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The canonical shock/isotropic turbulence interaction (SITI) at high shock Mach numbers (Ms) is studied by conducting direct numerical simulation (DNS) for thermally perfect gas (TPG) and calorically perfect gas (CPG). Combining DNS with linear interaction analysis (LIA), the amplification of vorticity variance across the shock wave is studied. It is found that the changes in vortical velocity fluctuation amplitude and turbulent length scales under vibrational excitation have a competitive effect on vorticity amplification. The latter is dominant and leads to the transverse vorticity amplification increasing by 32.2% at Ms = 6.0. Based on the LIA theory, a vorticity amplification model for SITI considering vibrational excitation is established. Furthermore, the impact of vibrational excitation on the downstream vorticity transportation is examined through an analysis of the transport equation. The vibrational excitation strengthens both the vortex stretching and viscous dissipation of streamwise vorticity but only alters the viscous dissipation of transverse vorticity. Then, the vorticity transportations of different turbulent structures for CPG and TPG are compared. The comparison indicates that the increment of vortex stretching for streamwise vorticity variance is sustained by the enhanced turbulent structures corresponding to the stable-node/saddle/saddle, and the rapid decay of transverse vorticity variance for TPG is associated with the enhanced viscous dissipation of the nonfocal turbulent structure. [ABSTRACT FROM AUTHOR]

Wall pressure fluctuations caused by turbulent boundary layers have a significant impact on aircraft structural vibration and cabin noise. This study aims to investigate the mechanism of turbulence-induced pressure fluctuations by focusing on the randomness of wall pressure fluctuations, analyzed in both the time–frequency and spatial-wavenumber domains using measured data obtained from a phase array in a wind tunnel. Three roughness elements were designed and installed upstream of the plate to manipulate the turbulent boundary layer at a specific Mach number. The results of the investigation demonstrate that the disturbance strength induced by the roughness element influences the randomness of wall pressure fluctuations, in addition to the parameters utilized for data analysis. Generally, stronger turbulence fluctuations tend to decrease the randomness of pressure fluctuations. Moreover, wall pressure fluctuations also exhibit certain statistical principles that cannot be precisely calculated using mathematical expressions, highlighting their inherent randomness. Further investigation into randomness in the spatial-wavenumber domain revealed the hydrodynamic modes of turbulence fluctuations with varying convection velocity analyzed through wavenumber maps computed using the beamforming algorithm. These modes with variable convective speed significantly contribute to the generation of randomness in wall pressure fluctuations. Both the time–frequency domain and the spatial-wavenumber domain affect the randomness characteristics of wall pressure fluctuations. However, such effects are not easily discernible through a rudimentary analysis of the space–time correlation of turbulence fluctuations. [ABSTRACT FROM AUTHOR]

The linear spatial stability of a parallel two-dimensional (2D) compressible boundary layer on an adiabatic plate is investigated by considering both 2D and three-dimensional (3D) disturbances. The compound matrix method is employed here, for the first time, for compressible flows, which, unlike other conventional techniques, can efficiently eliminate the stiffness of the equations governing the spectral amplitudes. The method is first validated with published results in the literature corresponding to spatial and temporal instability of flows ranging from low subsonic to high supersonic Mach numbers (M), which shows a good match depending upon the proper choice of free-stream temperature and the wall dispersion relation. Subsequently, flow compressibility effects and the spanwise variation of disturbances are also investigated for M ranging from low subsonic to high supersonic cases (from M = 0.1 to 6). Mack (AGARD Report No. 709, 1984) reported the existence of two unstable modes for M > 3 from viscous calculations (the so-called "second mode") that subsequently fuse to create only one unstable zone when M increases. Our calculations show a series of higher-order unstable modes for M > 3 in addition to the findings of Ma and Zhong ["Receptivity of a supersonic boundary layer over a flat plate. Part 1. Wave structures and interactions," J. Fluid Mech. 488, 31–78 (2003)], where such higher-order modes for supersonic boundary layers are all noted to be spatially stable. The number and the frequency extent of the corresponding unstable zones increase with an increase in M beyond 3 while propagating downstream at a higher speed than those corresponding to incompressible, subsonic, and low supersonic (M < 2) cases. [ABSTRACT FROM AUTHOR]

In order to investigate the effects of the throat heights of the air annular slot on the rotating detonation waves, experiments have been carried out while varying the throat height of the air annular slot (i.e., 0.5, 1.5, 2.5, and 4.0 mm), the mass flow rate of oxidizer (35–300 g/s), and the equivalence ratio (0.15–1.8). Air and ethylene have been utilized as oxidizer and fuel, respectively. The experimental results indicate that rotating detonation waves can be obtained under different inflow conditions, and two operating phenomena were observed, i.e., the detonation cases (the single-wave mode and the oscillation mode) and the deflagration cases (the deflagration mode and the chaotic mode). The effects of the throat height on the operating range of stable rotating detonations are related to the minimum mass flow rate of oxidizer. As the throat height increases, the minimum mass flow rate of oxidizer also increases. Furthermore, the relationship between the minimum Mach number to obtain rotating detonations at each throat height and the throat height has been obtained. Moreover, the effects of the eccentricity ratio of the air annular slot on the propagation modes have also been demonstrated. [ABSTRACT FROM AUTHOR]

Wedge tail airfoils refer to the addition of a wedge section to the tail of an airfoil. This modification has demonstrated greater efficiency in increasing the lift forces compared to the base airfoil design. Such airfoils are particularly important in scenarios where increased lift forces are required, such as in the fast and efficient maneuvering of marine and aerospace vehicles. This study aimed to predict the flow fields and aerodynamic coefficients of wedge tail airfoils using a multi-head perceptron (MHP) network and classical machine learning (ML) algorithms, including k-nearest neighbors, decision tree, and random forest. These predictions were based on airfoil sections, x–y grid coordinates, Mach number, and angle of attack, eliminating the need to solve the Navier–Stokes (NS) equations. The database required for training the MHP network and the ML models were generated using the Ansys solver, which solved the NS equations. The results of the models were compared, and it was observed that the MHP network consistently outperformed the others in terms of prediction accuracy. Additionally, the prediction process demonstrated a significant speed improvement of 125 times compared to conventional computational fluid dynamics techniques. Once the training was completed, the flow fields and aerodynamic coefficients for a wedge tail airfoil could be obtained within seconds. [ABSTRACT FROM AUTHOR]

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]

The basic philosophy of two-stage blast interrupter is combination of self-blast and puffer-interrupter principles. Gas circuit breaker (GCB) with two-stage blast interrupter has two gas chambers. High gas pressure in one chamber generates by using arc energy and in other chamber, by mechanical compression. In such type of breakers, interruption becomes quite difficult at low currents and hence optimization of gas chambers design is essential. In view of above, a numerical program has been developed to calculate SF6 gas pressure-rise in various interrupter volumes and gas flow rates across different passages of interrupter w.r.t. time. To validate model developed in the study, an experimental set-up is established and measured transient gas pressure-rise generated during 420 kV GCB operation. The effect of speed-time characteristics, stroke length and piston diameter on performance parameters of 420 kV GCB model is analyzed. Further, to understand gas flow pattern across insulated nozzle during GCB operation, authors developed a CFX model. By using this numerical model, effect of nozzle profile, gas flow passage areas and profile of gas discharge vents on transient gas pressure and Mach number distribution across inter-electrode gap and gas flow rates through various outlets has been discussed in detail in the paper. [ABSTRACT FROM AUTHOR]

