Adiabatic Wall Temperature Research Articles

Comprehensive Experimental Assessment Of Unsteady Pressure And Heat Flux In A Small-Core Turbine Over-Tip Shroud

Abstract For novel high-speed small core turbines, with tip clearance below 0.5mm, even small blade-to-blade tip clearance variation is significant. The assessment of these complex flows is pertinent to the design of the next generation of small-core turbines. This manuscript provides a thorough experimental analysis of the shroud?s unsteady heat flux and static pressure in a small-core squealer-tip blade turbine. Atomic Layer Thermopile sensors and fast response pressure transducers were used to perform high-frequency acquisition at 2MHz around the 51% axial blade chord on the shroud. Measurements were taken at engine-representative conditions at several operational conditions and tip clearances. The signals were phase-locked averaged over the revolution period and synchronized to identify individual blade and row signatures. The linear relationship between the rotational tip Reynolds and the static pressure ratio across the blade tip reveals the transition point to reverse over-tip flows. Total heat flux is decomposed into different steady and unsteady heat flux contributions. It is demonstrated that the adiabatic wall temperature governs the unsteady heat flux and contributes to one third of the total surface heat flux. A linear trend was observed between the unsteady heat flux and the tip clearance measured at the pressure and suction side rims. Similar trends were observed between the local heat flux and the pressure ratio across the tip. A comparison with CFD predictions highlights some limitations on resolving the detached and secondary flows, evidencing the necessity of complementary experimental data.

Film hole layout optimization for multi-row film cooling plate in continuous parameter space under non-uniform inlet temperature distribution

Film cooling is an advanced external cooling technology for protecting gas turbine blades from excessive temperatures. To counteract non-uniform heat loads caused by hot streak and swirl effects at the combustion chamber outlet, fine design of film hole positions is necessary. However, the complexity due to the numerous film holes and their positional parameters presents challenges in the independent optimization of film hole positions in a continuous parameter space. This study proposed an optimization method grounded in the divide and conquer approach to address optimization for film hole layout. The film hole layout was decomposed into multiple single rows and optimized systematically by iteratively refining the single-row film hole layout along the streamwise direction using the expectation maximization algorithm. Optimized layouts for five different inlet temperature distributions were obtained using the proposed method and adiabatic wall temperature distributions were compared and analyzed for optimized and staggered layouts. Additionally, the cooling performance of the optimized layout was evaluated against the staggered layout under conjugate heat transfer conditions and the robustness of the optimized layout under various blowing ratios and peak temperature positions was tested. The results demonstrated that the proposed method can optimize the positions of 52 film holes within 0.3h, and it was generalized to various inlet temperature distributions. By enhancing lateral interactions among downstream film jets, the lifting effect of kidney vortex in the optimized layout was weakened, bringing the film jets closer to the wall and significantly increasing coolant coverage during streamwise development. Comparative analysis reveals the superior cooling performance of the optimized layout at the blowing ratio of 1.0, evidenced by a 15.1% reduction in total input heat transfer rate and a 12.3% enhancement in overall cooling efficiency relative to the staggered layout. Under conditions with blowing ratios ranging from 0.5 to 1.5 and peak temperature position deviations, the average wall temperature of the optimized layout consistently remained lower than the staggered layout, indicating the robustness of the optimized layout to variations in blowing ratio and hot streak position.

Determination of Local Heat Transfer Coefficients and Friction Factors at Variable Temperature and Velocity Boundary Conditions for Complex Flows

Transient conjugate heat transfer measurements under varying temperature and velocity inlet boundary conditions at incompressible flow conditions were performed for flat plate and ribbed channel geometries. Therefrom, local adiabatic wall temperatures and heat transfer coefficients were determined. The data were analyzed using typical heat transfer correlations, e.g., Nu=CRemPrn, determining the local distributions of C and m. It is shown that they are closely linked. A relationship lnC=A−mB is observed, with A and B as modeling parameters. They could be related to parameters in log-law or power-law representations for turbulent boundary layer flows. The parameter m is shown to have a close link to local pressure gradients and, therewith, near wall streamlines as well as friction factor distributions. A normalization of the C parameter allows one to derive a Reynolds analogy factor and, therefrom, local wall shear stresses.

