**3. Maintenance process for fuselage structures**

#### **3.1 Corrective maintenance procedure**

The most widely used active sensor is the piezoelectric wafer active sensor (PWAS), which uses ultrasonic lamb waves. As an actuator, it converts the electric signal to mechanical motion to generate a longitudinal or transverse wave, which propagates on the panel and is reflected at a crack. As a sensor, it receives a wave reflected from a crack and converts it to electric signals. The location and size of damage are estimated by measuring the time, amplitude, or frequency of the

*Reliability and Maintenance - An Overview of Cases*

In general, two methods are used to detect damage [6]. In the pulse-echo method, one PWAS sends waves and receives waves reflected at a crack. In the pitch-catch method, one PWAS sends waves, and the other PWAS receives the waves. In addition, several PWAS, called a phase array, are used simultaneously to improve detection capability [12]. Although the abovementioned two methods require undamaged (pristine) state, the time reversal method [13] does not require it. Since the mechanism of detecting damage using PWAS is similar to conventional NDT ultrasonics, the detectable damage size is also similar to NDT. The most preferable feature of PWAS is its capability of detecting a remote damage from the sensor. Giurgiutiu [14] showed a lamb wave tuning method to detect a remote damage effectively. It has been shown that PWAS can be used for both metallic and composite panels [15]. In order to reduce the excessive number of wires to connect sensors, SMART layer [16] is developed by printing circuits of 30 sensors into a thin

Fiber Bragg grating (FBG) uses a series of parallel lines of optical fiber with different refractive indices [17]. When a local strain is produced due to the presence of a crack, it will change the spacing between gratings, which shifts the wavelength

Comparative vacuum monitoring (CVM) sensors are composed of alternating vacuum and atmospheric pressure galleries and detect cracks using pressure leakage between galleries. The testbed in Sandia National Laboratory showed that CVM could detect cracks in the size of 0.02 in [20]. Airbus [21] and Delta Airlines [22] also tested the feasibility of CVM on SHM. CVM sensors are lightweight made of polymer, and the gallery can be as small as 10 μm [23]. Even if CVM sensors

do not require undamaged (pristine) state, it can only detect damage

detect cracks that are relatively small and far away from the sensors.

underneath the sensor. Therefore, CVM is appropriate for hotspot monitoring. For fuselage damage monitoring, it would require a sensor layout with a very

There are other kinds of sensors, such as carbon nanotube sensors [24], printed sensors [25], and microelectromechanical systems sensors [26]. These sensors are, however, still in the research or development stage and take more time to be

As a summary, among different sensor technologies, it turned out that PWAS is the most appropriate for an SHM system for airplane fuselage monitoring as it can

of the reflected wave. FBG sensors detect damage by measuring the shift of reflected wavelength. It is small and lightweight. It was shown that a single optical fiber could incorporate up to 2000 FBG sensors [18]. The literature also showed that it could detect barely visible impact damage in a composite panel [19]. However, FBG sensors have a very short detection range because the local strain diminishes quickly as the distance increases. It would perform better for hotspot damage monitoring, where the damage location is already known. Since cracks in fuselage are opened during flight and closed on the ground, FBG is not appropriate for onground SHM. Lastly, since FBG measures the change in strains, it requires strains at the undamaged (pristine) state. If there is pre-existing damage, it can only measure

reflected wave.

dielectric film.

high density.

**30**

commercially available.

the change from the previous damage.

Repeated pressurization during takeoff and landing of an airplane can cause existing cracks on a fuselage skin to grow, for example, Aloha Airlines Flight 243. The rate of crack growth is controlled by, among other factors:


If left unattended, the cracks may grow to cause fatigue failure of the fuselage skin. In damage tolerance design, the less frequent the inspection, the lower the damage size threshold for repairing cracks in order to maintain a desired level of safety. The action of repairing cracks on fuselage skin to maintain a desired level of safety until the next scheduled maintenance is termed corrective maintenance. This section explains the modeling of the corrective maintenance procedure undertaken to prevent fatigue failure due to excessive crack growth.

