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Heat/Mass Transfer Characteristics on Turbine Shroud with Blade Tip Clearance
Abstract
: The present study is conducted to investigate the local heat/mass transfer characteristics on the shroud with blade tip clearances. The relative motion between blade and shroud is neglected, and four-bladed linear cascade is used in this study. naphthalene sublimation method is employed to determine the detailed local heat/mass transfer coefficients on the surface of shroud. The tip clearance is changed from 0.66%t to 2.85%t of the blade chord length. The flow enters the gap between the blade tip and shroud at the pressure side due to the high pressure side due to the high pressure difference. Therefore, the heat/mass transfer characteristics of the shroud are changed significantly from those with endwall. A first, high heat/mass transfer occurs along the profile of blade a the pressure side due to the entrance effect and acceleration of the gap flow. Then, the heat/mass transfer coefficients on the shroud increase along the suction side of the blade because tip leakage vortices are generated and interact with the main flow. The results show that the heat/mass transfer characteristics are changed largely with the gap distance between the tip of turbine blade and the shroud.
... Last, the effects of temperature and heat transfer coefficients (aerodynamic effects) are not decoupled. Rhee et al. [8] studied a stationary blade in a low-speed linear cascade. They used a naphthalene sublimation method to determine the local heat/mass transfer coefficients on the shroud surface. ...
... They also showed that the trajectory of the leakage flow moved away from the suction side as both the tip gap and the tip leakage increased. It must be noted that Rhee et al. [8] worked with a low-speed facility, which, according to Moore et al. [11], is inadequate in simulating the true flow structure in the tip gap. Another experimental study was made by Kwak and Han [9] and investigated the effect of the tip crown geometry. ...
... • There is a rise in Nusselt number right above the PS squealer rim, which corresponds well to flow entrance effects and to the acceleration of the flow. This feature can also be seen in the results of Rhee et al. [8] and Kwak and Han [9]. The CFD tends to overestimate this increase in Nusselt on the squealer rim. ...
Low emission requirements for heavy-duty gas turbines can be achieved with flat combustor temperature profiles, reducing the combustor peak temperature. As a result, the heat load on the first stage heat shield above the first stage blade increases. High lift airfoils cause increased thermal loading on the heat shield above the blade tip and impact the unavoidable secondary flows, including complex vortex flows. Cascade tests have been performed on a blade with a generic high lift profile and the results on the heat shield are presented. A transient thermochromic liquid crystal measurement technique was used to obtain heat transfer coefficients on the heat shield surface. Several variations of blade tip clearance were investigated, and the impact on heat transfer coefficients is shown. Computational fluid dynamics predictions are compared to the experimental data to interpret the data and validate the CFD.
The wakes that occur periodically owing to blade rotation, affect the HTCDs (heat transfer coefficient distributions) on the turbine blade tip and shroud. We investigated the wake effects on the HTCDs from the turbine blade tip with different configurations and the corresponding shroud. The NSM (naphthalene sublimation method) was used for measuring the local heat/mass transfer distributions (LHMTDs). Five different tip shapes were compared. The flow characteristics near turbine blade tip and shroud varied with respect to different tip shape. Especially, swirling flow and tip leakage flow patterns varied, which affected the HTD on the tip & shroud. In addition, the local heat transfer distributions (HTCDs) on the tip and the shroud varied with different tip configurations. However, when the wake effects were induced, the HTCDs on the tip & shroud became relatively uniform. Furthermore, the magnitudes of HTCDs on the tip & shroud were relatively similar in region of 0.3 < x/Cx < 0.8. owing to the wake effect. Therefore, the vertical rib case can be the best-adopted for the turbine blade tip by considering the reduced total PLC (pressure loss coefficients) and HTCDs.
A wake is induced in turbine passages due to the interaction of turbine stators and rotors. Each stage of the turbine blade has different geometric parameters, namely, the turning angle, leading edge radius, and solidity, which all vary. Comparison of flow and thermal characteristics between the first and second-stage blades was carried out. The flow and thermal characteristics of the first-stage blade endwall were investigated for various Strouhal numbers to determine the wake effect. Numerical simulations and experiments were performed to compare the flow and thermal characteristics of the different stage blades. Experiments were performed using a five-bladed linear cascade with moving cylindrical rods simulating the wake effect. The effective area and strength of secondary vortices differed in each stage of blade, which resulted in different local heat transfer distributions. In addition, the local heat transfer characteristics changed depending on the Strouhal number. The ‘without wake’ case (St = 0) showed non-uniform heat transfer distribution on the endwall with the occurrence of horseshoe, passage and corner vortices. The ‘with wake’ effect cases (St ≠ 0) showed a more uniform heat transfer distribution on the endwall. The wake effect disturbs the occurrence of secondary vortices. With an increase in the Strouhal number, the endwall is exposed to higher thermal load.
