Pressure Sensitive Paint Research Articles

A comprehensive deviation analysis of film cooling effectiveness measurements using pressure sensitive paint under various conditions

The application of pressure-sensitive paint (PSP) has been extensively utilized to measure the effectiveness of film cooling. During the experiment, the pressure fields for the air-jet case and the foreign-gas-jet case are usually assumed to be identical. However, when there is a disparity in density between the foreign gas and the air, the difference in pressure distributions cannot be disregarded in high-speed or complex flow conditions. The film cooling effectiveness deviation caused by pressure variation was analyzed by formula derivation. The deviation distribution was obtained through the numerical simulation of the experimental process. The results indicated that the deviation is universal under various conditions. For the film cooling on a flat plate, there is a butterfly-shaped negative deviation region and a narrow symmetrical positive deviation zone. The presence of the compound angle significantly increases the deviation and results in a positive deviation region upstream of the hole. The measurement deviation is widely distributed and exhibits significant magnitudes on the blade tip. The deviation range on the suction side is primarily concentrated near the hole. Specifically, the deviation on the pressure side can be almost neglected. A correction method is proposed, involving the use of an oxygen mixture with the same density as the foreign gas and the same oxygen mole fraction as air. This method proves effective in significantly reducing the deviation in all cases.

Experimental and numerical investigation on the metering and diffuser length effects of a laidback fan-shaped film cooling hole

In this study, the metering and diffuser length effects of a laidback fan-shaped (LBF) hole on the film cooling effectiveness (FCE) were investigated. The FCE was measured using the pressure sensitive paint (PSP) technique, with a density ratio (DR) of 2.0 and five blowing ratios (BR). The experimental results indicated that the FCE increased as the diffuser length increased. However, when the metering length ratio increased beyond the critical length, the coolant was shifted to a specific side of the diffusion section, resulting in a lower FCE. A computational analysis was performed to investigate the coolant shift at longer metering lengths. Numerical results indicate that the center velocity of the cylinder section is higher for the longer metering length cases, which leads to a large separation bubble in the diffusion section, these resulted in the increase of the flow unsteadiness and instability at the hole exit.

Mass-Heat Coupling in Film Cooling Experiments—The Influence of Temperature and Turbulence on the Dufour Effect

Abstract The use of foreign gases in laboratory film cooling experiments is attractive since variable density ratios can be achieved with coolant-to-freestream temperature ratios near unity, often reducing the cost and difficulty of the experimental campaign. In adiabatic effectiveness experiments employing pressure sensitive paint along with the mass transfer analogy to heat transfer, isothermal surfaces are often an experimental requirement. Furthermore, low-temperature laboratory experiments using thermal techniques often employ relatively close matches between the coolant and freestream temperatures. Using foreign gases, however, introduces off-diagonal couplings of heat and mass transport, which can produce unexpected results in film cooling experiments. In particular, the Dufour effect—also called the diffusion-thermo effect—which is the transfer of thermal energy by mass transfer processes, can manifest in surface temperatures that break the traditional bounds of thermal adiabatic effectiveness experiments: outside the upper and lower bounds of the coolant and freestream temperature. Beyond the expected confusion for the researcher, this effect can also be detrimental to those that assume that matching the coolant and freestream temperatures are the necessary and sufficient conditions to ensure isothermal surface conditions in traditional pressure sensitive paint experiments. In this work, the influence of cooling gas selection, experimental temperature, and experimental freestream turbulence conditions are explored on a simulated leading edge with compound injection from a cylindrical cooling hole. Air, argon, carbon dioxide, helium, and nitrogen coolants were analyzed due to their use in prior film cooling studies. The Dufour effect was found to be significant when using helium as the coolant, though temperature separation was also observed in argon and carbon dioxide cases. Additionally, elevated experiment temperatures generally increased temperature separation. Finally, high freestream turbulence intensity was found to reduce, but not eliminate, the Dufour effect in helium experiments.

