The ASETS-II experiment consists of three oscillating heat pipes (OHPs), an electronics box, and mounting structures that control boundary conditions. Each OHP consists of 34 channels in a typical single-layer closed loop design. Butane was selected as the working fluid for OHP #1 and #2 for its performance stability. R-134a was selected for OHP #3 in order to explore the Bond number limit’s influence on OHP operation in microgravity. The ASETS-II Flight and Flight Spare hardware were subjected to a comprehensive set of ground testing to baseline performance prior to flight testing. For most test conditions, the Flight and Flight Spare test results for OHP #1 and OHP #2 are within the margin of uncertainty in the measurements. OHP #3 on the Flight hardware performs similarly to OHP #3 on the Flight Spare hardware; however, the difference in performance is outside the margin of uncertainty in the measurements. This variation in performance may be attributable to the fact that OHP #3 is being pushed to operate near its Bond number limit.

A new advanced film cooling scheme, named high-aspect-ratio holes has been proposed. Four configurations were designed, and numerically simulated under density ratio of 2 and different blowing ratios. All configurations demonstrate extremely high film cooling effectiveness values, some are as high as the so-called perfect performance, while their mechanical strength are similar to the conventional schemes. The new scheme exhibits two traits distinctive from the conventional geometries: Its film cooling effectiveness is much higher than the coverage ratio ( t/P ), and the high film cooling effectiveness is obtained under strong counter-rotating vortex pair (CRVP). It has been found that, in the new scheme, along with the aspect ratio value increase, the CRVP move away from the coolant-mainstream interface, and the coolant laterally expands in the vicinity of the exit. Consequently, a continuous coolant film would occur near the trailing edge position if aspect ratio is high enough. The approach of high-aspect-ratio holes could be used to design the highest film cooling performance geometries.

In this work, an anti-icing “Piccolo tube” system is numerically analyzed, in order to investigate the efficiency of hot air impinging jets on internal wing surface. The model is two-dimensional and it is composed by the wing profile and an external volume. External air is considered a mixture, composed by dry air and water-vapor. A steady state problem with standard k-ε model is assumed. Numerical simulations are performed considering an average coefficient on the external surface and thermofluidynamic field inside the wing due to the anti-icing system. Numerical solutions are carried out by means of the Ansys-FLUENT code. Simulations are performed considering two values of the angle of attack, 0° and 4°, and two Mach number values, 0.576 and 0.700, related to the internal flow. esults are presented in terms of velocity and temperature fields, wall temperature profiles along the wing surfaces. As comparison, a case with only dry air in the external domain is performed. Results indicate that in presence of droplets the anti-icing system is less efficient than in the ideal case of dry air. Furthermore, a larger angle of attack improves the performance of the anti-icing system, but a larger Mach number deteriorates it.

This study compares several band models for calculating the thermal radiation from the plume of a simulated rocket motor. First, a non-scattering plume is considered allowing one to use the line-of-sight integration (LOS). The results for three band models are compared. Next we study a scattering plume. The radiative heat transfer equation is solved numerically using the axisymmetric finite volume method (FV). Detailed analysis of the numerical scheme of the FV method demonstrated that it has a large numerical diffusion term which, although formally of the first order, in practice can cause large errors in calculating thermal signatures of hot jets. The numerical scheme was hence modified using what we call the “quasi-Cartesian” approach. We demonstrate that for a non-scattering media in the first order the numerical diffusion term vanishes and the modified scheme formally reduces to the integration along the line of sight resulting in a very good agreement between the FV and LOS predictions. The new method is then used to calculate radiation from a scattering plume and the results for two band models are compared.

This study is dedicated to an innovative heatshield aeroshape intended to be the AEROFAST martian aerocapture demonstrator. The analysis of three representative Mars aerocapture trajectory points (maximum heat flux, maximum dynamic pressure and exit point) results in building a partial AeroThermo-DataBase including both convective and radiative heat flux evaluations carried out assuming CO 2 chemical non equilibrium.

The present study concerns the investigation of external natural convection driven around a Montgolfiere scientific balloon. Ground measurements are carried out on a small scale heated model located within an enclosure. Non intrusive planar Particle Image Velocimetry and thermocouple sensors are used to characterize both dynamical behavior of the flow generated around the balloon and heat transfers on its surface. It is shown that a relative good agreement is reached on wall heat transfers with simplified axisymmetrical 2D RANS simulations. However significant discrepancies exist regarding the analysis of the flow topology and dynamic quantities in comparison with the experimental data. The proposed numerical rebuilding is thus completed with a 3D Delayed Detached Eddy Simulation (DDES) in order to overcome conventional RANS approach inability to represent low frequency thermal plume instabilities.

