Numerical Investigation of the Flow in a Radial Vaned Diffuser Without and With Aspiration

2012 ◽  
Author(s):  
A. Marsan ◽  
I. Trébinjac ◽  
S. Coste ◽  
G. Leroy
Keyword(s):  
Skin Friction ◽  
Leading Edge ◽  
Major Change ◽  
Suction Side ◽  
Tip Clearance ◽  
Operating Point ◽  
Vaned Diffuser ◽  
Leading Role ◽  
Stable Operating ◽  
Blade Tip

An analysis of the flow in a centrifugal compressor vaned diffuser from the nominal operating point of the compressor stage to a point near surge was conducted. Study of performance coefficients, and then of the skin-friction pattern, reveals the growth of a corner stall between the hub wall and the suction side of the vane when moving the operating point towards surge. Considering the location of the skin-friction pattern singular elements, a boundary layer suction technique has been developed and then numerically tested. The hub wall corner stall was controlled, and performances predicted near surge have been significantly modified, as well as the flow structures when reaching the limit of numerical stable operating range: the major change in the topology of the flow occurs now within the impeller, in the splitter leading edge region, and let think about a leading role of the main blade tip clearance vortex in the instabilities release. The surge massflow seems to have been significantly reduced.

scholarly journals Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

Studies on the Outline of Flow Improvement With Speed Ratio in a Counter Rotating Turbine

2019 ◽  
Author(s):  
Rayapati Subbarao ◽  
M. Govardhan
Keyword(s):  
Skin Friction ◽  
Nodal Point ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Speed Ratio ◽  
Flow Topology ◽  
Current Effort ◽  
Pressure Surface ◽  
Flow Modification

Abstract Flow through the Counter Rotating Turbine (CRT) stage is more complex due to the presence of two rotors that rotate in the opposite direction, the spacing between them and the tip clearance provided on rotors. This flow aspect may change, if we change the parameters like speed, spacing and blade angles. Current effort contains simulation studies on the flow topology of CRT through dissimilar speed ratios in the range of 0.85–1.17. CRT components stator and the rotors are modelled. At nozzle inlet, stagnation pressure boundary condition is used. At the turbine stage or rotor 2 outlet, mass flow rate is specified. Skin friction lines are drawn on rotor 1 as well as rotor 2 on all over the blade. Not much variation of skin friction lines is witnessed in rotor 1 on the pressure side with exception to the position of the separation line close to leading edge. On suction side, skin friction lines are more uniform when the speed ratio is greater than 1. Skin friction lines on rotor 2 pressure surface show the presence of re-attachment lines. The position of the nodal point of separation near the hub remained same, but the strength is decreasing with speed ratio. On rotor 2 suction side, near the tip, all along the stream wise direction, line of re-attachment is observed that spreads from leading edge to trailing edge, whose strength is varying with speed ratio. Near the hub as well, line of re-attachment is observed, which is of more intensity in lower speed ratios. For the same region in rotor 1, there is proper reattachment as nodes are observed instead of lines, suggesting that more improved flow is occurring in rotor 1 than rotor 2. Thus, the present paper identifies the flow modification with speed ratio in a counter rotating turbine. Also, effort is made to see the consequence of flow change on the output of CRT.

scholarly journals Study and Control of a Radial Vaned Diffuser Stall

2012 ◽  
Vol 2012 ◽  
pp. 1-12 ◽  
Author(s):  
Aurélien Marsan ◽  
Isabelle Trébinjac ◽  
Sylvain Coste ◽  
Gilles Leroy
Keyword(s):  
Boundary Layer ◽  
Skin Friction ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Point Of View ◽  
Vaned Diffuser ◽  
Operating Range ◽  
Stable Operating ◽  
And Control ◽  
Friction Line

The aim of the present study is to evaluate the efficiency of a boundary layer suction technique in case of a centrifugal compressor stage in order to extend its stable operating range. First, an analysis of the flow pattern within the radial vaned diffuser is presented. It highlights the stall of the diffuser vanes when reaching a low massflow. A boundary layer separation in the hub-suction side corner grows when decreasing the massflow from the nominal operating point to the surge and finally leads to a massive stall. An aspiration strategy is investigated in order to control the stall. The suction slot is put in the vicinity of the saddle that originates the main separating skin-friction line, identified thanks to the analysis of the skin-friction pattern. Several aspiration massflow rates are tested, and two different modelings of the aspiration are evaluated. Finally, an efficient control is reached with a removal of only 0,1% of the global massflow and leads—from a steady-state calculations point of view—to an increase by 40% of the compressor operating range extent.

