Spanwise Distribution Research Articles

Experimental and Numerical Investigation of Active Flow Control of a Serpentine Diffuser

Total-pressure losses and distortion in an aggressive offset diffuser, induced by secondary-flow vortices and boundary-layer separation, are mitigated using fluidic actuation for the aerodynamic interface plane Mach number up to 0.57. Although total-pressure recovery is primarily impacted by separation in each of the diffuser’s turns, distortion is governed by counter-rotating streamwise vortices that advect low-momentum fluid from the wall region into the core flow. The present investigations have shown that secondary-flow vortices couple to the outboard segments of the separated-flow domains. Therefore, fluidic control of the scale and topology of the trapped vorticity within the internally separated flow can be leveraged to control the structure and strength of the ensuing secondary vortices, and thereby significantly reduce flow distortion and losses. In the present investigations, fluidic control is effected by a spanwise array of oscillating jets that are placed just upstream of the second-turn separation domain. Actuation directly affects the local flow separation, but it also alters the spacing and diminishes the strength of the base-flow streamwise vortices by forming and adjoining streamwise-vorticity concentrations of opposite sense. Additionally, the spanwise distribution of the actuator array was optimized to reduce circumferential distortion by 60% at actuation to a diffuser mass-flow-rate ratio lower than 0.5%.

High Dimensional Matching Optimization of Impeller–Vaned Diffuser Interaction for a Centrifugal Compressor Stage

Abstract The matching and interaction between the impeller and vaned diffuser is the most important aerodynamic-coupling between the components of a high-speed centrifugal compressor. Many research studies have been carried out during the last decade, both experimentally and numerically, on the flow mechanisms underlying impeller–vaned diffuser matching and interaction, with the aim of achieving a high-performance stage. However, the published work lacks any study that optimizes the matching of the impeller–vaned diffuser components in the environment of a full compressor stage due to two unresolved issues, i.e., identifying an effective matching optimization strategy and the high dimensional nature of the problem. To tackle these difficulties, four different optimization strategies (i.e., (1) integrated, (2) single component, (3) parallel, and (4) sequential optimization strategies) have been proposed and validated through a high dimensional matching optimization of the Radiver compressor test case published by the Institute of Jet Propulsion and Turbomachinery at Rheinisch-Westfälische Technische Hochschule (RWTH) Aachen University. Particular attention has been paid to the slope of the diffuser total pressure ratio characteristic near the surge point to further extend the stage surge margin. The results showed that the integrated optimization strategy was the most effective one for achieving good matching of the impeller–vaned diffuser interaction due to its inherently strong coupling optimization. Compared with the baseline compressor, the optimized stage achieved a gain of 1.2% in total-to-total isentropic efficiency at the peak efficiency point as well as a predicted 26.17% increase in stable operating range. For the stage examined in this study, a fore-loaded design of impeller blade as well as an increased vane angle for the diffuser vane was beneficial to the impeller–vaned diffuser matching. The more uniform spanwise distributions of the impeller discharge flow angle and the diffuser vane incidence presented the opportunity for a more optimized matching of the flow field between the 3D impeller and the 2D vaned diffuser. The outcomes of this work are particularly relevant for the advanced design of high-speed centrifugal compressors.

Experimental investigations on flow over a circular cylinder placed in a circular pipe

Flow over a circular cylinder placed inside a circular pipe is studied experimentally to understand the influence of Reynolds numbers (ReD = DUav/υ, where D is the diameter of the pipe, Uav is the average velocity in the pipe, and υ is the kinematic viscosity of the fluid) and blockage effects. In the present configuration, the influence of confinement, aspect ratio, upstream turbulence, shear, and end conditions coexists together. The wake dynamics of such a configuration are seldom reported in the literature. The Reynolds number range covered in the present study is ReD = 200–2.0 × 105. Four different flow regimes are defined based on the Reynolds number: steady, unsteady laminar, transition, and turbulent. In the unsteady laminar regime, the Strouhal number (St = fd/Uav, where f is the frequency of vortex shedding and d is the diameter of the bluff body) increases gradually. In the transition regime, a gradual fall in St is observed for all blockage ratios (d/D). In the turbulent regime, the upstream flow becomes fully turbulent, and the Strouhal number remains constant. The spanwise pressure distribution is influenced due to the blockage effects near the wall. The separation point moves 20° toward the rear stagnation point close to the wall compared to the center in the turbulent regime. A universal relation for the drag coefficient is proposed based on the pressure at the stagnation and separation points. The universal Strouhal number defined based on the wake width and velocity at the separation (Us) is shown to be independent of the blockage ratio. The results presented in the manuscript are relatively new in the domain of confined bluff body flows and will serve to enhance the general understanding of confined bluff body vortex dynamics.

