Abstract: Effects of transverse injection from bottom wall of a series of shallow cavity with the same length and different depth on flow field and combustion characteristics of supersonic combustor in altitude 30km, flying Mach number 6, under the condition with or without chemical reactions were simulated. Comparison with traditional wall transverse injection was carried out. It is found that introducing shallow cavity can effectively weaken shock system of flow field and obviously reduce total pressure loss of combustion flow field. Subsonic recirculation zone formed between the front wall of cavity and fuel jet column can strengthen mixing and achieve flame-holding and this effect will be more obvious with increasing depth of cavity.Employing shallow cavity can reduce transverse jet penetration depth, resulting in the decrease of combustion coefficient to a certain extent.

Key words: Cavity; Transverse injection; Recirculation zone; Total pressure loss; Thrust gain; Penetration depth

摘要： 模拟了高度30km,飞行马赫数6的超声速燃烧室流场和燃烧特性。通过对固定长度、不同深度的一组浅凹槽底壁燃料横向喷射的燃烧室的冷态与燃烧工况进行数值计算,并将其和传统壁面横向喷射方式进行比较,发现引入浅凹槽底壁喷射结构能有效减弱流场的激波系强度,明显降低燃烧流场的总压损失；凹槽前壁面和喷流柱之间形成稳定的亚声速回流区,能够稳定火焰,这在较大深度凹槽会更明显。引入浅凹槽一定程度降低了横向射流穿透深度,这也导致燃烧效率相比传统壁面横向喷射结构有一定下降。

关键词: 凹槽;横向喷射;回流区;总压损失;推力增益;穿透深度