摘要
掌握来流马赫数对超声通流风扇(STFF)叶型气动性能的影响对于保证STFF在宽速域下高效稳定运行具有重要意义。本文采用数值模拟方法,开展了宽广来流马赫数(Ma=0.10~2.36)对STFF叶型气动性能影响的研究,并着重讨论了大负攻角以及临界马赫数下叶栅流场结构的演变。研究发现:当来流从亚声速转变为跨声速时,叶栅通道内出现槽道激波,激波损失增加;同时槽道激波与相邻叶片吸力面附面层干扰诱发大尺度流动分离,黏性损失增加。Ma=0.6方案下流场结构随负攻角增加(−2º增至−4º)的演变与前述类似,不同点在于此时叶表并未发生大尺度流动分离。当来流从唯一攻角下的跨声速流态转变为0º攻角下的超声速流态,吸力侧前缘激波强度略微增大,同时前缘激波压力侧分支向叶栅通道上游移动,激波角增加,但波前马赫数降低,导致激波强度降低。在前述两激波的共同作用下,激波损失显著减小;前缘激波压力侧分支与相邻叶片吸力面附面层干扰强度减弱,黏性损失降低。当Ma=1.96、2.16且来流从负临界攻角增至负失速攻角,前缘激波吸力侧分支与相邻叶片压力面发生强烈干涉诱发了流动分离,黏性损失增加;分离区与主流构成的“虚拟压力面”使得前缘吸力侧激波由规则反射转变为马赫反射,前缘激波吸力侧分支在靠近压力面部分演变为马赫杆,激波损失增加。
It is of great significance to understand the influence of wide speed inflow on the aerodynamic performance of Supersonic Through-Flow Fan(STFF)cascades to ensure the efficient and stable operation of the STFF.In this pa⁃per,the influence of inlet Mach numbers ranging from 0.10 to 2.36 on the aerodynamic performance of the STFF cas⁃cade was studied by numerical simulation.The flow structure evolutions of the STFF cascade with large negative inci⁃dences and critical Mach numbers were emphatically discussed.The results are as follows.The change from sub⁃sonic to transonic incoming flow caused a passage shock in the flow passage and increased the shock loss.Simulta⁃neously,the interaction between the shock and the boundary layer of the adjacent blade suction surface induced large-scale flow separation and thus increased the viscosity loss.The evolution of the flow field structure with increasing negative incidences from−2°to−4°at a Mach number of 0.6 is similar to the above process.The difference was that no flow separation occurred on the blade surface in the latter case.The change of incoming flow from transonic flow regime under the condition of“unique incidence”to supersonic flow regime under the condition of 0°incidence slightly enhanced the shock intensity of the suction side branch of the leading-edge shock.Meanwhile,the pressure side branch of the leading-edge shock moved upstream resulting in a shock angle increase,while the Mach number ahead of the shock decreased.These two factors led to an intensity reduction of the pressure side branch of the leading-edge shock.Therefore,under the influence on both the suction and the pressure side of the leading-edge shock,the shock loss was eventually reduced.The interference between the pressure side branch of the leading-edge shock and the boundary layer of the adjacent blade suction surface was weakened,reducing the viscosity loss.In the cases of Ma=1.96 and 2.16 with an incidence varying from the negative critical value to the negative stall v
作者
孙士珺
李晓龙
刘艳明
王建华
王松涛
SUN Shijun;LI Xiaolong;LIU Yanming;WANG Jianhua;WANG Songtao(School of Aerospace Engineering,Beijing Institute of Technology,Beijing 100081,China;China State Shipbuilding Corporation Systems Engineering Research Institute,Beijing 100094,China;Scholl of Energy Science and Engineering,Harbin Institute of Technology,Harbin 150001,China)
出处
《航空学报》
EI
CAS
CSCD
北大核心
2023年第21期317-327,共11页
Acta Aeronautica et Astronautica Sinica
基金
国家自然科学基金(52006011)
北京理工大学青年教师学术启动计划(XSQD-202201002)。
关键词
轴向超声通流风扇
宽速域
来流马赫数
损失特性
激波结构
流动分离
axial supersonic through-flow fan
wide speed range
Mach number
loss characteristics
shock structure
flow separation
