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Turbine blade icing is a serious threat for the safety of wind power generation. Low energy ice protection techniques especially those using superhydrophobic surfaces, have attracted intensive interest recently. In this work, the anti-icing characteristics of wind turbine blade have been investigated in an ice wind tunnel, and the influence of surface wettabilities, preparation methods on the ice protection performance and the durability of different surface materials have been examined experimentally. It is found that the surface wettability can dramatically change the ice protection characteristics. Superhydrophobic surfaces prepared by both spraying and laser ablation methods can efficiently reduce the energy consumption of the electrothermal system, and the maximum reduction reaches 76.7% when the temperature of heating surface is lower than 15 °C . The surface tension induced flow pattern change of runback water should be responsible for the efficient anti-icing performance, where the runback water sheds from the superhydrophobic surface rapidly, preventing the runback icing on the surface. Droplets impacting can damage the wettabilities and the superhydrophobic surface prepared by laser ablation method shows a better durability during the ice wind tunnel test. A new concept of superhydrophobic-dry anti-icing has been proposed based on the results of this work, which is very promising in solving the icing issue of wind turbine blade with low energy consumption.

Considering the mass and energy sources carried by the accumulated ice layer, an unsteady heat and mass transfer model of the runback water film on the deicing surface is established to simulate aircraft electro-thermal deicing process. With the extension of the freezing coefficient to the transient calculation, the coupled heat transfer of the runback water and the solid skin is solved at each time step by a temperature-based method. Unsteady numerical simulation is carried out for the electro-thermal deicing system of a NACA 0012 airfoil. The temperature variations with time are in acceptable agreement with the literature data, and the unsteady temperature-based deicing model is verified. The calculation results of temperature, runback water flux and ice thickness on the deicing surface are analyzed at different time points, and it is shown that the unsteady electro-thermal deicing model can capture the main features of the icing, ice melting and re-freezing processes in the transient deicing simulations.

To study the effects of anisotropic thermal conductivity of composite aircraft skin on the heat transfer characteristics of electrothermal anti-icing system, the differential equation of anisotropic heat conduction was established using coordinate transformation of principal anisotropy axis. In addition, it was coupled with the heat and mass transfer model of the runback water film on the anti-icing surface to perform numerical simulation of the electrothermal anti-icing system. The temperature results of the vertical and cylindrical orthotropic thermal conduction in the rectangular and semi-cylindrical composite skin were consistent with those obtained by the traditional orthotropic model, which verified the anisotropic heat conduction model. The temperature distribution of anti-icing surface agreed well with the literature data, which validated the coupled heat and mass model of the runback water flow and the anisotropic skin. The anisotropic thermal conductivity of composite skin would make temperature change more gradual, and the effect was more significant where the curvature of the temperature curve was greater. However, the anti-icing surface of the electrothermal anti-icing system was slightly affected by the anisotropic heat conduction of the multilayered composite skin.

An experimental study was conducted to investigate the jet impingement thermal efficiency of a typical model for a multijet anti-icing system under different experimental conditions such as air mass flow rate at the jet exit, piccolo tube position and circumferential angle of jet holes. The experimental results showed that jet mass flow rate, velocity at the jet exits (whether choking conditions were established at the jet exits or not) and piccolo tube-to-leading edge distance within the cavity had significant influence on thermal efficiency of the anti-icing system. The heat flux on the external surface of leading edge, diameter of jet holes and circumferential angle of jet holes had a measurable but small effect on thermal efficiency of the anti-icing system.

A quantified parametric study for the heat transfer acting on a hypersonic blunt body with counterflowing jets is presented. Three-dimensional turbulent Navier-Stokes equations are solved to simulate freestream-jet interactive flowfields. The freestream and jet controlling parameters are treated as input sources of variation, and a point-collocation non-intrusive polynomial chaos (NIPC) method is utilized to quantify the variations in the output surface heat flux and total surface heat load acting on the blunt body by identifying the maximum and minimum of surrogate response values predicted by the NIPC. All of the sample cases are confirmed to form steady jet structures. Furthermore, through a global sensitivity analysis, Sobol indices evaluated by the NIPC, are used to rank the contributions of each input parameter to the variation in output quantities of interest. It is found that the designed upstream injection of baseline case effectively reduces the heat transfer to body surface compared with the no-jet case. The variations of input parameters induce remarkable variations of output heat flux and total heat load. The sensitivity analysis indicates that the jet-to-freestream total-pressure ratio is the top contributor to variations in heat flux, followed by the freestream Mach number. The jet total temperature is mainly important on the front part of forebody, while the contributions of jet Mach number and freestream temperature slightly increase downstream. The freestream density has the smallest effect. The sensitivity of total heat load to input parameters coincides with that of heat flux. This parametric study is expected to illustrate the significance of flow-controlling parameters to the heat transfer over blunt body and provide insight for aerothermal management by using counterflowing jets in hypersonic flows.

