International Journal of Modern Physics B

This paper presents a comprehensive aircraft icing simulation tool implemented in an in-house Navier–Stokes parallel multi-block solver. In detail, the droplet flow field is solved by Eulerian approach, and a Partial Differential Equation (PDE)-based ice accretion model is adopted to determine the runback water flow and icing rate. Numerical validations are performed on the two-dimensional (2D) NACA 0012 airfoil, where good agreements with the literature are observed. Additionally, the paper investigates the influence of droplet size on the final ice shape. Results show that droplets with greater Median Volume Diameter (MVD) are more likely to impact on the wall, which results in larger droplet impingement limit and icing limit.

When air passes through the hangar of a frigate, the unstable airwake appears in the rear of the hangar, which may significantly increase the workload of the ship-borne helicopter pilot. Therefore, there must be a profound understanding of the characteristics of the airwake. In this paper, the airwake was numerically studied by using the Improved Delayed Detached-Eddy Simulation (IDDES) turbulence model on structured grids. The flow fields of the different simplified frigate afterbody models, consisting of hangar and flight deck, were compared regarding the size of the recirculation zone. A parametric study was conducted by varying the hangar length to find the optimal afterbody model with minimal recirculation zone behind the hangar. The results show that the size as well as the location of the recirculation zone are significantly affected by the hangar length, and the optimal afterbody model has been obtained.

The recently proposed boundary condition-enforced immersed boundary-gas kinetic flux solver (IB-GKFS) is a new approach for simulation of compressible flows with curved and moving boundaries. In the previous application of IB-GKFS, only the Dirichlet boundary condition is considered, which cannot be applied directly to the Neumann boundary condition. In this paper, an auxiliary layer of Lagrangian points is introduced to tackle Neumann boundary condition. Two test cases, including flow around a circular cylinder and flow around a NACA0012 airfoil, are carried out for validation. The results obtained by the present scheme are compared with the reference data available in the literature. Good agreements are achieved, which indicate that the developed method can effectively simulate the compressible flow with Neumann boundary condition.

The nonlinear coupled constitutive relations (NCCR) has been studied intensively under structured finite-volume framework. For further engineering applications, it is necessary to study the potential capability of NCCR with the use of unstructured mesh. In this paper, a three-dimensional NCCR solver with hybrid unstructured techniques is developed on the basis of our in-house code. An upwind flux-splitting scheme with LU-SGS implicit time-marching scheme is employed. The accuracy of the developed NCCR solver is validated by two hypersonic cases, such as high-speed flows around 2D cylinder and 3D blunted cone tip in rarefied regimes. Computations show that the solver using hybrid unstructured techniques yields comparable good results as the structured one. Moreover, they can get solutions in better agreement with the Direct Simulation Monte Carlo (DSMC) than the NS results.

In this paper, a hybrid lattice Boltzmann flux solver (LBFS) is proposed for simulation of 3D integrated hypersonic fluid-thermal-structural problems. In the solver, the macroscopic 3D Navier–Stokes equations and structural heat transfer equation are discretized by the finite volume method and the numerical fluxes at the cell interface are reconstructed by the local solution of Boltzmann equation. To compute the numerical fluxes, two lattice velocity models are introduced. One is the D1Q4 discrete velocity model for calculating the inviscid flux across the cell interface of N–S equations, and the other one is the D3Q6 model for evaluating the flux of structural energy equation. Furthermore, a new dual-thermal-resistance model is proposed to calculate the thermal properties on the fluid-structure interface. To validate the accuracy and stability of the present solver, applications for hypersonic fluid-thermal-structural analysis are demonstrated on aerodynamically heated blunt cone body at $\mathrm{\text{Ma}}=10.6$. Numerical results showed that the present solver can predict accurately the thermal properties of hypersonic fluid-thermal-structural problems and offer the potential for significant improvements in predicting fluid-structural-thermal problems of long-endurance high speed vehicles.

Morphing aircraft has adjustable aerodynamic shapes and is suitable for variable flight conditions. And there have been growing interests in recent years. However, the forces and moments are highly nonlinear, bringing challenges in design and modeling of flight control system (FCS). In this paper, unsteady Computational Fluid Dynamics (CFD) simulations are performed by solving the Arbitrary Lagrangian–Eulerian (ALE) governing equations on unstructured dynamic mesh, then unsteady aerodynamic characteristics of a variable-sweep morphing aircraft at hypersonic speed are acquired. The nonlinearity index theory is employed to analyze the nonlinearity of pitching moments. FCS is designed using nonlinear dynamic inversion control, taking attitude angles and sweep angle into consideration. Unsteady flow simulation is performed using unsteady CFD coupled with FCS. Cases of open and close loop morphing flight over the variation of sweep angle corresponding to flight conditions are studied. Results show that the nonlinearity of unsteady forces and moments are significant. The combination of unsteady CFD and FCS provides a powerful approach to the study of morphing aircraft.

Hypersonic inlet is an important part of the propulsion system of hypersonic air-breathing vehicles. However, the performance of the two-dimensional hypersonic inlet, a major type of hypersonic inlets, is considerably deteriorated for lateral spillage. In this study, waverider-configuration chines mounted on the lateral sides of a two-dimensional three-staged external-compression hypersonic inlet for a Mach number of 6.0 are investigated to determine their ability to alleviate the lateral spillage. The chines are built by using a waverider design method. The numerical results suggest that a severe flow spillage induced by three-dimensional effect shows up near the lateral edge of the inlet without chines, which degrades the mass-flow ratio and flow uniformity. In contrast, the waverider-configuration chines effectively alleviate the lateral spillage. Consequently, the mass-flow ratio and the flow uniformity are both improved significantly.

