Parametric Study on Wing-Lambda-Shock Formation

2021 ◽  
Author(s):  
Sirikorn Chainok ◽  
Thanapol Rungroch ◽  
Pattarasuda Chairach ◽  
Prasert Prapamonthon ◽  
Soemsak Yooyen ◽  
...  
Keyword(s):  
Mach Number ◽  
Free Stream ◽  
Angle Of Attack ◽  
Shock Formation ◽  
Local Pressure ◽  
Drag Coefficients ◽  
Pressure Coefficients ◽  
Free Stream Mach Number ◽  
Lambda Shock ◽  
Lift And Drag

Abstract It is well-known that a wing is one of the most important parts of an aircraft as it is used to generate lift force. According to a wing moving at sufficiently high subsonic speeds, the flow speed on the wing’s upper surface can be supersonic due to acceleration through the curvature-created suction, thereby forming a shock wave in a lambda shape. Additionally, the lambda shock can interact with the boundary layer flow. These phenomena relate to disturbances in the flow field, including flow separation, thus causing undesirable effects on lift production. Hence, a better understanding of the phenomenon of wing-lambda-shock formation and its nature is essential. This study presents a numerical investigation of the lambda-shock formation on an ONERA M6 wing, which is known as a swept, semi-span wing with no twist, under parametric effects of angle-of-attack, and free-stream Mach number, which is increased up to the supersonic regime. The pressure coefficients obtained by simulations are validated by open data. Then, numerical results in terms of the local pressure coefficient, local Mach number, averaged lift and drag coefficients, and λ-shape characteristics based on Mach number and pressure coefficients are discussed under an investigated range of the parameters. Results show that the angle-of-attack and free-stream Mach number can affect the lambda shock formation on the wing upper surface physically. Specifically, an iso-sonic surface with lambda shock waves is disturbed when the angle-of-attack and free-stream Mach number vary in an investigated range. This also affects lift and drag coefficients of the wing.

Direct numerical simulation of flow past a transversely rotating sphere up to a Reynolds number of 300 in compressible flow

2018 ◽  
Vol 857 ◽  
pp. 878-906 ◽  
Author(s):  
T. Nagata ◽  
T. Nonomura ◽  
S. Takahashi ◽  
Y. Mizuno ◽  
K. Fukuda
Keyword(s):  
Numerical Simulation ◽  
Reynolds Number ◽  
Mach Number ◽  
Direct Numerical Simulation ◽  
Drag Coefficient ◽  
Rotation Rate ◽  
Free Stream ◽  
Rotating Sphere ◽  
Free Stream Mach Number ◽  
Low Reynolds

In this study, direct numerical simulation of the flow around a rotating sphere at high Mach and low Reynolds numbers is conducted to investigate the effects of rotation rate and Mach number upon aerodynamic force coefficients and wake structures. The simulation is carried out by solving the three-dimensional compressible Navier–Stokes equations. A free-stream Reynolds number (based on the free-stream velocity, density and viscosity coefficient and the diameter of the sphere) is set to be between 100 and 300, the free-stream Mach number is set to be between 0.2 and 2.0, and the dimensionless rotation rate defined by the ratio of the free-stream and surface velocities above the equator is set between 0.0 and 1.0. Thus, we have clarified the following points: (1) as free-stream Mach number increased, the increment of the lift coefficient due to rotation was reduced; (2) under subsonic conditions, the drag coefficient increased with increase of the rotation rate, whereas under supersonic conditions, the increment of the drag coefficient was reduced with increasing Mach number; and (3) the mode of the wake structure becomes low-Reynolds-number-like as the Mach number is increased.

scholarly journals Звуковой удар от тонкого тела и локальных областей нагрева сверхзвукового набегающего потока

2021 ◽  
Vol 91 (4) ◽  
pp. 558
Author(s):  
А.В. Потапкин ◽  
Д.Ю. Москвичев
Keyword(s):  
Mach Number ◽  
Supersonic Flow ◽  
Free Stream ◽  
Local Heating ◽  
Air Flow ◽  
Sonic Boom ◽  
Incoming Flow ◽  
Combined Method ◽  
Cold Flow ◽  
Free Stream Mach Number

The problem of a sonic boom generated by a slender body and local regions of supersonic flow heating is solved numerically. The free-stream Mach number of the air flow is 2. The calculations are performed by a combined method of phantom bodies. The results show that local heating of the incoming flow can ensure sonic boom mitigation. The sonic boom level depends on the number of local regions of incoming flow heating. One region of flow heating can reduce the sonic boom by 20% as compared to the sonic boom level in the cold flow. Moreover, consecutive heating of the incoming flow in two regions provides sonic boom reduction by more than 30%.

