Temperature Sensitive Paint Research Articles

Numerical Investigation of Nonlinear Boundary-Layer Transition for Cones at Mach 6

Direct numerical simulations were carried out for a straight and a flared cone at Mach 6 and zero angle of attack to investigate the effects of geometry on the linear and nonlinear stages of the laminar–turbulent transition process. The cone geometries and the flow conditions of the simulations are chosen to closely match those of the experiments conducted at the Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) at Purdue University. In the linear (primary) instability regime the flared cone resulted in larger integrated growth rates ( factors) and a higher dominant second mode frequency. Investigations of the secondary instability regime revealed that the flared cone also leads to a stronger fundamental resonance, characterized by larger integrated growth rates of the secondary instability waves after resonance onset. For both cases, the so-called controlled fundamental breakdown simulations were carried out, which showed the development of similar patterns of streamwise hot streaks on the surface of the cones. Streamwise streaks were also observed in the Purdue experiments (for the flared cone only) using temperature-sensitive paint. A detailed comparison of the results obtained for both geometries indicated that the streak development is caused by the same nonlinear mechanisms.

Measurement of boundary layer transition on oscillating airfoil using cntTSP in low-speed wind tunnel

cntTSP is a flow visualization technique that combines temperature-sensitive paint (TSP) and carbon nanotube (CNT). TSP is a sprayable temperature sensor, and CNT is used as a thin inner electric heater for the TSP. The cntTSP technique allows the visualization of the movement of the boundary layer transition on a moving airfoil. In this study, cntTSP was applied to a pitch-oscillating airfoil in a low-speed wind tunnel. Three methods were compared as methods for detecting the boundary layer transition position. The transition position was successfully evaluated using a heat transfer coefficient that was calculated by the surface temperature. Moreover, at a high-reduced frequency, the detected transition positions were different between the pitching-up and pitching-down procedure. Moreover, we observed a hysteresis loop of the transition position.

Thermophysical Properties of Temperature-Sensitive Paint.

The complex thermophysical property of temperature-sensitive paint (TSP) research is discussed. TSP is used for visualization of the surface temperature distribution in wind tunnel aerodynamic tests. The purpose of this research was to provide reliable, experimental, thermophysical data of the paint applied as a coating. As TSP is applied as thin surface layers, investigation of its final properties is challenging and demands the application of non-standard procedures. At present, most measurements were performed on composite specimens of TSP deposed onto a thin metallic film substrate or on TSP combined with a cellulose sheet support. The studies involved gravimetric,, thermogravimetric, and microcalorimetric analyses, transversal thermal diffusivity estimation from laser flash data and in-plane effective thermal diffusivity measurements done by the temperature oscillation technique. These results were complemented with scanning electron microcopy analysis, surface characterization and the results of dilatometric measurements performed on the TSP bulk specimens obtained from liquid substrate by casting. Complex analysis of the obtained results indicated an isotropic characteristic of the thermal diffusivity of the TSP layer and provided reliable data on all measured thermophysical parameters—they were revealed to be typical for insulators. Further to presenting these data, the paper, in brief, presents the applied investigation procedures.

Enhanced Fluorescence of La3+, Gd3+ doped EuW10 for Temperature sensing performance

AbstractUsing the luminescent properties of temperature‐sensitive paints (TSPs) to monitor temperature is a new type of temperature measurement technology. Herein, Na9[EuW10O36](EuW10) molecules were doped with non‐fluorescent rare earth metal ions (La3+ and Gd3+) to prepare Na9[EuxLa1‐xW10O36](EuxLa1‐xW10) and Na9[EuxGd1‐xW10O36](EuxGd1‐xW10) probe molecules. The doping of non‐fluorescent rare earth metal ions (La3+ and Gd3+) will not change the structure of the parent EuW10. As the relative content of La3+ ions increasing, the luminescent intensity of the EuxLa1‐xW10 series decreases non‐monotonically; as the relative content of Gd3+ ions increasing, the luminescent intensity of EuxGd1‐xW10 series decreases monotonically. Subsequently, EuxLa1‐xW10 and EuxGd1‐xW10 as probe molecules were added into methyl methacrylate (MMA) to obtain temperature‐sensitive paints via polymerization process. In the range of 50 °C to 100 °C, the maximum relative sensitivity of temperature‐sensitive paints EuW10/PMMA, Eu0.9La0.1W10/PMMA and Eu0.7Gd0.3W10/PMMA can reach 1.27 % °C−1, 1.43 % °C−1, 1.42 % °C−1, respectively. This study indicates that the temperature sensing characteristics can be enhanced by doping EuW10 with non‐fluorescent rare earth ions (La3+, Gd3+), which can be promising materials used in more areas with high‐precision temperature measurements.