Flow past a prolate spheroid, which is a representative simplified configuration for vehicles such as maneuvering ships, submarines and missiles, comprises a series of complex flow phenomena including pressure-induced flow separation, which results in unsteady forces and movements that may be detrimental to vehicles’ performance. In this paper, a Delayed Detached Eddy Simulation (DDES) method combined with a new high-order U-MUSCL scheme is proposed to more precisely and accurately capture the flow separation and vortex structure. This method is applied to simulate the aerodynamic performance of the 6:1 prolate spheroid at an AOA of $20^{\circ }$ with the Reynolds number of $4.2\times 10^6$. Axial pressure distribution of five individual chord wise sections and flow field structure of the aft body are analyzed. Numerical results agree well with the experimental data. It can be concluded that DDES combined with three-order U-MUSCL scheme demonstrates reliable performance since it captures the vortex structure of aft body distinctly and predicts the separation and reattachment points of the secondary vortex precisely.

This paper is focused on the gas properties over a cylinder from continuum to rarefied regimes based on the non-Newtonian constitutive model. This new constitutive model is first derived from Eu’s nonequilibrium ensemble method, which is intended for accurate description of nonequilibrium flows. Some assumptions and simplifications are made during the establishing progress of the new constitutive model by both Eu and Myong. To verify its accuracy, temperature contours and skin frictions around the cylinder are simulated by this new model. The inflow Mach number is equal to 10 and the Knudsen number ranges from 0.002 to 0.05. All simulation results are compared with Navier–Stokes (NS) and the direct simulation Monte Carlo (DSMC) methods in detail. The comparisons of friction around the surface show that the non-Newtonian constitutive models are better than the linear constitutive relations of NS equations for the prediction of nonequilibrium flow and much more close to DSMC simulation results.

A framework of numerical formulations for the aeroelastic analysis of helicopter rotor is presented in this paper. The blade structural dynamics are modeled by an open source multibody dynamic software MBDYN, which solves finite element equation of elastic bodies in general motions. Then the structural deformation is transformed to blade surface grid by radial base function (RBF) interpolation, and volume grids are regenerated by RBF and TFI methods. Lastly, the fluid governing equations are solved. By integrating the above methods, S76 hovering rotors are simulated and compared to the test data. Results show that elastic torsion decreases local angle of attack. For status at $M_{\mathrm{\text{tip}}}=0.65$ and $𝜃>10^0$, the shock and shock-induced separation are reduced on the outboard blade, which has remarkable effects on the prediction of rotor hovering performance.

The prediction of heat transfer for blunt bodies in hypersonic flows remains a great challenge. In particular, the uncertainties are larger in the leeside due to the complexity of the wake flow. Generally, the heat transfer is over-predicted using the Reynolds-averaged Navier–Stokes (RANS) models. In this paper, the improved delayed detached eddy simulation (IDDES) method is used to simulate the Mach 6 flow around a scaled spherical capsule model. In addition, a low dissipative WENO scheme is used for inviscid fluxes and dual-time stepping method is applied for time advancement. Results are compared to experimental data for mean and instantaneous heat transfer in the leeside of the aftbody. It is shown that the integrated error is 75.49% for RANS while 35.69% for IDDES method. Moreover, the multi-scale structures in the separation region are also resolved well by the IDDES method.

In this paper, a generalized sphere function-based gas kinetic scheme (GKS) is developed for simulation of two-dimensional compressible viscous flows. This work aims to improve the existing simplified GKS, in which a simple two-dimensional/three-dimensional distribution function is used as the equilibrium state. The present scheme applies the finite volume method to discretize Navier–Stokes equations and the numerical flux at cell interface is evaluated by the local reconstruction of solution for the continuous Boltzmann equation with $N$-dimensional sphere function. This local solution contains both the equilibrium part and nonequilibrium part, and the contribution of the nonequilibrium part is controlled by a switch function. It is found that the present scheme has the same expression as the circular function-based GKS for the two-dimensional case and the sphere function-based GKS for the three-dimensional case. Therefore, these two GKSs can be unified, and the developed method provides another way for the simple distribution function.

Mesh adaptation is a reliable and effective method to improve the precision of flow simulation with computational fluid dynamics. Mesh refinement is a common technique to simulate steady flows. In order to dynamically optimize the mesh for transient flows, mesh coarsening is also required to be involved in an iterative procedure. In this paper, we propose a robust mesh adaptation method, both refinement and coarsening included. A data structure of $k$-way tree is adopted to save and access the parent–children relationship of mesh elements. Local element subdivision is employed to refine mesh, and element mergence is devised to coarsen mesh. The unrefined elements adjacent to a refined element are converted to polyhedrons to eliminate suspending points, which can also prevent refinement diffusing from one refined element to its neighbors. Based on an adaptation detector for vortices recognizing, the mesh adaptation was integrated to simulate the unsteady flow around a tri-wedges. The numerical results show that the mesh zones where vortices located are refined in real time and the vortices are resolved better with mesh adaptation.

The lateral jet interaction on a slender body in supersonic flow was investigated by numerical simulation. The spatial and surface flow characteristics induced by jet interaction were shown. As a result, when the lateral jet is not in the longitudinal symmetry plane, the jet interaction causes asymmetric separation flow of surface and space, and destroys the pressure distributions of the slender body. With different angle of attack and circumferential positions of jet, the flow characteristic of the after body for jet in asymmetry plane changes greatly. The results with and without jet interaction also show that the far-field interaction played a major role in the lateral jet interaction.

