Abstract:
The paper presents results of a scramjet model test in a hypersonic pulsed adiabatic-compression wind tunnel AT-303 recently put into operation at ITAM. The model was tested at Mach number M ∞≈8,duration of runs was τ=50~60 ms, and a wide Reynolds number range of Re1 ∞=2.7×106~4.0×107 with boundary layer on the model surfaces developing naturally. Due to the model with fuel supply, the gaseous hydrogen was injected into the combustion chamber at air-to-fuel factors exceeding the stoichiometric ratio. The flow conditions sufficient for self-ignition of the hydrogen were provided. Lengthwise pressure and heat flux distribution along the inlet wedge ramp and along the whole engine duct were measured. The obtained results were compared to data of testing the same model in a hot-shot wind-tunnel IT-302M at M ∞≈6 and 8, τ=100~120ms, Re1∞=(1.3~1.8)×106, with boundary layer tripping on the surfaces of the inlet ramp and side compression wedges. It was demonstrated that the flow patterns of the same type developed in the model engine during testing in both the wind-tunnels. Immediately on starting of a wind-tunnel, a supersonic flow pattern formed in the inlet and downstream in the engine duct. After hydrogen injecting, firstly, a combustion flow pattern developed in the combustion chamber with the flow velocity being supersonic on average. After this flow pattern transformed into a flow pattern with thermal chocking at the combustion chamber exit and with a pseudo-shock wave developing in the inlet diffuser. The satisfactory agreement of flow characteristics of the inlet and the engine duct measured in both the wind-tunnels was obtained.