A SUPERPOSITION TECHNIQUE FOR PREDICTING OVERALL EFFECTIVENESS WITH MULTIPLE SOURCES OF INDIVIDUALLY CHARACTERIZED INTERNAL AND EXTERNAL COOLANT FLOWS

Abstract Gas turbine engines rely on intricate cooling schemes to protect turbine components from the high temperature freestream gas. Determining the optimal placement of film cooling holes while remaining mindful of the consumption of coolant air is therefore a design challenge. For several decades a method of superposition has been in use that provides a reasonable prediction of the adiabatic effectiveness in a region influenced by multiple rows of film cooling holes acting together if the adiabatic effectiveness distributions resulting from individual rows are already known. This classical superposition technique has been used to provide a first cut at determining where coolant holes might be placed. A shortcoming of the method is that it strictly applies only to prediction of adiabatic effectiveness which is indicative of the adiabatic wall temperature, not the actual surface temperature. The overall effectiveness, which is indicative of the actual surface temperature is somewhat more complicated as it is influenced not only by the external cooling, but also internal cooling. In the present work, a method of superposition of overall effectiveness data is proposed, allowing prediction of the actual surface temperature distribution on a component. Experimental results show that it is possible to use knowledge of the overall effectiveness distributions resulting from individual internal cooling and associated film cooling holes acting alone to determine how they are likely going to act when combined. This technique shows promise for a turbine designer to use superposition for actual surface temperature prediction and therefore higher quality initial cooling designs.

Experimental Investigation of Tip Design Effects on the Unsteady Aerodynamics and Heat Transfer of a High-Speed Turbine

Abstract While modern engine manufacturers devote significant efforts to the development of reliable and efficient machines, the introduction of novel, optimized components in the hot gas path represents a risky opportunity. Accurate experimental and numerical data are critical to assess the impact of new technologies on the harsh engine environment. The present study addresses the impact of a selection of high-performance rotor blade tips on the aerodynamic and heat flux field of a high-pressure turbine (HPT) stage. A combined numerical and experimental approach is employed to characterize the interaction of the tip leakage flow with the rotor secondary flows and the casing heat transfer mechanisms for each individual tip geometry. The turbine stage is tested at engine-scaled conditions in the rotating turbine facility of the von Karman Institute. In the present study, the turbine rotor is operated in rainbow configuration to allow the simultaneous testing of multiple blade tip geometries. Reynolds-averaged Navier–Stokes (RANS) simulations are employed to predict the aerodynamic and thermal fields of the individual profiles using test-calibrated boundary conditions. Isothermal steady computations are performed at different wall temperatures to compute the adiabatic wall temperature and the heat transfer convective coefficient. Low-order models are used to represent the over-tip thermal field and the driving heat transfer mechanisms. The time-resolved outlet flow is characterized using a vortex tracking technique and high-frequency aerodynamic measurements to identify the rotor secondary flow structures.

Heat transfer measurements on a flat plate and cylinders at high velocity conditions: Comparison of calculation models with infrared thermography

In this study, various heat transfer measurement models using infrared (IR) thermography were applied to a flat plate and cylinders to examine the accuracy and suitability of the measurements in a transonic blow-down test facility recently developed at Korea Aerospace University. Experiments were performed under three inlet velocity conditions for the flat plate, and for the cylinders, two diameters, two turbulence intensity levels, and five inlet velocity conditions. The surface temperature distribution during the experiment was acquired using an IR camera, and the surface heat transfer coefficient (HTC) was calculated using four HTC calculation models based on the one-dimensional semi-infinite (1D SI) solid assumption: step change (STEP), Duhamel's superposition (DUH), 1D finite volumetric model (1D FVM), and linear regression model (LRM). The LRM showed minimum deviation with reference correlation for the flat plate heat transfer measurements, while STEP, DUH, and 1D FVM showed increased deviations at higher mainstream velocity conditions. The measurements by LRM at the cylinder stagnation location showed the most comparable value with the reference correlation, whereas STEP and FVM showed about 10% deviation from the reference. It is concluded that the LRM is the most suitable heat transfer calculation model for high velocity tests because the model excludes the ramping data during the transient tests, and it can also obtain the local adiabatic wall temperature experimentally.

