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AERO 309 - Module 7 Homework In this assignment, you may work in either the SI or BGS system, but you must be consistent.
1. Performance curves a. List the aircraft you have been using for the previous assignments. List your
cruise altitude and the corresponding air density value at that altitude. List the wing area (if you have a multi-wing aircraft, only provide the information for one wing). Provide your cruise speed. Provide your takeoff weight. Provide your aspect ratio. b. In the Module 5 Homework, you created a drag polar for a wing (question 9).
What is the minimum CD for your wing? Multiply this minimum CD by 4. We will use this value as the CD0 for this assignment. What is your CD0? (If your question 9 in Module 5 was incorrect, make sure to correct it before finishing this problem and
homework) c. Assuming an Oswald efficiency factor of 0.8, you can now write the equation for
the drag polar of the full aircraft
CD=CD0+ [CL2] / [πeAR] What is the drag polar equation for your aircraft? (CL is not known and is part of the equation; you are just substituting in values [CD0, e, AR] that you know.)
d. For this question, assume CL=0, so CD=CD0. At standard sea level conditions and at cruise speed, what is the drag using CD=CD0? (Your answer is a force: N or lb)
e. We will assume that a turbojet engine is used, so the thrust available can be
considered constant with respect to the velocity. For this problem, let TA at sea level equal 2.5 times the drag you calculated in part d. What is your TA at sea
level?
f. Assume that the CLmax is 1.8, calculate the stall speed of your aircraft at takeoff weight and at sea level.
g. Use Excel to plot the TA and TR curves (together on one plot) for your aircraft at sea level and at takeoff weight. Plot your TR from stall speed to just past Vmax. Use a small enough interval with V so the curve is smooth. Remember that weight
(lift) is constant. The CL will change with airspeed. In Excel, have at least the following columns: V CL CD Required T Available T
h. From your thrust curves, what is Vmin and Vmax? How did you arrive at these values?
i. Use Excel to calculate and plot the PA vs V and PR vs V curves (together on one plot) for your aircraft at sea level and at takeoff weight.
j. Repeat f-i but at standard sea level and 90% takeoff weight.
k. Repeat f-i but at cruise altitude and 90% takeoff weight. Assume that TA at altitude is directly proportional to air density. [TS is the thrust available at sea level
that you calculated in part e]
TA=TS [ρ] / [ρS]