A robust method for reliable transition detection in temperature-sensitive paint data

This work presents a method to automatically detect boundary-layer transition on thermographic images of aerodynamic surfaces, acquired using the temperature-sensitive paint measurement technique. It is applied to surface temperature distributions mapped onto three-dimensional grids representing the surfaces of interest. Motivated and substantiated by earlier work, the transition location is determined at the location corresponding to the maximal value of the streamwise gradient of the surface temperature in the transitional region. The robustness of the transition detection method was proven for a challenging case with low signal-to-noise ratio and moderate streamwise temperature gradients. Moreover, the present method was validated against transition data obtained using an uncorrelated transition detection method based on the measurement of surface pressure fluctuations. The developed transition detection algorithm enabled the accurate, reliable, and consistent detection of the transition location under a variety of test conditions, including experiments on natural laminar flow airfoil and wing models at high Reynolds numbers and at both subsonic and transonic Mach numbers.

Inner structures of Görtler streaks

The hypersonic boundary layer transition over a concave wall is investigated in a Mach 6.5 quiet wind tunnel using temperature sensitive paint (TSP), CO2-enhanced filtered Rayleigh scattering flow visualization, PCB fast-response pressure sensors, and a high-frequency schlieren technique. The TSP shows that low- and high-temperature streaks are distributed in the spanwise direction. The wavelengths of naturally developing Görtler streaks are randomly distributed, with an average of approximately 7 mm, and change little as the unit Reynolds number increases. More importantly, three-dimensional waves are clearly visualized and quantitatively measured inside the Görtler streaks. This is the first time that the entire evolution of the Görtler instability has been visualized using the Rayleigh-scattering flow visualization in hypersonic flow. The results demonstrate that three-dimensional waves are amplified as a result of the Görtler instability, resulting in a localized high-shear layer around the interface of the three-dimensional waves, which contributes to the formation of hairpin vortices and mushroom-like structures. The three-dimensional waves grow and play an important role in Görtler instability-induced boundary layer transitions.

Experimental Investigation Into the Heat Transfer Mechanism of Oscillating Heat Pipes Using Temperature Sensitive Paint

Abstract Oscillating heat pipes (OHPs) represent a promising passive mechanism for the removal or spreading of heat. While simple to construct, the fluid and thermodynamics of these devices are still poorly understood. There is debate over whether the primary heat transfer mechanism is due to sensible heating of the liquid phase or due to latent heat transfer through phase change. To answer this question, an experimental apparatus was constructed to provide time- and space-resolved temperature and heat transfer data across the face of an operating OHP with HFE-7000 as the working fluid. This experiment utilized temperature sensitive paint (TSP) alongside visual recording of the fluid motion in order to determine the relative latent and sensible contribution to the overall heat transfer. The OHP was tested with input powers ranging from 2.6 W to 10.1 W. It found that latent heat transfer was the dominant heat transfer mechanism, accounting for between 65% and 83% of the total heat transferred in all cases.

