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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]

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 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]

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]

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]

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.

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]

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]

Introduction The aim of this article was to assess the flow capacity of end-to-side arterial anastomosis depending on the method of its implementation. Materials and Methods The study was conducted on 30 live Wistar rats in vivo, which were randomly divided into three groups. In each group of animals, an end-to-side microanastamosis was performed using three methods of donor artery preparation: 45 degrees (group A), 90 degrees (group B), and arteriotomy according to the "fish mouth" type (group C). The determination of flow capacity of anastomosis by measuring the blood volume flow with transonic flowmeter was performed. Results The obtained average values after the anastomosis were, respectively, 7.335 mL/s (standard deviation [SD]: 2.0771; min: 4.05; max: 10.85), 7.36 mL/s (SD: 0.836 mi: 6.15; max: 8.75), and 6.37 mL/s (SD: 1.247; min: 5.05; max: 9.05). No statistically significant difference in the blood volume flow velocity between all types of anastomoses was obtained (p = 0.251). Conclusion The flow capacity of end-to-side arterial anastomosis does not depend on the chosen method of anastomosis. [ABSTRACT FROM AUTHOR]

Some numerical simulations were used to study the effects of blowing ratio (0.25–0.5), hole diameters(0.65~1 mm), and hole inclination angle (30~60°) of the film cooling structure on the cooling and aerodynamic characteristics of the axisymmetric plug nozzle under the condition of transonic. The results showed that the bow shocks appear near the film holes on the plug wall in the supersonic region, which cause the wall cooling effectiveness of the plug to decrease. Compared with the baseline plug, the blowing ratio ranged from 0.25 to 0.5, the wall average temperature of the rear plug decreased by 34.4~48.1%, the thrust coefficient and total pressure recovery coefficient decreased by 0.31~0.61% and 0.52~0.93%, respectively. When the perforated percentage is constant, the wall cooling effectiveness increases with the decrease of the hole diameters. The increase in the inclination angle of the film holes lead to the decrease in cooling effectiveness and aerodynamic performance. This is because the penetration ability of the cooling air to the mainstream is enhanced, and the obstruction to the mainstream boundary layer is increased, resulting in the increase of bow shock intensity near the film holes in the supersonic region. [ABSTRACT FROM AUTHOR]

This study investigated the instability mechanisms within the blade-tip region of a transonic compressor. The variations in these mechanisms were explored under different operating conditions using methods such as spectral analysis and vortex identification. The results and analysis enhance the comprehension of self-excited unsteadiness in axial compressors, identifying its initiation characteristics and prevalence under different operational conditions. The regular occurrence of self-excited pulsations in transonic rotors was confirmed, especially when the flow rates fall blow specific levels, causing spontaneous pulsations within the tip region that are notably different from the blade-passing frequency. The findings highlight the crucial influence of interactions between separation flow, leakage flow, and shock waves in generating these pulsations. The effects of clearance size on flow structure intensity were also categorized, revealing three distinct self-excited unsteadiness mechanisms. These range from pulsations caused by shock–boundary-layer separation at small clearance sizes to those induced by interactions between leakage flow and shock waves at larger clearances, and finally to pulsations initiated directly by the clearance vortices at significantly larger clearances. This detailed analysis provides valuable insights into the dynamics of self-excited unsteadiness, supporting the advancement of the design and performance optimization of compressors. [ABSTRACT FROM AUTHOR]

The effects of back pressure on the transonic cascade operating state are crucial and can determine the structure of internal shock waves and secondary flows. In this paper, numerical methods validated by experiments were employed to investigate the evolution mechanisms of the inlet flow field, shock structure, secondary flow structure, and cascade performance under different back pressures. Analysis revealed that transonic cascade exhibited unique incidence characteristics in the inlet flow field under both subsonic and supersonic regimes, although these two regimes involved different physical mechanisms. The results revealed that the operating state of the transonic compressor cascade under the unique incidence condition was influenced by the outlet back pressure, and there existed a critical static pressure ratio. The critical static pressure ratio shifted from 1.61 for two-dimensional flow to 1.37 for three-dimensional (3D) flow at M1 = 1.1, due to the corner separation and the characteristics of 3D shocks. The 3D shock structure exhibited a non-uniform distribution along the spanwise direction due to the influence of back pressure and the separated boundary layer. The vortex structures analysis revealed that the secondary flow structure on the sidewalls of the transonic compressor cascade was primarily dominated by corner vortices, whose formation mechanism was related to the interaction between the shock wave and the sidewall boundary layer. Additionally, this interaction also led to the formation of detached shock and lip shock vortex structures. Finally, loss analysis indicated that the wake region of the transonic cascade primarily includes six types of loss, and the total loss of the cascade decreased with the rise in back pressure. [ABSTRACT FROM AUTHOR]

