Measurement Uncertainty Analysis in Determining Adiabatic Film Cooling Effectiveness by Using Pressure Sensitive Paint Technique

Over the past several decades, researchers from around the world have embraced the use of Pressure Sensitive Paint (PSP) to obtain detailed film cooling effectiveness distributions. Utilizing the mass transfer analogy along with the luminescence characteristics of PSP, the mixing behavior of the coolant and mainstream flows is easily quantified. As the method relies on tracking the concentration of various gases in contact with the paint, the measurements are not affected by thermal diffusion through the test surface. Therefore, researchers provide film cooling effectiveness distributions that are spatially, highly resolved, even in the near-hole regions. Such distributions are not only needed to visualize the three-dimensional mixing occurring near the holes, but the measurements can also be used to validate numerical simulations. When PSP was first introduced to the gas turbine and film cooling communities, steady, film cooling effectiveness distributions were primarily obtained in low-speed flows with simple geometries and coolant-to-mainstream density ratios of unity. However, as more researchers have integrated this measurement tool into their labs, the use of the paint has expanded to include flow conditions and cooling configurations relevant to advanced gas turbine engines. In this review, the mass transfer analogy for film cooling effectiveness measurements will be explained, and the recent, widespread deployment of the method will be discussed. Within the past decade, laboratories have successfully utilized traditional PSP for full coverage airfoil cooling over a range of coolant-to-mainstream density ratios simulating engine conditions. In addition, “binary” and “fast response” paints have gained popularity. The binary PSP is most often used in high speed flows, where viscous heating may affect the response of traditional PSP. A second luminophore in the binary paint allows for correction of the local temperature variation and improves the accuracy of the mass transfer analogy. Finally, fast response PSPs now allow researchers to visualize the unsteady interaction of the coolant and mainstream flows, yielding time-resolved film cooling effectiveness distributions. This review will discuss these developments and provide opportunities for further advancement into the future.

An experimental investigation has been completed to obtain detailed film cooling effectiveness distributions on a cooled turbine blade platform within a linear cascade. The platform has a labyrinth-like seal upstream of the blades to model a realistic stator-rotor seal configuration. Additional coolant is supplied to the downstream half of the platform via discrete film cooling holes. The coolant flow rate through the upstream seal varies from 0.5% to 2.0% of the mainstream flow, while the blowing ratio of the coolant through the discrete holes varies from 0.5 to 2.0 (based on the mainstream velocity at the exit of the cascade). Detailed film cooling effectiveness distributions are obtained using the pressure sensitive paint (PSP) technique under a wide range of coolant flow conditions and various freestream turbulence levels (0.75% or 13.4%). The PSP technique clearly shows how adversely the coolant is affected by the passage induced flow. With only purge flow from the upstream seal, the coolant flow rate must exceed 1.5% of the mainstream flow in order to adequately cover the entire passage. However, if discrete film holes are used on the downstream half of the passage, the platform can be protected while using less coolant (i.e. the seal flow rate can be reduced).

Detailed film cooling effectiveness distributions are measured on the leading edge of a rotating gas turbine blade with two rows (pressure-side row and suction-side row from the stagnation line) of holes aligned to the radial axis using the pressure sensitive paint (PSP) technique. Film cooling effectiveness distributions are obtained by comparing the difference of the measured oxygen concentration distributions with air and nitrogen as film cooling gas respectively and by applying the mass transfer analogy. Measurements are conducted on the first-stage rotor blade of a three-stage axial turbine at 2400rpm (positive off-design), 2550rpm (design), and 3000rpm (negative off-design), respectively. The effect of three blowing ratios is also studied. The blade Reynolds number based on the axial chord length and the exit velocity is 200,000 and the total to exit pressure ratio was 1.12 for the first-stage rotor blade. The corresponding rotor blade inlet and outlet Mach numbers are 0.1 and 0.3, respectively. The film cooling effectiveness distributions are presented along with discussions on the influence of rotational speed (off design incidence angle), blowing ratio, and upstream nozzle wakes around the leading edge region. Results show that rotation has a significant impact on the leading edge film cooling distributions with the average film cooling effectiveness in the leading edge region decreasing with an increase in the rotational speed (negative incidence angle).

