Trailing Edge Research Articles

Investigation of Reynolds number effects on flow stability of a transonic compressor using a dynamic mode decomposition method

With the continuous increase in aircraft flight altitude, low Reynolds number effects in high-altitude environments significantly impact the stability of compressor operation. In this study, dynamic mode decomposition (DMD) of the flow field at various time instances near stall conditions in a transonic compressor was performed to investigate the influence of Reynolds number (Re) effects on critical flow characteristics and instability mechanisms. The results indicate that under different Reynolds number index (RNI) conditions near stall conditions, the frequency of tip leakage flow varies. Under high RNI condition, the influence of tip leakage flow is concentrated mainly from the leading edge to the mid-chord. The stall mode is primarily caused by the interaction of shock with the Tip Leakage Vortex, resulting in the breakdown of the vortex. The development of vortex structures inside the passage leads to a delay in the flow field response to blade passing. Under low RNI condition, the tip clearance leakage flow is stronger, with a relatively reduced influence from the leading edge to the mid-chord and an increased impact from the mid-chord to the trailing edge. The main cause of the stall mode is the circumferential movement of vortex structures, causing secondary leakage over adjacent blade tip clearances. The blockage formed by vortex structures within the passage is weaker, and the flow field response to blade passing remains consistent with the blade passing frequency. However, as stall develops further, the flow field response significantly decreases.

Information transfer between tip leakage vortex and blade aerodynamic force of a compressor cascade

The compressor blade exhibits intricate flow patterns near stall conditions, with the tip leakage vortex and the main flow shedding vortex from the blade's trailing edge identified as factors influencing the blade's aerodynamic force oscillation. The tip leakage vortex is recognized as the primary contributor to flow structures during non-synchronous vibration (NSV), although existing compressor NSV reduced order models often attribute the fluid oscillation to the passage flow shedding vortex. A tool capable of distinguishing between flow structures and blade aerodynamic force is crucial for resolving this discrepancy. Detached eddy simulations (DES) in a compressor cascade form the basis for gathering data on vortex structures and blade aerodynamic force. This study employs an information thesis method to elucidate the cause-and-effect relationship between the tip leakage vortex and blade aerodynamic force. Through a series of unsteady single passage simulations and experimental validation, the mechanisms leading to variations in blade aerodynamic force are demonstrated. Notably, this research utilizes Proper Orthogonal Decomposition (POD) mode time coefficients in span and axial directions to represent dynamic information on vortex structures. Additionally, transfer entropy is introduced as a metric to evaluate the causal interaction between turbulent flow of the tip leakage and the blade's aerodynamic force, marking a significant advancement. The analysis of information transfer entropy aids in identifying the vortex structure with a stronger relationship to aerodynamic force, a critical finding. By applying this approach, the study examines tip leakage vortex structures in various tip clearance sizes concerning aerodynamic force. The Tip3C and Tip5C tip geometries exhibit tip leakage vortex features with radial vortex characteristics and backflow, akin to structures associated with non-synchronous vibration. These tip vortex structures in the Tip3C and Tip5C tip geometry cascades demonstrate substantial correlations with aerodynamic force oscillation on the blade surface. Conversely, the Tip1C tip geometry cascade displays distinct tip leakage vortex features compared to Tip3C and Tip5C, with weaker ties to blade aerodynamic force oscillation. As simulation models progress, consideration of vortex structures in the passage and leakage vortex at the leading edge is imperative for accurate predictions of aerodynamic force oscillation response.

Supercooled large droplet size distribution effects on airfoil icing: A numerical investigation based on a new coupled Eulerian method

Supercooled large droplets (SLDs) under natural icing conditions have the characteristics of easy deformation in motion and easy splashing on impact, and a bimodal droplet size distribution that has received less attention. The modified form of the Rosin-Rammler function was improved to achieve a more accurate nonlinear fitting of the SLD distribution curve. The droplet size distribution was divided into non-equipartition continuous multiple components. The drag source term of each component was coupled with the overall droplet size distribution, and the Eulerian equations of each component of SLDs were solved simultaneously. A new coupled Eulerian method for non-equipartition continuous multi-size droplets was proposed to simulate the impact characteristics of SLDs, and the SLD collection coefficients were validated. Effects of the ratio between the number of large and small droplet components and the number of all components on the simulation results were investigated to select a better combination based on stable convergence calculation steps and the calculation time. This new method was added to the multi-step icing numerical method, and the accuracy and robustness of the method in icing shape prediction were verified based on the freezing drizzle, median volume diameter < 40 μm (FZDZ, MVD < 40 μm) icing condition. Airfoil icing characteristics based on the bimodal and monomodal distribution were compared, and the icing shapes at the leading edge were similar. Still, the upper and lower limits of the icing shapes with the bimodal distribution were nearer to the trailing edge and the ice layer was thicker there.

