scholarly journals Uncertainty Quantification of Non-Dimensional Parameters for a Film Cooling Configuration in Supersonic Conditions

Fluids ◽  
2019 ◽  
Vol 4 (3) ◽  
pp. 155 ◽  
Author(s):  
Simone Salvadori ◽  
Mauro Carnevale ◽  
Alessia Fanciulli ◽  
Francesco Montomoli
Keyword(s):  
High Pressure ◽  
Uncertainty Quantification ◽  
Hermite Polynomials ◽  
Suction Side ◽  
System Response ◽  
General Description ◽  
Trailing Edge ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Shock Impingement

In transonic high-pressure turbine stages, oblique shocks originating from vane trailing edges impact the suction side of each adjacent vane. High-pressure vanes are cooled to tolerate the combustor exit-temperature levels, then it is highly probable that shock impingement will occur in proximity to a row of cooling holes. The presence of such a shock, together with the inevitable manufacturing deviations, alters the location of the shock impingement and of the performance parameters of each cooling hole. The present work provides a general description of the aero-thermal field that occurs on the rear suction side of a cooled vane. Computational Fluid Dynamics (CFD) is used to evaluate the deterministic response of the selected configurations in terms of adiabatic effectiveness, discharge coefficient, blowing ratio, density ratio, and momentum ratio. Turbulence is modelled by using both the Shear Stress Transport method (SST) and the Reynolds Stress Model (RSM) implemented in ANSYS ® FLUENT ® . The obtained results are compared with the experimental data obtained by the Institut für Thermische Strömungsmaschinen in Karlsruhe. Two uncertainty quantification methodologies based on Hermite polynomials and Padè–Legendre approximants are used to consider the probability distribution of the geometrical parameters and to evaluate the response surfaces for the system response quantities. Trailing-edge and cooling-hole diameters have been considered to be aleatory unknowns. Uncertainty quantification analysis allows for the assessment of the mutual effects on global and local parameters of the cooling device. Obtained results demonstrate that most of the parameters are independent by the variation of the aleatory unknowns while the standard deviation of the blowing ratio associated with the hole diameter uncertainty is around 12 % , with no impact by the trailing-edge thickness. No relevant advantages are found using either SST model or RSM in combination with Hermite polynomials and Padè–Legendre approximants.

Modeling and Impact of High-pressure Turbine Blade Trailing Edge Film Cooling Hole Variations

2020 ◽  
Author(s):  
Jan Kamenik ◽  
David J. Toal ◽  
Andy Keane ◽  
Lars Högner ◽  
Marcus Meyer ◽  
...  
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Film Cooling ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Hole

Numerical Investigations of Wake and Shock Wave Effects on Film Cooling Performance in a Transonic Turbine Stage: Part 1 — Methodology Development and Qualification Over Stationary Stators and Rotors

2013 ◽  
Author(s):  
Huazhao Xu ◽  
Jianhua Wang ◽  
Ting Wang
Keyword(s):  
Experimental Data ◽  
Mach Number ◽  
Film Cooling ◽  
Suction Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Numerical Investigations ◽  
Blowing Ratio

To understand the unsteady shock wave and wake effects on the film cooling performance over a transonic 3-D rotating stage, a series of numerical investigations have been conducted and are presented in this two-part paper. Part 1 is focused on the development of the computational model and methodology of the system setup and model qualification; Part 2 is to investigate the unsteady effects of shock waves and wakes on film cooling performance in a transonic rotating stage. In Part 1, the film cooling experimental conditions (non-rotating) and test sections of Kopper et. al. and Hunter are selected for model qualification. The numerical computation is carried out by the commercial software Ansys/Fluent using the pressure based compressible flow governing equations. The effects of four turbulence models are carefully compared with the experimental data. The Realizable k-ε turbulence model is found to match the experimental data better than the other models and is thus used for the rest of the study, including Part 2. The results show that 1) the weak shock emanating from the neighboring stator’s trailing edge results in a temperature rise and a reduction of film cooling effectiveness on the suction side near the trailing edge, 2) cooling ejection from the trailing edge reduces the shock strength in the stator passage, 3) an increase in Mach number from 0.84 to 1.50 can reduce the total pressure losses of fluid flow near the end-walls, 4) the film cooling effectiveness increases with increasing blowing ratio and becomes more even on the stator with a higher blowing ratio, and 5) an increase in Mach number from 0.84 to 1.50 gives rise to a higher cooling effectiveness in the region from the cooling holes to 80% of the chord length of the stator on the pressure side, but becomes lower after this up to the trailing edge. However, on the stator’s suction side, higher Mach number results in a lower cooling effectiveness region around the film holes from 30% to 55% of the chord length, but cooling effectiveness increases downstream.