Wake measurements of oscillating supercritical airfoil in compressible flow

2019 ◽  
Vol 43 (1) ◽  
pp. 112-121
Author(s):  
Behnaz Beheshti Boroumand ◽  
Mahmoud Mani
Keyword(s):  
Mach Number ◽  
Angle Of Attack ◽  
Time History ◽  
Mean Value ◽  
Hot Wire ◽  
Dynamic Tests ◽  
Flow Parameters ◽  
Free Stream Mach Number ◽  
Static And Dynamic Tests ◽  
Wake Characteristics

Boundary layer and wake behaviors are strongly affected by airfoil motion. Moreover, parameters like body oscillation frequency, oscillation type, Mach number, and angle of attack play main roles in wake characteristics. In this research, both static and dynamic tests were carried out in a tri-sonic wind tunnel to study wake profiles experimentally by hot wire anemometry. All data were recorded at a free stream Mach number of 0.4. Quarter-length and half-length of chord were also considered as downstream distances from the trailing edge in pitching motions of mean angle of attack of −0.4°. Frequencies of 3 Hz and 6 Hz with amplitude of 3° were chosen as oscillation parameters. Voltages at hot wire outputs were measured and analyzed qualitatively and statistically with root-mean-square, correlation, mean value distribution, time history, and frequency. Flow parameters were obtained by computational studies under similar experimental test conditions. The wake characteristics obtained from numerical and experimental methods were compared.

Subsonic Flow Over Open Cavities: Part 1 — Analytical Models and Numerical Investigations

2016 ◽  
Author(s):  
Weidong Shao ◽  
Jun Li
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Shear Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Feedback Loop ◽  
Subsonic Flow ◽  
Free Stream ◽  
Open Cavities ◽  
Free Stream Mach Number

The aeroacoustical oscillation and acoustic field generated by subsonic flow grazing over open cavities has been investigated analytically and numerically. The tone generation mechanism is elucidated with an analytical model based on the coupling between shear layer instabilities and acoustic feedback loop. The near field turbulent flow is obtained using two-dimensional Large Eddy Simulation (LES). A special mesh is used to absorb propagating disturbances and prevent spurious numerical reflections. Comparisons with available experimental data demonstrate good agreement in both the frequency and amplitude of the aeroacoustical oscillation. The physical phenomenon of the noise generated by the feedback loop is discussed. The correlation analysis of primitive variables is also made to clarify the characteristics of wave propagation in space and time. The effects of free-stream Mach number and boundary layer thickness on pressure fluctuations within the cavity and the nature of the noise radiated to the far field are examined in detail. As free-stream Mach number increases velocity fluctuations and mass flux into the cavity increase, but the resonant Strouhal numbers slightly decrease. Both the resonant Strouhal numbers and sound pressure levels decrease with the increase of boundary layer thickness. Results indicate that the instability of the shear layer dominates both the frequency and amplitude of the aeroacoustical oscillation.

Second-order boundary-layer effects in hypersonic flow past axisymmetric blunt bodies

1964 ◽  
Vol 20 (4) ◽  
pp. 593-623 ◽  
Author(s):  
R. T. Davis ◽  
I. Flügge-Lotz
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Free Stream ◽  
Practical Interest ◽  
Second Order ◽  
The Body ◽  
Specific Flow ◽  
Flow Past ◽  
Free Stream Mach Number ◽  
Stability And Accuracy

First- and second-order boundary-layer theory are examined in detail for some specific flow cases of practical interest. These cases are for flows over blunt axisymmetric bodies in hypersonic high-altitude (or low density) flow where second-order boundary-layer quantities may become important. These cases consist of flow over a hyperboloid and a paraboloid both with free-stream Mach number infinity and flow over a sphere at free-stream Mach number 10. The method employed in finding the solutions is an implicit finite-difference scheme. It is found to exhibit both stability and accuracy in the examples computed. The method consists of starting near the stagnation-point of a blunt body and marching downstream along the body surface. Several interesting properties of the boundary layer are pointed out, such as the nature of some second-order boundary-layer quantities far downstream in the flow past a sphere and the effect of strong vorticity interaction on the second-order boundary layer in the flow past a hyperboloid. In several of the flow cases, results are compared with other theories and experiments.

