Surface Heat Transfer Measurements Research Articles

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.

Transpiration-Cooling Heat Transfer Experiments in Laminar and Turbulent Hypersonic Flows

The design of a transpiration-cooled system requires detailed local heat transfer information on and in the vicinity of the porous injector; however, limited spatially resolved experimental studies exist, particularly in hypersonic flows. In this work, experiments were conducted in the University of Oxford’s high-density tunnel at Mach 6.1 in both laminar and turbulent regimes. Spatially resolved two-dimensional surface heat transfer measurements were acquired by imaging directly on and downstream of two microporous transpiration-cooled injectors (METAPOR® CE170 and zirconia) using high-speed infrared thermography. Whereas injection in the laminar regime results in a steady, monotonic reduction in heat transfer from the start of the injector, a flatter profile is present for the turbulent cases where turbulent mixing inhibits surface heat transfer reduction. It was found that a modification to existing relations from film theory successfully correlates the streamwise heat transfer distribution on the injector for different blowing rates of nitrogen and helium. A key result is that helium performs much better than reported in previous experiments. Finally, the downstream thermal effectiveness is characterized for turbulent flows. A collapse of the thermal effectiveness is achieved and a modified analytical correlation proposed.

Transient cooling performance of R134a flash-evaporation spray enhanced by cold air jet: A novel heat transfer enhancement method in laser dermatology

Nevus of Ota is a dermal hyperplastic pigmented disease with a morbidity rate of 0.1–0.6% in the Asian population. Laser surgery with the assistance of cryogen spray cooling (CSC) is a potential clinical treatment for nevus of Ota. However, for Asians with high melanin content, it is necessary to enhance cooling capability without the increment of refrigerant amount. In this study, a R134a flash-evaporation cryogen spray cooling coupled with cold air jet (CSC-CAJ) was developed to improve the cooling capacity. Epoxy resin was employed as a skin phantom to investigate the heat transfer characteristics of CSC-CAJ with different air jet parameters. Air jet parameters of CSC-CAJ are recommended to be Ta = -7 ℃, Va = 3 m3/h Za = 10 mm and Φa = 30°. Compared to CSC, CSC-CAJ could reduce the minimum surface temperature from -39 to -42.1 ℃, and the average value of convective heat transfer coefficient (h) of CSC-CAJ could increase from 8 to 12.9 W/(m2·k) or 61.3% for the recommended parameters. The liquid film residence time (τ) decreases from 301 to 234 ms or 23.3%. In addition, based on the visualization observation and surface heat transfer measurement, the frost formation phenomenon on the surface and the heat transfer enhancement mechanisms are well understood. Four modes, including Leidonfrost mode, no runoff mode, runoff mode and evaporation mode, are proposed to describe the four stages of CSC-CAJ. A non-dimensional correlation of normalized Nusselt number (Nu*) and time (t*) is summarized. The primary heat transfer enhancement mechanism of CSC-CAJ is found to be the increase in cryogen evaporation rate and the ratio of heat removal from the phantom surface to phase change latent heat.

Heat Transfer Measurements of a Transpiration-Cooled Stagnation Point in Transient Hypersonic Flow

This paper presents a novel experimental technique where infrared thermography is employed to directly measure the surface heat transfer of a transpiration-cooled porous material in transient hypersonic flow. Experiments were conducted in the Oxford High Density Tunnel on a flat-faced hemispherical probe at a single Mach 7 freestream condition () with nitrogen, air, argon, krypton, and helium injection gases and mass flow rates ranging from 0.01 to . Surface heat transfer measurements were extracted by imaging directly on the porous material using a FLIR A6751 high-speed long-wave infrared camera. Porous alumina was chosen due to its favorable thermal properties for infrared analysis and its very small pore sizes () enabling a uniform outflow. It was found that the surface Stanton number reduction matched to within 10% of both computational fluid dynamics results and correlations.

Turbulent spot transit of a hypersonic laminar separation

This paper presents an experimental study of the interaction of turbulent spots with an initially laminar separation, using schlieren visualisation and dynamic measurements of surface heat transfer and pressure. The separation is generated on a blunt cylinder-flare body, tested at Mach 9, and entropy layer effects result in a boundary layer edge Mach number of 3.43 at separation. A single roughness element is used to produce isolated spots, defined such that separation collapse and re-establishment is completed before arrival of the next spot. On average the spot axial length, at the start of the interaction, is 2.5 times the separation length, growing to 6.8 times as the spot base passes the original reattachment position. As such the spot comprises a large perturbation. The local separation responds almost immediately with passage of the spot, so that the downstream region may be unaffected while the upstream part has already collapsed. Re-establishment of separation is slow, taking up to four times the total transit time of the spot. A basic model is presented to explore the transient wave processes during spot passage.

