CN108263639A - Aircaft configuration key position fatigue life on-line monitoring method based on indirect measuring strain under spectrum carries - Google Patents
Aircaft configuration key position fatigue life on-line monitoring method based on indirect measuring strain under spectrum carries Download PDFInfo
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Abstract
本发明公开了谱载下基于间接测量应变的飞机结构关键部位疲劳寿命在线监测方法，涉及飞机结构疲劳裂纹萌生与扩展的监测技术领域，该方法的步骤为：(1)利用有限元分析方法计算结构件危险点并标定传感器位置和应力集中系数；(2)安装传感器进行在线应变监测；(3)使用雨流计数法进行对监测的应变进行循环计数，并求得危险点的局部应力应变；(4)计算循环的损伤累计，并判断是否开始裂纹萌生监测；(5)计算裂纹萌生评估参量，对比评估参量与判别参数判断疲劳裂纹的萌生情况后选择是否继续；(6)考察损伤累计结果后选择是否继续。监测结果说明该方法能较好的监测飞机结构件关键部位疲劳裂纹寿命情况。
The invention discloses an online monitoring method for the fatigue life of key parts of an aircraft structure based on indirect measurement of strain under spectral loading, and relates to the technical field of monitoring the initiation and growth of fatigue cracks in aircraft structures. Dangerous points of structural parts and calibration of sensor positions and stress concentration factors; (2) Install sensors for on-line strain monitoring; (3) Use the rainflow counting method to perform cyclic counting of monitored strains and obtain local stress and strain at dangerous points; (4) Calculate the cumulative damage of the cycle, and judge whether to start crack initiation monitoring; (5) Calculate the evaluation parameters of crack initiation, compare the evaluation parameters and discrimination parameters to judge the initiation of fatigue cracks, and choose whether to continue; (6) Investigate the cumulative damage results Then choose whether to continue. The monitoring results show that this method can better monitor the life of fatigue cracks in key parts of aircraft structures.
Description
技术领域technical field
本发明应用领域是疲劳寿命监测方向，特指一种谱载下基于间接测量应变的飞机结构关键部位疲劳寿命在线监测方法。The application field of the invention is the direction of fatigue life monitoring, in particular an online monitoring method for fatigue life of key parts of an aircraft structure based on indirect measurement of strain under spectral loading.
背景技术Background technique
飞机在国家的经济领域、交通领域、军事领域占据着重要地位，由飞机疲劳导致的结构断裂，会造成严重的事故发生。飞机结构断裂事故往往是由一条疲劳裂纹的萌生开始的，当疲劳裂纹扩展到一定长度，会直接引发结构破坏导致事故。因此，本发明提出了一种针对飞机结构关键部位的疲劳寿命在线监测方法，保证飞机安全可靠服役，具有重要的实际意义。Aircraft occupy an important position in the country's economic, transportation, and military fields. Structural fractures caused by aircraft fatigue will cause serious accidents. Aircraft structural fracture accidents often start with the initiation of a fatigue crack. When the fatigue crack expands to a certain length, it will directly cause structural damage and lead to accidents. Therefore, the present invention proposes an online fatigue life monitoring method for key parts of the aircraft structure to ensure safe and reliable service of the aircraft, which has important practical significance.
目前的寿命在线监测方法通常是在飞机关键部位应力集中部位周围安装应变传感器，以此来估算危险点的应力应变来计算寿命，但是这种传统的计算方法得到的寿命一般与实际寿命相差2倍左右，甚至误差更大，而裂纹监测往往需要更复杂的传感器并且太过于灵敏会出现误报警的情况，因此，在传统方法的基础上叠加一种同样基于应力集中部位周围应变在线测量的疲劳裂纹萌生监测方法，能够更加接近实际的在线监测寿命。所提出的方法也为其他机械结构关键部位的寿命在线监测提供了一项具有应用价值的技术。The current life on-line monitoring method usually installs strain sensors around the stress concentration parts of the key parts of the aircraft to estimate the stress and strain at the dangerous point to calculate the life, but the life obtained by this traditional calculation method is generally twice the actual life. Left and right, even larger errors, and crack monitoring often requires more complex sensors and is too sensitive to cause false alarms. Therefore, on the basis of traditional methods, a fatigue crack that is also based on online measurement of strain around stress concentration parts is superimposed The sprouting monitoring method can be closer to the actual online monitoring life. The proposed method also provides a valuable technology for the life online monitoring of key parts of other mechanical structures.
