Numerical investigation of high-pressure turbine blade tip flows: analysis of aerodynamics

2011 ◽  
Vol 225 (7) ◽  
pp. 940-953 ◽  
Author(s):  
P Vass ◽  
T Arts
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Three Dimensional ◽  
Suction Side ◽  
Point Of View ◽  
Internal Cooling ◽  
High Pressure Turbine ◽  
Blade Tip ◽  
Design Solutions ◽  
Exit Mach Number

The current contribution reports on the validation and analysis of three-dimensional computational results of the flow around four distinct high-pressure turbine blade tip geometries (TG1, 2, 3, and 4 hereinafter), taking into account the effect of the entire internal cooling setup inside the blade, at design exit Mach number: M = 0.8, and high exit Reynolds number: Re C = 900 000. Three of the four geometries represent different tip design solutions – TG1: full squealer rim; TG2: single squealer on the suction side; TG3: partial suction and pressure side squealer, and one (TG4) models TG1 in worn condition. This article provides a comparison between the different geometries from the aerodynamic point of view, analyses the losses, and evaluates the distinct design solutions. An assessment of the effect of the uneven rubbing of the blade tip was performed as well. TG1 was found to be the top performer followed by TG3 and TG2. According to the investigations, the effect of rubbing increased the kinetic loss coefficient by 10–15 per cent.

Multi-Disciplinary Analyses for the Design of a High Pressure Turbine Blade Tip

2016 ◽  
Author(s):  
Stefano Caloni ◽  
Shahrokh Shahpar
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Suction Side ◽  
High Pressure Turbine ◽  
Cooling Flow ◽  
Barrier Coating ◽  
Flow Features ◽  
Internally Cooled ◽  
Blade Tip

The design of a high pressure turbine blade is a challenging task requiring multiple disciplines to be solved simultaneously. Most recently, conjugate analyses are being developed to tackle such a problem; they are able to resolve both the fluid dynamics in a turbine passage and the thermal distribution in the solid part of the component. In this paper, the in-house Hydra CFD solver is used to analyse a high pressure shroudless turbine blade for a modern jet engine. The turbine is internally cooled and a Thermal Barrier Coating (TBC) is applied on the aerofoil surface. The coupling technique used at the interface in the presence of the TBC is described. The flow features at the tip of the turbine blade are the main focus of this study. Four different tip configurations are analysed. A flat tip and a squealer tip are chosen as reference designs; however the effects of opening the Trailing Edge (TE) on the Suction Side (SS) and the Pressure Side (PS) are also investigated. Both a cooled and an uncooled configuration of the turbine blade are analysed and the effect of the cooling flow on the over tip leakage is studied. Finally, conjugate analyses for the cooled turbine blades are used to predict the temperature reached by the different tip designs. The design with an opened TE on the SS shows a significant aerodynamic improvement over the others without increasing the temperature the tip has to withstand in operation.

Influence of Different Parametrizations on the Optimum Design of a High Pressure Turbine Blade Firtree

2016 ◽  
Author(s):  
Frank Wagner ◽  
Arnold Kühhorn ◽  
Thomas Weiss ◽  
Dierk Otto
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Complex Geometry ◽  
Three Dimensional ◽  
Parametric Models ◽  
Limiting Factor ◽  
High Pressure Turbine ◽  
The Creation ◽  
High Flexibility ◽  
Critical Regions