The impact of non-frozen turbulence on the modelling of the noise from serrated trailing edges

Serrations are commonly employed to mitigate the turbulent boundary layer trailing-edge noise. However, significant discrepancies persist between model predictions and experimental observations. In this paper, we show that this results from the frozen turbulence assumption. A fully developed turbulent boundary layer over a flat plate is first simulated using the large-eddy simulation method, with the turbulence at the inlet generated using the digital filter method. The space–time correlations and spectral characteristics of wall pressure fluctuations are examined. The simulation results demonstrate that the coherence function decays in the streamwise direction, deviating from the constant value of unity assumed in the frozen turbulence assumption. By considering an exponential decay function, we relax the frozen turbulence assumption and develop a prediction model that accounts for the intrinsic non-frozen nature of turbulent boundary layers. To facilitate a direct comparison with frozen models, a correction coefficient is introduced to account for the influence of non-frozen turbulence. The comparison between the new and original models demonstrates that the new model predicts lower noise reductions, aligning more closely with the experimental observations. The physical mechanism underlying the overprediction of the noise model assuming frozen turbulence is discussed. The overprediction is due to the decoherence of the phase variation along the serrated trailing edge. Consequently, the ratio of the serration amplitude to the streamwise frequency-dependent correlation length is identified as a crucial parameter in determining the correct prediction of far-field noise.

Experimental and numerical investigation on low-velocity impact damage and parametric study of composite rotor blades

Modern helicopter rotor blades are made of composite materials. Due to the harsh service environment, the blades may be impacted by hail and other falling objects at low velocity. The damage behavior of composite rotor blades subjected to low-velocity impact is studied here. Low-velocity impact experiment is conducted on a composite helicopter rotor blade, and the impact force and the damage are measured. Based on the experiment, a finite element model is developed and validated to predict the impact damage of the blade. With the model, a parametric study is performed to investigate the effects of impact energy, impact locations, and foam core on impact responses of the composite blades. The results show that the rib is the most vulnerable part to damage, associated with severe deformation, and delamination rapid expanding under low-velocity impact. The foam crush region is close to the blade top surface below the impact point. The interface between the skin and internal part is especially sensitive to increase in impact energy. Impacts near the leading edge produce higher impact force, while impacts near the trailing edge cause large deformation and small impact force. Better impact performance will be achieved when modulus and density of the foam are increased.

Thirteen vital factors for micro-scale radial turbine vane’s design of geo-solar-powered Brayton cycle applications

The ever-elevating importance of turbines makes enhancing their performance a crucial point. In most cases, the required performance can be achieved by improving the turbine’s components, i.e., its rotary part, stationary part, and volute. However, this requires a lot of effort, especially when dealing with rotary parts that are time-consuming and costly because of the rotor. Having said that, most of the previous studies have improved turbomachinery by using the latter method. The current study, on the other hand, highlights the importance of only one component, the cheaper one, as a key factor. So, once the numerical modelling was validated using a previous experimental published work, fourteen parameters, which represented all the aerofoil’s vane shape (the stationary part of the turbine only) of a micro-scale turbine i.e. 2–12 kW, were accurately studied with the aim of highlighting their influence on both the vane performance, stage operation, and the overall cycle’s performance as well. The results showed that while parameter 14 i.e. Linear Point Advanced Side 2 showed almost no influence and parameters 5 and 6 i.e. Trailing Edge Wedge Angle and Trailing Edge Beta Angle were the most influential ones, all the other parameters showed a significant effect on the overall turbines’ performance, especially at higher PR values. Moreover, the results have also confirmed that the best values of the mentioned parameters are 0.33 mm, 17 Deg., −25 Deg., 41 Deg., 1 Deg., 55 Deg., 0.3 mm, 0 Deg., 19 mm, 31 mm, 77 mm and 0 mm for all thirteen parameters, respectively. This results in enhancing the stator and the turbine efficiency by around 11 % and 5 % respectively, compared to the baseline design.

