Moderate Angles Of Attack Research Articles

Inviscid and RANS Simulation of the NASA-CRM Using the Cartesian Flow Solver UTCart

Inviscid and Reynolds-averaged Navier-Stokes (RANS) simulations of transonic flows around the NASA Common Research Model are conducted using the Cartesian flow solver UTCart. The immersed boundary method is used to represent the smooth geometry surfaces on the Cartesian grids. The wall function is combined with the immersed boundary method to reproduce the turbulent boundary layer on the geometry surface in the RANS simulations. In the inviscid calculations, the qualitative flow feature including the position on the shock-wave on the wing shows agreement with the reference result a body-fitted grid. In the RANS calculations, the trend of pitching moment and drag shows fair agreement with the reference result, while prediction of the flow separation at high angle of attack is still difficult. Compared with the reference result, the differences in the total drag coefficient at a moderate angle of attack on the medium grid (33 million cells) and the fine grid (99 million cells) are 31 drag counts (10%) and 20 drag counts (6.5%), respectively. Furthermore, each of the calculated aerodynamic coefficients shows a consistent trend of grid convergence toward the reference result.

RANS Simulations of NASA-CRM by the Finite-volume Method with the SST Turbulence Model

The paper reports numerical simulations of aerodynamic characteristics of a commercial aircraft model known as the NASA-CRM. Specific test cases and detailed experimental results have been provided by JAXA for the CFD workshop “Aerodynamic Prediction Challenge (APC).” We employed the cell-vertex finite-volume method on unstructured grids and carried out the RANS simulations with the k-ω SST turbulence model to present results in the APC-I. After the verification of the model in a benchmark problem, both alpha-sweep simulations from low- to high-angles of attack and the grid-convergence study at a cruise conditions are carried out. The simulations well reproduced the experimental results at moderate angles of attack by both types of computational grids used in the study but the results showed the dependency on the grid type and grid resolutions at the highest angle of attack of 5.72 degrees due to different predictions of flow separation near the root chord of the main wing.

Identification of flow phenomena over NACA 4412 wind turbine airfoil at low Reynolds numbers and role of laminar separation bubble on flow evolution

Objective of this study is to elucidate relationship with aerodynamic properties and vortex shedding from suction surface and wake of wind turbine blade at low Reynolds numbers. Force measurement of NACA 4412 airfoil was conducted at various angles of attack and Reynolds numbers. Furthermore, smoke-wire experiment was performed to clearly visualize flow patterns such as flow separation or laminar separation bubble (LSB) over the airfoil. Also, velocity measurements at near wake region were done to highlight coactions among LSB and trailing edge vortices. Hot-film experiment was performed to obtain more information with regards to progress and formation of LSB. Experiments indicated that location and formation of LSB were affected by variety of both Reynolds number and angle of attack. Besides, mean and rms velocities showed that flow characteristics near wake were influenced by LSB and trailing-edge separation. At lower angles of attack, short bubble occurred, and frequency of vortex shedding due to short bubble were high whereas shedding frequency of long bubble at moderate angles of attack were low. The bubbles can cause vibration and noise at the wind turbine blade, hence they should be better known and eliminated, which aerodynamic performance and energy efficiency of turbine can be increased.

High-fidelity numerical simulation of the flow field around a NACA-0012 aerofoil from the laminar separation bubble to a full stall

ABSTRACTIn the present work, large eddy simulations of the flow field around a NACA-0012 aerofoil near stall conditions are performed at a Reynolds number of 5 × 104, Mach number of 0.4, and at various angles of attack. The results show the following: at relatively low angles of attack, the bubble is present and intact; at moderate angles of attack, the laminar separation bubble bursts and generates a global low-frequency flow oscillation; and at relatively high angles of attack, the laminar separation bubble becomes an open bubble that leads the aerofoil into a full stall. Time histories of the aerodynamic coefficients showed that the low-frequency oscillation phenomenon and its associated physics are indeed captured in the simulations. The aerodynamic coefficients compared to previous and recent experimental data with acceptable accuracy. Spectral analysis identified a dominant low-frequency mode featuring the periodic separation and reattachment of the flow field. At angles of attack α ≤ 9.3°, the low-frequency mode featured bubble shedding rather than bubble bursting and reformation. The underlying mechanism behind the quasi-periodic self-sustained low-frequency flow oscillation is discussed in detail.

