Showing papers on "Drag divergence Mach number published in 2017"

Substantial drag reduction in turbulent flow using liquid-infused surfaces

Abstract: Experiments are presented that demonstrate how liquid-infused surfaces can reduce turbulent drag significantly in Taylor-Couette flow. The test liquid was water, and the test surface was composed of square microscopic grooves measuring 100 $\mu$m to 800 $\mu$m, filled with alkane liquids with viscosities from 0.3 to 1.4 times that of water. We achieve drag reduction exceeding 35\%, four times higher than previously reported for liquid-infused surfaces in turbulent flow. The level of drag reduction increased with viscosity ratio, groove width, fluid area fraction, and Reynolds number. The optimum groove width was given by $w^+ \approx 35$.

Aerodynamic drag of a transiting sphere by large-scale tomographic-PIV

Abstract: A method is introduced to measure the aerodynamic drag of moving objects such as ground vehicles or athletes in speed sports. Experiments are conducted as proof-of-concept that yield the aerodynamic drag of a sphere towed through a square duct in stagnant air. The drag force is evaluated using large-scale tomographic PIV and invoking the time-average momentum equation within a control volume in a frame of reference moving with the object. The sphere with 0.1 m diameter moves at a velocity of 1.45 m/s, corresponding to a Reynolds number of 10,000. The measurements in the wake of the sphere are conducted at a rate of 500 Hz within a thin volume of approximately 3 × 40 × 40 cubic centimeters. Neutrally buoyant helium-filled soap bubbles are used as flow tracers. The terms composing the drag are related to the flow momentum, the pressure and the velocity fluctuations and they are separately evaluated. The momentum and pressure terms dominate the momentum budget in the near wake up to 1.3 diameters downstream of the model. The pressure term decays rapidly and vanishes within 5 diameters. The term due to velocity fluctuations contributes up to 10% to the drag. The measurements yield a relatively constant value of the drag coefficient starting from 2 diameters downstream of the sphere. At 7 diameters the measurement interval terminates due to the finite length of the duct. Error sources that need to be accounted for are the sphere support wake and blockage effects. The above findings can provide practical criteria for the drag evaluation of generic bluff objects with this measurement technique.

Aircraft Lift and Drag Decomposition in Transonic Flows

Abstract: A vorticity-based exact theory for the analysis of the aerodynamic force is here applied to three-dimensional aircraft configurations in steady transonic flow by postprocessing numerical solutions....

Hypersonic Blunt-Body Shape Optimization for Reducing Drag and Heat Transfer

Abstract: A large design concern for high-speed vehicles such as next-generation launch vehicles or reusable spacecraft is the drag and heat transfer experienced at hypersonic velocities. In this paper, the optimized shapes for both minimum drag and minimum peak heat flux for an axisymmetric blunt body are developed using computational-fluid-dynamics software in conjunction with a genetic algorithm. For flowfield calculations, the commercial flow solver ANSYS Fluent is employed to solve the unsteady compressible Reynolds-averaged Navier–Stokes equations in conjunction with the shear-stress transport k-ω turbulence model. The hypersonic body shape is optimized using a multi-objective genetic algorithm to minimize both the drag and heat transfer. The multi-objective genetic algorithm creates a Pareto-optimal front containing the optimized shapes for various relative objectives of minimized drag and heat transfer. The results show a significant decrease in both the drag and peak heat flux and exhibit the expected chan...

Navier slip model of drag reduction by Leidenfrost vapor layers

Abstract: Recent experiments found that a hot solid sphere that is able to sustain a stable Leidenfrost vapor layer in a liquid exhibits significant drag reduction during free fall. The variation of the drag coefficient with Reynolds number deviates substantially from the characteristic drag crisis behavior at high Reynolds numbers. Measurements based on liquids of different viscosities show that the onset of the drag crisis depends on the viscosity ratio of the vapor to the liquid. Here we attempt to characterize the complexity of the Leidenfrost vapor layer with respect to its variable thickness and possible vapor circulation within, in terms of the Navier slip model that is defined by a slip length. Such a model can facilitate tangential flow and thereby alter the behavior of the boundary layer. Direct numerical and large eddy simulations of flow past a sphere at moderate to high Reynolds numbers (102≤Re≤4×104) are employed to quantify comparisons with experimental results, including the drag coefficient and the...

Assessment of the Aerodisk Size on Drag Reduction and Thermal Protection of High-Bluntness Vehicles at Hypersonic Speeds

Abstract: In this paper, numerical simulations are performed to investigate the influence of aerodisk size on the drag reduction and thermal protection of highly blunted bodies flying at hypersonic s...

