Failure Analysis of High-Pressure Compressor Blade in an Aero Gas Turbine Engine

Failure of high-pressure compressor rotor blade in an aero gas turbine engine is analyzed to determine its root cause. Forensic and metallurgical investigations are carried out on the blade and failed parts. The failure of the platform ladder is found to the first in the chain of events that led to the compressor blade failure. The mode of failure in the blade is found to be fatigue and has originated from the damaged region on the leading edge caused by dislodgement of platform ladder. The failure has caused extensive damages in high-pressure compressor module and also in downstream turbine blades as a secondary effect.     

Cyclic strength of turbine blades made of cast nickel-based alloys

We study the cyclic strength and durability of service-exposed turbine blades made of IN-738, ZMI-3U, and éP539LM nickel alloys, some of them with a protective coating, upon a long operating time in gas turbine compressor sets. The blades of IN-738 alloy are found to have the highest cyclic strength. The cyclic strength of uncoated IN-738 blades is about 10% higher than that of coated blades. __________ Translated from Problemy Prochnosti, No. 2, pp. 5–14, March–April, 2007.     

Vibration based damage detection of rotor blades in a gas turbine engine

This paper describes the problems concerning turbine rotor blade vibration that seriously impact the structural integrity of a developmental aero gas turbine. Experimental determination of vibration characteristics of rotor blades in an engine is very important from fatigue failure considerations. The blades under investigation are fabricated from nickel base super alloy through directionally solidified investment casting process. The blade surfaces are coated with platinum aluminide for oxidation protection. A three dimensional finite element modal analysis on a bladed disk was performed to know the likely blade resonances for a particular design in the speed range of operation. Experiments were conducted to assess vibration characteristics of bladed disk rotor during engine tests. Rotor blade vibrations were measured using non-intrusive stress measurement system, an indirect method of blade vibration measurement utilizing blade tip timing technique. Abnormalities observed in the vibration characteristics of the blade tip timing data measured during engine tests were used to detect the blade damage. Upon disassembly of the engine and subsequent fluorescent penetrant inspection, it was observed that three blades of the rotor assembly were identified to have damaged. These are the blades that exhibited vibration abnormalities as a result of large resonant vibration response while engine tests. Further, fractographic analysis performed on the blades revealed the mechanism of blade failures as fatigue related. The root cause of blade failure is established to be high cycle fatigue from the engine run data history although the blades were put into service for just 6 h of engine operation.     

Failure analysis of two sets of aircraft blades

This paper describes the methodology employed for the failure analysis of aircraft blades and its application to two premature failed sets. The first one corresponds to the high pressure compressor manufactured in a 718 nickel base superalloy. The failure analysis carried out on this blade points towards foreign object damage (FOD). The second set belongs to the high pressure turbine of another engine. Scanning electron microscopy attributes the first fail to the premature failure by a thermo-mechanical fatigue mechanism of one blade with an inadequate microstructure. The remaining blades of this set, which possess a correct microstructure, failed due to the impacts of the debris generated by the fracture of the first one.     

Failure of the J79 Engine Compressor Blade Due to Stall

The fractured and damaged compressor rotor blades of a J79 engine have been examined. Optical, stereoscopic, microhardness testing, and SEM examinations were carried out to find out the causes of the fracture. The material of the blades was STS403 and blades were used without coating. From the 15th through the 17th compressor stages the rotor blades, stator vanes, combustion, and turbine sections were damaged. The fractured surface of the 17th blade show multiple origins, secondary cracking in the blade’s convex surface, and extensive propagation before separation and rougher surfaces. The metallographic analysis of the microstructure suggested work hardening. Based on the results, the cause of the fractured blade was high-amplitude fatigue, due to severe stall. After normal engine usage of 5 months, the blade fractured and blade fragments hit the other compressor blades, stators, and casing, and caused damage in both combustion and turbine sections.     

某发动机一级涡轮叶片断裂分析

以某发动机涡轮一级事故叶片为研究对象,对断裂叶片进行了宏、微观形貌观察,硬度测试以及断口成分检查,分析了叶片裂纹的产生与发展过程,探讨了叶片断裂失效的原因。结果表明,叶片断裂模式为机械疲劳断裂。提出了预防措施。     

GE9FA燃气轮机进气加热系统的控制

当9FA燃气轮机压气机进口温度低时,为防止压气机进口结冰和超过运行极限,9FA燃气轮机的运行范围将受到限制。为了降低排放和防止压气机进口结冰,根据压气机进口可转导叶(IGV)来控制进气加热。压气机进气加热是将部分压气机的排气循环到压气机进口。该系统有三个作用:1)防止压气机进口结冰;2)通过降低IGV角度来满足低负荷时的排放要求,降低压气机压比;3)保证压气机有足够的安全裕度。进气加热系统保证了压气机的全范围运行。     

Possibility of increasing the service life of damaged blades

Results are given of an investigation of the effectiveness of a method of strengthening the working titanium alloy blades of a gas turbine engine compressor. It is established that the proposed strengthening method can lengthen to more than double the service life of damaged blades. In addition it is shown that in these blades, in comparison with nonstrengthened blades, cracks develop in a narrower zone and predominantly on the side of the inlet edge.Translated from Problemy Prochnosti, No. 4, pp. 45–49, April, 1992.     

