Showing papers on "Drag divergence Mach number published in 2006"

A theoretical prediction of friction drag reduction in turbulent flow by superhydrophobic surfaces

Abstract: We present a theoretical prediction for the drag reduction rate achieved by superhydrophobic surfaces in a turbulent channel flow. The predicted drag reduction rate is in good agreement with results obtained from direct numerical simulations at Reτ≃180 and 400. The present theory suggests that large drag reduction is possible also at Reynolds numbers of practical interest (Reτ∼105–106) by employing a hydrophobic surface, which induces a slip length on the order of ten wall units or more.

Turbulent Drag Reduction Using Superhydrophobic Surfaces

Abstract: Superhydrophobic surfaces are known to exhibit reduced viscous drag due to "slip" associated with a layer of air trapped at the liquid-solid interface. It is expected that this slip will lead to reduced turbulent skin-friction drag in external flows at higher Reynolds numbers in both the laminar and turbulent regimes. Results are presented from experiments exploring this effect. Large-area Superhydrophobic test surfaces have been fabricated and tested in a water tunnel, measuring drag in both the laminar and transitional regimes at velocities up to 1.4 m/s. Drag reduction of approximately 50% is observed for laminar flow. Lower levels of drag reduction are observed at higher speeds after the flow has transitioned to turbulence.

Drag reduction of a bluff body using adaptive control methods

Abstract: A classical actuator is used to control the drag exerted on a bluff body at large Reynolds number (Re=20000). The geometry is similar to a backward-facing step whose separation point is modified using a rotating cylinder at the edge. The slow fluctuations of the total drag are directly measured by means of strain gauges. As shown by visualizations, the actuator delays the separation point. The size of the low-pressure region behind the body is decreased and the drag reduced. It is found that the faster the rotation of the cylinder, the lower the drag. In a first study, the goal of the control is for the system to reach a drag consign predetermined by the experimentalist. The control loop is closed with a proportional integral correction. This adaptive method is shown to be efficient and robust in spite of the large fluctuations of the drag. In the second method, the system finds itself its optimal set point. It is defined as the lowest cost of global energy consumption of the system (drag reduction versus energy used by the actuator). For this purpose, an extremum seeking control method is applied in order to deal with the large background noise due to turbulence. It consists in a synchronous detection of the response measured in the drag measurements to a modulation of the actuator. The phase shift and amplitude of the modulation estimate the local gradient of the total energy function. With this gradient estimation, the system goes to the minimum of global power consumption by itself. The system is found to be also robust and reacts successfully to changes of the external mean flow. This experiment attests to the real efficiency of local active control in reducing autonomously the global energy consumption of a system under turbulent flow.

Counterflow drag reduction by supersonic jet for a blunt body in hypersonic flow

Abstract: Counterflow drag reduction by supersonic jet for a large angle blunt cone at hypersonic Mach number is investigated in a shock tunnel. The flowfields around the test model in the hypersonic flow with an opposing supersonic jet emanating from the stagnation point of the model are visualized by high speed schlieren technique. The aerodynamic drag force is measured using the accelerometer based force balance system. The experimental measurements show about 30%–45% reduction in drag coefficient for different jet pressures.

Computational Study of Shock Mitigation and Drag Reduction by Pulsed Energy Lines

Abstract: A series of numerical experiments were performed in which energy was deposited ahead of a cone traveling at supersonic/hypersonic speeds. The angle of attack was zero, and the cone half-angles ranged from 15 to 45 deg. The Mach numbers simulated were 2, 4, 6, and 8. The energy was deposited instantaneously along a finite length of the cone axis, ahead of the cone’s bow shock, causing a cylindrical shock wave to push air outward from the line of deposition. The shock wave would sweep the air out from in front of the cone, leaving behind a low-density column/tube of air, through which the cone (vehicle) propagated with significantly reduced drag. The greatest drag reduction observed was 96%. (One-hundred percent drag reduction would result in the complete elimination of drag forces on the cone.) The propulsive gain was consistently positive, meaning that the energy saved as a result of drag reduction was consistently greater than the amount of energy “invested” (i.e., deposited ahead of the vehicle). The highest ratio of energy saved/energy invested was approximately 6500% (a 65-fold “return” on the invested energy). We explored this phenomenon with a high-order-accurate multidomain weighted essentially nonoscillatory finite difference algorithm, using interpolation at subdomain boundaries. This drag-reduction/shock-mitigation technique can be applied locally or globally to reduce the overall drag on a vehicle.

