Purpose

This paper aims to study the influences of hydrogen jet pressure on flow features of a strut-based injector in a scramjet combustor under-reacting cases are numerically investigated in this study.

Design/methodology/approach

The numerical analysis is carried out using Reynolds Averaged Navier Stokes (RANS) equations with the Shear Stress Transport k-ω turbulence model in contention to comprehend the flow physics during scramjet combustion. The three major parameters such as the shock wave pattern, wall pressures and static temperature across the combustor are validated with the reported experiments. The results comply with the range, indicating the adopted simulation method can be extended for other investigations as well. The supersonic flow characteristics are determined based on the flow properties, combustion efficiency and total pressure loss.

Findings

The results revealed that the augmentation of hydrogen jet pressure via variation in flame features increases the static pressure in the vicinity of the strut and destabilize the normal shock wave position. Indeed, the pressure of the mainstream flow drives the shock wave toward the upstream direction. The study perceived that once the hydrogen jet pressure is reached 4 bar, the incoming flow attains a subsonic state due to the movement of normal shock wave ahead of the strut. It is noticed that the increase in hydrogen jet pressure in the supersonic flow field improves the jet penetration rate in the lateral direction of the flow and also increases the total pressure loss as compared with the baseline injection pressure condition.

Practical implications

The outcome of this research provides the influence of fuel injection pressure variations in the supersonic combustion phenomenon of hypersonic vehicles.

Originality/value

This paper substantiates the effect of increasing hydrogen jet pressure in the reacting supersonic airstream on the performance of a scramjet combustor.

中文翻译：

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基于支柱的超燃冲压发动机燃烧室氢射流压力变化的数值研究

目的

本文旨在研究氢气射流压力对超燃冲压发动机燃烧室中支柱式喷射器流动特性的影响，并在本研究中进行了数值研究。

设计/方法/方法

数值分析是使用雷诺平均纳维斯托克斯 (RANS) 方程和剪切应力传输 k-ω 湍流模型进行的，以理解超燃冲压发动机燃烧过程中的流动物理。三个主要参数，如冲击波模式、壁压力和燃烧室的静态温度都通过报告的实验进行了验证。结果符合范围，表明所采用的模拟方法也可以扩展到其他调查。基于流动特性、燃烧效率和总压力损失确定超音速流动特性。

发现

结果表明，通过火焰特征的变化来增加氢气射流压力会增加支柱附近的静压，并破坏正常冲击波的位置。实际上，主流流的压力将激波驱动向上游方向。该研究认为，一旦氢气射流压力达到 4 巴，由于正常冲击波在支柱前方的运动，进入的流动将达到亚音速状态。值得注意的是，与基线喷射压力条件相比，超声速流场中氢气射流压力的增加提高了流动横向上的射流穿透率，也增加了总压力损失。

实际影响

这项研究的结果提供了燃油喷射压力变化对高超音速飞行器超音速燃烧现象的影响。

原创性/价值

本文证实了在超音速反应气流中增加氢气射流压力对超燃冲压发动机燃烧室性能的影响。