Non-linear aerodynamic modelling of unmanned cropped delta configuration from experimental data

ABSTRACTThe paper presents the aerodynamic characterization of a low-speed unmanned aerial vehicle, with cropped delta planform and rectangular cross section, at and around high angles-of-attack using flight test methods. Since the linear models used for identification from flight data at low and moderate angles of attack become unsuitable for accurate parameter estimation at high angles of attack, a non-linear aerodynamic model has to be considered. Therefore, the Kirchhoff's flow separation model was used to incorporate the non-linearity in the aerodynamic model in terms of flow separation point and stall characteristic parameters. The Maximum Likelihood (ML) and Neural Gauss-Newton (NGN) methods were used to perform the parameter estimation on one set of low angle-of-attack and one set of near-stall flight data. It is evident from the estimates that the NGN method, which does not involve solving equations of motion, performs on a par with the classical ML method. This may be attributed to the reason that NGN method uses a neural network which has been trained by performing point to point mapping of the measured flight data. This feature of NGN method enhances its application over a wider envelope of high angles of attack flight data.

Periodic Filtering as a subgrid-scale model for LES of laminar separation bubble flows

ABSTRACTLaminar separation bubbles develop over many blades and airfoils at moderate angles of attack and Reynolds numbers ranging from 104 to 105. More accurate simulation tools are necessary to enable higher fidelity design optimisation for these airfoils and blades as well as to test flow control schemes. Following previous investigators, an equivalent problem is formulated by imposing suitable boundary conditions for flow over a flat plate which allows to use a high accuracy spectral solver. Large eddy simulation (LES) of such a flow were performed at drastically reduced resolution to assess the accuracy of several LES modelling approaches: the explicit dynamic Smagorinsky model, implicit LES, and the truncated Navier–Stokes approach (TNS). To mimic dissipation that occurs in implicit LES, the solution on a coarse mesh is filtered at every time step and two different filter strengths are used. In the TNS approach, the solution is filtered periodically, every few hundred time steps. The performance of each approach is evaluated against benchmark direct numerical simulation (DNS) data focusing on pressure and skin friction distributions, which are critical to airfoil designers. TNS results confirm that periodic filtering can act as an apt substitute for explicit subgrid-scale models, whereas filtering at every time step demonstrates the dependence of implicit LES on details of numerics.

Detailed analysis of the blade root flow of a horizontal axis wind turbine

Abstract. The root flow of wind turbine blades is subjected to complex physical mechanisms that influence significantly the rotor aerodynamic performance. Spanwise flows, the Himmelskamp effect, and the formation of the root vortex are examples of interrelated aerodynamic phenomena that take place in the blade root region. In this study we address those phenomena by means of particle image velocimetry (PIV) measurements and Reynolds-averaged Navier–Stokes (RANS) simulations. The numerical results obtained in this study are in very good agreement with the experiments and unveil the details of the intricate root flow. The Himmelskamp effect is shown to delay the stall onset and to enhance the lift force coefficient Cl even at moderate angles of attack. This improvement in the aerodynamic performance occurs in spite of the negative influence of the mentioned effect on the suction peak of the involved blade sections. The results also show that the vortex emanating from the spanwise position of maximum chord length rotates in the opposite direction to the root vortex, which affects the wake evolution. Furthermore, the aerodynamic losses in the root region are demonstrated to take place much more gradually than at the tip.

