Pre-stall Conditions Research Articles

Blade aerodynamic model for thick airfoils of the series NACA00xx in pre-stall conditions

Blade aerodynamic modeling is needed for design, control, and aeroelastic studies of wind turbines. The ultimate aim of this study is to establish a blade aerodynamic model with well-quantified accuracy for thick airfoils, predominant in wind turbine blades. The study is limited to pre-stall conditions, involving only attached and trailing-edge separated flows. The account of dynamic stall will be considered in further studies. The analysis of the Glasgow University database on the family of symmetrical airfoils NACA 00xx (xx = 12,…,30) has been made by considering particularly the 2D spatiotemporal contours of the surface pressure coefficients on the suction side. Such contours provide a clear visualization of flow regime type (attached, separated or stalled) and, therefore, allow the selection of oscillatory test cases in attached or trailing-edge separated flows. The aerodynamic model of the normal force coefficient is established by improving the Beddoes–Leishman BL model. An important modification is carried on the calculation of the delayed angle of attack using the Goman–Khrabrov model, instead of the complex original procedure. There is a new aerodynamic component for simulating the trailing-edge separation. The present model, although limited to pre-stall conditions, involves ten parameters for the unsteady aerodynamic behavior. They can be obtained with the global optimization of the deviations between experimental results and model predictions. Previous optimization studies of the parameters of the BL model involve all flow regimes for test cases and do not lead to conclusive results. The parameter values obtained in the present study show a coherent and physics-expected variation with airfoil thickness that is not negligible.

Influence of vortex interactions on the exergy transfer and performance of OWC wells turbine

To provide deep dives about aerodynamic loss mechanisms in Wells turbines for wave energy conversion, a loss audit analysis was performed by numerical experiments in a monoplane Wells turbine with guide vanes. The interactions between the tip-leakage and leading-edge vortices during the stall process were captured by an improved vortex identification method, which revealed the relationship between vortex interactions and stall mechanisms by identifying coherent structures and tracking the vortex core trajectory. Finally, the influence of vortex interactions on exergy transfer was quantified. The results indicate that the lost kinetic energy and mixing losses dominate the loss generation in the Wells turbine stage under stall conditions. Under the beneficial effect of tip leakage flow, leading-edge separation first begins at the equilibrium region between the tip-leakage and leading-edge vortices. As the leading-edge vortices expand toward the blade tip, the intensified leading-edge vortex interacts with the casing suction-side corner vortex and accelerates the dissipation of the tip-leakage vortices. Consequently, the contributions of viscous irreversibilities outweigh those of shaft work, being the dominant factor in the decrease in flow exergy, leading to a decrease in exergy utilization by 38.46% from the pre-stall condition to the stall condition.

Numerical simulation of NACA4412 airfoil in pre-stall conditions

PurposeThe purpose of this paper is to study turbulent flow separation at the airfoil trailing edge. This work aims to improve the knowledge of stall phenomenon by creating a QDNS database for the NACA412 airfoil.Design/methodology/approachQuasi-DNS simulations of the NACA 4412 airfoil in pre-stall conditions have been completed. The Reynolds number based on airfoil chord and freestream velocity is equal to 0.35 million, and the freestream Mach number to 0.117. Transition is triggered on both surfaces for avoiding the occurrence of laminar separation bubbles and to ensure turbulent mixing in the wake. Four incidences have been considered, 5, 8 10 and 11 degrees.FindingsThe results obtained show a reasonably good correlation of the present simulations with classical MSES airfoil simulations and with RANS computations, both in terms of pressure and skin-friction distribution, with an earlier and more extended flow separation in the QDNS. The database thus generated will be deeply analysed and enriched for larger incidences in the future.Originality/valueNo experimental or HPC numerical database at reasonable Reynolds number exists in the literature. The current work is the first step in that direction.

