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Spectral proper orthogonal decomposition(SPOD) has been recently used to capture temporal and spatial characteristics of complex flows. Applying SPOD analysis to such flow fields can deepen the understanding of flow mechanism. To further study the external flow at high Reynolds numbers, large eddy simulation(LES) is applied to acquire the flow field data while SPOD is utilized to extract important flow structures. The considered external flow includes incompressible flow around the cylinder and transonic shock oscillation. For the flow around the cylinder, flow field under three different Reynolds numbers is studied. Power spectral density(PSD) results of lift coefficient show that the vortex shedding frequency of low Reynolds number is S t = 0. 24 , which of moderate Reynolds number is S t = 0. 203 , and S t = 0. 27 for high Reynolds number flow. The increment of Reynolds number will lead to early instability of shear layer on the cylindrical surface. The modal results of SPOD and dynamic mode decomposition(DMD) at different Reynolds numbers show two antisymmetric structures at shedding vortex frequencies, indicating the resemblance in the process of vortex formation. For the shock wave/boundary layer interaction(SBLI) of transonic Royal Aircraft Establishment(RAE) 2822 airfoil, flow field under M ∞ = 0. 729 , R e c = 1. 94 × 1 0 7 is studied. The results show that the occurrence of unsteadiness is closely related to the normal shock wave/turbulent boundary layer interaction happening on the upper surface of the airfoil. Modal results of SPOD near the characteristic frequency show strong discontinuity near the shock wave. • Spectral proper orthogonal decomposition is developed to capture the spatiotemporal modes. • The mechanism of Reynolds-number effect on the flow around circular cylinder is revealed. • The low frequency of transonic buffet on airfoil is identified. [ABSTRACT FROM AUTHOR]

The transonic airflow around a supercritical wing with a shock wave is described via direct numerical simulations. Flow control for turbulent drag reduction is applied via streamwise traveling waves of spanwise velocity applied on a finite portion of the suction side. The near-field modifications caused by the forcing are studied via the analysis of the wake profile downstream of the trailing edge. Moreover, for the first time, the effects of spanwise forcing on aeroacoustic noise are considered to establish whether active flow control for drag reduction could possibly increase noise. By extracting the acoustic signals on a circumference placed in the near-field around the wing and by studying them in terms of sound intensity and frequency content, it is found that noise intensity is not significantly increased by spanwise forcing and that frequency content is only minimally altered. Furthermore, if the angle of attack is reduced to take into account the increased lift and the reduced drag made possible by the control action, changes in the noise characteristics become negligible. [ABSTRACT FROM AUTHOR]

The present study focuses on particular properties of transonic flows through a planar channel featuring a circular bump on the lower wall. The selected geometry is reminiscent of the region surrounding the trailing edge of an airfoil at zero angle of attack and the resulting flow pattern is indeed similar to the fishtail shock-pattern that characterizes airfoils flying at nearly sonic speed. Numerical simulations have been conducted by solving the inviscid Euler equations using both a commercial and an in-house CFD code; discontinuities are modeled using shock-capturing in the former and shock-fitting in the latter. Numerical experiments reveal different shock-patterns obtained by independently varying the inlet Mach number and the outlet-to-inlet static pressure ratio. When shock-interactions occur, shock-polar analysis reveals that the branching point can be modeled using either von Neumann's three-shock-theory or Guderley's four-wave-theory, depending on the inlet Mach number. Furthermore, for certain pairs of boundary conditions, multiple solutions have been observed. [ABSTRACT FROM AUTHOR]

Evacuated tube maglev train (ETMT) system aims to advance ultra-high-speed transportation, featuring unique high-speed flow phenomena and complex shockwave dynamics in low-pressure environments that demand further exploration. This paper examines the flow structures and aerodynamic loads of the ETMT over a range of Mach numbers from 0.8 to 2.0. Leveraging a compressible, density-based solver based on the Advection Upstream Splitting Method, extensive numerical simulations of the ETMT were conducted across transonic and supersonic regimes, revealing diverse aerodynamic characteristics under varying operational conditions. The research delineates how aerodynamic properties distinctively shift with operating Mach numbers. In supersonic conditions, distinct shockwave effects emerge prominently, and as the train's velocity escalates, there is a consistent reduction in overall drag and lift coefficients, resulting in a net reduction of 32% in the total train drag coefficient (a most economical Mach number of 1.8) and the lift diminished by 38%. However, notable disparities exist in the drag and lift coefficients among different train sections. These insights are instrumental in understanding the aerodynamic behavior of tube trains at ultra-high speeds and serve as a crucial guide for the train's exterior design. [ABSTRACT FROM AUTHOR]

The shock wave/boundary layer interaction at the tip of transonic shrouded rotor blades usually leads to severe flow separation, which seriously affects aerodynamic performance. In this paper, the effect of part-shroud treatment on the shock wave/boundary layer interaction was numerically investigated. The part-shroud treatment suppresses the shock wave/boundary layer interaction-induced flow separation by introducing the tip leakage vortex at an optimal location. Parametric studies reveal that positioning the tip leakage vortex onset upstream the shock wave effectively reduces flow separation on the suction side, with the best outcomes achieved when the onset is slightly upstream of the shock wave. Based on these findings, the optimal shroud range extends from the leading edge to just upstream of the shock wave under a near stall condition. In National Aeronautics and Space Administration Rotor 35, the 30% shrouded rotors demonstrate a 2.4% improvement in the stall margin over 100% shrouded rotors with little reduction in peak efficiency. [ABSTRACT FROM AUTHOR]