A novel 3-D unsteady model of in-flight electrothermal deicing process is presented in this paper to simulate the conjugate mass and heat transfer phenomena of water film runback, phase change, and solid heat conduction. Mathematical models of water film runback and phase change are established and solved by means of a loosely coupled method. At the current time step, solid heat conduction, water film runback, and phase change are iteratively solved until the heat boundary condition reaches convergence, then the temperature distribution and ice shape at the moment are obtained, and the calculation of the next time step begins subsequently. A deicing process is numerically simulated using the present model following an icing tunnel experiment, and the results match well with those in the literatures, which validate the present model. Then, an in-flight deicing process is numerically studied to analyze the effect of heating sequence.

The inflatable aerodynamic decelerator (IAD), which allows heavier and larger payloads and offers flexibility in landing site selection at higher altitudes, possesses potential superiority in next generation space transport system. However, due to the flexibilities of material and structure assembly, IAD inevitably experiences surface deformation during atmospheric entry, which in turn alters the flowfield around the vehicle and leads to the variations of aerodynamics and aerothermodynamics. In the current study, the effect of the static shape deformation on the hypersonic aerodynamics and aerothermodynamics of a stacked tori Hypersonic Inflatable Aerodynamic Decelerator (HIAD) is demonstrated and analyzed in detail by solving compressible Navier-Stokes equations with Menter's shear stress transport (SST) turbulence model. The deformed shape is obtained by structural modeling in the presence of maximum aerodynamic pressure during entry. The numerical results show that the undulating shape deformation makes significant difference to flow structure. In particular, the more curved outboard forebody surface results in local flow separations and reattachments in valleys, which consequently yields remarkable fluctuations of surface conditions with pressure rising in valleys yet dropping on crests while shear stress and heat flux falling in valleys yet rising on crests. Accordingly, compared with the initial (undeformed) shape, the corresponding differences of surface conditions get more striking outboard, with maximum augmentations of 379 pa, 2224 pa, and 19.0 W/cm2, i.e., 9.8%, 305.9%, and 101.6% for the pressure, shear stress and heat flux respectively. Moreover, it is found that, with the increase of angle of attack, the aerodynamic characters and surface heating vary and the aeroheating disparities are evident between the deformed and initial shape. For the deformable HIAD model investigated in this study, the more intense surface conditions and changed flight aerodynamics are revealed, which is critical for the selection of structure material and design of flight control system.

Experimental study of the local and average heat transfer characteristics of a single round jet impinging on the concave surfaces was conducted in this work to gain in-depth knowledge of the curvature effects. The experiments were conducted by employing a piccolo tube with one single jet hole over a wide range of parameters: jet Reynolds number from 27000 to 130000, relative nozzle to surface distance from 3.3 to 30, and relative surface curvature from 0.005 to 0.030. Experimental results indicate that the surface curvature has opposite effects on heat transfer characteristics. On one hand, an increase of relative nozzle to surface distance (increasing jet diameter in fact) enhances the average heat transfer around the surface for the same curved surface. On the other hand, the average Nusselt number decreases as relative nozzle to surface distance increases for a fixed jet diameter. Finally, experimental data-based correlations of the average Nusselt number over the curved surface were obtained with consideration of surface curvature effect. This work contributes to a better understanding of the curvature effects on heat transfer of a round jet impingement on concave surfaces, which is of high importance to the design of the aircraft anti-icing system.

In this paper, a model for simulating the rapid decompression of aircraft cabins is developed. The air flow between the cabins as well as between the damaged compartment and the external environment is simulated by assuming a reversible transformation of the insulation. Mathematical modelling for the active ventilation equipment is provided, with an emphasis on the impact of the opening process on the rapid decompression process. The numerical solution is introduced and the influence of the hole’s location on the decompression process is highlighted. The comparison with the literature results proves the effectiveness of the proposed method. This method can also be applied to the rapid decompression analysis of the actual aircraft.

