Analysis of Gas Turbine Rotor Blade Tip and Shroud Heat Transfer (original) (raw)
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Prediction of Unshrouded Rotor Blade Tip Heat Transfer
1995
The rate of heat transfer on the tip of a turbine rotor blade and on the blade surface in the vicinity of the tip, was successfully predicted. The computations were performed with a multiblock computer code which solves the Reynolds Averaged Navier-Stokes equations using an efficient multigrid method. The case considered for the present calculations was the SSME (Space Shuttle Main Engine) high pressure fuel side turbine. The predictions of the blade tip heat transfer agreed reasonably well with the experimental measurements using the present level of grid refinement. On the tip surface, regions with high rate of heat transfer was found to exist close to the pressure side and suction side edges. Enhancement of the heat transfer was also observed on the blade surface near the tip. Further comparison of the predictions was performed with results obtained from correlations based on fully developed channel flow. NOMENCLATURE A Row area Hydraulic diameter=2C • Gap spacing Cp Constant pressure specific heat Thermal conductivity in Mass rate of flow Nu Nusselt number Pr Prandtl number Local Radius Re Reynolds number Distance along the blade surface measured from the stagnation line St Stanton number • Temperature
NUMERICAL SIMULATION OF GAS TURBINE BLADE COOLING FOR ENHANCEMENT OF HEAT TRANSFER OF THE BLADE TIP
In today's industrial scenario, Gas Turbine is one of the most important components of auxiliary power plant system. In order maximize the overall performance and efficiency of all modern turbines, which theoretically operate according to Brayton cycle, they are operated at a very high temperature. These temperatures are so high that, which may fall in the region of turbine blade material melting point temperatures. Due to such high temperatures there is a possibility that the turbine blades may get damaged due to produced thermal stresses and presents a possible threat to the turbine system as well as the operators. Hence to ensure safe and reliable working of the turbines an effective and reliable cooling system is necessary. Currently available methods for cooling of the turbine blades include film cooling with impingement cooling for the leading edge, rib turbulated cooling using serpentine passages for the middle portion of the blade and pin fin cooling for the trailing edge of the turbine blades. The cooling mechanism for turbine blades must include cooling for all possible regions which are exposed to hot gas flow. The turbine blade tip is one of the critical regions which are severely exposed to hot gas flow occurring due to the leakage of gases from the clearance gap between the turbine tip and the shroud. Hence the tip of the turbine blade must be cooled effectively to prevent thermal expansion of the turbine blade tip due to heating. This cooling will eventually help to avoid rubbing of blades to the shroud which may cause their wear. In this paper the effect of provision of pins of two different diameters and heights over the turbine blade tip at the corners, on the heat transfer has been investigated. The results obtained were compared with the heat transfer of smooth tip two pass channels. Investigations were carried out at different Reynolds numbers ranging from 200000 to 450000. It was found that if pins are provided at the corners the local heating of the tip at corners is avoided. A heat transfer augmentation of about 1.3 times was observed as compared to a smooth surface with pressure drop of less than 6 %. Hence the proper arrangement and number of pin fins is recommended for augmentation of heat transfer over the turbine blade tip.
Heat Transfer and Pressure Distributions on a Gas Turbine Blade Tip
Journal of Turbomachinery, 2000
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 us...
Heat Transfer and Flow on the Squealer Tip of a Gas Turbine Blade
Journal of Turbomachinery, 2000
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 two-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 percent 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 percent of the blade span. Two different turbulence intensities of 6.1 and 9.7 percent at the cascade inlet are also considered for heat transfer measurements. Static pressure measurements are made in the midspan 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 ...
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. Computational fluid dynamics (CFD) predictions of blade tip heat transfer are compared with 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 10 6 , and total turning of 110 deg. Three blade tip configurations were considered; a 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 the blade span. The tip heat transfer results of the numerical models agree well with data. For the case in which side-by-side comparison with test measurements in the open literature is possible, the magnitude of the heat transfer coefficient in the " sweet spot " matches data exactly and shows 20–50% better agreement with experiment than prior CFD predictions of this same case.