Fast-Responding Pressure-Sensitive Paint Measurements of the IC3X at Mach 7.2

Global surface pressure measurements of a 5.7% scale AFRL Initial Concept 3.X vehicle (IC3X) were obtained using a fast-responding ruthenium-based pressure-sensitive paint (PSP) at the UTSA Mach 7 Ludwieg Tube Wind Tunnel at two different angles of attack, 0° and 2.5°. Static calibration of the paint was performed over a range of 0.386 kPa to 82.7 kPa to relate luminescent intensity to pressure. Details on the facility, paint preparation, application, calibration, and image processing techniques are provided in the manuscript. The results from statistical, spectral, and proper orthogonal decomposition (POD) analyses are presented to characterize the pressure field observed on the model. The experimental results qualitatively follow the expected trends and correspond to the occurrence of shock waves and expansion fans, which were visualized via Schlieren imaging. The theoretical pressure range obtained from conical shock analysis for 0° agrees with the experimentally derived pressure range for the model, and the outliers are attributed to errors in image registration. This study presents preliminary pressure measurements that pave the way for obtaining time-resolved global PSP measurements to train and validate aerothermodynamic machine learning models.

Thermophysical properties of pressure-sensitive paint

The paper deals with the problem of the thermophysical properties of pressure-sensitive paint (PSP). The paint was used to determine the pressure distribution on the surface of models examined in wind tunnel aerodynamic tests. The aim of the study was to provide reliable data on the values of thermophysical parameters and their dependence on temperature. PSP is a thin-film structure that is applied layer by layer a specific number of times on a metallic substrate, and determining the thermal properties requires non-standard test procedures to obtain reliable results. The measurements were carried out on paint fragments separated from the flat substrate of a metallic thin film (molybdenum), from which samples were obtained for complementary studies of the thermophysical properties. The study includes weight measurements, thermogravimetric analysis, micro-calorimetric analysis, and measurement of effective thermal diffusivity using the temperature oscillation technique. The findings are complemented with structural test data collected by means of scanning electron microscopy, surface characterisation, and density determination by the displacement method. The analysis of the structural test data indicates the micro-structural isotropy of the examined material with geometric anisotropy of the layer as a whole. In this way, the results of thermal diffusivity tests can be properly evaluated, and reliable data can be obtained on the complete thermophysical properties of the PSP layer. The indirectly determined thermal conductivity, i.e., by recalculating the specific heat, density, and thermal diffusivity data, allows the test material to be classified as an insulator. Apart from the presentation of the test results and calculations, this paper also contains a brief discussion of the test procedures.

Application of fast-response pressure sensitive paint to low-speed vortical flow at high angles of attack

Application of fast-response pressure sensitive paint to low-speed vortical flow at high angles of attack

Lifetime based pressure and temperature sensitive paint measurement on a civil aircraft wing under conditions near the onset of shock-induced separation

Lifetime-based pressure- and temperature-sensitive paint (PSP and TSP) measurements were conducted in a large-scale industrial transonic wind tunnel to obtain high-quality pressure data for validation of computational fluid dynamics (CFD) simulations. A wind tunnel test was performed using the NASA common research model in the JAXA 2 m × 2 m transonic wind tunnel. The freestream Mach number was varied in the range of 0.70–0.89, and aerodynamic forces and moments were acquired in order to obtain the pitch break condition related to the onset of shock-induced separation. Polymer-based PSP was coated on a main wing of the model, and TSP was used in tandem with PSP to correct the temperature dependence of PSP. The two-gate lifetime-based method was applied to obtain the PSP and TSP emissions. An a-priori/in-situ hybrid calibration was conducted to convert ratio-of-ratios to pressure, and measured pressure distributions were mapped onto a three-dimensional (3D) model grid. The root-mean-square error for the pressure measurements was evaluated by pressure tap data and was approximated to be 0.8 kPa for all Mach numbers tested. The obtained pressure distributions exhibited a significantly high signal-to-noise ratio and were used for comparison with CFD results on a 3D grid. The high-spatial-resolution PSP measurements helped to accurately localize the differences from the CFD simulation results and showed that the prediction of the shock location along the main wing is still a relevant challenge.

