Lecture 15 Flashcards
- What made the Comet 1 aircraft a very advanced concept for its time and why is this a relevant factor to why five of the 21 built, failed structurally
- Powered by 4 de Haviland ghost turbo-jet engines with a cursing altitude of 35000ft. intended for long-distance flights with high speeds with a maximum weight of 49 tonnes. This greater altidue and faster egninge reduce flight times.
- It was a sleek designed with swept-back wings and streamlined fuselage improved fuels
- It was made from a aluminium alloy that made the aircraft lighter.
- This was an advanced aircraft that was first of its time.
- The cabian was pressured at 0.57 bar which was 50% higher than other aircrafts at the time. This pressure allowed the aircraft to fly at many thousand feet higher.
- During each flight the fuselage functions as a pressure vessel and is subjected to both an axial and circumferential tensile (cyclic) stress.
- Using reverse engineering where did the forensic engineering investigators locate the start of the ‘fatal’ fatigue crack in the airframe of Comet G-ALYP
Failure was traced to a crack which had started at the corner of a window in the cabin roof. The windows housed the aerials for the Automatic Direction Finding system (ADF).
- Figure 62 is a close-up photograph showing the origin of the fatigue crack in Comet G-ALYP. Once the fatigue crack grew to its critical length of 46 mm there was ultimate failure due to fast brittle fracture. What would the investigators have observed on the crack surfaces local to the nucleation site to show that this was indeed where failure started
The sketch and photograph of the reconstruction for the rear ADF window show where the first crack probably originated. Fatigue markings were found on the fracture surface at this location: they had probably started at the edge of a countersunk hole which had been drilled though the aluminium skin to take a fastener.
When investigators examine the crack surfaces of a fatigue failure, particularly in the context of the Comet G-ALYP, they look for several specific features on the crack surfaces to confirm the origin of the failure:
Beach Marks or Striations: These are marks that indicate the progression of the crack over time. They appear as concentric rings or lines on the crack surface, showing the increment by which the crack advanced with each pressurization cycle.
Smooth, Shiny Area: This area, often called the “origin area” or “nucleation site,” is typically smooth and shiny. It is where the crack initially began and is distinct from the rougher texture seen in other areas of the fracture.
Ratchet Marks: These are step-like marks on the crack surface that indicate multiple origins or branches of the crack merging as it grows.
Fatigue Progression: The crack will show a progression from the origin point, with distinct regions of initial slow crack growth followed by faster crack growth as the crack lengthens and stress concentration increases.
Fast Fracture Zone: Beyond the smooth area of slow crack growth, investigators will observe a transition to a rougher surface where the final fast, brittle fracture occurred. This indicates the point at which the critical crack length was reached and the remaining material failed rapidly.
Investigators would have used the engineering tool box with tests on material specimen and look into design calculations.
- Explain what information the investigating team could use to identify that the cause of ultimate failure was by low cycle fatigue
Low cycle fatigue is when failure due to fatigue occurs within 10000 cycles of fatigue loading.
- Due to the pressure inside the aircraft each flight the fuselage functions as a pressure vessel and is subjected to both an axial and circumferential tensile (cyclic) stress. Comet G-ALYP had flown 1290 flights; G-ALYY a lower number at 900 flights. This comparatively small number of cycles had generated a fatigue crack which had grown to be long enough to cause fast (brittle) fracture.
- In 1954 fatigue of aircraft structures was growing in awareness in the UK. de Havillands in 1953 had tested a section, cycling it between zero and p. By 18000 cycles it failed from a fatigue crack; this gave a design life of 10000 flights. After 1230 flights Comet G-ALYU was fatigue tested. Water was used to pressurise the cabin between zero and p. It failed after 1830 cycles, and a total of 3060 “flights”.
- In 1953 what form of proof test was used by the de Havilland engineers, and after how many cycles did the Comet 1 aircraft section fail? What were the two fallacies in the design figures for their 10000 flight fatigue life that proved to be fatally wrong
- The test used in 1953 was testing a section of the aircraft between 0 and 0.57 bar. Comet 1 failed after 3060 flights which was 1830 cycles.
- Next to the window of the plane strain gauges where placed . At pressure p the maximum stress measured was 297 MPa. This was 50% more than from the design value of 195 MPa, and only 3% below minimum sy .
- When it was tested to give a design life of 10000 flights it was presureised to 2P when the actcual aircraft was pressurised to 1.33p.
- When the section was at 2p the bolt holes would have yielded in tension. On depressurising to zero pressure the elastic springback would have put the Al alloy around the bolt holes in compression. This plastic shakedown would have smoothed out the stresses in the test section and improved the fatigue resistance.
- What is a stress concentration (SC) factor and what determines its magnitude? We are told that the de Havillands engineers calculated the average stress value (difficult then!) in the skin near the origins of the crack to be about 195 MPa. In their design calculations do you think the de Havilland engineers took account of the SC of 2 for a bolt hole where failure initiated? As part of the investigation Comet G-ALYU was load tested to the cabin pressure of 0.57 bar and, using strain gauges, the maximum stress measured was 297 Mpa. Given this information show why we now know that some volume of the Al alloy next to the bolthole will have yielded
- SC is the elastic stress-concertration factor and is twice the calculated value of stress.
- A stress concentration (SC) factor, also known as the stress concentration factor (Kt), is a numerical value that quantifies the increase in stress at a geometric discontinuity or irregularity in a material, such as a notch, hole, or sudden change in cross-section.
- The SC factor is a ratio of the maximum stress at the discontinuity to the nominal stress in the rest of the material. It indicates how much higher the stress is at the point of discontinuity compared to the surrounding material.
- Given the investigation’s findings and the structural failures that occurred, it is possible that they either underestimated the SC factor or failed to account for it adequately in their design calculations.
- To show why some volume of the aluminum alloy next to the bolt hole would have yielded, we can compare the calculated stress values with the yield strength of the aluminum alloy:
- Calculated Stress Values: The maximum stress measured using strain gauges during the load test was 297 MPa.
- Yield Strength of Aluminum Alloy: The typical yield strength of aluminum alloy (such as 2024-T3) used in aircraft construction ranges around 345 MPa
- Given that the maximum stress measured (297 MPa) is close to but still lower than the yield strength of the aluminum alloy (around 345 MPa), the stress near the bolt hole might have been sufficient to cause localized yielding in the material. This might have been higher due to SC factor.
- The elastic Stress-Concentration (SC) factor for a bolt hole can be estimated to be be a least 2, giving an edge stress of 594 MPa. This elastic stress in the Comet was relieved by plasticity and truncated at sy . Even with the design value the elastic stress next to the hole is 390 MPa; 40 MPa more than the actual sy for the Al alloy