Numerical Simulations of a Feedback-Controlled Circular Cylinder Wake

Abstract: The effect of feedback flow control on the wake of a circular cylinder at a Reynolds number of 100 is investigated in direct numerical simulation. The control approach uses a low-dimensional model based on proper orthogonal decomposition (POD). The controller applies linear proportional and differential feedback to the estimate of the first POD mode. The range of validity of the POD model is explored in detail. Actuation is implemented as displacement of the cylinder normal to the flow. It is demonstrated that the threshold peak amplitude below which the control actuation ceases to be effective is in the order of 5% of the cylinder diameter. The closed-loop feedback simulations explore the effect of both fixed-phase and variable-phase feedback on the wake. Whereas fixed-phase feedback is effective in reducing drag and unsteady lift, it fails to stabilize this state once the low drag state has been reached. Variable-phase feedback, however, achieves the same drag and unsteady lift reductions while being able to stabilize the flow in the low drag state. In the low drag state, the near wake is entirely steady, whereas the far wake exhibits vortex shedding at a reduced intensity. A drag reduction of 15% of the drag was achieved, and the unsteady lift force was lowered by 90%.

Separate Effects of Mach Number and Reynolds Number on Jet Array Impingement Heat Transfer

Abstract: Limited available data suggest a substantial impact of Mach number on the heat transfer from an array of jets impinging on a surface at fixed Reynolds number Many jet array heat transfer correlations currently in use are based upon tests in which the jet Reynolds number was varied by varying the jet Mach number Hence, this data may be inaccurate for high Mach numbers Results from the present study are new and innovative because they separate the effects of jet Reynolds number and jet Mach number for the purposes of validating and improving correlations which are currently in use The present study provides new data on the separate effects of Reynolds number and Mach number for an array of impinging jets in the form of discharge coefficients, local and spatially-averaged Nusselt numbers, and local and spatially-averaged recovery factors The data are unique because data are given for impingement jet Mach numbers as high as 060 and impingement jet Reynolds numbers as high as 60,000, and because the effects of Reynolds number and Mach number are separated by providing data at constant Reynolds number as the Mach number is varied, and data at constant Mach number as the Reynolds number is varied As such, the present data are given for experimental conditions not previously examined, which are outside the range of applicability of current correlationsCopyright © 2006 by ASME

The unsteady drag force on a cylinder immersed in a dilute granular flow

Abstract: This paper presents results from hard-particle discrete element simulations of a two-dimensional dilute stream of particles accelerating past an immersed fixed cylinder. Simulation measurements of the drag force Fd are expressed in terms of a dimensionless drag coefficient, Cd=Fd∕[12ρνU2(D+d)], where ρ is the particle density, ν is the upstream solid fraction, U is the upstream instantaneous velocity, and D and d are the cylinder and particle diameters, respectively. Measurements indicate that the cylinder’s unsteady drag coefficient does not vary significantly from its steady (nonaccelerating) drag coefficient for both frictionless and frictional particles implying that the added mass for the flow is negligible. However, the drag coefficient is larger than its nominal value during an initial transient stage, during which a shock wave develops in front of the cylinder. Once the shock has developed, the drag coefficient remains constant despite the stream’s acceleration. The duration of the shock developme...