发明内容Contents of the invention
本发明目的在于为满足飞机结构疲劳寿命监测的需求，提出了一种谱载下基于间接测量应变的飞机结构关键部位疲劳寿命在线监测方法，该方法也适用于监测其它机械结构的关键部位的疲劳寿命。The purpose of the present invention is to meet the needs of fatigue life monitoring of aircraft structures, and propose an online monitoring method for fatigue life of key parts of aircraft structures based on indirect measurement of strain under spectral load. This method is also suitable for monitoring the fatigue of key parts of other mechanical structures life.
本发明所提供的技术方案为一种谱载下基于间接测量应变的飞机结构关键部位疲劳寿命在线监测方法，其步骤为：The technical solution provided by the present invention is an online monitoring method for fatigue life of key parts of an aircraft structure based on indirect measurement of strain under spectral load, the steps of which are as follows:
步骤1)：使用有限元方法确定所要监测结构关键部位危险点的位置，同时标定要在实际结构安装应变传感器的位置，1号传感器安装在危险点正对的背面，2号传感器安装在1号传感器同侧，与1号传感器的距离需要参考飞机结构件的实际厚度，保证两个传感器距离不小于飞机结构件厚度且不能大于飞机结构件厚度的两倍，并保证两个传感器位置的连线与预测裂纹扩展平面垂直，使用2号传感器和危险点的应力来确定这之间的应力集中系数Kt以用来后面计算监测部位的局部应力应变；Step 1): Use the finite element method to determine the position of the dangerous point in the key part of the structure to be monitored, and at the same time calibrate the position of the strain sensor to be installed in the actual structure. The No. 1 sensor is installed on the back facing the dangerous point, and the No. 2 sensor is installed on the No. 1 On the same side as the sensor, the distance from the No. 1 sensor needs to refer to the actual thickness of the aircraft structural part. Ensure that the distance between the two sensors is not less than the thickness of the aircraft structural part and cannot be greater than twice the thickness of the aircraft structural part, and ensure the connection between the positions of the two sensors. Perpendicular to the predicted crack propagation plane, use the stress of the No. 2 sensor and the dangerous point to determine the stress concentration factor K t between them, which will be used to calculate the local stress and strain of the monitoring part later;
步骤2)：根据步骤1)确定的传感器位置，在飞机结构件上安装应变传感器，用于实时监测1号传感器和2号传感器的应变值，分别记为ε1和ε2；Step 2): According to the sensor position determined in step 1), strain sensors are installed on the aircraft structure for real-time monitoring of the strain values of No. 1 sensor and No. 2 sensor, which are respectively denoted as ε 1 and ε 2 ;
步骤3)：使用雨流计数法对接收监测的一个载荷块中的ε2时间历程进行循环计数，同时使用该结构材料的循环应力应变曲线，根据Neuber法确定危险点的局部应力应变；Step 3): Use the rainflow counting method to cycle count the ε time history in a load block received and monitored, and use the cyclic stress-strain curve of the structural material to determine the local stress-strain at the dangerous point according to the Neuber method;
步骤4)：使用Smith公式计算由雨流计数对接收的载荷块ε2时间历程中所提取的第i个循环对应的疲劳损伤Di，如下式所示，Step 4): Calculate the fatigue damage D i corresponding to the i-th cycle extracted from the time history of the received load block ε 2 by using the Smith formula, as shown in the following formula,
σmax——该循环的最大应力；σ max ——the maximum stress of this cycle;
Δε——该循环的应变范围；Δε——the strain range of the cycle;
σ′f——疲劳强度系数；σ′ f — fatigue strength coefficient;
ε′f——疲劳塑性系数；ε′ f — fatigue plasticity coefficient;
E——杨氏模量；E - Young's modulus;
Ni——第i个循环对应的寿命；N i ——the life corresponding to the i-th cycle;
b——疲劳强度指数；b——fatigue strength index;
c——疲劳塑性指数；c——fatigue plasticity index;
使用Miner定理对所接收载荷块的总损伤D进行累积，其中The total damage D of the received load blocks is accumulated using Miner's theorem, where
当总损伤D累积到0.5时，继续下一步，当D＜0.5时，回到步骤3)继续采集下个载荷块进行循环计算；n表示循环总次数。When the total damage D is accumulated to 0.5, continue to the next step, and when D<0.5, return to step 3) and continue to collect the next load block for cyclic calculation; n represents the total number of cycles.