The effect of a misalignment between vane endwall and combustor exit in a gas turbine was investigated using a Computational Fluid Dynamics (CFD) simulation and experimental measurements. The misaligned endwall platform was simulated as a backward facing step in this study. The CFD simulation predicted two legs of the vortex, referred to as a step-induced vortex, created by the step flowing through nozzle passage. Heat transfer measurements demonstrated the effect of the step-induced vortex on the endwall and the vane surface indicated by locally increased heat transfer coefficients which corresponded to the locus of the vortex, as also predicted by the simulation. Although a boundary layer transition occurred early, the locally increased heat transfer persisted to the vane trailing edge. In summary, a misaligned endwall platform causes negative effects on the gas turbine with respect to the thermal design. A vortex was generated by the step, which caused a higher thermal load on the nozzle vane surfaces, especially near the endwall.
The present study aimed to investigate the effect of an unsteady wake on the heat transfer for the endwall surface of a linear turbine blade cascade. A naphthalene sublimation method was implemented to obtain the detailed heat/mass transfer distributions on the endwall surface. Tests were conducted on a five-passage linear cascade in a low-speed wind tunnel. The effects of unsteady wakes were simulated in the facility by a wake generator consisting of circular rods that were traversed across the inlet flow. The test conditions were fixed at a Reynolds number of 70,000 based on the inlet velocity and chord length. The flow coefficients were varied from 1.3 to 4.2. and range of Strouhal number was 0.1 to 0.3. The results showed that the heat transfer distributions differed between steady and unsteady test cases. The overall heat transfer for the unsteady test cases was higher, and the heat transfer was enhanced with increasing the Strouhal number due to the resulting thin boundary layer and high turbulence intensity. Therefore, a cooling system for the endwall of a rotor should focus on decreasing the high temperatures of the endwall surface induced by the unsteady wakes.
Detailed heat/mass transfer coefficients and film-cooling effectiveness were measured on the tip and inner rim surfaces of a rotor blade with a squealer rim. The blade was a two-dimensional version of a modern first-stage gas turbine rotor blade with a squealer rim. The experimental apparatus was equipped with a linear cascade of three blades, the axial chord length (Cx ) of which was 237 mm with a turning angle of 126°. The mainstream Reynolds number based on the axial chord was 1.5×105 . The turbulence intensity level at the cascade inlet was approximately 12%. Measurements were made at three different rim heights (H) of about 3%, 6%, and 9% of the axial chord length. The tip clearance (C) ranges were 1–3% of the axial chord length. Also, three different types of blade tip surfaces were equipped with a single row of film-cooling holes along the camber line, near the pressure and the suction side rim. In particular, a coolant was injected at an incline of 45° from near the suction side film cooling holes. The film cooling experiments were done with a fixed tip clearance and rim height at 1% and 6% of the axial chord length. The blowing rate was fixed at 1.5. High heat transfer rates were observed near the leading edge on the tip surface in some cases, due to the reattachment of tip leakage flow. The peak values moved toward the suction-side edge, and the magnitude and area of high heat transfer increased near the leading edge as the tip clearance increased. The heat transfer decreased on the tip surface with increases in the rim height. In the film-cooling cases, the high heat transfer and film-cooling effectiveness region appeared near the film-cooling holes.
Detailed heat/mass transfer coefficients and film-cooling effectiveness were measured on the tip and inner rim surfaces of a blade with a squealer rim. The test blade was a two-dimensional version of a modern first-stage gas turbine rotor blade with a squealer rim. The experimental apparatus was equipped with a linear cascade of three blades, and the axial chord length (Cx) was 237 mm with a turning angle of 126°, the mainstream Reynolds number based on the axial chord and inlet velocity was 1.5 × 105. In addition, three different types of blade tip surfaces were equipped with a single row of film-cooling holes along the camber line, near the pressure and suction-side rim. The blowing ratio was fixed at 1.5. High heat transfer rates were observed near the leading edge on the tip surface due to reattached flow. Furthermore, heat transfer on both inner side surfaces was higher than that on the tip surface. High film cooling effectiveness was observed in the middle region (0.1 < X/Cx < 0.6) due to stagnation of the film cooling. Ultimately, a proper cooling system is suggested to reduce the thermal load and enhance the film cooling effectiveness in the squealer tip.