Vortical flow prediction of a diamond wing with rounded leading edges

The objective of this work is to assess the potential and limitations of current practice in computational fluid dynamic modeling in predicting vortical flowfields over a generic 53-degree swept diamond wing with rounded leading edges. This wing was designed under STO AVT Task Group 183 and has a constant NACA 64A-006 airfoil section with a leading edge radius of 0.264 percent chord. CFD simulations were run for different angles of attack at a Mach number of 0.15 and a Reynolds number of 2.7×106 based on the mean aerodynamic chord to match experiments. The wind tunnel experiments of the diamond wing were carried out in the Institute of Aerodynamics and Fluid Mechanics of the Technische Universität München, Germany and include aerodynamic lift, drag, and pitch moment measurements as well as span-wise pressure distributions at different chord-wise locations. This data set is used to validate the CFD results. The results presented demonstrate that the CFD compare well with the experiments at small angles of attack; the pitch moments predicted by the SARC turbulence model provide a better match to experimental results than the SA model at moderate angles of attack; and at high angles of attack, CFD predictions are not as good. The flow visualization results show that a leading-edge vortex is formed above the upper surface of the wing at an angle of attack of about eight degrees. This vortex becomes larger and stronger when the angle of attack is increased. With increasing angle of attack, the vortex formation point moves upstream and the vortex core moves inboard towards the wing center. Finally, the computational results show that the flow over the diamond wing is relatively steady throughout the range of angles of attack.

Riblet drag reduction in mild adverse pressure gradients: A numerical investigation

Riblet films are a passive method of turbulent boundary layer control that can reduce viscous drag. They have been studied with great detail for over 30years. Although common riblet applications include flows with Adverse Pressure Gradients (APG), nearly all research thus far has been performed in channel flows. Recent research has provided motivation to study riblets in more complicated turbulent flows with claims that riblet drag reduction can double in mild APG common to airfoils at moderate angles of attack. Therefore, in this study, we compare drag reduction by scalloped riblet films between riblets in a zero pressure gradient and those in a mild APG using high-resolution large eddy simulations. In order to gain a fundamental understanding of the relationship between drag reduction and pressure gradient, we simulated several different riblet sizes that encompassed a broad range of s+ (riblet width in wall units), similarly to many previously published experimental studies. We found that there was only a slight improvement in drag reduction for riblets in the mild APG. We also observed that peak values of streamwise turbulence intensity, turbulent kinetic energy, and streamwise vorticity scale with riblet width. Primary Reynolds shear stresses and turbulence kinetic energy production however scale with the ability of the riblet to reduce skin-friction.

Performance of subgrid-scale models in coarse large eddy simulations of a laminar separation bubble

The flow over many blades and airfoils at moderate angles of attack and Reynolds numbers ranging from 104 to 105 undergoes separation due to the adverse pressure gradient generated by surface curvature. In many cases, the separated shear layer then transitions to turbulence and reattaches, closing off a recirculation region—the laminar separation bubble. An equivalent problem is formulated by imposing suitable boundary conditions for flow over a flat plate to avoid numerical and mesh generation issues. Recent work demonstrated that accurate large eddy simulation (LES) of such a flow is possible using only O(1%) of the direct numerical simulation (DNS) resolution but the performance of different subgrid-scale models could not be properly assessed because of the effects of unquantified numerical dissipation. LES of a laminar separation bubble flow over a flat plate is performed using a pseudo-spectral Navier-Stokes solver at resolutions corresponding to 3% and 1% of the chosen DNS benchmark by Spalart and Strelets (2000). The negligible numerical dissipation of the pseudo-spectral code allows an unambiguous assessment of the performance of subgrid-scale models. Three explicit subgrid-scale models—dynamic Smagorinsky, σ, and truncated Navier-Stokes (TNS)—are compared to a no-model simulation (under-resolved DNS) and evaluated against benchmark DNS data focusing on two quantities of critical importance to airfoil and blade designers: time-averaged pressure (Cp) and skin friction (Cf) predictions used in lift and drag calculations. Results obtained with explicit subgrid-scale models confirm that accurate LES of laminar separation bubble flows is attainable with as low as 1% of DNS resolution, and the poor performance of the no-model simulation underscores the necessity of subgrid-scale modeling in coarse LES with low numerical dissipation.

