The latest characterization of turbulent spot growth accounting for wall temperature and compressibility effects along with a formulation of the transitional Reynolds stresses are assessed in the present γ-α transition model. Additionally, previous model developments to reproduce the normal intermittency profile are now rederived based upon the exact Klebanoff profile. Furthermore, destruction and production terms of the intermittency equation are further refined assuring a more generic transition model. The construction of this transition model is primarily set up to be independent from the used turbulence model, allowing the most suitable to be selected by the user for the envisaged application. The newly developed γ-α model is validated with the classical SST turbulence model for subsonic (with and without pressure gradient) and supersonic (with and without shock wave boundary layer interaction) test cases and has shown to outperform the classical γ-Reθt model in onset location prediction, skin friction slope and overshoot.

As the development of hypersonic flight vehicles is rapidly accelerating, engineers are facing different challenges through-out the design cycle. One of these challenges is the avoidance of hot spots and induced boundary layer transition. Different correlations exist that could be used by designers, but they require specific parameters such as boundary layer thickness and edge properties, streamline length, and stagnation lines, which are not directly provided by a CFD solution. To extract these specific parameters, the Boundary Layer Identification and Transition Zone Detection (BLITZ) code was further extended to extract in a fully automatic way the necessary data all around a flight vehicle and at different instances along the flight trajectory. This paper presents a comparison between purely laminar, purely turbulent and transitional flow along the flight trajectory of a slender high-speed experimental flight test vehicle with respect to boundary layer state, integrated heat load and fluxes, aerodynamic forces, critical roughness and critical steps and gaps. Based on the results obtained from this analysis, an impact assessment of the vehicle weight is performed, showing significant gross take-off weight and fuel consumption reductions when transitional parameters are used rather than a more conservative approach purely based on fully turbulent flow.

As the instruments of satellites are getting more sophisticated, they are also getting more sensitive to external influences. Deposition of gases due to the low-temperature environment can be one of these external influences. For the European-Chinese SMILE mission, a Soft X-Ray Imager (SXI) is present which has CCD’s exposed to very low temperature. As water vapour is present in the air, as well as water absorbed into the paint of the enclosure, ice deposition could happen under specific conditions. To assess the outgassing of PU1 paint, an experimental outgassing study has been performed at room temperature. Additionally, numerical simulations are performed both in the continuum and the molecular regime to assess the proper functioning of the venting paths. A semi-theoretical correlation is provided which allows calculating the mass flow leaving the instrument based on the internal pressure and temperature. For the specific application of the SXI instrument, the venting rate of the enclosed air is assessed. Also, the outgassing of water vapour should not result in pressure conditions which could trigger ice deposition on the cold CCD surface. Based on the performed simulations and experiments, no issues are expected for the SXI.

Different correlations for boundary layer transition exist within literature. However, they often require boundary layer parameters (like displacement and momentum thickness) which are not automatically provided by standard CFD tools. To tackle this problem, the boundary layer identification and transition zone detection (BLITZ) code was developed. This work presents different improvements to the tool to further extend its capabilities and improve its accuracy. To ensure faster analysis of flight vehicles, parallelization is implemented combined with a new streamline tracing and interpolation methodology. Additionally, a blending methodology between the laminar and turbulent field solution is developed based on an intermittency factor to provide a better representation of the real flow acting on the vehicle. Together with the implementation of a standalone monitoring property calculation algorithm, it is now possible to make an improved estimate of the total heat load into the vehicle as well as the aerodynamic forces exerted on it considering the evolution of the transitional flow regime. Further, improvements on the boundary layer edge detection, streamline tracing and pressure gradient calculation are implemented. The results are validated on flat plate cases with and without pressure gradient. Finally, a first validation based on launch data found in the literature is provided. Consequently, an analysis of different possible fairing geometries is given to show the impact of continuity in radius of curvature on the boundary layer transition.

