INVESTIGATION OF BOUNDARY-LAYER TRANSITION ON SWEPT WINGS AT MACH NUMBERS 2.5 TO 5

1963 ◽  
Author(s):  
S. R. Pate ◽  
R. E. Brillhart
Keyword(s):  
Boundary Layer ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Mach Numbers ◽  
Swept Wings

The Effects of a Trip Wire and Unsteadiness on a High-Speed Highly Loaded Low-Pressure Turbine Blade

2004 ◽  
Vol 127 (4) ◽  
pp. 747-754 ◽  
Author(s):  
M. Vera ◽  
H. P. Hodson ◽  
R. Vazquez
Keyword(s):  
Boundary Layer ◽  
High Speed ◽  
Low Pressure ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Low Pressure Turbine ◽  
Roughness Elements ◽  
Unsteady Inflow ◽  
Mach Numbers ◽  
Inflow Conditions

This paper presents the effect of a single spanwise two-dimensional wire upon the downstream position of boundary layer transition under steady and unsteady inflow conditions. The study is carried out on a high turning, high-speed, low pressure turbine (LPT) profile designed to take account of the unsteady flow conditions. The experiments were carried out in a transonic cascade wind tunnel to which a rotating bar system had been added. The range of Reynolds and Mach numbers studied includes realistic LPT engine conditions and extends up to the transonic regime. Losses are measured to quantify the influence of the roughness with and without wake passing. Time resolved measurements such as hot wire boundary layer surveys and surface unsteady pressure are used to explain the state of the boundary layer. The results suggest that the effect of roughness on boundary layer transition is a stability governed phenomena, even at high Mach numbers. The combination of the effect of the roughness elements with the inviscid Kelvin–Helmholtz instability responsible for the rolling up of the separated shear layer (Stieger, R. D., 2002, Ph.D. thesis, Cambridge University) is also examined. Wake traverses using pneumatic probes downstream of the cascade reveal that the use of roughness elements reduces the profile losses up to exit Mach numbers of 0.8. This occurs with both steady and unsteady inflow conditions.

High-speed boundary-layer transition induced by a discrete roughness element

2013 ◽  
Vol 729 ◽  
pp. 524-562 ◽  
Author(s):  
Prahladh S. Iyer ◽  
Krishnan Mahesh
Keyword(s):  
Boundary Layer ◽  
Shear Layer ◽  
High Speed ◽  
Roughness Element ◽  
Transition Process ◽  
Significant Rise ◽  
Boundary Layer Transition ◽  
Streamwise Vortices ◽  
Layer Transition ◽  
Mach Numbers

AbstractDirect numerical simulation (DNS) is used to study laminar to turbulent transition induced by a discrete hemispherical roughness element in a high-speed laminar boundary layer. The simulations are performed under conditions matching the experiments of Danehy et al. (AIAA Paper 2009–394, 2009) for free-stream Mach numbers of 3.37, 5.26 and 8.23. It is observed that the Mach 8.23 flow remains laminar downstream of the roughness, while the lower Mach numbers undergo transition. The Mach 3.37 flow undergoes transition closer to the bump when compared with Mach 5.26, in agreement with experimental observations. Transition is accompanied by an increase in ${C}_{f} $ and ${C}_{h} $ (Stanton number). Even for the case that did not undergo transition (Mach 8.23), streamwise vortices induced by the roughness cause a significant rise in ${C}_{f} $ until 20$D$ downstream. The mean van Driest transformed velocity and Reynolds stress for Mach 3.37 and 5.26 show good agreement with available data. Temporal spectra of pressure for Mach 3.37 show that frequencies in the range of 10–1000 kHz are dominant. The transition process involves the following key elements: upon interaction with the roughness element, the boundary layer separates to form a series of spanwise vortices upstream of the roughness and a separation shear layer. The system of spanwise vortices wrap around the roughness element in the form of horseshoe/necklace vortices to yield a system of counter-rotating streamwise vortices downstream of the element. These vortices are located beneath the separation shear layer and perturb it, which results in the formation of trains of hairpin-shaped vortices further downstream of the roughness for the cases that undergo transition. These hairpins spread in the span with increasing downstream distance and the flow increasingly resembles a fully developed turbulent boundary layer. A local Reynolds number based on the wall properties is seen to correlate with the onset of transition for the cases considered.

