It is well-known that the phenylboronic acid derivatives have chemical interactions with sugars. Hence, stable nanoparticles with core-shell structure were formed by the covalent complexation between boronic acid groups of poly(3-acrylamidophenylboronic acid) (pAPBA) and hydroxyl groups of poly (2-lactobionamidoethyl methacrylate) (pLAMA). The image taken by transmission electron microscopy (TEM) shows that the nanoparticles had a size of about 200 nm and were irregular spheres. The methyl thiazolyl tetrazolium (MTT) assay suggests that the nanoparticles were non-cytotoxic on the human colorectal carcinoma (Caco-2) cells. The confocal laser scanning microscope (CLSM) images show that the nanoparticles could internalize into Caco-2 cells. The insulin-loaded nanoparticles by intranasal administration led to a significant decrease in the plasma glucose levels and the histological assessment revealed that the nanoparticles would not develop lesions in the nasal epithelium. The nanoparticles are promising carriers for peptide and protein drugs in nasal delivery.

To better understand the performance of gaseous film cooling near the injector region in the LOX/GH 2 thrust chambers, a methodology is employed to simulate the coupled heat transfer of the film cooling in thrust chambers with regenerative cooling. The conjugated flow and heat transfer behaviors of film coolant, hot gas, cooling channels and regenerative coolant are numerically investigated. A three-dimensional non-adiabatic flamelet model using real gas equation of state is developed to solve the combustion and validated against the experimental data. Film cooling performance is predicted for the conditions with different geometrical parameters and mass flow rates of the film coolant. The result shows that the reverse flow zones are formed and developed in the region near the injectors The occurrence of those zones is responsible for the significant reduction of the hot-gas-side wall temperature near the head plate of the injector. The coolant mass flow rate has a great influence on the film cooling performance due to the variety of coolant momentum in the exit of orifices and vorticity in the near-injector region. An optimum mass flow rate for maximizing the averaged effectiveness exists for a given film orifice geometric configuration. The high averaged effectiveness and the uniform flux distribution of hot-gas-side wall are observed at small orifices spacing. The film cooling effectiveness in the orifice exit region is obviously enhanced when the diameter of the orifice increased in the front part of the combustion chamber. The results would be useful for the analysis and optimization design of the straight cylindrical coolant orifices in the LOX/GH 2 thrust chamber.

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In order to understand heat transfer mechanisms in the combustion chamber with multi-elements injector and provide benchmark data of wall heat fluxes for the CFD code validation, experiments were carried out in a GH2/GO2 heat-sink combustion chamber. To obtain the transient heat flux in the experiments, the LevenbergMarquardt method was modified and applied to the rocket combustion chamber based on the temperature measured by single coaxial thermocouple (Named "single-point method"). The comparison between the singlepoint method and the two-point method with temperatures measured in two points shows that the single-point method could be used as heat flux measurements of the heat-sink combustion chamber in engineering. Heat flux distribution was obtained in experimental conditions of different work time and chamber pressure by the modified method, feasibly and efficiently. For the transient variations of the heat flux, an inverse variation trend of heat flux to time in the cylindrical segment and the nozzle segment was observed. A typical variation of the averaged heat flux was obtained under the conditions with different chamber pressure and work time. The results of wall heat flux scaled with pressure to the power 0.8 can be uniformized for all pressure levels. The heat fluxes obtained in the cylindrical chamber and nozzle section may be applied for the life cycle prediction of rocket engines.