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Pulsed Surface Arc Discharge (PSAD) actuators can improve the aerodynamic performance of aircraft due to their fast response speed and wide frequency band. Understanding the flow structure induced by arc discharge and its transfer process is the key to the sensible use of plasma excitation for flow control. This paper has investigated the dynamic heat and mass transfer process of charged particle clusters generated by arc discharge theoretically and experimentally. The results show that the clusters are composed of charged particles, and the thermal effect is the macroscopic manifestation of the thermal motion of charged particles. The heat and mass transfer processes of charged particle clusters can be expressed as a function of temperature. During heat and mass transfer, the transfer mode of the plasma cluster composed of charged particles will change. The plasma cluster will change from forced convection in the initial stage to natural convection in the final stage, and the Richardson number is different in different stages. The heat released by the gas discharge and the heat released by the plasma mass transfer will form accumulated heat in the space and change the temperature difference, thereby affecting the process of heat and mass transfer. This paper has constructed a model describing pulsed plasma clusters' dynamic heat and mass transfer. Furthermore, the dynamic mode decomposition results of the theoretical model function show that plasma clusters mainly conduct the mass transfer through low-frequency and heat transfer through high-frequency modes. [ABSTRACT FROM AUTHOR]

Nitrogen has been widely used as an expendable coolant in aviation and aerospace fields to cool the high-temperature components in a thermal management system. This paper presents a concept of a thermal management system for hypersonic vehicles that integrates various cooling technologies using supercritical nitrogen as heat sink. Since the thermophysical properties of supercritical nitrogen vary acutely with temperature, existing heat transfer empirical models fail to accurately evaluate the heat transfer coefficient of supercritical nitrogen during the parameter design of the thermal management system. Therefore, a new correlation specialized for supercritical nitrogen is proposed relying on a heat transfer experimental database including 784 data points compiled from seven published papers and the computer optimized dimensionless groups, which commendably solve the considerable deviation caused by the thermophysical property distortion and buoyancy effect. The new correlation has a mean absolute deviation (MAD) of 3.9% and improves the prediction accuracy of supercritical nitrogen heat transfer calculation remarkably by predicting 94.9% of the entire database within ±10% error band, based on which the heat transfer design method of the integrated thermal management system is discussed. [ABSTRACT FROM AUTHOR]

In this paper, a computational investigation of a rotorcraft avionics-bay cooling system is carried out. The introduced avionics cooling system utilizes a forced-convection method in which the ambient air is supplied to the avionics bay by a fan and then exhausted back into the ambient after cooling the equipment inside. The aim of this system is to keep the air temperature in the vicinity of the avionics equipment below the operational temperature limits. Depending on the locations of the fan and exhaust, local hot zones may form near some of the equipment. In order to overcome this issue, the fan capacity must be designed to provide a sufficiently high mass flow rate so that the temperature limits are not exceeded, while an excessive amount of cooling should be avoided to reduce power consumption. In this study, the effects of the fan and the exhaust locations on the amount of required mass flow rate are investigated. A prediction function is built by using Gaussian Process Regression method to predict the avionics surface temperatures. The Gaussian Process Regression method is trained by the results obtained from a large number of computational fluid dynamics (CFD) analyses. The prediction function is later used to determine the required mass flow rate depending on the fan and the exhaust locations. It is found out that the required mass flow rate changes significantly as the fan location changes. On the other hand, the exhaust location has a relatively lessened effect on the required mass flow rate. It is observed that depending on the locations of the fan and the exhaust, the required mass flow rate could be reduced to around half of its value. The amount of decrease in the electrical power consumption is even more significant. [ABSTRACT FROM AUTHOR]