Infrared temperature measurements on fast moving targets: A novel calibration approach

In this study, an infrared system is developed for accurate measurements of surface temperature and heat transfer on fast moving targets. The system was designed for the Oxford Turbine Research Facility, a world-leading experimental facility delivering highly engine representative, scalable heat transfer results for aerospace research. Infrared thermography is employed to acquire temperature maps of high-pressure turbine blades, allowing assessment of surface thermal conditions including heat transfer coefficient, adiabatic wall temperature, Nusselt number, cooling effectiveness, and metal effectiveness.Achieving accurate infrared thermography measurements in rotating turbomachinery experimental conditions is arduous due to reflections from the surroundings, low emissivity of metallic parts, and motion blur resulting from high speed. To overcome these challenges, calibration procedures were developed against a traceable standard using a bespoke steady experimental facility. A method to determine the reflected temperature from surroundings was also validated. Correction for all measurement disturbances is demonstrated to within the accuracy of the primary measurement thermocouple.Finally, the developed calibration method was validated on a fast-moving rotating geometry demonstrating accurate correction for all measurement disturbances, without the need for an in situ calibration. A detailed uncertainty analysis for each calibration step is also presented.

Dimensionless analysis of flow and heat transfer characteristics in a high-speed rotor–stator disk cavity based on similarity criteria

Turbine disk cavity system is of great importance for improving the rotating turbine blades cooling efficiency and preventing gas invasion. This study focuses on turbine disk cavity characteristics at high-speed rotational conditions due to complicated flow and heat transfer mechanisms. Then, a novelty method is proposed to reveal multiple similarity criterion numbers of rotor–stator disk cavity between actual engine and experimental case conditions, based on the dimensionless analysis. Especially, a new evaluation standard including the dimensionless adiabatic wall temperature and dimensionless heat flux is put forward and proved to be equally important with the Nusselt number. Results show that the dimensionless heat flux in the turbine disk cavity mainly depends on Reynolds number and Mach number, while the dimensionless adiabatic wall temperature is dominated only by the same Mach number. When only Reynolds numbers or Mach numbers are considered, the relative error can reach 14.5% on the similarity of dimensionless heat flux. Noteworthily, the effect of thermal boundary of turbine disk cavity should not be ignored in the rotor–stator disk cavity, which will bring 28.8% deviation to the similarity of the Nusselt number. Therefore, this study can provide a theoretical method and engineering application to obtain flow and heat transfer characteristics of actual engine turbine disk cavity under experiment case with normal temperature conditions.

Understanding Thermal Unsteadiness in Engine Representative Flows and Improved Methodologies for Derived Heat Transfer Calculations Using Thin-Film Gauges

Abstract The Oxford Turbine Research Facility (OTRF) is a high-speed rotating transient test facility, which allows unsteady aerodynamic and heat transfer measurements at engine representative conditions. In addition, a variety of inlet temperature profiles can be simulated in the rig including radial distortion, circumferential distortion, and swirl. However, the engine representative flows cause complications in the processing of heat transfer data. The unsteadiness in temperature data was found to significantly rise as temperature distortions were introduced in the nozzle guide vane (NGV) inlet profile, to model a lean-burn combustor exit. Using the NGV inlet temperature profile survey data, the thermal unsteadiness has been quantified and compared with a uniform inlet. The experiments with a radially varying NGV inlet temperature profile showed up to nine times higher thermal unsteadiness, compared to the uniform inlet. The second part of the paper is a continuation of the work presented in a previous paper by Singh et al. and describes improved methodologies for derived heat transfer calculations using thin-film gauges. In addition, the uncertainty associated with the derived heat transfer parameters, such as the heat transfer coefficient and adiabatic wall temperature has been quantified. The refined processing techniques have been demonstrated on casing heat transfer measurements, acquired in the OTRF with two inlet temperature profiles.