Comprehensive Improvement of Mixed-Flow Pump Impeller Based on Multi-Objective Optimization

The spanwise distribution of impeller exit circulation (SDIEC) has a significant effect on the impeller performance, therefore, there is a need for its consideration in the optimization design of mixed-flow pumps. In this study, a combination optimization system, including a 3D inverse design method (IDM), computational fluid dynamics (CFD), Latin hypercube sampling (LHS) method, response surface model (RSM), and non-dominated sorting genetic algorithm (NSGA-Ⅱ) was used to improve the performance of the mixed-flow pump after considering the effect of SDIEC on the performance of the impeller. The CFD results confirm the accuracy and credibility of the optimization results because of the good agreement the CFD results established with the experimental measurements. Compared with the original impeller, the pump efficiency of the preferred impeller at 0.8Qdes, 1.0Qdes, and 1.2Qdes improved by 0.63%, 3.39%, and 3.77% respectively. The low-pressure region on the blade surface reduced by 96.92% while the pump head difference was less than 1.84% at the design point. In addition, a comparison of the flow field of the preferred impeller and the original impeller revealed the effect of SDIEC on mixed-flow pump performance improvement and flow mechanism.

Aerodynamic analysis of wings in Chevron and V formation flights

In this paper, numerical analysis is conducted using a local camber correction approach called “decambering” to predict pre and post-stall aerodynamic characteristics of multiple lifting surfaces operating in formation. A three wings Chevron formation and five and nine wings V-formation are studied. NACA2412 wing section is used and experimental validation is provided with Cessna 172 aircrafts flying in Chevron formation. Effect of wing incidence and shifting of stall angles is looked at along with changes in geometric offsets between the wings in formation. The spanwise distribution of coefficients of lift and induced drag at different angles of attack, including high and post-stall angles of attack is studied for all the wings. The span efficiency factor, which represents the correction in drag due to change in lift as compared to that of an ideal wing of the same aspect ratio but with elliptical lift distribution is calculated. The maximum possible efficiency is then used to estimate the maximum reduction in drag possible for individual wings in different formations. The change in efficiency with number of lifting surfaces in a formation is also estimated.

A novel approach for the design of axial flow fan by increasing by-pass ratio in a constant-diameter turbofan

The present study introduces an innovative aerodynamic redesign of an axial flow fan based on constant diffusion factor and radial equilibrium. All input design parameters such as mass flow rate, hub to tip ratio, aspect ratio, tip diameter and angular velocity are taken from NASA Rotor 67 as a conventional axial flow fan. A computer program is developed to extract the three-dimensional geometry of a fan and to estimate the span-wise distribution of parameters. The new designed fan flow field is investigated in detail by CFD tool at both design and off design conditions. Finally, a turbofan cycle analysis is conducted based on thermodynamic and gas dynamic principles to evaluate the fan performance in a turbofan engine in comparison to NASA Rotor 67. Achieving a higher total pressure ratio, meeting the target pressure ratio in lower rotational speed with higher efficiency, delivering more bypass air in a constant diameter and less fuel consumption for the same specific thrust force are the main advantages for the new design strategy in comparison to the conventional designed fans such as Rotor 67. However, efficiency reduction in fan over speed is the main disadvantage.

