# 7ENT11111 Aerospace Aerodynamics Assignment Help

7ENT11111 Aerospace Aerodynamics Assignment Help
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**7ENT11111 Aerospace Aerodynamics Assignment - University of Hertfordshire, UK**

**Assignment Title - Aerodynamics Assignment**

**Learning Outcomes -**

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

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

**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 m^{2}/s.

**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/a^{2})x^{2}; 0 < x < a

y(x) = - ((2a-1)/(a-1)^{2})b + (2a/(a-1)^{2})bx - (1/(a-1)^{2})bx^{2}; 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 A_{0}, A_{1} and A_{2} 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, C_{L}, 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) = (C_{r}(y+s)-yC_{t})/s.

(2) Prove the aspect ratio for the wing is given by the equation: AR = 4s/(C_{r}+C_{t}).

(3) For this wing, C_{r} = 0.8m and the aerofoil lift curve slope a_{∞} = 2π. Using your values of incidence α, semispan s and tip chord C_{r} and assuming that two collocation points are used, calculate B_{1} and B_{3} in the 'monoplane equation' taking two values of φ_{1} as 45 and 90 degrees. The 'monoplane equation' is:

μαsinφ_{1} = _{n=1}∑^{2}(μ(2n-1) + sinφ_{1})B_{2n-1} sin (2n-1) φ_{1}

where μ = a_{∞}c/(8s) and y = - scosφ

(4) Hence calculate C_{L} and the Oswald Efficiency factor for your wing.

(5) Suggest a quick way of checking whether your C_{L} in Part (iv) is a reasonable value.

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