Version 2 Page 1 of 9 EG3121/EG4121/EG7038 All Candidates Semester 2 Examinations 2019 DO NOT OPEN THE QUESTION PAPER UNTIL INSTRUCTED TO DO SO BY THE CHIEF INVIGILATOR Department Engineering Module Code EG3121/EG4121/EG7038 Module Title Aerospace Materials Exam Duration (in words) Two Hours CHECK YOU HAVE THE CORRECT QUESTION PAPER Number of Pages 9 Number of Questions 5 Instructions to Candidates Answers are expected to three questions. Answers to only three questions will be marked. Attempted solutions which you do not wish to submit should be crossed out. If you do attempt more than three questions, and do not identify which three you want to be marked, only the first three in the answer book will be marked. For each question, the distribution of marks out of 20 is indicated in brackets. FOR THIS EXAM YOU ARE ALLOWED TO USE THE FOLLOWING: Calculators Permitted calculators are the Casio FX83 and FX85 models Books/Statutes provided by the University Yes Engineering Data Book Are students permitted to bring their own Books/Statues/Notes? No Additional Stationery No Version 2 Page 2 of 9 EG3121/EG4121/EG7038 All Candidates 1.) Question One [20 marks total] (a) Airbus uses aluminium alloy 2024-T351 for construction of the lower wing surface of the A320 airliner. The main alloying elements used in Al 2024 are given in Table Q1. In the –T351 temper the material is solution treated at 495°C, water quenched, cold worked then naturally aged at room temperature for 4 days. Table Q1. Alloying elements in Al 2024 Alloying element Composition range (wt%) Cu 3.8 – 4.9 Mg 1.2 – 1.8 Mn 0.3 – 0.9 Si Max 0.5 Fe Max 0.5 (i) State how Cu, Mg and Mn can strengthen aluminium, both in 2000 series alloys such as 2024 and in any other relevant aluminium alloy series. [3 marks] (ii) State the difference between the –T351 temper described above and the –T4, –T6 and –T7 tempers. [3 marks] (ii) Explain how the properties of a 2024-T351 material are likely to compare with the same alloy which has been heat treated to the –T4, – T6 and –T7 conditions. In your answer you should clearly link the changes in properties with differences in the microstructure and processing with reference to the age hardening process. [6 marks] (b) Critique the choice of this alloy for a lower wing surface and predict what materials might be used for this component in similar aircraft in the future. In your answer you should explain the main design drivers for the lower wing skins of large aircraft, describing what material properties are important and explaining the reasons why. [8 marks] Version 2 Page 3 of 9 EG3121/EG4121/EG7038 All Candidates 2.) Question Two [20 marks total] (a) For each titanium alloy composition below, state which class of alloys the composition belongs to (α alloys, near- α alloys, α/β alloys or β alloys) and identify the critical alloying elements and their role. (i) Ti-3Al-8V-6Cr-4Mo-4Zr [2 marks] (ii) Ti-8Al-1Mo-1V [2 marks] (iii) Commercially pure (ASTM grades 1-4) [1 mark] (b) For the following aerospace components, choose a class of titanium alloys which would have appropriate properties to meet the likely material properties. Explain your selections clearly. Where possible, propose specific alloys and appropriate processing, manufacturing or heat treatments. (i) The fan blade in a large turbofan engine. [3 marks] (ii) The high-pressure compressor discs of a large turbofan engine [2 marks] (iii) Small titanium fasteners to join other structural components. [2 marks] (c) The Boeing 777 main undercarriage legs and truckbeams are manufactured from a beta titanium alloy which is heat treated for improved strength. A sketch of the undercarriage components and a simplified phase diagram illustrating the effect of heat treatments are shown in Figure Q2 overleaf. Assess the choice of material and alloy for this component, explaining the solution treatment and aging process for α/β and β alloys with reference to the phase diagram. In your answer you should include a sketch of the microstructure and highlight any particular challenges with heat treating this component. [8 marks] Question continues overleaf. Version 2 Page 4 of 9 EG3121/EG4121/EG7038 All Candidates Figure Q2 – Boeing 777 undercarriage components, simplified β-isomorphous phase diagram and strength variation for titanium. Ref: Boyer RR 1996. Materials Science and Engineering A 213 103-114. Version 2 Page 5 of 9 EG3121/EG4121/EG7038 All Candidates 3.) Question Three [20 marks total] (a) A gas turbine engine was originally fitted with single-crystal high-pressure turbine blades manufactured from RR3000. The creep performance of this material is summarised in Figure Q3 overleaf. With the blades operating at a temperature of 1000°C, the available life was 500 hours. The engine has been donated to a museum, and they wish to run it for 20 minutes three times a year at