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Upon opening XFLR5, first go to the File menu and click new project (not necessary when firstopened, but good practice anyways). Then, in the Application menu, click Foil Direct Design. Thiswill open up a window with an airfoil shown (for designing your own airfoil). Hide this (removethe check on Show Foil in the lower right), open the Foil menu, and click Naca Foils (at thebottom). The code will ask for the number of the airfoil (0012) and the number of panels (use thedefault value 100). The NACA 0012 will appear on the screen and the listing on the bottom of the

page. Click on the listing to highlight it, indicating that you will be working on this airfoil.

Next, on the application menu, click Xfoil direct analysis. This will open a window which showsthe pressure profile ( c p v. x / c) plot (but no data yet) and the airfoil shape. To get some data, clickon the Polars menu and then Define Batch Polar. This window allows us to enter key datadepending on the Type of the analysis. We will just use Type 1, which means providing aReynolds and Mach number. Use Re = 1,000,000 (default is 100,000) and M = 0.00 (whichessentially means low Mach number). You can give this analysis a user-defined name or just usethe somewhat cryptic default. Dont worry about the transition model stuff, leave as default

values). Have supplied these values, we can now analyze the flow, either by using the batchanalysis under the Polars menu or the side menu on the right under Analysis. I will do the later, byspecifying (angle of attack) = 0 (no sequence), and hitting analyze. The result is now a pressureprofile is shown for this case. Now, redo this analysis, but set the angle of attack (in start) to 4.00(4 degrees). You will now see a second pressure profile which has two lines, one for the topsurface and the other for the bottom. The c p axis is inverted, with negative numbers at the top,because in this way the top line will typically correspond to the upper surface of the airfoil (wherewe want lower pressure). Actually, there is an upper and lower surface line for the 0 degree case aswell, but since the airfoil is symmetric there is no difference between the top and bottom profiles,so they overlap. Note also that the airfoil sketch is now tilted to show the angle of attack.

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To see other plots, go to Polars, then View, then choose. However, we need more points to makethis interesting, so first go back to the right-hand Analysis section and do a sequence starting at =8 and continuing to = 16 with a step ( ) of 4 degrees. This will yield three more pressureprofiles. Now, go to Polars, View, and then choose Cl vs. alpha. This will show the change in liftwith angle of attack. From angles 0-8, your lift line should be fairly linear. Above this, the liftcurve begins to flatten. The peak of the lift line corresponds to stall, after which increasing theangle of attack will reduce the lift. Next, choose Cl v. Cd this is the lift-drag polar. Finally, clickon user graph this will yield Cl/Cd v. alpha, with a peak at 8 degrees. If you right click on thisplot, you get another menu which allows you to control the plot display. Here, go to Graph and inthe submenu go to variables. This presents a list of variables with all the plot choices available.Construct a plot Cd v. alpha.

These are the tools you need there is plenty more (some of which will be available in the separateguide for XFLR5), but this should be sufficient to get you started. Dont be afraid to experiment

you can always reset by starting a new project.

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Exercise A: Effect of angle of attack

For this exercise, start with the NACA 0012 airfoil. Using Re = 1,000,000 and M = 0.00, analyze theairfoil for = -4 to 16 degrees with a step of 4 degrees. Determine the value of cl and cd at eachangle, and submit a plot with all six pressure profiles. Briefly discuss the trends seen what happensto the pressure profiles, lift, and drag as the angle of attack increases. Does this seem consistent witha Bernoulli approach to analysis?

Now, refine your plot, running from -4 to 18 degrees, with a step of 0.5 degrees. Based on thisanalysis, what is the maximum cl for this airfoil and what angle corresponds to the start of stall?What is the maximum cl / cd and at this point what is the value of cl, cd , and ? What anglecorresponds to zero lift? Provide plots of cl v. , cd v. , cl / cd v. , and the lift-drag polar using thisdata.

Exercise B: Effect of Reynolds number

Start with the NACA 0012 at = 4 degrees and M = 0.00. Create pressure profiles for Re = 100,000to 2.1 million with an increment of 400,000. In addition, do the case with Re = 4 million. Provide aplot with all these profiles. What key changes do you observe in the pressure profile as the Reynoldsnumbers change?