Extensive experimental studies of the heat transfer characteristics of jet impingement on a variable-curvature concave surface in a wing leading edge were conducted for aircraft anti-icing applications. The experiments were performed using a piccolo tube with three rows of aligned jet holes over a wide range of parameters: the jet Reynolds number (Rej) from 50,000 to 90,000, the relative tube-to-surface distance (H/d) from 1.74 to 20.0, the jet impingement angle (α) from 66° to 90°, and the relative chordwise arc length in the jet impingement zone (r/d) from 13.2 to 34.8. Experimental results indicated that the heat transfer performance at the stagnation point was enhanced with increasing Rej and α, and an optimal H/d existed to achieve the best heat transfer performance at the stagnation point corresponding to specific operating parameters. It was found that the attenuation coefficient curve of jet impingement heat transfer in the chordwise direction exhibited an approximate bell shape with the peak located at the stagnation point, affected only by r/d in the peak zone. In the non-peak zone, however it was affected significantly by a variety of factors including Rej, H/d and r/d. Experimental data-based correlations of the Nusselt number at the stagnation point and the distribution of the attenuation coefficient in the chordwise direction were developed and validated, which contributes significantly to the future design of a wing anti-icing system with three rows of aligned jet holes.

Nucleate boiling is one of the most efficient and effective heat transfer modes, but is limited by the critical heat flux (CHF). An innovative artery porous structure was proposed in this work to enhance the CHF based on the concept of “phase separation and modulation” by forming individual flow paths for liquid supply and vapor venting while keeping the liquid/vapor interface located in the porous structure. In the experiment, the porous structure was made of sintered copper microparticles, multiple arteries were machined directly on the heated surface, and water was employed as the working fluid. The experimental results were compared with those on a flat surface, and a unique evaporation/boiling curve for the artery porous structure was revealed. The experiment validated the principle proposed here for CHF enhancement, and a maximum heat flux of 416 W/cm2 on a heating area of 0.78 cm2 was achieved without the occurrence of any dryout. Further increase of heat flux was limited only by the design temperature of the electrical heater, and a much higher CHF can be expected. In addition, the effects of pore size, artery depth and contact condition on the evaporation/boiling heat transfer performance in the artery porous structure were also experimentally investigated, which can guide further design optimizations of this novel structure.

Cryogenic loop heat pipe (CLHP) is one of the key components in the future space infrared exploration system, which enables effective and efficient cryogenic heat transport over a long distance with a flexible thermal link. To realize reliable and long life operation, a CLHP-based thermal control system with redundancy design was proposed in this work, where two nitrogen-charged CLHPs were employed to provide cryocooling at 80–100 K. This study focused on the supercritical startup of the CLHPs with redundancy design, and an extensive experimental study under four possible working modes was conducted. Experimental results showed that with 2.5 W applied to the secondary evaporator, each CLHP could realize the supercritical startup successfully in the normal working mode; however, the required time differed a lot because the difference in the transport line diameter significantly affected the cryocooling capacity to the primary evaporator. In the backup conversion mode, instant switch of the two primary evaporators may cause an operation failure, and an auxiliary operation of the secondary evaporator in advance was necessary to make the primary liquid line filled with liquid. In the malfunction conversion mode, the simulated infrared detector had to be first shut down, but the time needed for the backup CLHP to realize the supercritical startup became obviously shorter than that in the normal working mode, because the primary evaporator of the backup CLHP was always in a cryogenic state. In the dual operation mode, the two CLHPs could realize the supercritical startup simultaneously, but a temperature oscillation phenomenon was observed, which can be eliminated by increasing the heat load applied to the secondary evaporator.