Turbine Blade Tip Cooling With Blade Rotation: Part II — Shroud Coolant Injection

2015 ◽  
Author(s):  
Onieluan Tamunobere ◽  
Sumanta Acharya
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Complementary Method ◽  
Blowing Ratio ◽  
Coolant Injection ◽  
Blade Tip

In this paper, blade-tip cooling is investigated with coolant injection from the shroud alone and a combination of shroud coolant injection and tip cooling. The blade rotates at a nominal speed of 1200 RPM, and consists of a cut back squealer tip with a tip clearance of 1.7% of the blade span. The blade consists of tip holes and pressure side shaped holes, while the shroud has an array of angled holes and a circumferential slot upstream of the rotor section. Different combinations of the three cooling configurations are utilized to study the effectiveness of shroud cooling as a complementary method of cooling the blade tip. The measurements are done using liquid crystal thermography. Blowing ratios of 0.5, 1.0, 2.0, 3.0 and 4.0 are studied for shroud slot cooling and blowing ratios of 1.0, 2.0, 3.0, 4.0 and 5.0 are studied for shroud hole cooling. For cases with coolant injection from the tip, the blowing ratios used are 1.0, 2.0, 3.0 and 4.0. The results show an increase in film cooling effectiveness with increasing blowing ratio for shroud hole cooling. The increased effectiveness from shroud hole cooling is concentrated mainly in the tip-region below the shroud holes and towards the blade suction side and the suction side squealer rim. Slot cooling injection results in increased effectiveness on the blade tip near the blade leading edge up to a maximum blowing ratio, after which the cooling effectiveness decreases with increasing blowing ratio. The combination of the different cooling methods results in better overall cooling coverage of the blade tip with the shroud hole and blade tip cooling combination being the most effective. The level of coolant protection is strongly dependent on the blowing ratio and combination of blowing ratios.

Impact of Tip Clearance Size and Rotation Speed on the Surge Onset in a High Pressure Centrifugal Compressor

2012 ◽  
Author(s):  
Nicolas Buffaz ◽  
Isabelle Trébinjac
Keyword(s):  
Centrifugal Compressor ◽  
Rotation Speed ◽  
Pressure Sensors ◽  
Leading Edge ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Tip Clearance ◽  
Vaned Diffuser ◽  
The Impact ◽  
Blade Leading Edge

The results presented in the paper aim at investigating the impact of tip clearance size and rotation speed on the surge onset in a transonic single-stage centrifugal compressor composed of a backswept splittered unshrouded impeller and a vaned diffuser. For that purpose, various slow throttle ramps into surge were conducted from 100% to 60% design speed of the compressor and two different tip clearance heights were investigated. The 1MW LMFA-ECL test rig was used to carry out the tests in the compressor stage. Unsteady pressure measurements up to 150 KHz were carried out in the inducer (i.e. the entry zone of the impeller between the main blade leading edge and the splitter blade leading edge) and in the diffuser thanks to nine and fifteen static pressure sensors respectively. At cruise rotation speed (92.7% of the nominal rotation speed), the surge is triggered by a boundary layer separation on the diffuser vane suction side whatever the tip clearance height may be. No precursor of surge or pre-surge activity has been recorded in the diffuser or in the impeller. The surge reveals a spike-type inception and the tip clearance increase does not change the path into instability. At lower rotation speeds high frequency disturbances (nearly half the BPF) have been recorded in the inducer before surge. These disturbances can be understood as “tip clearance rotating disturbances” because they are generated at the leading edge of the main blades and move along the tip clearance trajectory. These disturbances reveal a very unstable behavior while the compressor runs into a stable operating point even if the flow at the tip of impeller is dramatically affected by these disturbances. But these disturbances do not trigger the surge which always originates in the diffuser.

Turbine Blade Tip Cooling With Blade Rotation—Part II: Shroud Coolant Injection

2016 ◽  
Vol 138 (9) ◽  
Author(s):  
Onieluan Tamunobere ◽  
Sumanta Acharya
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Complementary Method ◽  
Blowing Ratio ◽  
Coolant Injection ◽  
Blade Tip

In this paper, blade tip cooling is investigated with coolant injection from the shroud alone and a combination of shroud coolant injection and tip cooling. With a nominal rotation speed of 1200 rpm, each blade consists of a cut back squealer tip with a tip clearance of 1.7% of the blade span. The blades also consist of tip holes and pressure side (PS) shaped holes, while the shroud has an array of angled holes and a circumferential slot upstream of the rotor section. Different combinations of the three cooling configurations (tip and PS holes, shroud angled holes, and shroud circumferential slot) are utilized to study the effectiveness of coolant injected from the shroud as a complementary method of cooling the blade tip. The measurements are done using liquid crystal thermography. Blowing ratios of 0.5, 1.0, 2.0, 3.0, and 4.0 are studied for shroud slot cooling, and blowing ratios of 1.0, 2.0, 3.0, 4.0, and 5.0 are studied for shroud hole cooling. For cases with coolant injection from the blade tip, the blowing ratios used are 1.0, 2.0, 3.0, and 4.0. The results show an increase in film cooling effectiveness with increasing blowing ratio for shroud hole coolant injection. The increased effectiveness from shroud hole coolant is concentrated mainly in the tip region below the shroud holes and toward the blade suction side and the suction side squealer rim. Slot coolant injection results in increased effectiveness on the blade tip near the blade leading edge up to a maximum blowing ratio, after which the cooling effectiveness decreases with increasing blowing ratio. The combination of the different cooling methods results in better overall cooling coverage of the blade tip with the shroud hole and blade tip coolant combination being the most effective. The level of coolant protection is strongly dependent on the blowing ratio and combination of blowing ratios.

