Showing papers on "Drag divergence Mach number published in 2014"

Summary of the Fourth AIAA Computational Fluid Dynamics Drag Prediction Workshop

Abstract: Results from the Fourth AIAA Drag Prediction Workshop are summarized. The workshop focused on the prediction of both absolute and differential drag levels for wing–body and wing–body/horizontal-tail configurations of the NASA Common Research Model, which is representative of transonic transport aircraft. Numerical calculations are performed using industry-relevant test cases that include lift-specific flight conditions, trimmed drag polars, downwash variations, drag rises, and Reynolds-number effects. Drag, lift, and pitching moment predictions from numerous Reynolds-averaged Navier–Stokes computational fluid dynamics methods are presented. Solutions are performed on structured, unstructured, and hybrid grid systems. The structured-grid sets include point-matched multiblock meshes and overset grid systems. The unstructured and hybrid grid sets comprise tetrahedral, pyramid, prismatic, and hexahedral elements. Effort is made to provide a high-quality and parametrically consistent family of grids for each g...

Influence of turbulence on the drag of solid discs and turbine simulators in a water current

Abstract: Laboratory experiments have been used to investigate the effects of turbulence on the drag of both solid discs and porous disc turbine simulators. These discs were introduced to turbulent flows, in a gravity-fed water flume, with various levels of turbulence intensity and integral length scales. The turbulence was generated using three different grid configurations, which produced intensities and scales comparable with previous wind tunnel studies. The drag measurements were taken with discs of two different diameters and porosities with and without the upstream grids. The experimental results have demonstrated that the drag coefficients, of all the discs tested, are significantly dependent on both the turbulence intensity and integral length scale. For small integral length scales, relative to the disc, the drag coefficients converged for turbulence intensities greater than 13 %, with an increase of around 20 % in drag coefficient over the low-intensity case. Experiments with turbulence intensities of 10 % demonstrated minimum drag coefficients when the integral length scale-to-disc diameter ratio was around 50 %. Significant variations in the drag coefficient of circular bluff bodies are therefore expected when operating in turbulent flows with different characteristics.

Reduced-Order Modeling of Unsteady Aerodynamics Across Multiple Mach Regimes

Abstract: A reduced-order model for unsteady aerodynamic calculations across a range of Mach regimes based on linear convolution and a nonlinear correction factor is developed. Separate investigations are conducted for the sub-, trans-, and supersonic Mach regimes, and overall good results are seen when reduced-order model results are compared with full-order computational-fluid-dynamics solutions, though the reduced-order model errors tend to decrease as the Mach number increases. To assist reduced-order model construction, the method-of-segments simplified model has been created and tested throughout these same Mach regimes. Finally, a practical example of the reduced-order model’s applicability is presented by following a single test case from subsonic up through supersonic flight.

Computational Fluid Dynamics Study of the Effect of Leg Position on Cyclist Aerodynamic Drag

Abstract: An experimental and numerical analysis of cycling aerodynamics is presented The cyclist is modeled experimentally by a mannequin at static crank angle; numerically, the cyclist is modeled using a computer aided design (CAD) reproduction of the geometry Wind tunnel observation of the flow reveals a large variation of drag force and associated downstream flow structure with crank angle; at a crank angle of 15 deg, where the two thighs of the rider are aligned, a minimum in drag is observed At a crank angle of 75 deg, where one leg is at full extension and the other is raised close to the torso, a maximum in drag is observed Simulation of the flow using computational fluid dynamics (CFD) reproduces the observed variation of drag with crank angle, but underpredicts the experimental drag measurements by approximately 15%, probably at least partially due to simplification of the geometry of the cyclist and bicycle Inspection of the wake flow for the two sets of results reveals a good match in the downstream flow structure Numerical simulation also reveals the transient nature of the entire flow field in greater detail In particular, it shows how the flow separates from the body of the cyclist, which can be related to changes in the overall drag

Investigation of aerodynamic characteristics of rarefied flow around NACA 0012 airfoil using DSMC and NS solvers