There has been a rapid advancement in deep learning models for diverse research fields and, more recently, in fluid dynamics. This study presents self-supervised transformers' deep learning for complex turbulent flow signals across various test problems. Self-supervision aims to leverage the ability to extract meaningful representations from sparse flow time-series data to improve the transformer model accuracy and computational efficiency. Two high-speed flow cases are considered: a supersonic compression ramp and shock-boundary layer interaction over a statically deformed surface. Several training scenarios are investigated across the two different supersonic configurations. The training data concern wall pressure fluctuations due to their importance in aerodynamics, aeroelasticity, noise, and acoustic fatigue. The results provide insight into transformers, self-supervision, and deep learning with application to complex time series. The architecture is extendable to other research domains where time series data are essential. [ABSTRACT FROM AUTHOR]

In this Letter, we have quantified the zones of turbulence in transonic shock diffraction across a 90 ° step corner using Lamb vector analysis. We have analyzed the nature of the three-dimensional turbulence structures using invariants of the velocity gradient tensor. From our results, we conclude that the vortex-induced shock, the lambda shock on the shear layer, and the Kelvin–Helmholtz instability forming on the shear layer are the main contributors in turbulence generation. In our study, we have shown that two-dimensional simulations can resolve the gasdynamic behavior accurately, though three-dimensional simulations are required to understand turbulent structures. [ABSTRACT FROM AUTHOR]

An experiment-supported simulation process chain is set up to perform numerical forced response analyses on a transonic high-pressure compressor front stage at varying operating conditions. A wake generator is used upstream of the rotor to excite a specific resonance within the operating range of the compressor. Thereby, extensive aerodynamic and structural dynamic experimental data, obtained from state-of-the-art rig testing at the Transonic Compressor Darmstadt test facility at the Technical University of Darmstadt, are used to validate numerical results and ensure realistic boundary conditions. In the course of this, five-hole-probe measurements at steady operating conditions close to the investigated resonance enable a validation of the steady aerodynamics. Subsequently, numerically obtained aeroelastic quantities, such as resonance frequency, and damping, as well as maximum alternating blade stresses and tip deflections, are compared to experimental blade tip timing data. Experimental trends in damping can be confirmed and better explained by considering numerical results regarding the aerodynamic wall work density and secondary flow phenomena. The influence of varying loading conditions on the resonance frequency is not observed as distinctly in numerical, as in experimental results. Generally, alternating blade stresses and deflections appear to be significantly lower than in the experiments. However, similar to the aerodynamic damping, numerical results contribute to a better understanding of experimental trends. The successive experimental validation shows the capabilities of the numerical forced response analysis setup and enables the highlighting of challenges and identification of potential further adaptations. [ABSTRACT FROM AUTHOR]

In rotating detonation engines the turbine inlet conditions may be transonic with unprecedented unsteady fluctuations. To ensure an acceptable engine performance, the turbine passages must be suited to these conditions. This article focuses on designing and characterizing highly diffusive turbine vanes to operate at any inlet Mach number up to Mach 1. First, the effect of pressure loss on the starting limit is presented. Afterward, a multi-objective optimization with steady RANS simulations, including the endwall and 3D vane design is performed. Compared to previous research, significant reductions in pressure loss and stator-induced rotor forcing are obtained, with an extended operating range and preserving high flow turning. Finally, the influence of the inlet boundary layer thickness on the vane performance is evaluated, inducing remarkable increases in pressure loss and downstream pressure distortion. Employing an optimization with a thicker inlet boundary layer, specific endwall design recommendations are found, providing a notable improvement in both objective functions. [ABSTRACT FROM AUTHOR]