A thin flat plate heat pipe (TFPHP) with a new structure of wick is fabricated and tested. The wick is formed by the narrow foils folded, the liquid passage and capillary force is provided by the folded clearance of the foils, the vapor passage is shaped by the periodic holes of the foils. One aluminum TFPHP of 260×60×4 mm 3 size is made and the test setup is established. The test results show that the spreader has good performance as, at the horizontal state, the power transferred can reach 18.5W, the corresponding thermal conductivity is 838W/(m·°C). For different tilt angle, we have the performance results. The effects of parameters to the performance are discussed. The spreader is significant for the spreading of point power.

Ice formation on airplane wing profile is a very dangerous condition because of the change in the profile aerodynamic, so it is necessary to avoid ice formation on the wings. The hardest condition for ice formation are at altitudes between 10.000 and 15.000 ft and at temperature between 0 °C and −15 °C, because they are particularly suitable for ice formation. In this paper an anti-icing system based on hot air impinging jets on internal wing surface is analyzed in order to check the efficiency of the system. A numerical model is given in order to evaluate the thermal and fluid dynamic behaviours of the impinging jet inside the wing panel. A wing profile with an angle of attack of 4.50° is taken into account with a free stream temperature of 258 K. A piccolo tube with a diameter of 1.00 inch and air temperature of 523 K is considered for anti-icing system. A structured mesh is used in the discretization of the computational domain for the two-dimensional and three-dimensional case. A steady state solution with k-ε RNG turbulent model has been found. Numerical simulations of a two and a three dimensional model of an aircraft wing has been carried out taking into account the external convective exchange by means of an average coefficient on the external surface and thermofluidynamic field inside the wing due to the anti-icing system. The analysis is performed by means of the FLUENT code in order to find the optimal geometrical configuration to avoid the ice formation on the external wing surface. Results are presented in terms of stream function and temperature fields, wall temperature profiles along the wing surfaces and air velocity and temperature profiles.

Carbon nanotube (CNT) forest/cluster synthesized by a thermal CVD process has millimeter growth height, large porosity and nano level pore size, plus high thermal conductivity of individual CNT, thus it is potentially a good wick structure material in developing micro heat transfer devices. However, thermal properties, including effective thermal conductivity (ETC) of a bulky CNT layer, may not be as good as the common metallic wick materials. In this paper, a Netzsch DSC 404 C Pegasus is used for measurement of the CNT heat capacity. CNT volume density is obtained by measuring the ratio of a bulky CNT weight and volume. Both the laser flash and 3-omega measurement methods are employed to measure ETC for CNT wick structures synthesized by the thermal CVD processes. For the laser flash method, measurement deviations caused by reflective silicon and thin substrate are corrected by surface treatment and increased sample thickness. Measurement results of the laser flash indicate that a 600μm thick CNT layer has ETC varying from 0.7–1.2W/m.K. For the 3-omega approach, the measurement system is validated on a quartz substrate. However, the test results yield larger ETC on 250μm CNT samples. Geometric and one dimensional thermal conduction analysis indicate that the bulky CNT thermal properties are tied to CNT synthesis processes. ETC of bulky CNT layer can be enhanced by straightening CNT growth and increasing CNT growth volume density.

With the increase of power consumption in compact electronic devices, passive heat transfer cooling technologies with high heat flux characteristics are highly desired in microelectronics industries. Carbon nanotube (CNT) cluster/forest has high effective thermal conductivity, nano pore size and large porosity, which can be used as wick structure in a heat pipe heatspreader and provides high capillary force for high heat flux thermal management. In this research, investigations of high heat flux cooling of the CNT bi-wick structure are associated with the development of a reliable thermometer and high performance/interface free heater. A 100nm thick and 600μm wide Z-shaped platinum wire resistor is fabricated on the backside of a CNT sample substrate to heat a 2×2mm 2 wick area. As a heater, it provides direct heating effect without thermal interface and is capable of over 800°C high temperature operation. As a thermometer, reliable temperature measurement is achieved by calibrating the resistance variation with temperature after the annealing process is applied. The CNT sample substrate is silicon. The backside of the silicon substrate is thermally oxidized to create a 2μm thick and pinhole-free SiO 2 layer so that the platinum heater and thermometer can survive from the server CNT growth environments and without any electrical leakage. For high heat flux cooling, the CNT bi-wick structure is composed of 250μm tall, 100μm wide stripe-like CNT clusters and 50μm empty space. Using 1×1cm 2 CNT bi-wick samples, experiments are completed in both the open and saturated environments. Testing results of CNT bi-wick structure demonstrate 600W/cm 2 heat transfer capacity and good thermal & mass transport characteristics in the nano level porous media.

