Film-cooled Blade Research Articles

Novel Strategy for Thermal Evaluation of Film-Cooled Blades Using Thermographic Phosphors at Simulated Engine Conditions

Abstract Practical strategy for the thermal evaluation of film-cooled blade is of great importance to the gas turbine community. Due to the physical or methodology limitations, it is difficult to evaluate the blade’s thermal performance at simulated engine conditions. The present study proposed novel focal-sweep-based phosphor thermometry for blade cooling inspection. While Mg4FGeO6:Mn (MFG) served as the temperature sensor to quantify the blade temperatures as well as simulated the thermal barrier coating (TBC) effect, the focal sweep method was adopted to overcome the optical constraints in cascade testing. The obtained MFG results of microstructures, jet impingement, and anti-erosion test demonstrated that the MFG phosphor is robust enough to simulate the thermal insulation effect of TBC and can withstand high-speed flow erosion. Furthermore, the proposed strategy clearly captured the blade temperature distributions (mainstream at T0,∞=∼850K) with high spatial resolution, which was then successfully remapped onto the three-dimensional twisted blade. Additional comparisons with the thermocouples demonstrated that the simulated TBC has a thermal insulation effect of about 68 K. This study addressed the common problems of phosphor thermometry in blade cooling evaluation, offering a practical strategy for future thermal diagnostics of the gas turbine.

Analysis of Rotor–Stator Interaction on the Aerothermal and TBC Insulation Performance of a Turbine Stage under Hot Streak Inlet Condition

Hot streaks and rotor–stator interaction have a great influence on the aerothermal performance of turbine blades. Previous investigations have conducted limited study of the film-cooled blade. To investigate the combined effects of a hot streak and rotor–stator interaction on the coated blade, an unsteady numerical simulation has been conducted with an efficient unsteady Navier–Stokes solver in this paper. The numerical results at four different relative stator–rotor locations (t = 0/4 T, 1/4 T, 2/4 T, and 3/4 T) have been investigated in one stator period. Compared with the stator, rotor–stator interaction exerts a significant impact on the cooling performance of the rotor blade under hot streak inlet conditions. The overall cooling effectiveness distribution of the coated rotor blade is similar to that of the uncoated blades in one stator period. Relatively lower overall cooling performance of the rotor blade can be observed in the 1/4 stator period. Then, the cooling performance begins to increase and relatively larger cooling effectiveness can be observed in the 3/4 stator period. The addition of a TBC is generally beneficial to the improvement of blade surface cooling performance, especially for the areas with low overall cooling performance. However, a negative cooling effectiveness increment can be observed at the trailing edge. It shows that for an area with poor cooling performance, the addition of thermal barrier coating will have the opposite effect. Therefore, it is necessary to enhance the design of cooling arrangements at the trailing edge to maximize the insulation performance of TBCs for the coated rotor blade.

Conjugate heat transfer of impingement cooling using conical nozzles with different schemes in a film-cooled blade leading-edge

In this paper, the effect of utilizing different schemes of conical nozzles for impingement cooling is studied in the conjugate heat transfer in a film cooled blade leading-edge. Three conjugate cooling schemes (tangential, inline normal, staggered normal) are analyzed and compared at different nozzle Reynolds numbers from 5000 to 25,000 and different temperature ratios from 0.5 to 0.85. The SST k-Omega turbulence model and a very fine unstructured mesh are validated and applied for all simulations. Based on the design, the staggered normal scheme can protect the most critical zone of a blade subjected to the highest temperature (stagnation area) better than the other cooling schemes. The internal cooling performance increases with the nozzle Reynolds number under a fixed temperature ratio. At identical nozzle Reynolds number, the internal heat transfer increases with the temperature ratios from 0.5 to 0.675 while it decreases over 0.675 up to 0.85. The staggered normal scheme achieves an increase in overall Nusselt number by 5.26–9% and by 9.78–21.27% compared to the tangential scheme and inline normal scheme, respectively over the applied range of nozzle Reynolds number. The staggered normal scheme achieves the highest cooling performance internally and externally among the other cooling schemes.