Film-cooling effectiveness and leakage flow play crucial roles in turbine blade heat transfer. In contrast to the conventional trailing edge cutback squealer blade tip (CB tip), this study aims to compare the film-cooling effectiveness of a suction side squealer tip (SS tip) under different blowing ratios and density ratios. This experiment is conducted in a three-blade cascade wind tunnel under transonic conditions. While applying the pressure-sensitive paint (PSP) technique for film-cooling effectiveness measurements, both top-view and side-view perspectives were employed to illustrate the distribution of local cooling effectiveness across various regions of the blade. Furthermore, PSP can be used to estimate the leakage flow across the blade tip by analyzing the local static pressure distribution. For the CB tip, the film-cooling effectiveness within the cavity increases as both the blowing ratio and density ratio increase. The PSP technique clearly captures the accumulation of coolant within the cavity and the coolant discharge near the trailing edge in the vicinity of the squealer cutback. When the blade tip is outfitted with only a suction side squealer, the film-cooling effectiveness on the blade tip is comparable to that of the full squealer with the trailing edge cutback. This similarity is achieved as the CB tip design requires 14–20% more coolant to achieve the same blowing ratios as the SS tip. The advantage of the SS tip is further demonstrated as leakage flow across the tip is reduced compared to the CB tip.

The paper is focused on the effect of leading edge airfoil geometry on endwall film cooling. Fillets placed at the junctions of the leading edge and the endwall are used in investigation. Three types of fillet profiles are tested, and the results are compared with baseline geometry without fillet. The design of the fillet is based on the suggestion by previous literature data indicating that sharp is effective in controlling the secondary flow. Three types of sharp slope fillet with the length to height ratio of 2.8, 1.2 and 0.5 are made using stereo lithography (SLA) and assessed in the experiment. Distributed with the approximately inviscid flow direction, four rows of compound angle laidback fan-shaped holes are arranged on the endwall to form full covered coolant film. The four rows of fanshaped holes are inclined 30 deg to the endwall surface and held an angle of 0, 30, 45 and 60 deg to axial direction respectively. The fanshaped holes have a lateral diffusion angle of 10 deg from the hole-centerline and a forward expansion angle of 10 deg to the endwall surface. The Reynolds number based on the axial chord and inlet velocity of the free-stream flow is 3.5*105, and the testing is done in a four-blade cascade with low Mach number condition (0.1 at the inlet) while the blowing ratio of the coolant through the discrete holes varies from 0.4 to 1.2. The film-cooling effectiveness distributions are obtained using the PSP (pressure sensitive paint) technique, by which the effect of different fillet geometry on passage induced flow and coolant is shown. The present paper compares the film cooling effectiveness distributions in a baseline blade cascade with three similar blades with different leading edge by adding fillets. The results show that with blowing ratio increasing, the film cooling effectiveness increases on the endwall. For specific blowing ratio, the effects of leading edge geometries could be illustrated as follows. The baseline geometry provides the best film cooling performance near leading edge pressure side. As for the leading edge suction side, the best leading edge geometry depends on the blowing ratio. The longfillet is the more effective in controlling horseshoe vortex at low blowing ratio, but for the high blowing ratio shortfillet and mediumfillet are better.

Effects of the presence of squealer, the locations of the film cooling holes, and the tip gap clearance on the film cooling effectiveness were studied and compared to plane tip. The film cooling effectiveness distributions were measured on the blade tip using Pressure Sensitive Paint technique. Air and nitrogen gas were used as the film cooling gases and the oxygen concentration distribution for each case was measured. The film cooling effectiveness information was obtained from the difference of the oxygen concentration between air and nitrogen gas cases by applying the mass transfer analogy. Plane tip and squealer tip blades were used while the film cooling holes were located (a) along the camber line on the tip or (b) along the span of the pressure side. The average blowing ratio of the cooling gas was controlled to be 0.5, 1.0, and 2.0. Tests were conducted in a five-bladed linear cascade with a blow down facility. The free stream Reynolds number, based on the axial chord length and the exit velocity, was 1,100,000 and the inlet and the exit Mach number were 0.25 and 0.59, respectively. Turbulence intensity level at the cascade inlet was 9.7%. All measurements were made at three different tip gap clearances of 1%, 1.5%, and 2.5% of blade span. Results show that the locations of the film cooling holes and the presence of squealer have significant effects on surface static pressure and film-cooling effectiveness.