Comparison of Corcos-Based and Experimentally Derived Coherence Factors for Buffet Forcing Functions

In this paper, high-spatial-resolution unsteady pressure-sensitive paint (UPSP) data are utilized to compare two methods for panel buffet forcing function (BFF) estimation for the Space Launch System (SLS). Such methods are based on discrete pressure measurements within a panel but employ coherence factors to account for partially correlated fluctuating pressures across the whole panel. In one method, coherence factors are derived based on the Corcos model, whereas the second method utilizes experimentally derived coherence factors. To simulate discrete measurements using UPSP data, suitable subsets of the data are extracted. When full UPSP resolution is retained, UPSP data provide a benchmark to assess discrete-measurement-based methods. The analysis focuses on the peak SLS buffet environment located downstream of the forward attachment hardware (FAH) between the core stage and solid rocket boosters. Trends of the Corcos-based and experimentally derived coherence factors are in reasonable agreement with the benchmark. However, at certain frequencies, experimentally derived coherence factors are sensitive to the separation distance between pressure measurements utilized to compute coherence lengths. Such sensitivity originates from deviation of the experimentally based coherence function from an exponential decay assumption. On the other hand, the present implementation of the Corcos model fails to capture certain nonturbulent boundary-layer-related environments, such as a subharmonic of FAH vortex shedding. For all methods presented in this paper, at near-transonic conditions, increased pressure coherence and spatial nonuniformity lead to BFF overestimation and sensitivity to the pressure measurement location within the panel.

Effects of structural parameters of laidback fan-shaped hole on film cooling effectiveness on convex surface

To burden higher turbine inlet temperature, shaped holes are introduced to the film cooling of turbine blade. The structural parameters of shaped holes are the key variables for the cooling design. Relevant literatures conduct studies mainly on the flat surface, which excludes the influence of blade curvature. And their conclusions are inconsistent, appealing for further mechanism study. In this paper, the effects of structural parameters of laidback fan-shaped hole are experimentally investigated on a convex surface. Pressure sensitive paint technique is used to measure the adiabatic effectiveness. Three parameters, including forward expansion angle (0–12 deg), lateral expansion angle (2–20 deg) and length of the diffuser (2–14 mm), at five blowing ratios (0.5, 0.8, 1.0, 1.3 and 1.5) are tested. Numerical simulation is also employed to present the mechanism of jet-mainstream interaction. It is shown that, at the blowing ratio over 0.8, the optimal forward and lateral expansion angle lies in the medium range. Cooling effectiveness increases before the optimal value due to better film attachment but deteriorates after then due to coolant separation within the hole. In contrast, at the blowing ratio of 0.5, the results become different because the coolant attaches well enough that the change of the coolant dissipation rate appears more dominant. The competition of two factors, weaker kidney vortex entrainment and larger shear surface, leads to an unpredictable trend. Finally, suggestions are provided for the parameter selection on the blade suction surface. As a comparison, the same studies are numerically conducted on a flat surface. Distinctly, at high blowing ratio, the effectiveness increases monotonously without a drop at too large expansion because the absence of radial pressure gradient worsens jet attachment and magnifies the effect of promoting jet attachment.

The Development and Application of Two-Color Pressure-Sensitive Paint in Jet Impingement Experiments

The Development and Application of Two-Color Pressure-Sensitive Paint in Jet Impingement Experiments

Turbulence Model Effects on Vortex Interaction Prediction on a Multiswept Delta Wing

In this study, vortical flows around a generic triple-delta wing geometry were investigated at transonic speeds. The focus of this work was the investigation of vortex interaction at medium to high angles of attack. To this end, numerical simulations were performed at two different Mach numbers, M=0.5 and M=0.85. To assess the performance of different turbulence models in predicting the complex flowfield, comparisons with experimental data were performed. In addition to the integral forces and moments, surface pressures from pressure-sensitive paint and velocity field data from particle image velocimetry were used. Across all investigated test cases, the best agreement with experimental results was achieved by the Menter shear stress transport turbulence model.