On the formation of three-dimensional separated flows over wings under tip effects

We perform DNS of flow over finite-aspect-ratio NACA 0015 wings to characterize the tip effects on the wake dynamics. This study focuses on the development of three-dimensional separated flow over the wing, and discuss flow structures formed on the wing surface as well as in the far field wake. Vorticity is introduced into the flow in a predominantly two-dimensional manner. The vortex sheet from the wing tip rolls up around the free end to form the tip vortex. At its inception, the tip vortex is weak and its effect is spatially confined. As the flow around the tip separates, the tip effects extend farther in the spanwise direction, generating noticeable three dimensionality in the wake. For low-aspect-ratio wings, the wake remains stable due to the strong downwash over the entire span. Increasing the aspect ratio allows unsteady vortical flow to emerge away from the tip at sufficiently high angles of attack. These unsteady vortices shed and form closed vortical loops. For higher-aspect-ratio wings, the tip effects retard the near-tip shedding process, which desynchronizes from the two-dimensional shedding over the mid-span region, yielding vortex dislocation. At high angles of attack, the tip vortex exhibits noticeable undulations due to the strong interaction with the unsteady shedding vortices. The spanwise distribution of force coefficients is related to the three-dimensional wake dynamics and the tip effects. Vortical elements in the wake that are responsible for the generation of lift and drag forces are identified through the force element analysis. We note that at high angles of attack, a stationary vortical structure forms at the leading edge near the tip, giving rise to locally high lift and drag forces. The analysis performed here reveals how the vortical flow around the tip influences the separation physics, the global wake dynamics, and the spanwise force distributions.

Influence of tubercles’ spanwise distribution on swept wings for unmanned aerial vehicles

In this work, a 3D numerical study on the influence of the spanwise distribution of tubercles on a unmanned aerial vehicle wing is presented. The idea of using tubercles in aeronautics comes from the humpback whale (Megaptera novaeangliae) which has a characteristic flipper, with a spanwise scalloped leading edge, creating an almost sinusoidal shape, consisting of bumps called tubercles. The whale uses this layout in order to achieve high underwater maneuverability. Early experimental research showed a great potential in enhancing the 3D aerodynamic characteristics of a wing. Most of the existing experimental results concern infinite wings (2D) models and are accompanied with substantial loss in lift and increase in drag in pre-stall region. On the other hand, 3D finite models have displayed a better overall aerodynamic performance (increased lift and moment, but also, decreased drag). At a range of Reynolds number between 500,000 and 1,000,000 (based on the mean chord of the flipper), tubercles act as virtual fences, introducing a pair of counter rotating vortices that delays the stall of the flipper, a phenomenon that the whales exploit to perform sharp turns and catch their prey. The aforementioned Reynolds number range is the same as the operational Reynolds number for typical unmanned aerial vehicles. To assess the influence of the tubercles installation on UAV wings, a full 3D computational study is carried-out with the use of CFD tools which at a first phase are validated and calibrated with the available literature experimental data. Then, computations are performed for different spanwise tubercles distributions. The results show that there is a noticeable potential on controlling the flow on the wings of a UAV operating in a Reynolds number range between 500,000 and 1,000,000 (based on UAV’s wing mean chord), which can lead to an aerodynamic performance and efficiency increase.

Modal Identification of Aerodynamic Models for Spanwise Lift Distribution Control

This work presents modal identification results for the unsteady three-dimensional flow around a wing featuring multiple trailing edge flaps. The unsteady spanwise lift distribution is modeled using the unsteady vortex lattice method, which yields a linear, time-invariant, high-order state-space model. The eigensystem realization algorithm is applied for model order reduction and modal identification, providing aerodynamic mode shapes and associated eigenvalues. The wing is considered rigid, so that the dynamic mode shapes associated with the spanwise lift distribution are purely aerodynamic. It is observed that the spatial-temporal discretization of the numerical model and flap dimensions inherently limit the number of modes that can be properly identified. Open-loop simulation results demonstrate the modal behavior of the aerodynamic system. Finally, aerodynamic mode shapes for different wing planforms are also presented for comparison. By expanding the system output (lift distribution) as a truncated superposition of aerodynamic mode shapes, a low-order multiple-input and multiple-output modal representation can be obtained, suitable for controller synthesis. Ultimately, this novel approach will be applied to gust alleviation, in order to keep a desired lift distribution profile using shape control techniques.

Preliminary assessment of surface pressure measurements on a metallic, additive manufactured wind tunnel model

Producing metallic wind-tunnel models with integrated pressure taps (i.e. taps formed during the manufacture process) via additive manufacturing (AM) techniques could present several advantages over conventionally manufacturing (CM) pressure-tapped models. However, there are limited reports of AM wind-tunnel models in general. This study describes, in brief, the design and manufacture of a metallic hemispherical wind-tunnel model with integrated pressure taps, produced by Selective Laser Melting (SLM). Mean and unsteady pressure distributions are acquired as a preliminary performance metric for the AM model. The measurements generally agree with previously published results for hemisphere flows at similar experimental conditions; but discrepancies, thought to arise from the model surface finish rather than the SLM process, are evident in the span-wise pressure distributions. Nevertheless, the results are encouraging and extend the bounds of possibility in the use of AM techniques for wind-tunnel model production.