Jet Installation Noise Modelling for Round and Chevron Jets

Wall-Modelled Large Eddy Simulations (LES) are conducted using a high-resolution CABARET method, accelerated on Graphics Processing Units (GPUs), for a canonical configuration that includes a flat plate within the linear hydrodynamic region of a single-stream jet. This configuration was previously investigated through experiments at the University of Bristol. The simulations investigate jets at acoustic Mach numbers of 0.5 and 0.9, focusing on two types of nozzle geometries: round and chevron nozzles. These nozzles are scaled-down versions (3:1 scale) of NASA’s SMC000 and SMC006 nozzles. The parameters from the LES, including flow and noise solutions, are validated by comparison with experimental data. Notably, the mean flow velocity and turbulence distribution are compared with NASA’s PIV measurements. Additionally, the near-field and far-field pressure spectra are evaluated in comparison with data from the Bristol experiments. For far-field noise predictions, a range of techniques are employed, ranging from the Ffowcs Williams–Hawkings (FW–H) method in both permeable and impermeable control surface formulations, to the trailing edge scattering model by Lyu and Dowling, which is based on the Amiet trailing edge noise theory. The permeable control surface FW–H solution, incorporating all jet mixing and installation noise sources, is within 2 dB of the experimental data across most frequencies and observer angles for all considered jet cases. Moreover, the impermeable control surface FW–H solution, accounting for some quadrupole noise contributions, proves adequate for accurate noise spectra predictions across all frequencies at larger observer angles. The implemented edge-scattering model successfully captures the mechanism of low-frequency sound amplification, dominant at low frequencies and high observer angles. Furthermore, this mechanism is shown to be effectively consistent for both M=0.5\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$M=0.5$$\\end{document} and M=0.9\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$M=0.9$$\\end{document}, and for jets from both round and chevron nozzles.

Hybrid LES/RANS Simulations of Compressible Flow in a Linear Cascade of Flat Blade Profiles

The paper reports on 3D numerical simulations of unsteady compressible airflow in a blade cascade consisting of flat profiles using a hybrid LES/RANS approach including a transition model. As a first step towards simulation of blade flutter in turbomachinery, various incidence angle offsets of the middle blade were modeled. All simulations were run for the flow regime characterized by outlet isentropic Mach number Mis=0.5 and zero incidence. The results of the LES/RANS simulations (pressure and Mach number distributions) were compared to a baseline RANS model, and to experimental data measured in a high-speed wind tunnel. The numerical results show that both methods overpredict flow separation taking place at the leading edge. In this regard, the hybrid LES/RANS method does not provide superior results compared to the traditional RANS simulations. Nevertheless, the LES/RANS results also capture vortex shedding from the blunt trailing edge. The frequency of the trailing edge vortex shedding in CFD simulations matches perfectly the spectral peak recorded during wind tunnel measurements.

Numerical investigation of elliptical trailing edge on cascade loss of transonic turbine airfoils

Abstract The shape and thickness of the trailing edge are important factors in the supersonic loss in transonic cascades. To investigate the influence of trailing edge shape, a prismatic turbine cascade and a three-dimensional cascade with an elliptic trailing edge were analyzed with numerical simulation by changing different elliptic axis ratios ER. It was indicated that the shock wave and its mixing with the wake were the important sources of loss in the transonic prismatic cascade. With the increase of exit Mach number, shock waves are generated rapidly and migrate downstream, and the shock loss and the total loss of the prismatic cascade are both accrescent. Increasing the elliptical axis ratio could weaken shock waves and delay shock waves. In the loss state region 3, the cascade loss of the ER3 blade is reduced by 8% compared with the round trailing edge cascade. As the elliptical axis ratio of the three-dimensional cascade is increasing, the profile loss and secondary loss are reduced. However, an excessive elliptical axis ratio could lead to an increase in tip leakage loss. Within the range of ER values from 0 to 3, the total loss of the transonic three-dimensional cascade reached minimization when ER was 3.0.

