timescale that the products remain at high temperatures.
The paper describes some improvements to a method of solving the time-dependent Reynolds-averaged Navier-Stokes equations for two-dimensional compressible flows for describing the coupling between viscous and inviscid regions in transonic flows. Mesh patching is implemented at the sublayer-boundary layer interfaces and the boundary layer-inviscid flow interface. Mesh stretching is geometric and is small in regions of large gradients and large in regions of small gradients, thus preserving second-order accuracy. Mesh lines are arranged so as to accommodate MacCormack's rapid solver for hyperbolic systems and a wider class of turbulent transport models. Results of computations of transonic flows over lifting airfoils are compared with experimental data and other solutions.
Diverse kinds of three-dimensional regions of separation in laminar and turbulent boundary layers are discussed that exist on lifting aerodynamic configurations immersed in flows from subsonic to hypersonic speeds. In all cases of three dimensional flow separation, the assumption of continuous vector fields of skin-friction lines and external-flow streamlines, coupled with simple topology laws, provides a flow grammar whose elemental constituents are the singular points: nodes, foci, and saddles. Adopting these notions enables one to create sequences of plausible flow structures, to deduce mean flow characteristics, expose flow mechanisms, and to aid theory and experiment where lack of resolution in numerical calculations or wind tunnel observation causes imprecision in diagnosing the three dimensional flow features.
A wind-tunnel investigation was made of the low-speed characteristics of a canard configuration having triangular wing and canard surfaces with an aspect ratio of 2. The exposed area of the canard was 6.9 percent of the total wing area. The canard hinge line was located at 0.35 of its mean aerodynamic chord and was 0.5 wing mean aerodynamic chord lengths forward of the wing apex. The ground effects, which made the lift more positive and the -Pitching moment more negative at a given angle of attack, were unaffected by the canard. The stability of the model at a constant canard hinge-moment coefficient decreased to 0 near a lift coefficient of 1.0. In addition, the maximum lift coefficient at which the canard could provide balance was decreased by ground effects to less than 1.0 if the moment center was as far forward as 0.21 of the wing mean aerodynamic chord. The relative magnitude of interference effects between the canard and the wing and body is presented.
Results of an investigation, conducted on the Langley helicopter test tower, of a rotor having an NACA 632A015 airfoil thickness distribution in combination with an NACA 230 mean line are presented. Comparison with a previously reported test of a symmetrical rotor blade efficiency was substantially improved over a wide range of tip Mach numbers. The maximum mean lift coefficient was essentially unchanged from that obtained with uncambered blades. Some data showing the effect of a distributed type of leading-edge roughness are also included.