Adaptive Airfoils for Drag Reduction at Transonic Speeds

Abstract: Adaptive airfoils and wings can provide superior performance at the expense of increased cost and complexity. In this paper, an aerodynamic optimization algorithm is used to assess an adaptive airfoil concept for drag reduction at transonic speeds. The objective is to quantify both the improvements in drag that can be achieved and the magnitude of the shape changes needed. In an initial study, a baseline airfoil is designed to produce low drag at a fixed lift coefficient over a range of Mach numbers. This airfoil is compared with a sequence of nine airfoils, each designed to be optimal at a single operating point in the Mach number range. Shape changes of less than 2% chord lead to drag reductions of 4-6% over a range of Mach numbers from 0.68 to 0.76. If the shape changes are restricted to the upper surface only, then changes of less than 1% chord lead to drag reduction of 3-5%. In a second study, a baseline airfoil is designed based on a multi-point optimization over eighteen operating points, including dive and low-speed off-design requirements. Adaptive airfoils are designed through single-point optimization for the operating points corresponding to cruise conditions, producing drag reductions ranging from 9.7 to 16.7% with shape changes on the order of a few percent chord.

Wing Drag Prediction and Decomposition

Abstract: A wake integration methodology to predict the aerodynamic characteristics of three-dimensional wings in viscous subsonic and transonic flows is presented. Results indicate that wake integration is viable as a method of drag prediction and, compared to surface integration, has the added benefit of being able to provide a decomposition of the total drag into its physical components of profile, induced, and wave drag for a wing in viscous, transonic flow. Future work will involve using wake integration on postprocess flows about more complex configurations for which this study provides a sound basic framework.

Application of theoretical principles to swimsuit drag reduction

Abstract: This study investigated the basic fluid mechanics associated with the hydrodynamic drag of a human. The components of drag (frictionD SF, pressureD P and waveD W) on a human swimmer were analysed by applying classical fluid dynamic fundamentals. General methods of reducing drag were considered and the most probable method identified, applied and tested on swimsuit hydrodynamic drag. This study employed turbulators, either one (upper back) or three (across the upper back, the chest and across the buttocks), that were compared to an identical full body suit with no turbulators. Male and female elite competitive swimmers (n = 7 each) were towed in an annular pool to determine passive drag at speeds from 0.4 to 2.2 m s−1. The total drag was reduced by 11–12% by one turbulator and 13–16% by three turbulators. The total drag was decomposed intoD SF, DP andD W to determine the mechanisms responsible for the reduced total drag by the turbulators. The presence of the turbulators did not significantly increase friction or wave drag; however, flow was attached to the body as there was a significant reduction in pressure drag (19–41%), with the greatest reduction being for three turbulators (chest, back, buttocks). This study demonstrated the importance of pressure drag in determining total drag at high human swimming speeds, and that drag reducing technology can significantly reduce it, in this case by appropriately sized and placed turbulators.

Experimental Investigation Of Starting Process for a Variable Geometry Air Inlet operating from Mach 2 to Mach 8

Abstract: A lot of key issues have to be addressed for developing the dual mode ramjet technology. But the main difficulty is to design a propulsion stream tube, which allows in the same time the constraints for obtaining a good performance at high flight Mach number (supersonic combustion) and those for ensuring a stable operating at low flight Mach number (subsonic combustion). If the overall Mach number range is limited (Mach 4-Mach 8), a completely fixed geometry can be defined for the propulsion stream tube with relatively good possible performance in the whole flight envelope. But, even if this limited Mach number range could be sufficient for missiles application, it is clear that airbreathing propulsion can not have any interest for reusable space launcher application if it is not able to power the launcher from Mach 2- to Mach 10+. This extension of the Mach number range requires some variation of the propulsion stream tube geometry to obtain acceptable performances. Indeed for higher Mach numbers, it will be necessary to use a high contraction ratio for the inlet in order to limit the supersonic Mach number at the entrance of the combustion chamber and a less and less diverging combustion chamber as Mach number will increase. At contrary, in order to extend the flight envelope to lower Mach numbers, the inlet contraction ratio must be reduced (with or without reduction of the capture area) to avoid a too large cowl spillage and the corresponding additive drag section of the combustion chamber must be open. Several approaches can be considered to fulfil this required adaptation of the propulsion stream tube geometry. One of the simplest somutions consists in obtaining the combined variation of the inlet and the combustion chamber geometry by a simple translation of the inlet/combustion chamber cowl along an axis close to horizontal axis of the vehicle. Thanks to this translation, it is possible to adapt the contraction ratio to the flight conditions in the Mach number range from 2 to 8 and then to obtain good performance at higher Mach numbers without penalty for the low Mach number regime. Such an air inlet has been designed and tested in ITAM blown-down and hot-shot windtunnels in the Mach number range from 2 to 8. The modularity of the model and its extensive instrumentation allowed acquiring a very valuable data base contributing to the knowledge of key parameters determining inlet performance and building a performance model. As the considered air inlet is mainly an internal compression one, a particular effort has been made to understand the self-starting limits and provide some improvement by using a porous bleed located on the compression ramp. Moreover, the effect of windtunnel test section starting on the controlled starting of the air inlet has been evaluated by comparing results obtained in blow-down windtunnel (long starting of windtunnel) and in hotshot windtunnel (short starting) and by using a specific starting device in blow-down windtunnel to quickly open the air inlet only when the flow is fully established in the test section.