Considering the influence of surface temperature distribution on air convective heat transfer coefficient, the robust tight-coupling method is firstly developed for aircraft thermal anti-icing simulations under icing conditions. To include the effects of the impinging water droplets on the conjugate heat transfer of thermal anti-icing systems, the Messinger thermodynamic model of runback water film is modified and added to the tightly coupled calculation of the external air flow and the internal solid skin heat conduction. Numerical simulations are carried out on an electro-thermal anti-icing system under both dry air and icing conditions, and the main conclusions below can be drawn. First, convective heat transfer coefficient changes slightly with surface temperature near the leading edge, but is obviously affected by temperature distribution in the downstream area. Second, the anti-icing temperature deviations between the predicted value and the experiment date are acceptable and comparable to the calculation results in the literature, verifying the feasibility and effectiveness of the tight-coupling method. Third, compared with the traditional decoupled loose-coupling method, the robust tight-coupling anti-icing method successfully captures the effect of surface temperature on convective heat transfer coefficient, and predicts higher temperature with lower drop rate on the downstream surfaces.

With the development of aviation military and civilian aircraft, the aviation oxygen supply equipment also has higher technical improvements and system innovations. In order to simulate the work of aviation oxygen supply equipment and accurately test the volume flow under high altitude condition, the flow detection method of damping equipment is adopted to detect the flow of oxygen equipment. The gas density correction is not needed during the testing process. The range setting, damper tube structure and other aspects are described, the test principle, identification/calibration methods were studied and verified and the flow test method has been applied for the new test equipment.

In order to support the future thermal control and energy conservation design for the Mars rover, numerical studies on natural convection in CO2 enclosures on Mars’ surface were conducted for both horizontal and vertical enclosures. The parameters are as follows: the atmospheric pressure was 1000 Pa, the gravitational acceleration was 3.62 m/s2, and the Prandtl number was 0.77. The heat flux, temperature, and velocity fields of the CO2 enclosures were obtained with the aspect ratio ranging from 5.56 to 200 and the Grashof number ranging from 430 to 2.6 × 104. It was found that natural convection formed more easily in the horizontal enclosures than that in the vertical enclosures when the enclosures had same thickness. With the increasing thickness of the enclosures, Rayleigh–Bénard convections formed in the horizontal enclosures, while only single-cell convections formed in the vertical enclosures. The heat flux through the horizontal enclosures was greater than that through the vertical enclosures with the same thickness when natural convection formed. The maximum difference between them reached 35.26%, which was illustrated by the field synergy principle. A hysteresis phenomenon of the natural convection dominating the heat transfer was found in the vertical enclosure on Mars’ surface. New values for the critical Grashof number and correlations for the average Nusselt number for both the horizontal and vertical CO2 enclosures on Mars’ surface were also developed.

In this paper, extensive experimental investigation of the heat transfer characteristics of jet impingement on a variable-curvature concave surface in a wing leading edge was conducted for aircraft anti-icing applications. The experiments were carried out over a wide range of parameters: the jet Reynolds number, Rej, from 51,021 to 85,340, the relative tube-to-surface distance, H/d, from 1.736 to 19.76, and the circumferential angle of jet holes on the piccolo tube, θ, from −60° to 60°. In addition, jet impingements with single one, two and three rows of aligned jet holes were all investigated. Experimental results revealed the effects of various parameters on the performance and characteristics of jet impingement heat transfer in the specific structure adopted here, and our insufficient understanding on the corresponding physical mechanism was presented. It was found that the jet impingement heat transfer performance was enhanced with the increase of jet Reynolds number. For single one row of jet holes, an optimal H/d of 4.5 was determined under Rej = 51,021 and d = 2 mm, for which the jet impingement achieved the best heat transfer performance. For two and three rows of aligned jet holes, the Nux curves in the chordwise direction exhibited much different shapes due to different intensity of the interference between adjacent air jets. This work contributes to a better understanding of the jet impingement heat transfer on a concave surface in a wing leading edge, which can lead to optimal design of the aircraft anti-icing system.

As a novel working fluid, microencapsulated phase change material suspension (MEPCMS) exhibits obvious superiority in both heat transfer and temperature control compared with traditional ones. In this paper, extensive experimental study on the performance of a liquid cooling garment (LCG) with the application of this novel working fluid was conducted for future space applications. The main task for a LCG is to efficiently collect, transport and dissipate the metabolic heat produced from the human body. In the experiment, a thermal manikin was employed to simulate the human body, and the performance of the LCG with MEPCMS as the working fluid was evaluated by a variety of aspects such as heat dissipation, temperature control, pump power consumption and thermal comfort under both steady state and transient conditions. Experimental results show that the inlet temperature, mass flowrate and volume concentration of the MEPCMS are three key parameters affecting the performance of the LCG, which can be enhanced significantly by a proper combination of these parameters. Otherwise, the performance of the LCG will deteriorate or even be worse than that using water as the working fluid. When the inlet temperature, mass flowrate and volume concentration of the MEPCMS were selected as 11 °C, 200 g/min and 20% respectively, the heat dissipation of the LCG was enhanced by up to 26% with no obvious increase of the pump power compared with that using water as the working fluid, the temperature distribution in the human body became more uniform, and the capability of the LCG to adapt large heat load change became stronger. This work helps better understand the performance characteristics and contributes greatly to the design of a LCG with the application of the MEPCMS.