Experimental validation of numerical decambering approach for flow past a rectangular wing

Experimental investigation on two rectangular wings with NACA0012 and NACA4415 profiles is performed at different Reynolds numbers to understand their aerodynamic behaviours at a high α regime. In-house developed numerical code VLM3D is validated using this experimental result in predicting the aerodynamic characteristics of a rectangular wing with cambered and symmetrical wing profile. The sectional coefficient of lift ([Formula: see text]) obtained from the numerical approach is used to study the variation in spanwise lift distribution. The lift and moment characteristics obtained from wind tunnel experiments are plotted, and change in the maximum coefficient of lift ([Formula: see text]) and stall angle ( α stall) are studied for both of the wing sections. A significant addition to the novelty of the present experiments is to provide some comparison of the numerical induced drag coefficient, [Formula: see text] with experimentally fitted model coefficients using least square technique. A novel method is used to examine the aerodynamic hysteresis at high angles of attack. The area included in the lift- Re curve loop is a measure of aerodynamic efficiency, and its variation with angle of attack and wing plan forms is studied.

Secondary currents and very-large-scale motions in open-channel flow over streamwise ridges

It is widely acknowledged that streamwise ridges on the bed of open-channel flows generate secondary currents (SCs). A recent discovery of meandering long streamwise counter-rotating vortices in open-channel flows, known as very-large-scale motions (VLSMs), raises a question regarding the interrelations between VLSMs and SCs in flows over ridge-covered fully rough beds. To address it, we conducted long-duration experiments using stereoscopic particle image velocimetry, covering a range of ridge spacings ( ) from to flow depths ( ). For a benchmark no-ridge case, the flow is quasi-two-dimensional in the central part of the channel, exhibiting a strong spectral signature of VLSMs, as expected. With ridges on the bed at , two SC cells are formed between neighbouring ridges and VLSMs are entirely suppressed, suggesting that ridge-induced SCs prevent the formation of VLSMs by absorbing their energy or overpowering their formation. At the same time, velocity auto- and cross-spectra reveal a new feature that can be explained by low-amplitude meandering of the alternating low- and high-momentum flow regions associated with instantaneous manifestations of SCs. Two-point velocity correlations and smooth velocity field reconstructions using proper orthogonal decomposition further support the validity of this effect. Its origin is probably due to the instability related to the presence of inflection points in the spanwise distribution of the streamwise velocity within the SC cells. These results have implications for bed friction in open channels, where the friction factor may increase if depth-scale SCs are present, or decrease under conditions of sub-depth-scale SCs and suppressed VLSMs.

Control and Entropy Analysis of Tip Leakage Flow for Compressor Cascade under Different Clearance Sizes with Endwall Suction

To investigate the influence of the change of tip clearance size on the control effect of the endwall suction, the effects of endwall suction on the aerodynamic performance of the axial compressor cascade were studied numerically. Three tip clearance sizes of 0.5% h, 1.0% h, and 2.0% h (h is the blade height) were mainly considered. The results show that the endwall suction scheme whose coverage range was 8–33% axial chord can reduce the leakage flow and improve the aerodynamic performance by directly influencing the structure of tip leakage vortex. The overall total pressure loss coefficients of the three clearance size schemes at 0° angle of incidence with 0.4 inlet Mach number are reduced by about 10.3%, 10.8%, and 6.0%, respectively, at the suction flow rate of 0.7%. Under the same suction flow rate, the onset position of the tip leakage vortex of the cascade with small clearance is shifted from the 15% of the axial chord length of original to the 48% of the axial chord length, which with large clearance is nearly no changed. The leakage flow rate and the distance from the leakage vortex to the suction slot are the main reasons for the different control effect of the endwall suction under different tip clearance sizes. The difference of the spanwise distribution of flow field parameters may also cause the difference of flow control effect.