special open days. Unfortunately the turbine blades require replacement, and the museum can only afford to replace them with blades machined from the Ni- superalloy Waspaloy. The business case requires the new blades to last 10 years. (i) Assuming that the stress on the new blades would be the same as those on the existing blades, determine whether a straight replacement of the blades with those manufactured from Waspaloy is feasible. If required, propose a blade temperature limit to achieve the required life. [7 marks] (ii) One of the museum volunteers looked up the static strength of Waspaloy and found it was around the same, or higher, than the single-crystal material. They were also intrigued that the strength of both seemed to be maintained or improve slightly at higher temperatures. Explain why this is with reference to the microstructure. [7 marks] (c) Explain how the configuration of a Bridgeman furnace used to manufacture single crystal turbine blades is used to overcome constitutional undercooling. You may find a simple diagram useful in your explanation. [6 marks] Question continues overleaf. Version 2 Page 6 of 9 EG3121/EG4121/EG7038 All Candidates Figure Q3 – Creep performance of aerospace alloys Please note that you are not required to submit copies of this figure with your answer. Version 2 Page 7 of 9 EG3121/EG4121/EG7038 All Candidates 4.) Question Four [20 marks total] A composite sandwich panel is to be used to support a number of avionics boxes in an aircraft. The applied loading and dimensions have been determined and are shown in Figure Q4 overleaf. The skins are to be manufactured from a laminate of unidirectional carbon-fibre reinforced polymer (CFRP) plies. Aluminium honeycomb and Rohacell closed cell foam have been proposed as materials for the core. Material data are given in Table Q4. (a) Show that the properties of the skin laminates can be estimated as Ex = 56.9 GPa, σ*xt = 878 MPa and σ*xc = 455 MPa using an appropriate method. Show all your working and assumptions clearly. [4 marks] (b) Assuming an Ultimate Factor of Safety of 1.5 in accordance with typical aerospace industry practice: (i) Calculate reserve factors for all applicable failure modes of the panel when manufactured with an Al 5052 Honeycomb 83 core. [8 marks] (i) For the panel manufactured using a Rohacell 110 IG-F core, state any failure modes that are not applicable and calculate reserve factors for any failure modes that differ from the panel with a honeycomb core. [3 marks] (c) Critique the two core options, selecting a most appropriate design choice in this case, but also outlining the advantages and disadvantages of each type of core (honeycomb and closed cell foam). [5 marks] Question continues overleaf. Version 2 Page 8 of 9 EG3121/EG4121/EG7038 All Candidates Table Q4 – Material properties Skin – Carbon Fibre Reinforced Epoxy Laminate Lay-up [0/45/-45/90]s Ply (lamina) thickness 0.125 mm Lamina fibre direction (0°) elastic modulus 175 GPa Lamina fibre direction (0°) tensile strength 2700 MPa Lamina fibre direction (0°) compressive strength 1400 MPa Core Option 1 - Al5052 honeycomb 83 Elastic modulus 1310 MPa Shear modulus 245 MPa Shear strength 1.8 MPa Cell size 6 mm Density 83 kg/m3 Core Option 2 – Rohacell 110 IG-F Elastic modulus 120 MPa Shear modulus 50 MPa Shear strength 2.4 MPa Density 110 kg/m3 Figure Q4 – Sandwich beam and design equations. Shear force and bending moment are given as limit loads, symbols have their usual meanings. Version 2 Page 9 of 9 EG3121/EG4121/EG7038 All Candidates 5.) Question Five [20 marks total] (a) A composite component for an aerospace application is manufactured using a resin transfer moulding processes. The laminate uses 10 plies of woven fabric with an aerial weight of 650 g/m2. The fibre is carbon with a fibre density of 1900 kg/m3. After production, the panel thickness is measured as 6.45 mm. (i) Estimate the fibre volume fraction, stating any assumptions you make, and comment on your answer. [4 marks] (ii) Describe the techniques that could be used to confirm the estimated fibre volume fraction and confirm the component quality. [4 marks] (b) Pre-preg/autoclave and Resin Transfer Moulding (RTM) are two major composite manufacturing processes. General Electric introduced CFRP fan blades on the GE90 engine used on the Boeing 777 large airliner in the mid-1990s. These blades were manufactured using pre-preg/autoclave. They have now introduced a CFRP fan on the LEAP engine for the Boeing 737 MAX. This is the latest version of the best-selling airliner in the world. These new blades are manufactured using an RTM process and include a 3D woven fibre form. Explain the difference between these manufacturing processes and analyse the reasons for this change in production method and material form. [12 marks]
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