Drag of Spheroid-Cone Shaped Airship

Abstract: Drag coefficients of streamlined bodies and complete airships are compared to confirm some general trends. Data obtained from flight testing an electrically powered, helium-filled, dirigible balloon with a spheroid-cone hull form are analyzed. Wind-tunnel tests indicate that placing a cone behind a sphere reduces its drag coefficient by about 50%, but flight-test data suggest the drag coefficient of the spheroid-cone airship is relatively high. The influence of atmospheric turbulence and nonsteady flow effects are discussed.

Drag reduction in turbulent channel flow with periodically arrayed heating and cooling strips

Abstract: A new technique giving significant drag reduction in turbulent shear flows has been proposed by using the buoyancy effect to generate periodic spanwise motion. Such spanwise motion can be obtained by arranging heating and cooling strips periodically aligned in the spanwise direction of a vertical channel, where the streamwise mean flow is perpendicular to the gravity vector. The strip size has been changed in order to obtain the optimum size corresponding to the maximum drag reduction. Series of direct numerical simulation have been performed where the bulk Reynolds number, Rem=Umδ∕ν is fixed at 2270 whereas the Grashof number is changed to between 106 and 107. At the lowest Grashof number, the buoyancy forces are not strong enough to disturb the flow and no apparent variation of drag compared to the plane channel is observed. However, as the Grashof number increases, considerable drag reduction can be obtained. At the highest Grashof number, an optimum strip size of about 250 wall units gives drag reduct...

Drag reduction by compressible bubbles.

Abstract: Drag reduction by bubbles in stationary turbulent flows is sensitive to the compressibility of the bubbles. Without this dynamical effect the bubbles only renormalize the fluid density and viscosity, an effect that by itself can only lead to a small percentage of drag reduction. We show in this paper that the dynamics of bubbles and their effect on the compressibility of the mixture can lead to a much higher drag reduction.

Aerodynamic Design of Biplane Airfoils for Low Wave Drag Supersonic Flight

Abstract: In this paper, aerodynamic design of biplane airfoils in supersonic flight is discussed based on Computational Fluid Dynamics (CFD). In supersonic flight, airfoils generate strong sonic booms and wave drags accompanied by shock waves. We propose a significant reduction of them, especially wave drags using a biplane-airfoil concept. The background of this concept is originated from Busemann biplane and Licher type biplane concepts. In order to focus on the shock wave characteristics around biplane configuration, inviscid flow (Euler) analyses are performed (which is particularly suitable for wave drag analyses). Design Mach number is 1.7. The aerodynamic design is conducted using an iterative inverse design method that is newly implemented. A biplane configuration having a desired performance has been obtained. Having 0.102 of total maximum thickness ratio (t/c), it has the lift to wave drag ratio (L/D) of 21.7 at a desired lift condition for supersonic flight, lift coefficient (Cl)=0.115. At the range of lift coefficient more than 0.144 this designed biplane has lower wave drag than that of a (zero-thickness) single flat plate airfoil.

