Tokitada Hashimoto

Experimental investigation of shock stand-off distance on spheres in hypersonic nozzle flows

In order to understand thermochemical non-equilibrium phenomena which occur during re-entry, shock stand-off distances on a simple hemi-spherical model were measured in a high enthalpy shock tunnel. Moreover, from the measurements, the unresolved question of how much influence un-recombinated oxygen atoms and vibrationally excited molecules related to the frozen phenomena of nozzle flows have on the flow was investigated. It was also discussed whether a binary scaling parameter which indicates a non-equilibrium process can accommodate the nozzle flows of a shock tunnel.

Observation of the Processes of Ignition and Combustion Flowfield Formation in a Supersonic Combustor with Presence of Stream wise Vortices

Quite different processes of ignition and combustion flowfield formation in a Mach 3 supersonic combustor depending on fuel injection and mixing schemes are observed by high speed schlieren video and wall pressure measurement. Experiments are conducted at a Mach 8 simulated flight condition using the High Enthalpy Shock Tunnel of JAXA, Japan. Parallel (12°) injection scheme of gaseous hydrogen fuel with generation of streamwise vortices by Hypermixer (HM) injector is examined and compared with the two typical normal (perpendicular) injection schemes, which are with and without backward facing step. Equivalence ratio is varied from Φ = 0.3 to 1.5. In the case of HM injector, the induction time increases as increasing Φ, to gradually push the ignition point far downstream of the injector. In larger Φ of 1.0 and 1.5, transient processes are observed, where a strong pressure wave generated by the explosive combustion of the well premixed gas at the combustor exit propagates upstream and decays around the injector to form a new quasi‐steady flowfield of a supersonic combustion mode downstream of the injector. In this new flowfield, combustion starts near downstream of the injector showing rather the features of diffusion flame combustion. The flow is choked upstream near the combustor exit and is accelerated to supersonic again in the nozzle. The properties of the observed pressure wave agreed well with those of Chapman‐Jouguet detonation wave on the basis of 1D‐flow analysis indicating that this pressure wave is a kind of detonation wave. For normal injectors, similar upstream propagation of the pressure wave is observed, but it propagates far upstream of the injector to form subsonic combustion mode. Comparing the characteristics of the pressure wave propagation for the injectors, some considerations are made on the mechanism of the transient processes, focusing on the roles of streamwise vortices to effectively operate the supersonic combustion.Quite different processes of ignition and combustion flowfield formation in a Mach 3 supersonic combustor depending on fuel injection and mixing schemes are observed by high speed schlieren video and wall pressure measurement. Experiments are conducted at a Mach 8 simulated flight condition using the High Enthalpy Shock Tunnel of JAXA, Japan. Parallel (12°) injection scheme of gaseous hydrogen fuel with generation of streamwise vortices by Hypermixer (HM) injector is examined and compared with the two typical normal (perpendicular) injection schemes, which are with and without backward facing step. Equivalence ratio is varied from Φ = 0.3 to 1.5. In the case of HM injector, the induction time increases as increasing Φ, to gradually push the ignition point far downstream of the injector. In larger Φ of 1.0 and 1.5, transient processes are observed, where a strong pressure wave generated by the explosive combustion of the well premixed gas at the combustor exit propagates upstream and decays around the inje...

Experimental Investigations of the Flow Field around a Generic Re-entry Capsule in HIEST

Viscous hypersonic flows with shock waves are of great importance in aerospace engineering applications. Acquisition of detailed experimental data like drag and heat transfer over hypersonic bodies are indispensable for the development of reliable space vehicles and thermal protection systems (TPS), and also for the validation and progress of numerical estimation. In the present study, a blunted double cone model (heat model) was designed and manufactured as the aerodynamic configuration design of re-entry capsules. The aim of this experiment is to obtain detailed data of heat transfer and pressure distribution as well as the flow structures over the heat model. Coaxial thermocouples and pressure sensors are mounted at the frontal and base region, and surface temperature and wall pressure are measured.