Recent experimental and theoretical work has demonstrated significant potential to tune the properties of silicon and germanium by adjusting the mesostructure, nanostructure, and/or crystalline structure of these group 14 elements. Despite the promise to achieve enhanced functionality with these already technologically important elements, a significant challenge lies in the identification of effective synthetic approaches that can access metastable silicon and germanium-based extended solids with a particular crystal structure or specific nano/meso-structured features. In this context, the class of intermetallic compounds known as Zintl phases has provided a platform for discovery of novel silicon and germanium-based materials. This review highlights some of the ways in which silicon and germanium-based Zintl phases have been utilized as precursors in innovative approaches to synthesize new crystalline modifications, nanoparticles, nanosheets, and mesostructured and nanoporous extended solids with properties that can be very different from the ground states of the elements.

The University of Queensland in 2014 school of mechanical and mining engineering centre for hypersonics abstract Realisation of the dream of airbreathing access-to-space requires the development of a scramjet engine that produces sufficient net thrust to enable acceleration over a wide Mach number range. With engines that are highly integrated with the airframe, the net performance of a scramjet powered vehicle is closely coupled with the vehicle attitude and is difficult to determine only from component level studies. This work investigates the influence of airframe integration on the performance of an airframe integrated scramjet through the measurement of internal pressure distribution and the direct measurement of the net lift, thrust and pitching moment using a three-component stress wave force balance. The engine chosen as the basis for this study was the Mach 12 rectangularto-elliptical shape-transition (m12rest) scramjet that was developed by Suraweera and Smart (2009) as a research engine for access-to-space applications. The inlet and combustor flowpath were integrated with a slender 6°wedge forebody, streamlined external geometry and three dimensional thrust nozzle. The scale of the engine was chosen so that the entire engine would fit within the core-flow diamond (bi-conic) produced by a Mach 10 facility nozzle. The Mach 10b facility nozzle was chosen because it is the largest nozzle current in use with the t4 Stalker Tube and because the offdesign performance of a scramjet engine is of interest for access-to-space vehicles that must accelerate over a range of Mach numbers.Freejet experiments were conducted within the t4 Stalker Tube. Two trueflight Mach 10 test conditions were used: a high pressure test condition that replicated flight at a dynamic pressure of 48 kPa and a low pressure test condition that replicated flight at a dynamic pressure of 28 kPa. Scaling of the test conditions according to the established binary scaling law was not completed due to facility operational limits.The engine featured two fuel injection stations from which gaseous hydrogen was injected. The first injection station was partway along the length of the inlet while the second injection station was at the start of the combustor behind a rearward facing circumferential step. In addition to investigating inlet-only and step-only injection, a combined scheme where 68 % of the fuel was injected from the step station and 32 % from the inlet station was also investigated.To support the analysis of the experimental results, numerical simulations of the engine with no fuel injection were conducted using the nasa code vulcan. Analysis of the simulations show that the mass capture ratio with respect to the projected inlet area is approximately 60 % at each test condition. The simulations also show that spillage of flow from the slender forebody accounts for just 12 % of the flow through the projected iii inlet area, a small but non-negligible fraction. By integrating the engine surface forces, the drag coefficient with respect ...

A method is proposed for the aerodynamic design of nozzles with uniform outflow for supersonic and hypersonic ground-test facilities. This method involves the coupling of an open-source Reynolds-Averaged Navier-Stokes CFD solver, Eilmer, with the simplex optimisation method of Nelder & Mead for the design of an expanding nozzle contour with the least variations in Mach number and flow angularity. Three nozzles were designed using this method to produce an exit flow of Mach 4, Mach 7 and Mach 10. Numerical simulations of the flow in these optimised nozzle contours showed excellent flow uniformity in the core flow-typical Mach number variations were less than 0.5%, flow angularity variations were less than 0.05 • , static temperature and flow velocity variations were less than 1%, and Pitot and static pressures variations were less than 2%. Experimental surveys of the Pitot pressures at several planes downstream of the exit of the three optimised nozzles showed excellent agreement with numerical simulations. The experimental measurements showed that there were good levels of uniformity in the core flow regions of all three nozzles, thus proving the validity of the proposed design method.

The mixing between the coolant and the boundary-layer gas downstream of an injector—for transpiration/film cooling—has been extensively studied for turbulent flows; however, only a handful of studies concerning laminar mixing exist, particularly in hypersonic flows. In this paper, the concentration of the coolant gas at the wall and the heat flux reduction downstream of a transpiring injector in a hypersonic laminar flow are experimentally measured and examined. Experiments are performed in the Oxford High Density Tunnel at Mach 7. A flat-plate model is coated with pressure-sensitive paint (PSP) to spatially resolve the film and obtain a film effectiveness based on coolant concentration. Thin-film arrays are installed to measure the heat flux reduction. Six different cases are studied featuring nitrogen and helium as the coolant gas, where the blowing ratio is varied from 0.0406% to $$0.295\%$$ 0.295 % . The unit Reynolds number of the flow is $$12.9\times 10^6\;\mathrm {m^{-1}}$$ 12.9 × 10 6 m - 1 . A coolant concentration of up to $$95\%$$ 95 % is achieved immediately (2 mm) downstream of the injector. The film concentration drops in a monotonic fashion farther downstream; however, a constant film coverage of 5–20 mm immediately downstream of the injector is observed in cases with a higher blowing ratio. A film coverage above 15% over three injector lengths is present even for the lowest blowing ratio. Heat flux reduction is achieved in all cases. The concentration effectiveness obtained from PSP is compared with the thermal film effectiveness calculated from the heat flux reduction. The latter is found to be higher than the former for all data points. Finally, a collapse of the thermal effectiveness is achieved and a modified analytical correlation is proposed. Graphical abstract

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