Radial migration flow structures on the blade suction surface in the rotor are commonly observed in transonic axial-flow fan/compressors. At sea level conditions, its intensity is relatively weak and causes very slight aerodynamic losses. However, under high-altitude low Reynolds number conditions, the radial migration flow will be significantly intensified. It will have strong interactions with other flow structures (such as shock wave, tip leakage flow, and wake) in the blade passage and lead to innegligible efficiency deficit in the entire compressor stage. To enhance the compressor performances at low Reynolds number conditions, it is crucial to achieve fundamental understandings of the origination, development, and downstream effects of radial migration flow. This paper conducted a detailed numerical investigation on the influences of radial migration flow near the rotor blade suction surface on the additional loss in a transonic compressor stage. The computational results confirm that under high-altitude conditions, radial migration flow is originated at the rotor blade hub corner and develops from the local separated flow and vortex. Its trajectory is largely determined by the interaction between the shock wave and the boundary layer flow, eventually mixing with the tip leakage flow and causing an accumulation of low-momentum flow near the casing. The migration of radial flow along the spanwise fraction interacts with the rotor blade wake and results in a larger total pressure and velocity deficit at the downstream stator inlet, deteriorating the aerodynamic loss in the stator blade. Meanwhile, the radial migration flow can be also triggered by the corner separation on the stator suction surface near the trailing edge, which exacerbates the local loss in the blade passage. Based on the features of radial migration flow, blade bend re-designs mainly near the hub and shroud ends have been applied to the rotor. It shows that the re-designed rotor has achieved an alleviated radial migration flow and a reduced wake zone at low Reynolds number conditions. Moreover, the re-designed compressor shows a notable higher adiabatic efficiency than the baseline one over the entire operating range, while the total pressure ratio and stable operating range are not sacrificed. [ABSTRACT FROM AUTHOR]

Recent studies show that non-radial structures arising from massive star shell convection play an important role in shaping core-collapse supernova explosions. During the collapse phase, convective vortices generate acoustic waves that interact with the supernova shock. This amplifies turbulence in the post-shock region, contributing to explosion. We study how various physical parameters influence the evolution of these convective vortices during stellar collapse using simplified simulations. We model the collapsing star with a transonic Bondi flow and represent convection as solenoidal velocity perturbations. Our results are consistent with previous studies, demonstrating that the peak perturbation amplitude scales linearly with the pre-collapse convective Mach number and inversely with the angular wavenumber of convection. While the shell radius and width primarily determine the time-scale of accretion, they have little impact on the peak perturbation amplitudes. Finally, we show that when the convective Mach number is below approximately 0.2, the dynamics remain within the linear regime. [ABSTRACT FROM AUTHOR]

This work is concerned with establishing the existence and uniqueness of the solution to a mixed-type degenerate boundary value problem for a relativistic magnetohydrodynamic system. We first consider a full relativistic magnetohydrodynamic system and reduce it to a simplified form under the assumption that the magnetic field vector is orthogonal to the velocity vector. We consider a boundary value problem for the steady part of this reduced system where the boundary data are prescribed on a sonic boundary and a characteristic curve. Here the main difficulty is the consideration of a relativistic system, with a general equation of state while considering the magnetic field effects as well, which we believe has never been analyzed before in the context of the analytical study of sonic–supersonic flows. Also, the degeneracy of the governing equations along the sonic curve is a crucial challenge. However, we employ the iteration method used in the work of Li and Hu [1] to prove the existence and uniqueness of a local classical supersonic solution in the partial hodograph plane first, and finally, we recover a local smooth solution to the boundary value problem in the physical plane by applying an inverse transformation. [ABSTRACT FROM AUTHOR]

To reduce the computational burden of aerodynamic design optimization, a multi-fidelity expected improvement (MFEI) method is developed, based on the error analysis of a multi-level hierarchical kriging (MHK) model for accelerating optimization convergence. By maximizing the MFEI function, a new sample of arbitrary fidelity level is determined to ameliorate the accuracy of the MHK model, and convergence to the high-fidelity optimum is ensured. The proposed optimization method based on MFEI and MHK is demonstrated by two analytical function cases and verified by two aerodynamic design optimizations: drag minimizations of an RAE2822 aerofoil and an ONERA M6 wing in transonic flows. It is shown that the MFEI method tends to infill more gainful samples of lower fidelities during optimization, so fewer highest-fidelity samples are required. This confirms that the proposed method can obtain optimal results within a limited computational budget and is more efficient than the existing single-fidelity or two-fidelity methods. [ABSTRACT FROM AUTHOR]

A new and simple implicit Green-Gauss gradient (IGG) scheme for unstructured meshes is proposed exploiting the ideas from the linearity-preserving U-MUSCL scheme to define values at cell faces. We construct an implicit one-parameter family of gradient schemes referred to as IGG(χ) where χ is a free-parameter. The computed gradients are at least first-order accurate on generic polygonal meshes and second-order accurate on uniform Cartesian meshes except when χ = 1 / 3 for which fourth-order accuracy can be realised. A theoretical analysis is carried out to understand the effect of the control parameter χ on accuracy and resolution of the gradients and numerical experiments on various mesh topologies confirm the theoretical findings. Finite volume simulations of the Poisson and Euler equations on Cartesian and unstructured meshes further highlight that the IGG(χ) is a versatile gradient scheme that gives second-order accurate solutions with the iterative convergence of the solver dependent on the choice of the χ parameter. The framework described in this study can also be employed to devise an implicit least-squares gradient scheme that applies equally well to unstructured finite volume and meshfree solvers. • Compact gradient scheme IGG(χ) derived from Gauss divergence theorem. • Uses linearity-preserving U-MUSCL scheme to compute values at cell faces. • New and simple implicit gradient scheme with a user-defined free-parameter. • Leads to family of schemes that give consistent gradients on unstructured meshes. • Efficacy demonstrated using Euler simulations of subsonic and transonic flows. [ABSTRACT FROM AUTHOR]