Validation of CFD simulations for X-31 wind-tunnel models

AbstractComputational Fluid Dynamics (CFD) has become an attractive method of choice in the design of many aerospace vehicles because of advances in numerical algorithms and convergence acceleration methods. However, the flow around an advanced fighter aircraft is complicated and usually unsteady due to the presence of vortex-dominated flows. The accuracy and predictability of conventional turbulence models for these applications may be questionable and therefore results obtained from these models must be validated and evaluated on the basis of experimental data from wind tunnels and/or flight tests. This work aims to validate CFD simulations of X-31 wind-tunnel models with and without a belly-mounted sting. The sting setup facilitates forced sinusoidal oscillations in one of three modes of: pitch, yaw, and roll. However, the results show that measured aerodynamic data are altered by the turbulent wake behind the sting, even at small angles of attack. The high angle-of-attack flow around the X-31 is also very complicated and unsteady due to canard and wing vortices. Therefore, validation of CFD models for predicting these complex flows can be a very challenging task. The X-31 wind-tunnel experiments were carried out in the German Dutch low-speed wind tunnel at Braunschweig and include aerodynamic force and moment measurement as well as span-wise pressure distributions at locations of 60% and 70% chord length. This data set is used to validate the Cobalt and Kestrel flow solvers and the results are similar and match quiet well with experiments for small to moderate angles of attack. The main discrepancies between CFD and measurements occur close to the wing tip, where leading-edge flaps are located.

Numerical investigation of laminar vortex shedding applying a discontinuous Galerkin Finite Element method

Two-dimensional vortex sheddings in the wakes of a circular cylinder and a NACA 0012 airfoil are numerically investigated by applying a discontinuous Galerkin finite element method (DG). The fluid flow is assumed to be laminar and incompressible. The in-house code BoSSS, in which the projection scheme of Karniadakis et al. (1991) is implemented, is employed for performing the numerical simulations. In order to validate and assess this new approach, it has been applied to several generic flow configurations (i.e., flow past a cylinder, flow over an airfoil). For the cylinder, the flow for the low Reynolds numbers up to 100 is considered and the onset of vortex shedding is investigated. For the airfoil, the Strouhal numbers are computed for low to moderate angles of attack. The computed flow patterns around the airfoil are in a good agreement with the available experimental and numerical data.

Wing Rock Induced by a Hemisphere–Cylinder Forebody

The wing rock induced by a hemisphere–cylinder forebody was investigated experimentally using an isolated delta wing and wing body. For the delta wing with sweep angle of 30 deg, no apparent limit-cycle wing rock occurs at angles of attack of 0–90 deg. The mean roll angles exhibit nonzero values at angles of attack of less than 15 deg and become zero at angles of attack of more than 15 deg. For the wing body with a hemisphere–cylinder forebody, apparent wing rock exists at angles of attack of 0–90 deg; the wing rock is induced by the forebody. The phase diagrams showed that the wing rock motion is a typical limit-cycle oscillation at lower and moderate angles of attack. With increasing angles of attack, however, stochastic components in wing rock motion are increased gradually, and the motion types start to deviate from a limit-cycle oscillation. The phase-locked particle-image-velocimetry measurements for the wing body revealed that the vortex pair over the forebody exhibits apparent dynamic hysteresis in position and strength during wing rock, and the dynamic hysteresis of forebody vortices further influence flowfields over the wing, resulting in asymmetric distributions of wing flow, which provides the driving moments sustaining the wing rock.