Numerical Analysis of the Rudder–Propeller Interaction

The proper evaluation of the Rudder–Propeller interactions is mandatory to correctly predict the manoeuvring capability of a modern ship, in particular considering the commonly adopted ship layout (rudder often works in the propeller slipstream). Modern Computational Fluid Dynamics (CFD) solvers can provide, not only the performance of the whole system but also an insight into the flow problem. In the present paper, an open-source viscous flow solver has been validated against available literature experimental measurements in different conditions. After an extensive analysis of the numerical influence of the mesh arrangement and the turbulent quantities on the rudder provided forces, the study focused its attention on the forces generated by the rudder varying the propeller loading conditions and the mutual position between the two devices. These analyses give a hint to describe and improve a commonly-used semi-empirical method based on the actuator disk theory. These analyses also demonstrate the ability of these numerical approaches to correctly predict the interaction behaviour in pre-stall conditions with quite reasonable computational requests (proper also for a design stage), giving additional information on the sectional forces distribution along the span-wise rudder direction, useful to further develop a new semi-empirical rudder model.

Aerofoil dipole noise due to flow separation and stall at a low Reynolds number

Aerofoil self-noise produced by flow separation and stall is relatively little understood regarding the underlying generation mechanisms. The focus of this work is to provide an improved level of understanding particularly with regard to the dipole noise sources utilising a high-fidelity direct numerical simulation. A NACA0012 aerofoil is considered under three different flow conditions at a Reynolds number Re∞=50,000 and a Mach number M∞=0.4. These include: a pre-stall condition with a laminar separation bubble (α=5∘), near-stall (α=10∘), and fully stalled (α=15∘). The noise radiation in the far-field is significantly increased at low frequencies for the full-stall case for all observer directions which is consistent with previous experimental observations. The dominant source regions for each configuration are identified for low, medium and high frequencies, separately. A number of key findings are made concerning the source characteristics in full-stall which differ considerably from the lower angle of attack cases. It is found that the location of the dominant sources changes more significantly with frequency for the full-stall case. Additionally, for medium to high frequencies the maximum acoustic source amplitude is weaker for the full-stall case, despite comparable levels observed in the far-field. This seemingly contradictory observation highlights the importance of phase variations in the wall pressure fluctuations. For the frequencies considered in this paper it is shown that the full-stall case usually produces a relatively more in-phase source distribution, resulting in a more efficient radiation despite the lower amplitude levels. The important flow structures which are responsible for the dipole sources are also identified through analysis of the pressure field at isolated frequencies. It is found that for low frequencies coherent structures in the shear layer are responsible for the scattering of the wall pressure fluctuations at the TE, which agrees with previous findings in the literature. However, at medium and high frequencies the shear layer structures are found to be relatively weak in the proximity of the TE. This indicates that the noise may be generated through other means, for example scattering of fluctuating pressure induced by vortices shed from the TE.

Numerical simulation of flapping airfoil with alula

Bionic micro aerial vehicles have become popular because of their high thrust efficiency and deceptive appearances. Leading edge or trailing edge devices (such as slots or flaps) are often used to improve the flight performance. Birds in nature also have leading-edge devices, known as the alula that can improve their flight performance at large angles of attack. In the present study, the aerodynamic performance of a flapping airfoil with alula is numerically simulated to illustrate the effects of different alula geometric parameters. Different alula relative angles of attack β (the angle between the chord line of the alula and that of the main airfoil) and vertical distances h between the alula and the main airfoil are simulated at pre-stall and post-stall conditions. Results show that at pre-stall condition, the lift increases with the relative angle of attack and the vertical distance, but the aerodynamic performance is degraded in the presence of alula compared with no alula, whereas at post-stall condition, the alula greatly enhances the lift. However, there seems to be an optimal relative angle of attack for the maximum lift enhancement at a fixed vertical distance considering the unsteady effect, which may indicate birds can adjust the alula twisting at different spanwise positions to achieve the best flight performance. Different alula geometric parameters may affect the aerodynamic force by modifying the pressure distribution along the airfoil. The results are instructive for design of flapping-wing bionic unmanned air vehicles.