Performance of radial–axial clearance rim seal (RACS) in realistic working condition is investigated and compared with axial-clearance rim seal (ACS), radial-clearance rim seal (RCS) in this paper. Authors use the numerical method of conjugate heat transfer for calculation to accurately take heat transfer between the rotor–stator cavity flow and the solid discs into account. Results show that seal effectiveness and cooling effectiveness of RACS are the best when compared with ACS and RCS, the minimum mass flow rate for seal of RACS is 75% of that of RCS, and 34.6% of ACS. RACS has higher air-cooled aerodynamic efficiency, minimizing the mainstream performance penalty when compared with ACS and RCS. Corresponding to the respective minimum mass flow rate for seal, the air-cooled aerodynamic efficiency of RACS is 23.71% higher than that of ACS, and 12.79% higher than the RCS. Finite element analysis in turbine disc shows that RACS minimizes the flow rate of cooling air required for suppressing the radial growth of turbine disc in the three rim seal types. Mass flow rate of the required cooling air of RACS is approximately 40.9–52.9% of that of ACS, and 70–75% of RCS. [ABSTRACT FROM AUTHOR]

The liquid droplet radiator (LDR) as a frameless heat removal system is considered as a promising solution for the waste heat dissipation of megawatt aerospace applications. The radiation heat transfer and evaporation characteristics of the LDR working fluid are tightly coupled with the operational performance of the liquid droplet generator. In this paper, the effects of operational parameters of droplet generator on the radiative heat transfer and evaporation loss rate are clarified, including the pressure disturbance frequency, pressure difference between the inside and outside of the droplet generator. It is observed that higher coolant mass flow rate does not assure higher heat transfer power for the LDR. For cases with higher pressure difference, where the coolant flow rate is higher consequently, the evaporation loss rate increases continually with pressure difference while the heat transfer power does not increase any more. The disturbance frequency has inconspicuous impact on the evaporation characteristics. Besides, cases with higher pressure differences have wider suitable frequency ranges according to the droplet formation restriction. Generally, a relatively higher pressure difference coupled with a higher disturbance frequency (such as 0.3 MPa and 12 kHz) is suggested to achieve a higher heat transfer power and low evaporation loss rate. This paper may provide favorable reference for determining the operational parameters of the droplet generator in consideration of radiation heat transfer and evaporation characteristics. [ABSTRACT FROM AUTHOR]

The study on the adhesion and rebound behavior of high-temperature alumina droplets impacting on the cold wall is very meaningful for the accurate prediction of slag deposition and internal flow field in solid rocket motors. In this paper, based on high-temperature alumina droplet impact experimental system, the experiment of alumina droplet impacting on the cold wall was carried out, and the temperature response of alumina droplet during the impact is obtained by the colorimetric thermometry. The study shows that the high-temperature alumina droplets impacting on the cold wall in a certain range of Oh number and Re number exhibit two kinds of behaviors: rebound and adhesion; based on experiments, the distribution laws of rebound and adhesion behaviors of alumina droplets impacting the cold wall are obtained; based on the energy conservation model, and introducing the heat transfer model between the high-temperature droplet and the cold wall, the maximum spreading coefficient model of alumina droplet impact on the cold wall and the prediction model of rebound and adhesion results were established, and the maximum spreading coefficient model prediction results were in good agreement with the experimental data. [ABSTRACT FROM AUTHOR]

In our previous paper (doi: 10.1016/j.ijheatmasstransfer.2019.119016), the gas-kinetic BGK scheme was used to simulate both hypersonic flow and structural heat transfer, between which loose coupling was adopted for solving two-dimensional fluid-thermal problems. To model the unitive and continuous fluid-thermal processes more actually, this paper presents a BGK scheme-based integrated algorithm and the research object is extended to three-dimensional hypersonic vehicles. To start with, an appropriate BGK model is constructed for the unified macroscopic governing equations, the effectiveness of which is verified by the Chapman-Enskog expansion analysis. Then following similar procedures in the original scheme, the gas distribution function at each cell interface is derived, from which the numerical fluxes can be evaluated. The dual-time stepping approach, together with an implicit JFNK–BGK method, is applied to efficiently obtain time-accurate solutions. With an effective parallelization strategy that is in line with the integrated algorithm, the computational efficiency is improved. As another aspect which distinguishes from the loosely-coupled methods, the fluid-structure interface conditions are also addressed. For validation, several test cases are investigated, including an individual hypersonic flow, an individual transient heat transfer, and two fluid-thermal problems. By comparison with other methods and analysis of the computed results, features of the developed method are demonstrated. [ABSTRACT FROM AUTHOR]

