高超声速脉冲风洞中带燃料的超燃冲压发动机模型试验(英文)

Testing a model scramjet with fuel supplying in hypersonic pulsed wind tunnel

给出了在ITAM最近投入使用的高超声速脉冲绝热压缩风洞AT-303中进行超燃冲压发动机模型实验的结果。实验马赫数M∞≈8，运行时间r=50—60ms，雷诺数范围Re1∞=2．7×10^6—4．0×10^7，模型表面的边界层自然转捩。在实验中，模型中有燃料供给：把气态氢以超过化学量比率的空气燃料因子注入到燃烧室。提供了足以发生氢燃料自点燃的流动条件。测量了沿进气道楔型压缩面和整个发动机通道上的纵向压力和热流分布。所获数据与同一模型在热射流风洞IT-302M（实验马赫数M∞≈6，8，运行时间τ=100—120ms，雷诺数范围Re1∞=（1．3-1．8）×10^6，进气道压缩面和侧压缩面进行了边界层转捩）。结果表明：实验模型发动机在两座风洞中进行实验所获得的流态类型相同。发动机刚刚启动时，在进气道入口及其下游的发动机通道内形成超声速流。注入氢后，首先在燃烧室内形成平均流速是超声速的燃烧流动。之后，在燃烧室出口出现热拥塞现象、在进气道扩压段产生伪激渡流态。在两座风洞中进行了进气道和发动机通道的流动特征试验，获得了令人满意的结果。

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 - 60ms, and a wide Reynolds number range of Re1∞ = 2.7 × 10^6 - 4.0 × 10^7 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） × 10^6, 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.