RBCC亚燃模态热力调节方式研究  

Research on thermal adjustment method of RBCC in ramjet mode

作　　者：王亚军[1] 李江[1] 秦飞[1] 何国强[1] 魏祥庚[1] 

机构地区：[1]西北工业大学航天学院,西安710072

出　　处：《固体火箭技术》2016年第2期151-159,共9页Journal of Solid Rocket Technology

摘　　要：为了进一步提升RBCC亚燃模态性能,利用一维分析模型,并基于燃烧室隔离段匹配工作,开展了热力调节方式研究。结果表明,影响RBCC扩张流道内热力喉道生成位置的主要因素为燃烧室扩张比、放热量、放热速率及放热位置,扩张比越大、放热量越少、放热越快和放热越靠前,热力喉道越早生成,通过一定范围参数的改变,热力喉道生成位置变化了9%~22%,当燃烧室扩张比和燃料总放热量相同时,热力喉道的生成位置越靠后燃烧室压力积分推力越大。相比扩张比和放热速率,改变燃料喷注位置是一种有效且易实现的热力调节方法。对于较为靠后的燃烧组织,应匹配较大扩张角的流道,能有效提高热力调节的可调范围,拓展发动机工作裕度。开展了亚燃模态来流Ma=3~5性能优化研究,相比固定喷注位置,采用改变燃料喷注位置的热力调节方式,发动机推力性能能够获得16.5%~24.1%的提升。In order to further enhance the performance of RBCC engine in ramjet mode, one-dimensional analysis model was used to study the methods of thermal adjustment based on the matching operation between eombustor and isolator.The results show that the main influencing factors of thermal throat in the expanding duct of RBCC were expansion ratio, exothermic quantity, exo- thermic rate and injection position.The bigger of expansion ratio, the less of exothermic quantity, the faster of heat release or the more front of injection location is, the earlier thermal throat forms in duct. Through the variation of above parameters in certain range, thermal throat position changes by 9% ~ 22%.The more after of thermal throat position is, the bigger of integral thrust can be obtained when the expansion ratio of comhustor and total exothermic quantity are the same.Compared to the expansion ratio and exo- thermic rate,changing the position of fuel injection is an effective and easy method to implement the thermal adjustment.More after of combustion organization should match the larger expansion angle of duct, which can effectively extend the range of thermal adjust- ment and the margin of engine operation. In the study of performance optimization, compared to the fixed position of injection, the thrust of engine can achieve 16.5%-24.1% improvements at fly condition of Ma= 3-5 by using the thermal adjustment method of changing fuel injection location.

关 键 词：火箭基组合循环 亚燃模态 热力喉道 热力调节 

分 类 号：V438[航空宇航科学与技术—航空宇航推进理论与工程]