航空发动机
航空髮動機
항공발동궤
AERO ENGINE
2011年
5期
29-32,37
,共5页
李瀚%索建秦%梁红侠
李瀚%索建秦%樑紅俠
리한%색건진%량홍협
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
燃燒室%火燄筒%流阻法%多項式擬閤法%流量分配%燃氣總溫%航空髮動機
연소실%화염통%류조법%다항식의합법%류량분배%연기총온%항공발동궤
combustor%liner%flow resistance method%polynomial fitting method%airflow distribution%gas total temperature%aeroengine
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。
以某型航空髮動機燃燒室為物理模型，改進瞭計算火燄筒流量分配的流阻法，併對其進行驗證，結果冷卻空氣量的相對誤差為5．7%；採用多項式擬閤法計算瞭火燄筒燃氣總溫沿軸嚮分佈。得到瞭主燃區總溫和燃燒室齣口總溫，併採用燃燒效率法對其進行瞭驗證，二者的相對誤差分彆為4．4％和1％。結果錶明：在初始設計階段，採用改進的流阻法和多項式擬閤法驗證火燄筒的沿程空氣流量分配和沿程燃氣總溫閤理有效。
이모형항공발동궤연소실위물리모형，개진료계산화염통류량분배적류조법，병대기진행험증，결과냉각공기량적상대오차위5．7%；채용다항식의합법계산료화염통연기총온연축향분포。득도료주연구총온화연소실출구총온，병채용연소효솔법대기진행료험증，이자적상대오차분별위4．4％화1％。결과표명：재초시설계계단，채용개진적류조법화다항식의합법험증화염통적연정공기류량분배화연정연기총온합리유효。
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.