A detailed computational analysis of the oxidizer preflow during startup of a storable propellant upper-stage rocket engine is presented. This low-pressure flow regime is controlled by various two-phase phenomena, particularly flash atomization and flash evaporation of superheated liquid oxidizer, leading to a significant temperature dropinthe combustion chamber.Toaccount for these phenomena, physical models are extendedtolow- pressure conditions and implemented into the framework of an iterative Euler-Lagrange method. A new semi- implicit concept of the spray source term linearization is employed, improving the robustness of the numerical solution procedure by accounting for intense coupling of vapor flow and spray. Following the analysis of the three- dimensional two-phase flowfield and spray deposit distribution in the combustion chamber and nozzle extension, the influence of liquid injection temperature and initial droplet size distribution is investigated by detailed parametric studies. The results indicate a particular importance of the spray-wall interaction and secondary-droplet breakup for the coarser sprays and a bimodal deposit distribution on the chamber walls. Computational results agree well with experimental data and are used to derive transfer functions describing global preflow dynamics on a system level.

At the end of its mission to the International Space Station, during its reentry into Earth atmosphere, the automated transfer vehicle is subject to high heat fluxes leading to structural heating and fragmentation of the vehicle. It has been concluded that, depending on the mode of release, onboard residual hypergolic propellants may ignite and explode upon exposure to the hot and reactive flow environment. Because an earlier explosion of the vehicle would change drastically the impact footprint of its fragments onto the Earth surface, this study proposes a reassessment of the explosion potential. From the trajectory analysis, several points of the reentry path have been computed using a Navier-Stokes solver accounting for nonequilibrium effects. Numerical simulations have been performed with and without perforation of the structure. In parallel, a comprehensive literature survey on ignition of monomethyl hydrazine and dimethyl hydrazine vapors with pure air or air mixed with nitrogen tetroxide has been performed to assess the autoignition potential of the mixture. Finally, the results of the computational fluid dynamics computations have been used to estimate the explosion risk in the presence of a propellant leakage. Analysis confirms the risk of a destruction of the automated transfer vehicle at higher altitude, which could induce a different footprint of the fragments on the ground.

A computational analysis of the oxidiser preflow during start-up of a storable propellant upper-stage rocket engine is presented. To account for flash-evaporation of the superheated liquid oxidiser new physical models and numerical solution techniques are developed and implemented into the framework of an iterative Euler-Lagrange method. The computed three-dimensional two-phase flow field and spray deposit distribution in combustion chamber and nozzle extension is analysed. The influence of liquid injection temperature, initial droplet sizes and wall temperature on the preflow is assessed by detailed parametric studies. Validated by experimental data, the results are used to derive transfer functions describing the preflow dynamics on a system level.

The application of computational fluid dynamics (CFD) methods which offer feasible solution for accelerating and refining the space-transportation design process, is discussed. CFD methods and simplfying analytical modeling are used to determine critical wind velocities for the launcher structure on the launch pad. They predict the effect of the complex flow in the ducts that divert the exhaust plumes on the launch pad. The three-dimensional computational results such as wall stream lines and velocity vector plots near the nozzle exit, as well as experimental data obtained both on ground and in flight, are essential for an in-depth understanding of the flow physics and ultimately for shape improvements that reduce the loads to an acceptable level.

The low-pressure oxidizer preflow during start-up of an upper-stage rocket engine is analyzed computationally. An iterative Euler-Lagrange approach is used to describe the three-dimensional two-phase flow of oxidizer vapour and spray which is characterized by intense phase transfer. Multipoint injection and flash-atomization of liquid oxidizer is modelled by stochastic droplet injection according to injector locations and spray characteristics. Based on the framework of classical D²-theory, a flash-evaporation model is developed to describe heat and mass transfer between superheated liquid droplets and vapour flow. The computed transsonic flow field agrees well with experimental characteristics such as pressure level and temperature decrease and further provides detailed local information on vapour flow homogeneity and liquid wall deposit.