Numerical investigation of particle deposition in film-cooled blade leading edge

This study numerically investigates film-cooling performance and particle trajectories in AGTB (two rows of cylindrical holes equipped on suction side (SS) and pressure side (PS) of the leading edge, respectively) turbine cascade. Particle deposition on a turbine blade is analyzed by calculations of capture efficiency and impact efficiency. The turbulent flow is modeled by the Realizable k-ε turbulence model, and the discrete phase model (DPM) with user-defined functions (UDFs) is used to simulate the particle motions. An invasion efficiency is proposed to analyze the possibility of particle invasion into the film hole. Comparisons of various particles with diameters of 1 µm, 5 µm, 10 µm, 20 µm, and 50 µm, respectively, are conducted for four blowing ratios (0.53, 0.93, 1.31, 1.63) and three inlet flow angles (123°, 133°, and 143°). It is observed that with a small inlet flow angle and a large blowing ratio, the capture efficiency on the PS decreases. It is found that smaller particle size results in lower invasion efficiency, and larger particles are more likely to invade into the film-cooling hole especially at a low blowing ratio.

Experimental measurements of the flow field around a film-cooled blade leading edge using particle image velocimetry

In this article, particle image velocimetry studies were conducted in a low-speed wind tunnel to investigate the effects of blowing ratio and blade span in terms of the characteristics of the flow field around a film-cooled blade leading edge. The measurements were performed at 20%, 40%, 60%, and 80% of blade span and blowing ratios of M = 0.5, M = 0.75, M = 1, M = 1.5, and M = 2. Velocity, turbulence intensity, and structure of vortices during the interaction between cooling flow and mainstream were analyzed in detail. The analysis shows a significant increase in mainstream velocity at low blowing ratios, M < 1. Peaks of turbulence were observed at low- and high-span locations. Aerodynamical losses are expected at higher blowing ratios due to the formation of secondary vortices near the outgoing jet. These vortices were a consequence of velocity gradients at this zone.

The Effects of Coolant Pipe Geometry and Flow Conditions on Turbine Blade Film Cooling

The performance of gas turbine engines can be improved by increasing the inlet gas temperature. Turbine blades can be damaged by high gas temperature, unless additional cooling mechanisms are incorporated to maintain the blades below an acceptable temperature limit. Film cooling techniques are often used to cool the blades to avoid damages. The performance of film cooling depends on several parameters, however. In this paper past research on film cooling is reviewed and areas in need of further investigation are identified. Computational fluid dynamics (CFD) simulations are then conducted on the widely-used single-hole film cooling arrangements in which coolant jets are injected into air flows inside a straight channel before issuing onto the blades. Cooling pipe-blade configurations and flow conditions are varied and the resulting flow hydrodynamics are examined. Counter rotating vortex pairs (CRVPs) formed in the flow strongly influence the film cooling performance. Small coolant inclination angles, exit holes enlargement in span wise direction, higher injected fluid density, and higher injectedambient fluid velocity ratios are all found to maintain the CRVPs away from each other and close to wall - both of which promote cooling. Pipe curvature can be used for enhancing cooling by exploiting the centrifugal force effect.

Unsteady Heat Transfer and Pressure Measurements on the Airfoils of a Rotating Transonic Turbine With Multiple Cooling Configurations

Measurements are presented for a high-pressure transonic turbine stage operating at design-corrected conditions with forward and aft purge flow and blade film cooling in a short-duration blowdown facility. Four different film-cooling configurations are investigated: simple cylindrical-shaped holes, diffusing fan-shaped holes, an advanced-shaped hole, and uncooled blades. A rainbow turbine approach is used so each of the four blade types comprises a wedge of the overall bladed disk and is investigated simultaneously at identical speed and vane exit conditions. Double-sided Kapton heat-flux gauges are installed at midspan on all three film-cooled blade types, and single-sided Pyrex heat-flux gauges are installed on the uncooled blades. Kulite pressure transducers are installed at midspan on cooled blades with round and fan-shaped cooling holes. Experimental results are presented both as time-averaged values and as time-accurate ensemble-averages. In addition, the results of a steady Reynolds-averaged Navier–Stokes computational fluid dynamics (RANS CFD) computation are compared to the time-averaged data. The computational and experimental results show that the cooled blades reduce heat transfer into the blade significantly from the uncooled case, but the overall differences in heat transfer among the three cooling configurations are small. This challenges previous conclusions for simplified geometries that show shaped cooling holes outperforming cylindrical holes by a great margin. It suggests that the more complicated flow physics associated with an airfoil operating in an engine-representative environment reduces the effectiveness of the shaped cooling holes. Time-accurate comparisons provide some insight into the complicated interactions that are driving these flows and make it difficult to characterize cooling benefits.