Aerodynamic Characteristics of Supercritical Outlet Guide Vanes at Low Reynolds Number Conditions

Abstract: As a part of an innovative aerodynamic design concept for a single stage low pressure turbine, a high turning outlet guide vane is required to remove the swirl from the hot gas. The airfoil of the vane is a highly loaded compressor airfoil that has to operate at very low Reynolds numbers (Re ∼ 120,000). Recently published numerical design studies and experimental analysis on alternatively designed airfoils showed that blade profiles with an extreme front loaded pressure distribution are advantageous for low Reynolds number conditions. The advantage even holds true for an increased inlet Mach number at which the peak Mach number on the airfoils reaches and exceeds the critical conditions (Mss > 1.0). This paper discusses the effect of the inlet Mach number and Reynolds number on the cascade performance for both a controlled diffusion airfoil (CDA) (called baseline) and a numerically optimized front loaded airfoil. The results show that it is advantageous to design the profile with a fairly steep pressure gradient immediately at the front part in order to promote early transition or to prevent too large laminar — even shock induced — separations with the risk of a bubble burst. Profile Mach number distributions and wake traverse data are presented for design and off-design conditions. The discussion of Mach number distributions and boundary layer behavior is supported by numerical results obtained from the blade-to-blade flow solver MISES.Copyright © 2006 by ASME

An airfoil for a helicoptor rotor blade

Abstract: An airfoil family for a helicopter rotor blade, designated SC362XX. SC362XX essentially removes the large lower surface suction peak associated with 'drag creep' at moderate lift coefficients while reducing the peak Mach number and shock strength at high lift/Mach number conditions. Another optional airfoil family for use at inboard regions of the helicopter rotor, which is designated SC3252XX airfoil family, is a relatively thicker airfoil section that includes a significant increase in thickness forward of the 30% x/c location to provide a relatively thick and rigid inboard section. The lift coefficient at which the drag divergence Mach number was optimized is the same in both families thereby readily providing application to a single rotor blade.

Mach 8 High Reynolds Number Static Stability Capability Extension Using a Hypersonic Waverider at AEDC Tunnel 9

Abstract: The recent increased need to collect accurate aerodynamic stability data at Mach 8 and high Reynolds numbers motivated the AEDC Tunnel 9 to develop a new Mach 8 facility. This upgrade modified an existing Mach 8 nozzle for use with the Mach 10/14 test cell and dynamically pitching model support hardware. The new capability allows investigators to collect an entire pitch polar during a single run at Mach 8 up to freestream unit Reynolds numbers of 47.7×10 6 /ft. In order to demonstrate this capability, a verification test entry acquiring force and moment, surface pressure, and heat-transfer data on a realistic hypersonic waverider configuration was conducted. Despite the challenges associated with fast pitch sweeps and extremely high aerodynamic loading, the test validates the feasibility of collecting accurate static stability data in the new facility. Additionally, independent numerical simulations were generated by the Air Force Research Laboratory at WrightPatterson AFB and the University of Minnesota in an effort to improve computation methods for high lift over drag hypersonic vehicles. The comparison of the experimental and computational dataset will be published at a later date.

Aerodynamic Shape Optimization Based on Drag Decomposition

Abstract: An advanced drag prediction method, mid-field drag decomposition method is applied in aerodynamic shape optimization problems. The drag decomposition method decomposes total drag into wave, profile, induced and spurious drag component, the latter resulting from the effect of numerical diffusion included in CFD results. Hence the more accurate drag prediction can be achieved by the elimination of the spurious drag component. This method is applied in transonic airfoil, planform and winglet shape optimizations. As the optimizer and flow solver, genetic algorithm and Euler/NS simulation are used, respectively. The results show that the optimizations based on the drag decomposition method are reliable and accurate. Moreover, precise investigation of the drag reduction mechanisms is achieved by using the drag decomposition method.

Nonlinear effects caused by pulsed periodic supply of energy in the vicinity of a symmetric airfoil in a transonic flow

Abstract: Changes in the structure of a transonic flow around a symmetric airfoil and a decrease in the wave drag of the latter, depending on the energy-supply period and on localization and shape of the energy-supply zone, are considered by means of the numerical solution of two-dimensional unsteady equations of gas dynamics. Energy addition to the gas ahead of the closing shock wave in an immediate vicinity of the contour in zones extended along the contour is found to significantly reduce the wave drag of the airfoil. The nature of this decrease in drag is clarified. The existence of a limiting frequency of energy supply is found.