Heat transfer is an important phenomenon that exists in many industrial applications, especially for gas turbines, aeronautical engines. In this work, two different turbulence models ($k-\omega$ and SAS model) are used to investigate the effects of inlet turbulence on wall heat transfer and the characteristics of flow field in a well-known turbine guide vane (LS89). In order to handle the transition, Menter’s $\gamma -{Re}_{𝜃t}$ transition model is used. The simulations show that the inlet turbulence has an apparent effect on the wall heat transfer of the vane. Not only the maximum wall heat transfer coefficient is increased, the distribution of wall heat flux at the suction side is also modified. The isentropic Mach number along the vane surface is insensitive to the variance of inlet turbulence intensity. Besides, a shock appears in the throat and a laminar-to-turbulence transition position moves forward after the main flow turbulence is enhanced. Moreover, the results indicate that SAS model is capable of capturing more flow structures such as reflecting pressure waves and shedding vortexes while the $k-\omega$ model misses them due to the dissipation.

To improve the predictive ability of computational fluid dynamics (CFD) on the transonic buffet phenomenon, NASA SC(2)-0714 supercritical airfoil is numerically investigated by noninstructive probabilistic collocation method for uncertainty quantification. Distributions of uncertain parameters are established according to the NASA wind tunnel report. The effects of the uncertainties on lift, drag, mean pressure and root-mean square pressure are discussed. To represent the stochastic solution, the mean and standard deviation of variation of flow quantities such as lift and drag coefficients are computed. Furthermore, mean pressure distribution and root-mean square pressure distribution from the upper surface are displayed with uncertainty bounds containing 95% of all possible values. It is shown that the most sensitive part of flow to uncertain parameters is near the shock wave motion region. Comparing uncertainty bounds with experimental data, numerical results are reliable to predict the reduced frequency and mean pressure distribution. However, for root-mean square pressure distribution, numerical results are higher than the experimental data in the trailing edge region.

An electronic state-to-state approach is developed to reproduce numerically the radiative processes in the experiments on the Electric Arc Shock Tube facility at NASA Ames Research Center. The experiments measured the spectral radiance of C₂ Swan band and CO 4th Positive band behind the strong shock wave in a Martian-like mixture. The present state-to-state approach solves the electronic states of strong radiators by integrating the collisional-radiative model into the master equations in the frame of Euler equations. Particularly, the electron impact dissociation of CO is included, and the rate coefficients are proposed for the electronic state-specific heavy-particle impact excitation and dissociation of C₂. The nonequilibrium radiation behind the shock is calculated by the line-by-line method, and it is then convoluted using the calibrated smearing function in order to compare with the experiments. The state-specific simulation results is found to agree well with both spectral and spatial measurement data. The simulated electronic state populations of CO and C₂ deviate from Boltzmann distributions significantly, which could explain the failure of previous two-temperature quasi-steady-state based simulations.

The hydrodynamic force (HF) evaluation plays a critical role in the numerical simulation of fluid–structure interaction (FSI). By directly using the distribution functions of lattice Boltzmann equation (LBE) to evaluate the HF, the momentum exchange algorithm (MEA) has excellent features. Particularly, it is independent of boundary geometry and avoids integration on the complex boundary. In this work, the HF of lattice Boltzmann simulation (LBS) is evaluated by using the MEA. We conduct a comparative study to evaluate two lattice Boltzmann models for constructing the flow solvers, including the LBE with single-relaxation-time (SRT) and multiple-relaxation-time (MRT) collision operators. The second-order boundary condition schemes are used to address the curve boundary. The test case of flow past a cylinder asymmetrically placed in a channel is simulated. Comparing the numerical solutions of Lattice Boltzmann method (LBM) with those of Navier–Stokes equations in the literature, the influence of collision relaxation model, boundary conditions and lattice resolution is investigated. The results demonstrate that the MRT-LB improves the numerical stability of the LBM and the accuracy of HF.

A new integral computational formulation is presented to evaluate the noncompact noise induced by low Mach number flows. Based on Howe’s Vortex Sound Theory and wave equation of Green’s function in free space, we obtain a new integral equation by choosing a permeable boundary as the integral boundary. The flow calculation is developed with second-order CFD solvers, and noise calculation is executed with a two-step method to solve scattered sources and far field pressure. Two- and three-dimensional circular cylinders are chosen as test examples. The pressure amplitude obtained with permeable boundary agrees well with that obtained with body-fitted high-order method. Numerical results indicate that the present method is valid and efficient to calculate noncompact noise, and the permeable boundary can replace the body surface as the integral boundary.

Green urbanism has stimulated more research on the aerodynamics of tree in recent years. The insight gained in studying wind load on trees would mitigate risk of tree falling and enable sustainable landscape planning. However, deciphering the effect of wind on trees is a daunting task because trees come in various species, shapes and sizes. In this study, we aim at conducting wind tunnel tests on various species of trees, including measuring the respective drag coefficient and turbulent flow field using a force balance and particle image velocimetry system. The wind tunnel experiment is conducted using scaled down fractal tree model at 10 and 15 m/s. The 3D-printed tree model is grown based on the data collected on the species-specific tree parameters, such as the height, trunk diameters, crown box dimensions, etc. In this paper, the wind tunnel result of Yellow Flame (Peltophorum pterocarpum) is presented. Results show that the drag coefficient for this inflexible tree model is not sensitive to wind speed. The Reynolds shear stress and turbulence kinetic energy are observed to be the largest at the top and bottom of the crown where the velocity gradients are the highest.

