Development of Higher Temperature Abradable Seals for Gas Turbine Applications

2004 ◽  
Author(s):  
Raymond E. Chupp ◽  
Yuk-Chiu Lau ◽  
Farshad Ghasripoor ◽  
Donald J. Baldwin ◽  
Chek Ng ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Field Tests ◽  
Rotating Blade ◽  
New Material ◽  
Blade Tip ◽  
Higher Temperature ◽  
Sprayed Coating ◽  
Thermally Sprayed ◽  
Unit Performance ◽  
Blade Tips

Improving sealing between rotating and stationary parts in industrial turbines can significantly increase unit performance. Abradable seals are being developed to reduce blade-tip clearances where an abradable material is placed on the stationary shroud or casing opposite the rotating blade tips to reduce clearances with minimum risk to the turbine components during rubs. A newly developed metallic abradable material is a thermally sprayed coating on the first stage of gas turbine shrouds to reduce tip clearances; thereby, improving turbine power and efficiency by about 1/2 percent. This material is an improvement over currently available metallic abradables increasing the maximum operating temperature from 750°C to about 950°C with no special treatments required on the rotating blade tips. Initial engine field tests have shown the new material to be abradable and on target to meet life requirements.

Numerical Simulation of Film Cooling on Rotating Blade Tips Within a High-Pressure Turbine

2013 ◽  
Author(s):  
K. Lu ◽  
M. T. Schobeiri ◽  
J. C. Han
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Flow Structures ◽  
Research Facility ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Rotating Blade ◽  
Blade Tip ◽  
Blade Tips

This paper numerically investigates the aerodynamics and film cooling effectiveness of high pressure turbine blade tips. Two different rotor blade tip configurations have been studied: the plane tip with tip hole cooling and the squealer tip with tip hole cooling. The geometry of the blades is determined based on the blade profiles within the three-stage multi-purpose turbine research facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL), Texas A&M University. Seven perpendicular holes along the camber line are used for the tip hole cooling. The clearance between the blade tip and casing is 1.0% of the blade span. For each blade tip configuration, the coolant is ejected through the cooling holes under blowing ratios of M = 0.5, 1.0 and 1.5. In this paper, a comparison between the plane tip and the squealer tip has been presented. The detailed flow structures and film cooling effectiveness are discussed.

Determination of Crack Initiation on L-1 LP Steam Turbine Blades: Part 1—Measurements on Rotor Train, Material Specimens and Blades

2011 ◽  
Author(s):  
Zdenek Kubin ◽  
Vaclav Cerny ◽  
Pavel Panek ◽  
Tomas Misek ◽  
Jan Hlous ◽  
...  
Keyword(s):  
Steam Turbine ◽  
Turbine Blades ◽  
Rotating Blade ◽  
New Material ◽  
Blade Tip ◽  
Manufacturing Accuracy ◽  
Torsional Excitation ◽  
Run Up ◽  
One Year

After 10 years of operation of a steam turbine with large output power there was an accident during the turbine run-up. One of rotating blade fell off. All 6 LP rotors (two machines) were checked and many cracks on the L-1 blades were found. Due to economic reasons, blades with an identical geometry were manufactured quickly and a new material was used. A better material was chosen in terms of yield limit. The egalization of rotor grooves was performed because of manufacturing accuracy. Tip-timing measurement was installed on two L-1 stages to monitor and protect the blades. After one year of smooth operation new inspections were made. Surprisingly, it was found that the blades made of the new material had comparatively more cracks than the original blades. A new investigation has been started. This article describes measurements including rotor torsional excitation, blade tip-timing measurements, modal analysis and material tests. A computational analysis is presented in Part 2. Application of both approaches revealed what hypotheses should be rejected and, on the other hand, which of them should be analyzed in a deeper way. Consequently, the unstalled flutter has been identified as the most probable cause of blade cracks.

scholarly journals A Study of the Prediction of Tip and Seal Clearances and Their Effects in Gas Turbine Transients

1984 ◽  
Author(s):  
P. Pilidis ◽  
N. R. L. Maccallum
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Computing Time ◽  
Tip Clearance ◽  
Blade Tip ◽  
Air Flows ◽  
Blade Tip Clearance ◽  
Blade Tips

Clearances of compressor and turbine blade tips and seals alter during and following speed transients. These changes affect the performance of the components and hence of the engine. This paper describes models for the prediction of blade tip clearance changes and seal clearance changes. These models have been applied to the H.P. Compressor and H.P. Turbine and to two seals controlling air flows in a Two-spool Bypass Engine. The predicted acceleration rates appear to be more influenced by the changes in the seal clearances than by the tip clearance changes. The increases in computing time in the engine transient program which result from inclusion of the model are acceptable.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

A Numerical Analysis of Gas Turbine Disks Incorporating Rotating Heat Pipes

2000 ◽  
Author(s):  
Y. Cao ◽  
J. Ling ◽  
R. Rivir ◽  
C. MacArthur
Keyword(s):  
Heat Pipe ◽  
Gas Turbine ◽  
Heat Pipes ◽  
High Heat ◽  
Thermal Dissipation ◽  
High Heat Flux ◽  
Heating And Cooling ◽  
Turbine Disks ◽  
Higher Temperature ◽  
Lower Temperature

Abstract Radially rotating heat pipes have been proposed for cooling gas turbine disks working at high temperatures. A disk incorporating the heat pipe would have an enhanced thermal dissipation capacity and a much lower temperature at the disk rim and dovetail surface. In this paper, extensive numerical simulations have been made for heat-pipe-cooled disks. Thermal performances are compared for the disks with and without incorporating the heat pipe at different heating and cooling conditions. The numerical results presented in this paper indicate that radially rotating heat pipes can significantly reduce the maximum and average temperatures at the disk rim and dovetail surface under a high heat flux working condition. In general, the maximum and average temperatures at the disk rim and dovetail surface could be reduced by above 250 and 150 degrees, respectively, compared to those of the disk without the heat pipe. As a result, a disk incorporating radially rotating heat pipes could alleviate temperature-related problems and allow a gas turbine to work at a much higher temperature.

