程序代写案例-EG4121/EG7038

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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
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

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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, –
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]

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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.

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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.

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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
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.

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Figure Q3 – Creep performance of aerospace alloys

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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
[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.

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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.

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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
[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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