Trailing Edge Research Articles

The flow control mechanism of trailing-edge Gurney flap on a 50°-swept delta wing in forced pitching

The flow control effect of the trailing-edge Gurney flap (TG) on the dynamic lift characteristics for a 50°-swept delta wing during large-amplitude pitching oscillations at various reduced frequencies (k = 0.072, 0.144, 0.287, and 0.575) was investigated via force, particle image velocity, and dye visualization measurements in a water channel facility. Numerical simulations were carried out to further understand the flow control mechanism of the TG in low and high reduced frequency cases (k = 0.072 and 0.575). It was found that as the reduced frequency increases, the lift increments brought by the TG are magnified and abated during the upstroke and downstroke processes, respectively. The breakdown of the leading-edge vortex (LEV) on the upper surface of the wing is promoted by the TG during the early stage of the pitching cycle. The lift enhancement being benefited by the TG is mainly contributed by the recovery of lower surface pressure along the trailing edge due to the blockage effect of TG, which also stimulates the spanwise flow and strengthens the LEV upon the upper surface. The significant lift increment contribution of the upper surface during the upstroke process can be maintained to higher angle of attack as the reduced frequency increases.

Analysis of the effect of cavitation on internal fluid excitation characteristics of centrifugal pumps

To explore the influence of cavitation on the internal fluid excitation characteristics of pumps, numerical simulations and performance testing evaluations were performed on the IS65-50-125 centrifugal pump. The prototype pump's exterior characteristic and cavitation performance curves, as well as its bubble volume distribution, were successfully replicated using numerical computations. The effect of cavitation on the internal pressure pulsation characteristics of the centrifugal pump under various operating situations was comprehensively investigated, indicating a relationship between the degree of cavitation and the root mean square values of pressure pulsation. Special emphasis was placed on the changes in features at intermediate and high frequencies, as well as the processes of rising bubble volume and vortex shedding at the impeller trailing edge on pressure pulsation. To validate the simulation results, a centrifugal pump vibration and noise testing platform was built, and studies on vibration intensity and internal sound field noise were conducted. The experimental results revealed that the vibration intensity and internal sound field sound pressure level of the centrifugal pump rose as cavitation conditions deteriorated, confirming the modeling results. This study's significant innovation is the precise identification of the pump's performance changes under different operating conditions by monitoring pressure pulsation changes at various frequencies, as well as an in-depth discussion of the impact mechanism of cavitation phenomena on the internal fluid excitation behavior of centrifugal pumps. The study demonstrates differences in pressure pulsation characteristics on the suction and pressure sides under various cavitation situations, as well as the process of vortex creation and shedding generated by bubbles in the impeller input channel during severe cavitation. This gives new theoretical basis for pump vibration and noise reduction, as well as significant improvements in centrifugal pump performance and stability.

An effective simulation scheme for the prediction of aerodynamic environment under hypersonic conditions characterized by NACA0012

Currently, aerodynamic environment prediction research into scramjet-propelled vehicles characterized by NACA0012 under hypersonic conditions is relatively sparse. Two-dimensional external flow field models are established, and then through validation tests, we perform a systematic investigation between simulation parameters and prediction accuracy, and an effective aerodynamic environment prediction simulation scheme under hypersonic conditions is proposed. Unlike under incompressible conditions, the maximum accuracy decline could be attributed to the inappropriate choice of the sharp trailing edge modeling method, but the definition formula is still preferred. In particular, for the two modeling data point sources, Airfoil tools and NACA4, the numerical performance of the latter is better than the former, and the calculation accuracy negatively correlates with the number of data points offered by both of them. Moreover, for the mesh cells near the shock, the cell Reynolds number and aspect ratio values should be no smaller than 16 and not exceed 380, respectively, and the recommended values for the far field distance, the turbulence model and flux type are 16L, Spalart-Allmaras, and ROE flux type. Under hypersonic conditions, the aerodynamic environment characterized by NACA0012 predicts a maximum temperature of approximately 1856.85 °C, with an average temperature change rate of 77 °C/s. Meanwhile, the top sound pressure level and the vibration acceleration could reach up to 145 dB and 182 g, respectively.

