{ "cells": [ { "cell_type": "markdown", "metadata": {}, "source": [ "###### Content provided under a Creative Commons Attribution license CC-BY 4.0; code under BSD 3-Clause license. (c)2015 L.A. Barba, Olivier Mesnard, Pi-Yueh Chuang, Natalia Clementi." ] }, { "cell_type": "markdown", "metadata": {}, "source": [ "# Vortex-source panel method" ] }, { "cell_type": "markdown", "metadata": {}, "source": [ "In [Lesson 9](09_Lesson09_flowOverCylinder.ipynb) of _AeroPython_, you learned to use a _source panel method_ to represent a circular cylinder, and in [Lesson 10](10_Lesson10_sourcePanelMethod.ipynb) we used it for a symmetric airfoil at zero angle of attack. But what if we want the airfoil to generate some lift? If we place the airfoil at a non-zero angle of attack, we _should_ get lift, but will a source-panel representation be able to give you lift? Remember the [_Kutta-Joukowski theorem_](http://en.wikipedia.org/wiki/Kutta%E2%80%93Joukowski_theorem)?\n", "\n", "\n", "Historically, the first panel method ever developed was a source-sheet method. At the time, Douglas Aircraft Company was concerned with calculating the flow around bodies of revolution, and it was only later that the method was extended to lifting surfaces. (See the reference below for a nice historical account.)\n", "\n", "A *source-panel method* leads to a solution with no circulation, therefore no lift. The objective of this lesson is to start with the source panel method we implemented in the previous lesson and add some *circulation* so that we may have a lift force. We introduce an important concept: the **Kutta-condition** that allows us to determine what the right amount of circulation should be." ] }, { "cell_type": "markdown", "metadata": {}, "source": [ "##### Reference\n", "\n", "* Smith, A.M.O., The Panel Method: Its Original Development. In _Applied Computational Aerodynamics_, Vol. 125, edited by P.A. Henne, published by AIAA (1990). [Read it on Google Books.](http://books.google.com/books?id=5Ov2tHj0wxoC&lpg=PA3&ots=SnUiqcdEnb&dq=The%20Panel%20Method%3A%20Its%20Original%20Development&pg=PA3#v=onepage&q&f=false)" ] }, { "cell_type": "markdown", "metadata": {}, "source": [ "## A lifting-body panel method" ] }, { "cell_type": "markdown", "metadata": {}, "source": [ "If we were to simply increase the angle of attack in the freestream and calculate the flow with a source sheet only, the rear stagnation point will not be located at the trailing edge. Instead, the flow will bend around the trailing edge and the stagnation point will be somewhere on the top surface of the airfoil. This is not a physically possible solution.\n", "\n", "For example, using the source-sheet panel method of [Lesson 10](10_Lesson10_sourcePanelMethod.ipynb) with an angle of attack $\\alpha=4^\\circ$ (using 40 panels), and plotting the streamlines in an area close to the trailing edge, we get the following plot:\n", "\n", "