Adiabatic wall temperature in the supersonic flow of moist air with spontaneous condensation

The expansion of moist air in supersonic nozzles is accompanied by spontaneous condensation of water vapor due to a sharp decrease in flow temperature and pressure. However, in the near-wall boundary layer due to the viscous dissipation, moist air temperature decreases slightly in comparison with initial one. This leads to the fact that the vapor remains overheated in the near-wall region and condensation does not spread to its entire depth. Droplets formed in the core flow as a result of condensation can penetrate into the near-wall region due to turbulent diffusion and evaporate due to higher temperature in the near-wall region. This process causes cooling of the near-wall gas layers and, consequently, a decrease in the adiabatic wall temperature. A decrease in the adiabatic wall temperature relative to the stagnation temperature leads to a decrease in aerodynamic heating and, consequently, in the heat flux from the gas to the body. In this paper, measurements of the adiabatic temperature of the supersonic nozzle wall for moist air flow with condensation are presented. Temperature measurements were made using an infrared thermal imager. Initial moisture content varied in the range of 1.5–20.2 g/(kg. dry air) and initial relative humidity in the range 20–90%. It is shown that in the region behind the zone of spontaneous condensation (“condensation shock”), adiabatic wall temperature depends on the initial moisture content and can be either greater or less than the value of a single-phase (dry air) flow with identical initial flow parameters. Estimates of the flow parameters behind the condensation region are obtained using one-dimensional diabatic flow equations and experimental static pressure distributions along the nozzle length. The correlation between the initial moisture content, the average droplet size and the temperature recovery factor is shown.

Asymptotic theory of Mack-mode receptivity in hypersonic boundary layers due to interaction of a heating/cooling source and a freestream sound wave

In this paper, we study the local receptivity of the inviscid Mack modes in hypersonic boundary layers induced by the interaction between a surface heating or cooling source (HCS) and a freestream acoustic wave. The asymptotic analysis reveals that among the three distinguished layers, i.e. the main, wall and Stokes layers, the leading-order receptivity is attributed to the interaction of the HCS-induced mean-flow distortion and the acoustic signature in the wall layer; the second-order contribution appears in the Stokes layer; the third-order contribution appears in both the main and wall layers. Interestingly, at a moderate Reynolds number, the third-order contribution to the receptivity efficiency may be quantitatively greater than the second-order one, but this does not lead to breakdown of this asymptotic theory. Assuming the HCS intensity to be sufficiently weak, the asymptotic predictions are made for four representative cases involving different Mach numbers and wall temperatures, which are compared with the results obtained by the finite-Reynolds-number theory based on either the extended compressible Orr–Sommerfeld equations or the harmonic linearised Navier–Stokes (HLNS) calculations. Taking into account the first three orders of the receptivity efficiency, the asymptotic predictions are confirmed to be sufficiently accurate even when the Reynolds number is a few thousands, and the agreement with the finite-Reynolds-number calculations is better when the wall temperature of the base flow approaches the adiabatic wall temperature. The HLNS calculations are also conducted for moderate HCS intensities. It is found that the nonlinearity does not affect the receptivity coefficient much even when the temperature distortion of the HCS reaches $80\,\%$ of the temperature at the wall.

Experimental Investigation of a High-Speed Turbine With Rainbow Rotor and Rim Seal Purge Flow