The aerodynamic performance of axial compressor rotors is negatively affected by the ingestion of boundary layer fluids upstream. As the boundary layer becomes thicker, the blade tip load increases and the local loss is aggravated, especially under off-design operating conditions. The major objective of this research is to evaluate the potential for novel blade sweep designs that can tolerate the ingested low-momentum boundary layer fluids. An optimization design approach using a surrogate model and genetic algorithm is employed. By altering the blade stacking line, the optimized sweep design is obtained. The flow mechanisms that enable the performance of the compressor rotor to be improved are fully analyzed, and the findings indicate that the aerodynamic advantages primarily stem from two key aspects. First, in the tip region, the blade loads are decreased at various chordwise locations and the interaction of the tip leakage flow with the mainstream is alleviated. As a result, the loss near the tip is reduced. Second, the blade sweep design alters the distribution of shock intensity across the spanwise direction, leading to a decrease in shock wave intensity in the mid-span region. This is beneficial in reducing the shock wave/boundary layer interaction strength at the trailing edge of the blade airfoil. Overall, after the sweep design has been optimized to ingest the upstream boundary layer, the compressor rotor experiences a 0.8% improvement in adiabatic efficiency compared with the baseline rotor, while preserving the total pressure ratio and stall margin. Additionally, the redesigned compressor retains the overall performance level under clean inlet conditions. This research provides a potentially effective blade sweep optimization design strategy that allows transonic compressor rotors to tolerate low-momentum upstream boundary layer incoming flows. [ABSTRACT FROM AUTHOR]

The aerodynamic performance of transonic compressors, particularly the stall margin, is significantly influenced by inlet distortion. While time-marching methods accurately simulate such unsteady flows, they can be time-consuming. To enhance the computational efficiency, two Fourier-based methods are proposed in this paper: the time-accurate method with interface filtering and the time–space collocation (TSC) method. The time-accurate method with interface filtering ignores the rotor–stator interaction effects, enabling a larger time step and faster convergence. In contrast, the TSC method accounts for harmonics of conservative variables and transforms the unsteady simulation into multiple steady-state calculations, thereby reducing computational costs. The two Fourier-based methods are validated using NASA Stage 67 and a two-stage transonic fan. Near the peak efficiency point, the results from both methods closely match that of URANS simulation and experimental data. The time-accurate method with interface filtering demonstrates a speed enhancement of 4 to 5 times as a result of a reduction in the iteration steps. In contrast, the TSC method exhibits a speed improvement of at least 20 times in two specific cases, attributable to the significantly smaller mesh size and iteration steps employed in the TSC method compared to the URANS method. Near the stall point, more harmonics for inlet distortion are necessary in TSC simulation to accurately capture flow separation. In the two-stage transonic fan simulations, the strong rotor–stator interaction effects lead to deviations from the URANS simulation; nevertheless, the Fourier-based simulations accurately reflect the trend of the stall margin under total pressure distortion. Overall, the Fourier-based methods show promising potential for engineering applications in estimating the performance degradation of compressors subjected to inlet distortion. [ABSTRACT FROM AUTHOR]

Large transport aircraft tend to adopt a wing layout with a high aspect ratio and swept-back angle due to the requirement of a high lift-to-drag ratio. Composite material is typically employed to ensure the light weight of the structure, causing serious static aeroelasticity problems to the aircraft. When the airplane is flying in the transonic region, its aerodynamic load is very complex, and the large load leads to large deformation of the wing, triggering geometric nonlinear effects, which further affects the static aerodynamic elasticity characteristics of the wing. In this study, in order to study the static aeroelastic characteristics of the transonic flow of a high-aspect-ratio airfoil, a new design method of the scaled similar optimization model is described, and the change in the model lift coefficient due to geometrically nonlinear static aeroelasticity effects when the angle of attack is changed was investigated by using simulation and wind tunnel test methods. In order to ensure the accuracy of the wing shape when the model was deformed greatly, this study employed the structural design scheme of the wing with the skin as the main stiffness component, and the thicknesses of different regions of the skin were used as the design variables for the stiffness optimization design. The engineering algorithm of nonlinear finite elements was used in this study to calculate the curve of lift with the angle of attack considering the geometric nonlinear static aeroelasticity effect. The results show that the similarity optimization process employed in this study can be used to complete the design of the high-speed aerostatic wing test model, and the wind tunnel test results show that geometric nonlinearity has a large impact on the lift coefficient of the wing. [ABSTRACT FROM AUTHOR]

This paper presents the results of numerical simulation of a transonic flow around a blunt body at a zero angle of attack in a test section with porous walls of the T-128 wind tunnel (WT) (Russia, Central Aerohydrodynamic Institute (TsAGI)). The calculations are performed using the EWT-128 TsAGI software package, which uses Darcy-type boundary conditions on the walls of the test section, whose porosity coefficient depends on the direction of the gas flow through the perforated surface. A comparison is made of the calculated and experimental drag coefficients of the model, as well as distributions of the isentropic Mach number on the model and on the perforated walls of the test section of the WT. A method is proposed for determining the correction to the Mach number of the oncoming flow to eliminate the wall interference of the test section in an aerodynamic experiment. The drag coefficients obtained as a result of calculations under WT conditions are corrected for the influence of the flow boundaries and compared with the free flow data. The calculated corrections are used to correct the experimental data. [ABSTRACT FROM AUTHOR]

Transonic flight presents significant challenges due to complex phenomena such as shock waves and turbulent boundary layer interactions, often leading to flow separation and instabilities. This study aims to evaluate the performance of various turbulence models in predicting these phenomena using Computational Fluid Dynamics (CFD) with ANSYS-Fluent 2020R1. The ONERA M6 wing model is analyzed under specific flow conditions (angle of attack, α = 3.06°, Mach number, M∞ = 0.8395, and Reynolds number, Re = 11.72 × 10^6) to assess five turbulence models: Spalart–Allmaras, k-ε Standard, k-ε Realizable, k-ω Standard, and k-ω SST. Results from the simulations closely matched experimental data, with all models demonstrating a margin of error within 5% compared to NASA's CFD benchmarks. Notably, the Spalart–Allmaras and k-ω SST models showed superior accuracy in predicting shock formation and pressure distribution, highlighting their potential for enhanced predictive capabilities in aerodynamic analysis. This study demonstrates the effectiveness of ANSYS-Fluent in aerodynamic performance evaluation and suggests pathways for advancing aircraft design efficiency and safety through improved turbulence modeling. [ABSTRACT FROM AUTHOR]