In a modern gas turbine engine the outer casing (shroud) of the shroudless high-pressure turbine is exposed to a combination of high flow temperatures and heat transfer coefficients. The casing is consequently subjected to high levels of convective heat transfer, a situation that is complicated by flow unsteadiness caused by periodic blade-passing events. In order to arrive at an over-tip casing design that has an acceptable service life it is essential for manfacturers to have appropriate predictive methods and cooling system configurations. It is known that both the flow temperature and boundary layer conductance on the casing wall vary during the blade-passing cycle. The current article reports the measurement of spatially and temporally resolved heat transfer coefficient (h) on the over-tip casing wall of a fully-scaled transonic turbine stage experiment. The results indicate that h is a maximum when a blade-tip is immediately above the point in question, while lower values of h are observed when the point is exposed to the rotor passage flow. Time-resolved measurements of static pressure are used to reveal the unsteady aerodynamic situation adjacent to the over-tip casing wall. The data obtained from this fully-scaled transonic turbine stage experiment are compared to previously published heat transfer data obtained in low-Mach number cascade style tests of similar high pressure blade geometries.
Detailed heat transfer distributions on the endwall and along the vane/blade surface are essential for component mechanical integrity and life predictions. Due to secondary flows, high gradients in heat transfer are present at the endwall and at the vane or blade surface itself where the passage vortex influences the mainstream flow. This paper documents the benchmarking of three turbulence models; (1) k-ε realizable with wall functions (2) k-ε realizable with two layer model, and (3) the V2F model for endwall and surface heat transfer and flowfield predictions. Benchmark experimental data from a scaled-up low speed rig for both a stator and rotor geometry are used for comparisons of heat transfer and flowfield. While the k-e realizable turbulence models give a good prediction of the secondary flow pattern, the heat transfer at the endwall and at the surface is not well predicted due to the inadequate modeling of near wall turbulence. The V2F model gives better agreement with the experiments on the endwall and vane midspan heat transfer is also well predicted, although transition occurs too far upstream on the suction surface. The results from this study represent the feasibility of CFD utilization as a predictive tool for local heat transfer distributions on a vane/blade endwall.
The effect of relative position of the stationary turbine blade for the fixed vane has been investigated on blade tip and shroud heat transfer. The local mass transfer coefficients were measured on the tip and shroud fur the blade fixed at six different positions within a pitch. A low speed stationary annular cascade with a single turbine stage was used. The chord length of the tested blade is 150 mm and the mean tip clearance of the blade having flat tip is 2.5% of the blade chord. A naphthalene sublimation technique was used for the detailed mass transfer measurements on the tip and the shroud. The inlet flow Reynolds number based on chord length and incoming flow velocity is fixed to 1.5{\times}10^5. The results show that the incoming flow condition and heat transfer characteristics significantly change when the relative position of the blade changes. On the tip, the size of high heat/mass transfer region along the pressure side varies in the axial direction and the difference of heat transfer coefficient is up to 40% in the upstream region of the tip because the position of flow reattachment changes. On shroud, the effect of tip leakage vortex on the shroud as well as tip gap entering flow changes as the blade position changes. Thus, significantly different heat transfer patterns are observed with various blade positions and the periodic variation of heat transfer is expected with the blade rotation.
Experiments were conducted in a low speed stationary annular cascade to investigate local heat transfer characteristics on the tip and shroud and the effect of inlet Reynolds number on the tip and shroud heat transfer. Detailed mass transfer coefficients on the blade tip and the shroud were obtained using a naphthalene sublimation technique. The turbine test section has a single stage composed of sixteen guide vanes and blades. The chord length and the height of the tested blade are 150 mm and about 125 mm, respectively. The blade has flat tip geometry and the mean tip clearance is about of the blade chord. The inlet flow Reynolds number based on chord length and incoming flow velocity is changed from to investigate the effect of Reynolds number. Flow reattachment after the recirculation near the pressure side edge dominates the heat transfer on the tip surface. Shroud surface has very intricate heat/mass transfer distributions due to complex flow patterns such as acceleration, relaminarization, transition to turbulent flow and tip leakage vortex. Heat/mass transfer coefficient on the blade tip is about 1.7 times as high as that on the shroud or blade surface. Overall averaged heat/mass transfer coefficients on the tip and shroud are proportional to respectively.