某型航空发动机压气机四级转子叶片失效分析

某型航空发动机压气机四级转子叶片是故障多发叶片.对叶片典型失效件进行了断口分析,发现叶片故障是由于叶片表面发生腐蚀导致疲劳强度降低,使得叶片在振动应力下发生疲劳失效.进一步对显微组织进行分析,发现叶片材质合乎要求,叶片腐蚀主要是由于使用环境因素造成的.研究结果对于叶片的故障分析及预防具有重要的意义.     

风力机叶片挥舞摆振气弹失稳分析

为避免风机叶片在强风作用下发生破坏,需对其采取停机保护措施。该文研究叶片处于非旋转状态时的挥舞摆振气弹失稳现象发生的条件。基于风力机叶片简化模型,采用迭代法求解叶片的自振频率及振型,建立了非旋转叶片挥舞摆振气弹效应响应的振型叠加法,该方法可以便捷地进行叶片多工况气动弹性响应分析。计算了在不同风速不同攻角条件下叶片的挥舞摆振气弹效应响应,得到了叶片挥舞摆振响应随风速和风攻角的变化规律以及不稳定风攻角的分布特征。结果表明:在某些风攻角下,风机叶片挥舞摆振失稳现象在风速较低的情况就有可能发生,其响应幅值与结构阻尼联系紧密。另外,挥舞摆振失稳会大大增加作用于叶片上的风荷载,并进一步造成叶片结构的损伤破坏。     

Investigation on the effects of blade corrosion on compressor performance

Corrosion is a common reason for marine gas turbine blade failure, and salt spray is largely responsible for corrosion of metallic objects near the coastline. It will directly lead to geometrical deviations of the compressor blade, forcing an increase in roughness height of the surface of the blade. By using a three-dimensional numerical method, this paper deals with the corrosion of a 1.5 stage axial compressor. Two different schemes, respectively based on blade geometrical model modification and turbulent near wall functions, were employed in numerical simulations for corrosion rate of the blades. The compressor map, derived from the numerical simulations by adjusting the back pressure, shows that there is a decline of compressor efficiency and pressure ratio when the blade is corroded, which will result in performance degradation of the compressor. The corrosion properties of stainless steel were studied, using the static salt spray corrosion test under laboratory conditions. Additionally, combined with numerical methods, the compressor performance parameters along the extension of operating time were simulated and a binomial relationship between performance degradation and corrosion time was established. This research provides a technical guide for compressor performance prediction under less than ideal conditions.     

Failure Prevention of Gas Turbine Blades Through Pack Aluminization

Pack aluminization of low-pressure turbine blade of an aero gas turbine engine has been carried out following a well-defined systematic procedure. The process parameters are first optimized on dummy blades, and optimized process is followed for the actual blades for evaluation and testing. Visual and binocular examination followed by metallurgical evaluation has been carried out to validate the process and to establish the adequacy and correctness of the coating. The coated blades are then evaluated through component-level test and engine-level test followed by field evaluation trials for performance and durability. The results of engine-level tests and inspection post-accelerated mission test cycles ensure that the blades with aluminide coating can withstand severe engine operating cycles without any damage or failure which otherwise would have experienced. The failure-free operation for an equivalent TBO life and post-AMT condition of blades are an indication of enhanced life of aluminide blades and prevention of failure of the turbine blades through pack aluminization.     

Investigation of the failure of the L-0 blades

A last stage (L-0) turbine blades failure was experienced at the 110 MW geothermal unit after 1 year of operation period. This unit has two tandem-compound intermediate/low-pressure turbines (turbine A and turbine B) with 23 in./3600 rpm last-stage blades. There were flexible blades continuously coupled 360 degrees around the row by loose cover segment at the tip and loose sleeve and lug at the mid-span (pre-twist design). The failed blades were in the L-0 row of the LP turbine B connected to the generator. The visual examination indicated that the group of 12 L-0 blades of rotor B on the generator side was bent and another group of 5 blades at 140 degrees from the first damaged group was also bent. The cover segments were spread out from the damaged blades and had cracks. Laboratory evaluation of the cracking in the cover segments indicates the failure mechanism to be high cycle fatigue (HCF), initiating at the cover segment holes outer fillet radius. The L-0 blades failure investigation was carried out. The investigation included a metallographic analysis of the cracked cover segments and bent blades, Finite Element Method (FEM) stress and natural frequency analysis (of blades/cover segments), fracture mechanics and crack propagation analysis. This paper provides an overview of the L-0 blades failure investigation, which led to the identification of the blades vibrations within the range 250–588 Hz induced due to unstable flow excitation (stall flutter) as the primary contribution to the observed failure.     

