Which part of an aircraft can most benefit from a structural health monitoring (SHM) system?

Question

In order to reduce operational cost and improve the safety and reliability of aircraft's the aerospace industry and government is increasingly evaluating new ideas such structural heath monitoring systems for aircraft's. Such a system will enable continuous monitoring, inspection and detection of damage to the aircraft with minimal human involvement. This will also help the airline industry to improve aircraft maintenance scheduling process.

_{Figure: Early Diagnosis for Microcracks in Aircraft}

Currently Delta airlines and an aircraft manufacture has teamed up with Sandia Laboratory to install sensors in aircraft to help understand flaw detection. Research also indicates that ultrasonic sensors are most widely used sensing element.

Question:
Which part of an aircraft can most benefit from structural health monitoring (SHM) system? Is it the fuselage, wings, landing gear, flaps, rudder, elevator, vertical stabilizer or another component? Why is this component the prime candidate?

References:

• In-flight sensor tests a step toward Structural Health Monitoring for safer flights
• Structural health monitoring: Composites get smart
• Early Diagnosis for Microcracks in Aircraft
• Air Asia QZ8501: Pilot response to equipment malfunction caused crash

Answer 1

• First you need a part that exceeds its fatigue limit under expected stresses because you need slow coalescence of cracks occurring uniformly in the measured area, not a local defect which might occur outside the sensor placement and resistant to modelling based on measurements from other pieces of the part. Steel/titanium etc parts stressed below their fatigue limit won't tell you anything.
• Parts that are easy to access can't lower maintenance costs as much as difficult to reach structures.
• Since you're trying to lower the cost without compromising safety of the procedures, the logical step is to put a monitor on the parts expected to show signs of fatigue first and for which fatigue is well understood. You want agreement with the model, proving that you can design the sensor, not that the sensor happened to find or not find crack growth.

Based on these considerations, problems I would expect are of interest are:

1. The fuselage. Compression cycles are a different fatigue mechanism than time spent in the air, requiring modelling and checks.
2. Landing gear attachment point. Landing hardness and its effects on the fatigue of connecting structures can potentially not match airframe life or duty cycles, requiring checks.
3. Wings. Wings encounter a large number of cycles and require access panels to measure internal structures. Fatigue in the middle of a large beam is probably relatively predictable.
4. Engine pylon attachment points. Compressor stalls, vibration in the engine, and awkward landings/maneuvers can all result in uncommon forces that would be difficult to treat based on empirical data, requiring checks
5. The tail surfaces that are far off the ground are a good place to perhaps save time on lifts.

Wings and fuselage are going to be best to demonstrate the sensor engineering principles. Pylons and landing gears (excepting any steel parts) would present the most insight perhaps.