摘要
根据高压压气机盘螺栓孔结构,设计中心孔板材疲劳试样.表征了孔挤压强化后的表面轮廓,分析了在多种交变载荷条件下孔挤压前后试样的疲劳寿命,并进行了断口观察和疲劳过程中孔挤压残余应力的演化分析.结果表明:孔挤压强化减小了孔壁表面粗糙度,并使孔结构在多种高温大应力条件下的高温疲劳性能提高1~3倍,但疲劳数据分散度略有增大.孔挤压残余应力在最大拉应力为663MPa,温度为600℃,应力比为0.1条件下20 000次疲劳试验中松弛到60%.原始试样的多源疲劳断口主要起源于孔边的加工刀痕,而挤压强化试样断口起源于孔挤压在倒角区域流动金属堆积处,为单源疲劳断口.
Fatigue specimen with central hole was designed based on the bolt-hole of a high pressure compressor disk.Of the as-received（AR）specimens and after hole-expansion（HE）,the surface profile and fatigue property in multiple alternating loads were characterized;moreover,failure analysis and the evolution of HE residual stress profile during fatigue cycle were investigated.Results showed that,after HE,the roughness of hole wall decreased,besides,the high temperature fatigue performances improved 1-3times under multiple large load condition with the slight rise of fatigue data dispersion.The HE residual stress released to 60% during 20 000 cycles at the maximum tensile stress of 663 MPa,temperature of 600℃and stress ratio of 0.1.Compared with the untreated specimens nucleating at machining marks,the HE crack initiated at metal accumulation by HE in the chamfer,presenting single source crack initiations.
作者
王欣
胡仁高
胡博
赵振业
罗学昆
古远兴
宋颖刚
WANG Xin HU Ren-gao HU Bo ZHAO Zhen-ye LUO Xue-kun GU Yuan-xing SONG Ying-gang(Surface Engineering Institute, Beijing Institute of Aeronautical Material, Aero Engine （Group） Corporation of China, Beijing 100095, China Strength Technology Laboratory, Sichuan Gas Turbine Establishment, Aero Engine （Group） Corporation of China, Chengdu 610500, China)
出处
《航空动力学报》
EI
CAS
CSCD
北大核心
2017年第1期89-95,共7页
Journal of Aerospace Power
关键词
高温合金
螺栓孔
孔挤压(HE)
应力集中
残余应力
高温疲劳性能
superalloy
bolt-hole
hole-expansion（HE）
stress concentration
residual stress
high-temperature fatigue property