Experimental Investigation of Temperature Difference Ratio of Trailing Edge Film Cooling Downstream of Pressure Side Split with Internal Spacer Arrays

2014 ◽  
Vol 1070-1072 ◽  
pp. 1743-1747
Author(s):  
Chang Ming Ling ◽  
Chun Hua Min ◽  
Ming Feng Chen ◽  
Jun Li ◽  
Dong Mei Zhao
Keyword(s):  
Reynolds Number ◽  
Temperature Difference ◽  
Film Cooling ◽  
Relative Length ◽  
Geometric Parameters ◽  
Pressure Side ◽  
Trailing Edge ◽  
Blowing Ratio ◽  
Pin Fins ◽  
Cooling Hole

This paper investigates experimentally the effects of spacer geometry on temperature difference ratio downstream of pressure side split of trailing edge film cooling. The internal pin fins arrayed staggered and pin fins in the last row face to the cooling hole one by one. The geometric parameters include relative length of trailing edge l/b, ratio of length to span of split a/c, ratio of length to width a/b, angle of split α. And the effects of blowing ratio and Reynolds number are also researched specially. The correlation of temperature difference ratio vs. Reynolds number, blowing ratio, and geometric parameters respectively was induced.

Wake–Wake Interaction and Its Potential for Clocking in a Transonic High-Pressure Turbine

2001 ◽  
Vol 124 (1) ◽  
pp. 69-76 ◽  
Author(s):  
Frank Hummel
Keyword(s):  
High Pressure ◽  
Flow Field ◽  
Flow Direction ◽  
Numerical Procedure ◽  
Flow Simulation ◽  
Suction Side ◽  
Frame Of Reference ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Time Resolved

Two-dimensional unsteady Navier–Stokes calculations of a transonic single-stage high-pressure turbine were carried out with emphasis on the flow field behind the rotor. Detailed validation of the numerical procedure with experimental data showed excellent agreement in both time-averaged and time-resolved flow quantities. The numerical timestep as well as the grid resolution allowed the prediction of the Ka´rma´n vortex streets of both stator and rotor. Therefore, the influence of the vorticity shed from the stator on the vortex street of the rotor is detectable. It was found that certain vortices in the rotor wake are enhanced while others are diminished by passing stator wake segments. A schematic of this process is presented. In the relative frame of reference, the rotor is operating in a transonic flow field with shocks at the suction side trailing edge. These shocks interact with both rotor and stator wakes. It was found that a shock modulation occurs in time and space due to the stator wake passing. In the absolute frame of reference behind the rotor, a 50-percent variation in shock strength is observed according to the circumferential or clocking position. Furthermore, a substantial weakening of the rotor suction side trailing edge shock in flow direction is detected in an unsteady flow simulation when compared to a steady-state calculation, which is caused by convection of upstream stator wake segments. The physics of the aforementioned unsteady phenomena as well as their influence on design are discussed.

