AERODYNAMICS_Airfoil_Analysis_Using_XFLR

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<p>Universidad Carlos III de Madrid</p><p>Aerodynamics</p><p>Aerospace Engineering</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>Ignacio Egido Garćıa</p><p>Enrique Labiano Laserna</p><p>Francisco Luis Ruiz Sánchez</p><p>Francisco Jesús Santos-Olmo Dı́az de</p><p>los Bernardos</p><p>100329532</p><p>100329870</p><p>100329495</p><p>100329884</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>Contents</p><p>1 Introduction 2</p><p>2 Results 3</p><p>2.1 NACA 633018 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3</p><p>2.2 NACA 631012 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4</p><p>2.3 NACA 63009 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6</p><p>2.4 NACA 64A006 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8</p><p>3 Conclusion 10</p><p>1</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>1 Introduction</p><p>The goal of this practice is to compare the behavior and stall of four different airfoils depending on its geometry,</p><p>which are shown in Figure 1.</p><p>Some of the characteristics that will be analyzed are the drag coefficient against the lift coefficient, the lift coefficient</p><p>against the angle of attack and the pressure coefficient for different angles of attack.</p><p>For this to be done, the flow will not be taken as ideal, so the boundary layer will appear and will be taken into</p><p>account. Also, the Mach number will be equal to zero and the Reynolds number will be fixed to be 5.8 · 106.</p><p>The software used to make the simulation will be XFLR5, which is able to give numerical results for many different</p><p>airfoils configurations under a large number of conditions. These results will be compared with the ones that appear</p><p>in the ’NACA Technical Note 2502’, a guide that contains experimental results of the given airfoils.</p><p>Figure 1: Profiles of the four airfoil sections</p><p>2</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>2 Results</p><p>2.1 NACA 633018</p><p>Figure 2 shows how the lift coefficient varies with respect to the angle of attack, compared with the experimental</p><p>result. Until an angle of attack of 13 degrees, both simulated and experimental results are almost the same,</p><p>increasing following a straight line that can be adjust to Cl = 2πα. But, when it reaches an angle of 13 degrees</p><p>and a lift coefficient of approximately 1.4, the experimental airfoil go through stall. The numerical line can be</p><p>predicted to undergo stall also but in a higher angle of attack. Stall happens at the moment when the boundary</p><p>layer separates, and the reason of this difference is the viscous effects, that in reality have a higher importance.</p><p>Figure 2: Lift coefficient vs angle of attack for NACA 633018</p><p>Polars are displayed in Figure 3, which is a graph that faces the lift coefficient against the drag coefficient. In</p><p>this case both numerical and experimental results behaves in the same way but with slightly lower values for the</p><p>experimental ones. The pattern of the lines increase fast until a point where stall occurs, and after that, the drag</p><p>coefficient has no significant change, while the lift coefficient starts to decrease gradually.</p><p>Figure 3: Polars for NACA 633018</p><p>3</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>Finally, for this airfoil geometry, the pressure coefficient along the chord is plotted for two different angles of attack,</p><p>6o and 16o. It is important to note that 6 degrees is before stall occurs and 16 degrees is after stall, which means</p><p>that the boundary layer has separated.</p><p>A suction peak at the leading edge is much more greater when the angle of attack is 16 degrees. And, as the</p><p>boundary layer separates at more or less 13 degrees; for 6 degrees the boundary layer is still joined, but for 16</p><p>degrees is clearly the opposite. This can be seen due to the shape of the graph, that starts at the trailing edge</p><p>straight but then it grows sharply until reaching the leading edge. This process is called trailing edge stall.</p><p>The last thing to be noted is that the lift that each angle of attack will produce can be seen by computing the area</p><p>of the graph. It can be seen that the area for 16 degrees is greater than the one for 6 degrees, so the lift will be</p><p>greater.</p><p>Figure 4: Pressure coefficient vs x/c for NACA 633018</p><p>2.2 NACA 631012</p><p>A thiner airfoil is now studied, but it is also thick compared to the following ones. In Figure 5 we can observe how</p><p>the lift coefficient varies with the angle of attack. Both, numerical and experimental results are plotted and it can</p><p>be observed how the graphs grow linearly until the stall occurs. In this case, an abrupt decrease in the lift occurs,</p><p>what is characteristic in a leading-edge stall, which consists of a flow separation in the upper surface near the leading</p><p>edge without subsequent reattachment. The behavior after the stall differs a bit between both representations, this</p><p>is due to the models used, but the maximum lift coefficient is approximately the same.