Today the design processes in the aero industry face many challenges. Apart from automation itself, a suitable parametric geometry setup plays a significant role in making workflows usable for optimization. At the same time there are tough requirements against the parametric model. For the lowest number of possible parameters, which should be intuitively ascertainable, a high flexibility has to be ensured. Within the parameter range an acceptable stability is necessary. Under these constraints the creation of such parametric models is a challenge, which should not be underestimated especially for a complex geometry. In this work different kinds of parametrization with different levels of complexity will be introduced and compared. Thereby several geometry elements will be used to handle the critical regions of the geometry. In the simplest case a combination of lines and arcs will be applied. These will be replaced by superior elements like a double arc construct or different formulations of b-splines. There will be an additional focus on the variation of spline degree and control points. To guarantee consistency a set of general parameters will be used next to the specific ones at the critical regions. The different parameter boundaries have a influence on the possible geometries and should therefore be tested separately before an optimization run. The analysis of the particular parametrization should be compared against the following points: • effort for the creation of the parametrization in theory • required time for the implementation in the CAD software • error-proneness/robustness of the parametrization • flexibility of the possible geometries • accuracy of the results • influence of the number of runs on the optimization • comparison of the best results Even though this assessment matrix is only valid for the considered case, it should show the general trend for the creation of these kinds of parametric models. This case takes a look at a firtree of a high pressure turbine blade, which is a scaled version of the first row from a small to medium aero engine. The failure of such a component can lead to a critical engine failure. For that reason, the modeling/meshing must be done very carefully and the contact between the blade and the disc is of crucial importance. It is possible to use scaling factors for three dimensional effects to reduce the problem to a two dimensional problem. Therefore the contact description is shortened from face-to-line to line-to-point. The main aim of the optimization is the minimization of the tension (notch stress) at the inner bends of the blade respectively at the outer bends of the disc. This has been the limiting factor in previous investigations. At this part of the geometry the biggest improvement are expected from a superior parametrization. Another important constraint in the optimization is the pressure contact (crushing stress) between blade and disc. Additionally the geometry is restricted with measurements of the lowest diameter at specific fillets to fulfill manufacturing requirements.

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.

Using CFD to Reduce Resonant Stresses on a Single-Stage, High-Pressure Turbine Blade

2002 ◽  
Author(s):  
J. P. Clark ◽  
A. S. Aggarwala ◽  
M. A. Velonis ◽  
R. E. Gacek ◽  
S. S. Magge ◽  
...  
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Guide Vane ◽  
Turbine Blades ◽  
Suction Side ◽  
Lessons Learned ◽  
Single Stage ◽  
Navier Stokes ◽  
High Pressure Turbine ◽  
Time Resolved

The ability to predict levels of unsteady forcing on high-pressure turbine blades is critical to avoid high-cycle fatigue failures. In this study, 3D time-resolved computational fluid dynamics is used within the design cycle to predict accurately the levels of unsteady forcing on a single-stage high-pressure turbine blade. Further, nozzle-guide-vane geometry changes including asymmetric circumferential spacing and suction-side modification are considered and rigorously analyzed to reduce levels of unsteady blade forcing. The latter is ultimately implemented in a development engine, and it is shown successfully to reduce resonant stresses on the blade. This investigation builds upon data that was recently obtained in a full-scale, transonic turbine rig to validate a Reynolds-Averaged Navier-Stokes (RANS) flow solver for the prediction of both the magnitude and phase of unsteady forcing in a single-stage HPT and the lessons learned in that study.

scholarly journals Erratum to: Dynamic probabilistic design approach of high-pressure turbine blade-tip radial running clearance

2016 ◽  
Vol 86 (1) ◽  
pp. 225-225
Author(s):  
Cheng-Wei Fei ◽  
Yat-Sze Choy ◽  
Dian-Yin Hu ◽  
Guang-Chen Bai ◽  
Wen-Zhong Tang
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Design Approach ◽  
Probabilistic Design ◽  
High Pressure Turbine ◽  
Blade Tip

Time-Averaged and Time-Accurate Aero-Dynamics for the Recessed Tip Cavity of a High-Pressure Turbine Blade and the Outer Stationary Shroud: Comparison of Computational and Experimental Results

2004 ◽  
Author(s):  
Brian R. Green ◽  
John W. Barter ◽  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Computational Models ◽  
Turbine Rotor ◽  
Single Stage ◽  
Cfd Modeling ◽  
Blade Surface ◽  
High Pressure Turbine ◽  
Pressure Transducers ◽  
Blade Tip