Damage behavior of cold-rolled sintered fiber felts

Porous Materials, such as sintered fiber felts (SFFs), can contribute to the reduction of aircraft noise by acting as a low-noise trailing edge (TE). A rolling process with a time-varying rolling gap was sucessfully used to tailor the aeroacoustic properties of SFFs. To ensure the suitability of rolled porous materials for application, the mechanical properties after rolling must be investigated. The objective of this study was to clarify the effect of the rolling process on the damage evolution during tensile loading. A SFF consisting of a functional layer and a support grid made of alloy 1.4404 was used for rolling. Interrupted tensile tests in combination with computed tomography (CT) were used to investigate the damage evolution of as-received material, uniformly rolled material and material rolled with a gradient in thickness reduction. Metallographic examination and fracture surface analysis using scanning electron microscopy (SEM) complemented the CT analysis. Uniformly rolled material with a low degree of deformation (−0.74 ≤ φ < 0) failed identically to the as-received material. The functional layer failed first, starting from the center. The support grid subsequently failed due to the incremental failure of individual wires. The material rolled with a gradient in thickness reduction failed accordingly. A significant change in damage evolution occured for material rolled at high degrees of deformation φ ≤ −1.53, where the support grid fails first and the functional layer second. The rolling process using a time-varying rolling gap does not result in a detrimental mechanical behavior under tensile load. The results improve the general understanding of the effects of rolling processes on the properties of porous materials and allow the damage behavior to be taken into account with regard to its application as a TE.

Deep-Learning-based damage detection of an experimentally tested full-scale wind turbine blade.

Reducing the Operation and Maintenance (O&amp;M) costs of Wind Farms can substantially contribute towards adopting environmentally sustainable and cost effective energy sources. O&amp;M costs in Wind Turbines can account up to 50% of the Turbine’s whole life cycle, with the majority of it being the frequent manual inspection needed for Wind Turbine Blades (WTB). Vibration-based structural health monitoring (SHM) approaches can be used for detecting WTB damages in an early state, before they lead to catastrophic failure, while they can be used to perform inspection on an informed basis, reducing O&amp;M costs. In the present study, a data-driven damage detection framework is developed for WTBs. A Finite Element Model (FEM) of a full-scale WTB (the Aventa AV-7 “Light Wind Turbine”, sold under the Leichtwindanlagen© name) is developed in ABAQUS FEA using sparce material properties and dimensions and subsequently modified to match a physical WTB. Damage is modelled in the form of a Trailing Edge (TE) longitudinal crack propagated in a stepwise manner. The FE model is updated based on the geometry altering and material property degrading effects of the TE crack. Numerical simulations are performed to obtain the dynamic responses of the blade, under white noise and chirp excitation. A damage detection framework is then developed using Auto-Regressive, with exogenous excitation (ARX) or without (AR), models and its Nonlinear counterparts NARX and NAR models. In addition, model dimensionality reduction is performed using Principal Component Analysis (PCA) and Deep-Autoencoders (D-AE). The results are presented in terms of Receiver Operating Characteristics (ROC) curves.

Analytic Prediction of Rotor Broadband Noise with Serrated Trailing Edges with Applications to Urban Air Mobility Aircraft

Theoretical predictions for broadband noise generated by rotors with serrated trailing edges are presented. This study extends Lyu and Ayton’s model for serrated trailing-edge noise to account for factors such as finite span length, the accurate acoustic attenuation rate of trailing-edge noise, modified spanwise wavenumber, and the proper scattering factor. This new model is implemented in the rotorcraft broadband noise tool, UCD‐QuietFly. The validation results demonstrate a satisfactory level of agreement with the experimental data for serrated trailing‐edge noise of a wing section, a hovering rotor, and a forward‐flight rotor. Next, the effects of serration parameters, such as the number of serrations, serrations starting radial location, serration geometry, and serration height, on rotor broadband noise are investigated. For example, it is observed that sine-wave, chopped peak, or sawtooth serration shapes provide the most significant noise reduction among the various shapes considered. Finally, broadband noise reduction through serrations is applied to two urban air mobility aircraft scenarios: a six‐passenger quadrotor and a quiet helicopter. For the quadrotor, serrated trailing edges show a noise reduction potential of over 9 dB, with higher reductions observed at higher tip speeds and fewer blades. A quiet helicopter equipped with serrated blades in forward flight achieves a maximum noise reduction of over 10 dB in the forward direction.