To investigate the static pressure distribution characteristics of a flying-wing model, an advanced binary pressure sensitive paint (PSP) technique is introduced. It has low-temperature sensitivity and can compensate the errors induced by temperature. The pressure measurement test was performed in 0.6 m trisonic wind tunnel at angles of attack ranging from 0${}^{\circ }$ to 12${}^{\circ }$ in supersonic condition, adopting a low-aspect-ratio flying wing model. The binary PSP is sprayed on the upper surface of the model while pressure taps are installed on the upper surface of the right wing. Luminescent images of two probes are acquired with a color charge-coupled-device camera system and processed with calibration results. During the test, the surface pressure is measured by PSP and transducer, respectively. The results obtained show that the binary paint is of advantage to the surface pressure measurement and flow characteristic analysis. The high-resolution pressure spectra at different angle of attack clearly reveal the impact of leading edge vortex on the upper surface pressure distributions. The pressure measured by PSP also agrees well with the pressure tap results. The root mean square error of pressure coefficient is 0.01 at $\mathrm{\text{Ma}}=1.5$, $\alpha =0^{\circ }$.

This paper describes an experimental investigation of the initial growth of flow asymmetries over a slender body of revolution at high angles of attack with natural and disturbed noses. Time-resolved particle image velocimetry (PIV) is used to investigate the flow field around the body. Flow visualization clearly shows the formation of the asymmetric vortices. Instantaneous PIV shows that the amplified asymmetric disturbances lead to Kelvin–Helmholtz instability appearing first on one side, which increases the momentum exchange crossing the layer. As a result, the separation region shrinks which creates the initial vortex asymmetry.

Cavity noise caused by the high-speed airflow has been paid much attention in the field of aerospace, and the study of cavity noise suppression has an important significance on improving the safety of aircraft. Effects of the leading-edge plate on the flow and the noise of the cavity model (the ratio of length to depth is 6) at a Mach number (Ma) of 0.6 are investigated with high-speed wind tunnel experiment. The acoustic and the flow field information in the cavity are obtained with the dynamic/static pressure measurement and oil flow experiment. The result shows that the leading-edge plate can greatly lift the shear layer, raise the impact position on the back-wall area, and reduce the flux rate and intensity in the cavity. With the controlling of leading-edge plate, the static pressure and backflow area are effectively suppressed and the SPL and peak noise also fall down significantly. The leading-edge plate provides a new method for cavity noise suppression in subsonic flow condition, which can be effectively applied to flow/noise controlling of cavity structure on aircraft.

In order to reveal the effect of aircraft surface concave roughness on the transition, several hypersonic quiet wind tunnel tests were carried out on the flat-plate model with independent concave roughness on the surface. The experimental Mach number was 6.5 and the total temperature was 430 K. Using infrared thermal imaging technology, a general rule of the effect of concave roughness on the boundary layer transition was obtained. The flow visualization results and the PCB pressure measurement results showed that the concave roughness promotes the transition by increasing the growth of the first mode wave rather than the growth of the second mode wave.

An experimental study is reported of supersonic jet surface flow structure visualization and wall shear stress field measurement issuing from a rectangular nozzle with aft-deck. The near-field surface flow structures evolution from over-expansion to under-expansion with the increase of nozzle pressure ratio (NPR) are successfully captured by surface oil flow visualization and shear sensitive liquid crystal coating (SSLCC) technique. The quantitative measurement result of shear stress vector field obtained by SSLCC shows that shear stress directions change significantly across the shock wave and expansion fans, while the magnitudes of shear stress have no obvious changes. Surface streamlines calculated by SSLCC image keep great consistency with the streamlines visualized using oil flow technique, which demonstrates the accuracy and potential application of SSLCC in supersonic jet surface flow visualization.

The influence of wavy wall on the hypersonic boundary layer instability and the related aerodynamic heating are investigated over a flared cone. Experiments are conducted in a Mach 6 wind tunnel using Rayleigh-scattering flow visualization, fast-response pressure sensors and infrared thermography. Three sizes of wavy wall have been placed on the model in comparison with the smooth wall cone. The results show that compared to the smooth-wall cone, all the wavy-wall cones can suppress the second-mode instability to a certain degree and eliminate the local heating spot before transition is completed. Also, it is found that with the decrease of wavy wall wavelength, the transition location is shifted downstream.

When aircraft pass through the clouds that contain super cooled water droplets during aviation, the droplets collide with the wing surface and ice is formed, which induces a significant threat to aviation safety. Studies on droplet movement in gaseous medium are prerequisite for deicing/anti-icing researches. So, in this work, an experimental study is performed on water droplet deformation as the droplets approach the leading edge of an airfoil. This experiment is carried out in a vertical wind tunnel, with a NACA0012 airfoil model assembled 4.3 m downstream of the droplet generator. The influence of Weber number (ranging from 0.2 to 36) on the deformation of a 2 mm diameter droplet is thoroughly investigated. The results indicate that droplets maintain initial form with Weber number under 10; after that droplet deforms into remarkable bag deformation as Weber number reaches to 19, and bag-stamen deformation mode as Weber number is above 20. This observed correlation between Weber number and deformation mode is validated through comparing with published simulation results. Furthermore, using the high-speed camera, clear images of the droplet structure during the deformation process are taken and are shown in detail in this work.

The amplification of Tollmien–Schlichting (T-S) wave plays an important role in the process of boundary-layer transition. This paper investigates the measurement of T-S wave using hot-wire anemometer (HWA) in a wind tunnel. To precisely acquire T-S wave, the vibration of hot-wire probe and the influence of electromagnetic interference (EMI) are considered. By introducing different amplitudes and frequencies of vibration ribbon, the development of T-S waves is obtained. Lift-up of low-speed fluid and downward of high-speed fluid are observed during the transition.

