AeroPy Documentation. Release Pedro Leal

Transcription

1 AeroPy Documentation Release Pedro Leal Nov 09, 2018

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3 Contents 1 Aerodynamic tools 3 2 Geometric tools Xfoil Module Documentation Aerodynamic Module Documentation Airfoil Module Documentation Tutorial To Do 13 4 Recommended Collaborations 15 5 Indices and tables 17 6 Examples 19 Python Module Index 21 i

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5 AeroPy is an library for calculating aerodynamic properties. The main feature of this library is the Python interface with XFOIL. The main objective of this library is to use XFOIL via Python iteratively in a total of 4 lines total (one line for most uses). Through this interface coupling with other softwares (Abaqus, Ansys, etc) is possible and iterative processes (optimization, design sensitivity) are also possible. For a thorough explanation please check the documentation and the tutorials. Contents 1

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7 CHAPTER 1 Aerodynamic tools xfoil_module: contains all the functions relating to XFOIL. aero_module: contains all functions related to aerodynamics, but not related to XFOIL. filehandling: constains functions to read and output in several formats 3

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9 CHAPTER 2 Geometric tools geometry: tools to generate and modify airfoils 2D_CST: contains libraries to develop 2D shapes using the Class/Shape Transformation equations 3D_CST: contains libraries to develop 2D shapes using the Class/Shape Transformation equations morphing: functions to generate structurally consistent morphed configurations Contents: 2.1 Xfoil Module Documentation The functions herein defined are purposely made for use with Xfoil Core Function aeropy.xfoil_module.call(airfoil, alfas= none, output= Cp, Reynolds=0, Mach=0, plots=false, NACA=True, GDES=False, iteration=10, flap=none, PANE=False, NORM=True) Call xfoil through Python. The input variables are: Parameters airfoil if NACA is false, airfoil is the name of the plain filewhere the airfoil geometry is stored (variable airfoil). If NACA is True, airfoil is the naca series of the airfoil (i.e.: naca2244). By default NACA is False. alfas list/array/float/int of angles of attack. output defines the kind of output desired from xfoil. There are four posssible choices (by default, Cp is chosen): 5

10 Cp: generates files with Pressure coefficients for desired alfas. Dump: generates file with Velocity along surface, Delta star,theta and Cf vs s,x,y for several alfas. Polar: generates file with CL, CD, CM, CDp, Top_Xtr, Bot_Xtr. Alfa_L_0: generates a file with the value of the angle of attack that lift is equal to zero. Coordinates: returns the coordinates of a NACA airfoil. Reynolds Reynolds number in case the simulation is for a viscous flow. In case not informed, the code will assume inviscid. Mach Mach number in case the simulation has to take in account compressibility effects through the Prandtl-Glauert correlation. If not informed, the code will not use the correction. For logical reasons, if Mach is informed a Reynolds number different from zero must also be informed. plots the code is able to save in a.ps file all the plots of Cp vs.alfa. By default, this option is deactivated. NACA Boolean variable that defines if the code imports an airfoil from a file or generates a NACA airfoil. GDES XFOIL function that improves the airfoil shape in case the selected points do not provide a good shape. The CADD function is also used. For more information about these functions, use the XFOIL manual. iteration changes how many times XFOIL will try to make the results converge. Speciallt important for viscous flows flap determines if there is a flap. In case there is the expected input is [x_hinge, y_hinge, deflection(angles)]. y_hinge is determined to be exactly in the middle between the upper and lower surfaces. PANE if there are more than 495 surface points, the paneling method will not be used. Need to use the PANE subroutine to solve this. It will find the best points that represent the geometry (only 160 of them). NORM For good results using the panel method, Xfoil requires normalized coordinates, so this option should always be True. Return type dictionary with outputs relevant to the specific output type. Usually x,y coordinates will be normalized. As a side note, it is much more eficient to run a single run with multiple angles of attack rather than multiple runs, each with a single angle of attack. Created on Sun Mar 9 14:58: Last update Fr Jul 13 15:38:40 Pedro Leal (Based on Hakan Tiftikci s code) Auxiliary Functions XFOIL Python interface. 6 Chapter 2. Geometric tools