The objective of this study is to examine the effect of the geometrical modification of the land on the overall film cooling effectiveness on the cutback region of a turbine blade model. A room-temperature experiment was conducted, in which nitrogen serves as the cooling stream, and the mainstream flow is air. The adiabatic film cooling effectiveness was mapped employing the pressure-sensitive paint (PSP) technique. Data was acquired at five different blowing ratios (from 0.45 to 1.65) for both the baseline and the modified model. Detailed film cooling effectiveness from PSP measurements in correlation with the flow map in streamwise and spanwise planes from particle image velocimetry (PIV) measurements was performed, characterizing the effect of rounding the edges of the lands. The results show that the rounded edges enable the coolant flow to reach the top surface of the land area more readily, especially at low blowing ratios. Superior coolant coverage on the land surface observed in the PSP measurements are well correlated with the PIV measurements. At the high blowing ratio of 1.65, the round edge of the lands helps regulate the mixing between the coolant and mainstream flows, therefore the film cooling effectiveness in the slot region is also improved.

Combined effects of unsteady wake and free-stream turbulence on turbine blade film cooling with laid-back fan-shaped holes using PSP technique

The impact of the purge flow injection on aerodynamics and film cooling effectiveness of a three-stage high pressure turbine with non-axisymmetric endwall contouring has been experimentally investigated. As a continuation of the previously published work involving stator-rotor gap purge cooling, this study investigates film cooling effectiveness on the first stage rotor contoured platform due to a coolant gas injection. Film cooling effectiveness measurements are performed on the rotor blade platform using a pressure sensitive paint (PSP) technique. The present study examines, in particular, the film cooling effectiveness due to injection of coolant from the rotor cavity through the circumferential gap between the first stator followed by the first rotor. Efficiency and performance experiments were conducted with and without cooling injection to show (a) the impact of endwall contouring on the turbine efficiency and (b) the impact of film cooling injection in association with the endwall contouring. The experimental investigation is carried out in a three-stage turbine facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL) at Texas A&M University. Its rotor includes non-axisymmetric endwall contouring on the first and second rotor row [1]. The turbine has two independent cooling loops. Film cooling effectiveness measurements are performed for three coolant-to-mainstream mass flow ratios of 0.5%, 1.0% and 1.5%. Film cooling data is also obtained for three rotational speeds, 3000 rpm (reference condition), 2550 rpm and 2400 rpm and compared with non-contoured endwall data.

Experimental study on film cooling performance of a turbine blade tip with a trapezoidal slot cooling scheme in transonic flow using PSP technique

The pronounced nonuniform temperature distribution in the core engine flow path is a recurring problem of gas turbine engine design process. Specifically, turbine entry conditions are usually characterized by severe temperature distortions, often referred to as hot and cold streaks, ascribed to combustor burners and combustor liners cooling systems. Temperature distortions remain an issue even at the exit section of the nozzle guide vane (NGV), where additional cold streaks coming from the vane film-cooling system are injected into the flow. A precise knowledge of thermal field and its evolution is thus essential to mitigate their impact on turbine performance and lifetime. Various studies focus on the description of streaks migration through a direct investigation of the thermal field, providing an effective evaluation of the global phenomenon. As a deeper understanding is often required, experimental techniques based on the detection of tracer gases can be successfully adopted. In this study, a realistic combustor outlet swirl profile was imposed on a fully cooled NGV cascade to analyze both film-cooling behavior and cold streaks migration and redirection. A concentration probe based on the fluorescence behavior of an oxygen sensor, fully characterized in a previous work, was here employed to track the position of the film cooling flows at the NGV cascade exit plane, while the adiabatic film-cooling effectiveness was evaluated on the NGV surfaces employing the pressure sensitive paint (PSP) technique. Overall, the swirling structure strongly affects both the film-cooling behavior and cold streaks migration through and downstream the vane. The importance of examining the unsteady aspect is also highlighted to better estimate actual components operating temperatures. A global understanding of the occurring phenomena is therefore provided, as well as significant pieces of information that can be extremely useful for the design phases of both the NGV and the following rotor cascade.

