Numerical study of blunt leading edge separation on a 53 degree swept diamond wing (STO AVT-183) using the Edge and Cobalt flow solvers

2015 ◽  
Author(s):  
Henrik Edefur ◽  
Magnus Tormalm ◽  
John Coppin ◽  
Trevor Birch ◽  
R K. Nangia
Keyword(s):  
Numerical Study ◽  
Leading Edge ◽  
Edge Separation ◽  
Blunt Leading Edge

Laminar hypersonic leading edge separation – a numerical study

2017 ◽  
Vol 821 ◽  
pp. 624-646 ◽  
Author(s):  
Amna Khraibut ◽  
S. L. Gai ◽  
L. M. Brown ◽  
A. J. Neely
Keyword(s):  
Wall Temperature ◽  
Numerical Study ◽  
Separated Flow ◽  
Leading Edge ◽  
Slip Boundary Condition ◽  
Navier Stokes ◽  
Recirculation Region ◽  
Slip Boundary ◽  
Limiting Case ◽  
Edge Separation

This paper describes laminar hypersonic leading edge separation. Such a configuration of separated flow was originally studied by Chapman et al. (NACA Tech. Rep. 1356, 1958) at supersonic Mach numbers as it is particularly amenable to theoretical analysis and assumes no pre-existing boundary layer. It can be considered as a limiting case of much studied generic configurations such as separation at a compression corner and separated flow behind a base. A numerical investigation is described using a compressible Navier–Stokes solver assuming perfect gas air, no slip boundary condition and a non-catalytic surface. A moderate enthalpy flow of $3.1\times 10^{6}~\text{J}~\text{kg}^{-1}$ with a unit Reynolds number of $1.34\times 10^{6}~\text{ m}^{-1}$ and a Mach number of 9.66 was considered. The resulting separated flow is analysed in the context of viscous–inviscid interaction and interpreted in terms of ‘triple-deck’ concepts. Particular emphasis is given to wall temperature effects. The effects of strong to moderate wall cooling on flow in the separated region as well as on processes of separation, reattachment and separation length, are highlighted. The numerical simulations have also shown the existence of a secondary eddy embedded within the primary recirculation region, with its size, shape and position, being strongly affected by the wall temperature.

The Effect of an Eroded Leading Edge on the Aerodynamic Performance of a Transonic Fan Blade Cascade

2014 ◽  
Author(s):  
A. Hergt ◽  
J. Klinner ◽  
W. Steinert ◽  
S. Grund ◽  
M. Beversdorff ◽  
...  
Keyword(s):  
High Speed ◽  
Numerical Study ◽  
Separation Bubble ◽  
Leading Edge ◽  
Suction Side ◽  
Chord Length ◽  
Blade Cascade ◽  
Shock Position ◽  
Fan Blade ◽  
Blunt Leading Edge

Especially at transonic flow conditions the leading edge shape influences the performance of a fan profile. At the same time the leading edge of a fan profile is highly affected by erosion during operation. This erosion leads to a deformation of the leading edge shape and a reduction of the chord length. In the present experimental and numerical study, the aerodynamic perfomance of an original fan profile geometry is compared to an eroded fan profile with a blunt leading edge and a chord length reduced by about 1 percent. The experiments are performed at a linear fan blade cascade in the Transonic Cascade Wind Tunnel of DLR in Cologne. The inflow Mach number during the tests is 1.25 and the Reynolds number 1.5 × 106. All tests are carried out at a low inflow turbulence level of 0.8 percent. The results of the investigation show that losses are increased over the whole operating range of the cascade. At the aerodynamic design point the losses raise by 25 percent. This significant loss increase can be traced back to the increase of the shock losses at the leading edge. The change in shock structure is investigated and described in detail by means of PIV measurements and Schlieren imaging. Additionally, the unsteady fluctuation of the shock position is measured by a high speed shadowgraphy. Then the frequency range of the fluctuation is obtained by a Fourier analysis of the time resolved shock position. Furthermore, liquid crystal measurements are performed in order to analyze the influence of the leading edge shape on the development of the suction side boundary layer. The results show that for the original fan blade the transition occurs at the shock position on the blade suction side by a separation bubble whereas the transition onset is shifted upstream for the fan blade with the blunt leading edge.

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