</p><p>4</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>Figure 5: Lift coefficient vs angle of attack for NACA 631012</p><p>In Figure 6 the lift coefficient against the drag coefficient is plotted. As occurs in the first airfoil, when the lift</p><p>increases, so does the drag, until the stall point is reached. However, in this case, the loss of lift is sharper because</p><p>of the abrupt separation of the flow. It can be also observed that after stall, the lift decreases as drag increases and</p><p>both plots shows a similar behavior, although the numerical lift increases more linearly.</p><p>Figure 6: Polars for NACA 631012</p><p>Finally, the Cp distribution for two different angles is shown in Figure 7. It can be observed that for a 7.8 degrees</p><p>angle of attack, a small suction peak occurs at the leading edge, which is due to the the formation of a small bubble</p><p>in that region of the airfoil. This phenomenon does not occur in the previous airfoil, which experienced a softer,</p><p>trailing-edge, stall, and this is a key to identify the actual type of stall of this NACA. At an angle of 13.8 degrees,</p><p>the suction peak is much higher than in the previous case, but it stabilizes earlier.</p><p>5</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>Figure 7: Pressure coefficient vs x/c for NACA 631012</p><p>2.3 NACA 63009</p><p>Figure 8: Lift coefficient vs angle of attack for NACA 63009</p><p>As it can be observed in Figure 8, the numerical and experimental lift coefficients shows a linear behavior until a</p><p>specific angle of attack. These results are consistent since according to thin airfoil theory, the lift coefficient for an</p><p>airfoil with no camber (symmetric) is Cl = 2πα. This linear tendency with a slope of 2π is appreciated until 8.5</p><p>degrees for the experimental case and until 15 degrees for the numerical one. At some specific point, there should</p><p>be a sudden change in the lift as it occurs in the experimental case. This abrupt change is produced due to the</p><p>separation of the boundary layer that will produce the aircraft to enter into stall conditions. In the numerical one,</p><p>this sudden decrease of the lift does not appear because viscous effects are not as prominent as in the experimental</p><p>one and therefore the boundary layer separation occurs for a higher angle of attack and in a smoothly way compared</p><p>to the experimental case.</p><p>6</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>Figure 9: Polars for NACA 63009</p><p>Now in Figure 9, the polar is plotted. For both cases there is a sharply increase in the lift until some specific point.</p><p>Once this point is reached, stall conditions take place. From this point on, the lift is maintained almost constant for</p><p>the experimental case. For the numerical one, the stall occurs a bit later and this is the reason why the increase of</p><p>lift is maintained a little more time than the experimental. Once stall occurs, the lift presents a smoothly decrease.</p><p>The results should have been more similar, but the differences in both cases could have been produced to the fact</p><p>that in the experimental the viscous effects</p><p>are more prominent. This is the reason why stall occurs always earlier,</p><p>because the drag effects are higher and so the separation of the boundary layer is produced for a lower value of the</p><p>angle of attack.</p><p>Figure 10: Pressure coefficient vs x/c for NACA 63009</p><p>Figure 10 shows the distribution of the pressure coefficient along the x coordinate of the airfoil. In this case, the</p><p>angles of attack studied will be 6 and 9 degrees. It can be appreciated that the pressure coefficient increases with the</p><p>angle of attack. There will be some key point in which the pressure coefficient will not grow up due to the separation</p><p>of the boundary layer. As in the previous cases, this point is known as the suction peak and it is produced at the</p><p>leading edge (x=0). This phenomenon can be explained as follows. According to thin airfoil theory, the value of θ</p><p>7</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>at the leading edge is zero. Due to this value of theta, the value of γ(θ) would go to infinity (see Equation 1) and</p><p>therefore Cp as well (Equation 2)</p><p>γ(θ) = 2U∞[A0</p><p>1 + cos(θ)</p><p>sin(θ)</p><p>+</p><p>∑</p><p>Ansin(nθ)] (1)</p><p>Cp = ∓</p><p>γ</p><p>U∞</p><p>(2)</p><p>In order to avoid this, viscous effects appear to produce the reattachment of the boundary layer. Since the simulation</p><p>is produced with a value of the Reynold number of 5800000, the flow can be considered turbulent. In theory,</p><p>a turbulent boundary layer is more energetic and so, more able to overcome adverse pressure gradient. As a</p><p>consequence, the flow is separated later resulting in a less drag. However, skin friction will be higher if compared</p><p>with the laminar boundary layer. In the laminar case, the boundary layer is separated before resulting in a higher</p><p>drag but lower skin friction.