Transonic aeroelasticity remains a significant challenge in aerospace. The coupling mechanism of aeroelastic problems involving the coexistence of fluid modes and multiple structural modes still needs further investigation. For this purpose, we analysed the dynamic characteristic of a two-degree-of-freedom (2DOF) NACA0012 airfoil in pre-buffet flow. First, we constructed an aeroelastic reduced-order model, which can represent near-unstable transonic flow using the dominant fluid mode. Then, the flutter mechanism was investigated by studying the main eigenvalues of the model that vary with the natural pitching frequency. The results revealed that the existence of the fluid mode transitions the transonic flutter type from coupled-mode flutter to single-DOF (SDOF) flutter, which leads to a reduction in the flutter boundary. Under the effect of the fluid mode, the system produces six aeroelastic phenomena at different structural natural frequencies, including SDOF heaving/pitching flutter, heaving/pitching instability within coupled-mode flutter, forced vibration and stable state. Moreover, we identified two types of SDOF flutter in the 2DOF system. The first type corresponds to the traditional SDOF flutter, where the coupling of other modes has a small impact on the system's stability in most cases. However, within specific ranges of natural frequencies, this type of SDOF flutter may disappear due to coupling with other modes. The second type of SDOF flutter is characterized by strong coupling dominated by the unstable mode. It arises from the interaction among the flow, heaving and pitching modes, and does not manifest in the absence of any of these modes. [ABSTRACT FROM AUTHOR]

In transonic flow conditions, buffeting associated with finite-amplitude lift fluctuations can limit the operational envelope of an aircraft. For both airfoils and wings, these oscillations have been linked to global flow instabilities that arise from a Hopf bifurcation. We employ a combination of numerical simulations and global stability analysis to investigate the near-critical behaviour of the oscillatory buffet-onset instability on airfoils. The flow is governed by the unsteady Reynolds-averaged Navier–Stokes equations, with a basic state provided by a steady-state solution. In the weakly nonlinear formulation, the disturbance amplitude is described by the Landau equation. The linear growth rate can be determined from either the simulations or the stability analysis, and the Landau constant is derived from simulations resulting in finite-amplitude equilibrium states. The results show that the Landau constant is nearly independent of Mach number and angle of attack for a given airfoil. Using the Landau constant derived from a small number of simulations, the stability analysis can be employed to efficiently capture the essential finite-amplitude behaviour needed to estimate the buffet-onset boundary. The stability analysis is shown to capture the envelope of lift oscillations during a continuous pitch of an airfoil, from pre-buffet through post-buffet lift levels. [ABSTRACT FROM AUTHOR]

The document delves into the flutter mechanism of damaged transonic rotor blades, utilizing high-fidelity fluid-solid coupling methods to analyze aeroelastic stability. It highlights the impact of structural damage on aeroelasticity, aiming to enhance fatigue prediction and optimize turbomachinery design. The study explores the dynamic response of blades under varying conditions, emphasizing the influence of damage parameters on stability and the role of boundary layer separation in inducing vibrations. The research provides valuable insights for designing rotor blades to mitigate flutter risk and predict damage, offering a collection of articles on aeroelasticity in turbomachinery and mechanical systems. [Extracted from the article]

For propeller-driven Mars airplanes operating at low Reynolds numbers, the speed of the rotor tips may reach transonic. To date, only a few studies have investigated the transonic buffet in flow past low-Reynolds-number airfoils. In this study, direct numerical simulations of high-speed flow (M=0.2 , 0.6, and 0.8) past a NACA 0012 airfoil are performed with a low Reynolds number of Rec=23,000 , where transonic buffet is observed at M=0.8. We first investigated the effects of Mach number on the aerodynamic performance and flow fields. Dynamic mode decomposition (DMD) and linear stability analysis (LSA) are used to analyze the flow instability mechanisms of the transonic buffet. Results showed that the multiple high-frequency oscillations are related to the vortex shedding at the trailing edge of the airfoil, and the low-frequency oscillation is caused by Type C shock motions. Both of them are confirmed to be self-sustained in feedback cycles. [ABSTRACT FROM AUTHOR]

This study addresses the technical challenge that conventional self-recirculating casing treatments predominantly enhance the compressor stability with minimal efficiency gains. Focusing on a transonic axial flow compressor, National Aeronautics and Space Administration Stage35, the research designed an innovative stator suction-type self-circulating casing treatment (SS-SCT), which features the suction slots on the stator instead of the traditional self-circulating casing treatment's suction port. Using multipassage unsteady numerical simulations, this paper investigated the underlying mechanisms by which the SS-SCT affects the compressor performance. The numerical results reveal that the SS-SCT not only increases the compressor stall margin by 3.36% but also enhances the efficiency by 0.13% at the design point and by 0.6% at the near-stall point. The jetting action of the SS-SCT mitigates the detrimental effects of the rotor tip leakage flow and alleviates the blockages in the rotor tip passage at the near-stall point, thereby improving the compressor stability. Concurrently, the suction slots extract the low-speed fluid generated by the boundary layer separation in the stator passage, reducing flow losses, and consequently enhancing the compressor efficiency. [ABSTRACT FROM AUTHOR]

Squealer tip has a significant influence on both aerodynamic and heat transfer characteristics of the high-pressure turbine. Among the geometric parameters of the squealer, squealer height is one of the essential parameters in the tip design. However, due to the complexity of parameterization and meshing of the squealer, the related research is usually carried out on the squealer with a constant height. In this paper, a parameterization strategy generates squealer of assigned heights at four key positions of the blade, the leading edge-pressure side, the leading edge-suction side, the trailing edge-pressure side, and the trailing edge-suction side. An in-house mesh generation platform (NuFlux) is adopted to automatically generate the structured meshes. The aerothermal performance of a transonic turbine stage is assessed using steady Reynolds-averaged Navier–Stokes simulations with the k − ω shear stress transport model for the turbulence closure. The main purpose is to obtain the squealer tip configuration with the lowest heat transfer coefficient. The results show that non-uniform squealer further reduces the cavity floor heat transfer on the basis of uniform squealer by changing the interaction process between the asymmetric vortex pair (the pressure-side corner vortex and the casing-driven scraping vortex), which provides a valuable reference for the design of the squealer tip of advanced high-pressure turbines. [ABSTRACT FROM AUTHOR]