步骤5)：计算所接收载荷块中ε1和ε2最大值差值的绝对值，求出该绝对值随载荷块数增加的曲线的导数作为疲劳裂纹萌生评估参量K，当疲劳裂纹萌生评估参量K超过判别参数Kc时，表明飞机结构关键部位有疲劳裂纹萌生，即疲劳裂纹形成，立即警告停止使用或进行检测维修，如果未超过判别参数Kc则继续步骤6)。Step 5): Calculate the absolute value of the difference between the maximum values of ε1 and ε2 in the received load block, and obtain the derivative of the curve of the absolute value increasing with the number of load blocks as the fatigue crack initiation evaluation parameter K, when the fatigue crack initiation evaluation When the parameter K exceeds the discriminant parameter Kc , it indicates that there is fatigue crack initiation in the key parts of the aircraft structure, that is, the formation of fatigue cracks. Immediately warn to stop using or carry out inspection and maintenance. If the discriminant parameter Kc is not exceeded, continue to step 6).
步骤6)：考察总损伤D累积情况，当D≥0.9时，表征预测的剩余寿命已不足10％，警报提示寿命消耗接近极限，如果选择继续使用则回到步骤3)继续计算，否则停止使用飞机并检测维修，当D＜0.9时，回到步骤3)继续计算。Step 6): Investigate the accumulation of total damage D. When D ≥ 0.9, the predicted remaining life is less than 10%, and the alarm prompts that the life consumption is close to the limit. If you choose to continue using it, go back to step 3) to continue the calculation, otherwise stop using it Aircraft and inspection maintenance, when D<0.9, return to step 3) to continue calculation.
将飞机一个飞行起落中待监测飞机结构件的受载情况即载荷谱定为一个载荷块。The loading condition of the structural parts of the aircraft to be monitored during one flight take-off and landing of the aircraft, that is, the load spectrum, is defined as a load block.
所述的判别参数预先由有限元模拟实际裂纹扩展件或采用模拟件疲劳试验进行标定。The discriminant parameters are calibrated in advance by finite element simulation of the actual crack growth part or fatigue test of the simulated part.
与现有技术相比，本发明具有如下有益效果。Compared with the prior art, the present invention has the following beneficial effects.