An aerofoil commonly used in aerospace engineering to produce lift is also employed in the motor sport industry to produce downforce for improving traction during cornering. This paper investigates aerofoil surface modification through ‘golf ball dimpling’, used to reduce flow separation behind a golf ball. The studies of other researchers have shown that this type of design can have a positive effect on improving aerofoil performance. However, no optimization information of dimple sizing is given in literature. Therefore, three types of dimpling sized at 5, 10 and 15 mm are applied to the surface of a NACA 6615 wing at 25% chord length from the leading edge in this study using Computational Fluid Dynamics (CFD) as an initial design process. Then a physical model, made through 3D printing additive manufacturing (AM), is tested at angles of attack (AoA) ranging from $0^{\circ }$ to $20^{\circ }$ and wind speed up to 30 m/s in a subsonic wind tunnel. Experimental and CFD results show that the smallest dimple size provides the most significant increase on lift to drag ratio at high AoA above $10^{\circ }$. This ratio increases further with the wind speed, indicating that a high AoA wing favors down force to improve drag reduction performance.

The mixing within a full depth lock-exchange gravity current was investigated experimentally with planar laser-induced fluorescence PLIF. The scalar edge of the gravity current was identified via an edge detection algorithm. A strong mixing region located along this edge was extracted from the flow field and analyzed. A dimensionless background potential energy was defined to characterize the local mixing rate in this region, which showed a two-stage behavior. In the section near the nose, mixing rate oscillates strongly under the influence of the lobe and cleft shifting, while in the upper part, the mixing rate grows gradually under the mechanism of billows rolling up.

The cavity flow oscillation is a typical unsteady, nonlinear flow phenomenon, accompanied by resonance with strong pressure oscillation and intense turbulence. A control method is proposed to suppress the cavity oscillation using high-speed mass injection. Wind-tunnel experiments are performed to validate the control method. A length-to-depth six cavity is employed in the experiments with a Mach number 1.8 of the incoming flow. The upstream injection is generated in the leading edge of the cavity with a large momentum coefficient (more than 10%). Effects of the high-speed mass injection on the cavity oscillation are validated by a significant reduction of the overall sound pressure level inside the cavity. The broadband cavity noise is reduced by more than 10 dB, with a great suppression of the cavity tones. Owing to the high-speed mass injection, the cavity oscillation has been effectively suppressed in the supersonic condition. Because there is no need of additional gas supply in the control system, the control strategy is considered to be a potential and practical candidate for engineering applications.

A two-stage multi-nozzle ejector experimental system is established, and the pressure distribution on the wall and in the flow field of the test model is measured when the total pressure of the two-stage primary flow changes. Results show that when the total pressure of primary flow increases, vacuum chamber static pressure decreases to its lowest limit of 1 kPa, and then begins to rise. Flow choking mechanism of two stage ejector is clarified. If the total pressure of the first stage primary flow is low, the flow in the second stage ejector first reaches the choking state when the total pressure of the second stage primary flow increases, which causes the rise of first stage back pressure, and then the first stage ejector reaches the choking state. If the total pressure of the first stage primary flow is high, the increase of total pressure of the second stage primary flow reduces the back pressure of the first stage ejector and therefore the first stage ejector first reaches the choking state, and then the second stage reaches the choking state.

Solid-state synthesis based on supersonic gas flow impingement at room-temperature is an alternative approach to traditional mechanochemical preparation. The supersonic airflow is generated by a convergent-diffusion nozzle with a design Mach number of 3.0. The solid material particles from the suction pipe in the coaxial nozzle can get very high kinetic energy in microseconds. Then the particles impact the target or collide with each other to achieve the transfer of kinetic energy to thermal or chemical energy. We utilize the infrared technology to successfully measure the solid particles’ temperature while they impact the target after gathered energy from the supersonic air gas. The results show that the average temperature of the silicon particles with an average particle diameter of 150 $\mu$m after impacting the target is about $60^{\circ }$C, and some of the temperature exceed $200^{\circ }$C. It dedicates that the kinetic energy of the particles during the collision translate into internal energy indeed. The work of this paper lays a good foundation for further research on the low-temperature solid-phase reaction processes.

Usually, force tests and pressure tests are conducted with different model in wind tunnel which could cause trouble for aircraft design as the pressure test results always could not match the force test results. So attempts to do force tests along with pressure tests are conducted. Results show that it is possible to do force test based on pressure-test-model when some improvements are done with the model and sting. Both force and pressure test results can be measured at the same time. The amount of pressure test results need not be reduced, and yet precision and accuracy of the force tests are acceptable for aircraft design, and this technology can be popularized with low cost.

To evaluate the unsteady nature of wall pressure in an over-expanded single expansion ramp nozzle, under fixed nozzle pressure ratio (NPR) of 5.92, 6.55 and 7.19, an experimental investigation has been conducted based on focusing Schlieren techniques and dynamic pressure measurement. For all cases, the results show fully formed restricted shock separation (RSS) on the upper wall, which experiences flow reattachment on the wall downstream of separation resulting in the formation of a separation bubble. The separation mode is also RSS on the lower wall at $\mathrm{\text{NPR}}=5.92$. However, the lower wall pressure is randomly larger or lower than ambient pressure near the nozzle exit at $\mathrm{\text{NPR}}=6.55$, the separated shear-layer can intermittently impinge on the lower wall, and the separation mode is partially restricted shock separation (pRSS). At $\mathrm{\text{NPR}}=7.19$, the separation flow does not reattach downstream of the lower wall. That is, it occurs free shock separation (FSS).

