CAJ | 학술논문

针对现有弹用固体火箭冲压发动机普遍采用的固定几何不可调节喷管,基于流量平衡的基本原理,建立了其理论设计及性能评估的数学模型.结合当前中远程空空导弹提出的Ma=2～3.5宽速度范围设计需求,运用所建立设计模型对实例设计方案开展了计算分析.结果表明,现有固定几何喷管本质上是为满足低速正常接力而折中设计出的,在高速巡航时,因扩张比偏小,不仅喷管出口气流速度和冲量小,而且导致燃烧室压强降低,还额外造成进气道结尾正激波总压损失加大,不能将进气道保有的捕获高速来流动能充分发挥出来.原设计方案在Ma=3.5高速巡航时,进气道实际总压恢复性能对比方案中的最大总压恢复性能水平,相对损失幅度高达42.67%,而且冲压发动机推力与其可能达到的最大值对比,相对损失幅度也高达31.8%.因此建议采用喷管调节技术来解决此类问题.
침대현유탄용고체화전충압발동궤보편채용적고정궤하불가조절분관,기우류량평형적기본원리,건립료기이론설계급성능평고적수학모형.결합당전중원정공공도탄제출적Ma=2～3.5관속도범위설계수구,운용소건립설계모형대실례설계방안개전료계산분석.결과표명,현유고정궤하분관본질상시위만족저속정상접력이절중설계출적,재고속순항시,인확장비편소,불부분관출구기류속도화충량소,이차도치연소실압강강저,환액외조성진기도결미정격파총압손실가대,불능장진기도보유적포획고속래류동능충분발휘출래.원설계방안재Ma=3.5고속순항시,진기도실제총압회복성능대비방안중적최대총압회복성능수평,상대손실폭도고체42.67%,이차충압발동궤추력여기가능체도적최대치대비,상대손실폭도야고체31.8%.인차건의채용분관조절기술래해결차류문제.
Aiming at fixod-geometry nozzle popularly used for existing ducted rockets, a mathematical model for analytical de-sign and performance evaluation of nozzle is established based on the principle of mass flow balance. To meet the design require- merits of Mach number 2 to 3.5 range for modem medium-to-long range air-to-air missile applications, nozzle design and analysis u- sing present model is carried out. The results show that existing fixed-geemetry design is essentially a compromise to satisfy low-speed takeover requirements. During high speed cruise process, lower nozzle expansion ratio causes not only lower exit flow velocity and impulse, but also lower pressure in combustion chamber. In addition, lower nozzle expansion ratio also results in higher total pressure loss of exit terminal normal shock wave. Therefore, free stream kinetic energy captured by the inlet can not be fully uti-lized. During Ma = 3.5 cruise process, actual total pressure recovery performance in original design has a relative loss of 42.67% compared to maximum potential value, and relative thrust loss of 31.8% compared to maximum achievable value of ducted rockets. Therefore, it is advised that nozzle regulation technology should be used to solve this problem.