The local heat/mass transfer characteristics on the tip and shroud were investigated using a low speed rotating turbine annular cascade. Time-averaged mass transfer coefficients on the tip and shroud were measured using a naphthalene sublimation technique. A low speed wind tunnel with a single stage turbine annular cascade was used. The turbine stage is composed of sixteen guide plates and blades. The chord length of blade is 150 mm and the mean tip clearance is about 2.5% of the blade chord. The tested Reynolds number based on inlet flow velocity and blade chord is 1.5×105 and the rotational speed of the blade is 255.8 rpm at design condition. The results were compared with the results for a stationary blade and the effects of incidence angle of incoming flow were examined for incidence angles ranging from −15 to +7 degree. The off-design test conditions are obtained by changing the rotational speed with a fixed incoming flow velocity. Flow reattachment on the tip near the pressure side edge dominates the heat transfer on the tip surface. Consequently, the heat/mass transfer coefficients on the blade tip are about 1.7 times as high as those on the blade surface and the shroud. However, the heat transfer on the tip is about 10% lower than that for the stationary case due to reduced leakage flow with the relative motion. The peak regions due to the flow reattachment are reduced and shifted toward the trailing edge and additional peaks are formed near the leading edge region with decreasing incidence angles. But, quite uniform and high values are observed on the tip with positive incidence angles. The time-averaged heat/mass transfer on the shroud surface has similar a level to that of the stationary cases.
A numerical analysis has been conducted in order to simulate the characteristics of complex flow through linear cascades of A numerical analysis has been conducted in order to simulate the characteristics of complex flow through linear cascades of
high performance turbine blade with/without tip clearance by using a pressure-correction based, generalized 3D incompressible high performance turbine blade with/without tip clearance by using a pressure-correction based, generalized 3D incompressible
Navier-Stokes CFD code. The development and generation of horseshoe vortex, passage vortex, leakage vortex, tip vortex within Navier-Stokes CFD code. The development and generation of horseshoe vortex, passage vortex, leakage vortex, tip vortex within
tip clearance, etc. are clearly identified through the present simulation which uses the RNG k-ε turbulent model with wall tip clearance, etc. are clearly identified through the present simulation which uses the RNG k-ε turbulent model with wall
function method and a second-order linear upwind scheme for convective terms. The present simulation results are consistent function method and a second-order linear upwind scheme for convective terms. The present simulation results are consistent
with the generally known tendency that occurs in the blade passage and tip clearance. A 3D model for secondary and leakage with the generally known tendency that occurs in the blade passage and tip clearance. A 3D model for secondary and leakage
flows through turbine cascades with/without tip clearance is also suggested from the present simulation results, including flows through turbine cascades with/without tip clearance is also suggested from the present simulation results, including
the effects of tip clearance height. the effects of tip clearance height.
This study was carried out to investigate the effect of relative blade position on heat transfer in a stationary blade and shroud. A low speed wind tunnel with a single stage stationary annular turbine cascade was used. The test section is composed of sixteen guide plates and sixteen blades. The chord length of the blade is 150 mm and the mean tip clearance of the blade is 2.5% of the blade chord. Detailed mass transfer measurements were conducted for the stationary blade fixed at six different relative blade positions within a single pitch using a naphthalene sublimation method. The Reynolds number based on blade inlet velocity and chord length ranged between 1.0×105 and 2.3×105 and mean turbulence intensity was about 3%. As the blade position changed, the incoming flow field condition also changed significantly due to a blockage effect. As a result, the heat transfer on the tip and the shroud was significantly affected by the blade position because the incoming flow condition is changed. Especially, the mass transfer coefficients in the upstream region of the tip vary up to ± 25% of their average values. On the shroud, the size and the level of peak regions due to flow acceleration, transition and tip leakage vortex were strongly affected by the relative blade position.
Effects of the rim height and the tip gap clearance on the heat transfer coefficients on the blade tip and near tip regions were measured with two different rim geometries. The heat transfer coefficient distributions were measured using the transient single color capturing liquid crystals technique. Rims were located along (a) the pressure and the suction side (full-rim case) and (b) the suction side of the blade tip (suction side rim case). The rim heights were (a) 2.1%, (b) 4.2%, and (c) 6.3% and the blade tip gap clearances were (a) 1.0%, (b) 1.5%, and (c) 2.5% of the blade span. Tests were performed on a five-bladed linear cascade placed in a blowdown facility. The overall pressure ratio, inlet total pressure to exit static pressure, was 1.2, and the Reynolds number based on the exit velocity and the axial cord length was 1.1 × 106. The turbulence intensity level at the cascade inlet was 9.7%, and the inlet and exit Mach number were 0.25 and 0.59, respectively. It was found that higher rims reduce the heat transfer coefficients on the tip and shroud, but the reduction on the pressure and suction sides was not significant. The suction side rim case provided lower heat transfer coefficients on the blade tip and near tip regions than the full-rim case.