In order to improve the data productivity of a wind tunnel test, the model under investigation in the wind tunnel is moved continuously with a predetermined constant angular speed in the so-called pitch–traverse mode. Alternatively, the wind tunnel model can be moved in the so-called pitch–pause mode, in which it keeps its position for a certain (measurement) time at a fixed pitch position, after which it is moved to the next pitch position. The latter procedure is more time-consuming, so, for the same time interval, the number of measured data points taken in the pitch–pause mode is less than that for the pitch–traverse mode. Since wind tunnel test time can be quite expensive, in most wind tunnel tests where only conventional forces and pressures are recorded with conventional measuring systems, the wind tunnel model is moved in the pitch–traverse mode in order to obtain as much aerodynamic data as possible during the tunnel runtime. The application of the Pressure-Sensitive Paint (PSP) technique has been widely used in wind tunnel testing for the purpose of providing pressure data on wind tunnel models with high spatial resolution. The lifetime-based PSP method has several advantages over the intensity-based method since it often has higher accuracy. Up until now, the lifetime-based PSP technique has mainly been used for wind tunnel testing, where the test model has been moved to the pitch–pause mode. The traditional lifetime method using on-chip accumulation requires multiple (~1000) excitation light pulses to accumulate enough luminescence (fluorescence or phosphorescence) photons on the camera sensor to provide acceptable signal-to-noise ratios and, therefore, it may seem to be not compatible with a continuously moving wind tunnel model. Nevertheless, the present study verifies the application of lifetime-based PSP utilizing on-chip accumulation with a continuously moving wind tunnel model which would make the entire PSP data acquisition compatible with that of the conventional measurements (forces and pressures), as mentioned above. In this paper, the applicability of the lifetime-based PSP technique to a continuously moving wind tunnel model (in pitch–traverse mode) is investigated with the help of measurements in the transonic wind tunnel in Göttingen (TWG). For this investigation, PSP was applied on the delta-wing model DLR-F22, which is to be tested in TWG. The pressure distribution on the wind tunnel model was measured using the PSP lifetime method for both model movement modes (pitch–pause and pitch–traverse mode) so that the corresponding PSP results could be directly compared with each other. In addition, an error analysis of the PSP results was carried out and compared with the conventional pressure measurement results, hence providing an assessment of the accuracy of the PSP results; finally, a recommendation for future PSP measurements could be given. [ABSTRACT FROM AUTHOR]

In this paper, we study the existence of transonic shocks in a cylindrical nozzle. More precisely, given the compatible outer normal momentum at the exit of a cylindrical nozzle, there exists a transonic shock solution such that the incoming supersonic flow is compressed to some subsonic flow. The shock position is uniquely determined by the exit normal momentum when we fix one point of the shock. This problem can be reduced to solve a nonlinear free boundary value problem for a hyperbolic–elliptic coupled system and the key ingredient is to solve a second‐order quasilinear elliptic equation with a nonlocal term by the method of separation of variables. [ABSTRACT FROM AUTHOR]

We investigate the driving of MHD turbulence by gravitational contraction using simulations of an initially spherical, isothermal, magnetically supercritical molecular cloud core with transonic and trans-Alfvénic turbulence. We perform a Helmholtz decomposition of the velocity field, and investigate the evolution of its solenoidal and compressible parts, as well as of the velocity component along the gravitational acceleration vector, a proxy for the infall component of the velocity field. We find that (1) In spite of being supercritical, the core first contracts to a sheet perpendicular to the mean magnetic field, and the sheet itself collapses. (2) The solenoidal component of the turbulence remains at roughly its initial level throughout the simulation, while the compressible component increases continuously, implying that turbulence does not dissipate towards the centre of the core. (3) The distribution of simulation cells in the B– ρ plane occupies a wide triangular region at low densities, bounded below by the expected trend for fast MHD waves (B ∝ ρ, applicable for high-local Alfvénic Mach number M A) and above by the trend expected for slow waves (B ∼ constant, applicable for low local M A). At high densities, the distribution follows a single trend |$B \propto \rho ^{\gamma _{\rm eff}}$|⁠ , with 1/2 < γeff < 2/3, as expected for gravitational compression. (4) The mass-to-magnetic flux ratio λ increases with radius r due to the different scalings of the mass and magnetic flux with r. At a fixed radius, λ increases with time due to the accretion of material along field lines. (5) The solenoidal energy fraction is much smaller than the total turbulent component, indicating that the collapse drives the turbulence mainly compressibly, even in directions orthogonal to that of the collapse. [ABSTRACT FROM AUTHOR]

The ability of streamwise-travelling waves of spanwise velocity to reduce the turbulent skin-friction drag is assessed in the compressible regime. Direct numerical simulations are carried out to compare drag reduction in subsonic, transonic and supersonic channel flows. Compressibility improves the benefits of the travelling waves, in a way that depends on the control parameters: drag reduction becomes larger than the incompressible one for small frequencies and wavenumbers. However, the improvement depends on the specific procedure employed for comparison. When the Mach number is varied and, at the same time, wall friction is changed by the control, the bulk temperature in the flow can either evolve freely in time until the aerodynamic heating balances the heat flux at the walls, or be constrained such that a fixed percentage of kinetic energy is transformed into thermal energy. Physical arguments suggest that, in the present context, the latter approach should be preferred. This provides a test condition in which the wall-normal temperature profile more realistically mimics that in an external flow, and also leads to a much better scaling of the results, over both the Mach number and the control parameters. Under this comparison, drag reduction is only marginally improved by compressibility. [ABSTRACT FROM AUTHOR]

