发动机低压涡轮导向叶片裂纹分析

低压涡轮导向叶片是发动机中重要热端部件之一,叶片在高温燃气环境下工作,服役条件十分恶劣。发动机工作结束后,发现低压涡轮导向叶片表面存有裂纹和基体缺失现象。通过外观检查、断口宏微观分析、材质分析、气膜孔检查及热模拟试验等手段,对低压涡轮导向叶片的裂纹性质及萌生原因进行分析研究。结果表明:故障低压涡轮导向叶片的裂纹性质为疲劳裂纹,叶片在工作过程中热障涂层脱落,导致叶片组织超温,使其抗疲劳性能下降并萌生疲劳裂纹。     

硅铝合金柴油机机体紧固面微动疲劳研究

目的针对柴油机机体和主轴承盖紧固面之间发生的微动疲劳失效现象,探讨微动状态下柴油机硅铝合金机体的裂纹萌生特性及寿命评价方法。方法建立机体紧固面组合结构有限元模型,计算机体紧固面上的应力/应变历史数据,在此基础上分析机体紧固面的接触状态,探讨摩擦系数及摩擦功对微动疲劳特性的影响。采用多轴疲劳参数(CCB、F、SSI、Ruiz参数)预测了机体微动裂纹萌生位置,对所用参数进行修正,建立适用于机体紧固面的微动疲劳寿命预测模型。结果预测结果与实验值对比可知,F、SSI参数与实验结果差异较大,CCB和Ruiz参数的寿命预测结果与实验值接近,在2.3倍公差带因子范围内。但由于CCB参数预测的裂纹萌生位置和机体实际断裂位置不符,所以不能用于机体微动疲劳寿命预测。结论在接触状态突变的区域容易萌生微裂纹,适当增大摩擦系数或者降低摩擦功可以抑制机体的微动疲劳损伤。Ruiz参数预测的机体微动疲劳寿命与实验值最为吻合,用Ruiz参数评估柴油机硅铝合金机体的微动疲劳寿命可以将误差控制在2.3倍公差带因子范围内。     

空心风扇叶片榫头裂纹原因分析

空心风扇叶片振动疲劳试验后在榫头表面出现裂纹。通过外观检查、断口宏微观分析、表面检查、成分分析、组织和硬度检测等试验,对裂纹性质和产生原因进行了分析研究。结果表明:叶片榫头表面裂纹为微动疲劳开裂,叶片与夹具间产生的微动磨损是导致该叶片过早萌生疲劳裂纹的主要原因,而产生微动磨损与叶片榫头的几何特征、夹具与其配合状态及榫头部位未采用表面处理措施有关。     

镍基铸造高温合金低周疲劳研究进展

镍基铸造高温合金具有优异的高温性能,广泛应用于航空发动机涡轮叶片等热端部件之中。航空发动机涡轮叶片是发动机中工作环境最为恶劣、结构最为复杂的零件之一,在发动机运行过程中所产生的高温交变应力的作用下,合金承受着严重的应力、应变循环损伤,裂纹往往在合金中的薄弱区域形成并扩展,使合金以低周疲劳的模式失效,严重影响了合金的服役寿命,因此对合金低周疲劳性能的研究尤为重要。本文详细阐述了影响镍基铸造高温合金低周疲劳性能的表面缺陷、内部组织及缺陷、晶体取向和低周疲劳试验条件等四方面因素,从位错运动方式和形态变化特点出发,研究了不同温度下镍基铸造合金的变形机制,最后总结了合金低周疲劳寿命预测的应力应变准则、能量准则、损伤累积准则及临界面和临界距离准则。     

Evaluation of service-induced damage and restoration of cast turbine blades

Conventionally cast turbine blades of Inconel 713C, from a military gas turbine aircraft engine, have been investigated with regard to service-induced microstructural damage and residual creep life time. For cast turbine blades, service life is defined by statistical values. The statistical methods can prove to be uneconomical, because safe limits must be stated with regard to the statistical probability that some blades will have higher damage than normal. An alternative approach is to determine the service-induced microstructural damage on each blade, or a representative number of blades, to better optimize blade us-age. Ways to use service-induced γ rafting and void formation as quantified microstructural damage pa-rameters in a service lifetime prediction model are suggested. The damage parameters were quantified, in blades with different service exposure levels, and correlated to remaining creep life evaluated from creep test specimens taken from different positions of serviced blades. Results from tests with different rejuvenation treatments, including hot isostatic pressing andJor heat treatment, are discussed briefly.     