A Probabilistic Uncertainty Estimation Method for Turbulence Parameters Measured by Hot-Wire Anemometry in Short-Duration Wind Tunnels

Abstract A robust and complete uncertainty estimation method is developed to quantify the uncertainty of turbulence quantities measured by hot-wire anemometry (HWA) at the inlet of a short-duration turbine test rig. The uncertainty is categorized into two macro-uncertainty sources: the measurement-related uncertainty (the uncertainty of each instantaneous velocity sample) and the uncertainty stemming from the statistical treatment of the time series. The former is addressed by the implementation of a Monte Carlo (MC) method. The latter, which is directly related to the duration of the acquired signal, is estimated using the moving block bootstrap (MBB) method, a nonparametric resampling algorithm suitable for correlated time series. This methodology allows computing the confidence intervals of the spanwise distributions of mean velocity, turbulence intensity, length scales, and other statistical moments at the inlet of the turbine test section.

Effects of heterogeneous surface geometry on secondary flows in turbulent boundary layers

The effect of spanwise heterogeneous surface geometry on turbulent boundary layer secondary flows and on skin friction is investigated experimentally. The surfaces consist of smooth streamwise-aligned ridges of different shapes and widths with spanwise wavelengths comparable to the boundary layer thickness ( ). Cross-stream stereoscopic particle image velocimetry combined with oil-film interferometry is used to investigate the flow field and assess the drag. Results show that the spanwise distribution of the skin friction varies as a consequence of the mean flow heterogeneity and is highly dependent on surface geometry. The swirling strength maps revealed remarkable changes in the secondary flow structures for different ridge shapes. For wide ridges, topological changes occur showing the appearance of tertiary vortices coexisting with the large-scale secondary structures. An imbalance in favour of these tertiary structures occurs over a certain width which take over the secondary structures, causing a swap in the locations of the low- and high-momentum pathways. Furthermore, the results indicate that the spanwise spacing alone is insufficient to characterise the surface heterogeneity. A new parameter ( ), which is based on the ratio of the perimeter over and below the mean surface height, is shown to adequately capture the changes in skin friction and streamwise circulation of the secondary motions. Triple decomposition allowed the quantification of the dispersive stresses for all these cases, which can contribute up to of the total shear stress when strong secondary motions occur.

The exploitation of CFD legacy for the meridional analysis and design of modern gas and steam turbines

In the last decades, the consolidation of 3D CFD approaches in the industrial design practices has progressively moved throughflow codes from the top of design systems to somewhere in between first development stages and the final aerodynamic optimizations. Despite this trend and the typical limitations of traditional throughflow methods, designers tend to still consider such methods as fundamental tools for drafting a credible aero-design in a short turnaround time. Recently a considerable attention has been devoted to CFDbased throughflow codes as suitable means to widen the range of applicability of these tools while smoothing the predictive gap with successive threedimensional flow analyses.The present paper retraces the development and some applications of a modern and complete CFD-based throughflow solver specifically tuned for multistage axial turbine design. The code solves the axisymmetric Euler equations with an original treatment of tangential blockage and body force. It inherits its numerical scheme from a state-of-the-art CFD solver (TRAF code) and incorporates real gas capabilities, three-dimensional flow features (e.g. secondary flows, tip leakage effects), coolant flow injections, and radial mixing models. Also geometric features of actual blades, like fillets, part-span shrouds, and snubbers, are accounted for by suitable models.The capabilities of the code are demonstrated by discussing a significant range of test cases and industrial applications. They include single stage configurations and entire multistage modules of steam turbines, with flow conditions ranging from subsonic to supersonic. Computational strategies for design and off-design analyses will be presented and discussed. The reliability and accuracy of the method is assessed by comparing throughflow results with 3D CFD calculations and experimental data. A good agreement in terms of overall performance and spanwise distributions is achieved in both design and off-design operating conditions.

The controlled periodic impact on the longitudinal vortex in the boundary layer at Mach 2

Experimental data on the development of controlled disturbances in the boundary layer with a longitudinal vortex are analysed. It is found experimentally that the mean flow nonuniformity can vary from 0.7 to 0.9 V in terms of the CTA mean voltage or 30–40% in ρU. Such flow nonuniformity complicates the estimation of the wave spectra, where normalization on the mean voltage of the anemometer is required. The applicability of the standard processing procedure is investigated. Comparative analysis of the data shows that both normalization types considered in the work can be applicable. However, in the case of flow nonuniformity in the transverse direction, it is preferable to use the normalization on the median value over the mean voltage spanwise distribution.