Aeroelastic flutter is a dynamically complex phenomenon that has adverse and unstable effects on elastic structures. It is crucial to better predict the phenomenon of flutter within the scope of aircraft structures to improve the design of their wings. This review aims to establish fundamental guidelines for flutter analysis across subsonic, transonic, supersonic, and hypersonic flow regimes, providing a thorough overview of established analytical, numerical, and reduced-order models as applicable to each flow regime. The review will shed light on the limitations and missing components within the previous literature on these flow regimes by highlighting the challenges involved in simulating flutter. In addition, popular methods that employ the aforementioned analyses for optimizing wing structures under the effects of flutter--a subject currently garnering significant research attention--are also discussed. Our discussion offers new perspectives that encourage collaborative effort in the area of computational methods for flutter prediction and optimization. [ABSTRACT FROM AUTHOR]

This study presents a design optimization of a circumferential feedback channel on the shroud casing between rotor trailing and leading edges to maximize the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 37, using three-dimensional Reynolds-averaged Navier–Stokes analysis (RANS) with the k-ε turbulence model. The location of the injection and bleeding ports, angles of injection, and bleeding were the most influential factors on the aerodynamic performance and were selected as the design variables. A hybrid genetic algorithm and particle swarm optimization algorithm coupled with three different surrogate models were utilized to maximize the peak adiabatic efficiency and stall margin of the transonic axial compressor. The optimization results for the circumferential feedback channel indicated that an increase in stall margin was achievable with a slightly reduced peak efficiency in comparison with that of the smooth casing. [ABSTRACT FROM AUTHOR]

It is difficult to accurately predict the flow field over an aircraft in the presence of shock waves due to its strong nonlinear characteristics. In this study, we developed an accuracy-enhanced flow prediction method that fuses deep learning and a reduced-order model to achieve accurate flow field prediction for various aerodynamic shapes. Herein, we establish a convolutional neural network/proper orthogonal decomposition (CNN-POD) model for mapping geometries to the overall flow field. Then, local flow regions containing nonlinear flow structures can be identified by the POD reconstruction to build the enhanced model. A CNN model is established to map geometries to the local flow field. The proposed method was applied to two cases involving the prediction of transonic flow over airfoils. The results indicate that the proposed accuracy-enhanced flow prediction method can reduce the prediction error for flow properties in regions with nonlinear flow structures by values ranging from 13% to 66.27%. Additionally, the proposed method demonstrates better efficiency and robustness in comparison to existing methods, and it can also address the prediction problem of complex transonic flow with multiple strong nonlinear structures. [ABSTRACT FROM AUTHOR]

In this study, the interaction between a turbulent wake and the boundary layer of a horizontal tail plane (HTP) in the transonic flow regime is investigated. The setup considered corresponds to a generic tandem wing configuration with an OAT15A airfoil as the main wing and a National Advisory Committee for Aeronautics (NACA) 64A-110 as an HTP. Due to the transonic flow, the suction side of the OAT15A exhibits buffet. The numerical approach consists of a two-stage procedure in which a detached eddy simulation (DES) provides unsteady inflow conditions for a subsequent zonal high-fidelity large eddy simulation (LES) performed for the HTP region only; the turbulent boundary layer is modeled using scale-resolved wall-modeled LES (WMLES). The study mainly pursues two objectives: first, to discuss the influence of wake turbulence on the flow characteristics of the NACA airfoil; and second, to evaluate the capability of WMLES as a low-cost but high-resolution numerical approach in challenging flow conditions. Essentially, the study confirms the expected result that the outer part of the HTP's boundary layer is dominated by the wake of the main wing. Mainly based on a discussion of turbulence spectra, the study further demonstrates the advantage of WMLES over DES, proving that WMLES is able to capture the effects of wake turbulence on boundary layer dynamics, and thus validates the WMLES approach as a cost-effective, high-resolution turbulence modeling approach. On a superordinate level, the study further sketches a possible way on how a flow problem with such a strongly unsteady behavior could be systematically evaluated. [ABSTRACT FROM AUTHOR]

