Wright State University Fall 2004

Department of Mechanical and Materials Engineering

ME 317: FLUID DYNAMICS LABORATORY

Experiment 6: Introduction to Airfoils

Objective: The objective of this lab is to analyze the basic flow physics of airfoils using XFOIL, which is an interactive program for the design and analysis of subsonic isolated airfoils. The program is to be used to analyze and then compare the performance of two Selig low-Reynolds number airfoils.

Motivation: A senior design project performed by students in the Department of Mechanical and Materials Engineering involved building a remote-controlled airplane, which was flown at the SAE AeroDesignEast Competition in Daytona Beach, Florida, in 2004. Dr. Thomas has entered the competition again this year, and analyzing two candidate wing sections would provide useful information that would improve the performance of the proposed airplane. For more information on last year’s competition, follow this link:

Reference: Fox, McDonald and Pritchard, Introduction to Fluid Mechanics, 6th edn., Section 9-7: Drag, and Section 9-8: Lift.

Preparation: Download the XFOIL program (version 6.9), and the s1210 and s1223 airfoil profiles to be analyzed from the author’s website. The appendix of this lab contains a customized basic operation manual for XFOIL. A complete documentation manual is also downloadable from the webpage.

Software Needed: XFOIL, Selig wing profiles, EXCEL

Required Analysis: For the S1210 and S1223 Selig airfoils calculate, using the XFOIL program, the coefficient of lift CL, the coefficient of drag CD, and pressure coefficient CP for angles of attack (AOA) ranging from α = –3º to 25º in 2º increments. The Reynolds number is Re = 310,000 with a Mach number of Ma = 0.1. Turn in the plots of the labeled airfoil, CL versus α, CD versus CL, and CP versus x/c, and complete answers to the questions.

  1. Plot the shape of each airfoil and label the chord, pressure and suction surfaces, mean camber line, trailing and leading edge, and stagnation point.
  2. Plot CL (on the y-axis) versus α (on the x-axis). Explain what is meant by airfoil stall using this plot.
  3. Plot coefficient of drag (CD) versus the coefficient of lift (CL). What can be inferred from this plot?
  4. Plot pressure coefficient CP versus x/c (the normalized chord) for a 0º angle of attack and the also the AOA at which the airfoils stalls, which would be obtained from the CL versus α plot. This plot is obtainable from the XFOIL program.
  5. From the CP versus x/c plot, what is the stagnation point on the leading edge for both airfoils?

XFOIL USER BASIC MANUAL

DOWNLOADS

1Download the XFOIL program (version 6.94) from only the software optimized for your system (Pentium 4 for CECS computers)

2NACA class airfoils are contained in the databank of Xfoil, but the SELIG airfoil profiles we would be analyzing are not in the databank but are downloadable from the same webpage. Under profile, click to open Selig, find and extract the required profiles (s1210 and s1223) from unzipped list, open the *.dat files in notepad and save as *.txt in the same path directory that contains the downloaded xfoil program.

PROGRAM EXECUTION

3Start the downloaded XFOIL program and a DOS command window with ‘Xfoil c>’ appears (including a list of possible commands within the ‘xfoil’ menu).

USEFUL NOTES:

  • At any menu-driven routine/subroutine prompt, type <?> for a list of command options executable within that routine/subroutine. In the list, the commands preceded by a period places the user in another lower-level menu while the lowercase letters i (for integer), r (for real), f (for filename) and s (for character string) following some commands are the type of argument(s) expected by the command
  • In this manual, the commands to be input would be contained in the parentheses “< >”. Do not enter these parentheses
  • Commands are not case sensitive
  • You are not limited to working with specifically the commands implemented below for this analysis. Explore other command options as well
  • Xfoils user manual is downloadable from the above listed website
  • To obtain pictures of plots, use print screen keyboard command

Upload airfoil profile to be analyzed:

Menu: “XFOIL c>”

1To obtain SELIG class airfoils, enter <LOAD f> where f is the filename of the airfoil text file saved in step 2 above (i.e. LOAD s1210.txt). To obtain NACA class airfoils if required, entering <NACA i> automatically invokes the paneling routine to create a buffer airfoil with a suitable paneling. The following ‘i’ is a 4 or 5 digit integer NACA airfoil designation required

  • In some cases, it is desirable to explicitly re-copy the NACA buffer airfoil into the current airfoil via <PCOP>

2Enter <PANE> to set the panel node to the default 140

  • To modify (i.e. increase paneling density) or display current airfoil paneling type, enter <PPAR> and follow prompts. Plots are shown in a ‘PltLib’ browser window that automatically appears. To exit back to previous menu (i.e. XFOIL c> ) simply press enter

Perform Analysis within Operation Analysis Subroutine OPER

Menu: “.OPERi c>”

3Type <OPER> to call the analysis subroutine for direct operating points. Remember, you can enter <?> for a list of possible commands within this menu.

