Jinping Li

The ablation product carbon dioxide (CO2) around the hypersonic vehicle may become an important source of radiation. Accurate simulation of the chemical reaction process between CO2 and air is of great significance for the design of thermal protection systems. A combination of ground experiments and numerical simulations is used in this study to investigate the reaction process of CO2–air mixtures with different proportions behind shock waves with velocities ranging from 3.6 to 4.1 km/s. Flow parameters are simulated based on the two-temperature (2-T) model and the Lee and Park chemical reaction mechanism, and the laser absorption spectroscopy measurements are carried out on the Φ800 mm shock tube at the Chinese Academy of Sciences, Institute of Mechanics, to obtain time-resolved post-shock gas temperatures and partial pressures of CO2 and carbon monoxide (CO). Measurement results indicate that the post-shock temperature is much higher than the equilibrium value when it is close to the shock front and then gradually develops towards equilibrium. The partial pressure of CO2 shows the same trend as temperature. Comparisons between experimental results and calculated results show that the used model can accurately predict the post-shock temperature. When the initial CO2 volume fraction is below 50%, the model underestimates the CO2 dissociation rate, meaning that the actual CO2 content is lower than the calculated value and the CO content is higher. Conversely, when the CO2 volume fraction is greater than or equal to 50%, the prediction of the CO2 dissociation rate tends to be overestimated.

This study investigates the radiative properties of hypersonic carbon dioxide (CO2) expanding flows. Spatially resolved spectral measurements for the expansion region of a two-dimensional model were conducted in the JF14 shock tunnel. Experiments were performed in a shock tube mode, where the shock wave with velocities ranging from 3.37 to 3.88 km/s directly propagates to the test section, forming a stable flow field around the model. Results showed that the radiation emission within the 250–710 nm range is primarily dominated by CO2 chemiluminescence and Schumann–Runge bands of molecular oxygen (O2). The intensity of the Schumann–Runge spectrum was found to be significantly higher than the O2 thermal radiation calculated under the two-temperature (2-T) model, indicating that the O2 excited state (O2(3B)) was overpopulated relative to the Boltzmann distribution in the expansion region. In regions with lower degrees of expansion, the CO2 chemiluminescence intensity was overestimated by about 3 to 6 times, which may be attributed to an overprediction of the atomic oxygen (O) and carbon monoxide (CO) number densities in the current thermochemical model. As the degree of expansion intensified, the CO2 chemiluminescence intensity gradually exceeded the predicted values, suggesting that the expansion process can accelerate the recombination of O and CO, leading to stronger CO2 chemiluminescence. This work enhances the understanding of radiative processes in hypersonic CO2 expanding flow and provides benchmarks for the development of thermochemical and radiative models.

This study investigates nitrogen molecular radiation in hypersonic expanding flow around a two-dimensional model. Experiments were conducted using the JF-14 shock tunnel in shock tube mode, generating a 3.25 km/s shock wave to create the required flow. A blunt model with a 25-mm nose radius was used to generate the expanding flow. Spatially resolved visible spectra were measured around the model's shoulder. Numerical simulations were then conducted to analyze flow properties and spectral distributions. The results show that the N2 first positive system dominates emissions in the 500–750 nm range. The spectral profiles matched well between experiment and simulation, but predicted intensities were 3–4 times higher, likely due to an overestimation of the excited state density in the model. A detected spectral line near 520 nm, close to the wall, suggests that the predissociation may have been underestimated. This work extends experimental data on hypersonic expanding flow, contributing to an improved understanding of radiation in such flows.

Solving the three-dimensional boundary layer equations carries theoretical significance and practical applications, which also poses substantial challenges due to its inherent complexity. In this paper, the laminar boundary layer equations for the symmetry plane of three-dimensional bodies are derived in an orthogonal curvilinear coordinate system associated with the principal curvatures. The derivation of the boundary layer equations is based not only on the common symmetric properties of the flow, as given by Hirschel et al. (Three-Dimensional Attached Viscous Flow, 2014, Academic Press, pp. 183–187), but also incorporates the geometric symmetry properties of the body. The derived equations are more representative and simplified. Notably, these equations can degenerate to a form consistent with or equivalent to the commonly used boundary layer equations for special bodies such as flat plates, cones and spheres. Furthermore, for hypersonic flows, the crossflow velocity gradient at the boundary layer edge on the symmetry plane is derived based on Newtonian theory. Subsequently, this parameter can provide the necessary boundary condition needed for solving the boundary layer equations using existing methods. Finally, as examples, the equations developed in this paper are solved using the difference-differential method for several typical three-dimensional blunt shapes that appeared on hypersonic vehicles. They prove to be useful in the analysis and interpretation of boundary layer flow characteristics in the symmetry plane of blunt bodies.

High enthalpy shock tunnel is a critical ground testing facility for evaluating the aerodynamic performance of hypersonic vehicles by simulating high-enthalpy, hypersonic flows. However, as the total temperature or pressure of the stagnant gas rises, the throat insert may be subjected to melting or oxidation, leading to a degradation in flow quality or a reduction in effective test time. This study investigates the unsteady heat transfer between the gas flow and the throat insert using a conjugate heat transfer model. The effects of test time, total temperature, and total pressure on the throat temperature are presented. An increase in any of these factors will raise the throat temperature. Moreover, oxidation will cause the copper throat insert to melt more quickly under the same conditions, which should be considered when addressing throat melting issues.