This study proposes a novel partial average fluctuation velocity (PAFV) turbulence based on the PAFV and the average strain rate tensor, which is used to calculate the shock wave boundary layer interaction flows, including cascade flow and transonic compressor flow. In order to illustrate the adequacy, the calculation results of several other turbulence models were also compared with PAFV, including k-ω, shear stress transport, SA-neg (negative Spalart-Allmaras), and stress-BSL (ω-based Reynolds stress model). The suitability of the PAFV turbulence model was initially tested by calculating the transonic flow in the L030-4 cascade. Subsequently, numerical computations were performed for the 3D (three-dimensional) National Aeronautics and Space Administration Rotor 67 transonic compressor rotor. In the numerical simulation, the flow control equations were transformed in a general curvilinear coordinate system for precision enhancement, and the spatial discretization of control equations was executed using the finite difference method. The convection term was processed using the Steger–Warming flux vector splitting method, the diffusion term was discretized using a second-order central difference scheme, and the time derivative term was discretized using the third-order Runge–Kutta method. A parallel computing strategy with multi-block partitioning was employed to speed up the calculations and facilitate faster convergence alongside a local time step method. All numerical computation procedures were written in Python, revealing the following: (1) The PAFV turbulence model aptly simulates the transonic flow within the cascade channel. The shock wave pattern in the cascade channel aligns well with the experimental schlieren images. (2) During the Rotor 67 transonic compressor rotor calculation, the flow–pressure ratio, flow efficiency, and inlet and outlet total temperatures and pressures align well with the experimental data. Furthermore, the resolution of the flow field structure related to the interaction between the tip leakage vortex and channel shock wave is high. (3) The turbulence model uses PAFV as the turbulence velocity scale, thereby suppressing the fluctuation velocity and turbulence viscosity near shock wave areas, preventing excessive turbulence viscosity. Moreover, PAFV inherently exhibits transport properties and anisotropic characteristics, making it well-suited for handling transonic compressor flows. [ABSTRACT FROM AUTHOR]

Unsteady pressure-sensitive paint (i-PSP) measurements were performed at a sampling rate of 30 kHz to investigate the near-endwall blade suction surface flow inside a low-pressure turbine cascade operating at engine-relevant high-speed and low-Re conditions. The investigation focuses on the interaction of periodically incoming bar wakes at 500 Hz with the secondary flow and the blade suction surface. The results build on extensive PIV measurements presented in the first part of this two-part publication, which captured the 'negative-jet-effect' of the wakes throughout the blade passage. The surface pressure distributions are combined with CFD to analyze the flow topology, such as the passage vortex separation line. By analyzing data from phase-locked PIV and PSP measurements, a wake-induced moving pressure gradient negative in space and positive in time is found, which is intensified in the secondary flow region by 33% with respect to midspan. Furthermore, two methods of frequency-filtering based on FFT and SPOD are compared and utilized to associate a pressure fluctuation peak around 678 Hz with separation bubble oscillation. [ABSTRACT FROM AUTHOR]

The compressor serves as a crucial component that influences the performance of the gas turbine engine. Researchers have been endeavoring to explore compressor types that possess a high loading level and high-efficiency characteristics concurrently. In this study, tandem blade technology was applied to a transonic ultra-highly loaded axial compressor, and the Baseline single-blade rotor was replaced by a tandem rotor to take into account the loading level and compressor performance. Detailed investigations were carried out to identify the effects on the aerodynamic performance of the ultra-highly loaded stage and the fundamental flow mechanism within the tandem-rotor stage. This paper presents original design maps for the tandem-rotor stage, and the selection criteria for tandem parameters in tandem-rotor stage are refined. The results indicate that the peak efficiency improved by 0.83%, the stall margin increased by 2.16%, and the choke flow rate rose by 0.30% for the optimal tandem-rotor configuration. The meridional division position of the rotor primarily affects the ratio of loading of the front and rear blades, while the circumferential relative position of the tandem rotor mainly influences the channel types formed by the front and rear blades. Larger values for the meridional division position parameter and smaller values for circumferential relative position parameter should be selected for the tandem rotor design to optimize both the isentropic efficiency and total pressure ratio. This investigation offers the theoretical foundation for the design of a transonic ultra-highly loaded tandem-rotor compressor. [ABSTRACT FROM AUTHOR]

