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Aerospace Aerodynamics Assignment Help

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7ENT11111 Aerospace Aerodynamics Assignment

Assignment Title - Aerodynamics Assignment

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Learning Outcomes -

Learning Outcome 1) Develop a knowledge and understanding of aerodynamic concepts and theories and their application in aerospace engineering.

Answer: Aerodynamics is the study of the interaction between air and objects moving through it. In aerospace engineering, understanding aerodynamic concepts is crucial for designing aircraft, spacecraft, and other flying vehicles. Key theories include lift and drag, which describe the forces acting on an object in flight. Bernoulli's principle explains how changes in air pressure can create lift, while drag is the resistance to motion caused by the object's interaction with the air. By applying these principles, engineers can design vehicles that are efficient, stable, and capable of achieving desired flight characteristics. Additionally, aerodynamic concepts are essential for understanding phenomena like turbulence and stall, which can significantly impact flight performance.

Learning Outcome 2) Apply aerodynamic theories in a range of aerospace applications and critically evaluate the predictions.

Answer: Once aerodynamic theories are understood, they can be applied to a variety of aerospace applications. This includes designing aircraft wings for optimal lift and drag, developing propulsion systems that minimize aerodynamic losses, and analyzing the stability and control characteristics of vehicles. To evaluate the accuracy of aerodynamic predictions, engineers often use computational fluid dynamics (CFD) simulations, wind tunnel testing, and flight testing. These methods allow for the comparison of theoretical predictions with experimental data, helping to refine models and ensure the safe and efficient operation of aerospace vehicles.

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Assignment Brief - This assignment will make use of fundamental potential flow theory, thin aerofoil theory and lifting line theory. Each student will have individual data.

Submission Requirements -

i. Present your results in a well-structured report.

ii. Present clearly the procedure, formulae used and working.

iii. It is acceptable to submit a scanned hand-written report, rather than type up your calculations - in fact this may be easier.

iv. The report should be anonymously submitted online via StudyNet by the deadline.

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Question 1 - Consider the two-dimensional, steady flow defines by a Cartesian axes system Oxy with velocities u along the x-axis and v along the y-axis.

For your Question 1 values of a and b Find the total flow speed and the angle by which the flow is inclined to the x-axis for the following flows:

(i) φ = ax + by

(ii) ψ = 2ax + 3by

The units of the potential function φ and streamline function ψ are m2/s.

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Question 2 - A thin cambered aerofoil has maximum camber b that occurs distance a from the leading edge (where both a and b are defined as are fractions of chord length, c, which will be taken as 1). The aerofoil is at an incidence α of zero degrees.

The camber line is given by the equation:

y(x) = 2(b/a)x - (b/a2)x2; 0 < x < a

y(x) = - ((2a-1)/(a-1)2)b + (2a/(a-1)2)bx - (1/(a-1)2)bx2; a < x < 1

(1) For your Question 2 values of a and b, write down the Equation for the camber line. IMPORTANT: Before moving on, check that y(a) = b and dy/dx(a) = 0.

(2) Evaluate the coefficients A0, A1 and A2 as defined in thin aerofoil theory for your values of a and b.

(3) Hence determine, for your values of a and b, the lift coefficient, CL, the pitching moment coefficient about the aerodynamic centre, Cm0, and the centre of pressure of the aerofoil.

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Question 3 - A tapered unswept wing has the planform (port side only).

The origin of axes Oxy is at the root chord as shown in Figure Q3. Hence the root is at y = 0 and the tip of the port wing is at y = -s.

(1) Prove that the chord at spanwise station y measured from the root is given by: C(y) = (Cr(y+s)-yCt)/s.

(2) Prove the aspect ratio for the wing is given by the equation: AR = 4s/(Cr+Ct).

(3) For this wing, Cr = 0.8m and the aerofoil lift curve slope a∞ = 2π. Using your values of incidence α, semispan s and tip chord Cr and assuming that two collocation points are used, calculate B1 and B3 in the 'monoplane equation' taking two values of φ1 as 45 and 90 degrees. The 'monoplane equation' is:

μαsinφ1 = n=1∑2(μ(2n-1) + sinφ1)B2n-1 sin (2n-1) φ1

where μ = a∞c/(8s) and y = - scosφ

(4) Hence calculate CL and the Oswald Efficiency factor for your wing.

(5) Suggest a quick way of checking whether your CL in Part (iv) is a reasonable value.

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