Accurate and rapid prediction of compressor performance and key flow characteristics is critical for digital design, digital twin modeling, and virtual–real interaction. However, the traditional methods of obtaining flow field parameters by solving the Navier–Stokes equations are computationally intensive and time-consuming. To establish a digital twin model of the flow field in a transonic three-stage axial compressor, this study proposes a novel data-driven deep attention symmetric neural network for fast reconstruction of the flow field at different blade rows and spanwise positions. The network integrates a vision transformer (ViT) and a symmetric convolutional neural network (SCNN). The ViT extracts geometric features from the blade passages. The SCNN is used for deeper extraction of input features such as boundary conditions and flow coordinates, enabling precise flow field predictions. Results indicate that the trained model can efficiently and accurately reconstruct the internal flow field of the compressor in 0.5 s, capturing phenomena such as flow separation and wake. Compared with traditional numerical simulations, the current model offers significant advantages in computational speed, delivering a three-order magnitude speedup compared to computational fluid dynamics simulations. It shows strong potential for engineering applications and provides robust support for building digital twin models in turbomachinery flow fields. [ABSTRACT FROM AUTHOR]

During the last few decades, plasma actuators have emerged as promising devices for aerodynamic flow control. This study focuses on the use of nanosecond plasma actuators for such purposes. A thermal phenomenological model is employed to simulate the effects of these actuators. The propagation of shock waves and their interactions for two specific geometries of plasma actuators, linear and annular plasma synthetic jet actuators, are examined here. A comparative analysis of the performance of these two configurations is presented. Furthermore, the geometric characteristics and temperature model are analyzed to provide insights that can be applied to practical problems. The influence of the actuators on a projectile in the transonic flow is also investigated. The results of the present study show that actuators placed in the conical and cylindrical regions of the object do not contribute to drag reduction. Conversely, actuators positioned at the boat-tail and base of the object effectively reduce drag. This drag reduction is primarily attributed to thermal disturbances in the separation area. Additionally, it is observed that the effects of shock waves and their interaction with stationary waves around the projectile are negligible in terms of drag force reduction. [ABSTRACT FROM AUTHOR]

Flexible aeroshells, featuring deformable membrane surfaces supported by pressurized gas, offer a novel re-entry system for high-altitude deceleration during re-entry. The flexibility of structural components introduces unique challenges in predicting aerodynamic behavior, as deformations can significantly alter the flow field, affecting pressure distribution and potentially causing aerodynamic instabilities. Therefore, coupled analysis is crucial to ensure the performance reliability of inflatable aeroshells. This study investigates the unsteady aerodynamic behavior, considering structural deformation in a two-way partitioned coupled manner in transonic flow. The dominant deformation patterns are identified using dynamic mode decomposition and the greedy compressive sensing algorithm. The findings reveal that the primary deformation pattern combines axial deformation and swing oscillation. The dominant oscillation frequencies remain different from the eigenfrequencies, indicating that the aerodynamic force is the primary driving force of such oscillations. The time evolution of these dominant modes indicates purely oscillatory behavior rather than increasing or decreasing trends. These insights are critical for the design and reliability assessment of inflatable aeroshells in varying aerodynamic environments. [ABSTRACT FROM AUTHOR]

Wind tunnel testing is a crucial method for studying aircraft flutter. Using heavy gas as the wind tunnel medium can mitigate the escalating issue of test models being overweight as advanced aircraft develop. This paper employs an analytical method for numerical calculations of three-dimensional (3D) wing flutter based on fluid–structure interaction (FSI). Flutter calculations for the Goland wing are conducted, and the results in the air medium are consistent with the literature. In contrast, significant differences in flutter behavior are observed in the heavy gas R134a medium. Compared to air, when the model reaches a critical state in R134a, the incoming flow velocity is lower, the incoming flow density is approximately 3 to 5 times air, and the incoming flow dynamic pressure is about 1.1 to 1.2 times that of air. The correction of heavy gas flutter data is crucial for wind tunnel testing. This paper proposes a correction method based on the unsteady transonic flow similarity law proposed by Bendiksen under quasi-steady conditions. Attempts are made to revise relevant published wind tunnel tests and heavy gas flutter calculation results. The transonic flutter similarity law effectively explains the flutter similarity of rigid models in both heavy gas and air media. Still, it fails in cases with highly reduced frequencies and low mass ratios, such as those encountered with flexible wings. [ABSTRACT FROM AUTHOR]

The performance of supersonic ejectors for heat recovery refrigeration cycles is optimised by means of fluid dynamic shape optimisation with a discrete adjoint method to efficiently evaluate gradients of an objective function with respect to an arbitrary number of design variables. The Reynolds-averaged Navier-Stokes equations and the discrete adjoint equations are solved using the finite volume solver SU2. The model is validated on cases from the literature. It is conclusively demonstrated that the method employed can generate a well performing ejector shape from a failed design despite concerns about the well-posedness of the problem given the apparent discontinuity of the entrainment ratio objective function at the critical point in the parameter space. This is achieved in only 10-20 iterations with a cost of one or two direct solutions each. The entrainment ratios predicted with the optimal shape exceed those announced by a leading manufacturer for identical cycle conditions by around 15 %. More importantly, discrete adjoint shape optimisation is here applied successfully for the first time to transonic ejector flows and is shown to be reliable and robust, paving the way for the application of topology optimisation methods to ejectors. [ABSTRACT FROM AUTHOR]

