摘要
以某型航空发动机燃烧室为物理模型,改进了计算火焰筒流量分配的流阻法,并对其进行验证,结果冷却空气量的相对误差为5.7%;采用多项式拟合法计算了火焰筒燃气总温沿轴向分布,得到了主燃区总温和燃烧室出口总温,并采用燃烧效率法对其进行了验证,二者的相对误差分别为4.4%和1%。结果表明:在初始设计阶段,采用改进的流阻法和多项式拟合法验证火焰筒的沿程空气流量分配和沿程燃气总温合理有效。
Taking an aeroengine combustor as physical model, the flow resistance method of liner was improved and validated, and the relative error of cooling airflow was 5.7% between the design value and prediction. The axial distribution of liner gas total temperature was calculated by the polynomial fitting method. The primary zone and the exit gas total temperature were obtained and the relative errors were 4.4% and 1% between the predictions obtained from the polynomial fitting method and combustion efficiency method. The results show that it is reasonable and effective to validate airflow distribution and gas total temperature along the liner by the improved flow resistance method and polynomial fitting method during preliminary design.
出处
《航空发动机》
2011年第5期29-32,37,共5页
Aeroengine
关键词
燃烧室
火焰筒
流阻法
多项式拟合法
流量分配
燃气总温
航空发动机
combustor
liner
flow resistance method
polynomial fitting method
airflow distribution
gas total temperature
aeroengine