Within the dynamic realm of variable-geometry shock wave/boundary layer interactions, the wall parameters of the flow field undergo real-time fluctuations. The conventional approach to sensing these changes in wall pressure through sensor measurements is encumbered by a cumbersome process, leading to diminished efficiency and an inability to provide swift predictions of wall parameters. This paper introduces a data-driven methodology that leverages non-contact schlieren imaging to predict wall pressure within the flow field, a technique that holds promise for informing the optimized design of variable-geometry systems. A sophisticated deep learning framework, predicated on Convolutional Neural Networks (CNN), has been engineered to anticipate alterations in wall pressure stemming from high-speed shock wave/boundary layer interactions. Utilizing an impulsive wind tunnel with a Mach number of 6, we have procured a sequence of schlieren images and corresponding wall pressure measurements, capturing the continuous variations induced by an attack angle from a shock wave generator. These data have been instrumental in compiling a comprehensive dataset for the training and evaluation of the CNN. The CNN model, once trained, has adeptly deduced the distribution of wall pressure from the schlieren imagery. Notwithstanding, it was observed that the CNN's predictive prowess is marginally diminished in regions where pressure variations are most pronounced. To assess the model's generalization capabilities, we have segmented the dataset according to different temporal intervals for network training. Our findings indicate that while the generalization of all models crafted was less than optimal, Model 4 demonstrated superior generalization. It is thus suggested that augmenting the training set with additional samples and refining the network architecture will be a worthwhile endeavor in subsequent research initiatives. [ABSTRACT FROM AUTHOR]

This work investigates the effectiveness of a shape-memory alloy (SMA) in controlling the instabilities of triangular composite plates under supersonic flow. Lagoudas' quadratic polynomial hardening theory models the SMA effect. First-order piston theory was used for the aerodynamic modeling, and the reference-temperature method was used for modeling the thermal heating. The buckling and post-buckling behaviors were studied for different boundary conditions with four different layups. In addition, buckling and post-buckling of the composite plate, with and without shape memory alloy wire, has been studied. The effect of SMA wire on aeroelastic instabilities is accurately studied. The embedded SMA wire significantly increased the stability region (postpone divergence and flutter velocities) and buckling temperature. Also, the time responses of the triangular composite plate are determined at different Mach numbers, showing that by increasing the Mach number, the SMA wire can control or decrease the vibration amplitudes. [ABSTRACT FROM AUTHOR]

Experiments have been carried out to investigate the over expanded free jet at Mach number 2.0 without and with tabs of different cross-sectional shapes, tab length penetration into the jet core defined as blockage ratio, and symmetrical and asymmetrical placement of tabs at the jet exit. The cross-sectional shapes such as triangular, square, and circular have been studied with the penetration depth of 3 %, 7 %, and 11 % blockage. The Pitot tube and Schlieren flow visualization were carried out in experiments. The basic features of the jet with Mach disk, shock cell, and its distortion with the adoption of tabs at various nozzle pressure ratios indicate a definite influence that helps in the reduction of the core jet length. This characteristic is also influenced by changes in tab cross-section and length of penetration. Since the basic jet has three-dimensional structures, the asymmetric placement of tabs is beneficial compared to the symmetric orientation of tabs. The jet width increases with adoption of these tabs and different shapes. Whereas, a maximum core jet length reduction of the order of 60 % could be achieved using the present techniques. [ABSTRACT FROM AUTHOR]

Understanding the occurrence of various feedback mechanisms of an under-expanded impinging supersonic jet is a crucial task in research. The presence of several jet modes is examined in this study for the flat and corrugated impinging plate geometries. The behavior of impinging plate configurations during mode switching is investigated by varying the flow state, such as the jet Mach number. The staging behavior at various jet Mach numbers is observed using acoustic spectral plots and schlieren flow visualization. To explore the presence of various types of modes during the jet impingement due to the modification of jet Mach number, ensemble averaging and Proper Orthogonal Decomposition of schlieren images are carried out. In the majority of situations, the corrugated design shows a reduction in tonal noise and overall sound pressure level. In exceptional cases, for the corrugated plates, the enhanced overall sound pressure level is caused by the existence of axisymmetric instability (A1, A2). [ABSTRACT FROM AUTHOR]

In this study, a CFD computational model has been established to properly simulate complex correctly expanded supersonic flow generated by a three-dimensional C–D nozzle for the purpose of Thrust Vector Control (TVC) simulations. Jet tabs affixed to the perimeter of the nozzle exit are used to achieve thrust vectoring. This study focuses on the estimation of the force on the tabs due to the jet issued from the C–D nozzle using Ansys 19.2 three-dimensional density based solver, implicit formulation with SST K–ω turbulence model. The force is calculated for different tab configurations based on tab sector angles of 60°, 90°& 120° for the designed Mach number and X/D ratio of 0.7. The nozzles with Mach 1.6, Mach 1.8 and Mach 2.0 designed using SolidWorks are compared. It has been observed that with an increase in sector angle, tab force increased by an average of 70 % for sector angle of 60° to 90° and 36 % for sector angle of 90° to 120°. [ABSTRACT FROM AUTHOR]

The effect of nozzle lip thickness and velocity ratio on coaxial subsonic jet mixing, at different Mach numbers, has been studied experimentally and numerically. Decay of coaxial subsonic jets emanating from coaxial nozzles of lip thickness 0.7, 1.7 and 2.65 mm with velocity ratio (VR) from 0.2 to 1.0 at primary jet exit Mach numbers of 0.6, 0.8 and 1.0 has been studied. Free jet without co-flow (VR0) was also studied for comparison. Jet centerline Mach number decay, turbulence and velocity variation in the radial direction are analyzed. The results show that mixing the coaxial jet at a low-velocity ratio is better than a high-velocity ratio, at all Mach numbers of the present study. The nozzle lip thickness has a significant influence on the secondary jet. Mixing of the jet in the presence of VR0.2 coaxial jet is found to be the highest. Characteristic decay of Mach 0.8 and 1.0 jet for lip thickness 1.7 and 2.65 mm is faster than lip thickness 0.7 mm. For a given lip thickness, increasing of velocity ratio is found to retard the mixing between primary and secondary jets. [ABSTRACT FROM AUTHOR]

This paper presents the numerical analysis of a convergent-divergent circular nozzle with the exit Mach number of 1.69 with and without passive control at the exit. The passive control method opted for this analysis was inward and outward ascending triangular protrusion. This paper explores the influence of the passive control geometry and its blockage area concerning the nozzle exit. The nozzle pressure ratio (NPR) used for carrying out the flow analysis were 3, 4.932, and 6. Two different inward and outward protrusions were used with a height of 1.5 mm and 3 mm. From the results, the potential core length of the protrusion 1.5 mm height was not much changed in the both outward and inward cases. But when the height of the protrusion was increased to 3 mm, there was a noticeable core length reduction at all NPR but with different cases. At the NPR of 6, the potential core length of the inward protrusions 3 mm was reduced by 44 % compared to the plain CD nozzle. [ABSTRACT FROM AUTHOR]