本发明的优点在于：提出了一种谱载下基于间接测量应变的飞机结构关键部位疲劳寿命在线监测方法。该方法所利用的应变传感器无需安装在飞机结构件危险点部位，而是通过在危险点部位另外一侧安装应变传感器来考察危险点周围的应变变化，以此来间接反映危险点部位的疲劳裂纹萌生和疲劳寿命情况，因此对应变传感器的尺寸和类型没有严格的限制，在严苛的工作环境下也可以根据需要选择合适的应变传感器，更利于应用到各种环境下进行疲劳裂纹的实时监测，例如，高温、燃油等环境。并且该方法在传统寿命预测计算的方法上增加了基于同一套应变传感器的疲劳裂纹萌生寿命的监测方法，这样既能够发挥传统疲劳寿命预测方法的预警作用，又能够规避前期的疲劳裂纹萌生监测的误报警现象，保证了寿命在线监测的准确性和可靠性。The invention has the advantages of proposing an online monitoring method for the fatigue life of key parts of the aircraft structure based on indirect measurement of strain under spectral load. The strain sensor used in this method does not need to be installed at the dangerous point of the aircraft structure, but by installing the strain sensor on the other side of the dangerous point to investigate the strain change around the dangerous point, so as to indirectly reflect the fatigue crack at the dangerous point Initiation and fatigue life conditions, so there are no strict restrictions on the size and type of strain sensors, and suitable strain sensors can also be selected according to needs in harsh working environments, which is more conducive to real-time monitoring of fatigue cracks in various environments , such as high temperature, fuel and other environments. Moreover, this method adds a fatigue crack initiation life monitoring method based on the same set of strain sensors to the traditional life prediction calculation method, which can not only play the early warning role of the traditional fatigue life prediction method, but also avoid the fatigue crack initiation monitoring in the early stage. False alarm phenomenon ensures the accuracy and reliability of online life monitoring.
附图说明Description of drawings
图1本发明方法实现疲劳裂纹在线监测的流程图。Fig. 1 is a flow chart of realizing online monitoring of fatigue cracks by the method of the present invention.
图2本发明方法的应变传感器安装示意图。Fig. 2 is a schematic diagram of installation of the strain sensor in the method of the present invention.
图3本发明方法应用到某飞机结构件的疲劳裂纹监测效果图。Fig. 3 is an effect diagram of the fatigue crack monitoring applied to an aircraft structural part by the method of the present invention.
具体实施方式Detailed ways
结合附图说明本发明的具体实施方式。The specific embodiment of the present invention will be described with reference to the accompanying drawings.
本发明通过飞机结构件疲劳试验对本发明作了进一步说明，The present invention has been further illustrated to the present invention by aircraft structure component fatigue test,
一种谱载下基于间接测量应变的飞机结构关键部位疲劳寿命在线监测方法，具体计算方法如下：An online monitoring method for fatigue life of key parts of aircraft structure based on indirect measurement of strain under spectral loading, the specific calculation method is as follows:
步骤1)：使用有限元方法确定本结构件关键部位危险点的位置，同时标定要在实际结构安装应变传感器的位置，1号传感器安装在危险点正对的背面，2号传感器安装在1号传感器同侧，与1号传感器的距离需要参考飞机结构件的实际厚度，保证两个传感器距离不小于飞机结构件厚度且不能大于飞机结构件厚度的两倍，并保证两个传感器位置的连线与预测裂纹扩展平面垂直，使用2号传感器和危险点的应力来确定这之间的应力集中系数Kt以用来后面计算部位的局部应力应变；Step 1): Use the finite element method to determine the position of the dangerous point of the key part of the structure, and at the same time calibrate the position of the strain sensor to be installed in the actual structure. The No. 1 sensor is installed on the back facing the dangerous point, and the No. 2 sensor is installed on the No. 1 On the same side as the sensor, the distance from the No. 1 sensor needs to refer to the actual thickness of the aircraft structural part. Ensure that the distance between the two sensors is not less than the thickness of the aircraft structural part and cannot be greater than twice the thickness of the aircraft structural part, and ensure the connection between the positions of the two sensors. Perpendicular to the predicted crack propagation plane, use the No. 2 sensor and the stress at the dangerous point to determine the stress concentration factor K t between them, which is used to calculate the local stress and strain of the position later;