The unsteady aero-dynamics of a single-stage high-pressure turbine blade operating at design corrected conditions has been the subject of a thorough study involving detailed measurements and computations. The experimental configuration consisted of a single-stage high-pressure turbine and the adjacent, downstream, low-pressure turbine nozzle row. All three blade-rows were instrumented at three spanwise locations with flush-mounted, high frequency response pressure transducers. The rotor was also instrumented with the same transducers on the blade tip and platform and the stationary shroud was instrumented with pressure transducers at specific locations above the rotating blade. Predictions of the time-dependent flow field around the rotor were obtained using MSU-TURBO, a 3D, non-linear, computational fluid dynamics (CFD) code. Using an isolated blade-row unsteady analysis method, the unsteady surface pressure for the high-pressure turbine rotor due to the upstream high-pressure turbine nozzle was calculated. The predicted unsteady pressure on the rotor surface was compared to the measurements at selected spanwise locations on the blade, in the recessed cavity, and on the shroud. The rig and computational models included a flat and recessed blade tip geometry and were used for the comparisons presented in the paper. Comparisons of the measured and predicted static pressure loading on the blade surface show excellent correlation from both a time-average and time-accurate standpoint. This paper concentrates on the tip and shroud comparisons between the experiments and the predictions and these results also show good correlation with the time-resolved data. These data comparisons provide confidence in the CFD modeling and its ability to capture unsteady flow physics on the blade surface, in the flat and recessed tip regions of the blade, and on the stationary shroud.

An Investigation on Aerodynamics Loss Mechanism of Squealer Tips of a High Pressure Turbine Blade Using URANS

2016 ◽  
Author(s):  
Jin-sol Jung ◽  
Okey Kwon ◽  
Changmin Son
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Wall Thickness ◽  
Suction Side ◽  
Leakage Flow ◽  
Pressure Side ◽  
Tip Leakage Flow ◽  
High Pressure Turbine ◽  
Loss Mechanism ◽  
Tip Leakage

The flow leaking over the tip of a high pressure turbine blade generates significant aerodynamic losses as it mixes with the mainstream flow. This study investigates the effect of blade tip geometries on turbine performance with both steady RANS and unsteady URANS analyses. Five different squealer geometries for a high pressure turbine blade have been examined: squealer on pressure side, squealer on suction side, cavity squealer, cavity squealer with pressure side cutback, and cavity squealer with suction side cutback. With the case of the cavity squealer, three different squealer wall thickness are investigated for the wall thickness (w) of 1x, 2x and 4x of the tip gap (G). The unsteady flow analyses using CFX have been conducted to investigate unsteady characteristics of the tip leakage flow and its influence on turbine performances. Through the comparison between URANS analyses, detailed vortex and wake structures are identified and studied at different fidelities. It is found that the over tip leakage flow loss is affected by the tip suction side geometry rather than that of the pressure side geometry. The unsteady results have contributed to resolve the fundamentals of vortex structures and aerodynamic loss mechanisms in a high pressure turbine stage.

Numerical Investigation of Rotation Effects on the Flow and Heat Transfer on the Turbine Blade Tip Underside With Bleed Hole at Different Locations

2019 ◽  
Author(s):  
Zhiqi Zhao ◽  
Lei Luo ◽  
Xiaoxu Kan ◽  
Dandan Qiu ◽  
Xun Zhou
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Structural Damage ◽  
Three Dimensional ◽  
Side Wall ◽  
Internal Cooling ◽  
Heat Transfer Augmentation ◽  
Smooth Channel ◽  
Blade Tip ◽  
Cooling Passage

Abstract High thermal load on the turbine blade tip surface leads to high temperature corrosion and severe structural damage. One common way is to deliver a part of coolant through bleed holes onto the tip portion for cooling purpose. In this study, numerical simulations are performed to investigate the effects of rotation on the internal tip heat transfer in a simplified rotating two-pass channel with a bleed hole, which is applicable to the internal cooling passage of typical gas turbine blade. The bleed hole is placed on the tip wall of a two-pass channel at different locations, i.e. the ratio of distance from the outlet-side wall to width of the tip wall is 0.07, 0.21, 0.5, 0.78, 0.93, respectively. A smooth channel without bleed hole is used as Baseline. The Reynolds number is fixed at 10,000. The Ro numbers are varied from 0 to 0.4. Results show that a three-dimensional vortex, which is induced by the Coriolis force, is found at the bend region. It transports the fluid from the trailing side to leading side, which is beneficial to enhance tip heat transfer. The middle-mounted hole shows a better heat transfer augmentation compared to other hole arrangement. The rotation have a notable effect on the heat transfer and flow structures. Compared to the smooth channel, the heat transfer augmentation is about 34%.