Improvement of Wall-Corrections at the National Research Council’s Trisonic Wind Tunnel

The development of flow separation on a model tested at subsonic flow conditions reaching stall is not currently accounted for by the wall-correction methodology used at the National Research Council’s (NRC’s) 5 ft trisonic wind tunnel. Consequently, although this method provides accurate wall corrections during tests in prestall conditions, it is unable to generate reliable corrections in stall. The present paper investigates the possible improvements to the singularity representation of the model, which are paramount to the “one-variable” wall-correction methodology. This revision of the model representation is based on high-fidelity free-air computational fluid dynamics solutions for the NASA Common Research Model (CRM), which are validated by experimental datasets obtained from a semispan version of the CRM tested at the NRC’s 5 ft trisonic wind tunnel. The improved potential representation is tested on solid-wall experimental wind-tunnel data to present more reliable behavior of the wall interference correction estimates when substantial wing stalling is encountered.

On the Effects of Tip Clearance and Operating Condition on the Flow Structures Within an Axial Turbomachine Rotor Passage

Abstract Effects of tip clearance size and flowrate on the flow around the tip of an axial turbomachine rotor are studied experimentally. Visualizations and stereo-particle image velocimetry (PIV) measurements in a refractive index-matched facility compare the performance, leakage velocity, and the trajectory, growth rate, and strength of the tip leakage vortex (TLV) for gaps of 0.49% and 2.3% of the blade chord, and two flowrates. Enlarging the tip clearance delays the TLV breakup in the aft part of the rotor passage at high flowrates but causes earlier breakup under pre-stall conditions. It also reduces the entrainment of endwall boundary layer vorticity from the separation point where the leakage and passage flows meet. Reducing the flowrate or tip gap shifts the location of the TLV detachment from the blade suction side (SS) upstream to points where the leakage velocity is 70–80% of the tip speed. Once detached, the growth rates of the total shed circulation are similar for all cases, i.e., varying the gap or flowrate mostly shifts the detachment point. The TLV migration away from the SS decreases with an increasing gap but not with the flowrate. Two mechanisms dominate this migration: initially, the leakage jet pushes the TLV away from the blade at 50% of the leakage velocity. Further downstream, the TLV is driven by its image on the other side of the endwall. Differences in migration rate are caused by the smaller distance between the TLV and its image for the narrow gap, and the increase in initial TLV strength with decreasing flowrate and gap.

Numerical Simulations of Low-Reynolds-Number Flow Past Finite Wings with Leading-Edge Protuberances

The use of various modeling approaches for the numerical simulation of low-Reynolds-number flow past wings of finite span with leading-edge protuberances is studied at pre- and poststall operating conditions. Variants of Reynolds-averaged Navier–Stokes simulations (including transitional modeling) and detached-eddy simulations (involving different options for shielding the boundary layer from the detached-eddy simulation limiter), as well as a low-Reynolds-number correction, are considered. An assessment of the capabilities exhibited by these modeling approaches is carried out, based on detailed comparisons against whole-flowfield and aerodynamic force measurements in a wind tunnel, at a Reynolds number of for wings with aspect ratios of 1 and 1.5. The results show that the use of detached-eddy simulation with an improved description of the boundary layer is required to reproduce the flow physics accurately at prestall conditions, where separation and subsequent reattachment are observed. At poststall regimes, massive flow separation occurs, and a choice of conservative shielding produces the best outcome with detached-eddy simulation. In addition, the present experimental dataset provides valuable information toward understanding the passive stall control mechanism generated by the use of leading-edge tubercles and its application to low-aspect-ratio wings.