步骤2)：根据步骤1)确定的传感器位置，在该飞机结构件上安装应变传感器，用于实时监测1号和2号传感器的应变值，分别记为ε1和ε2；Step 2): According to the sensor position determined in step 1), strain sensors are installed on the aircraft structure for real-time monitoring of the strain values of No. 1 and No. 2 sensors, which are respectively denoted as ε 1 and ε 2 ;
步骤3)：使用雨流计数法对接收的载荷块ε2进行循环计数，同时使用该结构材料的循环应力应变曲线，根据Neuber法确定危险点的局部应力应变；Step 3): Use the rainflow counting method to cycle count the received load block ε2 , and use the cyclic stress-strain curve of the structural material to determine the local stress-strain at the dangerous point according to the Neuber method;
步骤4)：使用Smith公式计算第i个循环对应的寿命，如下所示，Step 4): Use the Smith formula to calculate the life corresponding to the i-th cycle, as shown below,
σmax——该循环的最大应力；σ max ——the maximum stress of this cycle;
Δε——该循环的应变幅值；Δε——the strain amplitude of the cycle;
σ′f——疲劳强度系数；σ′ f — fatigue strength coefficient;
ε′f——疲劳塑性系数；ε′ f — fatigue plasticity coefficient;
E——杨氏模量；E - Young's modulus;
Ni——第i个循环对应的寿命；N i ——the life corresponding to the i-th cycle;
b——疲劳强度指数；b——fatigue strength index;
c——疲劳塑性指数；c——fatigue plasticity index;
使用Miner定理对总损伤D进行累积，其中The total damage D is accumulated using Miner's theorem, where
当总损伤D累积到0.5时，继续下一步，当D＜0.5时，回到步骤3)继续循环计算；When the total damage D is accumulated to 0.5, continue to the next step, and when D<0.5, return to step 3) and continue the cycle calculation;
步骤5)：计算所接收载荷块中ε1和ε2差值的绝对值，求出该值随载荷块数增加的曲线的导数作为疲劳裂纹萌生评估参量K，当疲劳裂纹萌生评估参量K超过判别参数Kc(该判别参数可预先由有限元模拟实际裂纹扩展件或由模拟件疲劳试验进行标定)时，表征飞机结构关键部位有疲劳裂纹萌生，疲劳到寿，警告立即停止使用飞机进行检测维修，如果未超过判别参数Kc则继续下一步。Step 5): Calculate the absolute value of the difference between ε1 and ε2 in the received load block, and obtain the derivative of the curve of this value increasing with the number of load blocks as the fatigue crack initiation evaluation parameter K, when the fatigue crack initiation evaluation parameter K exceeds When the discriminant parameter K c (this discriminant parameter can be calibrated in advance by finite element simulation of the actual crack growth part or by the fatigue test of the simulated part), it indicates that there is fatigue crack initiation in the key parts of the aircraft structure, and the fatigue reaches the end of life, and it is warned to immediately stop using the aircraft for detection Maintenance, if the judgment parameter Kc is not exceeded, continue to the next step.
步骤6)：考察总损伤D累积情况，当D≥0.9时，表征理论预测的剩余寿命已不足10％，警报提示寿命消耗接近极限，如果选择继续使用则回到步骤3)继续计算，否则停止使用飞机并检测维修，当D＜0.9时，回到步骤3)继续计算。。Step 6): Investigate the accumulation of total damage D. When D≥0.9, the remaining life predicted by the characterization theory is less than 10%, and the alarm prompts that the life consumption is close to the limit. If you choose to continue using, go back to step 3) to continue the calculation, otherwise stop Use the aircraft and check the maintenance, when D<0.9, go back to step 3) to continue the calculation. .