Influence of instability modes on cooling performance in hypersonic boundary layer with slot injection

A combined numerical-experimental investigation is presented with focus on the effects of boundary-layer instabilities and transition on the wall cooling performance in a Mach 5 low-enthalpy flow over a flat plate, with coolant injection achieved through a row of slots. The numerical study has been performed through direct numerical simulation (DNS) of the compressible Navier-Stokes equations, and is supported by results from linear stability analysis (LST) for the considered boundary layer. The experiments have been conducted in the High Density Tunnel (HDT) of the Oxford Thermofluids Institute, and include several blowing ratio conditions of injected air for the same freestream conditions. Surface heat transfer and pressure measurements, film effectiveness measurements, and Schlieren images are presented. The analysis links the wall cooling performance to the growth of imposed unstable boundary layer modes. Results indicate that 2D and 3D unstable modes, pertaining to the class of first instability modes, exist in the laminar boundary layer, and that imposition of these modes at different amplitudes leads to different states of the boundary layer, which we refer to as a perturbed state and a transitional state for medium and high amplitude respectively. As confirmed by comparison with experimental data, the perturbed and transitional states of the boundary layer significantly affect the wall cooling performance, providing an increase of the wall heat flux that results in a reduction of the beneficial effects of cooling.

Experimental investigation of the effects of leading edge bluntness on supersonic flow over a double compression ramp

Understanding the flow characteristics over a double compression ramp is crucial for high-speed vehicle design. Leading edge bluntness is a key factor influencing the formation of a separation region on a double compression ramp flow. In the present study, the effect of bluntness on a double compression ramp is investigated experimentally at a nominal Mach number of 4. The test model has 13° and 40° inclinations with respect to the freestream. Five different levels of leading-edge radius, varying from 0.0 to 2.0 mm, were subjected to supersonic wind tunnel tests. Shadowgraph and infrared thermography techniques were employed to visualize the flow features of the double ramp model. Measurements of surface heat-transfer along the centerline of the test model were obtained from the acquired infrared images. It is shown that the leading-edge radius alters the separation characteristics as well as the surface heat-transfer. Possible reasons for such flow characteristics are discussed.

Dynamic surface heat transfer and re-attachment flow measurement using luminescent molecular sensors

Luminescent molecular sensors, which can be used as Temperature Sensitive Paint (TSP) or Pressure Sensitive Paint (PSP) in respective temperature or pressure measurements, has been proved to be one promising surface quantity measurement technology in recent years. It is advantageous in the experiments with complicated surface and is able to give flow field information such as surface flux and wall shear-stress. The current study is focused on the dynamic heat transfer measurement of a backward-facing step model, using luminescent molecular sensors as temperature probe. The effects of flow separation and re-attachment after a back-step model were experimentally discussed in wind tunnel tests. The experimental system was consisted of a molecular sensor calibration system, a dynamic data recording system and a data processing system. It is found that the reattachment process will form a low temperature region, which then gives the clear temperature field of the flow. Dynamic temperature field data show a re-attachment position around x/h = 5.7, which agrees well with oil-flow measurements as well as previous experiments. The dynamic temperature fluctuation data is discussed with the vibrations of flow and transient heat transfer behaviors after the backward step, which then is used in the analysis of surface wall shear-stress variations. It is concluded that the current luminescent molecular sensor method is capable of quantitative measurement for surface heat transfer and fluid flows.

Multi-walled carbon nano-tubes for performance enhancement of thin film heat flux sensors

Surface heat transfer measurement is an important aspect in many research problems. Thin film heat flux sensor (TFHFS) is mostly considered in such situations for heat flux measurement due to its quick response and high accuracy. In the present studies, multi-walled carbon nanotubes (MWCNTs) are mixed with platinum while making the TFHFSs. Such addition is noticed to increase the sensitivity and decrease the temperature coefficient of resistance (TCR) of the thin film sensors. Improved sensitivity by 151% and 119% for Macor and Quartz sensors has led to increase in strength of the temperature response of the sensors during dynamic calibration experiments. Though heat flux recovery is seen to have encouraging agreement for all the sensors, sensitivity enhancement is noticed to be more prominent and advantageous for Macor based sensors. Present studies recommend adequately finished substrate for better adhesion. Further, use of MWCNTs is advisable especially for low heat flux measurement and also for Macor substrate since either situation demands the higher sensitivity to increase the output response.

