Effect of Surface Roughness on Conjugate Heat Transfer of a Turbine Vane

2016 ◽  
Author(s):  
Hongyang Li ◽  
Yun Zheng
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Surface Roughness ◽  
Suction Side ◽  
Transition Process ◽  
Considerable Effect ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Turbine Vane ◽  
Effect Of Surface

For the purpose of researching the effect of surface roughness on boundary layer transition and heat transfer of turbine blade, a roughness modification approach for γ-Reθ transition model was proposed based on an in-house CFD code. Taking surface roughness effect into consideration, No. 5411 working condition of Mark II turbine vane was simulated and the results were analyzed in detail. Main conclusions are as follows: Surface roughness has little effect on heat transfer of laminar boundary layer, while has considerable effect on turbulent boundary layer. Compared with smooth surface, equivalent sand roughness of 100μm increases the temperature for about 28.4K on suction side, reaching an increase of 5%. Under low roughness degree, effect of shock wave dominants on boundary layer transition process on suction side, while above the critical degree, effect of surface roughness could abruptly change the transition point.

Studies on Transitional Heat Transfer Characteristics Over Turbine Vane Surface Using a High Order LES Approach

2014 ◽  
Author(s):  
Debasish Biswas
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Gas Turbine ◽  
Free Stream ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Free Stream Turbulence ◽  
Turbine Vane ◽  
Turbine Airfoils

The boundary layer developing on a turbo-machinery blade usually starts as a laminar layer but in most situations it inevitably becomes turbulent. The transition from laminar to turbulent in the boundary layer, which often causes a significant change in operational performance of the machinery, is generally influenced by the free-stream turbulence level, the pressure gradient, and surface curvature, etc. Therefore, boundary layer transition is an important phenomenon experienced by the flow through gas turbine engines. A substantial fraction of the boundary layer on both sides of a gas turbine airfoil may be transitional. The extended transition zone exist due to strong favorable pressure gradients, found on both near the leading edge portion of the suction side and the pressure side, which serve to stabilize the boundary layer and consequently delay the transition process, even under high free-stream turbulence intensity (FSTI) in practical gas turbine. It is very important to properly model and predict the high FSTI transition mechanism, since boundary layer transition leads to substantial increase in friction coefficients and heat transfer rate. Boundary layer separation, which is expected to be a significant problem on the suction side of some high pressure turbine airfoils due to shock-boundary layer interaction, also depends strongly on the state of boundary layer with respect to transition. Acceleration rates, Reynolds numbers and FSTI play very important role in controlling the boundary layer transition on the pressure side of gas turbine airfoils. The main objective of the present work is to study the performance of a high order LES turbulence model in predicting the transitional heat transfer characteristics over turbine vane surface under high pressure turbine flow conditions. In this regard the model is assessed to the precise experimental data where measurements were carried out in moderate temperature using three-vane cascades under steady state conditions. Two types of vane configurations were used in the experiment. The aerodynamic configurations of the two vanes were carefully selected to emphasize fundamental differences in the character of suction surface pressure distributions and the consequent effect on surface heat transfer distributions. In both the experiments and the computations, principle independent parameters (Mach number, Reynolds number, turbulence intensity, and wall-to-gas temperature ratio) were varied over ranges consistent with actual engine operation. The computed results explained measured data very satisfactorily and helped to have a very good understanding of basic mechanism involved in the complex flow behavior and transition from laminar to turbulent flow.