11 aeropy.xfoil_module.alfa_for_file(alfa) Generate standard name for angles. This is mainly used by the file_name Pedro Leal aeropy.xfoil_module.create_input(x, y_u, y_l=none, filename= test, different_x_upper_lower=false) Create a plain file that XFOIL can read. XFOIL only reads file from the TE to the LE from the upper part first and then from the LE to the TE through the pressure surface. Inputs: x: list of coordinates along the chord y_u: list of coordinates normal to the chord for the upper surface. If y_l is not defined it is the y vector of the whole upper surface, y_l: list of coordinates normal to the chord for the lower surface file_name: label used for the file created Created on Thu Feb 27 Pedro Leal aeropy.xfoil_module.file_name(airfoil, alfas= none, output= Cp ) Create standard name for the files generated by XFOIL. Parameters Returns airfoil the name of the plain file where the airfoil geometry is stored (variable airfoil). alfas list/array/float/int of a single angle of attack for Cp and Dump, but the whole list for a Polar. Only the initial and the final values are used output defines the kind of output desired from xfoil. There are three posssible choices: Cp: generates files with Pressure coefficients for desired alfas Dump: generates file with Velocity along surface, Delta star and theta and Cf vs s,x,y for several alfas Polar: generates file with CL, CD, CM, CDp, Top_Xtr, Bot_Xtr Alpha_L_0: calculate the angle of attack that lift is zero The output has the following format (by default, Cp is chosen): for Cp and Dump: output_airfoil_alfa >>> file_name('naca2244', alfas=2.0, output='cp') >>> Cp_naca2244_0200 for Polar: Polar_airfoil_alfa_i_alfa_f >>> file_name('naca2244', alfas=[-2.0, 2.0], output='polar') >>> Polar_naca2244_n0200_0200 for Alpha_L_0: Alpha_L_0_airfoil 2.1. Xfoil Module Documentation 7

12 >>> file_name('naca2244', output='alpha_l_0') >>> Alpha_L_0_naca2244 Created on Thu Mar 16 Pedro Leal aeropy.xfoil_module.find_pressure_coefficients(airfoil, alpha, Reynolds=0, iteration=10, NACA=True, use_previous=false, chord=1.0, PANE=False, delete=false, Calculate the pressure coefficients of an airfoil. GDES=False) aeropy.xfoil_module.output_reader(filename, separator= \t, output=none, rows_to_skip=0, header=0, delete=false, structure=false, type_structure=none) Function that opens files of any kind. Able to skip rows and read headers if necessary. Inputs: Output: filename: just the name of the file to read. separator: Main kind of separator in file. The code will replace any variants of this separator for processing. Extra components such as end-line, kg m are all eliminated. Separator can also be a list of separators to use output: defines what the kind of file we are opening to ensure we can skip the right amount of lines. By default it is None so it can open any other file. rows_to_skip: amount of rows to initialy skip in the file. If the output is different then None, for the different types of files it is defined as: - Polar files = 10 - Dump files = 0 - Cp files = 2 - Coordinates = 1 header: The header list will act as the keys of the output dictionary. For the function to work, a header IS necessary. If not specified by the user, the function will assume that the header can be found in the file that it is opening. delete: if True, deletes file read. structure: the file that he is being read has a given structure. For a file with the following structure: For the case where the header: >> header = [ element, x1, y1, x2, y2 ] A possible structure is: >> structure = [[ element ], [ x1, y1 ], [ x2, y2 ]] type_structure: [ string, time, float, time, float ] Dictionary with all the header values as keys Created on Thu Mar 14 Pedro Leal aeropy.xfoil_module.prepare_xfoil(coordinates_upper, Coordinates_Lower, chord, reposition=false, FSI=False) Prepare XFOIL airfoil file. 8 Chapter 2. Geometric tools

13 The upper and lower functions will be the points in ordered fashion. Because of the way that XFOIL works the points start at the Trailing Edge on the upper surface going trough the Leading Edge and returning to the Trailing Edge form the bottom surface Utility Functions aeropy.xfoil_module.find_coefficients(airfoil, alpha, Reynolds=0, iteration=10, NACA=True, delete=false, PANE=False, GDES=False) Calculate the coefficients of an airfoil. Includes lift, drag, moment, friction etc coefficients. aeropy.xfoil_module.find_alpha_l_0(airfoil, Reynolds=0, iteration=10, NACA=True) Find zero lift angle of attack. Calculate the angle of attack where the lift coefficient is equal to zero. aeropy.xfoil_module.m_crit(airfoil, pho, speed_sound, lift, c) Calculate the Critical Mach. This function was not validated. Therefore use it with caution and please improve Pedro Leal 2.2 Aerodynamic Module Documentation Lifting Line Theory If θ 0 is an arbitrary span-wise location: α(θ o ) = 2b πc(θ o ) N A n sin(nθ o ) + α L=0 (θ o ) + 1 N 1 na n sin(nθ o ) sin(θ o ) (2.1) Each equation has N unknowns (A n ), so if there are N θ o, we have NxN system, which in Einstein notation can be written as: where, i = 0,..., N, j = 0,..., N and : ( C ij = 2b πc(j) + C ij A i = D i (2.2) ) n sin(nθ(i)) (2.3) sinθ(i) A i = A(i) (2.4) D i = α(i) α L=0 (i) (2.5) where n = 1, 3, 5,..., N 1. Since we are considering a symmetric wing, all of the even terms would cancel each other 2.2. Aerodynamic Module Documentation 9