1999
A combined experimental and computational study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first stage blade tip surface for a geometry typical of large power generation turbines (>100MW). This paper is concerned with the design and execution of the experimental portion of the study, which represents the first reported investigation to obtain nearly full surface information on heat transfer coefficients within an environment which develops an appropriate pressure distribution about an airfoil blade tip and shroud model. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57•10 6 , and total turning of about 110 degrees. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5% or 9%. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp-edge and rounded-edge tip geometries at each of the inlet turbulence intensity levels.
Experimental Evidence of Temperature Ratio Effect on Turbine Blade Tip Heat Transfer
Journal of Turbomachinery, 2018
Determination of a scalable Nusselt number (based on “adiabatic heat transfer coefficient”) has been the primary objective of the most existing heat transfer experimental studies. Based on the assumption that the wall thermal boundary conditions do not affect the flow field, the thermal measurements were mostly carried out at near adiabatic condition without matching the engine realistic wall-to-gas temperature ratio (TR). Recent numerical studies raised a question on the validity of this conventional practice in some applications, especially for turbine blade. Due to the relatively low thermal inertia of the over-tip-leakage (OTL) flow within the thin clearance, the fluids' transport properties vary greatly with different wall thermal boundary conditions and the two-way coupling between OTL aerodynamics and heat transfer cannot be neglected. The issue could become more severe when the gas turbine manufacturers are making effort to achieve much tighter clearance. However, there ...
THERMAL ANALYSIS OF A TURBINE ROTOR BLADE
The turbine entry temperature (TET) is a vital parameter for power output and the efficiency for an aero-engine. In present day gas turbines the TET exceeds the melting temperature of the blade material. Advanced blade cooling technologies have resulted in high TET with acceptable material temperatures. In this paper, the authors present the thermal analysis of a typical Turbine Rotor blade.1D coolant flow network was modeled in Flowmaster software for an arbitrary cooling configuration. Rotational speed and heat transfer effects have been taken in the 1D code. Heat Transfer and pressure loss correlations available in open literature have been used in the 1D code. The converged convective bulk temperatures and convective heat transfer coefficients obtained from 1D code have been applied on 2D FEM model to obtain nodal temperature distribution. Parametric analysis has been carried out with different source pressure and gas temperature to study the effect of blade metal temperature.
Thermal Analysis of a Cooled Turbine Blade
IOP Conference Series: Materials Science and Engineering, 2019
Using Computational Fluid Dynamic (CFD), a gas turbine with an air-cooled blade was analyzed thermally. In terms of design, the domain was divided into three regions. The first region is the blade to blade passage (external flow) governed by a quasi-3-D Euler equation in a conservative form. Mac-Cormack's technique algorithm based on the finite differences, was used for this region. The second region involves the coolant passage (internal flow). This region involved the use of 2-D axi-symmetric Navier Stokes equations (Finite volume with staggered grids). A 3-D Laplace heat transfer equation was applied for the third region which involves a blade metal, where the solution algorithm was based on the finite difference technique. Consequently, to achieve temperature distribution through blade metal, the three regions have been coupled via the external and internal boundaries. Six different cases were examined in same blade geometry. The effect of gas heat transfer coefficient was analyzed in cases 1 and 2. With regards to cases 3, 4, 5, and 6, gas and coolant temperatures were changed. The computational results showed that the blade surface (metal) temperature is cooler than the surrounding gases (external hot gases) by about 100-500 oC, depending on boundary condition. An increase in gas temperature by 100 oC resulted in 50-100 oC increase in metal temperature, while, an increase in coolant temperature by 100 oC resulted in an average 50 oC increase in blade temperature. The results also show a temperature difference in blade metal of 250-450 oC between the leading and trailing edges.