Experimental study of crossflow instability in a Mach 6 delta wing flow

In this paper, the traveling crossflow instability in the boundary layer on the windward side of a delta wing is studied. The experiments are carried out in a Mach 6 low-noise wind tunnel, with the angles of attack of the model being 5° and 10°, and the Reynolds number being in a range of 2.43 × 10<sup>6</sup>–14.21 × 10<sup>6</sup> m<sup>–1</sup>. The wall fluctuation pressure is measured by fast-response Kulite pressure transducers. The power spectrum density (PSD) analysis is conducted to obtain the disturbance waves' development process in the boundary layer. The temperature-sensitive paints (TSPs) and nano-tracer based planar laser scattering (NPLS) technique are also used. From the TSP results, the boundary layer transition near the leading edge of the delta wing is smooth and parallel to the leading edge. A peak around 10 kHz in power spectrum density is detected by the fast-response pressure sensor, which may be caused by the traveling crossflow waves. To verify this dominant mode, an NPLS image in the plane of <i>n</i> = 36 mm is obtained. The shapes of vortex structures correspond to the shapes of the crossflow vortices from the numerical simulation. Only when the boundary layer is laminar can the traveling crossflow wave signal be observed from the PSD curves. When the boundary layer is at a transitional or turbulent phase, the low-frequency component is dominant in the PSD curve. With the increase of Reynolds number, the characteristic frequency of the crossflow wave increases, and the wave’s amplitude first increases and then decreases. Moreover, the angle of attack effect is obtained. The increasing of the angle of attack can make the traveling crossflow wave grow faster and saturate, attenuate at the position closer to the leading edge of the delta wing or at a lower Reynolds number. By sensor pairs composed of three Kulite transducers, the phase velocity and the propagation angle of the traveling crossflow wave are investigated. The dimensionless phase velocities of the traveling wave are in ranges of 0.24–0.26 and 0.26–0.32 at 5° and 10° angles of attack, respectively. The propagation angles are at 50°–60° and 40°–55° at the angles of attack of 5° and 10°, respectively. At a larger angle of attack, the traveling wave’s propagation angel is smaller, but the phase velocity is bigger. It may be because the spanwise pressure gradient is higher at a larger angle of attack, and then the crossflow velocity is stronger.

Experimental research on crossflow instability for a hypersonic 4∶1 elliptic cone

In this work, the cross-flow instability on the surface of a blunt elliptical cone with a long-short-axis ratio of 4∶1 is studied experimentally in the Mach 6 hypersonic quiet wind tunnel. Comprehensive use of temperature sensitive paint (TSP) technology, nano-tracer-based planar laser scattering (NPLS) technology and Kulite sensor pressure test to measure the temperature distribution in the cross-flow area on the surface of the model, boundary layer flow structure and model surface pressure are tested. The mechanism of boundary layer transition in the cross-flow control area on the surface of the elliptical cone is studied, and the influence law of incoming flow unit Reynolds number and angle of attack on boundary layer transition is obtained, and some conclusions are obtained below. In the wind tunnel noise mode, the transition of the boundary layer in the cross-flow area between the surface center line and the leading edge of the elliptical cone model with a length-to-short-axis ratio of 4∶1 is controlled by the traveling waves, and no footprint of the steady vortex is found. The characteristic frequency of the traveling wave is about 20 kHz. When the unit Reynolds number of the incoming flow increases, the transition position will be advanced, and the frequency and amplitude of the traveling wave will increase. Within a certain angle of attack, the transition position of the upwind boundary layer is delayed, and the characteristic frequency of the traveling wave does not change much but the energy is weakened. When the angle of attack continues to increase, the transition phenomenon disappears.

Noise reduction for temperature-sensitive paint measurement contaminated by strong background radiation in a high enthalpy hypersonic tunnel

The strong background radiation in high enthalpy hypersonic shock tunnels has posed severe challenges for measurement using luminescent coatings. We proposed a solution for reducing background radiation from time-resolved temperature-sensitive paint (TSP) measurement in a hypersonic flow with Ma = 6.5 and T0 = 3525 K. The TSP was applied on an inlet ramp model, and the images were taken by a high-speed camera at 2 kHz under a modulated excitation. The strong background radiation led to a low signal-to-noise ratio and significant errors for the first half of the 130-ms test duration. Accordingly, three noise reduction methods were developed and evaluated based on temporal reconstruction, spatial reconstruction and robust principal component analysis (RPCA), respectively. The RPCA method showed the best performance that successfully recovered high-quality TSP data for a majority of test duration (t ≥ 40 ms).

