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The welding kinetics of interface evolution and mechanism inIN718 alloy solid-state welding are investigated systematically by comparing the hot compression behavior of the separated samples and conventional samples in the temperature range of 1000 °C–1160 °C and strain rate range of 0.1/s to 5/s. The separated samples obtain higher peak stress than conventional samples, and the kinetics analysis results show that separated samples provide higher structural factors and deformation activation energy. The activation energy of the interfacial microstructure evolution of the IN718 alloy was calculated to be 29.7948 kJ/mol. The evolution mechanism of the interface microstructure is characterized by the elimination of the interface second phase and the migration of straight interfacial grain boundaries (IGBs), which is inseparable from the occurrence of interfacial dynamic recrystallization. Interfacial dynamical recrystallization helps to regulate the chemical potential around the second phase for a continuous dissociation. The migration mechanism of the IGBs is influenced by the strain rates, which is continuous dynamic recrystallization (CDRX) for high strain rates and discontinuous dynamic recrystallization (DDRX) for low strain rates.

The improved delayed detached-eddy simulation (IDDES) method is used to simulate the reverse flows past an NACA0012 airfoil at medium (10°) and large (30°) angles of attack. The numerical results of the baseline configuration are compared with the available measurements. The effects of the undulating leading edge with four different amplitudes are compared and analyzed at angle of attack of 10°. Based on these analyses, the amplitude of A/C=0.04 yields the best performance. Compared with the uncontrolled case, the performances of the undulating leading edge are greatly improved with reducing of the aerodynamic fluctuations. Furthermore, the mechanisms of performance are explored by comparing the local flow structures near the undulations.

Hypersonic boundary layer transition from laminar to turbulent induced by three isolated roughness elements (ramp, diamond, and cylinder) was investigated by direct numerical simulations based on finite volume formulation with high order MDCD scheme. The effects of roughness shape on the transition mechanism, aerodynamic and aero-thermodynamic performances were quantitatively compared. The transition is dominated by three-dimensional shear layer instability and the streamwise transition locations are almost the same for all three roughness elements. The width of turbulent wake, the drag increment and the heat flux are the largest for cylinder. At the same time, the turbulent wake is the narrowest and drag and heat flux are the lowest for the ramp. [ABSTRACT FROM AUTHOR]

An unsteady Reynolds averaged Navier–Stokes (URANS) method combined with a rigid dynamic mesh technique was developed to simulate unsteady flows around complex configurations during pitching motion. First, a test case with the NACA0012 airfoil was selected to validate the numerical methods and our in-house codes. Then, we evaluated the unsteady flows around an advanced aircraft model during harmonic pitching motion at high incidence. The effects of pitching motion on the hysteresis of aerodynamic force, the evolution of the leading-edge vortex, and the distribution of pressure on the model’s surface were analyzed in detail. The roles of several significant parameters such as the reduced frequency and pitching amplitude were revealed. Several conclusions were found: pitching motion would delay the initiation of the leading-edge vortex, strengthen the vorticity, postpone the occurrence of vortex breakdown, and weaken the massively separated flows, thus causing additional aerodynamic force. Two categories of critical reduced frequency have been found, which divide the influence of reduced frequency on aerodynamic force into three stages, called the linear increasing range, slowly increasing range, and constant range. The first-order phase lag between the aerodynamic force and the incidence is a constant that is independent of the amplitude when the reduced frequency is sufficiently high. A scaled maximum value of CLis proposed; it depends only on the reduced frequency (instead of the amplitude), and increases linearly when the reduced frequency is sufficiently low.

Hypersonic boundary layer transition induced by an isolated cylindrical roughness element is investigated using direct numerical simulation method based on a finite volume formulation. To simulate the transition procedure by resolving the generation and evolvement of small-scale coherent structures, and capture the shock wave at the same time, high-order minimum dispersion and controllable dissipation scheme is validated and then applied. The results are compared with the available measurements in the quiet wind tunnel, such as the dominated frequency and root mean square of pressure. The computational dominated frequency of 19.23 kHz is very close to the experimental one, 21 kHz. Also, the disturbances of the roughness are mostly generated by the “jet” just before the roughness, and then they travel and develop downstream with the shear layer and vortex shedding. The transition is mainly dominated by the instabilities of both the horseshoe vortex and the shear layer.

An improved delayed detached eddy simulation (IDDES) method based on the k-ω-SST (shear stress transport) turbulence model was applied to predict the unsteady vortex breakdown past an 80°/65° double-delta wing (DDW), where the angles of attack (AOAs) range from 30° to 40°. Firstly, the IDDES model and the relative numerical methods were validated by simulating the massively separated flow around an NACA0021 straight wing at the AOA of 60°. The fluctuation properties of the lift and pressure coefficients were analyzed and compared with the available measurements. For the DDW case, the computations were compared with such measurements as the mean lift, drag, pitching moment, pressure coefficients and breakdown locations. Furthermore, the unsteady properties were investigated in detail, such as the frequencies of force and moments, pressure fluctuation on the upper surface, typical vortex breakdown patterns at three moments, and the distributions of kinetic turbulence energy at a stream wise section. Two dominated modes are observed, in which their Strouhal numbers are 1.0 at the AOAs of 30°, 32° and 34° and 0.7 at the AOAs of 36°, 38° and 40°. The breakdown vortex always moves upstream and downstream and its types change alternatively. Furthermore, the vortex can be identified as breakdown or not through the mean pressure, root mean square of pressure, or even through correlation analysis.