AbstractThe present paper addresses the experimental and numerical study of the unsteady flow established in a high-pressure turbine stage with rim seal purge. The HPT test section, operated at engine-relevant flow conditions in the high-speed turbine rig of the von Karman Institute, is heavily instrumented for high-resolution, high-bandwidth aerothermal measurements. A rainbow rotor setup allows the simultaneous testing of six different sectors, each hosting a specific tip and platform geometry optimized for enhanced aerodynamic performance. This paper focuses on the flow over the baseline sector, equipped with an axisymmetric hub platform and a squealer tip, with a purge flowrate matching 1.74% and 1% of the stage mass flow. The numerical study relies on Reynolds-averaged Navier–Stokes (RANS) computations with test-calibrated boundary conditions. Unsteady pressure and heat transfer measurements are performed at the rotor shroud. The maximum heat transfer is achieved along the front pressure side rim, whereas the squealer cavity generates a region of uniformly low static pressure and heat flux. The experimental adiabatic wall temperature is derived to quantify the thermal contribution to the global heat flux, demonstrating an increase in over-tip gas temperature up to 1.2 T01 above the pressure side rim. A Euler-based model is proposed to evaluate the temperature-driving work-exchange mechanism in the tip gap. The peak in casing heat transfer coefficient (750 W/m2K) is found at the tip leading edge. Time-resolved measurements of outlet total pressure, Mach number, and flow angle confirm the predicted phase and intensity of the tip leakage and upper passage vortex in the near-casing region. At 20% hr, the total pressure minimum is highly underpredicted by the RANS, indicating an inaccurate modeling of the interaction between rim seal purge and main flow. The measured impact of the purge flow variation is more significant than predicted by the RANS computations, with a negative offset of about 0.5% P01 in the lower 50% hr at lower purge flowrate.

Heat Transfer Coefficient and Adiabatic Wall Temperature Measurements on High-Pressure Turbine Nozzle Guide Vanes With Representative Inlet Swirl and Temperature Distortions

Abstract Combustor exit conditions in modern gas turbines are generally characterized by significant temperature distortions and swirl degree, which in turn is responsible for very high turbulence intensities. These distortions have become particularly important with the introduction of lean combustion, as a mean to control NOx pollutant emissions. For this reason, combustor–turbine interaction studies have recently gained a lot of importance. Past studies have focused on the description of the effects of turbulence, swirl degree, and temperature distortions on the behavior of the high-pressure stages of the turbine, both considering them as separated aspects and accounting for their combined impact. Aspects like pressure losses, hot streaks migration, and film-cooling behavior have been widely investigated. Even if some studies have focused on the characterization of the heat transfer coefficient (HTC) on the nozzle guide vane external surface, none of them have addressed this aspect from a purely experimental point of view. Indeed, when inlet conditions are characterized by both swirl and temperature distortions, they represent a severe challenge for the commonly adopted measurement techniques. The work presented in this paper was carried out on a non-reactive, annular, three-sector test rig made by a non-reactive combustor simulator and a nozzle guide vane cascade; it is able to create a representative combustor outflow, characterized by all the flow characteristics described before. A novel experimental approach, which was developed in a previous work, was exploited to experimentally retrieve the heat transfer coefficient and the adiabatic wall temperature distributions on a non-cooled nozzle guide vane. Temperature measurements on the cascade inlet and outlet planes were also used to provide boundary conditions and achieve a better understanding of the investigated phenomena. The results allowed to evidence the effect of the inlet swirl on the heat transfer coefficient distribution, as well as the evolution of the temperature distribution on the vane surface moving through the cascade, constituting the first attempt to evaluate these aspects from a purely experimental point of view.

Temperature recovery factor for gaseous nitrogen flow in a microtube

For gas flow over a flat plate, the temperature recovery factor is defined as a fraction of free-stream total temperature rise recovered at the wall or is the ratio of the actual temperature rise at the wall to the maximum possible temperature rise in the free-stream. Its value depends on Prandtl number and is independent of Mach number. The gas velocity and temperature are determined by the temperature recovery factor using the adiabatic wall temperature. It is an important parameter to be used to evaluate the gas velocity and temperature indirectly. In the present study, for internal flow, a methodology was introduced to estimate the temperature recovery factor numerically based on the total and bulk temperature for both laminar and turbulent flow regions in a microtube. The numerical simulations are based on the arbitrary Lagrangian-Eulerian method. The compressible momentum and energy equations for an ideal gas were solved to obtain the temperature recovery factor. To further validate the temperature recovery factor from numerical results, experiments were conducted using stainless steel microtubes. Although it is not a direct comparison, the gas bulk temperature and Mach number that were determined by the temperature recovery factor were compared with the experimental results and they were in good agreement.