Assessment of Binary Pressure-Sensitive Paint for Temperature and Heat Transfer Coefficient Measurement of Leading-Edge Film Cooling

Abstract Film cooling is a common technique for protecting gas turbine components from the hot combustor exhaust. Highly resolved film cooling effectiveness distributions are often obtained by measuring the mass transfer of a foreign gas coolant in mainstream air using pressure-sensitive paint (PSP). However, PSP is not able to measure the heat transfer coefficient, which is necessary to fully quantify the impact of film cooling. Instead, binary pressure-sensitive paint (BPSP) has an additional luminophore that is sensitive to temperature and can be used to measure the heat transfer coefficient. In this experiment, the film cooling effectiveness and heat transfer coefficient were measured using BPSP on the leading edge of a cylinder. The cylinder had a 7.62-cm diameter with two rows of cooling holes at ±15 deg from the leading edge. Each row contained ten holes with a 0.475-cm diameter, spaced 4 diameters apart in the spanwise direction, and angled 30 deg from the cylinder axis. The mainstream Reynolds number was 100,000 based on cylinder diameter with a turbulence intensity of 7.1%. The coolant-to-mainstream density ratio was 1.0, and the blowing ratio was 0.8. The heat transfer coefficient was measured in a transient heat transfer experiment using the reference signal from the BPSP. Despite the high uncertainty of the measurement, ranging from 24.0% to 71.1%, the results demonstrate the feasibility of the method and identify the best test methodology to minimize conduction errors.

Flow Topology of a Slanted Cylinder with Rounded-Edges Under Compressible Conditions

In this study, the slanted afterbody was modified to more closely represent cargo aircraft afterbodies through the addition of a basal edge fillet. Two slant angles ( and ) were explored over two Mach numbers ( and ) and two Reynolds numbers ( and ) using oil flow visualization and pressure-sensitive paint to identify surface flow phenomena, while the spanwise shadowgraph enabled the observation of mean wake phenomena. The rounded-edge afterbody surface measurements displayed a centerline-separated vortex state with flow features and regions of suction qualitatively similar to those of the baseline sharp-edge case, with some notable differences. For example, a separation bubble at the slant leading edge is nearly twice the size of the baseline sharp-edge model. This is accompanied by the presence of a “source point,” indicated by a region of high pressure where the flow reattaches, and a tighter counter-rotating vortex pair in comparison to the baseline. The rounded-edge prevented the previously seen compressible transition to a fully separated wake state within the Mach number range studied while showing an increase in separation bubble size with increasing Mach number. Also noteworthy was the existence of a new vortex state for the model, where the flow remains attached at the model centerline. This is in contrast to the previously observed flow topology of the centerline-separated vortex state in the sharp-edge model, which formed a separation bubble at the model centerline. The transition between the new centerline-attached vortex state and the traditional centerline-separated vortex state was shown to be dependent on both Mach number and Reynolds number, revealing the importance of boundary-layer development on the rounded afterbody flowfield. Overall, these observations indicate that the rounded afterbody produces flowfields with distinct features that vary from the previously described sharp-edge afterbody wake field under certain conditions.

An experimental and numerical study on the blade tip conjugate heat transfer performance

In the gas turbine, the blade tip is under extreme heat load and thereby it is necessary to adopt cooling method for the tip region. In this work, a blade tip conjugate heat transfer model was conducted in a linear cascade. The tip overall cooling effectiveness (φ) and adiabatic cooling effectiveness (η) were measured by the pressure sensitive paint (PSP) and the Infrared thermography (IR), respectively. Impacts of coolant blowing ratios (M = 1.0 ∼ 2.5), tip gap size (τ = 1.98%, 3.97%C, chord length), and film holes locations on the tip φ distributions were investigated. Then, the numerical simulations were performed to obtain the tip solid temperature distributions, and to reveal the conjugate heat transfer characteristics. Within the scope of this work, both the tip φ distributions and the tip solid temperature distributions indicate that the convection in film holes is a major factor of the tip conjugate heat transfer process. The tip area-averaged φ for the CL_H case are increased evidently by 13.1%∼38.9% with the increased coolant M, whereas the increase of η is not obvious. The tip gap size has less effect on the tip φ distributions than the coolant M, especially at a low M. The higher φ and the lower tip solid temperature region are detected near the film holes for any film hole location. The area-averaged φ values of the cases with film holes along the chamber line are 10.6% and 3.0% higher than those of the cases with film holes along the pressure-side or suction-side rim, respectively. This work revealed the major factor in the tip conjugate heat transfer process and may serve as a guide for tip cooling design.