This study investigates the application of deep learning models—specifically Deep Neural Networks (DNN), Long Short-Term Memory (LSTM), and Long Short-Term Memory Neural Networks (LSTM-NN)—to predict panel flutter in aerospace structures. The goal is to improve the accuracy and efficiency of predicting aeroelastic behaviors under various flight conditions. Utilizing a supersonic flat plate as the main structure, the research integrates various flight conditions into the aeroelastic equation. The resulting structural vibration data create a large-scale database for training the models. The dataset, divided into training, validation, and test sets, includes input features such as panel aspect ratio, Mach number, air density, and decay rate. The study highlights the importance of selecting appropriate hidden layers, epochs, and neurons to avoid overfitting. While DNN, LSTM, and LSTM-NN all showed improved training with more neurons and layers, excessive numbers beyond a certain point led to diminished accuracy and overfitting. Performance-wise, the LSTM-NN model achieved the highest accuracy in classification tasks, effectively capturing sequential features and enhancing classification precision. Conversely, LSTM excelled in regression tasks, adeptly handling long-term dependencies and complex non-linear relationships, making it ideal for predicting flutter Mach numbers. Despite LSTM's higher accuracy, it required longer training times due to increased computational complexity, necessitating a balance between accuracy and training duration. The findings demonstrate that deep learning, particularly LSTM-NN, is highly effective in predicting panel flutter, showcasing its potential for broader aerospace engineering applications. By optimizing model architecture and training processes, deep learning models can achieve high accuracy in predicting critical aeroelastic phenomena, contributing to safer and more efficient aerospace designs. [ABSTRACT FROM AUTHOR]

The accurate prediction of blade loading is crucial to designing efficient turbomachinery, but traditional methods often neglect the impact of compressibility, leading to inaccuracies at high speeds. This study investigates the effect of compressibility on the blade loading parameters, particularly the Zweifel coefficient (Zw) and lift coefficient (CL), in turbine cascades. A novel intermediate method (IM), with averaged flow properties derived from both inlet and outlet conditions, is proposed to enhance the accuracy of Zw and CL calculations in compressible flow regimes. This method is based on the extended Kutta–Joukowski theorem for compressible flow and incorporates the Mach number directly into the modified definitions of Zw and CL. The analysis reveals that the averaged flow angle (αm), calculated by using a velocity-weighted approach, serves as a crucial parameter for blade similarity studies. The proposed correction method is applied and validated based on CFD simulations of the VKI-RG turbine cascade. The IM and modified definitions provide a robust framework for accurately predicting blade loading at high speeds, enabling improved design and analysis of turbomachinery. [ABSTRACT FROM AUTHOR]

This study seeks to develop a fundamental comprehension of the noise challenges encountered by commercial aircraft fuselage surface attachments, such as blade antennas and pitot tubes. The study examines the flow characteristics and far-field noise directivity of a wall-mounted NACA0012 airfoil with various sweep angles (−35°, −15°, 0°, +15°, and +35°) and an aspect ratio of 1.5. The Mach numbers of the incoming flow range from 0.8 to 0.9 with a Reynolds number of about 7 × 105. Delayed Detached Eddy Simulation (DDES) and the Ffowcs Williams–Hawkings (FW-H) equation are utilized. The results show that the shock wave intensity at the junction between the airfoil and the bottom wall is enhanced by the forward-swept angle. The shock wave moves and changes into a λ-type structure, while the boundary layer separates and produces shedding vortices in the junction at a smaller Mach number on the forward-swept airfoil compared to the straight airfoil and the backward-swept airfoil. These phenomena cause significant surface pressure fluctuations in the junction and result in a significant dipole noise in the far field, which is the primary source of noise in the far field. In addition, the normal Mach number and the absolute sweep angle also contribute to the far-field noise. [ABSTRACT FROM AUTHOR]

Aircraft secondary flow systems are small-flow circulation devices that are used for thermal and cold management, flow control, and energy generation on aircraft. The aerodynamic characteristics of main-flow-based inlets have been widely studied, but the secondary-flow-based small inlets, jets, and blowing and suction devices have seldom been studied. Two types of secondary flow systems embedded in a supersonic aircraft wing, a ram-air intake and a submerged intake, are researched here. Firstly, wind tunnel tests under subsonic, transonic, and supersonic conditions are carried out to test the total pressure recovery and total pressure distortion. Secondly, numerical simulations are used to analyze the flow characteristics in the secondary flow systems. The numerical results are validated with experimental data. The calculating errors of the total pressure recovery on the ram-air and submerged secondary flow systems are 8% and 10%, respectively. The simulation results demonstrate that the total pressure distortion tends to grow while the total pressure recovery drops with the increasing Mach number. As the Mach number increases from 0.4 to 2, the total pressure recovery of the ram-air secondary flow system decreases by 68% and 71% for the submerged system. Moreover, the total pressure distortion of the ram-air and submerged secondary flow systems is increased by 19.7 times and 8.3 times, respectively. Thirdly, a detailed flow mechanism is studied based on the simulation method. It is found that the flow separation at the front part of the tube is induced by adverse pressure gradients, which primarily determine the total pressure recovery at the outlet. The three-dimensional vortex in the tube is mainly caused by the change in cross-sectional shape, which influences the total pressure distortion. [ABSTRACT FROM AUTHOR]

In order to determine the installation position of the front fin on the example rocket forebody, an optimized method based on a comprehensive evaluation indicator and variable step search is presented. The comprehensive indicator consists of four weight coefficients, two lateral aerodynamic forces and two aerodynamic moments. The computational fluid dynamics method based on a shear stress transport turbulence model is established to analyze the flow field characteristics of the forebody. The results indicate that under equal weight coefficients, the presented search algorithm can provide an optimized solution for the front fin to achieve the minimum value of the comprehensive evaluation indicator. When the range of the current wing movement changes or the weight coefficient distribution changes, this search algorithm can still provide the optimal solution and some feasible solutions. Under the given conditions, there is a difference between the optimal solution of the aerodynamic force priority and that of the aerodynamic moment priority. For the case of the aerodynamic moment priority, the mean level of the pressure coefficient corresponding to the optimal solution on the given observation plane is low. The presented method is effective in learning the appropriate installation position of the rocket's front fins. [ABSTRACT FROM AUTHOR]

A self-aligned focusing schlieren (SAFS) system combines the field of view of a conventional schlieren system with the defocus blur of a focusing schlieren system away from the object plane. It can be assembled in a compact form, measuring 1.2 m (4 ft) in length in the described case. The depth of field is sufficiently shallow to distinguish specific spanwise features in a supersonic flow field within a 76.2 mm (3 in) wide test section. As a result, the boundary-layer perturbations on windows and window-material defects and surface imperfections are blurred. Analytical forms are derived for depth of field and vignetting of the SAFS system. A laser spark velocity measurement in Mach 2 flow is performed by tracking the blast wave of a laser spark using 500 kHz SAFS imaging with a 200 ns optical pulse width. The flow Mach number and stagnation temperature are measured by comparing the blast-wave dynamics to an analytical solution. Additionally, schlieren image velocimetry is performed by analyzing natural flow perturbations in 500 kHz SAFS images using a self-correlation method. Comparing the spectra of gas density perturbations from the core flow and a near-wall region reveals a significant difference, with high-frequency prevalence at the boundary-layer location. [ABSTRACT FROM AUTHOR]

