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Homework #1: Due January 30 From Textbook: Chapter 3, Problems 3.1, and 3.3
Additional Problems: 1) Go to the web site of a general aviation aircraft manufacturer (Cessna, Piper, Mooney, Beach/Raytheon, Cirrus, etc) and collect geometry and performance information on a single engine airplane of your choice. Use this data to generate Thrust and Power versus Velocity plots and use these plots to determine: (a) max. speed as sea level (b) max. speed at 10,000 ft (c) equivalent airspeeds for (L/D) max and ( CL3/2/CD)max.
In addition to the web site data, you will need to make assumptions about the span efficiency factor (between 0.75 and 0.85), the propeller efficiency (use 0.80 for a fixed pitch prop, 0.85 for constant speed), and the zero lift drag coefficient. A good correlation for CDo is based upon the take off weight: CDo = 10^(a + 1.089 +0.515*log10(WTO))/S The factor, a , depends upon how aerodynamically "clean" the airplane is. Use a = -2.046 for fixed landing gear planes a = -2.222 for retractable landing gear with aluminum fuselage, or a = -2.301 for retractable landing gear with composite fuselage. Homework #2: Due February 6 From Textbook: Chapter 3, Problems 3.4, 3.5, and 3.7. (Neglect ground effect for these aircraft since the given CL's are so low.)
Additional Problems: 1) For a Cessna 172 (using data from the Cessna website), determine the level runway take-off and landing ground distances both sea level and 10,000 ft. Use the average take-off thrust as the thrust available at Vavg. For landing, thrust is zero. Use the take-off and landing CL's at the velocities corresponding to 1.20 and 1.15 times the stall speed for take-off and landing, respectively. Also, use the following data: Take-off Landing mr 0.02 0.30 CLmax 2.00 2.80 CL CLopt 1.50 CDo 0.0350 0.0650 e 0.75 0.65 h/b 0.20 0.20 Homework #3: Due February 13 From Textbook: Chapter 3, Problems 3.12, 3.14, 3.15 and 3.18. Homework #4 : Due March 6th
From Book: Chapter 7, Problems 5 and 9 Chapter 8, Problems 4, 7, 10, 11 and 15 Other
Problems: 1) A transonic airfoil has a normal shock on the upper surface at a flight Mach number of 0.95 as shown below. The freestream pressure and temperature are 10 psia and 70oF, respectively. The upper surface shock occurs at a local Mach number of 1.3. If the trailing edge pressure (upper and lower surface) is 12 psia, find the Mach numbers and velocities at the trailing edge. Note that these values will be different between upper and lower surface since one path is isentropic and the other is not.
Homework #5 : Due March 15th
From Book: Chapter 9: Problems 2, 4, 8, 10, 11 Other
Problems: 1) Consider the 2-D ramp inlet as shown below. What are the Mach number and pressure behind the first shock? Behind the first reflection?
Homework #6 : Due March 31st
From Book: Chapter 9: Problems 12 and 13. Use shock-expansion theory
Other problems 1) Repeat problems 13 using Thin Airfoil Theory and comment on the differences between the two methods.
2) Calculate the Thin Airfoil Theory lift and drag coefficients for an airfoil shaped like that shown below at Mach = 1.6 and a t/c = 5% at an angle of attack of 3 degrees.
Homework #7 : Due April 24th
From Book: Chapter 10: Problems 1, 2, 4, 9, and 12
Chapter 19: Problems 1 and 2.
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Contact: gallyt@erau.edu