核电汽轮机60Si2MnA弹簧钢垫片的老化行为

核电厂汽轮机叶片叶根底部的60Si2MnA弹簧垫片在服役过程中受到潜在的腐蚀、疲劳、应力松弛等多种老化效应的影响。针对某核电厂汽轮机低压转子末级、次末级叶片根部弹簧垫片,综合分析其服役过程中潜在的腐蚀、磨损及疲劳等方面影响,对垫片老化行为和服役寿命进行研究。结果表明,已服役10年旧弹簧垫片表面存在微动磨痕,微观上未见明显的腐蚀产物及腐蚀形貌,在长期服役过程中,腐蚀效应对垫片寿命的影响有限,不是限制其服役寿命的主要因素。由于材料在初始循环阶段表现出明显的强化和硬化特性,拆卸后的弹簧垫片与原始状态之间存在一定差异,电厂应明确初始安装指导文件不能用于已安装过的弹簧垫片。基于当前的运行状态和试验保守分析预测,已服役10年旧弹簧垫片至少可以继续服役5年,可延长现有检修和更换周期,节约大修工期及成本。     

Residual stresses as a factor in the selection of tungsten carbide coatings for a jet engine application

Tungsten carbide thermal spray coatings are important to the aerospace industry for the mitigation of midspan damper wear on jet engine fan and compressor blades. However, in some cases the coating can fail due to spallation and cracking, and in other situations the fatigue life of a fan or compressor blade is reduced when a coating is applied. Coating failures can result in decreased engine performance and costly maintenance time. A comprehensive experimental research program was conducted to evaluate coating crack resistance in bending, low-cycle fatigue properties of the coating and substrate, coating performance in jet engine tests, and microstructures for a wide range of coating compositions and application processes. Coating residual stress distributions also were evaluated. Eleven coatings were ranked according to their performance relative to the other coatings in each evaluation category. Results from the bend and low-cycle fatigue evaluations were compared to the experimentally evaluated residual stresses. Comparisons of rankings indicate a strong correlation between performance and the residual stresses in the coatings. Results from the program were used to select a suitable coating system for final in-service use based on two important criteria: (1) the coating must not fail while in service, and (2) the coating must not induce crack propagation into the substrate of the midspan damper.     

PVD TBC experience on GE aircraft engines

The higher performance levels of modern gas turbine engines present significant challenges in the reli-ability of materials in the turbine. The increased engine temperatures required to achieve the higher per-formance levels reduce the strength of the materials used in the turbine sections of the engine. Various forms of thermal barrier coatings have been used for many years to increase the reliability of gas turbine engine components. Recent experience with the physical vapor deposition process using ceramic material has demonstrated success in extending the service life of turbine blades and nozzles. Engine test results of turbine components with a 125 μm (0.005 in.) PVD TBC have demonstrated component operating tem-peratures of 56 to 83 °C (100 to 150 °F) lower than non-PVD TBC components. Engine testing has also revealed that TBCs are susceptible to high angle particle impact damage. Sand particles and other engine debris impact the TBC surface at the leading edge of airfoils and fracture the PVD columns. As the impacting continues, the TBC erodes in local areas. Analysis of the eroded areas has shown a slight increase in temperature over a fully coated area ; however, a significant temperature reduc-tion was realized over an airfoil without TBC.     

Development of a simulated bird-strike test method

An inexpensive test method has been developed to predict the dynamic impact response (bird-strike resistance) of aircraft turbine engine fan blade materials. The test utilizes an instrumented pendulum to transfer energy to a machined fan blade cross section. To validate the test, the relative performance of one wrought and four cast Ti-6Al-4V materials subjected to the simulated bird-strike test was compared to that of the same materials tested at Wright-Patterson Air Force Base by conventional bird-strike testing. The relative ranking of crack initiation loads determined from the simulated test corresponded to fan blade failure velocities determined by the Air Force test, thereby validating the simulated method as a screening tool for fan blade material development.     