Research on Theoretical Model of Combined Micro-Machining of Laser and Electrolysis of Thermal Barrier Coated Turbine Blade Film Cooling Holes

Research on Theoretical Model of Combined Micro-Machining of Laser and Electrolysis of Thermal Barrier Coated Turbine Blade Film Cooling Holes

Numerical Computation and Optimization of Turbine Blade Film Cooling

The effect of film cooling parameters on the cooling effectiveness of an actual turbine blade is studied numerically. Film cooling parameters such as the hole shape, holes distribution, blowing ratio, streamwise angle, and spanwise angle are investigated to select the appropriate cooling parameters. Unstructured finite volume technique is used to solve the steady, three-dimensional, and compressible Navier-Stokes equations. Using one cooling holes array indicates that the average overall film cooling effectiveness is enhanced by decreasing the streamwise angle for high blowing ratio on the suction side of the turbine blade. The film cooling effectiveness is enhanced on the pressure side for a blowing ratio of unity. In addition, the cooling effectiveness increases by increasing the lateral and forward diffusion angles.The computations reveal that the efficiency of cooling is decreased at the leading edge due to the large surface curvature of the blade. The presence of compound shape (spanwise angle) enhanced the film cooling effectiveness on the two sides. Multistagger cooling hole arrays are investigated and the results indicate that five-stagger cooling arrays on the pressure side and three-stagger cooling arrays on the suction side with LFDCA-9.3-14.6 hole shape are enough to have good cooling of the two sides using 2.17% bleed air of the engine.

Influence of Coolant Density on Turbine Blade Film-Cooling With Axial and Compound Shaped Holes

Adiabatic film-cooling effectiveness is examined systematically on a typical high pressure turbine blade by varying three critical flow parameters: coolant blowing ratio, coolant-to-mainstream density ratio, and freestream turbulence intensity. Three coolant density ratios 1.0, 1.5, and 2.0 are chosen for this study. The average blowing ration and the turbulence intensity are 1.5% and 10.5%, respectively. Conduction-free pressure sensitive paint (PSP) technique is used to measure film-cooling effectiveness. Foreign gases are used to study the effect of coolant density. Two test blades feature axial angle and 45 deg compound-angle shaped holes on the suction side and pressure side. Both designs have 3 rows of 30 deg radial-angle cylindrical holes around the leading edge region. The inlet and the exit Mach number are 0.27 and 0.44, respectively. Reynolds number based on the exit velocity and blade axial chord length is 750,000. Overall, the compound angle design performs better film coverage that axial angle. Greater coolant-to-mainstream density ratio results in lower coolant-to-mainstream momentum and prevents coolant to lift-off.

Conjugate calculation of a film-cooled blade for improvement of the leading edge cooling configuration

Great efforts are still put into the design process of advanced film-cooling configurations. In particular, the vanes and blades of turbine front stages have to be cooled extensively for a safe operation. The conjugate calculation technique is used for the three-dimensional thermal load prediction of a film-cooled test blade of a modern gas turbine. Thus, it becomes possible to take into account the interaction of internal flows, external flow, and heat transfer without the prescription of heat transfer coefficients. The focus of the investigation is laid on the leading edge part of the blade. The numerical model consists of all internal flow passages and cooling hole rows at the leading edge. Furthermore, the radial gap flow is also part of the model. The comparison with thermal pyrometer measurements shows that with respect to regions with high thermal load a qualitatively and quantitatively good agreement of the conjugate results and the measurements can be found. In particular, the region in the vicinity of the mid-span section is exposed to a higher thermal load, which requires further improvement of the cooling arrangement. Altogether the achieved results demonstrate that the conjugate calculation technique is applicable for reasonable prediction of three-dimensional thermal load of complex cooling configurations for blades.

Comparative Thermodynamic Analysis of Gas Turbine Systems with Turbine Blade Film Cooling

Since the gas turbine systems require active cooling to maintain high operating temperature while avoiding a reduction in the system operating life, turbine blade cooling is very important and essential but it may cause the performance losses in gas turbine. This paper deals with the comparative thermodynamic analysis of gas turbine system with and without regeneration by using the recently developed blade-cooling model when the turbine blades are cooled by the method of film cooling. Special attention is paid to investigating the effects of system parameters such as pressure ratio and turbine inlet temperature on the thermodynamic performance of the systems. In both systems the thermal efficiency increases with turbine inlet temperature, but its effect is less sensitive in simpler system