An experimental study of airfoil instability tonal noise with trailing edge serrations

This paper presents an experimental study of the effect of trailing edge serrations on airfoil instability noise. Detailed aeroacoustic measurements are presented of the noise radiated by an NACA-0012 airfoil with trailing edge serrations in a low to moderate speed flow under acoustical free field conditions. The existence of a separated boundary layer near the trailing edge of the airfoil at an angle of attack of 4.2 degree has been experimentally identified by a surface mounted hot-film arrays technique. Hot-wire results have shown that the saw-tooth surface can trigger a bypass transition and prevent the boundary layer from becoming separated. Without the separated boundary layer to act as an amplifier for the incoming Tollmien–Schlichting waves, the intensity and spectral characteristic of the radiated tonal noise can be affected depending upon the serration geometry. Particle Imaging Velocimetry (PIV) measurements of the airfoil wakes for a straight and serrated trailing edge are also reported in this paper. These measurements show that localized normal-component velocity fluctuations that are present in a small region of the wake from the laminar airfoil become weakened once serrations are introduced. Owing to the above unique characteristics of the serrated trailing edges, we are able to further investigate the mechanisms of airfoil instability tonal noise with special emphasis on the assessment of the wake and non-wake based aeroacoustic feedback models. It has been shown that the instability tonal noise generated at an angle of attack below approximately one degree could involve several complex mechanisms. On the other hand, the non-wake based aeroacoustic feedback mechanism alone is sufficient to predict all discrete tone frequencies accurately when the airfoil is at a moderate angle of attack.

Forebody Flow During the Wing Rock Over Low Swept Wing/Body Configuration in High Rate Pitching-up

Over the configuration of a pointed ogive-cylindrical body with 30° swept wing, wind tunnel experiments are conducted to investigate the wing rock motion in pitching-up as well as the flow structure responsible for it. Results show that wing rock would present as a sinusoid like motion during angle of attack 20°∼85° in high rate pitching-up.And the evolvement of forebody vortices with angle of attack should be responsible for the sinusoid like motion. Firstly, the wing rock is triggered by forebody asymmetric vortices at moderate angle of attack (about 30°). Secondly, the variation of forebody asymmetric vortices with angle of attack sustains the wing rock during angle of attack 30°∼60°. At last, the wing rock after angle of attack 60° is a convergent process of rolling due to the breakdown of forebody asymmetric vortices.

Computational and experimental results in the open test section of the aeroacoustic windtunnel Braunschweig

This paper discusses results of numerical simulations for low speed open tunnel test section in connection to wind tunnel experiments aiming to assess the computational accuracy. The investigations are performed for the aeroacoustic wind tunnel Braunschweig with and without a mounted experimental model. RANS simulations are performed within a CFD procedure developed to simulate open test sections. In the empty tunnel, measurements of the free shear layers and of the flow fields downstream the nozzle show a good agreement with the predicted flows. The jet core size and its position, as well as the velocity gradients in the mixing layer are simulated mostly accurate. Afterwards, the in-tunnel simulations with a 2D-high lift model are discussed. These reproduce very well the measured global lift force and local flow pressure distributions at moderate angles of attack but show discrepancies at maximum lift. In overall a good agreement of the numerical results compared to the experiments is achieved.

Study of the hydrodynamic of a flapping foil at moderate angle of attack

The mechanics of a rigid flapping foil animated by a combination of harmonic heave translation and pitch rotation is examined numerically and experimentally by Digital Particle Image Velocimetry (DPIV). The Arbitrary-Lagrangian–Eulerian (ALE) technique associated with a r-refinement grid adaption algorithm provides a good solution for studying the foil in a fixed reference frame while maintaining the grid quality over the whole simulation. Quantitative predictions were made, and showed very good agreement with the experimental data for a reduced frequency ranging in (0–0.6) and with the fixed values of Reynolds (4×103), heave amplitude (3c/4) and maximum angle of attack (20°). At low Strouhal numbers we observed a linear relationship between the thrust coefficient and the instantaneous angle of attack; for higher reduced frequency a non-linear behavior is observed, linked to added mass effects. Under the assumption of small angle of attack, these added mass effects correspond to a pitch stiffening and a heave damping.