The size of cracks in fuselage structures in a fleet of airplanes is modeled as a random variable characterized by a probability distribution that depends on manufacturing and the loading history of the airplane. The corrective maintenance procedure changes this distribution by repairing large-sized cracks as illustrated in **Figure 1**. **Figure 1** is presented as a probability density function (PDF) versus crack length. The solid curve represents the crack size distribution of an airplane entering the maintenance hangar. Different cracks grow at different rates because of random distribution of the Paris-Erdogan model parameters. The maintenance process is designed to repair fuselage skin with cracks larger than a repair threshold. Since crack detection is not perfect due to inspector's capability [27], maintenance only partially truncates the upper tail of the distribution, as represented by the dashed curve in **Figure 1**. It is noted that while there is uncertainty in damage detection, it is assumed that the size of the detected damage is known without any error/noise.

**Figure 1.** *The effect of inspection and repair process on crack size distribution.*

The shaded area represents the fraction of cracks missed during maintenance because of detection imperfection. The cracks that are missed during maintenance and happen to grow beyond the critical crack size before the next maintenance affects the safety of the aircraft.

at a different one. The analysis of the change in a guided wave's shape, phase, and amplitude yields indications about crack presence and extension. The probability of detection of the SHM method is comparable with that of conventional ultrasonic and eddy current methods [28]. Crack size and location can be displayed on ground equipment when connecting to onboard sensors and actuators after landing. Onground equipment can reduce the flying weight and thus may lower the lifecycle

*Advantages of Condition-Based Maintenance over Scheduled Maintenance Using Structural…*

The abovementioned process is called herein maintenance assessment. SHMbased maintenance assessment can be performed as frequently as every flight. However, as the crack increases by only a small amount in each flight cycle, it is unnecessary to perform this assessment after every flight. Also, maintenance assessment is not completely cost-free but requires a small amount of time and personnel. Typically, this assessment frequency ð Þ *Nshm* is assumed to coincide with the A check of scheduled maintenance, which is about 180 flight cycles (250 flight

**Figure 3** delineates the condition-based maintenance process. During the assessment, maintenance is requested if the crack size on a fuselage skin exceeds a specified threshold ð Þ *ath* . This threshold is performed, so as to repair all detected cracks on fuselage skins with threatening crack sizes. Additionally, the threshold for threatening crack size *arep*�*shm* is set substantially lower than the threshold for requesting maintenance ð Þ *ath* to prevent too-frequent maintenance trips for that

Condition-based maintenance is controlled by the following parameters:

• The thickness of fuselage skin ð Þ*t* , which affects the crack growth rate

• The thickness ð Þ*t* , along with the frequency of assessment ð Þ *Nshm* , and the threshold for requesting maintenance ð Þ *ath* affect the safety of the airplane

• The threshold for repair *arep*�*shm* determines the number of cracks needed to be repaired on fuselage skin. It is also set to prevent frequent maintenance trips

fuel cost.

airplane.

**Figure 3.**

**33**

*Flowchart of the condition-based maintenance.*

hours with average flight length of 1.4 h [4]).

*DOI: http://dx.doi.org/10.5772/intechopen.83614*

## **3.2 Scheduled maintenance**

The flowchart in **Figure 2** depicts the scheduled manual maintenance, in which maintenance is programmed at specific predetermined intervals (every *Nman* flight cycles) and corrective action is taken to ensure the airworthiness of the airplane until the next scheduled maintenance.

As all detected cracks on fuselage skins are repaired, the desired level of safety is determined by detection resolution/capability of GVI, *agvi*. It is expected that trained inspectors are able to detect cracks larger than 0.5 in (12.7 mm) in GVI. This is also the threshold for repair in scheduled maintenance.