Abstract: In this study, rarefied supersonic and subsonic gas flow around a NACA 0012 airfoil is simulated using both continuum and particle approaches. Navier–Stokes equations subject to the first order slip/jump boundary conditions are solved under the framework of OpenFOAM package. The DSMC solver of the package, i.e., dsmcFoam, has been improved to include a newly presented “simplified Bernoulli trial (SBT)” scheme for inter-molecular collision modeling. The use of SBT collision model permits to obtain accurate results using a much lower number of simulator particles. We considered flow at different angles of attacks and Knudsen numbers at both the subsonic and supersonic regimes. The computed density and surface pressure distributions are compared with the experimental and numerical data and suitable accuracy was observed. We investigate variations of the lift and drag coefficients with the Knudsen number and angle of attack. At low Kn number in supersonic regime, our results for lift coefficient agree well with the linearized theory; however, the deviation starts as soon as the angle of attack goes beyond 15 ° or shock wave forms above the airfoil. Along with this, we have observed that drag coefficient increases with the Kn number increasing. We also investigated the effect of Kn number on the leading edge shock position and structure, drag polar ( C L / C D ) , and slip velocity over the airfoil.

Accuracy of high-order density-based compressible methods in low Mach vortical flows

Abstract: SUMMARY At low Mach numbers, Godunov-type approaches, based on the method of lines, suffer from an accuracy problem. This paper shows the importance of using the low Mach number correction in Godunov-type methods for simulations involving low Mach numbers by utilising a new, well-posed, two-dimensional, two-mode Kelvin–Helmholtz test case. Four independent codes have been used, enabling the examination of several numerical schemes. The second-order and fifth-order accurate Godunov-type methods show that the vortex-pairing process can be captured on a low resolution with the low Mach number correction applied down to 0.002. The results are compared without the low Mach number correction and also three other methods, a Lagrange-remap method, a fifth-order accurate in space and time finite difference type method based on the wave propagation algorithm, and fifth-order spatial and third-order temporal accurate finite volume Monotone Upwind Scheme for Conservation Laws (MUSCL) approach based on the Godunov method and Simple Low Dissipation Advection Upstream Splitting Method (SLAU) numerical flux with low Mach capture property. The ability of the compressible flow solver of the commercial software, ANSYS FLUENT, in solving low Mach flows is also demonstrated for the two time-stepping methods provided in the compressible flow solver, implicit and explicit. Results demonstrate clearly that a low Mach correction is required for all algorithms except the Lagrange-remap approach, where dissipation is independent of Mach number. © 2013 Crown copyright. International Journal for Numerical Methods in Fluids. © 2013 John Wiley & Sons, Ltd.

Drag increase at the very start of drag reducing flows in a rotating cylindrical double gap device

Abstract: In this note we present some results concerning the time required for turbulent structures to achieve their steady state, called here the developing time. Notably, there is a drag increase at the very start of the test. Such a drag increase is strongly dependent on the concentration, molecular weight, temperature, Reynolds number, and molecule conformation before the test start-up. The analysis conducted here improves the understanding of the way drag reduction evolves over time, which was considered in Pereira et al. (2013).

Investigation of three-dimensional dynamic stall on an airfoil using fast-response pressure-sensitive paint

Abstract: Dynamic stall on a pitching OA209 airfoil in a wind tunnel is investigated at Mach 0.3 and 0.5 using high-speed pressure-sensitive paint (PSP) and pressure measurements. At Mach 0.3, the dynamic stall vortex was observed to propagate faster at the airfoil midline than at the wind-tunnel wall, resulting in a “bowed” vortex shape. At Mach 0.5, shock-induced stall was observed, with initial separation under the shock foot and subsequent expansion of the separated region upstream, downstream and along the breadth of the airfoil. No dynamic stall vortex could be observed at Mach 0.5. The investigation of flow control by blowing showed the potential advantages of PSP over pressure transducers for a complex three-dimensional flow.

A three-dimensional numerical investigation on drag reduction of a supersonic spherical body with an opposing jet

Abstract: A three-dimensional numerical simulation of a supersonic free-stream at Mach 2.5 over a spherical body with a sonic opposing jet from its stagnation point is carried out by solving the three-dimens...