In recent times, steam ejectors have garnered significant interest among researchers due to their environmental friendliness and the utilization of low-grade energy sources. However, a key drawback of the ejector refrigeration cycle (ERC) is its relatively low coefficient of performance (COP). Understanding the behavior of ejectors under various operating conditions is crucial for addressing this concern. This study specifically focuses on investigating the flow characteristics of ejectors in the single-choking mode. Both dry steam model (DSM) and wet steam model (WSM) are employed to analyze and evaluate the performance in this study. Based on the findings, it is evident that the discharge pressure (DP) significantly influences the flow characteristics. With increasing DP, there is a decrease in the Mach number and liquid mass fraction (LMF) within the ejector, while the temperature distribution shows an upward trend. Additionally, as the DP increases, there is a notable decline in the entrainment ratio (ER) and production entropy. With an increase in the DP, both the DSM and WSM exhibit similar trends. However, in the DSM, the ER reaches zero at an earlier stage compared to the WSM. Specifically, when the DP rises from 5000 Pa to 5600 Pa, there is a 12.6 % increase in the production entropy in the WSM, while the DSM experiences a slightly higher increase of 12.9 %. [ABSTRACT FROM AUTHOR]

The impact of upstream conditions on magnetopause reconnection has been an intriguing question in solar wind‐magnetosphere coupling. In this study, we conduct a statistical analysis of plasma properties in the reconnection outflow region and the associated upstream solar wind/magnetosheath. We observe that the normalized ion density (N/Nsw) decreases and the flow speed (V/Vsw) increases in the upstream magnetosheath with distance from the subsolar point, consistent with previous models and observations. The magnetic field strength (|B|), ion density (N), and ion bulk speed (|V|) in the upstream magnetosheath exhibit close correlations with those in the reconnection outflow region. This upstream‐downstream correlation likely arises from the process of forming reconnection outflows, where most upstream ions cross the separatrix and mix with ion outflow already accelerated near the X‐line. High‐speed part of reconnection outflow is mostly located on the magnetosphere side of the magnetopause current layer, with outflow velocities peaking close to the upstream magnetosheath Alfvén speed. The spatial extent of high‐speed outflow is greater in conditions of lower solar wind Alfvén Mach number (MA,sw). Additionally, the southward magnetic field in the magnetosheath and |B| of magnetopause current layer are larger in the cases of lower MA,sw. These findings indicate a close connection of plasma properties between the outflow region of magnetopause reconnection and the upstream magnetosheath. Key Points: Plasma and magnetic fields in the upstream magnetosheath/solar wind correlate with those in the outflow region of magnetopause reconnectionIn the upstream magnetosheath and the outflow region, ion density decreases and flow speed increases with distance from the subsolar pointThe spatial extent of reconnection outflow in the magnetopause current layer is larger at lower solar wind Mach numbers [ABSTRACT FROM AUTHOR]

The dispersion relation of the magnetized Kelvin–Helmholtz (KH) instability driven by shear flow with zero thickness of the shear layer is derived theoretically based on a set of magnetohydrodynamic equations in the presence of a transverse magnetic field and a radiation field. The influence of the magnetic field strength, the radiation field strength, and the density ratio of the two sides of the shear layer on KH instability is analyzed by solving the dispersion equation. The results indicate that the presence of radiation and transverse magnetic fields can destabilize the KH instability due to the resulting increase in Mach number, which in turn reduces the compressibility of the system. Also, the extent of the destabilizing effect of the magnetic field can be affected by the magnetoacoustic Mach number M 1 f and the Mach number M 2. The growth rates vary more significantly for relatively small values of both parameters. Finally, the stabilizing effect of a large density ratio is considered, and it is found that as the density ratio increases, the effect of the radiation field is more significant at larger Mach number M 2. These results can be applied to astrophysical phenomena with velocity shear, such as flows across the transition layer between an H ii region and a molecular cloud, accretion flows, and shear flows of cosmic plasmas. [ABSTRACT FROM AUTHOR]

Context. Maps of polarized dust emission of molecular clouds reveal the morphology of the magnetic field associated with star-forming regions. In particular, polarization maps of hub-filament systems show the distortion of magnetic field lines induced by gas flows onto and inside filaments. Aims. We aim to understand the relation between the curvature of magnetic field lines associated with filaments in hub-filament systems and the properties of the underlying gas flows. Methods. We consider steady-state models of gas with finite electrical resistivity flowing across a transverse magnetic field. We derive the relation between the bending of the field lines and the flow parameters represented by the Alfvén Mach number and the magnetic Reynolds number. Results. We find that, on the scale of the filaments, the relevant parameter for a gas of finite electrical resistivity is the magnetic Reynolds number, and we derive the relation between the deflection angle of the field from the initial direction (assumed perpendicular to the filament) and the value of the electrical resistivity, due to either Ohmic dissipation or ambipolar diffusion. Conclusions. Application of this model to specific observations of polarized dust emission in filamentary clouds shows that magnetic Reynolds numbers of a few tens are required to reproduce the data. Despite significant uncertainties in the observations (the flow speed, the geometry and orientation of the filament), and the idealization of the model, the specific cases considered show that ambipolar diffusion can provide the resistivity needed to maintain a steady state flow across magnetic fields of significant strength over realistic time scales. [ABSTRACT FROM AUTHOR]

Core-collapse (CC) supernova remnants (SNRs) are the nebular leftovers of defunct massive stars that died during a supernova explosion, mostly while undergoing the red supergiant phase of their evolution. The morphology and emission properties of those remnants are a function of the distribution of circumstellar material at the moment of the supernova, as well as the intrinsic properties of the explosion and those of the ambient medium. By means of 2.5-dimensional (2.5D) numerical magneto-hydrodynamic (MHD) simulations, we modelled the long-term evolution of SNRs generated by runaway rotating massive stars moving into a magnetised interstellar medium (ISM). Radiative transfer calculations reveal that the projected non-thermal emission of SNRs decreases over time, namely: older remnants are fainter than younger ones. Older (80 kyr) SNRs whose progenitors were moving with a space velocity corresponding to a Mach number of M = 1 (v⋆ = 20 km s−1) in the Galactic plane of the interstellar medium (nISM = 1 cm−3) are brighter in synchrotron than when moving with a Mach number of M = 2 (v⋆ = 40 km s−1). We show that runaway red supergiant progenitors first induce an asymmetric non-thermal 1.4 GHz barrel-like synchrotron SNRs (at the age of about 8 kyr), before further evolving to adopt a Cygnus-loop-like shape (at about 80 kyr). It is conjectured that a significative fraction of SNRs are currently in this bilateral-to-Cygnus loop evolutionary sequence. Therefore, this population should be taken into account with repect to interpreting the data as part of the forthcoming Cherenkov Telescope Array (CTA) observatory. [ABSTRACT FROM AUTHOR]