The effects of the tip shape of V-groove on drag reduction and flow field characteristics by numerical analysis

Abstract: The flow in turbulent boundary layer and the viscous drag over V-groove surface are numerically simulated using the RANS formula and RNG k-e turbulence model.The effects of the tip shape of V-groove on the drag reduction,the velocity profile in the turbulent boundary layer and shear stress on groove surface are studied.It is shown that,the smaller the fillet radius of the tip is,the better the effect of drag reduction is.And the best ratio of drag reduction that can be attained is 6.6%.As the fillet radius of the tip becomes smaller,the wall shear stress on the middle and low part of the groove surface becomes less,but the local wall shear stress on the tip becomes greater.The secondary vortices generated at the groove peaks are the fundamental reasons of the drag reduction,and the variation of the drag reduction is also caused by the secondary vortices.

Optimization of mass bleed control for base drag reduction of supersonic flight bodies

Abstract: The minimization of base drag using mass bleed control is examined in consideration of various base to orifice exit area ratios for a body of revolution in the Mach 2.47 freestream. Axisymmetric, compressible, mass-averaged Navier-Stokes equations are solved using the standard k-θ turbulence model, a fully implicit finite volume scheme, and a second order upwind scheme. Base flow characteristics are explained regarding the base configuration as well as the injection parameter which is defined as the mass flow rate of bleed jet non-dimensionalized by the product of the base area and freestream mass flux. The results obtained through the present study show that for a smaller base area, the optimum mass bleed condition leading to minimum base drag occurs at relatively larger mass bleed, and a larger orifice exit can offer better drag control.

Constant Altitude-Constant Mach Number Cruise Range of Transport Aircraft with Compressibility Effects

Abstract: An approximate solution of the constant altitude‐constant Mach number cruise range for high subsonic speed flight of the turbojet/fan aircraft is proposed. The solution considers cambered wing drag polar of modern transport aircraft, dependence of the specific fuel consumption on Mach number, and compressibility effects on aerodynamic characteristics of the aircraft. The method aims for a quick assessment of the cruise range during conceptual or preliminary design phase. An application of the method to a known type of aircraft is also presented.

Nonlinear effects of pulsed periodic energy supply on shock wave structure of transonic flow around airfoils

Abstract: The work deals with an investigation of possibilities of controlling the aerodynamic characteristics of airfoils with the aid of external local pulsed-periodic energy supply at transonic flight regimes. The alteration of flow structure near a symmetric airfoil and its wave drag has been studied on the basis of a numerical solution of two-dimensional nonstationary gas dynamics equations versus the energy supply period, the localization and shape of the energy supply zone. The energy supply upstream of the closing shock in the immediate proximity of the contour in the zones extended along it is found to result in a considerable reduction of the profile wave drag. The nature of such a drag reduction is elucidated. The existence of a limiting frequency of energy supply is established.

Three-dimensional SBLI control for transonic airfoils

Abstract: The application of shock control to transonic airfoils and wings has been demonstrated widely to have the potential to reduce wave drag. Most of the suggested control devices are two-dimensional, that is they are of uniform geometry in spanwise direction. Examples of such techniques include contour bumps and passive control. Recently it has been observed that a spanwise array of discrete three-dimensional controls can have similar benefits but also offer advantages in terms of installation complexity and drag. This paper describes research carried out in Cambridge into various three-dimensional devices, such as slots, grooves and bumps. In all cases the control device is applied to the interaction of a normal shock wave (M=1.3) with a turbulent boundary layer. Theoretical considerations are proposed to determine how such fundamental experiments can provide estimates of control performance on a transonic wing. The potential of each class of three-dimensional device for wave drag reduction on airfoils is discussed and surface bumps in particular are identified as offering potential drag savings for typical transonic wing applications under cruise conditions.