Heat transfer characteristics of an impinging air jet are experimentally investigated under a fixed pumping power condition. The effects of dimensionless pumping power on the Nusselt number are considered. The focus is on cases where the nozzle-to-plate spacing is equal to or less than one nozzle diameter. The results show that the Nusselt number is independent of the nozzle-to-plate spacing under fixed pumping power conditions, while the Nusselt number increases with decreasing the nozzle-to-plate spacing under fixed flow rate conditions. Based on the experimental results, new correlations for the stagnation and average Nusselt numbers of the impinging jet are developed as a function of the pumping power alone.

Owing to the persistent interest in the approximate methods, the order of magnitude analysis is revisited in the present work and a method is proposed for the quick estimation of drop in temperature of a fluid flowing through a circular duct. The principle involves judicious guessing of temperature drop for one of the investigated cases, while the temperature estimates for rest of the cases are deduced from the initial guess. A problem of duct flow with widely varying inlet temperature and mass flow rate conditions, typical of a high altitude simulation ground test facilities for scramjet combustors, re-entry vehicles, etc., are considered which vary from 400 K to 1200 K with corresponding mass flow rates of 300 kg/s to 0.5 kg/s. The results of this quick estimation method portray remarkable agreement with the exit temperatures of the fluid as predicted by solving the problem using transient quasi-one dimensional codes and two dimensional CFD techniques over the entire operating regime and hence provide a conservative estimate of vital parameters for any design exercise as well as to rate an existing system at off-design conditions. The significant advantages of the proposed method over the others are viz. (a) simplicity (b) lesser computational effort and (c) reasonable accuracy, will be elucidated in detail in this article.

An Advanced Vapor Compression Loop (AVCL) for high temperature lift for heat rejection to hot lunar surface during lunar daytime was developed. The loop consists of an evaporator, a compressor, a condenser, and an electronic expansion valve. Different types of evaporators were evaluated in this study: a circular tube evaporator, a circular tube evaporator with a twisted tape, a circular tube evaporator with a wick, and a circular tube evaporator with a wick and a twisted tape. The evaporators were tested with two different compressors. The first was a 0.5hp oil-less compressor and the second was a 5.3hp compressor that used oil as lubricant. A heat exchanger (recuperator) was used to subcool the high pressure liquid and to superheat the low pressure vapor. Tests were performed with and without the recuperator. Vapor superheat during the tests was controlled with an electronic expansion valve controller. The working fluid was R134a. The results show that the heat source-to-working fluid thermal resistance of the circular tube evaporator with the wick and the twisted tape was one-third of that of the circular tube evaporator. The recuperator was able to decrease the vapor quality at the evaporator inlet and increase the vapor superheat at the compressor inlet. The evaporators without wicks were able to operate at a heat flux of 5.7W/cm 2 with the recuperator and vapor superheat set at 5°C. Evaporators with wicks reached dryout at lower heat fluxes when maintaining superheat at 5°C. However, the wicked evaporators reached a heat flux of 7.6W/cm 2 when decreasing superheat below 5°C. A temperature lift of 70°C was achieved with the 5.3hp compressor.

A loop heat pipe (LHP) analytical model that simulates the steady state and transient thermal behaviors of LHPs with multiple evaporators and multiple condensers has recently been developed. It can be used as a stand-alone computer code or as a subroutine to general spacecraft thermal analyzers. Multi-evaporator and multi-condenser LHPs are more complex in their operation when compared to single-evaporator LHPs because of the thermal and fluid interactions among the evaporators, compensation chambers, and condensers. This analytical model has been used to simulate the thermal performance of a miniature loop heat pipe (MLHP) with two evaporators and two condensers in laboratory and thermal vacuum tests. In addition, the MLHP was tested in the laboratory under five different configurations where the relative elevations and tilts among loop components were varied so as to investigate the gravity effects on the loop performance and to verify the analytical model’s capability to predict such effects. The MLHP performance tests that were simulated included start-up, high power, heat transport limit, and heat load sharing between the two evaporators. In all tests that were modeled, the LHP analytical model accurately predicted the steady state and transient behaviors of the LHP. Furthermore, the model was run-time efficient and yielded stable solutions in all cases.