Comparison of the Linear and Spanwise-Segmented DBD Plasma Actuators on Flow Control Around a NACA0015 Airfoil

Abstract: In this paper, linear and spanwise-segmented plasma actuator implementations as well as base airfoil with no plasma cases are presented. These approaches are used to control the flow around a NACA0015 airfoil. The experiments were conducted in a wind tunnel at Reynolds number of $3.6 \times 10^{4}$ . The plasma actuators mounted on the leading edge of the airfoil at chord position of 0.1 ( $x/C$ ). The electrical parameters used for the plasma generating device are set to constant values of 6-kVpp applied voltage and 3.5-kHz excitation frequency. It is observed that the use of spanwise-segmented plasma actuators converts the 2-D flow structure around the airfoil into 3-D forcing flow structure. The change of the wake region width of the airfoil is visualized by using the smoke-wire method. The flow visualization is performed at attack angles of 0°, 5°, 10°, and 15°. In addition, necessary measurements are also made to determine drag and lift forces. A comparative study on the drag and lift forces for the NACA0015 airfoil is performed. As a part of the conclusions, linear and spanwise-segmented plasma actuator implementations as well as base airfoil with no plasma cases are compared and related results on flow control are presented.

Comparative investigation on the aerodynamic effects of combined use of underbody drag reduction devices applied to real sedan

Abstract: To reduce the aerodynamic drag, the performance of the underbody aerodynamic drag reduction devices was evaluated based on the actual shape of a sedan-type vehicle. An undercover, under-fin, and side air dam were used as the underbody aerodynamic drag reduction devices. In addition, the effects of the interactions based on the combination of the aerodynamic drag reduction devices were investigated. A commercial sedan-type vehicle was selected as a reference model and its shape was modeled in detail. Aerodynamic drag was analyzed by computational fluid dynamics at a general driving speed on highway of 120 km/h. The undercover reduced the slipstream area through the attenuation of the longitudinal vortex pair by enhancing the up-wash of underflow, thereby reducing the aerodynamic drag by 8.4 %. The under-fin and side air dam showed no reduction in aerodynamic drag when they were solely attached to the actual complex shape of the underbody. Simple aggregation of the effects of aerodynamic drag reduction by the individual device did not provide the accurate performance of the combined aerodynamic drag reduction devices. An additional aerodynamic drag reduction of 2.1 % on average was obtained compared to the expected drag reduction, which was due to the synergy effect of the combination.

Mach Number Effect on Supersonic Drag Reduction using Repetitive Laser Energy Depositions over a Blunt Body

Abstract: Supersonic drag reduction performance using repetitive pulse energy depositions over blunt bodies was experimentally studied under twoMach numbers. The normalized drag reduction and energy deposition efficiency of Mach-1.92 over a 10-mm-dia. blunt-cylinder model were 8% and 1.2 at most, respectively. On the other hand, these values at Mach-3.20 over the same model were 22% and 6.2, respectively. The shock-wave deformation period using single-pulse energy deposition at Mach-3.20 was 64 ®s. This duration was shorter than that of 80 ®s at Mach-1.92, but the deformation magnitude on the model center axis of 40% at Mach-3.20 was larger than that of 15% at Mach-1.92. These experimental characteristics were consistent as solutions of the Riemann problem. Moreover, a drag reduction performance was much improved with a larger model diameter of 20 mm. Therefore, it has been experimentally demonstrated that the drag reduction performance due to energy deposition improves much at a high Mach number and with large model dimensions.

Impact of frontal-porous-insert temperature on the cylinder drag in supersonic flow

Abstract: In the present paper, we report on experimental and simulated data on the impact of cellular-porous-material heating on the wave drag of a cylinder with a frontal gas-permeable porous insert streamlined by a supersonic flow (М∞ = 4.85, Re1∞ = 3.3×106 m-1). Weighing data obtained in the supersonic wind tunnel are compared with data simulated using a discrete model of the cellular-porous material. An increase of the wave drag with a growth of porous-insert temperature and its decrease occurring upon decreasing the temperature are demonstrated.

Using Surface Sensitivity from Mesh Adjoint for Transonic Wing Drag Reduction

Abstract: The placement of a flow control device is highly dependent on the designers’ experience and their view of the area where the device will be most effective. In this paper, the mesh adjoint approach is proposed to help the designer to identify the regions where drag is most sensitive to a change of the surface of a transonic wing and to assess the success of an optimization. An array of shock control bumps are deployed in the areas of high sensitivity and optimized using a gradient-based approach. In addition to the sensitivity in the shock regions, a nonshock region is also identified using the sensitivity map on the wing. This region is not apparent from surface flow properties, such as pressure or skin friction, and could be overlooked by a designer without the sensitivity map. The results show that the mesh adjoint approach successfully identifies drag sensitive areas on the wing and assists in the deployment of the bump arrays. The bumps are parameterized using class/shape function transformation, whic...