The zonal detached-eddy simulation (ZDES) method divides the flow field into different computational modes, offering flexibility in the simulation of complex flows. The transonic shock buffet on the upper surface of the OAT15A supercritical airfoil is numerically simulated by ZDES based on the k-ω SST model. The results show that the error in the low-frequency characteristics of the transonic shock buffet predicted by ZDES compared to experiments is within 3%. Its predictions of pressure fluctuations and averaged velocity agree better with the experimental data. Overall, ZDES is effective in predicting the periodic oscillations of shock waves along the streamwise direction, flow separation induced by shock wave motion, and the small-scale vortex structures in the wake region. [ABSTRACT FROM AUTHOR]

The article in the Journal of Fluid Mechanics delves into the sensitivity of flows over three-dimensional swept wings at low Reynolds numbers, using linear stability analysis to study flow instability and receptivity. The research identifies regions of maximum flow receptivity to momentum forcing, which vary with Reynolds number and wing sweep angle, impacting flow control applications. The document also references studies on topics like vortex shedding, cavity oscillations, and three-dimensional flows around flat plate wings, exploring aspects such as frequency selection and the effects of aspect ratio on flow behavior. Researchers like Theofilis, Sipp, and Zhang have contributed to this body of work, available online through Cambridge University Press. [Extracted from the article]

This paper presents a novel framework for the active control of transonic airfoil flutter using synthetic jets through deep reinforcement learning (DRL). The research, conducted in a wide range of Mach numbers and flutter velocities, involves an elastically mounted airfoil with two degrees of freedom of pitching and plunging oscillations, subjected to transonic flow conditions at varying Mach numbers. Synthetic jets with zero-mass flux are strategically placed on the airfoil's upper and lower surfaces. This fluid–structure interaction (FSI) problem is treated as the learning environment and is addressed by using the arbitrary Lagrangian–Eulerian lattice Boltzmann flux solver (ALE-LBFS) coupled with a structural solver on dynamic meshes. DRL strategies with proximal policy optimization agents are introduced and trained, based on the velocities probed around the airfoil and the dynamic responses of the structure. The results demonstrate that the pitching and plunging motions of the airfoil in the limited cycle oscillation (LCO) can be effectively alleviated across an extended range of Mach numbers and critical flutter velocities beyond the initial training conditions for control onset. Furthermore, the aerodynamic performance of the airfoil is also enhanced, with an increase in lift coefficient and a reduction in drag coefficient. Even in previously unseen environments with higher flutter velocities, the present strategy is achievable satisfactory control results, including an extended flutter boundary and a reduction in the transonic dip phenomenon. This work underscores the potential of DRL in addressing complex flow control challenges and highlights its potential to expedite the application of DRL in transonic flutter control for aeronautical applications. [ABSTRACT FROM AUTHOR]

The uncertain factors in wind tunnel experiments, especially the interference effects, will cause the flow around the test model to differ from that under the free flow condition. The wind tunnel wall interference effects are challenging to be eliminated, especially in transonic flow regions where classical linear correction methods are ineffective. To address this issue, an efficient wind tunnel wall interference correction framework based on data assimilation is proposed for steady and shock buffet flows around the airfoils in the transonic region in this work. The ensemble Kalman filter (EnKF) is used to find the combination of inflow conditions, which minimizes the differences between the pressure coefficients measured in the wind tunnel experiment and those estimated by computational fluid dynamics (CFD). The polynomial chaos expansion method is used to propagate uncertainty in the forecast step of the EnKF in order to improve efficiency. Typical wind tunnel wall interference correction problems are studied, including two steady cases: the transonic flow around the Royal Aircraft Establishment 2822 (RAE2822) and the National Aerospace Laboratory 7301 (NLR7301) airfoils, and one unsteady case: the transonic shock buffet on the National Aeronautics and Space Administration supercritical (phase 2)-0714 (NASA SC(2)-0714) airfoil. It is demonstrated that with the corrected Mach number and angle of attack, the estimated CFD results are in better agreement with the measured experimental data than traditional linear methods and the computational efficiency is improved compared to the original EnKF. [ABSTRACT FROM AUTHOR]

Unmanned aerial vehicles (UAV) on Mars operate in a thin atmosphere of carbon dioxide, where the flow Reynolds number and local sound velocity are significantly reduced compared to the Earth's atmosphere because of low gas density and temperature. Therefore, a Mars UAV with a rotary-wing design can operate in a state of low-Reynolds number (Re = 1-5 × 104) and transonic flow (Ma tip = 0.7–1.2), where flow separation and shock waves occur simultaneously. The interaction between flow separation and shock waves not only makes the entire flow field more complex but also seriously affects the aerodynamic performance of the airfoil, thereby reducing the payload of UAVs, which previous aircraft designs based on the thick atmospheric environment of the Earth have not considered. Therefore, in this study, a numerical simulation method is used to conduct airfoil optimization research on a characteristic airfoil under low-Reynolds number transonic flow (Re = 33200, Ma = 0.85) conditions to acquire additional insight, improve aerodynamic performance of the airfoil, and enhance the payload. The results were compared with the optimization results of subsonic flow (Re = 19500, Ma = 0.50). The Hicks-Henne bump functions method was selected for airfoil parameterization, and non-dominated sorting genetic algorithms (NSGA)-Ⅱ were used for optimization. Through an analysis of the pressure coefficient (C p), surface friction coefficient (C f), and density gradient, the number and intensity of oblique shock waves on the upper and lower surfaces were observed to significantly affect the aerodynamic performance of airfoils under low-Reynolds number flow. The changes in flow separation and transition positions do not directly affect the aerodynamic performance but weaken the oblique shock wave intensity induced by the separation point or directly transform strong shock waves into weaker compression waves because of the absence of flow separation. The optimization results for both subsonic and transonic flow point to a thinner overall shape with a flatter surface with the drag greatly reduced due to the disappearance of shock waves on the lower surface, which significantly improves aerodynamic performance. Compared with the original airfoil, at Ma = 0.85, this design can increase the lift coefficient (C l) by at least 124 % and lift drag ratio (C l /C d) by 177 %. • Aerodynamics issues at low Reynolds numbers and transonic in the Martian environment. • Complex wave system issues occurring at low Reynolds numbers. • Numerical simulation and optimization of airfoil design for Mars Unmanned aerial vehicles. [ABSTRACT FROM AUTHOR]