A novel multi-cavity slashface gap with different cavity obstruction modes, compared with a typical single-cavity slashface, is designed to arrange coolant flow on a gas turbine endwall. Solving the steady Reynolds averaged Navier–Stokes equations, the interaction condition between the coolant jet and cascade secondary vortices is studied. The corresponding film cooling and aero-thermal performances of the endwall are evaluated. The obtained results indicate that the dual-cavity slashface urges the coolant from the fore cavity to stagnate on the obstruction and to be redirected toward the spanwise direction. Consequently, the coolant downstream migration tendency near the obstruction is restricted, and the cooling effectiveness downstream of the obstruction increases. When the blowing ratio (M) is 0.5, the cavity obstruction is most effective when placed at 30% L s and 50% L s in redirecting the coolant to eject out more quickly. The endwall cooling effectiveness in the vicinity is increased from 0.2 to 0.4. The influence of the obstruction is the most moderate when M = 1.0. The cooling effectiveness upstream of the obstruction is only increased by 0.06 in the DC-10% case. Another mainstream ingestion region appears downstream of the obstruction, and the lack of coolant decreases the average endwall cooling effectiveness by 0.04. At a high blowing ratio (M = 1.5), the ingestion and ejection velocities at the slashface are the highest, and the obstruction can only influence the region of z / C ax < 0.4. After penetrating directly into the mainstream of the coolant, an uncovered region is formed on the endwall, and its average cooling effectiveness is increased by up to 0.08 with an increasing net heat flux reduction of 0.2 in the DC-30% case. This paper provides insights into elaborate purge flow control for turbine cooling designers. [ABSTRACT FROM AUTHOR]

Heat pipe technologies have numerous applications under microgravity conditions in aerospace, from miniature devices in laptops used aboard the International Space Station (ISS) to heat transport systems in satellites. Advances in understanding the behavior of two-phase flow systems under microgravity conditions could lead to higher-efficiency devices and improved heat-exchanger designs. Under normal gravity conditions, heat pipes can function properly without any kind of capillary structure (thermosyphons, vertical position) or they can be aided by a wick material in the horizontal position, allowing the return of the condensate from the condenser to the evaporator. However, gravitational acceleration can significantly alter the flow regime, masking some effects that are not observed under microgravity conditions. This paper presents an experimental analysis of three miniature heat pipe technologies developed in Brazil for thermal management and heat dissipation of electronic devices under gravity and microgravity conditions. Different geometries of heat pipes charged with methanol were tested under microgravity conditions aboard a sounding rocket. Results demonstrated that the mini heat pipes present low thermal resistances for both gravity and microgravity conditions. [ABSTRACT FROM AUTHOR]

This paper introduces and develops a novel curved pin fin for heat transfer enhancement of aeroengine turbine blade. The effects of inclination angle and vertical segment length of curved pin fin on heat transfer performance are studied under stationary and rotating condition. Four cases with different inclination angle (90°, 75°, 60° and 45°) and four cases with different vertical segment length (9, 7, 5 and 3) are studied for rotation number varying from 0 to 0.5. Numerical results show that curved pin fins with smaller inclination angle and shorter vertical segment are characterized by lower friction factor and better thermal performance. Compared to upright pin fin, the thermal performance factors for curved pin fins with inclination angle γ = 45 ° and vertical segment length H v = 3 are improved by 6.1% and 14.4%, respectively. Under rotating frame, the effects of Coriolis force are suppressed by curved pin fin, and lower heat transfer enhancement is produced consequently. Moreover, curved pin fin with smaller inclination angle and shorter vertical segment is featured with relatively lower sensitivity with respect to rotation number. [ABSTRACT FROM AUTHOR]