Effect of Upstream Leakage Flow on Film Cooling Characteristic of a Turbine Convex Endwall

Abstract This paper investigates the effect of upstream leakage coolant flow on the film cooling performance of the convex endwall. The experiment is conducted in an annular passage with four vanes, and the endwall film cooling effectiveness distribution is measured by the pressure-sensitive paint technique. The near-wall flow field distribution predicted by computational fluid dynamics is employed to gain a deeper understanding of the influence of the passage secondary flow on the film cooling characteristic of the endwall. Furthermore, the effect of mass flow ratios (MFR = 0.75%, 1.0%, 1.25%, and 1.5%), leakage slot inclination angles (α = 30 deg, 45 deg, and 60 deg), density ratios (DR = 1.0, 1.5, and 2.0), as well as Reynolds number (Re = 2.0 × 105, 3.0 × 105, and 4.0 × 105) on the endwall film cooling is also investigated. Results indicate that the mass flow ratio has a great influence on the endwall film cooling effectiveness distribution, and a larger uncooled region can be observed when MFR is less than 1.25% due to the cooling air more easily being entrained by the secondary flow near the endwall. Reducing the inclination angle of the leakage slot enhances the axial velocity component of the coolant flow, thereby weakening the secondary flow near the endwall and improving the film cooling effectiveness distribution of the endwall. As the density ratio increases, the jet momentum of the leakage flow is reduced, particularly for higher MFR cases, which results in a significant reduction in the film cooling performance of the endwall. Besides, thinning the incoming boundary layer of the passage reduces the strength and size of the secondary flow in the passage and improves the endwall adiabatic effectiveness distribution when the Reynolds number is increased.

Investigation of aerothermal performance of blade tip with conical holes in transonic flow

The blade tip with conical holes is proposed in the current paper to improve the film cooling performance on the turbine blade tip. The film cooling performance of the blade tip with cylindrical holes is studied as a reference case. The pressure sensitive paint technique is adapted to obtain the tip film cooling effectiveness in transonic flow. The numerical simulation is used to obtain the tip flow field and aerodynamic loss. The cascade exit Mach number and inlet Reynolds number are 1.05 and 370000, respectively. The experiment and calculation are carried out at two tip clearance gaps of 0.7% and 1.5% and four mass-flow-ratios. The results show that at low mass flow ratios, the film cooling effectiveness of the blade tip with conical holes is significantly higher than that of the blade tip with cylindrical holes at the tip mid-chord region. Increasing the tip clearance gap from 0.7% to 1.5% shows a negative effect on the film cooling performance of the blade tip with conical holes in the tip mid-chord region, particularly for the low mass flow ratio. The pitch-wise averaged aerodynamic loss coefficient of the blade tip with cylindrical holes is lower than that of the blade tip with conical holes near the shroud.

Large Eddy simulation in optimizing trench configuration to improve cooling effectiveness of a laidback fan-shaped hole

In modern gas turbines, a thermal barrier coating (TBC) is used to protect the components of the turbine from the hot flow coming from the combustor, which results in the cooling holes on the blade surfaces being embedded in a 2D trench slot. In this study, large eddy simulations (LES) were carried out to investigate the effects of trench geometry on the film-cooling effectiveness of a shaped cooling hole. The shaped cooling hole is located on a flat surface with an injection angle of 30 degrees, and the blowing and density ratios are fixed at 1.5. Thirteen different designed cases were sampled using Box-Behnken method based on three different geometry parameters of the trench slot, trench height and upstream and downstream trench lengths. A genetic aggregation algorithm was applied to achieve the optimal trench configuration that maximizes the overall cooling effectiveness. The film-cooling performance computed by the LES method was validated for a reference case without a trench and three different trenched cases with the experiments using the PSP (pressure sensitive paint) technique. The cooling effectiveness of the trench cavity and flat plate was found to be significantly enhanced in cases of trenches with relatively greater trench height and less downstream trench length. The optimal trench configuration demonstrated less turbulent fluctuations in the mixing region, and the overall area-averaged adiabatic film-cooling effectiveness was improved by 23% in comparison with a reference hole without a trench.