The steady and unsteady laminar flow and heat transfer characteristics for a pair of opposing confined impinging slot jets in 2D and 3D were evaluated numerically at two Reynolds numbers. The present study continues the authors’ earlier work [1] and identifies the main similarities and differences arising from the expansion to the third dimension. At lower Reynolds number jet (Re = 300), the flow interaction produces a symmetric, steady flow hydrodynamic pattern with the jets being deflected laterally for the 2D flow. At Re = 300, the 3D slot jet produces almost the same values as the 2D case, yet the flow is slightly asymmetrical and unsteady. However, by further increasing the Reynolds number to 750, a complex and highly unsteady flow develops for both 2D and 3D simulations. The symmetry of both the 2D and 3D flows is disrupted and the resulting complex flow patterns reveal the vortex pairing effects, leading to the jet “buckling and sweeping” motion, enabling the enhanced local heat transfer. The convective heat transfer coefficients and the unsteady flow development between the jets are thoroughly investigated, with the flow unsteadiness also characterized by analyzing the stagnation point displacement on the channel walls. The comparison between the 2D and 3D flow patterns indicate that the 3D opposite jets enhance the unsteady effects compared to the 2D unsteady opposite jets. The complex vortex patterns resulting from the unsteady jets interaction, as well as the velocity, vorticity and temperature fields for both 2D and 3D cases are thoroughly evaluated. The comparison between the 2D and 3D impinging air jets is documented and the impact on chip/microelectronics cooling is highlighted.

In recent years there has been an ever increasing need to launch small payloads (∼1–100 kg) into low earth orbit (LEO). Examples include the defense and telecommunications industries. Permanent human presence in LEO, such as the international space station, requires continual re-supply from earth. Additionally, NASA’s stated mission of launching a manned mission to Mars requires many tonnes of raw materials to be economically launched into LEO and assembled there. Conventional rocket launch from earth is prohibitively expensive for small mass payloads. Estimates range from $7000–$20,000 to launch 1 kg of mass into low earth orbit. Several concepts have been proposed to economically launch small payloads from earth, including light gas guns, electromagnetic launchers and the so called “slingatron” concept. The goal of these concepts is to reduce the cost per kg (to under $1000) to achieve LEO. Each of these concepts involves launching small payloads that traverse the atmosphere and then placed into low earth orbit using thrusters to turn the velocity vector into a stable circular orbit. As the launch vehicle traverses the dense lower portion of the atmosphere it experiences severe thermal heating loads that must be absorbed by a thermal protection system (TPS) if the payload is to survive the transit. The University of Texas is currently heading a multi-university research initiative (MURI) to study the feasibility of launching small payloads into low earth orbit from an electromagnetic gun housed in an airborne platform. As part of the study, the aerothermal issues associated with traverse through the atmosphere and propellant mass required to achieve a stable circular orbit are investigated. The effort focuses on quantifying the required parasitic mass of the thermal protection system (TPS) and propellant need to place a nominal 10 kg launch mass into a circular low earth orbit from an electromagnetic launcher at 16 km altitude. The TPS is assumed to be graphite or carbon-carbon composite. In this effort, we consider ballistic trajectories only. Circular orbit is achieved using rocket thrusters at the terminal altitude. Total parasitic mass (TPS + propellant) is estimated for various launch angles.