The intra-cluster medium is prone to turbulent motion that will contribute to the non-thermal heating of the gas, complicating the use of galaxy clusters as cosmological probes. Indirect approaches can estimate the intensity and structure of turbulent motions by studying the associated fluctuations in gas density and X-ray surface brightness. In this work, we aim to constrain the gas density fluctuations occurring in the CHEX-MATE sample to obtain a detailed view of their properties in a large population of clusters. To do so, we use a simulation-based approach to constrain the parameters of the power spectrum of density fluctuations, assuming a Kolmogorov-like spectrum and including the stochastic nature of the fluctuation-related observables in the error budget. Using a machine-learning approach, we learn an approximate likelihood for each cluster. This method requires clusters not to be too disturbed, as fluctuations can originate from dynamic processes such as merging. Accordingly, we removed the less relaxed clusters (centroid shift w > 0.02) from our sample, resulting in a sample of 64 clusters. We defined different subsets of CHEX-MATE to determine properties of density fluctuations as a function of dynamical state, mass, and redshift, and we investigated the correlation with the presence or not of a radio halo. We found a positive correlation between the dynamical state and density fluctuation variance, a non-trivial behaviour with mass, and no specific trend with redshift or the presence of a radio halo. The injection scale is mostly constrained by the core region. The slope in the inertial range is consistent with the Kolmogorov theory. When interpreted as originating from turbulent motion, the density fluctuations in R500 yield an average Mach number of ℳ3D ≃ 0.4 ± 0.2, an associated non-thermal pressure support of Pturb/Ptot ≃ (9 ± 6)%, or a hydrostatic mass bias bturb ≃ 0.09 ± 0.06. These findings align with expectations from existing literature. [ABSTRACT FROM AUTHOR]

We have mapped the magnetic field (B -field) for a region of about 30 pc around the centre of our Galaxy, which encompasses the circumnuclear disc (CND), the minispiral, and the 20 and 50 km s−1 molecular clouds, using thermal dust polarization observations obtained from SOFIA/HAWC+ and JCMT/SCUPOL. We decompose the spectra of 12CO (J = 3 → 2) transition from this region into individual cloud components and find the polarization observed at different wavelengths might be tracing completely different layers of dust along the line of sight. We use modified Davis–Chandrasekhar–Fermi methods to measure the strength of B -field projected in the plane of the sky (⁠|$B_{{}_{\mathrm{POS}}}$|⁠). The mean |$B_{{}_{\mathrm{POS}}}$| of the CND and the minispiral, probed at 53 μ m is of the order of ∼2 mG. |$B_{{}_{\mathrm{POS}}}\!\!\!\lt \!1$| mG close to the Galactic Centre, in the region of the ionized mini-cavity within the CND, and increases outwards. However, the longer wavelength polarization at 216 μ m appears to come from a dust layer that is cooler and behind the CND and has a stronger B -field of about 7 mG. The B -field strength is lowest along the Eastern Arm of the minispiral, which is also the only region with Alfvén Mach number, |$\mathcal {M}_{\mathrm{A}}\gt 1$| and mass-to-flux ratio, λ ≳ 1. Such an observed weak B -field could be a result of the low resolution of the observation, where the tangled B -fields due to the strong turbulence in the high density clumps of the CND are lost within the beam size of the observation. [ABSTRACT FROM AUTHOR]

This paper presents a numerical study conducted to analyze the aerodynamic performance of supersonic missiles consisting of a cylindrical body and four flat-plate rear fins arranged uniformly, equipped with conical and ogive heads. Computational Fluid Dynamics (CFD) simulations were performed using the ANSYS Fluent 17.1 solver, along with the Gambit grid generation software. The objective was to compare the aerodynamic characteristics of these two head designs in terms of drag, lift, and stability at supersonic speeds. Various flow parameters, including Mach number and angle of attack, were investigated to comprehensively assess the performance of the missile configurations. The results indicate clear differences in the aerodynamic behavior of conical and ogive heads. Specifically, there was a 2--11 percent increase in the lift coefficient of the conical heads compared to the ogive heads, and an increase in the drag coefficient of both conical and ogive heads. [ABSTRACT FROM AUTHOR]

The present study explores the characteristics of reacting flow in a scramjet combustor with struts, focusing particularly on implementing different injection strategies. A three-dimensional DLR scramjet combustor is utilised to assess the impact on the system, incorporating multiple injections and varying injection angles on the triangular wedge. The analysis considers three injectors with parallel, upward and downward injections at angles of 15° and 30°. The numerical investigation is conducted under a constant total pressure of 7.82 bar, a temperature of 340 K, and an airspeed of Mach 2 at the inlet. The results highlight the significance of injector location and shape in promoting flame stabilization. Furthermore, injection angles play a crucial role in mitigating shockwave intensity. The numerical analysis involves a steady-state Reynolds-averaged Navier- Stokes equation with the shear stress transport k--ω turbulence model. The obtained results were analyzed by examining the critical variables such as Mach number, static pressure and combustion efficiency across the combustor. Based on the computational results, injecting fuel upward not only increases the overall pressure loss but also enhances the subsonic regime downstream of the strut, which leads to better mixing and combustion efficiencies. This is primarily due to shockwave generation from the edges of the strut and the interactions with the fuel stream shear layers. [ABSTRACT FROM AUTHOR]

This study explores the efficiency of streamline-tracing intake in scramjets within the hypersonic domain, where the Mach number (Ma) exceeds five. Focusing on a rectangular-to-elliptical shape transition intake following Busemann streamlines with a contraction ratio of 4.67, tailored for Ma 7.1, this study innovates beyond traditional design method, employing numerical simulations to interpolate iso- Ma points and shape the leading edge of the intake to enhance the performance. Three distinct leading-edge designs—using a sinusoidal function to closely track all iso- Ma points, a convex second-order polynomial to bolster compression and mass flow, and a concave polynomial to widen the starting envelope—were conceptualized. The research included experiments with 0.06 scaled models for validation and performance assessments of full-scale intakes using three-dimensional viscous simulations. Key performance metrics, such as the pressure ratio, total pressure recovery, mass flow rate, and exit Ma , were evaluated to reveal the capability of all three intake designs to operate effectively over a broad Ma range (3–7). A key finding of this study is the ability to selectively enhance either the compression capability or the total pressure recovery by strategically designing the leading edge of the intake through interpolation methods. This study underscores the importance of using numerical simulation results in the notching process as an effective design strategy to handle a wide range of off-design conditions with fixed geometry and achieve performance improvements in scramjet intakes. • Design of streamline-tracing inlets with various leading-edge shapes. • Similar wave structures observed for various off-design conditions. • A wide range of operating conditions without deteriorating total pressure recovery. • A wide Mach number range of inlet starting. [ABSTRACT FROM AUTHOR]