Coherent structure dynamics and sediment erosion mechanisms around an in‐stream rectangular cylinder at low and moderate angles of attack

Local scour around elongated in‐stream structures (e.g., high–aspect ratio rectangular bridge piers) is mainly driven by the interactions between the erodible bed and the large‐scale coherent structures generated by the presence of the flow obstruction. The present investigation uses eddy‐resolving numerical simulations to study the mean flow and turbulence structure around a high–aspect ratio rectangular cylinder placed in a flat bed channel. Simulations are conducted for three angles of attack (α = 0°, 15°, and 30°) at a channel Reynolds number of 2.4 × 105. This paper focuses on the dynamics of the large‐scale coherent structures forming around the rectangular cylinder and their role in controlling sediment entrainment at conditions corresponding to the start of the scour process. Simulation results show that most of the sediment is entrained from the bed by the eddies shed inside the separated shear layers (SSLs), by the legs of the necklace vortices, and by the strongly accelerated flow on the outer side of the SSLs. For α = 0° and 15°, the horseshoe vortex (HV) system plays a relatively minor role in the entrainment of sediment in front of the cylinder, and the passage of the wake vortices (rollers) results in a small amplification of the bed friction velocity. In contrast, for α = 30°, the unsteady dynamics of the main necklace vortices part of the HV system and of the roller vortices results in a significant amplification of the instantaneous bed friction velocity. The mean flux of sediment entrained from the bed calculated on the basis of the mean flow field is found to underestimate by 2–3 times the same quantity when estimated more correctly on the basis of the instantaneous flow fields. The primary reason for this underestimation is sediment entrainment in regions where the mean flow bed friction velocity is smaller than the critical value for entrainment. Quantitative information on the extent of the regions of high values of the bed friction velocity and its standard deviation and on the sediment entrainment flux as a function of the angle of attack is essential to guide scour protection measures around the in‐stream noncircular obstructions.

Aerodynamic and aerothermodynamic trade-off analysis of a small hypersonic flying test bed

This paper deals with the aerodynamic and aerothermodynamic trade-off analysis aiming to design a small hypersonic flying test bed with a relatively simple vehicle architecture. Such vehicle will have to be launched with a sounding rocket and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry technologies such as boundary-layer transition and shock–shock interaction phenomena. The flight shall be conducted at hypersonic Mach number, in the range 6–8 at moderate angles of attack. In the paper some design analyses are shown as, for example, the longitudinal and lateral-directional stability analysis. A preliminary optimization of the configuration has been also done to improve the aerodynamic performance and stability of the vehicle. Several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper. The aerodynamic model of vehicle is also provided.

Framework for Establishing Limits of Tabular Aerodynamic Models for Flight Dynamics Analysis

This paper describes the use of Computational Fluid Dynamics for the generation and testing of tabular aerodynamic models for flight dynamics analysis. Manoeuvres for the AGARD Standard Dynamics Model wind tunnel geometry for a generic fighter are considered as a test case. Wind tunnel data is first used to validate the prediction of static and dynamic coefficients at both low and high angles, featuring complex vortical flow, with good agreement obtained at low and moderate angles of attack. Then the generation of aerodynamic tables is described based on an efficient data fusion approach. An optimisation is used to define time optimal manoeuvres based on these tables, including level flight trim, pull-ups at constant and varying incidence, and level and 90 degree turns. The manoeuvre description includes a definition of the aircraft states and also the control deflections to achieve the motion. The main point of the paper is then to assess the validity of the aerodynamic tables which were used to define the manoeuvres. This is done by replaying them, including the control surface motions, through the time accurate CFD code. The resulting forces and moments can be compared with the tabular values to assess the presence of inadequately modelled dynamic or unsteady effects. The agreement between the tables and the replay is demonstrated for slow manoeuvres. A study for the pull-up at increasing rates shows discrepancies which are ascribed to vortical flow hysteresis at elevated motion rates.