Sensitivity of Adiabatic Cooling Effectiveness to Mainstream and Coolant Temperatures in Transonic Flow Over an Idealized Blade Tip Model

Abstract Film cooling, as a key technology to ensure turbine survival in new generation gas turbines, has been studied profusely in subsonic flows. But in transonic flow, the determination of adiabatic cooling effectiveness faces a dilemma in the state of the art. Specifically, derivation of reference temperature, or local recovery temperature, in adiabatic effectiveness is disputable. Some researchers designate it to be the adiabatic wall temperature for the uncooled model (linear regression method (LRM)), but others calculate it from an iterative procedure based on a pair of cooling tests (dual linear regression technique (DLRT)). As the first of the kind effort to explore this dilemma, this article carried out transient thermal measurements by infrared thermography, for transonic flow over an idealized blade tip model. Heat transfer experiments were conducted for the uncooled and cooled cases, at two mainstream temperatures of 340 K and 325 K and two coolant temperatures of 276 K and 287 K. Data from these six experimental groups were processed by LRM and DLRT, respectively, to obtain heat transfer coefficient and adiabatic effectiveness, whose sensitivity to mainstream and coolant temperatures is tested and compared. It is found that the heat transfer coefficient is basically insensitive to temperature boundary conditions and data reduction methods, as expected. However, for adiabatic effectiveness, LRM results are sensitive to the 11 K decrease of coolant temperature in areas confined to the upstream of cooling injection, and much less so to the 15 K rise in mainstream temperature. DLRT result, derived from the test pair with two coolant temperatures, reduces globally and conspicuously with 15 K increase in mainstream temperature. Furthermore, adiabatic effectiveness obtained by LRM is qualitatively different from that by DLRT, which is mainly attributed to the large discrepancy in reference temperature between the two methods.

Effects of fuel conversion ratio on cooling and drag reduction performance for supersonic film using gaseous hydrocarbon fuel

Film cooling especially combined with regenerative cooling is an effective solution to meet the heat protection and the drag reduction demands for hydrocarbon-fueled scramjet engines. Typically, fuel is the only available coolant to organize the active cooling for hypersonic vehicles, and the coolant used in the film cooling is extracted from the regenerative cooling channels; however, the fuel pre-cracking in the regenerative cooling process must be considered as it can change the initial composition and energy level of the coolant in the film cooling. Numerical simulations based on the RANS method are conducted to evaluate the effects of film pre-cracking on its cooling and drag reduction performance, by changing the fuel conversion ratio from 0 to 40 %. Results show that the influence of film inlet cracking on its drag reduction performance is limited, and the relative change of the wall friction under different conversion ratios is less than 5 %. However, the film inlet cracking has a notable negative influence on its thermal protection performance. Specifically, for every 10 % increase in the film conversion ratio, the average adiabatic wall temperature will rise by about 20 K. Further analysis reveals that although the total amount of the fuel chemical heat sink is comparable to its physical heat sink, only around 25 % of the film chemical heat sink can be effectively used to enhance the fuel film cooling performance. The loss of the effective chemical heat sink during the pre-cracking process is the main reason for the decrease of film cooling performance as the film inlet conversion ratio increases.

Compressible gas-droplet flow and heat transfer behind a condensation shock in an expanding channel

A two-fluid mathematical model of a supersonic gas-droplet mixture flow with account of droplet growth due to vapor condensation behind a condensation shock in a plane expanding channel is constructed. The flow structure in both the inviscid core and the two-phase boundary layer assumed to be laminar is studied numerically. The range of flow parameters is considered in which the spontaneous condensation zone (“condensation shock”) is fairly narrow, and just behind this shock thermal parameters of the phases almost reach thermodynamic equilibrium. Downstream, the droplet radius continues to grow due to the flow expansion and further vapor condensation. Of primary concern is the flow region where the droplets become sufficiently inertial to travel with a velocity slip, and hence to model adequately the velocity and temperature fields in both phases a two-fluid model is required. In addition to the Stokes drag, the Saffman lifting force exerted on the droplets in the boundary layer is taken into account. The Saffman force results in droplet deposition and the formation of a thin liquid film on the channel walls. A parametric numerical analysis of the two-phase flow structure in the inviscid flow region and in the near-wall boundary layer is performed, and the effects of phase transitions (droplet condensation and evaporation, as well as the film formation) on the reduction of the adiabatic channel wall temperature are analyzed.