Synthetic Jet Actuator (SJA) works cyclically with the directionally transportation of fluid near the exit, and the paper presents the performance of a loudspeaker driven SJA in static flow field. Time-Resolved Particle Image Velocimetry (TR-PIV) system is used to measure the flow field characteristics near the SJA slot exit with input voltage changing, and the flow field snapshots obtained by TR-PIV are modality analyzed by Dynamic Mode Decomposition (DMD) method. The PIV experiments show that by varying input voltage at fixed oscillating frequency, the loudspeaker diaphragm vibration displacement is the parameter that affects the jet velocity and the performance of the SJA. The traveling vortex vanishes at high voltage due to the interaction between vortex structures and the synthetic main jet. In DMD method, the first three-order modes can characterize the main information of the original flow field with reverting the flow field snapshot sequence. It indicates that the DMD method is applicable in the SJA flow field research and the reduced order model can effectively simplify the analysis of flow field.

The ejection test technology is studied in a sub-transonic supersonic wind tunnel using a single cylinder to provide ejection velocity. The angular velocity adjusting device of ejection mechanism is designed, which can adjust the ejection velocity and angular velocity of the model independently. When the ejection cylinder moves downward, the angular velocity adjusting mechanism works at the same time, so that the model has the preset ejection velocity and angular velocity at the moment of leaving the ejection frame. The ejection velocity error is less than 5%, the angular velocity error is less than 10%, and the repetition rate is more than 95%. The new technology has been verified by wind tunnel tests under complex aerodynamic conditions of sub-transonic supersonic and multi-body interference. All parameters have reached or surpassed the existing technical specifications. It has served for model tests many times and met the needs of wind tunnel test research on ejection of embedded weapons in aircraft.

Aiming at the severe aerodynamic heating problem in the leading edge of the hypersonic vehicle, in order to ensure the sharp shape of the leading edge of the wing, a dredging thermal protection structure is proposed, and the built-in high-temperature heat pipe structure is used to provide thermal protection for the leading edge of the wing. By means of numerical simulation and arc wind tunnel test, the dredging thermal protection structure of the leading edge of the wing is analyzed, and the thermal protection effect of the built-in high-temperature heat pipe is obtained. The numerical results show that under certain thermal conditions, the temperature at the leading edge of the wing decreases by 304 K, and the minimum temperature of the tail increases by 130 K. The heat flow is dredged from the high-temperature zone to the low-temperature zone, and the thermal load of the leading edge of the wing is weakened. The same result can be obtained by the arc wind tunnel test, which verifies the accuracy of the numerical method and the feasibility of the dredging thermal protection structure with high-temperature heat pipe embedded in the leading edge of the wing.

The electrothermal anti-icing system is one of the commonly used ice protection systems. In this paper, the heat and mass transfer process on three-dimensional surface of the electrothermal anti-icing system is analyzed. The mass and energy conservation equations are given. A calculation method of the convective heat transfer coefficient on three-dimensional surface is proposed, and the mathematical model of the electrothermal anti-icing system is established. The model is applied to calculate the temperature distribution of the anti-icing system in different conditions. The numerical results are compared with experimental data, and the good agreement between them proves that the developed method is reliable. The results also show that only part of droplets impacted on the leading edge evaporate immediately, while the rest of droplets move downstream in the form of liquid water and evaporate gradually.

Recently, the Gurney Flap (GF) has been used to improve the performance of Horizontal Axis Wind Turbine (HAWT) by enhancing its lift coefficient. Compared to HAWT, research on GF application for Vertical Axis Wind Turbine (VAWT) is very limited. Moreover, most works studied a GF geometry attached to the trailing edge of a stationary airfoil, without considering the rotating effect experienced by VAWT. For this reason, a three-straight-bladed VAWT rotating blade with GF is studied by transient RANS simulation together with a stress-blended eddy simulation (SBES) turbulence model to investigate the GF height effect and the flow characteristics near the blade trailing edge. Results have shown that by introducing the blade rotating, an optimum GF height is found to be 3% of the blade chord, slightly higher than 2% chord in a stationary airfoil case. In addition, the presence of GF can delay deep stall of VAWT blades, thus eliminating negative instantaneous moment coefficient and improving the turbine performance.

The inflatable leading edge (ILE) as a dynamic stall control concept for helicopter rotor blades was investigated numerically on a dynamically pitching airfoil. A fluid–structure interaction (FSI) numerical method for the elastic membrane structure was constructed based on unsteady Reynolds-averaged Navier–Stokes (URANS) equations and mass spring damper (MSD) structural dynamic model. The numerical results indicate that the ILE can change the radius of curvature of the airfoil leading edge, which could reduce the streamwise adverse pressure gradient and suppress the formation of dynamic stall vortex (DSV). Although the maximum lift coefficient of the airfoil is reduced by 8.2%, the maximum drag and pitching moment coefficients of the airfoil are reduced by up to 50.1% and 55.3%, respectively.

During a whole-day period, profiles of mean wind speed, wind shear and turbulence level shows great variability due to continuously varying atmospheric stability. Clearly understanding the spatial and temporal behaviors of the atmospheric wind flow is of great importance for science purposes. Large-eddy simulation (LES) technique is employed here to reproduce the evolution of atmospheric flow during a diurnal cycle. With the obtained LES results, wind characteristics in terms of wind speed, wind shear, turbulence intensity and turbulent kinetic energy can be examined referring to the stability classification. Besides, wind profiles obtained using currently available engineering models are also included for comparison. Disparities between the model predictions and the LES results illustrate that the standard engineering models cannot well capture the wind characteristics driven by the varying atmospheric stability solely, and a further improvement in models is highly needed.

General rigid bead-rod theory [O. Hassager, J. Chem. Phys. 60, 4001 (1974)] explains polymer viscoelasticity from macromolecular orientation. By means of general rigid bead-rod theory, we relate the complex viscosity of polymeric liquids to the architecture of axisymmetric macromolecules. In this paper, we explore the complex viscosities of different axisymmetric diblock copolymer configurations. When nondimensionalized with the zero-shear viscosity, the diblock copolymer complex viscosity depends on the dimensionless frequency and the sole dimensionless architectural parameter, the macromolecular lopsidedness. In this paper, through this way, we thus compare the dimensionless relaxation time of different diblock macromolecular chains. We explore the effects of linear density, macromolecular length, and bead number ratio.