In this work, a simulated aircraft fuel tank inerting system has been successfully established based on a model tank. Experiments were conducted to investigate the influences of different operating parameters on the inerting effectiveness of the system, including flow rate of the inert gas (nitrogen-enriched air), inert gas concentration, fuel load of the tank and different inerting approaches. The experimental results show that under the same operating conditions, the time span of a complete inerting process decreased as the flow rate of inert gas was increased; the time span using the inert gas with 5% oxygen concentration was much longer than that using pure nitrogen; when the fuel tank was inerted using the ullage washing approach, the time span increased as the fuel load was decreased; the ullage washing approach showed the best inerting performance when the time span of a complete inerting process was the evaluation criterion, but when the decrease of dissolved oxygen concentration in the fuel was also considered to characterize the inerting effectiveness, the approach of ullage washing and fuel scrubbing at the same time was the most effective.

Deployable radiator is a radiation cooling device which is folded next to side wall of spacecraft while launching, and can expand itself after the spacecraft is on track. It can effectively improve the cooling capacity of the spacecraft. In this paper, the current research and development situations of the deployable radiator were studied and analyzed. Depending on this, design and simulation of the deployable radiator based on the loop heat pipe (LHP) was completed, and the experiment was verified in the vacuum environment. The result showed that the thermal experimental results and the thermal analysis data were in good accordance under hot case as well as cold case. Moreover, the temperature oscillations at the shell of the LHP evaporator and the outlet of saturated vapor and the condenser pipeline were found in the experiment. The phenomenon of temperature oscillations was eliminated experimentally by three measures, which included increasing the input heat load, heating the compensation chamber to control the temperature and cooling the compensation chamber.

Regarding the hypersonic flow over an Earth-reentry inflatable decelerator from 80 km to 50 km altitude, the influence of surface temperature on the aerodynamics and aerothermodynamics of the vehicle were investigated. By applying surface radiative equilibrium condition and several fixed surface temperatures (300 K to 2000 K) as boundary conditions, a system of Navier-Stokes equations incorporating turbulent effects was numerically solved based on finite volume method. The results show that the surface pressure on the forebody and afterbody demonstrates almost coincident distribution at different surface temperatures. With the surface temperature rising, the shear stress on the forebody and afterbody presents most augmentation by 47.35% and 40.77% respectively. The axial force coefficient decreases while the normal force coefficient and pitch moment coefficient increase with a maximum difference of 9%. Additionally, the surface heat flux varies differently on each surface region with the surface temperature rising whereas the total heat load decreases, and the maximum discrepancy of heat flux exceeds 10% on the nose, cone body and centerbody bottom.

Numerical simulation have been undertaken to investigate the forced convention heat transfer of carbon dioxide flow over the Mars rover under the Martian conditions. The study is carried out based on a simplified Mars rover model which is a cuboid with longer and wider top surface and it is performed under the conditions of temperature 160 K and 290 K, Reynolds numbers of 6.64×103 to 7.01×104 and the pressure of 1000 Pa. Different directions of inflow are investigated in order to better reveal the complicated flow fields around the Mars rover. Results show that the heat transfer rate mainly depends on the inflow direction and there is a special angle that the heat transfer rate can reach its maximum. Another point of the conclusion is that the heat transfer coefficient is influenced by the separation and reattachment of turbulent flow. Additionally, the evolution of vortices cause the heat transfer coefficient increasing slightly. Based on the numerical simulation data, a correlation of the force convection heat transfer of carbon dioxide flowing over the Mars rover is developed. Compared with the empirical correlation from literature, the present correlation is more applicable under the Martian conditions.