The flow near the stagnation streamline of a blunt body is often attracted and analyzed by using the approximation of local similarity, which reduces the equations of motion to a system of ordinary differential equations. To efficiently calculate the stagnation-streamline parameters in hypersonic magnetohydrodynamic (MHD) flows, an improved quasi-one-dimensional model for MHD flows is developed in the present paper. The Lorentz force is first incorporated into the original dimensionally reduced Navier–Stokes equations to compensate for its effect. Detailed comparisons about the shock standoff distance and the stagnation point heat flux are conducted with the two-dimensional Navier–Stokes calculations for flows around the orbital reentry experiment model, including gas flows in thermochemical nonequilibrium under different magnetic field strengths. Results show that the shock curvature should be considered in the quasi-one-dimensional model to prevent accuracy reduction due to the deviation from the local similarity assumption, particularly for hypersonic MHD flows, where the shock standoff distance will increase with larger magnetic strength. Then, the shock curvature parameter is introduced to compensate for the shock curvature effect. A good agreement between the improved quasi-one-dimensional and the two-dimensional full-field simulations is achieved, indicating that the proposed model enables an efficient and reliable evaluation of stagnation-streamline quantities under hypersonic MHD flows.

The starting process of the flow in the nozzle of the JF-14 shock tunnel (1.6 m in length, 500 mm in outlet diameter) in the State Key Laboratory of High Temperature Gas Dynamics is analyzed by calculation and experiment. Two key factors which directly affect the duration of the nozzle starting are the velocity of the expansion wave and the low-velocity zone generated by the interaction between the secondary shock wave and boundary layer on the wall surface. In the process of the nozzle starting, the flow field stabilizes at the center of the nozzle outlet first, and then gradually stabilizes along the radius direction, thus defining the central startup and complete startup of the nozzle. It is found that there is a critical initial pressure. When the initial pressure is lower than the critical pressure, the airflow can reach stability in the nozzle outlet center with the shortest time, otherwise, the time required is much longer. The time required for the airflow to stabilize in the whole outlet section is mainly affected by the size of the low-velocity zone. It is also found that only at a very low initial pressure can the airflow simultaneously reach stability at the entire outlet of the nozzle.

The need to increase the payload capacity of the rockets motivates the development of high-power rocket engines. For a chemical propulsion system, this results in an increasing thermal load on the structure, especially the combustion chamber and nozzle must be able to withstand the extreme thermal load caused by high-temperature and high-pressure combustion gas. In order to protect the structure from the effect of increasing heat flux, it is necessary to counteract such effect with more advanced thermal management technology. This requires us to accurately predict the aerodynamic heating of the structure by high-temperature and high-speed combustion gas. In this study, a high-temperature combustion gas tunnel developed in the laboratory is used to produce high-speed combustion gas. Combined with the results of numerical calculation, the flow and aerodynamic heating characteristics of air and hydrogen–oxygen combustion gas under the same total temperature and pressure are analyzed and compared. The comparison revealed that the combustion gas flow in the nozzle has higher static temperature, velocity, and smaller Mach number. When the combustion gas flows around the sphere, the shock standoff distance and stagnation pressure are smaller than those of air, and the wall heat flux is much larger than that of air. The active chemical reaction in the combustion gas makes the aerodynamic heating of the structure more severe. Finally, through the analysis of a large amount of data, a semi-empirical formula for the heat flux of the stagnation point heated by a high-speed hydrogen and oxygen equivalent ratio combustion gas is obtained.

Coaxial thermocouples have the advantages of fast response and good durability. They are widely used for heat transfer measurements in transient facilities, and researchers have also considered their use for long-duration heat transfer measurements. However, the model thickness, transverse heat transfer, and changes in the physical parameters of the materials with increasing temperature influence the accuracy of heat transfer measurements. A numerical analysis of coaxial thermocouples is conducted to determine the above influences on the measurement deviation. The minimum deviation is obtained if the thermal effusivity of chromel that changes with the surface temperature is used to derive the heat flux from the surface temperature. The deviation of the heat flux is less than 5.5% when the Fourier number is smaller than 0.255 and 10% when the Fourier number is smaller than 0.520. The results provide guidance for the design of test models and coaxial thermocouples in long-duration heat transfer measurements. The numerical calculation results are verified by a laser radiation heating experiment, and heat transfer measurements using coaxial thermocouples in an arc tunnel with a test time of several seconds are performed.

Robust fast-response transient calorimeters with novel calorimeter elements have attracted the attention of researchers as new synthetic materials have been developed. This sensor uses diamonds as the calorimeter element, and a platinum film resistance is sputtered on the back to measure the temperature. The surface heat flux is obtained based on the calorimetric principle. The sensor has the advantages of high sensitivity and not being prone to erosion. However, non-ideal conditions, such as heat dissipation from the calorimeter element to the surroundings, can lead to measurement deviation and result in challenges for sensor miniaturization. In this study, a novel transient calorimeter (NTC) with two different sizes was developed using air or epoxy as the back-filling material. Numerical simulations were conducted to explain the complex heat exchange between the calorimeter element and its surroundings, which showed that it deviated from the assumption of an ideal calorimeter sensor. Accordingly, a dynamic correction method was proposed to compensate for the energy loss from the backside of the calorimeter element. The numerical results showed that the dynamic correction method significantly improved the measurement deviation, and the relative error was within 2.3% if the test time was smaller than 12 ms in the simulated cases. Detonation shock tunnel experiments confirmed the results of the dynamic correction method and demonstrated a practical method to obtain the dynamic correction coefficient. The accuracy and feasibility of the dynamic correction method were verified in a single detonation shock tunnel and under shock tube conditions. The NTC calorimeter exhibited good repeatability in all experiments.