Experimental Investigation on the Effectiveness of Active Control Mechanism on Base Pressure at Low Supersonic Mach Numbers

Abstract: In the current investigation, the experiments were carried out to evaluate the effectiveness of microjets in controlling the base pressure from a convergent-divergent nozzle at low supersonic Mach at different expansion level. Tests were carried out for low supersonic Mach numbers 1.25, 1.3, 1.48, and 1.6 while nozzle pressure ratio ranges from 3 to 11. The jets are augmented abruptly into an axisymmetric circular channel with different cross-sectional areas as that of nozzle exit area. The results show that the proficiency of the microjets is only marginal in controlling the base pressure even under the influence of favorable pressure gradient at lower NPRs namely 3 and 5. It was also observed that for higher values of the NPRs such as 7, 9, and 11, the dynamic control by very small jets results in rise of base pressure for the different values of the L/D ratios of these investigations. For NPRs 5 and 7, the trend differs due to the level of expansion, nature of waves present in the base region, relief available to the flow, length to diameter ratio of the enlarged duct, and the Mach numbers. It is seen that most of the cases exhibit similar behavior for higher as well as the lower length to diameter ratios, which means; that the back pressure has not adversely influenced the flow field in the base region as well as in the duct. With this it can be stated that the microjets can be an alternative for the experimentalist for base pressure control in the form of microjets.

An experimental study of the vortex wake at Mach number of 3

Abstract: The results of experimental study of the flow in the wing wake at Mach number of 3 are presented. These experiments extends the data obtained in the same experimental setup at Mach numbers of 2.5 and 4 [1]. Experiments were carried out in supersonic wind tunnel T–325 of ITAM SB RAS. Rectangular half-wing with sharp edges with a chord length of 30 mm and semispan of 95 mm was used to generate vortex wake. Experimental data were obtained in two cross sections located 1.5 and 6 chord length downstream of the trailing edge at wing angle of attack of 10 degrees. Constant temperature hot-wire anemometer was used to measure disturbances in supersonic flow. Hot-wire aemometer was made of a tungsten wire with a diameter of 10 µm and length of 1.5 mm. Shlieren flow visualization were performed. As a result, the position and size of the vortex core in the wake of a rectangular wing were determined. For the first time mass flow distribution and its pulsations in the supersonic longitudinal vortex was measured at Mach number of 3.

Transonic Drag Reduction Through Trailing-Edge Blowing on the FAST-MAC Circulation Control Model

Abstract: A third wind tunnel test of the FAST-MAC circulation control semi-span model was completed in the National Transonic Facility at the NASA Langley Research Center where the model was configured for transonic testing of the cruise configuration with 0deg flap detection to determine the potential for transonic drag reduction with the circulation control blowing. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged ap. Recent upgrades to transonic semi-span flow control testing at the NTF have demonstrated an improvement to overall data repeatability, particularly for the drag measurement, that allows for increased confidence in the data results. The static thrust generated by the blowing slot was removed from the wind-on data using force and moment balance data from wind-o thrust tares. This paper discusses the impact of the trailing-edge blowing to the transonic aerodynamics of the FAST-MAC model in the cruise configuration, where at flight Reynolds numbers, the thrust-removed corrected data showed that an overall drag reduction and increased aerodynamic efficiency was realized as a consequence of the blowing.

Aerodynamic characteristics of breathing blunt nose configuration at hypersonic speeds

Abstract: Breathing blunt nose technique is one of the promising methods for reducing the drag of blunt-nosed body at hypersonic speeds. The air, traversed by the bow shock positioned ahead of the nose, at the stagnation region is allowed to enter through a hole at the blunt-nose and ejected at the rear part (base region) of the body. This manipulation reduces the positive pressure over the stagnation regions of the nose and increases the pressure at the base, resulting in reduced suction at the base. The simultaneous manifestation of reducing the compression at the nose and suction at the base regions results in reduction of the total drag. The drag reduction caused by the breathing blunt nose technique has been measured in a Mach 7 tunnel. Also, the drag and flow field around the blunt-nosed body, with and without breathing hole, has been computed. The aerodynamic characteristics of the breathing blunt nose model obtained experimentally are compared with the CFD results. It is found that the breathing results in ...