Averaged and Time-Dependent Aerodynamics of a High Pressure Turbine Blade Tip Cavity and Stationary Shroud: Comparison of Computational and Experimental Results

2004 ◽  
Vol 127 (4) ◽  
pp. 736-746 ◽  
Author(s):  
Brian R. Green ◽  
John W. Barter ◽  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Computational Models ◽  
Single Stage ◽  
Time Dependent ◽  
Cfd Modeling ◽  
Blade Surface ◽  
High Pressure Turbine ◽  
Pressure Transducers ◽  
Blade Tip

The unsteady aero-dynamics of a single-stage high-pressure turbine blade operating at design corrected conditions has been the subject of a thorough study involving detailed measurements and computations. The experimental configuration consisted of a single-stage high-pressure turbine and the adjacent, downstream, low-pressure turbine nozzle row. All three blade-rows were instrumented at three spanwise locations with flush-mounted, high-frequency response pressure transducers. The rotor was also instrumented with the same transducers on the blade tip and platform and the stationary shroud was instrumented with pressure transducers at specific locations above the rotating blade. Predictions of the time-dependent flow field around the rotor were obtained using MSU-TURBO, a three-dimensional (3D), nonlinear, computational fluid dynamics (CFD) code. Using an isolated blade-row unsteady analysis method, the unsteady surface pressure for the high-pressure turbine rotor due to the upstream high-pressure turbine nozzle was calculated. The predicted unsteady pressure on the rotor surface was compared to the measurements at selected spanwise locations on the blade, in the recessed cavity, and on the shroud. The rig and computational models included a flat and recessed blade tip geometry and were used for the comparisons presented in the paper. Comparisons of the measured and predicted static pressure loading on the blade surface show excellent correlation from both a time-average and time-accurate standpoint. This paper concentrates on the tip and shroud comparisons between the experiments and the predictions and these results also show good correlation with the time-resolved data. These data comparisons provide confidence in the CFD modeling and its ability to capture unsteady flow physics on the blade surface, in the flat and recessed tip regions of the blade, and on the stationary shroud.

Evolution of Surface Deposits on a High Pressure Turbine Blade: Part I — Physical Characteristics

2006 ◽  
Author(s):  
James E. Wammack ◽  
Jared Crosby ◽  
Daniel Fletcher ◽  
Jeffrey P. Bons ◽  
Thomas H. Fletcher
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Three Dimensional ◽  
Surface Treatments ◽  
Surface Topology ◽  
Laboratory Simulation ◽  
Thermal Barrier ◽  
High Pressure Turbine ◽  
Average Roughness ◽  
Barrier Coating

Turbine blade coupons with three different surface treatments were exposed to deposition conditions in an accelerated deposition facility. The facility simulates the flow conditions at the inlet to a first stage high pressure turbine (T = 1150°C, M = 0.31). The combustor exit flow is seeded with dust particulate that would typically be ingested by a large utility power plant. The three coupon surface treatments included: (1) bare polished metal, (2) polished thermal barrier coating with bondcoat, and (3) unpolished oxidation resistant bondcoat. Each coupon was subjected to four successive 2 hour deposition tests. The particulate loading was scaled to simulate 0.02 ppmw (parts per million weight) of particulate over three months of continuous gas turbine operation for each 2 hour laboratory simulation (for a cumulative one year of operation). Three-dimensional maps of the deposit-roughened surfaces were created between each test, representing a total of four measurements evenly spaced through the lifecycle of a turbine blade surface. From these measurements the surface topology and roughness statistics were determined. Despite the different surface treatments, all three surfaces exhibited similar non-monotonic changes in roughness with repeated exposure. In each case, an initial build-up of isolated roughness peaks was followed by a period when valleys between peaks were filled with subsequent deposition. This trend is well documented using the average forward facing roughness angle in combination with the average roughness height as characteristic roughness metrics. Deposition-related mechanisms leading to spallation of the thermal barrier coated coupons are identified and documented.

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