Co-Flow Jet Airfoil Trade Study Part II : Moment and Drag

Abstract: This paper is Part II of a parametric study on CFJ airfoils. In the first part of the paper, the CFJ airfoil suction surface shape is modified to reduce or overcome the nose-down moment. In the second part of the paper, the injection and suction sizes and Cµ are varied to increase the CFJ airfoil thrust generation. For both parts, the resulting effects on the lift, drag, moment and energy consumption is analyzed. The two dimensional flow is simulated using steady and unsteady Reynolds Average Navier-Stokes (RANS). A 5th order WENO scheme for the inviscid flux, a 4th order central differencing model for the viscous terms and the one equation SpalartAllmaras model for the turbulence are used to resolve the flow. The Mach number is 0.15 and Reynolds number is 6.4 × 10 6 . The nose-down moment of the CFJ airfoils was successfully reduced with the use of reflex camber while negative drag was achieved with a thinner airfoil, and a reduced injection size. Increasing Cµ further reduces the drag, but at the cost of a much higher energy consumption and reduced corrected aerodynamic efficiency. The minimum drag achieved isCD = 0.033 and the highest moment achieved is CM = 0.031.

Numerical characterization of high-temperature filament interaction with blunt cylinder at Mach 3

Abstract: Energy deposition is a robust technique for various high-speed flow control applications including drag reduction. A numerical study of energy deposition is performed for perfect gas flow approaching a blunt cylinder at Mach 3. The energy deposition is simulated by a high-temperature filament injected instantaneously in front of the cylinder. The effect of important dimensionless parameters is studied to characterize the drag modification. The results indicate a saturation effect on maximum drag reduction at higher magnitudes of energy deposition. The computations reveal that the discharge location of the filament does not significantly impact the drag. A phenomenological examination of the interaction is performed. The effectiveness and efficiency of the filament on drag reduction are investigated. A one-dimensional analytical approach is studied to describe the numerical results.

Numerical Investigation of Actively Cooled Structures in Hypersonic Flow

Abstract: This paper presents a numerical evaluation of active-cooling thermal-protection systems in hypersonic flows. The aerothermodynamic model used herein consists of 1) an aerodynamic model based on the Reynolds-averaged Navier–Stokes equations, 2) a thermal-diffusion finite-element model, and 3) a solution methodology that couples the thermal-diffusion and aerothermal components. Hypersonic validation cases are performed on blunt-body and flat-plate geometries. A double-wedge airfoil with a rounded leading edge is simulated at speeds from Mach 3 to Mach 8 at altitudes ranging from sea level to 45 km. Coupled aerodynamic–thermal analysis is performed at a speed of Mach 5 at an altitude of 45 km and at a speed of Mach 8 at an altitude of 25 km with several chordwise-position-dependent cooling distributions on the interior of the airfoil. Active cooling using a piecewise continuous cooling distribution results in sufficient temperature reduction but also results in significant chordwise temperature gradients.

Aerodynamic behaviour of NREL S826 airfoil at Re=100,000

Abstract: This paper presents wind tunnel measurements of the NREL S826 airfoil at Reynolds number Re = 100,000 for angles of attack in a range of -10° to 25° the corresponding Large Eddy Simulation (LES) for selected angles of attack. The measurements have been performed at the low speed wind tunnel located at Fluid Mechanics laboratory of the Technical University of Denmark (DTU). Lift coefficient is obtained from the forge gauge measurements while the drag is measured according to the integration of the wake profiles downstream of the airfoil. The pressure distribution is measured by a set of pressure taps on the airfoil surface. The lift and drag polars are obtained from the LES computations using DTU's inhouse CFD solver, EllipSys3D, and good agreement is found between the measurement and the simulations. At high angles of attack, the numerical computations tend to over-predict the lift coefficients, however, there is a better agreement between the drag measurements and computations. It is concluded that LES computations are able to capture the lift and drag polars as well as the pressure distribution around the airfoil with an acceptable accuracy.

Effect of the vortex dynamics on the drag coefficient of a square back Ahmed body: Application to the flow control

Abstract: A vortex generated behind a simplified vehicle induces a pressure force at the back wall that contributes to a significant part of the drag coefficient. This pressure force depends on two parameters: the distance of the vortex to the wall and its amplitude or its circulation. Therefore there are two ways to reduce the drag coefficient: pushing the vortices away from the wall and changing their amplitude or their dynamics. Both analytical studies and numerical simulations show that these two actions decrease the pressure force and consequently reduce the drag coefficient. The first action is achieved by an active control procedure using pulsed jets and the second action is achieved by a passive control procedure using porous layers that change the vortex shedding. The best drag coefficient reduction is obtained by coupling the two procedures.