14 2.2.2 The code Current funcionatilities: - Lifting line theory - generate field pressures for Abaqus or other softwares - air properties calculator - Reynolds calculator Created on Mon Jul 20 17:26:19 Pedro Leal aeropy.aero_module.llt_calculator(alpha_l_0_root, c_d_xfoil, N=10, b=10.0, taper=1.0, chord_root=1, alpha_root=0.0, V=1.0) Calculate the coefficients for a Wing. TODO : - Include elliptical wing When alpha_l_0_root = zero, nan! Include non rectangular wings something else? aeropy.aero_module.reynolds(height, V, c) Simple function to calculate Reynolds for a given Pedro Leal Created in Jul aeropy.aero_module.air_properties(height, unit= feet ) Function to calculate air properties for a given height (m or ft). Sources: Created on Thu May 15 14:59:43 Pedro Leal aeropy.aero_module.calculate_moment_coefficient(x, y, Cp, alpha, c=1.0, x_ref=0.25, y_ref=0.0, flap=false) Calculate the moment coeffcient. Inputs are x and y coordinates, and pressure coefficients (Cp). Inputs can be 10 Chapter 2. Geometric tools

15 in a list in xfoil format (counterclockwise starting from the trailing edge, in case necessary, check create_input function from xfoil_module) or dictionaries with upper and lower keys. Parameters flap if true, also calculates the moment contribution from the trailing edge and the panels in front of the flap (that are not directly in contact with the air) aeropy.aero_module.pressure_shell(data, half_span, chord= MAX, air_density=0, Velocity=0, N=10, thickness=0, txt=false, llt_distribution=false, distribution= Uniform, amplifier=1) Converts pressure coefficient data, usually 2D, into a 3D presurre field that Abaqus understands. Can be used for shells (considers thicknesses), but also for any surface. Can do Lifting Line Theory (LLT), Elliptical, and Uniform distributions. If chord= MAX, the maximum value for vector x is used as chord. If data in non-dimensional, use a numerical value. If txt==true, an output textfile is generated. aeropy.aero_module.pressure_shell_2d(data, chord, thickness, half_span, height, Velocity, N, txt=false) Calculate pressure field for a 2D Shell. 2.3 Airfoil Module Documentation The functions herein defined are purposely made for generating NACA airfoils, flapped airfoils and CST airfoils Core Function Auxiliary Functions Utility Functions 2.4 Tutorial 2.3. Airfoil Module Documentation 11

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17 CHAPTER 3 To Do Include asymmetric wing Create airfoil and wing classes 13

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19 CHAPTER 4 Recommended Collaborations Please use and adapt this library to your needs. There are several functionalities I wished to implement, but did not have the time. Hence I am strongly recommending the following collaborations: Airfoil generator with a GUI Atmospheric module (use the library already available in aero_module) Extend aero_module for wings with non-constant cross sections 15

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21 CHAPTER 5 Indices and tables genindex modindex search 17

22 18 Chapter 5. Indices and tables

23 CHAPTER 6 Examples A simple example demonstrating how to use AeroPy to calculate lift, drag and moment coefficients for a NACA0012 airfoil at angle of attack of 12 degrees. from aeropy.xfoil_module import find_coefficients find_coefficients(airfoil='naca0012'alpha=12.) >>> {'CM': , 'CL': , 'Top_Xtr': , 'CD': , 'CDp': , 'alpha': 12.0, 'Bot_Xtr': 1.0} Another example showing how to use AeroPy to calculate pressure coefficients for a NACA0012 in viscous flow at Reynolds = and angle of attack of 12 degrees from aeropy.xfoil_module import find_pressure_coefficients find_pressure_coefficients(airfoil='naca0012', Reynolds = 1e6, alpha=12.,naca=true) >>> 'y': [ ,..., ], 'x': [1.0,..., 1.0], 'Cp': [ ,..., ]} 19

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25 Python Module Index a aeropy.aero_module, 10 aeropy.xfoil_module, 6 21

26 22 Python Module Index

27 Index A aeropy.aero_module (module), 10 aeropy.xfoil_module (module), 6 air_properties() (in module aeropy.aero_module), 10 alfa_for_file() (in module aeropy.xfoil_module), 6 C calculate_moment_coefficient() (in module aeropy.aero_module), 10 call() (in module aeropy.xfoil_module), 5 create_input() (in module aeropy.xfoil_module), 7 F file_name() (in module aeropy.xfoil_module), 7 find_alpha_l_0() (in module aeropy.xfoil_module), 9 find_coefficients() (in module aeropy.xfoil_module), 9 find_pressure_coefficients() (in module aeropy.xfoil_module), 8 L LLT_calculator() (in module aeropy.aero_module), 10 M M_crit() (in module aeropy.xfoil_module), 9 O output_reader() (in module aeropy.xfoil_module), 8 P prepare_xfoil() (in module aeropy.xfoil_module), 8 pressure_shell() (in module aeropy.aero_module), 11 pressure_shell_2d() (in module aeropy.aero_module), 11 R Reynolds() (in module aeropy.aero_module), 10 23