【摘要】：系統(tǒng)梳理了影響固體火箭沖壓發(fā)動機二次燃燒性能的關鍵參數,通過數值仿真研究了各因素對二元雙下側進氣道布局固沖發(fā)動機二次燃燒性能的耦合影響規(guī)律,在此基礎上,完成了高性能固體火箭沖壓發(fā)動機地面試驗樣機設計和驗證。研究結果表明,采用軸向一次進氣方案時,進氣角度在45°~60°、頭部距離在0.5D~0.75D時,二次燃燒效率較高,設計點余氣系數2.0時,在該范圍內應盡可能選用大的進氣角度和頭部距離;設計點余氣系數2.0時,應優(yōu)先選用軸向兩次進氣的結構形式,二次進氣的間距在0.5D附近、二次進氣角度在45°~60°時,可達到較高的燃燒效率;經優(yōu)化設計,空燃比14~20范圍內變化時,發(fā)動機的燃燒效率達到0.924以上。
[Abstract]:The key parameters affecting the secondary combustion performance of solid rocket ramjet are systematically combed, and the coupling law of each factor on the secondary combustion performance of binary double lower inlet layout solid impact engine is studied by numerical simulation. The ground test prototype of high performance solid rocket ramjet is designed and verified. The results show that when the axial primary intake scheme is used, the intake angle is 45 擄~ 60 擄and the head distance is 0.5D~0.75D, the secondary combustion efficiency is higher and the residual gas coefficient of the design point is 2.0. In this range, a large intake angle and head distance should be selected as much as possible. When the residual gas coefficient of the design point is 2.0, the structure of the axial and secondary intake should be preferred. When the distance between the secondary intake and the angle of the secondary intake is about 0.5D and the angle of the secondary intake is 45 擄~ 60 擄, the combustion efficiency can be achieved. After optimization design, the combustion efficiency of the engine is more than 0.924 when the air-fuel ratio varies within the range of 14 鈮，

本文編號：2483627