Film Cooling on Turbine Vane Endwalls With Different Inlet Cooling Configurations: Experimental and Computational Results

Abstract A comparative study of turbine endwall film cooling resulting from three different cooling configurations in front of the passage inlet was conducted. The three inlet cooling configurations considered in this study were a conventional continuous slot, double rows of discrete film holes, and an interrupted slot with a backward-facing step. Pressure sensitive paint (PSP) was sprayed over the endwall surfaces to map adiabatic cooling effectiveness contours and five-hole probe and thermocouple measurements were implemented to assess aero-thermal fields at the passage exit. The evolution of cooling effectiveness was compared across a full range span of coolant flow rates of 0.3−1.8% in a real engine. Additionally, complex flow structures inside the passage due to interactions of coolant with mainstream flows were visualized using computational fluid dynamics (CFD) simulations to support the observed coolant coverage patterns. A side-by-side comparison of CFD simulations against experiments was made as well to evaluate the reliability of conventional turbulence modeling methods in such a complex flow. The continuous slot was found to produce the highest cooling effectiveness values while the double-row holes offered the most uniform coolant coverage, leading to spreading out of the coolant jets downstream of the passage throat and even beyond the trailing edge of the passage for coolant injection rates higher than 1.0%. Supported by the CFD-predicted flow structures and measured aerodynamic losses, the injection from the double-row holes weakened the passage secondary flows and thus caused the lowest aerodynamic loss. On the contrary, the interrupted slot enhanced the secondary flows due to the backward step, resulting in the highest flow loss.

Experimental Investigation of Effusion Film Cooling on a Cylindrical Leading Edge Model

Abstract Effusion film cooling is effective for cooling high-temperature turbine blades because it requires less coolant and produces a more uniform temperature distribution than conventional film cooling. Effusion cooling for a cylindrical model representing the leading edge of a gas turbine blade was investigated. The experiment was performed in a low-speed wind tunnel at a Reynolds number of 100,000. Pressure-sensitive paint was used to measure the adiabatic film cooling effectiveness. Additive manufacturing was used to fabricate a porous structure on the test cylinder for effusion cooling. Both simple and compound angles were used for cooling injection. The effects of streamwise and spanwise hole spacings, turbulence intensities (1% and 8.7%), and blowing ratios (0.075, 0.15, 0.3, and 0.6) were studied at a fixed density ratio of 1. The effusion hole diameter was 0.1 cm, and the spanwise hole pitch-to-diameter ratio was either 2 or 4. Compared with conventional film cooing, effusion cooling achieved a higher cooling effectiveness and produced a better coolant coverage. Increasing the streamwise spacing noticeably reduced the cooling effectiveness for the simple-angle design due to film lift-off; the compound-angle designs thus achieved higher effectiveness. The simple-angle holes were more sensitive to changes in the mainstream turbulence intensity; increases in the turbulence intensity promoted the mixing of the coolant with the mainstream. Moreover, effusion cooling was more resistant to coolant lift-off at high blowing ratios.

Film Cooling Effectiveness on Pressure Surface and Suction Surface of Turbine Guide Vane With Diffusion Slot Holes

Abstract This paper investigated the film-cooling effectiveness of diffusion slot holes in a turbine nozzle guide vane. The pressure-sensitive paint measurement technique was employed to obtain the film-cooling effectiveness at a density ratio of DR = 1.5. The mainstream Reynolds number based on the axial chord length and the exit velocity was 6 × 105. The mainstream turbulence intensity was approximately 3.7%. Three diffusion slot hole geometries with cross-sectional aspect ratios (ASs) of 2.3, 3.4, and 4.9 were tested and compared with a typical fan-shaped hole. The experiments were performed at three typical hole row locations on the pressure surface (PS) and suction surface (SS). The average blowing ratios varied from M = 0.5 to 2.5. The results showed that throughout the blowing ratio range, on the PS, a substantially higher film-cooling effectiveness than the fan-shaped hole is always obtained from the diffusion slot hole with a large aspect ratio (AS = 4.9); on the SS, the diffusion slot hole with a small AS (AS = 2.3). The influence of hole row positioning is inconsistent for diffusion slot holes with different ASs. The diffusion slot hole is less affected by the PS when the AS is moderate and less affected by the SS when the AS is large. The film-cooling effectiveness of the diffusion slot holes is basically the lowest where the PS has a maximum concave curvature and the highest where the SS has a large favorable pressure gradient.