Aero-propulsion coupling effects and installation characteristics of a distributed propulsion system

The ducted fan, as a quintessential propulsion device utilized in Distributed Electric Propulsion (DEP) aircrafts, exhibits a significant influence on the aero-propulsion coupling effects due to its installation characteristics on the wing. In this paper, we employed steady-state numerical analysis techniques in conjunction with actuator disk models and overset mesh to investigate aero-propulsion coupling effects and installation characteristics of a DEP configuration with 5 ducted fans distributed along the wing's trailing edge. We compared various conditions to the wing baseline airfoil under hover condition and cruise condition of the fan tip speed ratio λ from 1.44 to 4.32 and fan installation angle δT from 0 to 90°. The results indicate that DEP system affects the flow over the wing downstream of the maximum thickness point of the baseline airfoil profile at an installation angle of 0. A higher tip speed ratio the flow is accelerated in the upper surface region near the wing's trailing edge, whereas the low energy fluid fails to be accelerated at lower λ, instead forming a local flow blocking. With the increase of the fan installation angle, the overall system lift increases, but the thrust decreases, and both trends increase with the increase of the tip speed ratio. Specifically, the thrust provided by the distributed fans is reduced, and the inlet becomes one of the drag sources when the installation angle is high.

Enhanced cooling performance of blade tip slot cooling: Effect of slot open length

Addressing the critical challenge of thermal load-induced damage in turbine blade tips, this study aims to elucidate the effects of slot open length on film cooling effectiveness (FCE) across turbine blade tips with Pressure Side (PS) and Suction Side (SS) slots considering the application of additive manufacturing to turbine blade tips. Employing a linear cascade setup and the pressure sensitive paint (PSP) technique, we examined and quantified cooling performance across turbine blade tips with PS or SS slots at various slot open lengths. We revealed that while short slot open lengths enhance FCE at the leading edge (LE) due to increased averaged blowing ratios, they fail to adequately cover the trailing edge (TE) region. This decrease in cooling performance is largely attributable to the mixing of coolant with hot gas. In contrast, longer slot open lengths achieve a consistent FCE distribution across the blade tip despite a slight decrease in FCE at the LE region. The study suggests that integrating additional thermal design strategies, such as partition walls with slots, could effectively manage hot gas behaviors, thus enhancing the overall cooling design of turbine blade tips. This research makes a significant contribution to the field of turbine blade tip cooling design, providing insights into optimizing cooling performance through advanced slot configurations enabled by additive manufacturing technologies.

Adaptive control method for morphing trailing-edge wing based on deep supervision network and reinforcement learning

The morphing trailing-edge wing can adaptively change its shape according to the different flight conditions, improving cruising efficiency. The control of the morphing trailing edge is a high-dimensional variable and nonlinear problem, facing the difficulty of the high training cost and the difficulty in real-time continuous control. To overcome these difficulties, a novel adaptive control method based on the couple of Deep Supervision Net (DSN) and Deep Reinforcement Learning (DRL) is proposed to solve a control problem of the morphing trailing-edge wing. DSN uses distributed loads to improve aerodynamic prediction accuracy compared to the traditional surrogate models, while DRL conducts real-time control and improves the performance at non-design points compared to the traditional multi-design points optimization. After control, the lift-to-drag ratio at the start point can be increased by 1.36%, and can be increased by 2.61% at the end point tested on a commercial wide-body aircraft model.

A developed fatigue analysis approach for composite wind turbine blade adhesive joints using finite-element submodeling technique

The intricate shapes and complex structures of composite wind turbine blades pose significant challenges in conducting local-fatigue simulation analyses. Moreover, adhesive joints, commonly found in composite wind turbine blades, are often neglected in current finite-element methods, leading to a dearth of empirical evidence to substantiate design enhancements for these blades. To tackle these issues, this study introduces a fatigue analysis methodology based on a finite-element submodeling approach to investigate the local fatigue behavior of composite wind turbine blades, with particular emphasis on the adhesive joints. The research begins with a comprehensive fatigue simulation analysis of composite wind turbine blades to establish boundary conditions for subsequent submodeling. A specialized submodel focusing on the trailing edge, a common adhesive joint, is constructed with the inclusion of extra adhesive regions for a more accurate depiction of their actual configuration. This submodel is subsequently utilized to investigate the localized fatigue characteristics unique to the trailing edge. Moreover, an exploration into the influence of the submodeling technique and its parameters on the predicted outcomes is carried out to enhance the applicability of the proposed methodology. These results significantly advance the understanding of local static and fatigue properties in large-scale structures through finite element analysis.

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.