Abstract: In the present paper, extremely unsteady shock wave buffet induced by strong shock wave/boundary-layer interactions (SWBLI) on the upper surface of an OAT15A supercritical airfoil at Mach number of 0.73 and angle of attack of 3.5 degrees is first numerically simulated by IDDES, one of the most advanced RANS/LES hybrid methods. The results imply that conventional URANS methods are unable to effectively predict the buffet phenomenon on the wing surface; IDDES, which involves more flow physics, predicted buffet phenomenon. Some complex flow phenomena are predicted and demonstrated, such as periodical oscillations of shock wave in the streamwise direction, strong shear layer detached from the shock wave due to SWBLI and plenty of small scale structures broken down by the shear layer instability and in the wake. The root mean square (RMS) of fluctuating pressure coefficients and streamwise range of shock wave oscillation reasonably agree with experimental data. Then, two vortex generators (VG) both with an inclination angle of 30 degrees to the main flow directions are mounted in front of the shock wave region on the upper surface to suppress shock wave buffet. The results show that shock wave buffet can be significantly suppressed by VGs, the RMS level of pressure in the buffet region is effectively reduced, and averaged shock wave position is obviously pushed downstream, resulting in increased total lift.

Abstract: The crossflow instability of a three-dimensional boundary layer is a main factor affecting the transition around the swept-wing. The three-dimensional boundary layer flow affected by the saturated crossflow vortex is very sensitive to the high frequency disturbances, which foreshadows that the swept wing flow transition will happen. The primary instability of the compressible flow over a swept wing is investigated with nonlinear parabolized stability equations (NPSE). The Floquet theory is then applied to the analysis of the influence of localized steady suction on the secondary instability of crossflow vortex. The results show that suction can significantly suppress the growth of the crossflow mode as well as the secondary instability modes.

Abstract: The crossflow instability of a three-dimensional (3-D) boundary layer is an important factor which affects the transition over a swept-wing. In this report, the primary instability of the incompressible flow over a swept wing is investigated by solving nonlinear parabolized stability equations (NPSE). The Floquet theory is applied to study the dependence of the secondary and high-frequency instabilities on curvature, Reynolds number and angle of swept (AOS). The computational results show that the curvature in the present case has no significant effect on the secondary instabilities. It is generally believed that the secondary instability growth rate increases with the magnitude of the nonlinear mode of crossflow vortex. But, at a certain state, when the Reynolds number is 3.2 million, we find that the secondary instability growth rate becomes smaller even when the magnitude of the nonlinear mode of the crossflow vortex is larger. The effect of the angle of swept at 35, 45 and 55 degrees, respectively, is also studied in the framework of the secondary linear stability theory. Larger angles of swept tend to decrease the spanwise spacing of the crossflow vortices, which correspondingly helps the stimulation of ‘z’ mode.

Abstract: Flows around vortex generators (VGs), which serve as one of the important flow control methods, are investigated by solving Reynolds-averaged Navier-Stokes (RANS) equations. The influences on the main flow of VGs are intended to explore. Firstly, the flow around a single VG on a flat plane is computed to validate the schemes and to acquire basic knowledge of this kind of flow. Secondly, transonic flow past a standard model, named by ONERA-M6 wing, is predicted to investigate the flow features of shockwave/boundary-layer interactions (SWBLI). Thirdly, the effects of a row of VGs mounted about 25% local chord on a supercritical wing are analyzed in transonic condition with strong SWBLI. Lastly, VGs are mounted more upwind (about 3.5% local chord) to explore the effects at low speed and high incidence condition. The numerical results show that seven VGs can effectively suppress the separations behind the strong SWBLI and decrease spanwise flow and wing-tip vortex in transonic condition. VGs also can decrease the large scope of separation over the wing at low speed with high angle of attack.

• Validation of IDDES with the grid scale considering vorticity vectors. • Flight velocity effect on the mixing layer properties. • Flight velocity effect on the instantaneous end of potential core. • Flight velocity effect on the intermittency of jet centerline. The flight velocity effect on a cold jet flow is numerically predicted by the improved delayed detached eddy simulation (IDDES) with a modification to the grid length scale, mainly focusing on the near-field flows and intermittency properties in the jet centerline region. First, the modification in IDDES is validated using three meshes and available measurements. This method can reduce the modeled eddy viscosity effectively and improve the performance of original IDDES. Then, the flight velocity effects on the mean flow, turbulence, and intermittency are studied. From the results, the fluctuation intermittency near the end of potential core is caused by the intrusion of turbulent shear-layer structures into the jet core. When the flight velocity increases, the intermittency factor in the jet centerline region is reduced, and its peak moves downstream. The skewness and kurtosis factors perform in a similar manner. [ABSTRACT FROM AUTHOR]