Freejet Testing of the HIFiRE 7 Scramjet Flowpath at Mach 7.5

Results are presented on the freejet testing of a 75%-scale replica of the HIFiRE 7 scramjet. The HIFiRE 7 scramjet flowpath includes a two-dimensional forebody, a rectangular-to-elliptical shape transition inlet, an elliptical combustor, and a thrust nozzle. Only porthole injectors are used, giving the internal flowpath a clean configuration suitable for high-Mach-number operation with no physical obstructions to the flow. Furthermore, the structure of the shock waves in the scramjet is tailored to introduce a shock-wave/boundary-layer interaction in the combustor to promote fuel ignition. The objective of the experiments is to investigate the performance of this scramjet flowpath at a simulated flight Mach number of 7.5 and an altitude of 29.5 km with gaseous hydrogen as the fuel. Static pressure measurements show that robust combustion can be sustained in the flowpath at a range of equivalence ratios between 0.48 and 0.84. Based on these experiments, a fuel equivalence ratio of 0.8 is recommended for the flight. Corresponding surface heat transfer measurements reveal that, when the fuel–air mixtures ignite and burn, the surface heat transfer levels in the combustor and nozzle increase to as much as three times the fuel-off levels. A quasi-one-dimensional cycle analysis of the tests shows that the overall fuel-based combustion efficiency of this engine is 89% at the fuelling conditions planned for the flight.

Determination of Cooling Parameters for a High-Speed, True-Scale, Metallic Turbine Vane

Facilities such as the Turbine Research Facility (TRF) at the Air Force Research Laboratory have been acquiring uncooled heat transfer measurements on full-scale metallic airfoils for several years. The addition of cooling flow to this type of facility has provided new capabilities and new challenges. Two primary challenges for cooled rotating hardware are that the true local film temperature is unknown, and cooled thin-walled metallic airfoils prohibit semi-infinite heat conduction calculation. Extracting true local adiabatic effectiveness and the heat transfer coefficient from measurements of surface temperature and surface heat transfer is therefore difficult. In contrast, another cooling parameter, the overall effectiveness (ϕ), is readily obtained from the measurements of surface temperature, internal coolant temperature, and mainstream temperature. The overall effectiveness is a normalized measure of surface temperatures expected for actual operating conditions and is thus an important parameter that drives the life expectancy of a turbine component. Another issue is that scaling ϕ from experimental conditions to engine conditions is dependent on the heat transfer through the part. It has been well-established that the Biot number must be matched for the experimentally measured ϕ to match ϕ at engine conditions. However, the thermal conductivity of both the metal blade and the thermal barrier coating changes substantially from low-temperature to high-temperature engine conditions and usually not in the same proportion. This paper describes a novel method of replicating the correct thermal behavior of the thermal barrier coating (TBC) relative to the metal turbine while obtaining surface temperature measurements and heat fluxes. Furthermore, this paper describes how the ϕ value obtained at the low-temperature conditions can be adjusted to predict ϕ at high-temperature engine conditions when it is impossible to match the Biot number perfectly.

Roughness-induced turbulent wedges in a hypersonic blunt-body boundary layer

AbstractThis paper uses measurements of surface heat transfer to study roughness-induced turbulent wedges in a hypersonic boundary layer on a blunt cylinder. A family of wedges was produced by changing the height of an isolated roughness element, providing conditions in the following range: fully effective tripping, for the largest element, with a turbulent wedge forming immediately downstream of the element; a long wake, in length several hundred times the boundary layer thickness, leading ultimately to transition; and retention of laminar flow, for the smallest element. With appropriate element size, a fully intermittent wedge formed, comprising a clear train of turbulent spots.