The capability for solving compressible fluid flows in the rotating frame of reference is added to an existed open-source CFD solver, namely, HiSA solver. The new implementation is explained and validated using the experimental data of the Sikorsky S-76 rotor. A comparison is presented between the moving mesh results obtained from the original HiSA code and the single rotating frame results achieved through the new implementation. The comparison includes an analysis of torque and thrust values, as well as computational costs. The results imply that, for evaluating the performance of an isolated rotor or for shape optimization purposes at the transonic regime, the single rotating frame method, like the one introduced in the current work, can provide accurate results within an acceptable computational budget. Furthermore, the results show that, at least 25 revolutions are required for the transient analysis to reach an acceptable steady-state converged solution like the one obtained by the single rotating frame method. [ABSTRACT FROM AUTHOR]

In variable geometry turbine vanes, tip clearance height and shape vary with the rotation of the vane, which affect the aerodynamic performance significantly. However, these issues are rarely considered in published studies. The current paper investigated the flow field features of transonic variable geometry turbine vanes with non-uniform partial clearance induced by the vane rotating. The results show that: The influence of guide vane rotation on the clearance height and its distribution cannot be ignored. At the same turning angle, the maximum clearance difference is up to 0.79 mm (0.8 % vane height). The height and shape variation of the non-uniform clearance leads to the change in the leakage flow rate, secondary flow structure, and aerodynamic loss of the variable guide vane. Under the combined effect of pressure difference on both sides of the clearance, axial and circumferential non-uniformity of clearance height, the total pressure loss coefficient is up to 9.44 % when the turning angle is −10°. The effect of the pivot on the clearance flow was also analyzed. The pivot increases the pressure in the gap flow field and reduces leakage flow velocity. However, a backflow region appears at the suction side of the pivot, which increases the aerodynamic losses. [ABSTRACT FROM AUTHOR]

To reveal the influence of distortion of S-duct on the flow field and performance of transonic fan, full annulus unsteady numerical simulation was carried out under S-duct/fan integrated condition. This study focuses on the coupling flow of S-duct/fan integrated condition under peak efficiency (PE) condition and near stall (NS) condition. Results show that, compared with the condition that S-duct or fan is investigated alone, the distortion of S-duct is suppressed under S-duct/fan integrated condition. The curved structure of S-duct is the important reason for the flow migration from undistorted region to distortion region. Higher inlet relative Mach number and more work input of the rotor are observed in the counter-swirl region. The flow separation in S-duct is weakened under NS condition compared with that of PE condition. Counter-swirl region of blade tip has larger region of blockage, and rotating stall inception is most likely to occur in this region. [ABSTRACT FROM AUTHOR]

In this work, a combined design of blade and endwall using the Extended Free Form Deformation technique is proposed and implemented on a transonic compressor rotor (the NASA Rotor 67). The best combined design is explored by numerical investigations and optimization via Kriging surrogate model method. As a result, by only modifying the hub and blade under 15 % span, the peak adiabatic efficiency is increased by 0.4 % and the overall aerodynamic performance for all operating conditions is also improved. An analysis shows that the enlarged dihedral angle between the hub and blade restrains boundary layer intersection and development. Meanwhile, the optimized hub changes the pressure distribution and prevents migration of cross-flow. As a result, the low-energy flow in the corner is accelerated downstream and the separation is limited, leading to an effective improvement in throughflow capability and aerodynamic performance in the corner region. [ABSTRACT FROM AUTHOR]

Background: The same principle behind pulse wave analysis can be applied on the pulmonary artery (PA) pressure waveform to estimate right ventricle stroke volume (RVSV). However, the PA pressure waveform might be influenced by the direct transmission of the intrathoracic pressure changes throughout the respiratory cycle caused by mechanical ventilation (MV), potentially impacting the reliability of PA pulse wave analysis (PAPWA). We assessed a new method that minimizes the direct effect of the MV on continuous PA pressure measurements and enhances the reliability of PAPWA in tracking beat-to-beat RVSV. Methods: Continuous PA pressure and flow were simultaneously measured for 2–3 min in 5 pigs using a high-fidelity micro-tip catheter and a transonic flow sensor around the PA trunk, both pre and post an experimental ARDS model. RVSV was estimated by PAPWA indexes such as pulse pressure (SVPP), systolic area (SVSystAUC) and standard deviation (SVSD) beat-to-beat from both corrected and non-corrected PA signals. The reference RVSV was derived from the PA flow signal (SVref). Results: The reliability of PAPWA in tracking RVSV on a beat-to-beat basis was enhanced after accounting for the direct impact of intrathoracic pressure changes induced by MV throughout the respiratory cycle. This was evidenced by an increase in the correlation between SVref and RVSV estimated by PAPWA under healthy conditions: rho between SVref and non-corrected SVSD – 0.111 (0.342), corrected SVSD 0.876 (0.130), non-corrected SVSystAUC 0.543 (0.141) and corrected SVSystAUC 0.923 (0.050). Following ARDS, correlations were SVref and non-corrected SVSD – 0.033 (0.262), corrected SVSD 0.839 (0.077), non-corrected SVSystAUC 0.483 (0.114) and corrected SVSystAUC 0.928 (0.026). Correction also led to reduced limits of agreement between SVref and SVSD and SVSystAUC in the two evaluated conditions. Conclusions: In our experimental model, we confirmed that correcting for mechanical ventilation induced changes during the respiratory cycle improves the performance of PAPWA for beat-to-beat estimation of RVSV compared to uncorrected measurements. This was demonstrated by a better correlation and agreement between the actual SV and the obtained from PAPWA. [ABSTRACT FROM AUTHOR]