Supersonic aerodynamic performance of truncated cones with repetitive laser pulse energy depositions

Abstract: We investigate the drag characteristics of truncated cones in Mach 1.94 flow with repetitive laser pulse energy depositions with a frequency of up to 80 kHz. The drag decrement is almost in proportion to the laser pulse repetition frequency, and scales with a greater-than-square power of the truncation diameter. The performance of the latter is associated with the effective area of pressure modulation and the effective residence time of vortices which are baroclinically generated after the interaction between laser-heated gas bubbles and the bow shock wave. With employing a concave head, the drag decrement is enhanced. With increasing the truncation diameter, the efficiency of energy deposition becomes higher; yet, within the operation range of this study the drag coefficient still remains high.

Extended Formation Flight at Transonic Speeds

Abstract: Aircraft flown in formation can realize significant reductions in drag by flying in regions of wake upwash. However, most transports fly at transonic speeds where the impact of compressibility on formation flight is not well understood. This study uses an Euler solver to analyze the inviscid aerodynamic forces and moments of transonic wing/body configurations flying in a two-aircraft formation. Formations with large streamwise separation distances (10–50 wingspans) are considered. This work indicates that compressibility-related drag penalties in formation flight may be eliminated by slowing 2–3% below the nominal out-of-formation cruise Mach number, either at fixed lift coefficient or fixed altitude. The latter option has the additional benefit that the aerodynamic performance of the formation improves slightly at higher lift coefficients. Although optimal in-formation lift coefficients are not as high as those estimated by incompressible analyses, modest increases in altitude can yield further improveme...

Design and aerodynamic investigation of a parallel vehicle on a wide-speed range

Abstract: In order to design a hypersonic vehicle for a wide-ranged Mach number, a novel parallel vehicle for a wide-speed range has been proposed. In this paper, we employ a numerical method to investigate a parallel vehicle’s aerodynamic performance and flow field characteristics. The obtained results show that the aerodynamic performance of the novel parallel vehicle is better than that of the waverider designed with a single Mach number for the wide-speed range. With the increase in Mach number, the lift-to-drag ratio of the novel parallel vehicle first increases and then decreases. When the Mach number is 7 and the angle of attack is 3°, the lift-to-drag ratio is the largest, and its value is 3.968. When the angle of attack is 3°, the lift-to-drag ratio is not lower than 3.786 in the range considered in the current study, and the novel parallel vehicle’s aerodynamic performance is good. The wing changes the drag performance of the parallel vehicle remarkably, and results in the decrease of the lift-to-drag ratio. Meanwhile, the wing can enhance the pitching moment performance.

Far-Field Induced Drag Prediction Using Vorticity Confinement Technique

Abstract: The goal of this study is to show the efficiency of the wake-integral technique coupled with vorticity confinement method for the prediction of induced drag by inviscid numerical modeling. While the vorticity confinement method has previously been used to capture trailing vortices, its quantitative effect on improving wake-integral drag predictions has not been investigated. In the present study, parameters of vorticity confinement are tuned and the advantages of using a variable confinement parameter are demonstrated. A combination of vorticity confinement with the wake-integral technique indicates that the computational fluid dynamics-predicted induced drag for wing with and without winglet becomes insensitive to wake-integration plane location, the computed drag approaches its theoretical lifting-line value, and the spurious entropy drag is suppressed.