Effect of a Cutback Squealer and Cavity Depth on Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2008 ◽  
Vol 130 (2) ◽  
Author(s):  
Shantanu Mhetras ◽  
Diganta Narzary ◽  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cavity Depth ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blade Tip ◽  
Film Cooling Holes

Film-cooling effectiveness from shaped holes on the near tip pressure side and cylindrical holes on the squealer cavity floor is investigated. The pressure side squealer rim wall is cut near the trailing edge to allow the accumulated coolant in the cavity to escape and cool the tip trailing edge. Effects of varying blowing ratios and squealer cavity depth are also examined on film-cooling effectiveness. The film-cooling effectiveness distributions are measured on the blade tip, near tip pressure side and the inner pressure side and suction side rim walls using pressure sensitive paint technique. The internal coolant-supply passages of the squealer tipped blade are modeled similar to those in the GE-E3 rotor blade with two separate serpentine loops supplying coolant to the film-cooling holes. Two rows of cylindrical film-cooling holes are arranged offset to the suction side profile and along the camber line on the tip. Another row of shaped film-cooling holes is arranged along the pressure side just below the tip. The average blowing ratio of the cooling gas is controlled to be 0.5, 1.0, 1.5, and 2.0. A five-bladed linear cascade in a blow down facility with a tip gap clearance of 1.5% is used to perform the experiments. The free-stream Reynolds number, based on the axial chord length and the exit velocity, was 1,480,000 and the inlet and exit Mach numbers were 0.23 and 0.65, respectively. A blowing ratio of 1.0 is found to give best results on the pressure side, whereas the tip surfaces forming the squealer cavity give best results for M=2. Results show high film-cooling effectiveness magnitudes near the trailing edge of the blade tip due to coolant accumulation from upstream holes in the tip cavity. A squealer depth with a recess of 2.1mm causes the average effectiveness magnitudes to decrease slightly as compared to a squealer depth of 4.2mm.

Wake-Wake Interactions and its Potential for Clocking in a Transonic High Pressure Turbine

2001 ◽  
Author(s):  
Frank Hummel
Keyword(s):  
High Pressure ◽  
Flow Field ◽  
Flow Direction ◽  
Numerical Procedure ◽  
Flow Simulation ◽  
Suction Side ◽  
Frame Of Reference ◽  
High Pressure Turbine ◽  
Time Step ◽  
Trailing Edge

Two-dimensional unsteady Navier-Stokes calculations of a transonic single stage high pressure turbine were carried out with emphasis on the flow field behind the rotor. Detailed validation of the numerical procedure with experimental data showed excellent agreement in both time-averaged and time-resolved flow quantities. The numerical time-step as well as the grid resolution allowed the prediction of the Kármán vortex streets of both stator and rotor. Therefore the influence of the vorticity shed from the stator on the vortex street of the rotor is detectable. It was found that certain vortices in the rotor wake are enhanced while others are diminished by passing stator wake segments. A schematic of this process is presented. In the relative frame of reference the rotor is operating in a transonic flow field with shocks at the suction side trailing edge. These shocks interact with both rotor and stator wakes. It was found that a shock-modulation occurs in time and space due to the stator wake passing. In the absolute frame of reference behind the rotor a 50% variation in shock strength is observed according to the circumferential or clocking position. Furthermore a substantial weakening of the rotor suction side trailing edge shock in flow direction is detected in an unsteady flow simulation when compared to a steady state calculation which is caused by convection of upstream stator wake segments. The physics of the mentioned unsteady phenomena as well as their influence on design are discussed.

Influence of tip clearance and cavity depth on heat transfer in a cutback squealer tip

2020 ◽  
pp. 095441002096469
Author(s):  
Weijie Wang ◽  
Shaopeng Lu ◽  
Hongmei Jiang ◽  
Qiusheng Deng ◽  
Jinfang Teng ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Suction Side ◽  
High Heat ◽  
Tip Clearance ◽  
Tip Leakage Flow ◽  
Trailing Edge ◽  
Cavity Depth ◽  
High Heat Transfer ◽  
Blade Tip ◽  
High Heat Transfer Coefficient

Numerical simulations are conducted to present the aerothermal performance of a turbine blade tip with cutback squealer rim. Two different tip clearance heights (0.5%, 1.0% of the blade span) and three different cavity depths (2.0%, 3.0%, and 6.0% of the blade span) are investigated. The results show that a high heat transfer coefficient (HTC) strip on the cavity floor appears near the suction side. It extends with the increase of tip clearance height and moves towards the suction side with the increase of cavity depth. The cutback region near the trailing edge has a high HTC value due to the flush of over-tip leakage flow. High HTC region shrinks to the trailing edge with the increase of cavity depth since there is more accumulated flow in the cavity for larger cavity depth. For small tip clearance cases, high HTC distribution appears on the pressure side rim. However, high HTC distribution is observed on suction side rim for large tip clearance height. This is mainly caused by the flow separation and reattachment on the squealer rims.

