Application requirements for avionics are often very strict. For example, the heat sinks of avionics need very good temperature uniformity, but the flow rate of coolant is very restricted. In addition, the use of micro-channels is not recommended due to the potential clogging issue. Considering these design requirements, we will discuss a multiple-objective optimal design method to obtain a good stacked mini-channel structure for avionics applications. In our thermal design, the design variables are the mini-channel geometry parameters. Temperature uniformity, entropy generation, max temperature of heat sink and pump work are chosen as the objective functions. A Multi Objective Genetic Algorithm (MOGA) and Fluent solver are used together to minimize multiple objective functions subject to constraints, and locate the Pareto front. By analyzing the multiple objective optimal results, we can draw the conclusion that the objective functions of T_max and s_g have same effect on the optimization, and the multiple optimal results are a set and not a single value. If mostly focusing on the temperature uniformity, we can recommend some optimal structures to design a stacked mini-channel heat sink.

This paper deals with the modeling of the thermal load and the simulation of thermal dynamic characteristics for the hydraulic system of a large airliner in a full mission profile. Firstly, the formation mechanism of the thermal load in the hydraulic system is analyzed, and thermal dynamic modeling is conducted of the hydraulic components of an hydraulic system with an immersed heat exchanger employing the lumped parameter thermal node method and oil temperature and power loss of each key node within the hydraulic system within a full mission profile. Then, a thermal dynamic simulation model based on MATLAB/Simulink is established, and the temperatures at the nodes of different components and the absorptive capacity of the fuel heat sink in the thermal management module are calculated. The simulation results show that the thermal management scheme of the heat exchanger, located in the return oil pipeline of the hydraulic piston pump housing and immersed in the central fuel tank, can dissipate the thermal load of the system. This work is of important significance for temperature analysis and thermal load dissipation of the airliner hydraulic system.

Silica gel is a typical porous desiccant material. Its adsorption performance for gaseous air pollutants was investigated to determine its potential contribution to reducing such pollutants. Three gaseous air pollutants, toluene, carbon dioxide, and methane, were investigated in this paper. A thermogravimetric analyzer was used to obtain the equilibrium adsorption capacity of gases on single silica gel particles. The silica gel adsorption capacity for toluene is much higher than that for carbon dioxide and methane. To understand gas pollutant thermal desorption from silica gel, the thermogravimetric analysis of toluene desorption was conducted with 609 ppm toluene vapor at 313 K, 323 K, and 333 K. The overall regeneration rate of silica gel was strongly dependent on temperature and the enthalpy of desorption. The gas pollutant adsorption performance and thermal desorption on silica gel material may be used to estimate the operating and design parameters for gas pollutant adsorption by desiccant wheels.

The dust storm on the surface of Mars is a severe threat to Mars exploration missions. Taking adequate measures to avoid the impact of the harsh wind-blown dust environment is indispensable. Ground simulation of the Martian high-speed windblown dust environment is helpful for analysis of the environmental effects and evaluations of the suitability of the components and materials. In this paper, a novel reflux subsonic low-density dust wind tunnel is presented to simulate the high-speed windblown dust environment of the Martian atmosphere with a velocity of more than 100 m/s. The sand and dust are fed into the wind tunnel through the ejector assembly together with the compressed gas, resulting in high uniformity of particles in the test section. The construction design of the Mars wind tunnel is introduced. The key parameters, which are the nozzle parameters and the contraction curve, are discussed in detail. The convergent nozzle is most suitable for the ejector assembly. Moreover, the bicubic curve is selected as the contraction curve. The gas-particle two-phase computational fluid dynamic (CFD) simulations demonstrate the rationality of the wind tunnel design.

The aircraft engine bleed air simulation thermodynamic laboratory simulation parameters include the bleed air pressure and temperature. However, existing laboratories cannot carry out the dynamic test of the engine bleed air simulation. In the engine bleed air simulation dynamic test, the temperature control has the characteristics of strong coupling and nonlinear and large inertia. The conventional control strategy cannot solve the contradictions of the response speed and stability of the system. Moreover, the dynamic control of the pressure and temperature involve strong coupling. That often leads to the failure of control decisions. Therefore, there is still no relevant report on the laboratory equipment used for the engine dynamic bleed air simulation. According to the above problem, this study adopted heat exchangers for indirect heating to reduce the coupling of dynamic control between temperature and pressure. Specifically, to take into account the rapid response and stability of the system, this study used the lookup table-based PID (LPID) controller to control the temperature and pressure of the bleed air simulation test. The dynamic test errors were within 10%, and the steady-state accuracies were within ±2%. The simulation software results and the engine bleed air simulation test results showed that temperature and pressure control systems based on the LPID controller have advantages: high control precision, a low overshoot amount, a fast response, and a high stability.

Hypersonic vehicles are an important area of research in the aerospace field today. One of the important issues is the power of the engine. In order to achieve large-span flight speeds, a more efficient approach is to use combined power systems. However, the problem of pre-cooler icing can occur in combined engine applications. The flow in the pre-cooler is extremely complex. Outside the tube is the high-temperature wet air entering from the engine intake, and the tube cooling is the ultra-low temperature cooling medium. Icing not only increases the heat exchange resistance of the pre-cooler during operation and affects the heat exchange performance of the pre-cooler, but also causes a large total pressure loss, resulting in a degradation of the engine performance. There is a lack of research on the icing law of the pre-cooler under different parameters. Therefore, it is necessary to conduct a corresponding numerical calculation study on pre-cooler icing and explore the influence of various influencing factors on icing. In this paper, a mathematical model of icing (frost) is established for the frosting phenomenon that may occur during the operation of the pre-cooler. Additionally, the principle of heat and mass transfer in the icing process is described by the mathematical model, and the influence of different parameters on the frosting parameters is explored by using the computational fluid dynamics (CFD) method. The law of tube bundle icing under different parameters was calculated, and the variation laws of frost layer morphology and wet air pressure drop were obtained. The laws of tube bundle icing under different parameters were calculated, and the changes in frost layer pattern and wet air pressure drop when each parameter was changed, which can provide guidance for the design and application of pre-coolers in the future.