Characteristic unsteady pressure field on a civil aircraft wing related to the onset of transonic buffet

Unsteady pressure fields on a civil aircraft wing near the onset of transonic buffet have been investigated experimentally to identify flow unsteadiness that relates to the buffet onset determined by global criteria. An 80%-scale NASA Common Research Model was tested in the JAXA 2 m × 2 m Transonic Wind Tunnel at a Mach number of 0.85 and a chord Reynolds number of 2.27 × 106. The angle of attack was varied in small increments around the buffet onset angle determined by global criteria based on the lift curve and wing-root strain-gauge data. Unsteady pressure fields over the wing were measured using unsteady pressure-sensitive paint (PSP) with temperature-effect correction by temperature-sensitive paint (TSP). Characteristic pressure fluctuations, known as “buffet cells”, were observed under the off-design conditions at a bump Strouhal number of 0.2–0.5. The PSP results showed that the buffet cells arise at the mid-span wing at η ≈ 0.45, where a strong shock wave causes an initial boundary-layer separation. Phase shift distributions indicated that a pressure perturbation propagates from the inboard wing toward the outboard wing. The convection velocity and spanwise wavelength were approximately 0.5U∞ and 1.3cMAC, respectively. The angle of attack at which buffet cells first appear was found to be approximately equal to the buffet onset determined by the global criteria, indicating that the occurrence of the buffet cells is deeply related to the buffet onset for the present wing geometry.

Feasibility of skin-friction field measurements in a transonic wind tunnel using a global luminescent oil film

The feasibility of skin-friction field measurements using the global luminescent oil-film skin-friction field estimation method was evaluated for a challenging case of a supercritical airfoil model under transonic wind-tunnel conditions (freestream Mach number of 0.72) at a high Reynolds number (10 million, based on the model chord length). The oil-film thickness and skin-friction coefficient distributions were estimated over the airfoil model upper surface for a range of angles of attack (from -0.4^{circ } to 2.0^{circ }), thus enabling the study of different boundary-layer stability situations with laminar–turbulent transition, including cases with shock-wave/boundary-layer interaction. Conventional pressure measurements on the surface and in the wake of the model as well as Schlieren flow visualizations were conducted to support the oil-film based investigations. In the laminar-flow regions, the oil-film thickness could be generally kept below the critical limit of roughness that would induce premature boundary-layer transition. The skin friction in this region could be estimated with a moderate confidence level, as confirmed for portions of the chord by the reasonable agreement with numerical data obtained via laminar boundary-layer computations. Moreover, the location of transition onset was evaluated from the skin-friction estimations with relatively low uncertainty, thus enabling the examination of the transition location evolution with varying angle of attack. The estimated locations of transition onset were shown to be in general agreement with reference transition locations measured via temperature-sensitive paint. On the other hand, the oil-film thickness in the turbulent-flow regions was larger than the height of the viscous sublayer, which led to an hydraulically rough surface with increased skin friction, as compared to the clean configuration. For this reason, quantitative skin-friction estimations were not feasible in the turbulent-flow regions. The global effects of the oil-film setup on the flow around the airfoil were evaluated from the estimations of the aerodynamic coefficients. In particular, it was shown that the presence of the specific base coat used for the application of the oil film already induced a significant increase in airfoil drag, as compared to the clean configuration, whereas a thin oil film led to negligible or small additional increases in drag. Based on the present observations, considerations for the further improvement of the global luminescent oil-film skin-friction field estimation method in transonic flow experiments at high Reynolds numbers are elucidated.Graphic abstract

A luminescent temperature sensor based on rhodamine B on a polymer-ceramic substrate for aerothermal measurements

Temperature-sensitive paint (TSP) is a type of luminescent temperature sensor that can provide surface temperature measurements on the surface of a test article. These surface temperature measurements can be used to determine heat flux and visualize boundary layer transition on an aerodynamic body. This paper presents the static response characteristics of a TSP based on rhodamine B (RhB) applied on a polymer-ceramic supporting matrix. Six solvents with varying polarity indices were used to apply the RhB to the sensors, which were then evaluated in terms of luminescent signal level and temperature sensitivity. The results are presented and also compared to previous results for RhB on an anodized aluminum matrix. The temperature range was between 150 K and 365 K. The study confirms that the influence of the solvent on the sensor’s final characteristics is significant, and shows temperature sensitivities as high as −4.8% K−1 at 150 K when dichloromethane is used as the application solvent.