Investigation of Effect of Unsteady Interaction on Turbine Blade Film Cooling

Film cooling is extensively used to provide protection against the severe thermal environment in gas turbine engines. It is known that the cooling effectiveness of the downstream rotor is influenced by the unsteady wake of the stator. Presence of wake due to an upstream vane is studied by placing a periodic set of rods upstream of the blade in most of the investigations. In order to explore the influence of the unsteady stator wake on the cooling effectiveness of the rotor blade under real situation, the film cooling of moving turbine blades, with preceding stationary vanes, was investigated numerically, and the influence of the unsteady interaction between the rotor and the stator on the film cooling effectiveness of the flow field surrounding the cooling holes on moving blades was discussed. Results show that the unsteady interaction between the stator wake and the cooling jet flow becomes stronger for the case of larger blowing ratio under the same stator wake condition, and the fluctuation of the film cooling effectiveness is increased correspondingly. Because of the wider overlying range of the cooling jet flow at the time of the stator wake passing through the cooling holes, better film cooling effectiveness achieves downstream of the 30% arc length. Cooling jet flow can hardly reach the endwall zone due to the effect of the vortex at this zone. It is entrained downstream of the jet position, resulting in better effectiveness in this zone.

Improvement of a Film-Cooled Blade by Application of the Conjugate Calculation Technique

The conjugate calculation technique has been used for the three-dimensional thermal load prediction of a film-cooled test blade of a modern gas turbine. Thus, it becomes possible to take into account the interaction of internal flows, external flow, and heat transfer without the prescription of heat transfer coefficients. The numerical models consist of all internal flow passages and cooling hole rows, including shaped holes. Based on the results, deficiencies of the test configuration close to the leading edge region and in the blade tip region have been detected, which lead to hot spots and surface areas of high thermal load. These regions of high thermal load have been confirmed by thermal index paint measurements in good agreement to the conjugate calculation results. Based on the experimental and numerical results, recommendations for the improvement of the blade cooling were derived and an improved blade-cooling configuration has been designed. The conjugate calculation results, as well as new measurement data, show that the changes in the cooling design have been successful with respect to cooling performance. Regions of high thermal load have vanished, and effective cooling is reached for all critical parts of the test blade.

Modeling film-coolant flow characteristics at the exit of shower-head holes

Abstract The coolant flow characteristics at the hole exits of a film-cooled blade are derived from an earlier study where the hole pipes and coolant plenum were also discretized. The blade chosen is the VKI rotor with three staggered rows of shower-head holes. The present analysis applies these flow characteristics at the shower-head hole exits. A multi-block three-dimensional Navier–Stokes code with Wilcox's k – ω model is used to compute the heat transfer coefficient on the film-cooled turbine blade. A reasonably good comparison with the experimental data as well as with the more complete earlier analysis where the hole pipes and coolant plenum were also gridded is obtained. If the 1/7th power law is assumed for the coolant flow velocity and temperature distributions at the hole exits, considerable differences in the heat transfer coefficient on the blade surface, especially in the leading-edge region, are observed even though the span-averaged values of h match well with the experimental data. This calls for spatially-resolved experimental data near film-cooling holes on a blade for better validation of the code.

Effect of Unsteady Wake With Trailing Edge Coolant Ejection on Film Cooling Performance for a Gas Turbine Blade

The effect of unsteady wakes with trailing edge coolant ejection on surface heat transfer coefficients and film cooling effectiveness is presented for a downstream film-cooled turbine blade. The detailed heat transfer coefficient and film effectiveness distributions on the blade surface are obtained using a transient liquid crystal technique. Unsteady wakes are produced by a spoked wheel-type wake generator upstream of the five-blade linear cascade. The coolant jet ejection is simulated by ejecting coolant through holes on the hollow spokes of the wake generator. For a blade without film holes, unsteady wake increases both pressure side and suction side heat transfer levels due to early boundary layer transition. Adding trailing edge ejection to the unsteady wake further enhances the blade surface heat transfer coefficients particularly near the leading edge region. For a film-cooled blade, unsteady wake effects slightly enhance surface heat transfer coefficients but significantly reduces film effectiveness. Addition of trailing edge ejection to the unsteady wake has a small effect on surface heat transfer coefficients compared to other significant parameters such as film injection, unsteady wakes, and grid generated turbulence, in that order. Trailing edge ejection effect on film effectiveness distribution is stronger than on the heat transfer coefficients.