LZ50车轴钢在复合微动条件下的研究

微动疲劳损伤广泛存在于各种机械接触变载荷作用的构件上,如螺栓,轴承,键槽和榫槽等.微动疲劳会加速受微动作用构件的接触处表面及表层裂纹的萌生和扩展,在微动疲劳的早期阶段裂纹生长速率较高,导致了在微动条件下金属构件过早失效,大幅度降低构件寿命.本文以LZ50车轴钢为主要研究对象,实现了在圆形和椭网形路径加载下的拉扭复合微动...     

Effect of nitrogen ion implantation dose on torsional fretting wear behavior of titanium and its alloy

Various doses of nitrogen ions were implanted into the surface of pure titanium, Ti6Al7Nb and Ti6Al4V, by plasma immersion ion implantation. Torsional fretting wear tests involving flat specimens of no-treated and treated titanium, as well as its alloys, against a ZrO₂ ball contact were performed on a torsional fretting wear test rig using a simulated physiological medium of serum solution. The treated surfaces were characterized, and the effect of implantation dose on torsional fretting behavior was discussed in detail. The results showed that the torsional fretting running and damage behavior of titanium and its alloys were strongly dependent on the dose of the implanted nitrogen ions and the angular displacement amplitude. The torsional fretting running boundary moved to smaller angular displacement amplitude, and the central light damage zone decreased, as the ion dose increased. The wear mechanisms of titanium and its alloys were oxidative wear, abrasive wear and delamination, with abrasive wear as the most common mechanism of the ion implantation layers.     

40CrMnMo活塞连杆稀土离子氮碳共渗工艺

对40CrMnMo钢制活塞连杆试样进行不同温度、时间、气压及稀土含量下的稀土离子氮碳共渗工艺试验,测定了不同工艺参数下表面强化层的硬度、厚度,并对比了稀土离子氮碳共渗工艺与其他表面强化工艺处理后的磨损性能和弯曲疲劳强度等力学性能.结果表明:5%RE、520℃×2 h、1.05 kPa气压为该工艺的最佳参数.40CrMnMo钢制高速增压柴油机活塞连杆用该工艺处理后,完全能满足其使用性能要求.     

Research Progress on Surface Impact Strengthening Mechanisms and Application of Nickel-Based SuperalloysEI北大核心CSCD

为满足不断攀升的两机涡轮动力系统的快速发展,表面冲击强化技术在涡轮转子用高温合金表面强化的应用及相应机制的研究受到了广泛关注。然而,高温合金表面硬化层在高温服役环境下的回复、再结晶行为难以避免,由此引起的表面强韧化、抗疲劳效果的退化,成为制约表面冲击强化技术在先进高温合金关键部件深入应用的瓶颈。本文总结了近年来镍基高温合金表面冲击强化机制及应用研究进展,分析了表面冲击强化对镍基高温合金表面强韧性及抗疲劳的作用规律,探究了高温合金表面冲击硬化层在高温及长期时效过程中的显微组织、微结构演化及其对高温稳定性的作用机理。以期为发展镍基高温合金表面冲击强化、提高两机涡轮转子疲劳抗力提供基础。     

The effects of multiple re-shot peening on fretting fatigue behavior of Al7075-T6

Fretting fatigue is a consequence of small oscillatory movement between two contacting parts. This type of damage may give rise to significant reduction in fatigue life of the components. The resistance of materials against fretting fatigue can be improved by surface treatment such as shot-peening. In this work, the effect of multiple re-shot peening on resistance of Al7075-T6 against fretting fatigue was investigated. After each re-shot peening, the specimen was subjected to 80% of the fatigue life corresponding to the cycles to failure of the specimen after re-shot peening. This process continued until the effect of any further re-shot peening became insignificant. The results showed almost a 100% enhancement in fatigue life for the first re-shot peening. The enhancement, however, reduced logarithmically for the next re-shot peenings such that for the 4th re-shot peening the increase of fatigue life fell below 2% which was negligible. On the whole, the fretting fatigue life increased by 600% after shot peening and 4 times re-shot peenings.     