Fast Low-Fidelity Wing Aerodynamics Model for Surrogate-Based Shape Optimization

Abstract: Variable-fidelity optimization (VFO) can be efficient in terms of the computational cost when compared with traditional approaches, such as gradient-based methods with adjoint sensitivity information. In variable-fidelity methods, the direct optimization of the expensive high-fidelity model is replaced by iterative re-optimization of a physics-based surrogate model, which is constructed from a corrected low-fidelity model. The success of VFO is dependent on the reliability and accuracy of the low-fidelity model. In this paper, we present a way to develop a fast and reliable low-fidelity model suitable for aerodynamic shape of transonic wings. The low-fidelity model is component based and accounts for the zero-lift drag, induced drag, and wave drag. The induced drag can be calculated by a proper method, such lifting line theory or a panel method. The zero-lift drag and the wave drag can be calculated by two-dimensional flow model and strip theory. Sweep effects are accounted for by simple sweep theory. The approach is illustrated by a numerical example where the induced drag is calculated by a vortex lattice method, and the zero-lift drag and wave drag are calculated by MSES (a viscous-inviscid method). The low-fidelity model is roughly 320 times faster than a high-fidelity computational fluid dynamics models which solves the Reynolds-averaged Navier-Stokes equations and the Spalart-Allmaras turbulence model. The responses of the high-and low-fidelity models compare favorably and, most importantly, show the same trends with respect to changes in the operational conditions (Mach number, angle of attack) and the geometry (the airfoil shapes).

Aerodynamic performance of an over-the-wing propeller configuration at increasing Mach number

Abstract: Over-the-wing propeller configurations and particularly channel wing concepts show increased climb performance, and through effective acoustic shielding, reduced noise emissions when compared to a conventional tractor configuration. The main aerodynamic mechanisms could be identified by steady flow simulations of a simplified wing geometry and actuator disk. At take-off, where the thrust coefficient is very high, the drag of the wing decreases much stronger than the thrust of the propeller. This paper investigates the cruise conditions where the thrust coefficient is by one order of magnitude lower. The numerical results give evidence that, even at a moderate flight Mach number of 0.6, the beneficial influence of the over-the-wing propeller on the drag coefficient of the wing is negligibly small. On the other hand, the amount of propeller efficiency that is lost through high inflow velocity above the wing increases with Ma due to compressibility effects. As a result, the propulsive efficiency of an over-the-wing configuration is 16 % smaller than the reference (tractor). Semi-empirical correlations show that even at very low Mach numbers a drawback of at least 5 % remains. Although repositioning the propeller at the wing trailing edge may recover 4 % of the propulsive efficiency at Ma = 0.6, it is not advisable to give up most of the noise-shielding effect at take-off which is an important advantage of the channel wing.

Kinetic theory of drag on objects in nearly free molecular flow

Abstract: Using an analogy between the density expansion of the transport coefficients of moderately dense gases and the inverse-Knudsen-number expansion of the drag on objects in nearly free molecular flows, we formulate the collision integrals that determine the first correction term to the free-molecular drag limit. We then show how the procedure can be applied to calculate the drag coefficients of an oriented disc and a sphere as a function of the speed ratio.

Effect of Angle of Attack on Stability Derivatives of a Delta Wing with Straight Leading Edge in Supersonic Flow

Abstract: In the Present paper effect of angle of incidence on pitching derivatives of a delta wing with Straight leading edges for attached shock case in Supersonic Flow has been studied. A Strip theory is used in which strips at different span wise location are independent. This combines with similitude to give a piston theory. From the results it is seen that with the increase in the Mach number, there is a decrement of stiffness as well as the damping derivatives in pitch for all the Mach number tested, however, the magnitude of decrement for different inertia level will differ. It is seen that with the increase in the angle of attack both stiffness and damping derivatives increase linearly, nevertheless, this linear behavior limit themselves for different Mach numbers. For Mach number M = 2, this limiting value of validity is fifteen degrees, for Mach 2.5 & 3, it is twenty five degrees, whereas, for Mach 3.5 & 4 it becomes thirty five degrees. When these stability derivatives were considered at various pivot positions, namely h = 0.0, 0.2, 0.4, 0.6, 0.8, and 1.0. After scanning the results it was observed that with the shift of the pivot position from the leading edge to the trailing edge, the magnitude of both the stability derivatives were decreasing progressively with the pivot position. Results have been obtained for supersonic flow of perfect gases over a wide range of angle of attack and Mach number. The effect of real gas, leading edge bluntness of the wing, and secondary wave reflections are neglected.