scholarly journals Investigation of Separation Phenomena in a Radial Pump at Reduced Flow Rate by Large-Eddy Simulation

2016 ◽  
Vol 138 (12) ◽  
Author(s):  
Antonio Posa ◽  
Antonio Lippolis ◽  
Elias Balaras
Keyword(s):  
Reverse Flow ◽  
Leading Edge ◽  
Suction Side ◽  
Secondary Flows ◽  
Vaned Diffuser ◽  
Flow Rates ◽  
Turbulent Statistics ◽  
Large Eddy ◽  
Radial Pump ◽  
Reduced Flow

Turbopumps operating at reduced flow rates experience significant separation and backflow phenomena. Although Reynolds-Averaged Navier–Stokes (RANS) approaches proved to be usually able to capture the main flow features at design working conditions, previous numerical studies in the literature verified that eddy-resolving techniques are required in order to simulate the strong secondary flows generated at reduced loads. Here, highly resolved large-eddy simulations (LES) of a radial pump with a vaned diffuser are reported. The results are compared to particle image velocimetry (PIV) experiments in the literature. The main focus of the present work is to investigate the separation and backflow phenomena occurring at reduced flow rates. Our results indicate that the effect of these phenomena extends up to the impeller inflow: they involve the outer radii of the impeller vanes, influencing significantly the turbulent statistics of the flow. Also in the diffuser vanes, a strong spanwise evolution of the flow has been observed at the reduced load, with reverse flow, located mainly on the shroud side and on the suction side (SS) of the stationary channels, especially near the leading edge of the diffuser blades.

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.

scholarly journals Influence of Large Relative Tip Clearances for a Micro Radial Fan and Design Guidelines for Increased Efficiency

2020 ◽  
Author(s):  
Patrick H. Wagner ◽  
Jan Van herle ◽  
Lili Gu ◽  
Jürg Schiffmann
Keyword(s):  
Pressure Rise ◽  
Leading Edge ◽  
High Sensitivity ◽  
Design Guidelines ◽  
Tip Clearance ◽  
Small Scale ◽  
Blade Tip ◽  
Dynamics Simulations ◽  
Experimental Findings ◽  
Blade Tip Clearance

Abstract The blade tip clearance loss was studied experimentally and numerically for a micro radial fan with a tip diameter of 19.2mm. Its relative blade tip clearance, i.e., the clearance divided by the blade height of 1.82 mm, was adjusted with different shims. The fan characteristics were experimentally determined for an operation at the nominal rotational speed of 168 krpm with hot air (200 °C). The total-to-total pressure rise and efficiency increased from 49 mbar to 68 mbar and from 53% to 64%, respectively, by reducing the relative tip clearance from 7.7% to the design value of 2.2%. Single and full passage computational fluid dynamics simulations correlate well with these experimental findings. The widely-used Pfleiderer loss correlation with an empirical coefficient of 2.8 fits the numerical simulation and the experiments within +2 efficiency points. The high sensitivity to the tip clearance loss is a result of the design specific speed of 0.80, the highly-backward curved blades (17°), and possibly the low Reynolds number (1 × 105). The authors suggest three main measures to mitigate the blade tip clearance losses for small-scale fans: (1) utilization of high-precision surfaced-grooved gas-bearings to lower the blade tip clearance, (2) a mid-loaded blade design, and (3) an unloaded fan leading edge to reduce the blade tip clearance vortex in the fan passage.

Blade Tip Heat Transfer and Aerodynamics in a Large Scale Turbine Cascade With Moving Endwall

2006 ◽  
Author(s):  
P. Palafox ◽  
M. L. G. Oldfield ◽  
P. T. Ireland ◽  
T. V. Jones ◽  
J. E. LaGraff
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Large Scale ◽  
Infrared Camera ◽  
Suction Side ◽  
Constant Heat Flux ◽  
Tip Clearance ◽  
Airfoil Surface ◽  
Nusselt Number Distribution ◽  
Blade Tip

High resolution Nusselt number (Nu) distributions were measured on the blade tip surface of a large, 1.0 meter-chord, low-speed cascade representative of a high-pressure turbine. Data was obtained at a Reynolds number of 4.0 × 105 based on exit velocity and blade axial chord. Tip clearance levels ranged from 0.56% to 1.68% design span or equally from 1% to 3% of blade chord. An infrared camera, looking through the hollow blade, made detailed temperature measurements on a constant heat flux tip surface. The relative motion between the endwall and the blade tip was simulated by a moving belt. The moving belt endwall significantly to shifts the region of high Nusselt number distribution and reduces the overall averaged Nusselt number on the tip surface by up to 13.3%. The addition of a suction side squealer tip significantly reduced local tip heat transfer and resulted in a 32% reduction in averaged Nusselt number. Analysis of pressure measurements on the blade airfoil surface and tip surface along with PIV velocity flow fields in the gap give an understanding of the heat transfer mechanism.

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