In order to meet the demanding drag reduction requirements of high-speed aircraft, the drag reduction mechanism of array microslit-opposed jets in the passivation leading edge is studied using a numerical method based on the shear stress transfer (SST) k-ω two-path turbulence model. The computations are performed with the commercial CFD solver FLUENT 13.0 code. The interference characteristics of the array jets are analyzed, the influence laws of the microslit array mode and the angle of attack on drag reduction performance are explored, and a prediction method of passivation leading edge drag characteristics based on radial basis function (RBF) neural network is constructed. The research conditions are as follows: the incoming Mach number is 2.0, the angle of attack ranges from 0° to 30°, and the number of microslits is 0 to 5. The results show that when there is interference between microslit jets, the interaction between adjacent reverse reflux regions will change the shape of the reflux regions. When there is a microslit jet in the center line, the position of the detached shock wave is more affected by the angle of attack. The drag characteristics of the leading edge are determined by the location of the detached shock wave and the characteristics of the reflux regions. The mode of microslits determines the drag reduction efficiency and sensitivity to the angle of attack. There may be contradictions in the design of array microslit-opposed jet schemes regarding the overall drag reduction efficiency, local load, and drag reduction stability. The drag prediction model based on the RBF neural network has an accuracy of 5%, which can provide support for the design of array microslit-opposed jets. Practical Applications: The array microslit-opposed jets provide a new technical approach for reducing drag on the passivation leading edge. The paper studies its flow characteristics and laws and provides relevant conclusions. It can serve as a theoretical reference during design. The number of microslits is not the more the better. When designing microslit patterns, many factors need to be considered. [ABSTRACT FROM AUTHOR]

The main objective of this research work is focused on the acoustic response analysis of functionally graded material (FGM) sandwich cylindrical shell structure immersed in convected fluids. In the analysis, two common types of FGM sandwich cylindrical shells are considered, which include one with FGM face shell and homogeneous ceramic or metal core, and the other with FGM core and homogeneous face shell. Given the structure is immersed in a mean flowing fluid of velocity tangential to the acoustically deformed boundary and it is excited by an incidence oblique plane wave. Fluid–structure coupling is considered by imposing velocity continuity condition at fluid–structure interfaces. By comparing the numerical results with the existing solutions in open literature, the validity of the proposed theoretical model is verified. Finally, based on the developed theoretical mode, the influences of the gradient index, sandwich type, external mean flow, and skin-core-skin ratios on the structural sound transmission performance of FGM sandwich cylindrical shell have been investigated, wherein the first resonance dip shift noticeably to high frequency as the Mach number is increased, and case II has the largest amplitude fluctuation and the lowest STL curve magnitudes, in addition, when M 1 = M 2 > 1.2 , two branches, including positive refraction branch and the negative refraction branch, are found in the contour plots under opposite flow direction conditions, and the large skin-core-skin ratio is beneficial to improve the sound insulation characteristics of structures. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this study is to perform an off-design analysis of the inverted Brayton cycle engine. Design/methodology/approach: The off-design analysis equations of the inverted Brayton cycle engine were first derived in this study and the control parameters of the inverted Brayton cycle engine were first determined and investigated. Findings: It is observed that by controlling the total temperature decrease in cooling section, it is possible to adapt the engine for low specific fuel consumption working conditions or high thrust working conditions. Specific fuel consumption is reduced by 27.1 % by stopping cooling in the cooling section and thrust is increased by 27.6 % by working with full load of the cooling section (500 K temperature decrease in cooling section). It is observed that thrust depending on the flight Mach number increases with an increase in flight Mach number and reaches a peak value at 5.21 flight Mach number and reduces by 80.8 % at 6 flight Mach number relative to the peak value. The specific fuel consumption increases rapidly as the Mach number increases, and the specific fuel consumption is 49.0 g/[kN.s] at Mach 1, reaches 70.4 g/[kN.s] at Mach 5 and increases to 412 g/[kN.s] at Mach 6. The specific fuel consumption increases from 68.1 to 73.0 g/(kN.s) and the thrust decreases from 16.5 to 13.3 kN as the total preburner exit temperature increases from 1,500 to 2,000 K. Specific fuel consumption decreases from 83.1 to 64.8 g/(kN.s) and thrust increases from 4.60 to 11.08 kN depending on afterburner exit total temperature increase from 1,800 to 2,500 K. Research limitations/implications: The cooling section reduces total temperature of the gas flow to lower values to increase the compressor total pressure ratio. The compressor increases the total pressure of the gas flow to the optimum total pressure ratios to increase the nozzle exit Mach number and gain more thrust. The afterburner increases the total temperature of the gas flow to increase the sound speed in the nozzle exit to increase thrust. The nozzle expands the gas flow to reduce the static pressure of the gas flow to near the optimum value, atmosphere pressure, to increase thrust and reduce specific fuel consumption. Practical implications: Hypersonic and supersonic air vehicles can use the current engine model for the its own propulsion systems. Social implications: After first heavier than air flight, aero engines was designed for only used for aero vehicle. Internal combustion engines were used for propelled propeller aircraft at the first term of aircraft. However, propeller-propelled aircrafts are not sufficient to increase aircraft velocity to supersonic Mach numbers due to the shock losses of propeller, so the supersonic era was only introduced by revolution in propulsion systems with new concept. A jet engine was developed to be candidate for supersonic flight. Originality/value: Off-design analysis equations of an inverted Brayton cycle engine were first derived in this study. Furthermore, the control parameters of the inverted Brayton cycle engine were first determined and investigated in this paper. [ABSTRACT FROM AUTHOR]

This study focuses on optimizing the lateral jet efficiency of THAAD-like (Terminal High Altitude Area Defense) missiles operating under hypersonic rarefied flow conditions. We employ the DSMC-QK algorithm to simulate the three-dimensional lateral jet flow field, accounting for thermochemical non-equilibrium effects. The analysis investigates how the force/momentum amplification coefficient varies with the angle of attack, jet pressure ratio, jet Mach number, and jet gas composition. Subsequently, we develop an artificial neural network (ANN) proxy model using the pyrenn toolbox, achieving an average prediction error of 0.866% and a maximum error of 1.60%. Utilizing this ANN model, we perform single- and multi-objective optimizations with a genetic algorithm to determine the optimal jet parameters. The results reveal that in multi-objective optimization, the proportion of helium in the jet gas composition increases, leading to a slight reduction in the force amplification coefficient but a substantial 61.4% decrease in the mass flow rate. This demonstrates that a judicious selection of jet gas composition can significantly reduce mass flow while maintaining high jet efficiency, thus achieving efficient lateral jet control. [ABSTRACT FROM AUTHOR]

The features of dust ion acoustic waves (DIAWs) in a magnetized plasma composed of cold ions, nonthermal nonextensive electrons as well as stationary dust particles are presented. By considering the dust-ion collisions, the damped Zakharov–Kuznetsov (dZK) equation is derived by the standard reductive perturbation technique (RPT). In the presence of external periodic force, the solitary wave solution of dZK equation is given. The effect of system parameters on the phase velocity, Mach number, amplitude and width of DIAWs is discussed. It is found that nonthermal and nonextensivity parameters modify the properties of DIAWs. The phase velocity, Mach number and amplitude increase with increasing nonthermal parameter, but decrease with increasing nonextensivity parameter. [ABSTRACT FROM AUTHOR]

This work presents a coupled numerical methodology, based on a hybrid Navier-Stokes (NS) and Direct Simulation Monte Carlo (DSMC) method, for high-speed compressible gas flows—involving a combination of continuum and rarefied flow regions. Implementation details of coupling cycle are presented along with the validation and performance studies for three cases (Mach number M a 1 = 3.38 , 6.5 and 9) of normal shock waves in argon and three cases (Inlet pressure pin = 1.0 atm, 0.5 atm and 0.1 atm) of flow through a micro-scale convergent-divergent (CD) nozzle. In comparison with our DSMC method-based solutions in the complete domain, the hybrid NS-DSMC method solutions achieve almost similar accuracy in a substantially reduced computational time—44% to 50% for the normal shock problem and 67% to 85% for the micro-scale CD nozzle problem. For the first time in literature, the hybrid NS-DSMC method is presented as a Cartesian-grid method, which is becoming popular nowadays. [ABSTRACT FROM AUTHOR]