This work is a numerical and experimental study of a rectangular thin plate undergoing stall flutter at Mach 0.8. This constitutes one of the first studies of this kind where three-dimensionality is fully implemented in a numerical simulation including the test-section effects characterizing wind-tunnel experiments. In order to break down the fluid–structure interaction to its main driving phenomena, an aerodynamic model is proposed that is based on computationally inexpensive steady-state simulations. Two types of dynamic instability are observed in the numerical simulations; Flutter by mode coalescence is promoted at zero flow incidence, however, high bending precludes this from happening for higher values of angle of attack. Stall flutter is instead a nonlinear one-degree type of instability. Both of these instability mechanisms can be explained in terms of hysteretic behaviour of the pressure distribution, which becomes more pronounced at high angles of attack, when a large separation region is formed. Tests were conducted employing titanium alloy plates in order to survive the aerodynamic loads characterizing the wind-tunnel initial transient. However, due to wall interference, high bending was promoted so that the internal stress exceeded the yield values before flutter could be measured. Numerical simulations were in general agreement with the experiment in terms of both amplitude and oscillation frequency. [ABSTRACT FROM AUTHOR]

This paper describes a new method for performing the time-frequency analysis of particle image velocimetry data with account for surface constant-temperature anemometer sensor readings. This method is based on calculating correlation coefficients between particle image velocimetry data and constant-temperature anemometer data, previously filtered in a given frequency range. This approach makes it possible to obtain data on the unsteady characteristics of flows throughout the entire measurement range in a frequency range whose boundaries extend beyond the boundaries of the particle image velocimetry frequency range. The proposed method is applied to study nonstationary processes in a region where a shock wave interacts with a boundary layer at a Mach number M = 1.43. [ABSTRACT FROM AUTHOR]

This research paper presents a thorough examination of shock wave train phenomena in various duct structures, utilizing a machine learning-optimized k-ω model. The focus of the study is to apply machine learning techniques to adapt the constant coefficients of the k-ω turbulence model, improving its accuracy and computational efficiency in capturing the dynamics of shock waves. An important aspect of this investigation is the impact of diverging sections within the duct, specifically how changes in the divergence angle, while maintaining a constant ratio of the exit area to the throat area, affect compressible flow parameters, shock wave positions, and other related characteristics of shock trains. The study systematically explores these effects and provides fresh insights into the behavior of shock wave trains under different divergence conditions. In a departure from the usual practices in supersonic research, this study compares the phenomena of shock wave trains in rectangular and circular duct geometries. The findings contribute valuable data and analyses to a field that has traditionally focused on rectangular configurations. The results indicate that the shape of the duct has a significant influence on the shock wave train, emphasizing the importance of considering geometric diversity in the study of supersonic flows and shock wave phenomena. This research sets the stage for further investigations into the interaction between duct geometry, shock wave propagation, and the dynamics of compressible flows. • Improved k-ω SST model accuracy to 95.5 % with machine learning. • Longer duct diverging sections intensify the first lambda shock. • Duct shape critically affects shock wave and boundary layer interactions. [ABSTRACT FROM AUTHOR]