</p><p>So, it can be concluded that for the NACA 63009 a leading-edge stall occurs. The separation bubble is formed at</p><p>the leading edge, where the boundary layer starts to separate. The width of the boundary layer increases as we</p><p>move backward, towards the trailing edge.</p><p>2.4 NACA 64A006</p><p>Figure 11: Lift coefficient vs angle of attack for NACA 64A006</p><p>In Figure 11, as in the previous cases, it can be observed that both numerical and experimental cases grow at</p><p>the same time following the linear distribution stated by the thin airfoil theory Cl = 2πα. On the one hand, the</p><p>numerical case reaches a maximum value for the lift coefficient of about 0.75 when the angle of attack is about 7</p><p>degrees, that indicates that for values of the angle of attack higher than 7 degrees the airfoil will enter in stall. On</p><p>the other hand, for the experimental case the angle of attack which produces stall is slightly higher, being about 9</p><p>degrees and the lift coefficient near 0.9. In this case, opposite to the previous, the experimental values are higher</p><p>than the numerical ones. This difference between the experimental and the simulation may be produced due to a</p><p>poor mesh in the software, since the higher importance that the viscous effects have in the reality in comparison</p><p>with the model should make that the detachment of the boundary layer,and then stall is produced for a smaller</p><p>angle of attack in reality than in the model.</p><p>8</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>Figure 12: Polars for NACA 64A006</p><p>The polars (Cl vs Cd) are plotted in Figure 12. The values of the lift coefficient grow sharply for both cases until the</p><p>moment when the airfoils enter into stall. Again, it can be checked that stall is reached later by the experimental</p><p>case, differing with the previous cases. Once the maximum values of the lift coefficient are reached and the airfoils</p><p>enter into stall, the value of the lift coefficient decreases smoothly for an increasing drag coefficient.</p><p>Figure 13: Pressure coefficient vs x/c for NACA 64A006</p><p>The relationship of the pressure coefficient along the span of the core is shown in Figure 13. Two different angles</p><p>of attack has been selected: 4.5 degrees corresponds to an angle of attack smaller than the one which produces stall</p><p>and 8 degrees corresponds to an angle of attack higher than the one which produces stall. The lift produced can</p><p>be calculated as the area enclosed between the curves of the pressure coefficient for the lower and upper surfaces.</p><p>Although the area for angle of attack of 8 degrees is greater than the one for angle of attack of 4.5 degrees, it must</p><p>be taken into account that the airfoil has entered into stall for 8 degrees. So, the lift will increase with the angle of</p><p>attack until the airfoil enters into stall, and then if the angle of attack continues increasing, the lift will decrease.</p><p>It can be clearly noticed that the value of the angle of attack that produces stall for this NACA is notably smaller</p><p>than for the previous cases, so it can be stated that it suffers thin airfoil stall. This type of stall is characterized by</p><p>the formation of a leading edge bubble for small angle of attack, producing then a premature leading edge stall.</p><p>9</p><p>Laboratory Session II:</p><p>Airfoil Analysis Using XFLR5</p><p>3 Conclusion</p><p>It can be observed from the obtained numerical results that the model and software that have been used for this</p><p>practice are such a good approximation since they do not differ so much with respect to the real experimental data</p><p>from ’NACA Technical Note 2502’.</p><p>The main idea that can be extracted is that the geometry of the airfoil can vary dramatically its performance, being</p><p>very clear that increasing the thickness of the airfoil it is increased the maximum angle of attack and the lift but</p><p>also the drag.</p><p>Figure 14: Comparison of the lift coefficient vs angle of attack for the different NACA</p><p>In Figure 14, the lift coefficient versus the angle of attack is displayed for the four different airfoils studied. Together</p><p>with the previous commented results this graph shows the different type of stall that each of the airfoils suffer:</p><p>- NACA 633018: Trailing-edge stall (t>16%c): It is characterized by a progressive and gradual movement of</p><p>separation of the boundary layer from the trailing edge to the leading edge as the angle of attack increases. The</p><p>stall is soft.</p><p>- NACA 631012: Leading-edge stall (9%c</p>