Next, solve for = -4 to 20 with a step of 0.5 degrees for Re = 100,000, 500,000, 900,000,2,100,000, and 4,000,000. For each Reynolds number, what is the maximum cl, what anglecorresponds to the start of stall, what is the maximum cl / cd and at this point what is the value of cl,cd , and ? Provide plots of cl v. , cd v. , cl / cd v. , and the lift-drag polar using this data, with thecurve for all five Reynolds numbers on each plot. Briefly describe the trends, noting the behavior atlow angles of attack and at high angles of attack.

Exercise C: Effect of Mach number

Start with the NACA 0012 at = 4 degrees and Re = 2 million. Create pressure profiles for M = 0.0,0.1, 0.2, and 0.3. Provide a plot with all these profiles. What key changes do you observe in thepressure profile as the Mach numbers change?

Next, solve for = -4 to 20 with a step of 0.5 degrees for these four Mach numbers. For each Machnumber, what is the maximum cl, what angle corresponds to the start of stall, what is the maximumcl / cd and at this point what is the value of cl, cd , and ? Provide plots of cl v. , cd v. , cl / cd v. , andthe lift-drag polar using this data, with the curve for all four Mach numbers on each plot. Brieflydescribe the trends, noting the behavior at low angles of attack and at high angles of attack.

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Exercise D: Effect of Thickness

For this exercise, assume Re = 2,000,000 and M = 0.00. Do an analysis running from -4 to 20degrees, with a step of 0.5 degrees, for the NACA 0004, 0008, 0012, 0016, and 0020 airfoils. Basedon this analysis, what is the maximum cl for these airfoils and what angle corresponds to the start ofstall? What is the maximum cl / cd and at this point what is the value of cl, cd , and ? Provide plots ofcl v. , cd v. , cl / cd v. , and the lift-drag polar using this data, with all five airfoil curves on each

plot. Briefly describe the trends you observe.

Further, reexamine the convergence/non-convergence and the shape of the two thinnest airfoils. Upto what angle of attack do you believe the observed curves? Show these limits on your plots.

Exercise E: Effect of camber, magnitude

For this exercise, assume Re = 2,000,000 and M = 0.00. Do an analysis running from -4 to 20degrees, with a step of 0.5 degrees, for the NACA 0012, 0312, 1312, 2312, and 3312 airfoils. Presenta plot of all six airfoil profiles. Based on this analysis, what is the maximum cl for these airfoils andwhat angle corresponds to the start of stall? What is the maximum cl / cd and at this point what is thevalue of cl, cd , and ? What angle corresponds to zero lift? Provide plots of cl v. , cd v. , cl / cd v. ,and the lift-drag polar using this data, with all six airfoil curves on each plot. Briefly describe thetrends you observe what are the key differences between symmetric airfoils and asymmetricairfoils, and what does adding camber do to the aerodynamics?

Exercise F: Effect of camber, location

For this exercise, assume Re = 2,000,000 and M = 0.00. Do an analysis running from -4 to 20degrees, with a step of 0.5 degrees, for the NACA 0012, 2112, 2312, 2512, and 2712 airfoils. Presenta plot of all five airfoil profiles. Based on this analysis, what is the maximum cl for these airfoils and

what angle corresponds to the start of stall? What is the maximum cl / cd and at this point what is thevalue of cl, cd , and ? What angle corresponds to zero lift? Provide plots of cl v. , cd v. , cl / cd v. ,and the lift-drag polar using this data, with all five airfoil curves on each plot. Briefly describe thetrends you observe what are the key differences between symmetric airfoils and asymmetricairfoils, and what does changing the camber location do to the aerodynamics?

Exercise G: Finding the best values for a NACA airfoil

For this, your goal is to locate the NACA airfoil that achieves the stated goals below. You may allowthe camber to range from 0%-9%c, the camber location to be from 10%-70%c, and the thickness torange from 3%-30%c. Using your knowledge of the trends from the previous exercises find:

a) the airfoil with the highest cl,maxb) the airfoil with the highest cl at = 0c) the airfoil with the largest stall angled) the airfoil with the best cl / cd ratio at = 4 degreese) the airfoil with the best cl / cd ratio at = 4 degrees and Re = 200,000

For all these (except as noted), assume Re = 2 million and M = 0. Watch for inaccurate results or alack of convergence. Note that the results may still not be the best design as there are many otherconsiderations we are not covering.

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