Recent Development in Turbine Blade Film Cooling

Gas turbines are extensively used for aircraft propulsion, land‐based power generation, and industrial applications. Thermal efficiency and power output of gas turbines increase with increasing turbine rotor inlet temperature (RIT). The current RIT level in advanced gas turbines is far above the .melting point of the blade material. Therefore, along with high temperature material development, a sophisticated cooling scheme must be developed for continuous safe operation of gas turbines with high performance. Gas turbine blades are cooled internally and externally. This paper focuses on external blade cooling or so‐called film cooling. In film cooling, relatively cool air is injected from the inside of the blade to the outside surface which forms a protective layer between the blade surface and hot gas streams. Performance of film cooling primarily depends on the coolant to mainstream pressure ratio, temperature ratio, and film hole location and geometry under representative engine flow conditions. In the past number of years there has been considerable progress in turbine film cooling research and this paper is limited to review a few selected publications to reflect recent development in turbine blade film cooling.

Heat transfer on a film-cooled blade – effect of hole physics

Abstract A multiblock, three-dimensional Navier–Stokes code has been used to study the within-hole and near-hole physics in relation to heat transfer on a film-cooled blade. The flow domain consists of the coolant flow through the plenum and hole-pipes for the three staggered rows of shower-head holes on the VKI rotor, and the main flow over the blade. A multiblock grid is generated that is nearly orthogonal to the various surfaces. It may be noted that for the VKI rotor the shower-head holes are inclined at 30° to the spanwise direction, and are normal to the streamwise direction on the blade. Wilcox's k–ω turbulence model is used. The present study provides a much better comparison for the span-averaged heat transfer coefficient on the blade surface with the experimental data than an earlier analysis wherein coolant velocity and temperature distributions were specified at the hole exits rather than extending the computational domain into the hole-pipe and plenum. Details of the distributions of coolant velocity, temperature, k and ω at the hole exits are also presented.

Effect of Unsteady Wake on Detailed Heat Transfer Coefficient and Film Effectiveness Distributions for a Gas Turbine Blade

Unsteady wake effects on detailed heat transfer coefficient and film cooling effectiveness distributions from a gas turbine blade with film cooling are obtained using a transient liquid crystal technique. Tests were performed on a five-blade linear cascade at a axial chord Reynolds number of 5.3 × 105 at cascade exit. Upstream unsteady wakes are simulated using a spoke-wheel type wake generator. The test blade has three rows of film holes on the leading edge and two rows each on the pressure and suction surfaces. Air and CO2 were used as coolants to simulate different coolant-to-mainstream density ratio effect. Coolant blowing ratio for air injection is varied from 0.8 to 1.2 and is varied from 0.4 to 1.2 for CO2. Results show that Nusselt numbers for a film-cooled blade are much higher compared to a blade without film injection. Particularly, film injection causes earlier boundary layer transition on the suction surface. Unsteady wakes slightly enhance Nusselt numbers but significantly reduce film cooling effectiveness on a film-cooled blade compared with a film-cooled blade without wakes. Nusselt numbers increase slightly but film cooling effectiveness increase significantly with an increase in blowing ratio for CO2 injection. Higher density coolant (CO2) provides higher effectiveness at higher blowing ratios (M = 1.2) whereas lower density coolant (Air) provides higher effectiveness at lower blowing ratios (M = 0.8).

Impact of Rotor–Stator Interaction on Turbine Blade Film Cooling

The goal of this study is to quantify the impact of rotor–stator interaction on surface heat transfer of film cooled turbine blades. In Section I, a steady-state injection model of the film cooling is incorporated into a two-dimensional, thin shear layer, multiblade row CFD code. This injection model accounts for the penetration and spreading of the coolant jet, as well as the entrainment of the boundary layer fluid by the coolant. The code is validated, in the steady state, by comparing its predictions to data from a blade tested in linear cascade. In Section II, time-resolved film cooled turbine rotor heat transfer measurements are compared with numerical predictions. Data were taken on a fully film cooled blade in a transonic, high pressure ratio, single-stage turbine in a short duration turbine test facility, which simulates full-engine nondimensional conditions. Film cooled heat flux on the pressure surface is predicted to be as much as a factor of two higher in the time average of the unsteady calculations compared to the steady-state case. Time-resolved film cooled heat transfer comparison of data to prediction at two spanwise positions is used to validate the numerical code. The unsteady stator–rotor interaction results in the pulsation of the coolant injection flow out of the film holes with large-scale fluctuations. The combination of pulsating coolant flow and the interaction of the coolant with this unsteady external flow is shown to lower the local pressure side adiabatic film effectiveness by as much as 64 percent when compared to the steady-state case.