Towards convective heat transfer enhancement: surface modification, characterization and measurement techniques

In this work, heat transfer surface modification and heat transfer measurement technique is developed. Heat transfer investigation was aimed to study the effect of carbon nano fibers (extremely high thermal conductive material) on the enhancement level in heat transfer. Synthesis of these carbon nano structures is achieved using thermal catalytic chemical vapor deposition process (TCCVD) on a 50 μm pure nickel (Ni270) wire. The micro wire samples covered with CNF layers were subjected to a uniform flow from a nozzle. Heat transfer measurement was achieved by a controlled heat dissipation through the micro wire to attain a constant temperature during the flow. This measurement technique is adopted from hot wire anemometry calibration method. Synthesis of carbon nano structures, heat transfer surface characterization and measurement technique are evaluated. Preliminary results indicate that an average enhancement in Nusselt Number of 17% is achieved.

HP Vane Aerodynamics and Heat Transfer in the Presence of Aggressive Inlet Swirl

Modern lean burn combustors now employ aggressive swirlers to enhance fuel-air mixing and improve flame stability. The flow at combustor exit can therefore have high residual swirl. A good deal of research concerning the flow within the combustor is available in open literature. The impact of swirl on the aerodynamic and heat transfer characteristics of an HP turbine stage is not well understood, however. A combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK. The swirl simulator is capable of generating an engine-representative combustor exit swirl pattern. At the turbine inlet plane, yaw and pitch angles of over ±40 deg have been simulated. The turbine research facility used for the study is an engine scale, short duration, rotating transonic turbine, in which the nondimensional parameters for aerodynamics and heat transfer are matched to engine conditions. The research turbine was the unshrouded MT1 design. By design, the center of the vortex from the swirl simulator can be clocked to any circumferential position with respect to HP vane, and the vortex-to-vane count ratio is 1:2. For the current investigation, the clocking position was such that the vortex center was aligned with the vane leading edge (every second vane). Both the aligned vane and the adjacent vane were characterized. This paper presents measurements of HP vane surface and end wall heat transfer for the two vane positions. The results are compared with measurements conducted without swirl. The vane surface pressure distributions are also presented. The experimental measurements are compared with full-stage three-dimensional unsteady numerical predictions obtained using the Rolls Royce in-house code Hydra. The aerodynamic and heat transfer characterization presented in this paper is the first of its kind, and it is hoped to give some insight into the significant changes in the vane flow and heat transfer that occur in the current generation of low NOx combustors. The findings not only have implications for the vane aerodynamic design, but also for the cooling system design.

High-resolution hot-film measurement of surface heat flux to an impinging jet

To investigate the complex coupling between surface heat transfer and local fluid velocity in convective heat transfer, advanced techniques are required to measure the surface heat flux at high spatial and temporal resolution. Several established flow velocity techniques such as laser Doppler anemometry, particle image velocimetry and hot wire anemometry can measure fluid velocities at high spatial resolution (µm) and have a high-frequency response (up to 100 kHz) characteristic. Equivalent advanced surface heat transfer measurement techniques, however, are not available; even the latest advances in high speed thermal imaging do not offer equivalent data capture rates. The current research presents a method of measuring point surface heat flux with a hot film that is flush mounted on a heated flat surface. The film works in conjunction with a constant temperature anemometer which has a bandwidth of 100 kHz. The bandwidth of this technique therefore is likely to be in excess of more established surface heat flux measurement techniques. Although the frequency response of the sensor is not reported here, it is expected to be significantly less than 100 kHz due to its physical size and capacitance. To demonstrate the efficacy of the technique, a cooling impinging air jet is directed at the heated surface, and the power required to maintain the hot-film temperature is related to the local heat flux to the fluid air flow. The technique is validated experimentally using a more established surface heat flux measurement technique. The thermal performance of the sensor is also investigated numerically. It has been shown that, with some limitations, the measurement technique accurately measures the surface heat transfer to an impinging air jet with improved spatial resolution for a wide range of experimental parameters.

Heat Transfer Characteristics of an Oblique Jet Impingement Configuration in a Passage With Ribbed Surfaces

Heat transfer measurements of a confined impingement cooling configuration with ribs on the target surfaces are presented. The assembly consists of four nonperpendicular walls of which one holds two rows of staggered inclined jets, each impinging on a different adjacent wall. The ribs are aligned with the inclined jet axes, have the same pitch, and are staggered to the impinging jets. The flow exhausts through two staggered rows of holes opposing the impingement wall. The passage geometry is related to a modern gas turbine blade cooling configuration. A transient liquid crystal technique was used to take spatially resolved surface heat transfer measurements for the ground area between the ribs. A comparison with the smooth baseline configuration reveals local differences and a generally reduced heat transfer for the rib-roughened case. Furthermore, lumped heat capacity measurements of the ribs yielded area averaged heat transfer information for the ribs. From the combination of ground and rib heat transfer measurements, it is concluded that the overall performance of the ribbed configuration depends on the Reynolds number. Of the five investigated jet Reynolds numbers from 10,000 to 75,000, only for the highest Re the averaged Nusselt numbers increase slightly compared with the smooth baseline configuration.