Physics of Prestall Propagating Disturbances in Axial Compressors and Their Potential as a Stall Warning Indicator

Axial compressors in aero engines are prone to suffering a breakdown of orderly flow when operating at the peak of the pressure rise characteristic. The damaging potential of separated flows is why a safe distance has to be left between every possible operating point and an operating point at which stall occurs. During earlier investigations of stall inception mechanisms, a new type of prestall instability has been found. In this study, it could be demonstrated that the prestall instability characterised by discrete flow disturbances can be clearly assigned to the subject of “Rotating Instabilities”. Propagating disturbances are responsible for the rise in blade passing irregularity. If the mass flow is reduced successively, the level of irregularity increases until the prestall condition devolves into rotating stall. The primary objective of the current work is to highlight the basic physics behind these prestall disturbances by complementary experimental and numerical investigations. Before reaching the peak of the pressure rise characteristic flow, disturbances appear as small vortex tubes with one end attached to the casing and the other attached to the suction surface of the rotor blade. These vortex structures arise when the entire tip region is affected by blockage and at the same time the critical rotor incidence is not exceeded in this flow regime. Furthermore, a new stall indicator was developed by applying statistical methods to the unsteady pressure signal measured over the rotor blade tips, thus granting a better control of the safety margin.

Visualizations of Flow Structures in the Rotor Passage of an Axial Compressor at the Onset of Stall

Experiments preformed in the JHU refractive index matched facility examine flow phenomena developing in the rotor passage of an axial compressor at the onset of stall. High-speed imaging of cavitation performed at low pressures qualitatively visualizes vortical structures. Stereoscopic particle image velocimetry (SPIV) measurements provide detailed snapshots and ensemble statistics of the flow in a series of meridional planes. At prestall condition, the tip leakage vortex (TLV) breaks up into widely distributed intermittent vortical structures shortly after rollup. The most prominent instability involves periodic formation of large-scale backflow vortices (BFVs) that extend diagonally upstream, from the suction side (SS) of one blade at midchord to the pressure side (PS) near the leading edge of the next blade. The 3D vorticity distributions obtained from data recorded in closely spaced planes show that the BFVs originate form at the transition between the high circumferential velocity region below the TLV center and the main passage flow radially inward from it. When the BFVs penetrate to the next passage across the tip gap or by circumventing the leading edge, they trigger a similar phenomenon there, sustaining the process. Further reduction in flow rate into the stall range increases the number and size of the backflow vortices, and they regularly propagate upstream of the leading edge of the next blade, where they increase the incidence angle in the tip corner. As this process proliferates circumferentially, the BFVs rotate with the blades, indicating that there is very little through flow across the tip region.

Numerical Study of Reduced Frequency Effect on Longitudinal Stability Derivatives of Airfoil under Pitching and Plunging Oscillations

In this study, incompressible, unsteady and turbulent flow over an airfoil with pitching and plunging oscillations is numerically studied in order to investigate the effect of reduced frequency on stability derivatives of oscillating airfoil. Linear k – e model called Launder-Sharma and Rhie and Chow model are used for turbulence modeling and overcoming pressure checkerboard problem. This means that a co-located approach is proposed in this paper to study a moving grid problem, and the results demonstrate the high accuracy of the method. Control volume and Crank-Nicholson discretization method are also used for the numerical solution. It is shown that the longitudinal stability derivatives of plunging and pitching motions trend change intensively beyond the stall angle of attack while pitching rate has a completely opposite behavior. The results also show that increasing reduced frequency leads to stability reduction in plunging oscillation but it does not have significant effect on pitching oscillation case in pre-stall, stall and post-stall conditions. Grid convergence is examined to assess the accuracy of the numerical method that shows the high accuracy of it and this is a prominent achievement of the present study. The results of the proposed method in forces and moment show a good agreement with the experimental data.

Visualization and Time-Resolved Particle Image Velocimetry Measurements of the Flow in the Tip Region of a Subsonic Compressor Rotor