The self-circulating casing treatment can effectively expand the stable working range of the compressor, with little impact on its efficiency. With a single-stage transonic axial flow compressor NASA (National Aeronautics and Space Administration) Stage 35 as the research object, a multi-channel unsteady numerical calculation method was used here to design three types of self-circulating casing treatment structures: 20% Ca (axial chord length of the rotor blade tip), 60% Ca, and 178% Ca (at this time, the bleed position is at the stator channel casing) from the leading edge of the blade tip. The effects of these three bleed positions on the self-circulating stability expansion effect and compressor performance were studied separately. The calculation results indicate that the further the bleed position is from the leading edge of the blade tip, the weaker the expansion ability of the self-circulating casing treatment, and the greater the negative impact on the peak efficiency and design point efficiency of the compressor. This is because the air inlet of the self-circulating casing with an air intake position of 20% Ca is located directly above the core area of the rotor blade top blockage, which can more effectively extract low-energy fluid from the blockage area. Compared to the other two bleed positions, it has the greatest inhibitory effect on the leakage vortex in the rotor blade tip gap and has the strongest ability to improve the blockage at the rotor blade tip. Therefore, 20% Ca from the leading edge of the blade tip has the strongest stability expansion ability, achieving a stall margin improvement of 11.28%. [ABSTRACT FROM AUTHOR]

The steady and unsteady flow properties of the National Aeronautics and Space Administration (NASA) single-stage transonic compressor stage 35 are numerically investigated through the Reynolds-averaged Navier–Stokes (RANS) simulation method. The main purpose of the present paper is twofold. One is to validate the capability of the Spalart–Allmaras (S–A)-Helicity-γ model in unsteady RANS (URANS) simulation for rotor–stator interaction of a whole-stage compressor, and another is to figure out the roles played by helicity modification and transition augmentation in improving the performance of traditional S–A model. The results of steady RANS simulation indicate that the S–A model with helicity regulation is more sensitive to vortex structures in the end wall and blade-tip regions, while the inclusion of transitional intermittency factor can help predict the important transition phenomenon in the midspan region. As the back pressure increases, the inaccuracy of flow fields exchange caused by the mixing plane method becomes nonnegligible for the prediction of aerodynamic performance and thermodynamic quantities. Thus, URANS simulation seems to be necessary for more accurate prediction of the flow details of the stator subjected to periodic sweep of the rotor wake. On the suction surface of the stator, a turbulent separation bubble near the leading edge is transported downstream and merges into another separation bubble with much lower turbulence intensity at the trailing edge periodically. The combination of helicity and transition modifications provides advantages over traditional S–A model in describing this process, quite similar to its performance in steady simulations. [ABSTRACT FROM AUTHOR]

Numerical simulations of a T-tail flutter in subsonic and transonic flow regimes were conducted using a 3D Navier-Stokes code which considers the in-plane motion of the horizontal tail plane (HTP). In our subsonic flutter simulations, we confirmed that the flutter velocity of the T-tail is dependent upon the mean angle of attack of the HTP, and also confirmed that the results show a satisfactory agreement with those predicted using a lifting-surface theory. Thereby we were able to verify the reliability and accuracy of the 3D Navier-Stokes code. In transonic simulations, the sharp transonic-dip phenomena of the dynamic pressure of flutter and limit cycle oscillation (LCO) were predicted. It can be shown that the characteristics of LCO and flutter are considerably different depending on Mach number and the angle of attack of the HTP. As a possible cause of these phenomena, we examined the behavior of the shock wave generated on the upper or lower surface of the HTP, which varies strongly depending on the Mach number and the angle of attack of the HTP. [ABSTRACT FROM AUTHOR]

The unsteady pressure field over an axisymmetric backward-facing step was investigated experimentally at transonic freestream Mach numbers of 1.05, 1.2, and 1.4. The study was aimed at examining the influence of transonic freestream Mach numbers on the spatio-temporal character of the unsteady pressure field and on the dominant modes/mechanisms driving it. Surface flow visualization, schlieren, and unsteady pressure measurements were carried out as part of the experimental investigation. From oil flow visualization and schlieren, the reattachment region was identified, and consequently, the mean reattachment length was estimated. The mean reattachment length shows an increase with the increase in freestream Mach number. The coefficient of mean pressure along the rearbody imitates a classical backward-facing step flow profile and can be divided into three distinct regions. The peak values of the coefficient of mean pressure and the coefficient of root mean square of the fluctuation are seen to decrease with an increase in the freestream Mach number. Conventional spectral analysis reveals that as the freestream Mach number increases, the dominant peak in the spectra shifts to lower frequencies. From the spectra, three dominant fluid dynamic mechanisms depending on the freestream Mach number have been identified. Proper Orthogonal Decomposition (POD) analysis shows that 79–84 % of the total energy contribution comes from the first six modes. The temporal dynamics of the POD modes indicate three prominent mechanisms are responsible for the unsteady pressure field. Spectral analysis of POD modes indicates that the spectra are primarily driven by the first three POD modes for freestream Mach number of 1.05 and the first two modes for freestream Mach numbers of 1.2 and 1.4. Moreover, it reveals the presence of three dominant modes, and the freestream Mach number strongly dictates the dominant mode that is driving the pressure field. [ABSTRACT FROM AUTHOR]