4Input <VISC> to toggle from default invisid flow mode to viscous mode. The subsequent command should read “.OPERv c>”. Enter Reynolds number at the prompt.

5Enter <MACH r> to input air speed mach number

6Enter <TYPE 1> to specify the dependence of Re and Mach on coefficient of lift. Type 1 implies that the airfoil chord and airspeed are held constant while lift is varied

7A polar point contains a set of subsequently calculated parameters when one parameter is incrementally varied. <PACC> thus enables polar point accumulation. Invoke <PACC> and input an arbitrary save filename and dump filename (dump file is not necessary), both of which would automatically create output files in the path directory from which the xfoil program is run. The command should now read ‘.OPERva c>’

  • Other Polar commands to explore are: PLIS (List saved polar), PWRT (Write individual polar to save file), PGET, PREM, PSUM etc. Again, type <?> for complete list of options

8Vary the angle of attack (AOA) (called alfa in xfoil) individually with <ALFA r> (where r is in degrees), or vary as a list with <ASEQ r1 r2 r3>. Aseq prescribes a sequence of AOA from r1 to r2 in r3 increments. The program runs when you press enter

9The plot for every run is shown in the ‘PltLib’ browser window. If this window is not up it can be displayed using <CPX> which plots Cp Vs x (normalized x/c)

10If a program run for a sequential alfa input (aseq) does not converge (this can be determined from reading the last line following a calculation run or the plot in PltLib (<CPX> to display Cp Vs x/c plot) calculation would stop at the failed polar set and xfoil would not write subsequent calculated data set (CL, CD etc) to the polar file. To converge:

  • Rerun calculation by inputting alfa(s) again. A faster way is to rerun a calculation using the previous or updated inputs, is to type <!> then enter
  • Or increase iterations performed (default is 10 iterations) using <ITER>
  • Or vary the angle of attack (AOA) individually with <ALFA r> (where r is in degrees) and rerun or increase iterations

11Enter <PLIS> to list stored polar and confirm that calculated polar set for each alfa increment is read into polar file. If polar sets are missing (convergence problem) repeat 15. Hint: Inputting alfa values individually, and rerunning until convergence increases chances of convergence. Continually check PLIS to confirm

12Sort polars for plotting using <PSOR> and follow prompt (use 1 at prompt)

13Enter <CPX> to show Cp Vs x plot in PltLib

  • Resize PltLib window with <SIZE>. Follow prompt and replot for required alfa.
  • For a given Cp Vs x/c plot, run with the required alfa and display with <CPX>. Enter <HARD> to save a *.ps file type hardcopy of the current plot shown

14Enter <CPV> to show pressure vectors on suction and pressure surface of airfoil. Cool

15Enter <APLO ii> to see airfoil CS plot for stored airfoil(s) for each polar set

16To output x vs. Cp data set to be used to plot in EXCEL (see following section), enter <CPWR f> and input filename or use default

17Quit with <QUIT> (must be in the ‘xfoil c>’ menu so press enter to exit any submenu) and perform same procedure for next airfoil (1 through 16)

Obtain Output Data (Polar Points and Plots)

Menu: “.OPERi c>”

Although, you can obtain Output Data in Xfoil using:

18Enter <PPLO f> to plot stored polar values in PltLib browser

19Enter <HARD> to save a *.ps file type hardcopy of whatever current plot is shown in the PltLib window. As with all newly created files, the output file is created in the path directory from which the xfoil program is run

20To label the plot (please show appropriate labeling for plotted data) use <ANNO> and follow the prompts

21Other useful plots (relating to boundary layer quantities) can be plotted by xfoil from the <VPLO> submenu invoked within the <OPERva> menu. Enter <?> for a list of possible commands within this submenu. Play around.

It is recommended that you obtain the required plots using EXCEL:

22In a new excel worksheet, import the xfoil file(s) for each data file previously saved when the <PACC> command was invoked in the <OPER> menu.

  • Data Menu > Import External Data > Import Data. Hint: Use delimited, check space and tab
  • Import data for both airfoils to one spreadsheet for plotting. Remember to have alpha listed sequentially. Sort if not.

23Use the data for the required plots and appropriately label it

24 A plot of the airfoil section can be obtained by typing in GDES in the ‘xfoil c>’ menu. Use this plot for question 1.