Gas Turbine Oxidation Life Assessment and Monitoring

Gas turbine hot-section parts can be very expensive components with a finite lifetime. Their durability is strongly dependent on the operating service conditions which control the maintenance intervals and associated expenses. Blade damage is the most frequent reason for failures in gas turbine engines and also oxidation is one of the most critical degradation mechanisms when the power system operates in partially loading condition. This paper describes the methodology of oxidation life assessment for uncoated and coated blades with various metallic coatings in the first stage of a typical gas turbine engine. The resultant weight-change curves were validated by comparing them with the experimental and numerical data from the various references. Also a computer code was developed for real-time monitoring of remaining oxidation life and damage on the maximum blade temperature where the maximum damage accumulation occurs. Results indicated that the oxidation life consumption increases if the duration of cycle decreases or the temperature of the metal is increased during real-time operational condition. Also, considering a duplex metallic coating on the first-stage blade, the rate of oxidation damage reduces about 10 times compared to an uncoated blade.     

Numerical simulation on vacuum solution heat treatment and gas quenching process of a low rhenium-containing Ni-based single crystal turbine blade

Numerical heat-transfer and turbulent lfow model for an industrial high-pressure gas quenching vacuum furnace was established to simulate the heating, holding and gas fan quenching of a low rhenium-bearing Ni-based single crystal turbine blade. The mesh of simpliifed furnace model was built using ifnite volume method and the boundary conditions were set up according to the practical process. Simulation results show that the turbine blade geometry and the mutual shielding among blades have signiifcant inlfuence on the uniformity of the temperature distribution. The temperature distribution at sharp corner, thin wal and corner part is higher than that at thick wal part of blade during heating, and the isotherms show a toroidal line to the center of thick wal. The temperature of sheltered units is lower than that of the remaining part of blade. When there is no shelteration among multiple blades, the temperature distribution for al blades is almost identical. The lfuid velocity ifeld, temperature ifeld and cooling curves of the single and multiple turbine blades during gas fan quenching were also simulated. Modeling results indicate that the loading tray, free outlet and the location of turbine blades have important inlfuences on the lfow ifeld. The high-speed gas lfows out from the nozzle is divided by loading tray, and the free outlet enhanced the two vortex lfow at the end of the furnace door. The closer the blade is to the exhaust outlet and the nozzle, the greater the lfow velocity is and the more adequate the lfow is. The blade geometry has an effect on the cooling for single blade and multiple blades during gas fan quenching, and the effects in double layers differs from that in single layer. For single blade, the cooing rate at thin-waled part is lower than that at thick-waled part, the cooling rate at sharp corner is greater than that at tenon and blade platform, and the temperature at regions close to the internal position is decreased more slowly than that close to the surface. For multiple blades in single layer, the temperature at sharp corner or thin wal in the blade that close to the nozzles is much lower, and the temperature distribution of blades is almost paralel. The cooling rate inside the air current channel is lower than that of at the position near blade platform and tenon, and the effect of blade location to the nozzles on the temperature ifeld inside the blade is lower than that on the blade surface. For multiple blades in double layers, the lfow velocity is low, and the lfow is not uniform for blades in the second-layer due to the shielding of blades in the ifrst-layer. the cooling rate of blades in the second-layer is lower than that in the ifrst-layer. The cooling rate of blade close to the nozzles in the ifrst-layer is the higher than that of blade away from the nozzles in the second-layer, and the temperature distribution on blades in the same layer is almost paralel. The cooling rate in thin wal position of blade away from the nozzles is larger than that in tenon of the blade closer to the nozzles in the same layer. The cooling rate for blades in the second-layer is much lower both in thin wal and tenon for blades away from the nozzles.     