The present study focuses on heat transfer behaviour in aircraft compartments. The objectives of the analysis were to investigate the transient effects on fluid flow structures and heat transfer mechanisms in aircraft wing boxes relative to aircraft turnaround times. Experimental methods employed were flow visualisation and thermocouple measurements. A simplified air filled aluminium rectangular enclosure of aspect ratio 0.25 was constructed to replicate an aircraft wing box. Rubber insulation was used between wall surfaces to maintain separate thermal boundary conditions at each wall. Flow visualisation was used to illustrate the transient evolution of full field flow structures and thermocouple measurements were recorded to investigate the full field transient temperature distribution. Experiments were carried out from time zero to steady state and were conducted for Rayleigh numbers of 2.87×10 6 , 4.81×10 6 and 7.39×10 6 based on enclosure height. Fluid flow patterns revealed the presence of two counter acting flows in the cavity with a downward motion adjacent to the front and rear sides of the enclosure. The downward motions were present due to the cooling of the warm air inside the cavity by the adjacent cooler aluminium side walls. Secondary flows existed in the lower region of the cavity specifically prominent to the front and rear surfaces where the height of detachment of the flow varied with time and temperature of the adjacent side walls. Transient spatial temperature distribution plots confirmed the approach to steady state was gradual, that a vertically thermal stratified distribution existed in the cavity and the presence of the thermal boundary layers along the horizontal and vertical walls were clearly evident. It was concluded that transient effects are significant relative to aircraft turnaround times as it was observed that the flow structures change over the initial transient phase of analysis and remain similar thereafter. Wing box material thermophysical properties influenced the developed flow structures causing two secondary flows to exist along the front and rear spar walls. Vertical temperature distribution results were generalised using a power law based on dimensionless time (t*), dimensionless temperature (T*) and Rayleigh number.

Current state-of-the-art thermoregulatory models do not predict body temperatures with the accuracies that are required for the development of automatic cooling control in liquid cooling garment (LCG) systems. Automatic cooling control would be beneficial in a variety of space, aviation, military, and industrial environments for optimizing cooling efficiency, for making LCGs as portable and practical as possible, for alleviating the individual from manual cooling control, and for improving thermal comfort and cognitive performance. In this paper, we adopt the Fiala thermoregulatory model, which has previously demonstrated state-of-the-art predictive abilities in air environments, for use in LCG environments. We compare the model’s tissue temperature predictions with analytical solutions to the bioheat equation, and with experimental data for a 700 W rectangular type activity schedule. The thermoregulatory model predicts rectal temperature, mean skin temperature, and body heat storage (BHS) with mean absolute errors of 0.13°C, 0.95°C, and 11.9 W·hr, respectively. Even though these accuracies are within state-of-the-art variations, the model does not satisfy the target BHS accuracy of ±6.5 W·hr. We identify model deficiencies, which will be addressed in future studies in order to achieve the strict BHS accuracy that is needed for automatic cooling control development.

Enclosure natural convection has applications in many engineering disciplines such as electronic component cooling, building ventilation and heating, and renewable energy systems such as solar collectors. To date there has been little research on its application in the aeronautical industry where it plays an important role in aircraft heat transfer. Standard civil aircraft comprise of a number of vented and non-vented compartments. One such compartment is the leading edge which houses the main aircraft bleed duct and several electronic devices. Natural convection heat transfer in the leading edge compartment is important during aircraft turn around while systems are running and the exterior of the aircraft is subjected to solar loading. The objective of this paper is to investigate the heat transfer in a leading edge. The effect of an internal heat generating source, a bleed duct, is also investigated. A dimensional analysis of the governing equations is carried out to determine the appropriate scaling groups for the natural convection cases and a numerical simulation is run on a commercial CFD package. The numerical results are compared with experimental measurements obtained from a leading edge test facility. The effect of bleed duct placement is also investigated both experimentally and numerically with good agreement achieved. It was found that bleed duct placement can have a significant effect on the overall temperature distribution within a leading edge compartment. The results provide a basis for the optimisation of natural convection in such aircraft wing compartments.

For many applications the optimisation of natural convection cooling is a major design consideration due to factors such as weight, accessibility, cost and power consumption. In aircraft wing compartments, natural convection is the dominant mode of heat transfer due to high wall temperatures resulting from solar loading and heat dissipating internal electronics. This paper investigates the flow structures in a leading edge compartment subject to various thermal boundary conditions. The experimental configuration consisted of two leading edge enclosures; the first is a single compartment while the second has an attached wing box. Particle image velocimetry (PIV) was employed to obtain velocity measurements of the flow in both leading edge enclosures. The second compartment investigated the effect of an adjacent fluid filled enclosure on the flow regime in the leading edge compartment. Higher local velocities were found in the second compartment due to an increase in buoyancy forces resulting from a lower of the average fluid temperature within the compartment. The introduction of a heat dissipating component gave rise to two separate convection structures and in general increased the fluctuations in the both temperature and velocities within the compartment.