Performance study of an inlet in supersonic flow

2012 ◽  
Vol 227 (1) ◽  
pp. 159-174 ◽  
Author(s):  
Mohammad R Soltani ◽  
Mohammad Farahani
Keyword(s):  
Mach Number ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Free Stream ◽  
Characteristic Curve ◽  
Pressure Recovery ◽  
Internal Boundary ◽  
Performance Study ◽  
Free Stream Mach Number

The performance characteristics of an axisymmetric inlet at its design and off-design operational conditions are experimentally investigated. The model is tested for wide ranges of free stream Mach numbers, M∞ = 1.5–2.5, and mass flow rates. For each test, the pressure recovery, the mass flow passing through the inlet and the pressure distribution over the spike and the cowl are measured. In addition, the shock wave formed in front of the inlet is visualized. The characteristic curve of the inlet is then obtained for each free stream Mach number. As the Mach number is increased, the pressure recovery is reduced, but the maximum value of the mass flow rate grows up. Variations of the mass flow affect the surface pressure over both the front portion of the cowl and the entire surface of the spike. Further, it has changed both pressure and Mach number at the end of the diffuser, which would consequently affect the performance of the propulsion system. In addition, contrary to the internal boundary layer, the external one far from the cowl lip has been found to be almost independent of the inlet mass flow rate for a constant free stream Mach number.

scholarly journals Nonlinear unsteady streaks engendered by the interaction of free-stream vorticity with a compressible boundary layer

2017 ◽  
Vol 817 ◽  
pp. 80-121 ◽  
Author(s):  
Elena Marensi ◽  
Pierre Ricco ◽  
Xuesong Wu
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Wind Tunnel ◽  
Boundary Layers ◽  
Free Stream ◽  
Nonlinear Interactions ◽  
Compressible Boundary Layer ◽  
Displacement Effect ◽  
Free Stream Mach Number ◽  
First Time

The nonlinear response of a compressible boundary layer to unsteady free-stream vortical fluctuations of the convected-gust type is investigated theoretically and numerically. The free-stream Mach number is assumed to be of $O(1)$ and the effects of compressibility, including aerodynamic heating and heat transfer at the wall, are taken into account. Attention is focused on low-frequency perturbations, which induce strong streamwise-elongated components of the boundary-layer disturbances, known as streaks or Klebanoff modes. The amplitude of the disturbances is intense enough for nonlinear interactions to occur within the boundary layer. The generation and nonlinear evolution of the streaks, which acquire an $O(1)$ magnitude, are described on a self-consistent and first-principle basis using the mathematical framework of the nonlinear unsteady compressible boundary-region equations, which are derived herein for the first time. The free-stream flow is studied by including the boundary-layer displacement effect and the solution is matched asymptotically with the boundary-layer flow. The nonlinear interactions inside the boundary layer drive an unsteady two-dimensional flow of acoustic nature in the outer inviscid region through the displacement effect. A close analogy with the flow over a thin oscillating airfoil is exploited to find analytical solutions. This analogy has been widely employed to investigate steady flows over boundary layers, but is considered herein for the first time for unsteady boundary layers. In the subsonic regime the perturbation is felt from the plate in all directions, while at supersonic speeds the disturbance only propagates within the dihedron defined by the Mach line. Numerical computations are performed for carefully chosen parameters that characterize three practical applications: turbomachinery systems, supersonic flight conditions and wind tunnel experiments. The results show that nonlinearity plays a marked stabilizing role on the velocity and temperature streaks, and this is found to be the case for low-disturbance environments such as flight conditions. Increasing the free-stream Mach number inhibits the kinematic fluctuations but enhances the thermal streaks, relative to the free-stream velocity and temperature respectively, and the overall effect of nonlinearity becomes weaker. An abrupt deviation of the nonlinear solution from the linear one is observed in the case pertaining to a supersonic wind tunnel. Large-amplitude thermal streaks and the strong abrupt stabilizing effect of nonlinearity are two new features of supersonic flows. The present study provides an accurate signature of nonlinear streaks in compressible boundary layers, which is indispensable for the secondary instability analysis of unsteady streaky boundary-layer flows.

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