Energetic positrons have been observed in the interstellar medium, and high-energy positrons with relativistic energies up to approximately 1 TeV have been detected in Galactic cosmic rays. We conducted a study on the acceleration of particles, specifically positrons, in a nonrelativistic quasi-parallel collisionless shock induced by a plasma consisting of protons, electrons, and positrons. The positron-to-proton number density ratio in the plasma is 0.1. We focused on a representative shock with a sonic Mach number of 17.1 and an Alfvénic Mach number of 16.8 in the rest frame of the shock. To investigate the acceleration mechanisms of particles including positrons in the shock, we utilized 1D particle-in-cell simulations. It was found that all three species of particles in the shock can be accelerated and exhibit power-law spectra. At the shock front, a significant portion of incoming upstream particles are reflected and undergo significant energy increases, and these reflected particles can be efficiently injected into the process of diffusive shock acceleration (DSA). Moveover, the reflected positrons can be further accelerated by an electric field parallel to the magnetic field when they move along the magnetic field upstream of the shock. As a result, positrons can be preferentially accelerated to be injected in the DSA process compared to electrons. [ABSTRACT FROM AUTHOR]

Mitigation of gust-induced separation and aerodynamic loads using a high-frequency blowing/suction slot is demonstrated using high-order implicit large eddy simulation (LES). This approach was previously shown to be effective at alleviating dynamic stall on pitching wings. A NACA0012 wing section operating at a transitional chord-based Reynolds number of Rec=500,000⁠, subsonic freestream Mach number of M∞=0.1 ⁠, and angles of attack of α = 4 deg and 12 deg is subjected to discrete 1-cos transverse gusts. Gust-induced stall is demonstrated and then active flow control (AFC) is applied to cases vulnerable to gust-induced stall. The flow control strategy is shown to be effective at stall suppression during gust encounter, thereby providing partial alleviation of gust induced loads. This approach is most effective at attenuating pitching moment increment. [ABSTRACT FROM AUTHOR]

In this work, to comprehensively analyze the flow field characteristics of a normal slot plasma synthetic jet actuator, three-dimensional simulation models are established for both normal slot and normal orifice actuators. A detailed comparative analysis of the three-dimensional flow field characteristics of these two actuators is performed. The results indicate that the motion shockwaves and jets generated by the normal slot actuator cover a larger and more uniform region, showing planar characteristics and excellent flow control uniformity. The total pressure ratio for the normal slot actuator is 3.59, significantly higher than the value of 3.50 for the normal orifice actuator, indicating lower pressure loss in the former. Additionally, the normal slot has a larger average exit Mach number (Ma), indicating a stronger flow control capability. It also achieves the peak Ma in a shorter time, indicating a faster momentum output response. Therefore, compared with the normal orifice actuator, the normal slot actuator has better potential for flow control. [ABSTRACT FROM AUTHOR]

The study of laminar incompressible fluid flow in the boundary layer revealed, even earlier, that the condition of complete adhesion of fluid particles to the surface (non-slip condition) of the moving body (half-plane) is not met in the flow development (formation) region. The assumption of constancy of the fluid velocity on the surface of a moving body, hence non-slip, leads, in the flow development region, to the complete absence of the normal component of the velocity field. And this contradicts the very concept of the flow development region, where there should be two velocity components - longitudinal (primary) and normal (secondary) ones. In the previous works of the authors, analytical solutions were obtained for the velocity field in the region of development of incompressible fluid flow in the boundary layer. Since the use of the incompressible fluid flow model is restricted by the Mach number, to further expand the speed range, the problem of the of slightly compressible fluid flow development region in the boundary layer was considered. It is analytically proven that all considerations regarding the impossibility of complete non-slip in the flow development region can be applied to a slightly compressible flow. Slight compressibility at the same time means the subsonic nature of the flow and the neglect of temperature effects due to friction. On the basis of a critical analysis of the existing approaches, which consider the flow of a fluid around a immobile plate in the framework of non-gradient flow (which is just impossible due to the lack of a mechanism for creating the motion of the fluid), it is shown that the system of equations is actually non-closed. For the region of flow development, where the longitudinal pressure gradient is not a constant value, one equation is missing. This equation, as in previous works, is obtained from the necessary condition for the extreme of the fluid rate functional. And although the complete solution for the longitudinal component of the velocity contains four constants of integration, to obtain the asymptotics near the solid surface it is sufficient to know only two quantities - the velocity and its first derivative (gradient). These values, as it turns out from the asymptotic solution, coincide with the case of incompressible flow, which allows us to expand the scope of the previously obtained results for a wider domain of Mach numbers, for example . And such values already correspond to the speeds of modern civil aircraft. The dimensionless distribution of pressure in the slightly compressible flow development region is presented and its significant heterogeneity is shown, which, in turn, indicates the importance of the obtained results. [ABSTRACT FROM AUTHOR]

We are concerned with the long time existence of strong solutions to the initial boundary value problem with slip boundary conditions for the compressible Navier-Stokes equations in three-dimensional simply connected bounded domains. We verify that the barotropic Navier-Stokes equations admit a unique strong solution on the time interval where the solution to the incompressible Navier-Stokes equations exists and is reasonably smooth, when the Mach number is sufficiently small. Furthermore, we obtain the uniform convergence of strong solutions for the compressible system toward those for the corresponding incompressible system on that time interval. [ABSTRACT FROM AUTHOR]

Bypass diffusers are used to drain the excess steam generated in the steam generator in case of sudden load reduction or shutdown of the steam turbine. However, the steam at the orifice outlet with the high flow velocity may reverse into the space of last-stage blades and cause forced vibration of the turbine blades. For this study, a full-scale CFD calculation model which couples the last stage and the second-last stage with the bypass diffusers was constructed. The fluid dynamic characteristics of the high-speed steam discharged from the outlet of the bypass diffusers and the effect of steam on the last-stage rotating blades were analyzed comprehensively via both steady and transient numerical methods. The steady results show that the steam at the orifice outlet of the bypass diffusers presents a typical jet flow with some steam flowing back into the last-stage blades region through the exhaust of the cylinders. This results in a notable disturbance to the last-stage rotating blades, characterized by a non-uniform circumferential pressure distribution. The transient numerical simulation results reveal that the outlet mass flow rate of the steam from the second-last-stage stationary blades has a significant effect on both the Mach number distribution and the surface forces acting on the last-stage rotating blades. The higher outlet mass flow rate of the steam escalates the instability of the flow field. The difference in the Mach number at the same position can reach as high as 60% or more under different operating conditions. The forces acting on the last-stage rotating blades in different directions change periodically with time, and the magnitude and period of the variation correlate with the outlet mass flow rate of the steam. [ABSTRACT FROM AUTHOR]