Comparisons of boundary-layer transition measurement techniques at supersonic Mach numbers

AIAA Journal ◽  
1991 ◽  
Vol 29 (6) ◽  
pp. 865-871 ◽  
Author(s):  
R. M. Hall ◽  
C. J. Obara ◽  
D. L. Garraway ◽  
C. B. Johnson ◽  
R. E. Wright ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Measurement Techniques ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Mach Numbers

Effects of Large Scale High Freestream Turbulence, and Exit Reynolds Number on Turbine Vane Heat Transfer in a Transonic Cascade

2007 ◽  
Author(s):  
S. Nasir ◽  
J. S. Carullo ◽  
W. F. Ng ◽  
K. A. Thole ◽  
H. Wu ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Reynolds Number ◽  
Large Scale ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Layer Transition ◽  
Mach Numbers

This paper experimentally and numerically investigates the effect of large scale high freestream turbulence intensity and exit Reynolds number on the surface heat transfer distribution of a turbine vane in a 2-D linear cascade at realistic engine Mach numbers. A passive turbulence grid was used to generate a freestream turbulence level of 16% and integral length scale normalized by the vane pitch of 0.23 at the cascade inlet. The baseline turbulence level and integral length scale normalized by the vane pitch at the cascade inlet were measured to be 2% and 0.05, respectively. Surface heat transfer measurements were made at the midspan of the vane using thin film gauges. Experiments were performed at exit Mach numbers of 0.55, 0.75 and 1.01 which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 9 × 105, 1.05 × 106, and 1.5 × 106, based on true chord. The experimental results showed that the large scale high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the vane as compared to the low freestream turbulence case and promoted slightly earlier boundary layer transition on the suction surface for exit Mach 0.55 and 0.75. At nominal conditions, exit Mach 0.75, average heat transfer augmentations of 52% and 25% were observed on the pressure and suction side of the vane, respectively. An increased Reynolds number was found to induce earlier boundary layer transition on the vane suction surface and to increase heat transfer levels on the suction and pressure surfaces. On the suction side, the boundary layer transition length was also found to be affected by increase changes in Reynolds number. The experimental results also compared well with analytical correlations and CFD predictions.

Boundary-Layer Transition Measurements at Mach Numbers from 5.4 to 7.4

1960 ◽  
Vol 27 (9) ◽  
pp. 719-720 ◽  
Author(s):  
Fred Landis ◽  
Martin R. Fink ◽  
Murray H. Rosenberg
Keyword(s):  
Boundary Layer ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Mach Numbers

Effects of Large Scale High Freestream Turbulence and Exit Reynolds Number on Turbine Vane Heat Transfer in a Transonic Cascade

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
Shakeel Nasir ◽  
Jeffrey S. Carullo ◽  
Wing-Fai Ng ◽  
Karen A. Thole ◽  
Hong Wu ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Reynolds Number ◽  
Large Scale ◽  
Boundary Layer Transition ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Layer Transition ◽  
Turbine Vane ◽  
Mach Numbers

This paper experimentally and numerically investigates the effects of large scale high freestream turbulence intensity and exit Reynolds number on the surface heat transfer distribution of a turbine vane in a 2D linear cascade at realistic engine Mach numbers. A passive turbulence grid was used to generate a freestream turbulence level of 16% and integral length scale normalized by the vane pitch of 0.23 at the cascade inlet. The base line turbulence level and integral length scale normalized by the vane pitch at the cascade inlet were measured to be 2% and 0.05, respectively. Surface heat transfer measurements were made at the midspan of the vane using thin film gauges. Experiments were performed at exit Mach numbers of 0.55, 0.75, and 1.01, which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 9×105, 1.05×106, and 1.5×106 based on a vane chord. The experimental results showed that the large scale high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the vane as compared to the low freestream turbulence case and promoted a slightly earlier boundary layer transition on the suction surface for exit Mach 0.55 and 0.75. At nominal conditions, exit Mach 0.75, average heat transfer augmentations of 52% and 25% were observed on the pressure and suction sides of the vane, respectively. An increased Reynolds number was found to induce an earlier boundary layer transition on the vane suction surface and to increase heat transfer levels on the suction and pressure surfaces. On the suction side, the boundary layer transition length was also found to be affected by increase changes in Reynolds number. The experimental results also compared well with analytical correlations and computational fluid dynamics predictions.

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