A new optically index matched facility has been constructed to investigate tip flows in compressor-like settings. The blades of the one and a half stage compressor have the same geometry, but lower aspect ratio as the inlet guide vanes (IGVs) and the first stage of the low-speed axial compressor (LSAC) facility at NASA Glenn. With transparent blades and casings, the new setup enables unobstructed velocity measurements at any point within the tip region and is designed to facilitate direct measurements of effects of casing treatments on the flow structure. We start with a smooth endwall casing. High speed movies of cavitation and time-resolved PIV measurements have been used to characterize the location, trajectory, and behavior of the tip leakage vortex (TLV) for two flow rates, the lower one representing prestall conditions. Results of both methods show consistent trends. As the flow rate is reduced, TLV rollup occurs further upstream, and its initial orientation becomes more circumferential. At prestall conditions, the TLV is initially aligned slightly upstream of the rotor passage, and subsequently forced downstream. Within the passage, the TLV breaks up into a large number of vortex fragments, which occupy a broad area. Consequently, the cavitation in the TLV core disappears. With decreasing flow rate, this phenomenon becomes more abrupt, occurs further upstream, and the fragments occupy a larger area.

Low Reynolds number airfoil aerodynamic loads determination via line integral of velocity obtained with particle image velocimetry

The small magnitude lift forces generated by both a NACA 0012 airfoil and a thin flat plate at Re = 29,000 and 54,000 were determined through the line integral of velocity, obtained with particle image velocimetry, via the application of the Kutta–Joukowsky theorem. Surface pressure measurements of the NACA0012 airfoil were also obtained to validate the lift coefficient C l. The bound circulation was found to be insensitive to the size and aspect ratio of the rectangular integration loop for pre-stall angles. The present C l data were also found to agree very well with the surface pressure-determined lift coefficient for pre-stall conditions. A large variation in C l with the loop size and aspect ratio for post-stall conditions was, however, observed. Nevertheless, the present flat-plate C l data were also found to collectively agree with the published force-balance measurements at small angles of attack, despite the large disparity exhibited among the various published data at high angles. Finally, the ensemble-averaged wake velocity profiles were also used to compute the drag coefficient and, subsequently, the lift-to-drag ratio.

波状前縁付き矩形翼周りの流れ解析

The objective of the present study is to investigate the flow mechanism of delaying stall by use of a wavy leading edge on rectangular wing. Flows around rectangular wings with wavy leading edge were simulated by large eddy simulation with the Dynamic Smagorinsky Model (DSM). The Reynolds number based on the chord length of the wing and the uniform-flow velocity was 1:38X105. The predicted lift and drag coefficients for poststall condition agreed qualitatively with experimental data by using DSM, moreover those for prestall condition were in very good agreement with the experimental data. The mechanism of delaying stall was revealed by the flow simulation. The predicted flows around the wavy leading edge indicated existence of longitudinal vortices and the flows around wavy leading edge remained attached at poststall angle of attack. The longtiudinal vortices were carried to wing tips and the flow separation was suppressed in the area from the protuberances to the wing tips.

Leading-Edge Separation Control by Means of Pulsed Vortex Generator Jets

DOI: 10.2514/1.26176 An airfoil was designed to stall with an abrupt turbulent leading-edge separation and equipped with a pneumatic system for active flow control. The system pulses compressed air through 45 deg skewed slots. Depending on the position of the slots, two very different behaviors were found. With the actuators positioned in the region of the separation line, the control system could not prevent separation. However, a very advantageous influence on the separation resulted with even higher normal forces than in prestall condition. The duty cycle was found to be the major control parameter for this kind of control. When the actuation was positioned well in front of the separation line the influence of the actuation system on the stall behavior changed completely. Here the system was able to prevent the leading-edge separation. The influence of the duty cycle becomes much weaker and maximum normal force scaled mainly with the mean ejected momentum, varied either by the duty cycle or the supply pressure.

Flow Instability at the Inlet of a Centrifugal Compressor

Using laser Doppler anemometry, two components of the gas velocity have been measured at the inlet of a centrifugal compressor impeller, operated at speeds typical of service conditions for a medium-sized turbocharger. The flow was found to be unstable, especially adjacent to the suction side of the blades, such that two predominant conditions existed in the flow. The unstable flow is illustrated in the paper by distributions of relative velocity and relative flow angle, and the effects of different operating conditions on these distributions are examined. The instability is believed to be caused by a pre-stall condition as the compressor operating point approaches a fully stalled condition which occurs during surge.