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

scholarly journals Heat Transfer and Flow on the Squealer Tip of a Gas Turbine Blade

2000 ◽  
Author(s):  
Gm S. Azad ◽  
Je-Chin Han ◽  
Robert J. Boyle
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Rotor Blade ◽  
Static Pressure ◽  
Cavity Surface ◽  
Edge Region ◽  
Blade Tip ◽  
The Heat Transfer Coefficient

Experimental investigations are performed to measure the detailed heat transfer coefficient and static pressure distributions on the squealer tip of a gas turbine blade in a five-bladed stationary linear cascade. The blade is a 2-dimensional model of a modern first stage gas turbine rotor blade with a blade tip profile of a GE-E3 aircraft gas turbine engine rotor blade. A squealer (recessed) tip with a 3.77% recess is considered here. The data on the squealer tip are also compared with a flat tip case. All measurements are made at three different tip gap clearances of about 1%, 1.5%, and 2.5% of the blade span. Two different turbulence intensities of 6.1% and 9.7% at the cascade inlet are also considered for heat transfer measurements. Static pressure measurements are made in the mid-span and near-tip regions, as well as on the shroud surface opposite to the blade tip surface. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and an exit Reynolds number based on the axial chord of 1.1×106. A transient liquid crystal technique is used to measure the heat transfer coefficients. Results show that the heat transfer coefficient on the cavity surface and rim increases with an increase in tip clearance. The heat transfer coefficient on the rim is higher than the cavity surface. The cavity surface has a higher heat transfer coefficient near the leading edge region than the trailing edge region. The heat transfer coefficient on the pressure side rim and trailing edge region is higher at a higher turbulence intensity level of 9.7% over 6.1% case. However, no significant difference in local heat transfer coefficient is observed inside the cavity and the suction side rim for the two turbulence intensities. The squealer tip blade provides a lower overall heat transfer coefficient when compared to the flat tip blade.

Back Flow in Turbines of Nuclear Closed-Cycle Gas Turbine Plants in Case of Circuit Pipe Rupture

1975 ◽  
Vol 97 (2) ◽  
pp. 189-194 ◽  
Author(s):  
K. Bammert ◽  
P. Zehner
Keyword(s):  
Gas Turbine ◽  
Rotor Blade ◽  
Back Flow ◽  
Blade Tip ◽  
Stator Blade ◽  
Flow Through ◽  
High Temperature Reactor ◽  
Turbine Cascades ◽  
Characteristic Features ◽  
Safety Considerations

For operation of a gas turbine in single-cycle arrangement with a high-temperature reactor, rupture of a main circuit pipe has to be included in the safety considerations. In the event of such an accident there may be a back flow through the turbo machines or a forward flow up to the choking limit. This paper is a report on tests carried out in a two-dimensional cascade wind tunnel on turbine cascades under back flow conditions. By the example of three selected representative cascades the characteristic features in turbine cascades with back flow are discussed. These cascades are a rotor blade tip section with aerofoil-like profiles and a wide pitch, a stator blade or rotor blade mean section with an usual deflection and a rotor blade root section with a narrow pitch and a large deflection.

scholarly journals Heat Transfer and Pressure Distributions on a Gas Turbine Blade Tip

2000 ◽  
Vol 122 (4) ◽  
pp. 717-724 ◽  
Author(s):  
Gm. S. Azad ◽  
Je-Chin Han ◽  
Shuye Teng ◽  
Robert J. Boyle
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Turbulence Intensity ◽  
Rotor Blade ◽  
Static Pressure ◽  
Pressure Distributions ◽  
Blade Tip

Heat transfer coefficient and static pressure distributions are experimentally investigated on a gas turbine blade tip in a five-bladed stationary linear cascade. The blade is a two-dimensional model of a first-stage gas turbine rotor blade with a blade tip profile of a GE-E3 aircraft gas turbine engine rotor blade. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1×106. The middle 3-blade has a variable tip gap clearance. All measurements are made at three different tip gap clearances of about 1, 1.5, and 2.5 percent of the blade span. Heat transfer measurements are also made at two different turbulence intensity levels of 6.1 and 9.7 percent at the cascade inlet. Static pressure measurements are made in the midspan and the near-tip regions as well as on the shroud surface, opposite the blade tip surface. Detailed heat transfer coefficient distributions on the plane tip surface are measured using a transient liquid crystal technique. Results show various regions of high and low heat transfer coefficient on the tip surface. Tip clearance has a significant influence on local tip heat transfer coefficient distribution. Heat transfer coefficient also increases about 15–20 percent along the leakage flow path at higher turbulence intensity level of 9.7 over 6.1 percent. [S0889-504X(00)00404-9]

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