This paper deals with the topic of reducing drag force acting on aircraft wings by incorporating novel winglet designs, such as multi-tip, bird-type, and twisted. The high-speed NASA common research model (CRM) was selected as the baseline model, and winglet designs were retrofitted while keeping the projected wingspan constant. Computational analysis was performed using RANS coupled with the Spalart–Allmaras turbulence model to determine aerodynamic coefficients, such as CL and CD. It was observed that the multi-tip and bird-type designs performed exceptionally well at a low angle of attack (0°). A parametric study was conducted on multi-tip winglets by tweaking the parameters such as sweep angle (Λ), tip twist (Є), taper ratio (λ), and cant angle (Φ). The best combination of parameters for optimal aerodynamic performance while maintaining the wing root bending moment was determined using both the Taguchi method and Taguchi-based grey relational analysis (T-GRA) coupled with principal component analysis (PCA). Also, the percentage contribution of each parameter was determined by using the analysis of variance (ANOVA) method. At the design point, the optimized winglet design outperformed the baseline design by 18.29% in the Taguchi method and by 20.77% in the T-GRA coupled with the PCA method based on aerodynamic efficiency and wing root bending moment. [ABSTRACT FROM AUTHOR]

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This paper investigates the aerodynamic performance of a system of multiple vehicles (three) propelled by an electric fan/compressor system in a closed low-pressure environment. The study was conducted on a 2-D axisymmetric model of multiple vehicles traveling at the same cruising speed inside a tube. The flow field was computed using the unsteady RANS model. The results demonstrate a significant improvement in aerodynamic performance when using a multiple-vehicle system equipped with an air-breathing electrical engine. The combined use of multiple-vehicle and the air-breathing propulsion system eliminated the piston effect at low Mach numbers and reduced transonic range effects at higher Mach numbers. Consequently, a considerable enhancement in aerodynamic performance over the single-vehicle model was attained. A minimum of 11% and a maximum of 57% reduction of further in average system drag was achieved for M = 0.3 and 0.7, respectively. An increase in thrust was also observed in the multiple-vehicle system over the single-vehicle system. The results of the study demonstrated the advantage of using an air-breathing multiple-vehicle system as an efficient and environmentally friendly high-speed transportation system. [ABSTRACT FROM AUTHOR]

Vibration causes compressor rotor blade failure that affects aeroengine safety. Therefore, a plane cascade model of rotor blades was developed to capture the vibration phenomenon. The blade vibration, vortex shedding, and aerodynamic noise were measured. The results indicate that flow and noise induce a natural vibration mode, resulting in torsional vibration fatigue damage. The frequency of trailing edge vortex shedding is consistent with the frequency of tip clearance leakage vortex pulsation and is close to the second-order modal frequency of the blade, which induces torsional vibrations. The triple frequency of the background noise peak frequency in the wind tunnel is consistent with the first standing wave acoustic resonance modal frequency of the cascade experimental section duct and is close to the third-order natural modal frequency of the blade, which induces bending and torsional vibrations. This research provides guidance for the study of rotor blade multifield coupling-induced vibration. [ABSTRACT FROM AUTHOR]

The complex tip flow instability and its induced non-synchronous vibration have become significant challenges, especially as aerodynamic loading continues to increase. This study investigates the effects of tip clearance on non-synchronous propagating flow disturbances of compressor rotors under high aerodynamic loading conditions by conducting full-annulus unsteady numerical simulations with three typical tip clearance values for a 1-1/2 stage transonic compressor. The non-synchronous aerodynamic excitation frequency, circumferential mode characteristics, and annular unstable flow structures are analyzed under near stall conditions. The results show that the total pressure ratio and normalized mass flow parameters first increase and then decrease as the tip clearance increases from 0.5%C (where C represents the tip chord length) to 2%C under high aerodynamic loading conditions, instead of constantly decreasing. For the 0.5%C tip clearance case, the traveling large-scale tornado-like separation vortices cause a low non-synchronous aerodynamic excitation frequency and severe pressure fluctuations. The periodic shedding and reattachment processes of the rotor blades separated by 2 – 3 pitches result in 19 dominant mode orders in the circumferential direction. As the tip clearance increases from 1%C to 2%C, the difference of tip flow structures in each blade passage is significantly weakened, and the dominant mode order of the disturbance is equal to the rotor blade-passing number. The pressure fluctuation is mainly caused by cross-channel tip leakage flow, and the aerodynamic excitation frequency exhibits evident broadband hump characteristics, which has been reported as a rotating instability phenomenon. [ABSTRACT FROM AUTHOR]

In this paper, theoretical and numerical analysis of a transonic flow field of critical flow Venturi nozzles according to the ISO 9300 standard is performed. Deviations of the flow field from an estimate based on one dimensionality are clarified. While the theoretical analysis allows prediction of these deviations, the numerical analysis allows quantification of their influence. The main studied phenomena include the local supersonic compression in transonic expansion and the Prandtl-Meyer expansion. The tendency of the flow field to spatial oscillations is shown, alongside the ability of the system to damp these oscillations.

AbstractThe paper evaluates the application of several multi-physical fields reconstruction approaches based on the Tomographic Particle Image Velocimetry (PIV) in three-dimensional supersonic flows with shock waves, including the pressure, temperature, and density. The MacCormack method and the Flux Vector Splitting method with outstanding performance in two-dimensional supersonic flow are extended into three-dimensional forms. The conventional Poisson method is also considered because of its widespread application and high accuracy in subsonic and transonic flow. All of these approaches are established by the conservative Navier-Stokes equations and solved by the time-marching iteration process, combined with the perfect gas law and adiabatic flow assumption. The performances are evaluated by numerical velocimetry data. To simulate the typical three-dimensional features, the cases of uniform freestream at Mach 0.78 and 3 past a cone with a 20° half-angle are selected to obtain a sufficient spanwise velocity component. The results confirm the feasibility of the PIV-based reconstruction methods in the conical flow field. The conventional Poisson method performs well only in the subsonic case while the Flux Vector Splitting method has a better performance in supersonic flow, including higher accuracy, stability, and efficiency, with a low-level root-mean-square error of 0.449% and local maximum relative error of 1%. [ABSTRACT FROM AUTHOR]