Thrust Removal Scheme for the FAST-MAC Circulation Control Model Tested in the National Transonic Facility

Abstract: A second wind tunnel test of the FAST-MAC circulation control semi-span model was recently completed in the National Transonic Facility at the NASA Langley Research Center. 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 flap. The model was configured for transonic testing of the cruise configuration with 0deg flap deflection to determine the potential for drag reduction with the circulation control blowing. Encouraging results from analysis of wing surface pressures suggested that the circulation control blowing was effective in reducing the transonic drag on the configuration, however this could not be quantified until the thrust generated by the blowing slot was correctly removed from the force and moment balance data. This paper will present the thrust removal methodology used for the FAST-MAC circulation control model and describe the experimental measurements and techniques used to develop the methodology. A discussion on the impact to the force and moment data as a result of removing the thrust from the blowing slot will also be presented for the cruise configuration, where at some Mach and Reynolds number conditions, the thrust-removed corrected data showed that a drag reduction was realized as a consequence of the blowing.

Effect of angle of attack on stiffness derivative of an oscillating supersonic delta wing with curved leading edges

Abstract: In the Present paper effect of angle of incidence on Stiffness derivative of a delta wing with Curved leading edges for attached shock case in Supersonic Flow has been studied. A Strip theory is used in which strips at different span wise location are independent of each other. This combines with similitude to give a piston theory which gives closed form solutions for stiffness derivatives at low supersonic to high supersonic Mach numbers. From the results it is seen that with the increase in the Mach number, there is a continuous decrease in the magnitude of stiffness derivatives for all the Mach number tested, however, the magnitude of decrement for different inertia level will differ. It is seen that with the increase in the angle of attack the stiffness derivative increases linearly, nevertheless, this linear behavior limit themselves for different Mach numbers. For Mach number M = 2, this limiting value of validity is fifteen degrees, for Mach 2.5 & 3, it is twenty five degrees, whereas, for Mach 3.5 & 4 it becomes thirty five degrees, when these stability derivatives were considered at various pivot positions; namely at h = 0.0, 0.4, 0.6, and 1.0. After scanning the results it was observed that with the shift of the pivot position from the leading edge to the trailing edge, the magnitude of stiffness and the damping derivatives continue to decrease progressively. Results have been obtained for supersonic flow of perfect gases over a wide range of angle of attack and Mach number. The effect of real gas, leading edge bluntness of the wing, shock motion, and secondary wave reflections are neglected.

Pitching Airfoil Performance Enhancement Using Co-Flow Jet Flow Control at High Mach Number

Abstract: Pitching airfoils with Co-Flow Jet (CFJ) flow control are simulated using Unsteady Reynolds Average Navier-Stokes (URANS) at Mach number 0.4 with reduced frequency of 0.1. The flow is transonic with shock wave boundary layer interaction. A 5th order WENO scheme for the inviscid flux, a 4th order central differencing model for the viscous terms and the one equation SpalartAllmaras model for the turbulence are used to resolve the flow. The airfoil oscillate around its mean AoA of 10 ◦ with amplitude of 5 ◦ , 7.5 ◦ and 10 ◦ . The study demonstrates that the CFJ pitching airfoil is very effective to remove dynamic stall at high Mach number of 0.4. The performance is significantly enhanced with radically increased lift, reduced drag, and decreased moment variation.

Coefficient of Drag and Trajectory Simulation of 130 mm Supersonic Artillery Shell with Recovery Plug or Fuze

Abstract: In the present study, the drag variation and trajectory elements estimation of a supersonic projectile having two different nose shapes are made numerically. The study aims at finding the coefficient of drag and shock wave pattern for 130 mm artillery shell fitted with recovery plug or with fuze, when travelling at zero angle of attack in a supersonic flow of air. The coefficient of drag (CD) obtained from the simulation is used as an input parameter for estimation of trajectory elements. The numerical results, i.e., the coefficient of drag at different Mach numbers and trajectory elements are validated with the data recorded by tracking radar from an experimental firing. Based on numerical results and data recorded in experimental firing, the coefficient of drag in the case of the shell with recovery plug is 2.7 times more than for the shell with fuze. The shock wave in the case of the shell with recovery plug is detached bow shock wave, whereas in the case of a shell with fuze, the shock is attached. The results indicate that the coefficient of drag increases with detached shock wave and an increase in the radius of the shell nose. Good agreements were observed between numerical results and experimental observations. Defence Science Journal, Vol. 64, No. 6, November 2014, pp.502-508, DOI:http://dx.doi.org/10.14429/dsj.64.8110