Abstract: This paper deals with design activities performed by CIRA in the frame of its flying test bed vehicle development, named FTB-X. The re-entry scenario with the corresponding loading environment for the proposed vehicle, devoted to research activities in re-entry technologies, is reported and analyzed. The hypersonic environment characteristics of FTB-X are investigated on a trajectory-based design approach by means of several engineering and CFD analyses. Being a phase A design, the air is modeled only as a perfect gas and the analyses were based on laminar flow conditions, while only a preliminary analysis of real gas and turbulence effects has been performed in time. The results show that the aerodynamics derived from engineering design approach is sufficiently accurate for preliminary analysis purpose; moreover, surface heat loads were computed on the vehicle configuration for thermal shield design scopes. [Copyright &y& Elsevier]

Abstract: Simulation of the temperature distribution on the rocket nozzle wall in high temperature environment is most important for assessing safety and reliability of the nozzle. For that purpose, the coupling simulation of heat transfer and transient temperature of the rocket nozzle wall is carried out in this paper. Equations of radiative heat flux on surfaces in an enclosure with inhomogeneous, participating media is developed to compute the radiative heat flux. The convective heat transfer considering ablation of the C/C material has been studied. And radiative transfer between the plume and nozzle wall has been analyzed with the net radiation computing method in one enclosure. Also the differential equation for two-dimensional transient conduction in the cylindrical coordinate has been established, discretized and solved. The convection and radiation heat fluxes, also the transient temperature along the wall length and thickness of the composite nozzle wall after startup of the engine are investigated. Two benchmark problems have been taken, one is the nozzle of an experimental rocket engine fabricated with stainless substrate and a ZrO2 coating, the other is constructed with multi-layer composite materials. The computing results show that: plume heating on the throat section is most serious and wall temperature in that place is much higher than the average temperature, and heat insulation of the throat is very important for rocket nozzle. The typical character of transient temperature is steady increase with time after startup of the rocket engine, both on the internal and outside walls of the rocket nozzle while the former increases rapidly. On the same section, the temperature decreases along the thickness from internal to outside of the wall, the difference is over 1000 K due to the heat insulating layer in the nozzle material. The coupled simulation method and related results in the present study can be best used to predict the safety performance of the rocket nozzle in startup stage. Results of coupling simulation method are compared with available results in references which show reasonable agreement. [Copyright &y& Elsevier]

The solid-fuel ramjet (SFRJ) motor is the most favorable air-breathing propulsive system and a suitable option for extending the speed and range of some military applications as in ramjet-powered missiles. However, SFRJ encountered severe challenges associated with low regression rate, as this latter plays an important role on the performance of the solid-fuel ramjet motor. To improve the solid-fuel ramjet's regression rate, this paper proposes new configuration which employs two solid fuel tubular grains with keeping overall simplicity of classic SFRJ motor, then the new design is numerically investigated by a developed in-house CFD code. The developed code is used to estimate Reynolds-averaged Navier–Stokes equations of reacting, unsteady, turbulent, and swirling flow which is validated and verified for several test cases. The numerical results for proposed designs are compared with classic SFRJ. The results reveal significant increase in solid fuel regression rate and heat transferred into grain's interior surface. Moreover, the thrust, specific impulse, and characteristics velocity are increased as well. [ABSTRACT FROM AUTHOR]

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

Abstract: In this paper, a finite volume method was developed to compute the infrared radiative signatures of both liquid and solid rocket plumes, and to investigate the radiative signatures on two wavebands, 2–6 and 8–12 μm bands, respectively. This method is based on the radiative transport equation for local thermodynamic equilibrium, which accounts for radiative heat transfer in inhomogeneous emitting, absorbing, and non-isotropic scattering media. The radiative properties of combustion gases are computed by using the Elsasser narrow band model, and that of particles are computed by using the Mie scattering theory. A numerical method is applied to solve the following three problems. The first one is the computation of the radiative heat transfer in an absorbing, emitting and isotropic scattering homogeneous cylindrical medium in which the radiative flux is presented. The other two are the computations of the infrared radiation of plumes from the liquid and solid rockets in which both the source and apparent radiant intensities are presented as results. The atmospheric absorption effect on the radiation of plumes is also considered by calculating the atmospheric transmittance using MODTRAN. Some of these results, compared with those available in the literature, have shown satisfactory agreement in most cases. [Copyright &y& Elsevier]