Experimental investigation of nose-tip bluntness effects on the hypersonic crossflow instability over a cone

The nose-tip bluntness effects on the hypersonic crossflow instability were experimentally investigated over a 7∘ half-angle straight cone at 6∘ angle of attack. The experiments were conducted in a Mach 6 wind tunnel at Reynolds number of 7.94×106 m−1 using temperature-sensitive paint (TSP) and fast-response pressure transducers. The wavelet transform was applied to obtain the most amplified wavenumber of the stationary crossflow vortices. The stationary crossflow vortices with wavenumber of 40–60 and the traveling crossflow waves with characteristic frequencies of 9–20 kHz have been detected. TSP results indicate that the bluntness has little influence on the boundary layer transition location with a small nose-tip radius (smaller than 1.5 mm), while the transition front moves backward obviously with a larger nose-tip radius (larger than 2.5 mm). The developments of two types of crossflow instability, the stationary mode and the traveling mode, are both not sensitive to the bluntness when the nose-tip radius is below 1.5 mm. And the growth and saturation of the two types crossflow mode are both inhibited under larger bluntness. Besides, possible explanations for how the bluntness affects the transition location and crossflow instability is discussed.

Low-Reynolds-Number Flowfield of Wing with Control Surface in Propeller Slipstream

An experimental study is conducted on a NACA 0012 airfoil with a control surface to investigate the influence of a propeller slipstream under a low-Reynolds-number condition. To investigate the influence of the propeller slipstream on a wing surface, the flowfield on the wing surface is visualized using a temperature-sensitive paint when the chord-based Reynolds number is equal to . The control surface performance improves in the installed-propeller condition because the propeller slipstream suppresses the laminar separation bubble. Finally, we suggest that the propeller should be installed ahead of the control surface to benefit from the effect of the propeller slipstream under the low-Reynolds-number flight condition.

A swept fin-induced flow field with different height mounting gaps

In order to apply the air fin successfully and ensure the maneuverability of hypersonic vehicle, a key problem to be studied urgently is the heat flux brought by the fin mounting gap. The appearance of mounting gap and fin shaft can induce many complex flow structures which need more attentions to be investigated. Under Ma 6, Nano-tracer-based Planar Laser Scattering (NPLS) and Temperature Sensitive Paints (TSP) were applied to visualize and measure transient flow structures and heat flux distribution of a swept fin-induced flow field with different height mounting gaps. Complementarily, Reynolds-averaged N-S equations were solved with k-ω SST turbulent model. The heat flux distribution results of numerical simulation and TSP observed the change of high heat flux region with different mounting gap, both in position and magnitude. The streamlines based on Computational Fluid Dynamics (CFD) and flow visualization results obtained by NPLS revealed the cause of high heat flux region. The high heat flux region in this flow field is mainly related to the reattachment of vortex and flow stagnation. The increase of gap height can lead to stronger gap overflow and shaft-induced horseshoe vortex, which are source of the high heat flux around the fin. The case with the highest mounting gap (4 mm) en-counters the most severe aerodynamic heating, both on the surface of fin and plate. Thus, under the premise of ensuring the flexibility of the fin, the gap should be set as small as possible.