The visualization and measurements of aerodynamic effects on a 3D aircraft model were conducted using an optical pressure measurement system, based on the pressure-sensitive paint (PSP) technique. PSP technology provides a good understanding of the flow around the wind tunnel model. The PSP technique can be used to carry out absolute pressure measurements on a surface of the model and to determine additional aerodynamic data using scientific-grade cameras and image processing techniques. Surface pressures from the top, bottom, left, and right viewing directions were obtained using the DLR-PSP system on the entire surface, which can be observed by eight CCD cameras. Finally, the measured pressures can be integrated to calculate the forces and moments of the complete model, or parts thereof.

A detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform. The platform was cooled by purge flow from a simulated stator–rotor seal combined with discrete hole film-cooling. The cylindrical holes and laidback fan-shaped holes were accessed in terms of film-cooling effectiveness. This paper focuses on the effect of coolant-to-mainstream density ratio on platform film-cooling (DR = 1 to 2). Other fundamental parameters were also examined in this study—a fixed purge flow of 0.5%, three discrete-hole film-cooling blowing ratios between 1.0 and 2.0, and two freestream turbulence intensities of 4.2% and 10.5%. Experiments were done in a five-blade linear cascade with inlet and exit Mach number of 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 750,000 and was based on the exit velocity and chord length of the blade. The measurement technique adopted was the conduction-free pressure sensitive paint (PSP) technique. Results indicated that with the same density ratio, shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The optimum blowing ratio of 1.5 exists for the cylindrical holes, whereas the effectiveness for the shaped holes increases with an increase of blowing ratio. Results also indicate that the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity.

The nonuniformity of the Hp turbine inlet flow field put forward higher requirements for NGV (Nozzle Guide Vanes) leading edge and gill region film cooling. The assumption of design condition in most of the experiments couldn’t reflect the true operation environment in the Hp turbine NGV. The factor of off-design condition was incorporated into the experiment in this research. The GE-E3 Hp turbine nozzle guide vanes were used in the experiment to investigate the cooling performance of injection from leading edge and gill region with inlet Reynolds numbers of Re = 3.5×105 and inlet Mach number of Ma = 0.1. The compound angle fan-shaped film cooling hole configuration was applied. The cooling characteristics at off-design condition were analyzed and compared in the paper. The leading edge and gill region film cooling performance was assessed with the incidence angle varying from i = −10deg to i = +10deg. The blowing ratio varying from M = 0.7 to M = 1.3, was also selected as an experimental variable. Film cooling effectiveness distribution was measured using PSP (Pressure Sensitive Paint) technique. The film cooling performance of the compound angle fan-shaped holes was assessed at both design and off-design conditions. The object of this research is to change the concept that NGV leading edge film cooling experiment only needs the data at design condition. Through the comparative analysis of experimental results at different inlet flow angle, the influence of off-design condition on NGV leading edge and gill region film cooling could be illustrated at a reasonable level.

Several studies have previously been conducted to improve the cooling performance of film cooling. However, most of the research has conducted experiments with film cooling holes on flat plates, and thus, the results of these studies do not encompass the influence of the complex mainstream behavior within the turbine passage on film cooling. In this study, three different film cooling hole configurations were installed on the endwall of a turbine linear cascade to measure adiabatic film cooling effectiveness and evaluate cooling performance. The film cooling holes compared in the experiment for film cooling effectiveness were a 7-7-7 fan-shaped hole (Baseline), a Baseline with a double-step structure at the hole exit (Staircase), and a Baseline with an additional expanded passage at the hole leading edge (Compound Expansion). A total of nine holes were manufactured on the turbine endwall to assess film cooling performance, as various factors, such as mainstream acceleration, secondary flow within the turbine passage, and so on, can influence film cooling. Adiabatic film cooling effectiveness was measured using the pressure-sensitive paint (PSP) technique. Mass flow ratios ranging from 0.25% to 1.25% of the mass flow rate of a single turbine passage were supplied to the plenum chamber within the test rig. As a result, all experimental results confirmed the impact of secondary flow within the turbine passage on film cooling. In the case of the Staircase, it exhibits an overall cooling trend similar to the Baseline. It shows small cooling performance degradation compared with Baseline due to lift-off, and its double-step structure laterally expanding results in better cooling performance at high mass flow ratio (MFR) conditions. For the Compound Expansion, at low MFR, the momentum of the coolant is lower compared with other configurations, leading to lower cooling performance due to the influence of secondary flow. However, at high MFR, the Compound Expansion provides wider protection compared with other hole geometries and shows high cooling performance.