Failure of Low-Pressure Turbine Blades in Military Turbofan Engines: Causes and Remedies

Failure of low-pressure (LP) turbine rotor blades in low bypass military turbofan engines is a great concern for designers, manufactures, repair and overhaul agencies, operators, and airworthiness authorities. The present paper analyzes the LP turbine blade failure cases to determine its root cause. Forensic and metallurgical investigations are carried out on the failed blades. In most cases, the failure was originated from the leading edge and had propagated toward the trailing edge. Intergranular features and high oxidation on the fractured surface have been found as the cause of fatigue failure. Operation at elevated temperatures for considerable time was found responsible for these fatigue failures. Malfunction of fuel system, failure in control sensors, and nonuniformity in atomizer characteristics were the root cause of high temperature in turbine leading to the failure of blades. The paper also presents various remedial measures to address the blade failures from manufacturing and operational points of view.     

Coupled mechanical,metallurgical and FEM based failure investigation of steam turbine blade

Steam turbine blades are the critical component in power plant, specifically low pressure blades are generally found to be more susceptible to failure. A mechanical, metallurgical and FEM based coupled methodology is used in the present failure investigation of low pressure steam turbine blade. The results of each investigation of turbine blade failure were then interpreted that leads to find the location of primary failure, sequence of failure and the root cause of its failure. All the three aspects of failure investigation are important in answering the questions raised for the failure and to avoid any future miss-happening.     

基于Campbell理论的航空发动机涡轮叶片共振裕度分析

针对航空发动机涡轮叶片的共振特性会导致叶片发生疲劳断裂、振动失效等问题,本文以某型号航空发动机涡轮叶片为研究对象,开展共振裕度分析研究。首先基于试验自锤击法和有限元物理仿真计算两种方法同步分析叶片的振动特性,通过提取叶片在两种工作状态下前6阶的模态分析结果,验证了该模型的正确性与实用性。其次在已有模型基础上,通过绘制不同工作转速下的Campbell共振曲线图,结合该型号发动机的实际工况参数,进行了转速共振裕度的校核分析,对叶片上可能发生共振的工作转速进行了解析并提出优化及改进方案。本方法主要是为叶片的前期设计制造及共振安全性检验问题提供了充足的科学依据和方法。     

Identification of failure mechanisms in nickel base superalloy turbine blades through microstructural study

Failures in gas turbine blades can occur by various mechanisms that are operative at high temperatures. Most frequent ones are the creep and stress rupture. These mechanisms are strongly dependent on the microstructure of the blades and often, they are facilitated due to microstructural degradation that occurs during service exploitation. The study of these failures, therefore, requires detailed microstructural examination. This paper reports the basic microstructural features in cast nickel base superalloys that control the failure mechanisms in gas turbine blades followed by analysis of two service failures. A methodology has been suggested for effective microstructural analysis of turbine blades after service exposure. It has been shown how systematic microstructural study on the turbine blades can be helpful in establishing engine operating conditions.     

On the Problem of Improvement of the Damping Ability of Rotor Blades of Contemporary Gas-Turbine Engines

We present the results of experimental investigations of the influence of structural and technological factors on the damping ability of lock joints of compressor blades carried out at the Pisarenko Institute of Problems of Strength of the National Academy of Sciences of Ukraine, study the possibility of using dampers mounted in shroud joints of turbine blades, and formulate some recommendations concerning the possibility of im proving the damping ability of rotor blades of contemporary gas-turbine engines.     

Failure analysis of a second stage blade in a gas turbine engine

The failure of a second stage blade in a gas turbine was investigated by metallurgical and mechanical examinations of the failed blade. The blade was made of a nickel-base alloy Inconel 738LC. The turbine engine has been in service for about 73,500 h before the blade failure at 5:50 PM on 14 August 2004. Due to the blade failure, the turbine engine was damaged severely. The investigation was started with a thorough visual inspection of the turbine and the blades surfaces, followed by the fractography of the fracture surfaces, microstructural investigations, chemical analysis and hardness measurement.

The observation showed that a serious pitting was occurred on the blade surfaces and there were evidences of fatigue marks in the fracture surface. The microstructural changes were not critical. It was found that the crack initiated by the hot corrosion from the leading edge and propagated by fatigue and finally, as a result of the reduction in cross-section area, fracture was completed.

An analytical calculation parallel to the finite element method was utilized to determine the static stresses due to huge centrifugal force. The dynamic characteristics of the turbine blade were evaluated by the finite element modal and harmonic analyses. Finally according to the log sheet records and by using a Campbell diagram there was a good agreement between the failure signs and FEM results which showed the broken blade has been resonated by the third vibrational mode occasionally before the failure occurred.