Comparative Surface Heat Transfer Measurements inHypervelocity Flow

Reliable prediction of the high heat transfer rates experienced during the hypersonic portion of planetary entry and descent is critical to vehicle survival. Two types of sensors that can be used to measure surface heat flux are coaxial thermocouple gages and thin film resistance thermometers. Individually, both types of gages have been used successfully in a number of studies [1–19]. Both thermocouple and thin film gages measure surface temperature from which heat transfer can be calculated. Both have µs response times, and can be flush-mounted in models. Coaxial thermocouples are robust, can survive challenging experimental conditions, and are typically used in higher enthalpy flows. Thin film resistance gages typically provide improved signal levels, but are less robust, have to be individually calibrated, and are typically used in lower enthalpy flows. As a result, there are few studies which directly compare measurements from the two types of gages. In the present work, we perform experimental measurements at a range of intermediate enthalpies in hypervelocity flow and make direct comparisons between temperature histories and heat flux data obtained from thermocouple and thin film gages.

Transient surface heating rates from a nickel film sensor using inverse analysis

Convective surface heat transfer measurements play an important role in many industrial, environmental and aerodynamic problems. In most of the cases, the flow is unsteady which results in temperature variation in the body. The surface heating rates are then predicted from the measured temperature histories by suitable heat transfer modeling. In this paper, the temperature history obtained from a nickel film sensor during a flight test is considered to study the effect of sensor thickness on surface heat flux measurements during the flight measurement. Inverse methods using analytical solutions as well as control volume approximations are used to infer the surface heat flux. The experimental temperature data are discretized using cubic-spline method to obtain the closed form solution which is used for inverse analysis. The results are compared with that of standard bench mark results with thin film gauge analysis based on semi-infinite one dimensional medium. No significant change in surface heat flux is observed between inverse and thin film analysis. However, when the thickness of nickel film is increased by 100 times during numerical simulation of inverse method, it is seen that peak surface heat flux increases by 20%.

The Effects of Freestream Turbulence, Turbulence Length Scale, and Exit Reynolds Number on Turbine Blade Heat Transfer in a Transonic Cascade

This paper experimentally investigates the effect of high freestream turbulence intensity, turbulence length scale, and exit Reynolds number on the surface heat transfer distribution of a turbine blade at realistic engine Mach numbers. Passive turbulence grids were used to generate freestream turbulence levels of 2%, 12%, and 14% at the cascade inlet. The turbulence grids produced length scales normalized by the blade pitches of 0.02, 0.26, and 0.41, respectively. Surface heat transfer measurements were made at the midspan of the blade using thin film gauges. Experiments were performed at the exit Mach numbers of 0.55, 0.78, and 1.03, which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 6×105, 8×105, and 11×105, based on true chord. The experimental results showed that the high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the blade as compared with the low freestream turbulence case. At nominal conditions, exit Mach 0.78, average heat transfer augmentations of 23% and 35% were observed on the pressure side and suction side of the blade, respectively.

Evaluation of Alternatives for Two-Dimensional Linear Cascade Facilities

This paper presents two low-cost alternatives for turbine blade surface heat transfer and fluid dynamics measurements. These models embody careful compromises between typical academic and full-scale turbomachinery experiments and represent a comprehensive strategy to develop experiments that can directly test shortcomings in current turbomachinery simulation tools. A full contextual history of the wide range of approaches to simulate turbine flow conditions is presented, along with a discussion of their deficiencies. Both models are simplifications of a linear cascade: the current standard for simulating two-dimensional turbine blade geometries. A single passage model is presented as a curved duct consisting of two half-blade geometries, carefully designed inlet and exit walls and inlet suction. This facility was determined to be best suited for heat transfer measurements where minimal surface conduction losses are necessary to allow accurate numerical model replication. A double passage model is defined as a single blade with two precisely designed outer walls, which is most appropriate for flow measurements. The design procedures necessary to achieve a desired flow condition are discussed.