Energy loss evaluation of a radial inflow turbine for organic Rankine cycle application using hierarchical entropy production method

To evaluate the location and main sources of energy loss in radial inflow turbines for organic Rankine cycle application, this study proposed a hierarchical entropy production method, which is superior to the traditional pressure drop method. The method includes three levels: local entropy production, split of turbine total entropy production, and split of component entropy production. The energy loss of the radial inflow turbine under design condition and different pressure ratios is presented. The results indicate that the high-entropy production zone is primarily located at the stator trailing edge and the rotor tip clearance. The proportion of turbulent entropy production and wall entropy production in the total energy loss of the turbine is about 77% and 20%, respectively. Among the components of the radial inflow turbine, the energy loss of the rotor and diffuser is the highest, accounting for 71.9% and 13.6% of the total entropy production of the turbine, respectively. However, the stator and rotor have higher volume average entropy generation rate and area average entropy generation rate. The high-entropy production region is mainly located in the stator outlet zone and the rotor tip zone. When the pressure ratio increases from 3 to 5, the turbine efficiency decreases by 13.44%. The pressure ratio has a significant effect on the turbulent entropy production of the rotor. This method can provide insight into the energy loss characteristics of radial inflow turbines for organic Rankine cycle applications.

Analysis of the Effect Variation of Sides Racing Bike Helmets on Flow Characteristics Using CFD

The selection of materials and aerodynamic design forms needs to be considered in designing a comfortable helmet. The emergence of the drag force on the helmet causes a compressive force, which causes the rider discomfort when driving at high speeds. This research was conducted on 4 variations of helmet models consisting of 1 time-trial type helmet as a reference and 3 other helmets which were developed from the basic model. This simulation was conducted in a 2D helmet model at Re 7.14 x 104, 1.00 x 105, and 1.16 x 105. In this study, Turbulence Intensity was varied as 0.1%, 0.2%, and 1%. The results are shown through the drag coefficient's value and visualization of streamlined flow. The results showed that the 2D simulation on the helmet produced a lower error by using Set Up Turbulence Intensity 0.1%. The emergence of the Oscillating Karman vortex in large size indicates the higher the drag force that occurs on the helmet. In this study, helmet type 3, with a trailing edge and has a smooth curve, has the lowest drag force at Re 1.16 x 105 with a drag coefficient value of around 0.38.

Large Eddy Simulation of Laminar and Turbulent Shock/Boundary Layer Interactions in a Transonic Passage

Abstract Shock/boundary layer interactions (SBLI) are a fundamental fluid mechanics problem relevant in a wide range of applications including transonic rotors in turbomachinery. This paper uses wall-resolved large eddy simulation (LES) to examine the interaction of normal shocks with laminar and turbulent inflow boundary layers in transonic flow. The calculations were performed using GENESIS, a high-order, unstructured LES solver. The geometry created for this study is a transonic passage with a convergent-divergent nozzle that expands the flow to the desired Mach number upstream of the shock and then introduces constant radius curvature to simulate local airfoil camber. The Mach numbers in the divergent section of the transonic passage simulate single stage commercial fan blades. The results predicted with the LES calculations show significant differences between laminar and turbulent SBLI in terms of shock structure, boundary layer separation and transition, and aerodynamic losses. For laminar flow into the shock, significant flow separation and low-frequency unsteadiness occur, while for turbulent flow into the shock, both the boundary layer loss and the low-frequency unsteadiness are reduced. The time period of the unsteadiness was hypothesized to align with the time it takes for turbulent structures to convect from the shock to the trailing edge and acoustic disturbances to travel back from the trailing edge to the shock, and this is supported by the LES results.

Towards extending the aircraft flight envelope by mitigating transonic airfoil buffet

In the age of globalization, commercial aviation plays a central role in maintaining our international connectivity by providing fast air transport services for passengers and freight. However, the upper limit of the aircraft flight envelope, i.e., its operational limit in the high-speed (transonic) regime, is usually fixed by the occurrence of transonic aeroelastic effects. These harmful structural vibrations are associated with an aerodynamic instability called transonic buffet. It refers to shock wave oscillations occurring on the aircraft wings, which induce unsteady aerodynamic loads acting on the wing structure. Since the structural response can cause severe structural damage endangering flight safety, the aviation industry is highly interested in suppressing transonic buffet to extend the flight envelope to higher aircraft speeds. In this contribution, we demonstrate experimentally that the application of porous trailing edges substantially attenuates the buffet phenomenon. Since porous trailing edges have the additional benefit of reducing acoustic aircraft emissions, they could prospectively provide faster air transport with reduced noise emissions.