表面再结晶晶粒对涡轮叶片DZ4合金疲劳性能的影响

目的 了解表面再结晶晶粒影响涡轮叶片DZ4合金疲劳失效的机制,从而找出提高疲劳寿命的途径.方法 采用我国自主研发的涡轮叶片用定向凝固合金DZ4.对合金表面采用不同压力(0.1、0.3、0.5 MPa)进行喷丸处理,随后进行1220℃的热处理,合金表面塑性变形层发生再结晶形成再结晶晶粒.通过扫描电子显微镜等对不同条件下获得的再结晶微观组织进行观察分析.对于含表面再结晶层的DZ4试样和原始DZ4试样分别开展低周疲劳测试,以获得其循环变形行为和低周疲劳寿命.结果 在较低的喷丸压力下(如0.1 MPa),表层再结晶组织以离散的再结晶晶粒出现.在较高喷丸压力下(0.3 MPa和0.5 MPa),获得了完全的表面再结晶层,而且随喷丸压力的增大,再结晶层厚度增加.疲劳试验结果 显示,在较低喷丸压力(0.1 MPa)下获得的再结晶试样,其疲劳寿命比原始DZ4合金的明显降低;在0.5 MPa下获得的含再结晶层的DZ4试样,具有比原始DZ4合金更高的疲劳寿命.结论 定向凝固合金DZ4表面的再结晶层并不一定会降低其疲劳寿命.在较低喷丸压力(如0.1 MPa)及退火条件下形成的离散状再结晶晶粒对疲劳性能有害,疲劳寿命降低明显.在较高喷丸压力(如0.5 MPa)和退火条件下形成的致密细晶再结晶层,可提高DZ4合金的低周疲劳性能.晶体塑性模型有助于理解和预测再结晶层对疲劳性能的影响.     

Estimation of spallation life of thermal barrier coating of gas turbine blade by thermal fatigue test

Plasma-sprayed thermal barrier coatings (TBCs) are applied to protect the blades of a gas turbine system from high-temperature gas and to lower the surface temperature of the blades. The failure of TBC is directly connected to the failure of the blades because the spallation of a ceramic layer leads to the acceleration of local corrosion and oxidation at the location of failure. Therefore, the spallation life of TBC is very important in the evaluation of the reliability of a gas-turbine blade.In this study, thermal fatigue tests were performed at 1100 °C and 1151 °C. Then, c-scanning and bond strength tests were performed for TBC specimens that were thermally aged by thermal fatigue tests. From the results, an empirical equation based on the ratio of the delamination area and the thermal cycle number was presented and the spallation life of a TBC specimen could be roughly estimated using the relationship between the delaminated area and the number of cycles.     

叶片质量不平衡对基础环式风机基础服役性能影响研究

风力发电机叶片长期暴露于恶劣的高空环境，极易产生质量不平衡故障，对风力机的稳定可靠运行影响极大。传统的叶片质量不平衡研究主要集中于对风力发电机组功率和塔筒的分析。基础环式基础广泛应用于风力发电机，在循环荷载作用下易发生损伤破坏。因此，基于叶素动量等理论建立考虑叶片质量不平衡故障的基础环式风机基础服役性能评估模型，模型利用解析方法计算质量失衡叶片对基础造成的附加荷载，结合平衡条件和连续性条件，选用大型有限元软件ABAQUS模拟分析叶片质量不平衡故障对风力机基础应力、疲劳寿命以及侧壁混凝土裂缝发展规律的影响。结果发现：质量失衡叶片在水平方向对风力机基础疲劳寿命的影响最大，叶片质量失衡故障会加快基础混凝土损伤过程，增加基础环的水平度，进而降低风力机基础整体服役的可靠性。此外，建议风力机发生较小叶片质量失衡损伤时通过降低转速维持风力机运行，并定期监测叶片质量失衡比。     

TC4钛合金空心风扇叶片面板成形缺陷的蠕变修复

航空发动机用钛合金宽弦空心风扇叶片具有独特的结构特点,其外部为等厚面板,内部减重空腔与加强筋交替排列。成形中当叶片承受弯扭变形作用时,其局部面板会发生失稳产生凹陷,导致叶型超差。为了解决这一问题,提出了一种基于蠕变成形的钛合金空心风扇叶片面板修复方法,该方法采用气压加载,在较高的温度下,使凹陷部位发生蠕变变形,恢复到理论外形。针对某型TC4钛合金宽弦空心风扇叶片试验件进行面板气胀蠕变修复工艺的研究,对经历2次热循环的基体材料进行了高温蠕变拉伸试验,建立了材料蠕变模型,利用有限元软件ABAQUS对面板蠕变过程进行了数值模拟,考察了保压压力以及蠕变时间对于修复效果的影响,并选取了最佳的工艺参数进行实验。结果表明,运用气压加载蠕变修复工艺,采用最大保压压力为4 MPa的分段加压曲线,在750℃下经过1 h的气压加载,能够有效修复钛合金空心风扇叶片面板在弯扭预成形中产生的凹陷,可以作为辅助该类零件制造的一种有效方法。