Electronic oxygen regulator has a high sensitivity to respiratory regulation, which is the development direction of future aircraft oxygen system. China has carried out the principle research of the electronic oxygen regulator, the pressure control cavity is a key component of the electronic oxygen regulator in the establishment of the excessive pressure. This paper introduces the operating principle of the pressure control cavity and the structure design of double solenoid valve control adopted the pressure control cavity. Theoretical calculation and experiment verification show that the control structure of pressure control cavity is reasonably designed, and meets the need of the electronic oxygen regulator when establishing the excessive pressure.

The flexible inflatable aeroshell naturally deforms under aerodynamic loads during atmospheric entry. In this paper, we investigate the hypersonic aerodynamics of an inflatable aeroshell forebody with undulated surface deformation in the continuum and near-continuum flow regimes, where flow physics is complex involving rarefaction and thermochemical nonequilibrium effect. The aeroshell forebody shape is originated from a model of 8.3 m diameter stacked-tori hypersonic inflatable aerodynamic decelerator, and a series of surface deflections is generated by a parametric method based on the consistency of undulation with underlying structure. The trajectory points of 90, 80, and 68 km altitudes in a low Earth orbit reentry are considered. The thermochemical nonequilibrium Navier-Stokes equations with slip and no-slip boundary conditions are numerically solved to characterize flowfields and provide aerodynamic predictions. The results demonstrate that the undulated surface deflection induces local fluctuations of pressure and local Knudsen number, while the local variations of aerodynamic quantities create minor difference in total drag. The surface friction and drag are dependent on boundary conditions, i.e., either slip or no-slip model. The results also indicate that it is required to employ slip boundary conditions for the accurate prediction of friction and drag in the near continuum regime.

A tight-coupling heat transfer method is proposed in this paper for the analysis of the performance of the hot air ice protection system. The Eulerian method is used for the calculation of the local collection efficiency. The external heat transfer coefficient was computed using the boundary layer integrated method. The thermal conductivity within the wing skin and the internal heat transfer between the hot air flow and the skin were computed using the computational fluid dynamic method. At the same time, the external heat flux boundary condition and the surface temperature were updated automatically by user-defined function to drive the iteration of the tight-coupling heat transfer calculation. The surface temperature results in dry air condition are compared with flight test data and show agreement. The maximum temperature difference between the simulation and the test is 11.5 K. In addition, the method proposed in this paper is applied to the wing hot air ice protection system of a real aircraft in icing conditions. It is found that the surface temperature ranges from 3 to 30 ℃ under certain flight and icing conditions. Larger droplet diameter or larger liquid water content leads to more runback water which changes the surface temperature not much when the parameters of the bleeding air are the same.

To calculate ice shapes at engine inlet, a fully three-dimensional ice accretion model based on the classical Messinger thermodynamic model is established, which takes the effect of runback water into consideration. Considering the air frictional force as the main factor to drive the surface runback water streaming, the shear stress on the inlet surface is used to determine the film flow direction and the mass flux distribution of the runback water. To solve the ice accretion model of runback water on the three-dimensional surface, an effective approach of circularly checking the condition of control volume is developed, which can quickly perform the icing calculation of all grids on the surface. The ice accretion calculation of a certain three-dimensional engine inlet is carried out and the ice shapes are obtained using the code written in this paper. Comparisons between computational results and FENSAP-ICE results show that the ice profiles of both codes are in good agreement except for slight differenc...

Many flight and icing conditions should be considered in order to design an efficient ice protection system to prevent ice accretion on the aircraft surface. The anti-icing heat load is the basic knowledge for the design of a thermal anti-icing system. In order to help the design of the thermal anti-icing system and save the design time, a fast and efficiency method for prediction the anti-icing heat load is investigated. The computation fluid dynamics (CFD) solver and the Messinger model are applied to obtain the snapshots. Examples for the calculation of the anti-icing heat load using the proper orthogonal decomposition (POD) method are presented and compared with the CFD simulation results. It is shown that the heat loads predicted by POD method are in agreement with the CFD computation results. Moreover, it is obviously to see that the POD method is time-saving and can meet the requirement of real-time prediction.