The Effect of Pulsed Blowing on the Boundary Layer of a Highly Loaded Low Pressure Turbine Blade

2013 ◽  
Author(s):  
Marion Mack ◽  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Highly Loaded

The performance of the low pressure turbine (LPT) can vary appreciably, because this component operates under a wide range of Reynolds numbers. At higher Reynolds numbers, mid and aft loaded profiles have the advantage that transition of suction side boundary layer happens further downstream than at front loaded profiles, resulting in lower profile loss. At lower Reynolds numbers, aft loading of the blade can mean that if a suction side separation exists, it may remain open up to the trailing edge. This is especially the case when blade lift is increased via increased pitch to chord ratio. There is a trend in research towards exploring the effect of coupling boundary layer control with highly loaded turbine blades, in order to maximize performance over the full relevant Reynolds number range. In an earlier work, pulsed blowing with fluidic oscillators was shown to be effective in reducing the extent of the separated flow region and to significantly decrease the profile losses caused by separation over a wide range of Reynolds numbers. These experiments were carried out in the High-Speed Cascade Wind Tunnel of the German Federal Armed Forces University Munich, Germany, which allows to capture the effects of pulsed blowing at engine relevant conditions. The assumed control mechanism was the triggering of boundary layer transition by excitation of the Tollmien-Schlichting waves. The current work aims to gain further insight into the effects of pulsed blowing. It investigates the effect of a highly efficient configuration of pulsed blowing at a frequency of 9.5 kHz on the boundary layer at a Reynolds number of 70000 and exit Mach number of 0.6. The boundary layer profiles were measured at five positions between peak Mach number and the trailing edge with hot wire anemometry and pneumatic probes. Experiments were conducted with and without actuation under steady as well as periodically unsteady inflow conditions. The results show the development of the boundary layer and its interaction with incoming wakes. It is shown that pulsed blowing accelerates transition over the separation bubble and drastically reduces the boundary layer thickness.

Conjugate Heat Transfer Analysis of a High Pressure Air-Cooled Gas Turbine Vane

2010 ◽  
Author(s):  
Zhenfeng Wang ◽  
Peigang Yan ◽  
Hongyan Huang ◽  
Wanjin Han
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Boundary Conditions ◽  
Gas Turbine ◽  
Turbulence Model ◽  
Conjugate Heat Transfer ◽  
Suction Side ◽  
Turbulence Characteristic ◽  
Transition Flow ◽  
Characteristic Boundary

The ANSYS-CFX software is used to simulate NASA-Mark II high pressure air-cooled gas turbine. The work condition is Run 5411 which have transition flow characteristics. The different turbulence models are adopted to solve conjugate heat transfer problem of this three-dimensional turbine blade. Comparing to the experimental results, k-ω-SST-γ-θ turbulence model results are more accurate and can simulate accurately the flow and heat transfer characteristics of turbine with transition flow characteristics. But k-ω-SST-γ-θ turbulence model overestimates the turbulence kinetic energy of blade local region and makes the heat transfer coefficient higher. It causes that local region temperature of suction side is higher. In this paper, the compiled code adopts the B-L algebra model and simulates the same computation model. The results show that the results of B-L model are accurate besides it has 4% temperature error in the suction side transition region. In addition, different turbulence characteristic boundary conditions of turbine inner-cooling passages are given and K-ω-SST-γ-θ turbulence model is adopted in order to obtain the effect of turbulence characteristic boundary conditions for the conjugate heat transfer computation results. The results show that the turbulence characteristic boundary conditions of turbine inner-cooling passages have a great effect on the conjugate heat transfer results of high pressure gas turbine. ANSYS is applied to analysis the thermal stress of Mark II blade which has ten radial cooled passages and the results of Von Mises stress show that the temperature gradient results have a great effect on the results of blade thermal stress.

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