Aerodynamic Analysis Over Double Wedge Airfoil

Abstract: Aeronautical studies are being focused more towards supersonic flights and methods to attain a better and safer flight with highest possible performance Aerodynamic analysis is part of the whole procedure, which includes focusing on airfoil shapes which will permit sustained flight of aircraft at these speeds Airfoil shapes differ based on the applications, hence the airfoil shapes considered for supersonic speeds are different from the ones considered for Subsonic The present work is based on the effects of change in physical parameter for the Double wedge airfoil Mach number range taken is for transonic and supersonic Physical parameters considered for the Double wedge case with wedge angle (ranging from 5 degree to 15 degree Available Computational tools are utilized for analysis Double wedge airfoil is analysed at different Angles of attack (AOA) based on the wedge angle Analysis is carried out using fluent at standard conditions with specific heat ratio taken as 14 Manual calculations for oblique shock properties are calculated with the help of Microsoft excel MATLAB is used to form a code for obtaining shock angle with Mach number and wedge angle at the given parameters Results obtained from manual calculations and fluent analysis are cross checked

The calculation of drag on nano-cylinders

Abstract: Summary The study of molecular flows at low Knudsen numbers (∼0.1–0.5), over nano-scaled objects of 20–100 nm size is becoming an important area of research. The simulation of fluid–structure interaction at nano-scale is important for understanding the adsorption and drag resistance characteristics of nano-devices in the fields of drug delivery, surface cleaning and protein movement. A novel formulation has been proposed that calculates localised values for both the kinetic and configurational parts of the Irving–Kirkwood stress tensor at given fixed positions within the computational domain. Macroscopic properties, such as streaming velocity, pressure and drag coefficients, are predicted by modelling the fluid–structure interaction using a moving least-squares method. The gravitation-driven molecular flow is examined over three different cross-sectional shapes—i.e. diamond-, circular- and square-shaped cylinders—confined within parallel walls and has been simulated for rough and smooth surfaces. The molecular dynamics formulation has allowed, for the first time, the calculation of localised drag forces over nano-cylinders. The computational simulation has shown that existing methods, including continuum-based approaches, significantly underestimate drag coefficients over nano-cylinders. The proposed molecular dynamics formulation has been verified on simulation based tests, as experimental and analytical results are unavailable at this scale. Copyright © 2016 John Wiley & Sons, Ltd.

Suction and Blowing Flow Control on Airfoil for Drag Reduction in Subsonic Flow

Abstract: Lift force is produced from a pressure difference between the pressures acting in upper and lower surfaces. Therefore, flow becomes detached from the surface of the airfoil at separation point and form vortices. These vortices affect the aerodynamic performance of the airfoil in term of lift and drag coefficient. Therefore, this study is investigating the effect of suction and jet blowing in boundary layer separation control on NACA 0012 airfoil in a subsonic wind tunnel. The experiment examined both methods at the position of 25% of the chord-length of the airfoil at Reynolds number 1.2 × 105. The findings show that suction and jet blowing affect the aerodynamic performance of NACA 0012 airfoil and can be an effective means for boundary layer separation control in subsonic flow.

Numerical Study on Mechanism of Drag Reduction by Microblowing Technique on Supercritical Airfoil

Abstract: Numerical studies on the applications of the microblowing technique (MBT) on a supercritical airfoil are performed based on a microporous wall model (MPWM) to represent the macroscaled collective characteristics of the huge number of microjets. The influences on the aerodynamic characteristics by microblowing with a MBT zone on different locations are analyzed. It is found that a MBT zone near the leading edge of the airfoil could achieve more reduction of skin-friction drag than a zone near the trailing edge. While for pressure drag, microblowing does not always result in a pressure drag penalty but could even reduce the pressure drag if the MBT porous zone is arranged on the region near the trailing edge. For the flow field without a shock wave, the MBT zone should be arranged on the lower wall and near the trailing edge. The typical configuration followed this guideline could simultaneously decrease the pressure drag and skin-friction drag while also increasing the lift. Numerical results indic...

Airfoil Theory at Supersonic Speed

Abstract: A theory is developed for the airfoil of finite span at supersonic speed analogous to the Prandtl airfoil theory of 1918-1919 for incompressible flow In addition to the profile and induced drags, account must be taken at supersonic flow of still another drag, namely, the wave drag, which is independent of the wing aspect ratio Both wave and induced drags are proportional to the square of the lift and depend on the Mach number, that is, the ratio of flight to sound speed In general, in the case of supersonic flow, the drag-lift ratio is considerably less favorable than is the case for incompressible flow Among others the following examples are considered: 1) lifting line with constant lift distribution (horseshoe vortex); 2) computation of wave and induced drag and the twist of a trapezoidal wing of constant lift density; 3) computation of the lift distribution and drag of an untwisted rectangular wing