The Dufour Effect in Film Cooling Experiments With Foreign Gases

Abstract In typical film cooling experiments, the adiabatic wall temperature may be determined from surface temperature measurements on a low thermal conductivity model in a low-temperature wind tunnel. In such experiments, it is generally accepted that the adiabatic wall temperature must be bounded between the coolant temperature and the freestream recovery temperature as they represent the lowest and highest temperature introduced into the experiment. Many studies have utilized foreign gas coolants to alter the coolant properties such as density and specific heat to more appropriately simulate engine representative flows. In this paper, we show that the often ignored Dufour effect can alter the thermal physics in such an experiment from those relevant to the engine environment that we generally wish to simulate. The Dufour effect is an off-diagonal coupling of heat and mass transfer that can induce temperature gradients even in what would otherwise be isothermal experiments. These temperature gradients can result in significant errors in the calibration of various experimental techniques, as well as lead to results that at first glance may appear non-physical such as adiabatic effectiveness values not bounded by zero and one. This work explores Dufour effect induced temperature separation on two common cooling flow schemes, a leading edge with compound injection through a cylindrical cooling hole, and a flat plate with axial injection through a 7–7–7-shaped cooling hole. Air, argon, carbon dioxide, helium, and nitrogen coolant were utilized due to their usage in recent film cooling studies.

The effect of the initial swirl of the supersonic flow of humid air on the adiabatic wall temperature

The paper presents the results of measuring the adiabatic wall temperature of an axisymmetric channel during acceleration of the moist air flow in it to supersonic speed. The initial swirl was imparted to the flow (swirling parameter S = 0.5, 1.0, 2.5). The relative initial humidity (RH) of the flow varied in the range of 10 ÷ 90%. When the flow was accelerated to supersonic speeds, part of the moisture condensed, which influenced the wall temperature. It is shown experimentally that with an increase in the initial moisture content of the flow to certain values, the distribution of the wall temperature for a flow without initial swirl (S = 0) and with swirl with S = 0.5, 1.0 practically coincide. However, from a certain value of RH, the wall temperature in the case of a swirling flow decreases in comparison with a flow without swirling. The maximum decrease in the wall temperature was achieved at RH = 90%. An increase in the initial swirl to S = 2.5 led to a greater decrease in the wall temperature, while the mass air flow through the channel decreased by 26% at an identical pressure drop.

Measurement of surface heat transfer caused by interaction of sonic jet and supersonic crossflow near injection hole

This paper investigates the surface heat transfer caused by interaction of a jet and a supersonic crossflow near the jet injection hole. A sonic jet with different momentum ratios (J=0.510, 1.018, 1.477) was injected perpendicularly into a crossflow with a Mach number of 3.0 in a supersonic wind tunnel. Surface temperature through time measured by infrared thermography was used to deduce surface heat flux. In addition, heat transfer coefficients and adiabatic wall temperatures were derived from time histories of surface heat flux and temperature. In order to consider an effect of conduction from the inner hole surface, a three-dimensional energy conservation is considered in the deduction process of the heat flux. As a result, the characteristics of the heat transfer near the hole and the change in the heat transfer with momentum ratios are presented. The separation vortex and recirculation vortex are found to be dominant flow features in terms of the augmentation of the heat transfer. The maximum heat transfer is observed at the immediate vicinity of the hole due to the flow oscillation from a jet-mixing layer. This oscillation resulted in a 390% of augmentation of the heat transfer near the hole compared to the freestream even at the lowest momentum ratio. Also, the augmentation near the hole is more susceptible to change of momentum ratio compared to the augmentation on the overall interaction area.