A Review of Surface Roughness Effects in Gas Turbines

2010 ◽  
Vol 132 (2) ◽  
Author(s):  
J. P. Bons
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Surface Roughness ◽  
Gas Turbines ◽  
Engine Performance ◽  
Boundary Layer Transition ◽  
Roughness Effects ◽  
Layer Transition ◽  
Laminar Separation

The effects of surface roughness on gas turbine performance are reviewed based on publications in the open literature over the past 60 years. Empirical roughness correlations routinely employed for drag and heat transfer estimates are summarized and found wanting. No single correlation appears to capture all of the relevant physics for both engineered and service-related (e.g., wear or environmentally induced) roughness. Roughness influences engine performance by causing earlier boundary layer transition, increased boundary layer momentum loss (i.e., thickness), and/or flow separation. Roughness effects in the compressor and turbine are dependent on Reynolds number, roughness size, and to a lesser extent Mach number. At low Re, roughness can eliminate laminar separation bubbles (thus reducing loss) while at high Re (when the boundary layer is already turbulent), roughness can thicken the boundary layer to the point of separation (thus increasing loss). In the turbine, roughness has the added effect of augmenting convective heat transfer. While this is desirable in an internal turbine coolant channel, it is clearly undesirable on the external turbine surface. Recent advances in roughness modeling for computational fluid dynamics are also reviewed. The conclusion remains that considerable research is yet necessary to fully understand the role of roughness in gas turbines.

Numerical Simulation of Turbulent Impingement Cooling

2001 ◽  
Author(s):  
Andreas Abdon ◽  
Bengt Sundén
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Turbulence Models ◽  
Local Heat Transfer ◽  
Transition Process ◽  
Local Heat ◽  
Boundary Layer Transition ◽  
Wall Jet ◽  
Layer Transition ◽  
Stagnation Region

Simulations of turbulent impinging jet heat transfer for different nozzle configurations using Reynolds averaged governing equations and two-equation turbulence models have been conducted. The considered nozzle configurations are a square-edged orifice and a pipe exit. The results for a jet Reynolds number of 10000 and dimensionless nozzle-to-plate distance of 2 show that the heat transfer is well predicted for the pipe configuration but underpredicted for the orifice. The disagreement may be partly explained by underprediction of turbulence in the stagnation region and inaccurate treatment of the wall jet boundary layer transition. An investigation of the local heat transfer distribution for the orifice reveals two local maxima. These are related to an accelerating laminar boundary layer and the transition process of the wall jet, respectively, for the calculations. The application of a realizability constraint on the models leads to reduced turbulence levels, not only in the stagnation region, but also in the throttled flow of the orifice configuration. This improves the prediction of heat transfer and nozzle exit turbulence levels significantly.

Roughness, bluntness, and angle-of-attack effects on hypersonic boundary-layer transition

1966 ◽  
Vol 24 (1) ◽  
pp. 1-31 ◽  
Author(s):  
H. T. Nagamatsu ◽  
B. C. Graber ◽  
R. E. Sheer
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Surface Roughness ◽  
Reynolds Number ◽  
Mach Number ◽  
Angle Of Attack ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Number Range ◽  
Transfer Data

An investigation was conducted in a hypersonic shock tunnel to study the laminar boundary-layer transition on a highly cooled 10° cone of 4 ft. length over the Mach-number range of 8·5 to 10·5 with a stagnation temperature of 1400 °K. The effects on transition of tip surface roughness, tip bluntness, and ± 2° angle of attack were investigated. With fast-response, thin film surface heat-transfer gauges, it was possible to detect the passage of turbulent bursts which appeared at the beginning of transition. Pitot-tube surveys and schlieren photographs of the boundary layer were obtained to verify the interpretation of the heat-transfer data. It was found that the surface roughness greatly promoted transition in the proper Reynolds-number range. The Reynolds numbers for the beginning and end of transition at the 8·5 Mach-number location were 3·8 × 106−9·6 × 106and 2·2 × 106−4·2 × 106for the smooth sharp tip and rough sharp tip respectively. The local skin-friction data, determined from the Pitot-tube survey, agreed with the heat-transfer data obtained through the modified Reynolds analogy. The tip-bluntness data showed a strong delay in the beginning of transition for a cone base-to-tip diameter ratio of 20, approximately a 35% increase in Reynolds number over that of the smooth sharp-tip case. The angle-of-attack data indicated the cross flow to have a strong influence on transition by promoting it on the sheltered side of the cone and delaying it on the windward side.

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.

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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