Optimal section design of surface piercing propeller using artificial neural networks algorithm

ABSTRACT Section optimisation of a surface piercing propeller (SPP) with the combination of artificial neural networks and NSGA II algorithm is investigated in this research. The blade section have a sharp leading edge and thick trailing edge. The geometry of this propeller is mostly designed experimentally, and comprehensive studies have not been performed to examine its parameters. In this regard, the propeller's parametric geometry is designed based on the cubic B-spline method. The SPP was first fabricated and tested, and the simulated data of the tested propeller were verified. Then, limited data were extracted according to a verified CFD model based on which the ANN was trained. The ANN results indicated the mean square error of the trained network is 3e-9, 1e-7, and 1e-7 for training, validation, and test data, respectively. comparing the optimisation and experimental results showed a relative error amounting to 3.5% in thrust and 4.24% in torque.

Effects of blowing ratios and lip thicknesses on film cooling for trailing edge cutback with latticework ducts

In this study, a trailing edge cutback configuration with latticework ducts is proposed to provide more effective thermal protection for high-pressure turbine blades. Detached eddy simulation, as a LES/RANS hybrid method, is utilized to investigate the film cooling efficiency and turbulent behavior of the trailing edge cutback. The flow dynamics mechanism of cutback film cooling is elucidated by analyzing five blowing ratios (0.2 ∼ 1.25) and four lip thicknesses (0.5 ∼ 2) schemes. Time-averaged results indicate that the counter-intuitive efficiency decline and recovery behavior is identified again. Quantitatively, the area-averaged film cooling efficiency decreases by 14.3 % in the efficiency decline interval and increases by 16.3 % during the recovery interval. The distinct deflection of the swirling coolant jet leads to an efficiency distribution characterized by higher values on the sides and lower values in the center. Moreover, the mixing of the swirling jets and crossflow produces the Asymmetric Counter-rotating vortex pairs and Spiral-vortex system, which exacerbate heat transfer near the wall. At the highest blowing ratio, the excess coolant is forced against the wall by a rotating streamwise vortex system, providing better film coverage. Additionally, the film cooling efficiency of the cutback structure with a lip-to-slot ratio of 2 decays by 27 % as the blowing ratio decreases. The detrimental influence of the lip thickness increase on the film cooling efficiency is considerably mitigated by the swirling jet, especially in the maximum lip thickness.

Tree rings reveal the transient risk of extinction hidden inside climate envelope forecasts

Given the importance of climate in shaping species' geographic distributions, climate change poses an existential threat to biodiversity. Climate envelope modeling, the predominant approach used to quantify this threat, presumes that individuals in populations respond to climate variability and change according to species-level responses inferred from spatial occurrence data-such that individuals at the cool edge of a species' distribution should benefit from warming (the "leading edge"), whereas individuals at the warm edge should suffer (the "trailing edge"). Using 1,558 tree-ring time series of an aridland pine (Pinus edulis) collected at 977 locations across the species' distribution, we found that trees everywhere grow less in warmer-than-average and drier-than-average years. Ubiquitous negative temperature sensitivity indicates that individuals across the entire distribution should suffer with warming-the entire distribution is a trailing edge. Species-level responses to spatial climate variation are opposite in sign to individual-scale responses to time-varying climate for approximately half the species' distribution with respect to temperature and the majority of the species' distribution with respect to precipitation. These findings, added to evidence from the literature for scale-dependent climate responses in hundreds of species, suggest that correlative, equilibrium-based range forecasts may fail to accurately represent how individuals in populations will be impacted by changing climate. A scale-dependent view of the impact of climate change on biodiversity highlights the transient risk of extinction hidden inside climate envelope forecasts and the importance of evolution in rescuing species from extinction whenever local climate variability and change exceeds individual-scale climate tolerances.

Aeroacoustic source localization using the microphone array method with application to wind turbine noise

The deconvolution DAMAS algorithm can effectively eliminate the misconceptions in the usually-used beamforming localization algorithm, allowing for a more accurate calculation of the source location as well as the intensity. When solving a linear system of equations, the DAMAS algorithm takes into account the mutual influence of different locations, reducing or even eliminating sidelobes and producing more accurate results. This work first introduces the principles of the DAMAS algorithm. Then it applies both the conventional beamforming algorithm and the DAMAS algorithm to simulate the localization of a single-frequency source from a 1.5 MW wind turbine, a complex line source with the text “UCAS” and a line source downstream of an airfoil trailing edge. Finally, the work presents experimental localization results of the source of a 1.5 MW wind turbine using both the conventional beamforming algorithm and the DAMAS algorithm.