Plate-fin heat exchangers are widely used in industry at present due to their compact structure and high efficiency. However, there is a problem of flow maldistribution, resulting in poor performance of heat exchangers. The influence of the header configuration on fluid flow distribution is studied by using CFD software FLUENT. The numerical results show that the fluid flow inside the header is seriously uneven. The reliability of the numerical simulation is validated against the published results. They are found to be basically consistent within considerable error. The optimal number of the punch baffle is investigated. Various header configuration with different opening ratios have been studied under the same boundary conditions. The gross flow maldistribution parameter (S) is used to evaluate flow nonuniformity, and the flow maldistribution parameters of different schemes under different Reynolds numbers are listed and compared. The optimal header with minimum flow maldistribution parameter is obtained through the performance analysis of headers. It is found that the flow maldistribution of the improved header is significantly smaller compared with the conventional header. Hence, the efficiency of the heat exchanger is effectively enhanced. The conclusion provides a reference for the optimization design of plate-fin heat exchanger.

High-subsonic and low-Reynolds-number flow is a special aerodynamic problem associated with near space propellers and Mars aircrafts. The flow around airfoils and the corresponding aerodynamic performance are different from the incompressible flow at low-Reynolds-number, due to complex shock wave-laminar separation bubble interaction. The objective of this paper is to figure out the effect of Mach number on aerodynamic performance and special flow structure of airfoil. An in-house Reynolds-averaged Navier–Stokes solver coupled with $\gamma -{Re}_{𝜃t}$ transition model is employed to simulate the flows around the E387 airfoil. The results show that the lift slope is larger than $2\pi$ in the linear region. No stall occurs even at an attack angle of $20^{\circ }$. With increase of Mach number, lift coefficient decreases when attack angle is below $10^{\circ }$. However, once the angle of attack exceeds $12^{\circ }$, higher Mach number corresponds to higher lift coefficient. In addition, the strength and number of shock waves are very sensitive to Mach number. With increase of Mach number, the region of reverse flow vortex near transition location becomes smaller and finally disappears, while a new reverse flow vortex appears near the trailing edge and becomes larger.

Electromagnetic energy harvester has been widely concerned in recent years due to its advantages of small size and high sensing frequency. In this paper, the stochastic behaviors of a nonlinear electromagnetic energy harvesting system (NEEH) with fractional damping are investigated under the additive and multiplicative stochastic excitation. Firstly, by applying the stochastic average method to the NEEH system, the mean square of the output current, and the steady-state probability density of vibration amplitude, displacement and velocity are obtained. Meanwhile, the validity of the theoretical results is verified by comparing with the numerical results given by the Monte Carlo method. Secondly, by investigating the theoretical and numerical results, the influences of noise intensity and fractional order on the NEEH system are explored. It is obvious that a higher output voltage can be obtained by the larger intensity of the stochastic excitation, and the smaller coefficient and fractional order of the fractional damping.

Aiming at accurate decomposition and identification of drag, the drag prediction technology based on the mid/far-field integral method is developed. The method decomposes the far-field drag into entropy drag and induced drag according to its physical mechanism, and introduces an appropriate entropy correction to eliminate the numerical dissipation by analyzing the influence of the trailing integral section position on the entropy drag calculation. Based on the analysis of thermodynamic reversible processes and irreversible processes, the drag is refined into viscous drag, shock wave drag, induced drag and pseudo-drag. The mid-field integral method is used to calculate the separate contribution of viscous drag, shock wave drag and induced drag by calculating the limited integral domain. Numerical results show that the developed method is feasible in accurately reflecting the physical mechanism and predicting the drag ratio. Thus, it provides a reliable tool for drag reduction of large passenger aircraft.

Robust design optimization has a great potential application in many engineering fields. In the conventional robust aerodynamics design optimization method, the main difficulty is expensive computational cost related to a large number of function evaluations for uncertainty quantification (UQ). To alleviate the expensive burden for UQ, two levels Kriging surrogate model was introduced. The first level is for the mean value and the second level is for the variances. Through the second level Kriging surrogate models, the method of Monte Carlo Simulation (MCS), which requires a huge number of function evaluations, can be effectively applied to the analysis of variance. Efficient Global Optimization algorithm (EGO) was employed to achieve the global optimized results. To validate the performance of the design method, both one-dimensional function and two-dimensional function were applied. Finally, robust aerodynamics design optimization was applied for a low-drag airfoil. The results show that the optimal solutions obtained from the uncertainty-based optimization formulation are less sensitive to uncertainties to small manufacturing errors.

This study aims to investigate the aeroelastic behaviors of cylindrical composite panels in a high-speed flow. A fluid-structure coupling program is applied to simulate the response of the interaction with the presence of both fluid and structural nonlinearities at $Ma=0.99$. The cylindrical composite panels with different layer orientations, cross-ply and angle-ply are modeled by an assemblage of triangular finite elements and the nonlinearity caused by large displacement is described by an updated tangent stiffness matrix. The inviscid unsteady equations are discretized by an AUSMpw+ flux splitting scheme. The static divergence, flutter characteristic and post-flutter response are computed and analyzed with the increase of the dynamic pressure. The results show that static and dynamic responses vary with the layer orientation. Particularly, the dynamic response of the cross-ply panel presents as high-frequency resonant oscillation; on the contrary, the dynamic response of the angle-ply panel appears as low-frequency single mode oscillation. Nonlinearities involved with aeroelastic response, such as the movement of the multiple-shock structure, and large displacement in structure, are discussed.

