Numerical Analysis of the ONERA-M6 Wing with Wind Tunnel Wall Interference

Nambu, Taisuke; Hashimoto, Atsushi; Aoyama, Takashi; Satô, Tetsuya

doi:10.2322/tjsass.58.7

Cited by 6 publications

(2 citation statements)

References 16 publications

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“…His results indicated that the swept leading edge and the small curvature of an airfoil arc produce the formation of a λconfiguration of the local supersonic regions, thereby forming the λ-shape of pressure contours on the upper surface of the wing. Furthermore, the ONERA M6 wing has been widely used to validate computational results obtained by computation fluid dynamics, including further studies on the complicated phenomena of the transonic flow field and others [9][10][11]. So far, investigations on transonic flow also revealed the presence of a double supersonic-flow regime, including a formation of these regimes on airfoils that are comprised of a nearly flat arc [12][13].…”

Section: Introductionmentioning

confidence: 99%

Parametric Study on Wing-Lambda-Shock Formation

Chainok

Rungroch

Chairach

et al. 2021

Volume 1: Aerospace Engineering Division Joint Track; Computational Fluid Dynamics

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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.

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Section: Introductionmentioning

confidence: 99%

Parametric Study on Wing-Lambda-Shock Formation

Chainok

Rungroch

Chairach

et al. 2021

Volume 1: Aerospace Engineering Division Joint Track; Computational Fluid Dynamics

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“…The researchers in Johan et al (1995) have examined the solution of steady laminar viscous flow over ONERA M6 wing using parallel CFD computations and unstructured mesh. The Navier–Stokes equations for the flow around ONERA M6 wing have been solved in Nambu et al (2015). In this CFD simulations study, the effects of wind tunnel wall interference have been considered and the same shock wave as in the experimental results has been nearly captured.…”

Section: Introductionmentioning

confidence: 99%

Taper stacking for the aerodynamic performance of wings

Kaya

Elfarra

2020

AEAT

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Purpose The critical Mach number, lift-to-drag ratio and drag force play important role in the performance of the wings. This paper aims to investigate the effect of taper stacking, which has been used to generalize wing sweeping, on those parameters. Design/methodology/approach The results obtained are based on steady-state turbulent flowfields computations. The baseline wing is ONERA M6. Various wing planforms are generated by linearly or parabolically varying the spanwise stacking location. The critical Mach number is determined by changing the freestream Mach number for a fixed angle of attack. On the other hand, the analysis of the drag force is carried out by changing the angle of attack to keep the lift force constant. Findings By changing the stacking location, the critical Mach number and the corresponding lift-to-drag ratio have increased by around 7 and 3%, respectively. A reduction of 12.8% in total drag force has been observed in one of the analyzed cases. Moreover, there exist some cases in which the values of drag reduce significantly while the lift is the same. Practical implications The results of this new stacking approach have implied that the drag force can be decreased without decreasing the lift. This outcome is valuable for increasing the range and endurance of an aircraft. Originality/value This work generalizes wing sweeping by modifying the taper stacking along the span. In literature, wing sweep is enhanced using segmented stacking of taper distribution. The present study is further enhancing this concept by introducing continuous stacking (infinite number of stacking segments) for the first time.

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