The integrated design of waverider forebodies and inlets is considered a critical challenge in high Mach number vehicle development. To facilitate the rapid construction of integrated geometrical models for waverider forebodies and inlets during the conceptual design phase, a method based on discrete point cloud reconstruction has been proposed. In this method, the geometries of the waverider body and inlet are used as inputs and decomposed into the point cloud under discrete rules. This point cloud is refitted to generate new section lines, which are then lofted into an integrated shape under the constraints of guide curves. By modifying the coordinates of the point cloud positions, the geometric configuration of the integrated shape can be rapidly adjusted, providing initial support for subsequent aerodynamic optimization and thermal protection. Using this method, an integrated approach was applied to a waverider forebody and inward-turning inlet in a tandem configuration. This achieved body-inlet matching and integration, resulting in a 15.6% improvement in the inlet's total pressure recovery coefficient. The integration time was reduced to just 3.18% of the time required for traditional manual adjustments. Additionally, optimization based on the discrete point cloud enhanced the lift-to-drag ratio by 7.83%, demonstrating the feasibility of the proposed method. [ABSTRACT FROM AUTHOR]

Many UAVs today have an S-bend inlet for the sake of stealth; however, the majority of them have a relatively gentle transition of the flow channel. This study presents an experimental investigation of the aerodynamic performance and swirl flow characteristics of a UAV inlet with double 90° bends, which is also integrated with an aircraft fuselage as well as a volute. The influences of angle of attack, sideslip angle, AIP Mach number and freestream speed are explored in detail. The influences of the deflectors installed ahead of the first 90° bend of the inlet and the baffle installed at the bottom of the volute are revealed. It is found that both the deflectors and the baffle are beneficial in enhancing the aerodynamic performance of the inlet and alleviating the intensity of the swirl flow inside the volute. [ABSTRACT FROM AUTHOR]

This paper outlines an experimental investigation conducted at the INCAS trisonic wind tunnel, focusing on the determination of pitch damping coefficient. The model used for this investigation is the Basic Finner Model, a standard model for dynamic tests which consists in a cone-cylinder body with four rectangular fins. The study aims to evaluate the influence of various parameters—including the Mach number, angle of attack, reduced frequency, center of rotation, and roll angle—on pitch damping coefficient. The employed method for determining these coefficients is the free oscillation method which consists in measuring the model oscillation in free stream after an initial perturbation. In order to perform these dynamic tests in the wind tunnel, a dedicated rig was developed to initiate the model's oscillation using a linear servo-actuator and to record its oscillation using a strain gauge. The results obtained from the experiments illustrate how each parameter impacts the pitch damping coefficient, highlighting the precision of the measurements. The paper's conclusion presents that the developed rig and the method used provide accurate results, and the variation in different parameters can change the damping coefficient. [ABSTRACT FROM AUTHOR]

A micro-pulse detonation engine (μPDE) was designed and installed to a direct-connect scramjet combustor of Pusan National University (PNU-DCSC). The active excitation on the scramjet combustor was experimentally studied using the μPDE operating at frequencies of 10–20 Hz. A vitiation air heater (VAH) was used to supply high-enthalpy vitiated air to the isolator and the scramjet combustor at a Mach number of 2.0, with a total temperature of 1600 K and a total pressure of 1.68 MPa. The exit of μPDE was located at the center of the cavity of the scramjet combustor. Active excitation was performed at equivalence ratios of 0.111 and 0.163, and characteristics were analyzed through Schlieren recording and bottom wall pressure measurement. As a result, when the detonation emitted from the μPDE entered the scramjet combustor, it instantly formed a shock train and moved forward within the scramjet combustor. The flame instantaneously changed from the cavity shear layer flame to the cavity flame. Through bottom wall pressure measurement, it was also observed that active excitation resulted in a significant pressure increase near the cavity compared to when active excitation was not performed. This revealed combustion characteristics, indicating improved combustion efficiency from the pressure increase in the scramjet combustor. [ABSTRACT FROM AUTHOR]

Conventional forward bleed slots reduce the hypersonic inlet starting Mach number but suffer from excessive flow leakage after restart. This paper proposes a novel reverse bleed slot design method for curved axisymmetric inlets of a solid-fuel scramjet. Leveraging the Kantrowitz criterion and detailed flow analysis, the method optimizes bleed slot placement, number, area, and angle. Results show superior aerodynamic performance by placing slots in the non-starting region of the internal compression section, considering both unstarted flow and separation bubble dynamics during restart. Each bleed slot area is calculated successively down-stream based on the Kantrowitz criterion. Finally, the effects of bleed slot angle have been extensively studied. The key inlet performance reaches its optimum at a slot angle of approximately 130°, achieving a significant reduction in the starting Mach number (from 4.80 to 3.65) and a 50% decrease in bleed flow rate compared to the forward slot design. This method demonstrates its feasibility and effectiveness, enabling substantial improvement in inlet starting performance with minimal flow loss. [ABSTRACT FROM AUTHOR]

Being a direct computational aeroacoustics method, Lattice Boltzmann method (LBM) has great potential and broad application perspective in the field of numerical simulation of aerodynamic noise due to its low dispersion and low dissipation. A series of numerical algorithms and the related improvements based on the standard LBM method are proposed and developed in this paper to adapt to the airfoil noise calculation with complex grid at middle‐high Reynolds number. First, a new mathematical‐boundary‐recognition algorithm based on Green's formula is proposed to deal with complex curved geometric models, which is validated by three‐element airfoil 30P30N benchmark. Then, in order to reduce grid redundancy and improve computing efficiency, the grid refinement technique of domain decomposition model (DDM) is adopted and also improved, which is verified by calculating the flow and sound fields around 2D and 3D cylinders at Reynolds number equal to 90,000. Finally, three different LES turbulence models are combined with the standard MRT‐LBM method, where different finite difference schemes are used to solve Reynolds stress tensor which is different from the traditional one. Through the direct acoustic numerical simulation of NACA0012 airfoil at Reynolds number equal to 200,000, the effects of Smagorinsky models and Wall‐adapting local eddy‐viscosity (WALE) model on aerodynamic noise prediction are compared and analyzed. Overall, the proposed methodology is shown to be appropriate for predicting the aerodynamic noise at low Mach number and can successfully simulate the generation and propagation of far field acoustics. [ABSTRACT FROM AUTHOR]

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Localized dynamic pressure enhancements arising from kinetic processes are frequently observed downstream of the Earth's bow shock. These structures, called jets, modify their plasma surroundings and participate in particle energization. Here, we report the first observations of jet-like structures in a non-planetary shock environment: downstream of interplanetary shocks. We introduce an analysis approach suitable for such conditions and apply it to Wind spacecraft data. We present one event with a Mach number similar to the Earth's bow shock as a benchmark, as well as two low Mach number, low beta shocks: a parameter range that is difficult to access at planets. The jet-like structures we find are tens of ion inertial lengths in size, and some are observed further away from the shock than in a limited magnetosheath. We find that their properties are similar to those of magnetosheath jets: in the frame of the shock these structures are fast, cold, and most have no strong magnetic field variations. All three interplanetary shocks feature foreshock activity, but no strongly compressive waves. We discuss the implications, these findings have for the proposed jet formation mechanisms. [ABSTRACT FROM AUTHOR]