Abstract: This paper presents RANS simulations of the low-Reynolds-number flow past an SD7003 airfoil at , where transition takes place across a laminar separation bubble. The transition prediction procedure using an approximate envelope method as well as a linear stability solver is discussed. The numerical results are validated against PIV- and force measurements obtained in several wind- and watertunnels and are also compared to XFOIL results. Good agreement is found within the operational range of the airfoil. [Copyright &y& Elsevier]

An improved local correlation-based intermittency transition model is developed for high-speed flow heat transfer in this paper. Based on the intermittency transition model, two improvements have been proposed. One is introducing a new correlation of the momentum thickness Reynolds number for high-speed boundary layer into the transition onset function. In this new correlation, the free stream Mach number and the free stream Reynolds number are used to reflect the compressible effects. The other improvement is the modification of the intermittency factor transport equation to include physical information regarding the transition processes. Several test cases containing supersonic and hypersonic flat plates, a slender cone at different Reynolds numbers, the X-51A forebody, and the Hypersonic Inflatable Aerodynamic Decelerator are conducted to assess the applicability of the improved intermittency transition model. The numerical results show the ability of the improved model to describe the high-speed boundary layer transition processes. However, further research is needed to reproduce the overshoot phenomena and verify the capacity of the improved model on complex cases. [ABSTRACT FROM AUTHOR]

This paper aims to predict the transient temperature fields of rocket combustion chambers. This is important for design and lifetime estimation. A thermally coupled simulation (conjugate heat transfer) model was developed and validated using a combustion chamber experiment with rectangular cross-section and the propellant combination gaseous oxygen–methane. The simulation is based on a thermal finite element model of the structure, Reynolds-averaged Navier–Stokes simulations of nine cooling channel flows and a Reynolds-averaged Navier–Stokes simulation with flamelet extension of the hot-gas flow. A pseudo-transient, implicit Dirichlet–Neumann scheme is used for calculating the equilibrium of structure and fluids. A comparison with experiment shows good agreement with the axial, vertical and transverse temperature distribution in the structure, the transient behavior and the wall heat flux through the cooling channels. It is shown that wall functions in the coolant domains can efficiently decrease the computational effort. [ABSTRACT FROM AUTHOR]

Huge aerodynamic drag and severe aeroheating are the inevitable challenges for hypersonic vehicles, this paper proposes a novel combined spike-aerodome and lateral jet strategy for drag reduction and thermal protection. Meanwhile, the hypersonic flow characteristics and structure thermal response are numerically investigated by an in-house code. The obtained results demonstrate that this novel strategy not only further reduces the total drag, but also provides excellent thermal protection for both the blunt body and spike rod. Then the effects of lateral jet pressure ratio, spike length and aerodome diameter on the flow properties, drag and heat reduction performance are thoroughly explored. Increasing the lateral jet pressure ratio can effectively enhance the thermal protection. Surprisingly, in a certain range, higher lateral jet pressure ratio even leads to larger pressure peak value on the blunt body. Increasing the spike length would enhance the heat transfer along the blunt body surface, and the total drag changes non-monotonically as the spike length varies. The total drag rises significantly when the aerodome diameter increases, and it can be increased by 40.1% when the aerodome diameter ratio increases from 0.15 to 0.5. Besides, increasing the aerodome diameter is beneficial to remarkably enhance thermal protection. [ABSTRACT FROM AUTHOR]