The aim of this paper is to develop a profile of axisymmetric minimum length nozzle giving a uniform and parallel flow at the exit section. The study is done at high temperature, lower than the dissociation threshold of the molecules. The design is made by the method of characteristics (MOC). The variation of the specific heats with the temperature is considered. The numerical results have been validated with CFD simulation Ansys-Fluent software. The second part of this study is to calculate and analyze the transonic flow field of this supersonic nozzle. The computation of the flow field characteristics at the throat is thus essential to the nozzle developed thrust value and therefore to the aircraft or rocket it propels. An investigation was conducted to analyze the effects of parameters on the position of the sonic line. These parameters include stagnation temperature T0, radius of the nozzle, types of gases, and exit Mach number ME. [ABSTRACT FROM AUTHOR]

A series of implicit large eddy simulations have been conducted to implement thermal control on the transonic shock-boundary layer interactions (SBLIs) over a natural laminar flow airfoil, operating beyond the drag divergence Mach number. The study focuses on the SHM1 airfoil, where the baseline flow exhibits shock-induced separation under specific conditions: free stream Mach number M ∞ = 0.78 , angle of attack α = 0.38 ° , and Reynolds number R e = 8 × 10 6 . A time-periodic surface heat flux is introduced, strategically located near the shock structures of the unaltered flow, and the impact of heating vs cooling is investigated through instantaneous Schlieren visualizations and vorticity dynamics, and time-averaged load distributions and boundary layer parameters. Time-averaged Mach contours are utilized to measure the shock strength and extent, revealing that thermal control effectively mitigates the detrimental impact of transonic SBLI on the airfoil's performance. Cooling control emerges as the more effective method, and a case featuring multiple cooling controls near the shock structures demonstrates superior efficacy in controlling shock waves and suppressing shock-induced separation. This optimized configuration results in an improved aerodynamic efficiency of 12.65% compared to the baseline flow. [ABSTRACT FROM AUTHOR]

This article presents single- and multi-disciplinary shape optimizations of a generic business jet wing at two transonic cruise flow conditions. The studies performed are based on two high-fidelity gradient-based optimization tools, assisted by the adjoint method (following both discrete and continuous approaches). Single discipline and coupled multi-disciplinary sensitivity derivatives computed from the two tools are compared and verified against finite differences. The importance of not making the frozen turbulence assumption in adjoint-based optimization is demonstrated. Then, a number of optimization runs, ranging from a pure aerodynamic with a rigid structure to an aerostructural one exploring the trade-offs between the involved disciplines, are presented and discussed. The middle-ground scenario of optimizing the wing with aerodynamic criteria and, then, performing an aerostructural trimming is also investigated. [ABSTRACT FROM AUTHOR]

The presence of bleed in an aero engine's compressor can significantly impact its flow characteristics and contribute to rotating instability. This study focuses on the impact of a typical cavity bleed structure on the internal flow characteristics of a compressor, specifically its circumferential non-uniformity and aerodynamic stability. A numerical simulation involving multiple flow passages was conducted on the transonic high-pressure rotor of the E3 compressor, considering its typical bleed structure. The study delves deep into the non-uniform flow characteristics and the mechanisms behind their generation in the compressor flow field. Furthermore, the influence of bleed on the rotating instability of the compressor is explored by comparing changes in compressor instability and adiabatic efficiency under uniform and non-uniform flow field conditions. The findings indicate that the axial position of the cavity bleed structure plays a crucial role in influencing key parameters such as rotor stall margin, peak efficiency, and total pressure ratio under near-stall conditions. The circumferential non-uniformity, resulting from the presence of the cavity bleed, intensifies with higher bleed air flow rates. For the upstream bleed configuration applied to the rotor, with a total bleed rate of 5%, the maximum variation in absolute flow angle at the inlet of different rotor channels can reach up to 1°. Additionally, the maximum difference in inlet flow coefficient can reach 0.0392. These findings demonstrate that the non-uniformity caused by the typical bleed structure leads to a loss in stall margin for the rotor when compared to a uniform flow field scheme. [ABSTRACT FROM AUTHOR]

A reduced-order model based on the proper orthogonal decomposition (POD) method is used to assess whether it can predict the highly unsteady flow in turbomachinery cascades. Three cases are considered: a transonic fan rotor called the National Aeronautics and Space Administration Rotor 67, a cascade of oscillating compressor blades called the 10th Standard Configuration, and a cascade of oscillating turbine blades called the 11th Standard Configuration. The governing equations used in this POD model, called zeta-POD, are written using the specific volume instead of density. This formulation of the governing equations allows the precomputation of the coefficients of the system of ordinary differential equations prior the flow simulation. Consequently, a significant speedup is obtained compared to the full-order model and a traditional, density-based POD. In addition, the accuracy of the zeta-POD model is compared against the full-order model and the density-based POD. [ABSTRACT FROM AUTHOR]