Three parameters affect the lifecycle cost and safety of an aircraft undergoing scheduled maintenance: the maintenance interval, *Nman*; the threshold for repair (detection capability), *agvi*; and the thickness of the fuselage skin, *t*. To achieve a certain desired level of safety, *Nman* and *agvi* are correlated with each other. These three parameters together determine the number of maintenance trips and the number of cracks needed to be repaired on fuselage skins.

#### **3.3 Condition-based maintenance**

The condition-based maintenance process tracks crack growth continuously and requests maintenance when the crack threatens safety. In this chapter, the condition-based maintenance is considered to be performed using SHM technique. This technique employs onboard sensors and actuators, which are embedded in the structure, to monitor existing crack condition. In doing so, they detect cracks in metallic structures using guided waves transmitted from one location and received

**Figure 2.** *Flowchart of the scheduled maintenance.*

*Advantages of Condition-Based Maintenance over Scheduled Maintenance Using Structural… DOI: http://dx.doi.org/10.5772/intechopen.83614*

at a different one. The analysis of the change in a guided wave's shape, phase, and amplitude yields indications about crack presence and extension. The probability of detection of the SHM method is comparable with that of conventional ultrasonic and eddy current methods [28]. Crack size and location can be displayed on ground equipment when connecting to onboard sensors and actuators after landing. Onground equipment can reduce the flying weight and thus may lower the lifecycle fuel cost.

The abovementioned process is called herein maintenance assessment. SHMbased maintenance assessment can be performed as frequently as every flight. However, as the crack increases by only a small amount in each flight cycle, it is unnecessary to perform this assessment after every flight. Also, maintenance assessment is not completely cost-free but requires a small amount of time and personnel. Typically, this assessment frequency ð Þ *Nshm* is assumed to coincide with the A check of scheduled maintenance, which is about 180 flight cycles (250 flight hours with average flight length of 1.4 h [4]).

**Figure 3** delineates the condition-based maintenance process. During the assessment, maintenance is requested if the crack size on a fuselage skin exceeds a specified threshold ð Þ *ath* . This threshold is performed, so as to repair all detected cracks on fuselage skins with threatening crack sizes. Additionally, the threshold for threatening crack size *arep*�*shm* is set substantially lower than the threshold for requesting maintenance ð Þ *ath* to prevent too-frequent maintenance trips for that airplane.

Condition-based maintenance is controlled by the following parameters:


**Figure 3.** *Flowchart of the condition-based maintenance.*

The shaded area represents the fraction of cracks missed during maintenance because of detection imperfection. The cracks that are missed during maintenance and happen to grow beyond the critical crack size before the next maintenance

The flowchart in **Figure 2** depicts the scheduled manual maintenance, in which maintenance is programmed at specific predetermined intervals (every *Nman* flight cycles) and corrective action is taken to ensure the airworthiness of the airplane

As all detected cracks on fuselage skins are repaired, the desired level of safety is

Three parameters affect the lifecycle cost and safety of an aircraft undergoing scheduled maintenance: the maintenance interval, *Nman*; the threshold for repair (detection capability), *agvi*; and the thickness of the fuselage skin, *t*. To achieve a certain desired level of safety, *Nman* and *agvi* are correlated with each other. These three parameters together determine the number of maintenance trips and the

The condition-based maintenance process tracks crack growth continuously and

condition-based maintenance is considered to be performed using SHM technique. This technique employs onboard sensors and actuators, which are embedded in the structure, to monitor existing crack condition. In doing so, they detect cracks in metallic structures using guided waves transmitted from one location and received

requests maintenance when the crack threatens safety. In this chapter, the

determined by detection resolution/capability of GVI, *agvi*. It is expected that trained inspectors are able to detect cracks larger than 0.5 in (12.7 mm) in GVI. This

is also the threshold for repair in scheduled maintenance.

number of cracks needed to be repaired on fuselage skins.

affects the safety of the aircraft.

*Reliability and Maintenance - An Overview of Cases*

until the next scheduled maintenance.

**3.3 Condition-based maintenance**

**Figure 2.**

**32**

*Flowchart of the scheduled maintenance.*

**3.2 Scheduled maintenance**