A closed-loop two-phase mini-channel-based heat sink driven by a micro-gear pump was developed in this work. Using water as an example, experiments were performed in two micro-channel heat sinks under the conditions of initial pressure of Pin = 34–113 kPa, mass velocity of G = 19–468 kg/m2s, outlet quality of xe,out = −14–66%, and heat flux of q″ = 0–230 W/cm2, which covered single-phase flow, subcooled flow boiling, and saturated flow boiling regions. The results showed distinctive differences between the subcooled and saturated boiling regime and revealed that the influence of the system pressure. The experimental data were also compared to a boiling mechanism demarcation map and assessed against some empirical correlations, which suggests some uniqueness of the current heat sink associated with flow boiling at the mesoscale.

To satisfy the cooling demand from a particular space application, a closed-loop two-phase minichannel heat sink driven by a micro-gear pump was developed in this work. The cooling system was made of oxygen-free copper containing 14 parallel channels of the dimension of 0.8 mm (width) × 2 mm (depth) × 20 mm (length). The cooling performance of FC-72 and water was evaluated by the heat transfer coefficient and the pressure drop under different conditions. The analysis shows that both coolants can be used to achieve the design criteiron, that is, a maximum heat dissipation rate of 100 W/cm2 while keeping the peak heating surface temperature below 90°C. However for FC-72, the limitation is in the critical heat flux, which occurs easily under low-flow-rate conditions that deteriorates the performance of the cooling system. The comparion of boiling heat trnsfer coefficient shows that none of the compared correlations can predict both FC-72 and water data well, with the corrleation developed in 1996 by Tran, Wa...

A flow and heat transfer model in the combined thermal and acoustic environment was established. The temperature distributions of the test pieces in different heat flux densities with two installation locations of the graphite heater (inside the duct and on the duct) were calculated and analyzed by the CFD method combined with DO radiation model. The results show that the established model can well simulate the flow field and heat transfer in the combined thermal and acoustic environment simulation test-bed. It is shown that the temperature of test piece increases with the increase of the heat flux, but to reach the same temperature, the placement that the heater and the test piece are placed on top surface of the duct requires less heating power, meanwhile the temperature of the test piece has better uniformity. When the heater is placed inside the wind duct, the temperature of the test piece requires heating heat flux 300 kW/m2 to reach 1000 °C. While when the heater is placed on the duct, the test piece temperature reaches 1000 °C with only 50 kW/m2 heat flux density needed. The numerical simulation method and results of flow filed and heat transfer under the combined thermal and acoustic environment in this paper may provide a guidance for building test-beds and setting the parameters of the tests and experiments.

With the need of airplane for long endurance flight, the pilot raises new demand for the gas-water content of the product generated by OBOGS. For this reason, this paper conducted experiments with the system on the two types of oxygen-rich molecular sieves — NaX type and LiX type, which are used currently for domestic OBOGS under different operating conditions, conducted sampling analysis of the bleed air and product gas and obtained H 2 O content of the product gas, which provides reference for OBOGS researches.

The mathematical models and a numerical code for numerical simulation of a thermal anti-icing system are presented in this article. Mass conservation is applied to the runback water flow. An energy balance is imposed on the airfoil skin including the water flow. The heat transfer coefficient distributions are obtained using the boundary layer integral method. The external flowfield and the local water collection efficiency data are predicted using an Eulerian method, based on a computation fluid dynamic code and its user-defined functions. Given the input of the electrical power density distribution, the numerical code is able to calculate airfoil equilibrium surface temperature, mass flux of runback water, runback ice mass flux, and range if happens. A user interface is developed to integrate the computation fluid dynamic code to achieve a method for the analysis of a thermal anti-icing system. All the numerical results are compared with both experimental data and other numerical results presented in the literature.

A prediction method of surface temperature and runback ice for a three-dimensional hot air anti-icing system was proposed. Computational approach to realize this method was introduced. Both the external and internal flows were separately calculated, results of which were set as boundary conditions of heat conduction computation in airfoil skin. The results of external and internal flow calculations show that the effect of surface temperature on convective heat transfer coefficients and local droplet collection efficiency is negligible and the calculations can be decoupled. The prediction method based on heat flux was used to calculate surface temperature and runback ice results. The results show that, the effects of LWC and Mach number are much more significant than the effect of external flow temperature. The surface temperature at impinging interaction point is more sensitive to the change of external conditions than that at stagnation point. The surface temperature changes significantly with changing Mach number because both the mass rate of droplet and the impact limit are changed.