In this paper, four turbulence models are compared on the basis of the predictions they produce for the flow and heat transfer in the avionics bay area of a rotorcraft. The turbulence models studied are the standard k - ε , Renormalization Group (RNG) k - ε , realizable k - ε , and Shear-Stress Transport (SST) k - ω model. The avionics bay used in the study houses avionics equipment mounted on the floor and on a rack inside the bay to mimic a realistic distribution of equipment in actual rotorcraft. The avionics bay incorporates a cooling system consisting of a fan that intakes ambient air via a forced-convection method and an exhaust, with the purpose of keeping the avionics equipment temperatures below their operational limits. The turbulence levels, flow and thermal fields are investigated and compared in order to quantify the differences between the predictions of the turbulence models used in computations. It is observed that the use of the standard and realizable k- ε models mostly produce similar flow and thermal results across the bay area, while the largest differences between the predictions are found with the RNG k - ε and the SST k - ω turbulence models especially at the locations in the vicinity of the fan impingement. Comparing the avionics equipment surface temperature predictions, the standard and realizable k - ε turbulence models are observed to produce slightly more conservative results. The average variation between the surface temperature predictions by all models is less than 3.5 K, indicating that this parameter is considerably insensitive to the selected turbulence model. A close examination of the fan jet region starting from the fan inlet revealed generally similar predictions of flow and thermal features by all turbulence models. The maximum variation in the velocity and temperature predicted by all turbulence models is 2.2% and 0.1% of the fan inlet conditions, respectively. Large differences in turbulence intensity were observed depending on the location downstream of the fan jet and the turbulence model used. [ABSTRACT FROM AUTHOR]

Understanding of flow boiling heat transfer under hypergravity is needed due to its applications to modern flight vehicles. Few investigations on this issue have been reported. The present paper presents the experimental results of R134a flow boiling heat transfer in 1.002 and 2.168 mm ID tubes under hypergravity up to 3.16 g. The results reveal effects of gravity, channel size, heat flux, mass flux, quality, and pressure on and their interrelations in flow boiling heat transfer. Gravity has strong effects on flow boiling heat transfer. The heat transfer coefficients (HTCs) under hypergravity levels up to 3.16 g are normally greater than those under Earth's gravity. The heat transfer characteristics in the 1.002 and 2.168 mm tubes are very different, indicating that gravity effects on flow boiling heat transfer strongly interrelate with channel size. Also, gravity effects on flow boiling heat transfer are influenced by heat flux, mass flux, and pressure, and somewhat related to quality. [ABSTRACT FROM AUTHOR]

Heat transfer of the hydrocarbon fuel at a supercritical pressure is a key process in the aerospace engine cooling applications. A CFD model is applied in this paper to study flow dynamics and heat transfer of the aviation kerosene RP-3 in a horizontal square cooling channel under asymmetric heating and buoyancy effects at various supercritical pressures. The turbulent fluid flows are handled by the standard k – ε turbulence model with an enhanced wall treatment, and the strong thermophysical property variations are calculated using the extended corresponding state approaches and a four-component surrogate model of RP-3. The effects of secondary flows and heat flux redistribution induced by buoyancy on supercritical-pressure heat transfer are analyzed. Results indicate that drastic variations of the fuel density with temperature at a supercritical pressure of 3 MPa induce strong buoyancy effect on heat transfer. In the top heated case, secondary flows carry the relatively cold fluid to the side and opposite walls, thereby increasing their fin effectiveness in heat transfer. As the operating pressure increases from 3 to 5 MPa, the density variation of RP-3 is decreased, consequently leading to the weakened buoyancy effect. As the solid thermal conductivity increases from 20 to 100 W/(m-K), the surface heat flux redistribution is dictated by heat conduction in the channel walls, and therefore, the buoyancy effect is significantly reduced. Numerical results herein could help elucidate the heat transfer characteristics of RP-3 in practical engine cooling processes. [ABSTRACT FROM AUTHOR]

This is an erratum to our published paper [1] . [ABSTRACT FROM AUTHOR]