本发明的优点在于：提出了一种谱载下基于间接测量应变的飞机结构关键部位疲劳寿命在线监测方法。该方法所利用的应变传感器无需安装在飞机结构件危险点部位，而是通过在危险点部位另外一侧安装应变传感器来考察危险点周围的应变变化，以此来间接反映危险点部位的疲劳裂纹萌生和疲劳寿命情况，因此对应变传感器的尺寸和类型没有严格的限制，在严苛的工作环境下也可以根据需要选择合适的应变传感器，更利于应用到各种环境下进行疲劳裂纹的实时监测，例如，高温、燃油等环境。并且该方法在传统寿命计算的方法上增加了基于同一套应变传感器的疲劳裂纹萌生的监测方法，既能够发挥传统疲劳寿命预测方法的预警作用，又能够规避前期的疲劳裂纹萌生监测的误报警现象，保证了寿命在线监测的准确性。The invention has the advantages of proposing an online monitoring method for the fatigue life of key parts of the aircraft structure based on indirect measurement of strain under spectral load. The strain sensor used in this method does not need to be installed at the dangerous point of the aircraft structure, but by installing the strain sensor on the other side of the dangerous point to investigate the strain change around the dangerous point, so as to indirectly reflect the fatigue crack at the dangerous point Initiation and fatigue life conditions, so there are no strict restrictions on the size and type of strain sensors, and suitable strain sensors can also be selected according to needs in harsh working environments, which is more conducive to real-time monitoring of fatigue cracks in various environments , such as high temperature, fuel and other environments. Moreover, this method adds a fatigue crack initiation monitoring method based on the same set of strain sensors to the traditional life calculation method, which can not only play the early warning role of the traditional fatigue life prediction method, but also avoid the false alarm phenomenon of fatigue crack initiation monitoring in the early stage , to ensure the accuracy of online life monitoring.
为了验证本发明提出的基于应变传感器间接测量的飞机结构关键部位疲劳寿命在线监测方法的效果，将本方法所得的监测的结果与实际观察测量的裂纹萌生情况比较，如图2所示。结果表明，当监测的总损伤D超过0.9后，并未发现裂纹继续使用，当监测的疲劳裂纹萌生评估参量K大于判别参数0.1时，停止使用，此时观测到了疲劳裂纹萌生，长度为2.22mm，对应寿命为2875个载荷块(飞行起落)，实时理论计算疲劳裂纹萌生寿命约2176个载荷块(裂纹刚萌生时的寿命，对应总损伤D累积到1时的寿命)，说明该方法及时地捕捉了疲劳裂纹的萌生，准确的达到了实时在线监测寿命的目的，同时也保留了传统疲劳寿命预测方法的预警作用，因此，提出的谱载下基于间接测量应变的飞机结构关键部位疲劳寿命在线监测方法能够精准的监测疲劳裂纹萌生寿命情况。In order to verify the effect of the online monitoring method for the fatigue life of key parts of the aircraft structure based on the indirect measurement of the strain sensor proposed by the present invention, the monitoring results obtained by this method are compared with the actual observation and measurement of crack initiation, as shown in Figure 2. The results show that when the monitored total damage D exceeds 0.9, no cracks are found to continue to be used. When the monitored fatigue crack initiation evaluation parameter K is greater than the discrimination parameter 0.1, the use is stopped. At this time, fatigue crack initiation is observed, and the length is 2.22mm. , the corresponding life is 2875 load blocks (flight take-off and landing), and the real-time theoretical calculation fatigue crack initiation life is about 2176 load blocks (the life when the crack is just initiated, corresponding to the life when the total damage D accumulates to 1), which shows that the method is timely It captures the initiation of fatigue cracks, accurately achieves the purpose of real-time online life monitoring, and also retains the early warning function of traditional fatigue life prediction methods. Therefore, the proposed online fatigue life of key parts of aircraft structures based on indirect measurement of strain The monitoring method can accurately monitor the life of fatigue crack initiation.
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Application publication date: 20180710 Assignee: Unmanned Intelligence (Beijing) Technology Co.,Ltd. Assignor: Beijing University of Technology Contract record no.: X2024980043439 Denomination of invention: Online monitoring method for fatigue life of key parts of aircraft structures based on indirect strain measurement under spectral load Granted publication date: 20181113 License type: Open License Record date: 20241230 |
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