A novel adaptive Kriging controller for a class of discrete-time nonlinear systems with known control gains is proposed. The dynamic models of the systems may involve uncertainties that can be described as realizations of Gaussian random functions. The Kriging system is used to approximate the uncertainties, and an adaptive control law is established so that the tracking error remains bounded. To eliminate the infinite increase of the data size in the Kriging system, which usually occurs in non-parametric estimator-based adaptive controllers, a recursive Kriging estimation algorithm is proposed. In addition, the application of the proposed control scheme is extended to the case of uncertain control gain. Finally, some numerical examples are given to show the viability and effectiveness of the proposed control approach.

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.

Environmental wind tunnels for high-speed trains play a significant role in their development. The cooling system of the wind tunnel poses a challenge as it requires lower temperatures and a higher cooling capacity during operation. The conventional approach to wind tunnel refrigeration uses evaporative cooling, which is less efficient at low temperatures and comes with environmental and safety risks. In this study, we propose an innovative air compression refrigeration method based on the Brayton cycle. This method converts high-pressure air into low-temperature air at atmospheric pressure for wind tunnel refrigeration. The new cooling system has reduced energy usage by 3.72 MW, leading to a 13.15% improvement. The return cooler of the system is modeled using the effective number of heat transfer units and the mean temperature difference design method. Additionally, the turbine within the system is analyzed using one-dimensional flow characteristic analysis and the principle of similarity. This method has been validated by comparing it to other published papers. Subsequently, we perform a thorough sensitivity analysis on the key design parameters of the system. We observe that with a sufficient heat transfer area of the recooler, the cooling efficiency of the system exhibits a gradual decline from 64% to 60% as the mass flow rate of the system rises. For a fixed turbine, the cooling efficiency of the system rises from 20% to 62% and subsequently declines to 37%, with an increase in the mass flow rate. As a result, we conclude that the design parameters of the turbine have a more significant influence on the cooling efficiency of the system than the recooler. Our study will establish a foundation for selecting parameters to optimize the refrigeration system in the future.

Researches have focused on the thermal protection system (TPS) of hypersonic vehicles under severe aerodynamic heat. According to the second law of thermodynamics, heat transfer needs to consume a heat sink, but the cold energy provided by the airborne heat sink is limited. Therefore, it is necessary to explore new available heat sinks during hypersonic cruise. This paper numerically calculated the wall temperature distribution of hypersonic vehicle X-51A with different Mach numbers, altitudes and angles of attack using ANSYS Fluent 19.0. A dimensionless parameter, relative temperature coefficient r_t, was proposed to characterize the relative value of local wall temperature in the whole wall temperature range. The distribution regularity and influencing factors of wall temperature were summarized. The low-temperature region which is less affected by flight conditions was divided as the heat sink and its heat dissipation capacity (Q) and characteristics were studied. The angle of attack has great influence on the temperature distribution. In XY view the rear side of leeward surface is least affected by flight conditions and its r_t is less than 0.2, which can be used as a low-temperature region. Taking this region as a heat sink to dissipate heat, it is found that the Q of the new heat sink is 30 kW/m² at Ma 3, 90 kW/m² at Ma 4 and 200 kW/m² at Ma 5. The Q increases greatly with the increase in Mach number, and the convective heat transfer coefficient (h) also increases. At the same Q, the h decreases with the increase in Mach number. The exploration of low-temperature region as heat sink has an important reference for reducing the dependence on consumable heat sink and alleviating the energy shortage of the hypersonic vehicle.