An entropy decrease phenomenon in the shock wave is studied. The statistical entropy of a unit mass system (UMS) is constructed based on statistical mechanics. Two terms in the microscopic statistical entropy are connected with the macroscopic entropy increment. In order to obtain the number density and velocity distribution function of argon gas, the direct simulation Monto Carlo method is adopted. The physical mechanism for entropy decrease phenomenon in the shock layer is revealed as compression work winning internal energy to produce the heat loss of a UMS.

In this study, we numerically investigate the shock-induced bubble collapse phenomenon near a tissue-like material. Here, a sharp-interface model is applied to represent a 3-material system containing the liquid, the air and the gelatin. We present the results of the complex interface dynamics and compare them with the experimental data. The goal of this study is to understand the mechanism of the shock wave–bubble interaction near tissues in a more accurate manner with advanced numerical methods.

Coriolis mass flowmeter (CMF) is widely used in the industrial field. In mass flow measurement, there are many impurities in measured fluids that will adhere to the inner wall of the vibrating tube of CMF. The vibration characteristics of CMF would change due to the structural change, i.e., wall clung state, which will generate the wall clung state fault. In this paper, aiming at the wall clung state fault of CMF, the finite element model of CMF is established based on ANSYS. The velocity distribution of fluid in the vibrating tube of CMF is analyzed, considering the fluid–structure interaction. The location of the wall clung state in a vibrating tube is determined. Then, the fault model is established. The mechanism of the vibration transmission characteristics outwards of CMF caused by the wall clung state is analyzed by harmonic response analysis. Finally, the failure mode of CMF is investigated.

A numerical investigation is presented for a free-molecular flow past two parallel circular flat plates using the test particle Monte Carlo (TPMC) method, with flow shadowing and multiple reflections analyzed. Flow shadowing and multiple reflections have a significant effect on aerodynamic forces, and the distance-to-radius ratio of the two plates has a considerable influence on flow shadowing and multiple reflections. As the distance-to-radius ratio increases, flow shadowing and multiple reflections become weaker. Different distance-to-radius ratios result in different surface pressure distributions.

Vortex identification is important for understanding the physical mechanism of turbulent flow. The common vortex identification techniques based on velocity gradient tensor such as $Q$ criterion will consume a lot of computing resources for processing great quantity of experimental data. To improve the vortex identification efficiency and achieve real-time recognition, we present a novel vortex identification method using segmentation with convolutional neural network (CNN) based on flow field image data, which is named “Butterfly-CNN”. Considering that the view of flow field is small, it is necessary to integrate both the local and global feature maps to achieve higher precision. The architecture consists of an encoded–decoded path, which is similar to $U$-net but with different superimposed network part. In the Butterfly-CNN, the cross-expanding paths are designed with the global information to enable precise localization, and the feature maps after each convolution are regarded as the original pictures, then convolute to the size of the last feature map and upsample to the original size again. Finally, the decoded and cross-expanding networks are added up. The Butterfly-CNN can be trained end-to-end from a few images, and it is useful and efficient for vortex identification.

As a result of the increasing demand for aircraft performance, the high-aspect-ratio wing (HARW) becomes the primary structural type of the emerging aircraft as it improves the lift-to-drag ratio. However, HARWs result in higher deflections, leading to a geometrical nonlinear behavior and nonlinear aeroelastic problems. In this work, the preliminary study on aeroelastic behavior of two classic types of HARWs, namely, Goland wing and High Altitude Long Endurance (HALE) wing, are investigated. Aeroelastically scaled models using two scaling methods, called classic scaling method and modified scaling method, are applied and compared with each other. Froude number is considered in the modified scaling method. Both static aeroelastic and flutter analyses are conducted on each scaled model by using different scaling methods. The results show that the Froude number is important to the HALE wing, where both static deflects and flutter velocity obtained by the scaled model agree well with the full model when the Froude number similarity is met. Thus, the Froude number may be the key similar parameter for the HALE wing scaling model design. For the Goland wing scaling model, the Froude number could be ignored.

In high-density data center, energy consumption is increasing dramatically. For reducing the energy consumption, CFD software, Fluent 15.0, is used to simulate the flow and temperature field distribution with $k-𝜀$ turbulence model and fluid–solid coupling method. Fans on the back of racks are simplified as walls with a certain pressure jump. Severs are treated as solid heat sources and porous media. Simulation results reveal that the temperature distribution on the back of racks is not uniform when air conditioners are arranged face-to-face, and local high temperature points emerge near the side wall of air conditioners. Factors affecting cooling efficiency, such as location of air conditioners, speed of inlets, distance of racks, etc., need to be improved. Geometric model is optimized by using a diagonal rack arrangement and drilling holes on the side wall. Based on this, four different cases with various hot aisle distance are proposed. Single and double modular data center are both simulated. Results of new model are better than those of baseline model.

The nonlinear increase of the lift of the double swept waverider at high angles of attack is of vital interest. The aerodynamic performance of the double swept waverider is calculated and compared with that of single swept waveriders. Results suggest that the lift nonlinearity of the double swept waverider is stronger than that of equal-planform-area single swept one, and the nonlinearity increases as Mach number increases. Some scholars have proposed the “vortex lift” to explain the nonlinear lift increase, but it is questionable as the main lift of the waverider comes from the lower surface rather than the upper surface. This paper proposes another explanation that the nonlinear lift increase is related to the attachment of shock wave, influenced by the leading-edge sweep angle. The shock wave is more inclined to attach under the lower surface with smaller swept than that of larger swept as angle of attack increases. When the shock wave attaches, the pressure increase via angle of attack is nonlinear, leading to the nonlinearity of lift increase.