Design approaches of performance-scaled rotor for wave basin model tests of floating wind turbines

The Froude scaling law is usually utilized in the wave basin model tests of Floating Wind Turbines (FWTs). However, the Froude-Scaled Rotor (FSR) cannot generate desired aerodynamic loads due to the Reynolds-Number Scaling Effect (RNSE). To mitigate the adverse effects of RNSE, two approaches are proposed to design Performance-Scaled Rotors (PSRs) in this paper. Taking DTU 10 MW baseline wind turbine as an example, the SD2030 airfoil is selected to replace the original FFA-W3-xx airfoils. Maximum Lift Tracking (MLT) and Load Distribution Matching (LDM) algorithms are proposed to assign the chord lengths and twist angles. Herein, MLT leads all airfoils to operate at the optimal angle of attack that corresponds to the maximum lift coefficient and afterwards increasing the chord lengths. LDM simultaneously adjusts the chord length and twist angle, aiming to match the span-wise distribution of normal force at the design point. Results show that both approaches can generate desired rotor thrusts in a range of tip speed ratios, which seems to outperform prior PSRs in the existing publications. The blade mass and inertia can be preserved with careful manufacturing procedures. The redesigned PSRs are helpful to improve the accuracy and reliability of FWT model tests in the wave basin.

Reducing Secondary Flow Losses in Low-Pressure Turbines: The “Snaked” Blade

This paper presents an innovative design for reducing the impact of secondary flows on the aerodynamics of low-pressure turbine (LPT) stages. Starting from a state-of-the-art LPT stage, a local reshaping of the stator blade was introduced in the end-wall region in order to oppose the flow turning deviation. This resulted in an optimal stator shape, able to provide a more uniform exit flow angle. The detailed comparison between the baseline stator and the redesigned one allowed for pointing out that the rotor row performance increased thanks to the more uniform inlet flow, while the stator losses were not significantly affected. Moreover, it was possible to derive some design rules and to devise a general blade shape, named ‘snaked’, able to ensure such results. This generalization translated in an effective parametric description of the ‘snaked’ shape, in which few parameters are sufficient to describe the optimal shape modification starting from a conventional design. The “snaked” blade concept and its design have been patented by Avio Aero. The stator redesign was then applied to a whole LPT module in order to evaluate the potential benefit of the ‘snaked’ design on the overall turbine performance. Finally, the design was validated by means of an experimental campaign concerning the stator blade. The spanwise distributions of the flow angle and pressure loss coefficient at the stator exit proved the effectiveness of the redesign in providing a more uniform flow to the successive row, while preserving the original stator losses.

Validation of a Model for Estimating the Strength of a Vortex Created from the Bound Circulation of a Vortex Generator

A hypothesis was tested and validated for predicting the vortex strength induced by a vortex generator in wall-bounded flow by combining the knowledge of the Vortex Generator (VG) geometry and the approaching boundary layer velocity distribution. In this paper, the spanwise distribution of bound circulation on a vortex generator was computed by integrating the pressure force along the VG height, calculated using Computational Fluid Dynamics (CFD). It was then assumed that all this bound circulation was shed into a wake to fulfill Helmholtz’s theorem which then curls up into one primary tip vortex. To validate this, the trailed circulation estimated from the distribution of the bound circulation was compared to the one in the wake behind the vortex generator, determined directly from the wake velocities at some downstream distance. In practical situations, the pressure distribution on a vane is unknown and consequently other estimates of the spanwise force distribution on a VG must instead be applied, such as using 2D airfoil data corresponding to the VG geometry at each wall-normal distance. Such models have previously been proposed and used as an engineering tool to aid preliminary VG design. Therefore, it is not the purpose of this paper to refine such engineering models, but rather to validate their assumptions, such as applying a lifting line model on a VG that has a very low aspect ratio and is placed in a wall boundary layer. Herein, high Reynolds number boundary layer measurements of VG-induced flow were used to validate the Reynolds-Averaged Navier–Stokes (RANS) model circulation results, which were used for further illustration and validation of the hypothesis.