Detailed measurements of the subsonic flow in a large-scale, plane turbine cascade were made to evaluate the three-dimensional nature of the flow field. Tests were conducted at a passage aspect ratio of 1.0 with a collateral inlet boundary layer. Flow visualization was done on airfoil and endwall surfaces. Velocity and pressure measurements were taken before and behind the cascade and in six axial planes within a cascade passage, using a five-hole probe. Hot wire measurements were taken in the endwall boundary layer within the cascade passage. The characteristics of the endwall boundary layer are presented, showing that three-dimensional separation is an important feature of end-wall flow. A large part of the endwall boundary layer was found to be very thin when compared to the cascade inlet boundary layer. Data showing the growth of aerodynamic loss through the passage are discussed.
The complex three-dimensional flow in the endwall region near the base of a turbine blade has an important impact on the local heat transfer. The initial horseshoe vortex, the passage vortex, and resulting corner vortices cause large variations in heat transfer over the entire endwall region. Due to these large surface gradients in heat transfer, conventional measurement techniques generally do not provide an accurate determination of the local heat transfer coefficients. In the present study, the heat/mass transfer analogy is used to examine the local transport coefficients for two different endwall boundary layer thicknesses and two free-stream Reynolds numbers. A linear turbine blade cascade is used in conjunction with a removable endwall plate. Naphthalene (C10H8) is cast into a mold on the plate and the rate of naphthalene sublimation is determined at 6000+ locations on the simulated endwall by employing a computer-aided data acquisition system.
A numerical computation of the leakage flow is used to link a simple
flow and heat transfer model to a specific turbine geometry and
operating point for which a unique set of measured local tip and shroud
heat fluxes is available. Clearance leakage flow and accompanying
heat-transfer are of interest because of long-obvious effects on
structural durability and aerodynamic performance. This model should be
useful in the understanding and interpretation of future measurements
and also for providing early design estimates of the levels of tip and
shroud heat transfer that need to be compensated for by active turbine
cooling.
A visualization of the flow on the suction side and end-wall of a passage between two neighboring turbine blades is compared with mass (heat) transfer measurements on the same surfaces. Besides the horseshoe and passage vortices, there are several smaller vortices formed near the junction of blade and end-wall whose origins are discussed. The vortices detach from the end-wall and move up the blade's span. These vortices, sometimes in counter rotating pairs, are responsible for substantial local variations of heat transfer.Der Durchflu auf der Saugseite und an der Endwand eines Kanals zwischen zwei benachbarten Turbinenschaufeln wird sichtbar gemacht und mit gemessenen Massen-(Wrme)strmen der selben Oberflchen verglichen. Neben den huffrmigen Wirbeln und Durchfluwirbeln werden mehrere kleinere Wirbel in der Nhe der Verbindungsstelle von Schaufel und Endwand gebildet und die Ursache ihrer Entstehung wird diskutiert. Die Wirbel lsen sich von der Endwand ab und bewegen sich ber die Spannweite der Schaufel. Diese Wirbel, manchmal als gegenlufige rotierende Paare, sind fr die wesentlichen lokalen Variationen des Wrmeflusses verantwortlich.
The naphthalene sublimation method can be used to study mass and heat transfer with confidence for a variety of applications, but with certain restrictions. This method is particularly useful in complex flows and geometries and for flows with large gradients in wall transport rate. Mass transfer boundary conditions analogous to isothermal and adiabatic walls in convective heat transfer can be easily imposed. Furthermore, the nature of mass transfer allows one to impose these boundary conditions such that errors analogous to conductive losses in a wall are not present. The test specimen can be easily prepared by several methods, including dipping, machining spraying, and casting. The local transfer coefficients can be determined with high accuracy and in detail by automated measurement systems that eliminate most human errors and reduce extraneous sublimation losses during the measurement process. The heat transfer coefficient, which is often desired, can be readily determined from the measured mass transfer results with good confidence via a heat/mass transfer analogy. However this method cannot generally be used in certain flow situations such as high-velocity flows because of recovery temperature effects and natural convection due to the thermal buoyancy effects of sublimation latent heat.
A combined computational and experimental 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. 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 x 10(exp 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.
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