Boundary Layer Transition Preflight Experiments in a Mach-6 Quiet Tunnel

Experiments on a subscale model of the U.S. Air Force Office of Scientific Research Boundary Layer Transition flight experiment were made in the Boeing/Air Force Office of Scientific Research Mach-6 Quiet Tunnel. Infrared thermography and temperature-sensitive paint were used to measure the surface temperature and heat transfer and provide images of streamwise streaks with the model at 0, 2, and 4 deg angle of attack. Variations in heating patterns were measured for yaw and roll angles. All smooth body cases remained laminar under quiet flow, but symmetric turbulent wedges were seen in noisy flow. Discrete surface pressure sensors measured possible instabilities on the model surface at locations that correspond to computed instability growth. It was discovered that the geometry was sensitive to forward- and rearward-facing steps near the swept leading edges. Controlled forward- and rearward-facing steps were created and tested. No transition was measured under quiet flow for any rearward-facing step at 0 deg angle of attack. Early transition was measured for the forward-facing steps, and correlations were developed for scaling step effects to the flight vehicle.

Apparent temperature in temperature-sensitive paint measurement and its effect on surface heat flux determination for hypersonic flows

Non-zero temperature gradients in temperature-sensitive paint (TSP) cause the apparent temperature (the temperature measured by the TSP) to correspond to the temperature somewhere inside the TSP, which does not equal the top surface temperature. Treating the apparent temperature as the average temperature across the TSP layer is not always accurate, especially when there is a large temperature gradient in the TSP. In this paper, the apparent temperature is theoretically derived by integrating the luminescent intensity across the TSP layer and numerically determined using a Monte Carlo ray-tracing method. The results of a simulation show that the difference between the apparent and average temperatures increases with the temperature gradient in the TSP layer, which leads to a non-negligible error when determining the surface heat flux based on the average temperature. A modified Levenberg–Marquardt algorithm is used to accurately recover the surface heat flux based on the apparent temperature for typical optical conditions in the measurement system. The results are compared with those obtained using the average temperature of the TSP. The effects of heat flux, TSP thickness and base material on the errors in the determination of the heat flux are investigated in detail.

Flow Structure and Heat Flux Distribution of a Backswept Fin Induced Flow Field at M = 6

Transient flow structures and surface heat flux distribution of the interfered flow field induced by a backswept fin were obtained by Nano-tracer-based Planar Laser Scattering and temperature sensitive paints under M = 6. Meanwhile, the Reynolds-averaged Navier–Stokes equations with k-ω Shear-Stress Transport turbulence model were solved to simulate the flow field. The numerical results were compared and analyzed with the experimental results. Typical flow structures in interference areas of the swept fin were observed in experiment, including the detached bow shock wave, separation zone, thin boundary layer around the fin and horseshoe vortex, etc. The numerical results were in good agreement with flow visualization images. For heat flux distribution, the experiment observed the high heat flux region near the side of the fin leading edge caused by separated flow and the region with heat flux increase caused by bow shock. The numerical results of the heat flux caused by bow shock were pretty good, but the results of the high heat flux region near leading edge due to reattachment had much difference with the experimental results, which still need further improvement.

Experimental study on shock and turbulent boundary layer interactions under different incident shock waves

In this paper, the interactions between shock wave and the turbulent boundary layer of an incoming wall were experimentally studied. The experiments were performed in a Mach 3.4 supersonic low-noise wind tunnel at the unit Reynolds number of 6.30×106 m−1. Under different incident shock wave conditions, fine flow field structure, wall pressure, and wall temperature distribution were determined. Relationship was also analyzed. Instantaneous fine structures of typical flow fields with shock wave and turbulent boundary layer interactions were investigated using the nano-tracer planar laser scattering technique, and the transient flow images were analyzed statistically. The oscillation position of the induced shock wave satisfies a normal distribution. Simultaneously, the wall pressure distribution under different incident shockwave conditions of the interaction region was determined using the flush air data sensing system, and the relationship between flow the structure of the separation region and the wall pressure distribution with different incident shock wave intensities is analyzed. Using a temperature sensitive paint technology, the wall temperature distribution under different incident shockwave conditions of the interaction region model was determined. Also, the relationship between the strip structure caused by Gortler vortices and the incident shock wave intensity in the shockwave/turbulent boundary layer interaction was analyzed.