Investigation and Comparison of Airfoils

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1 AENG 360 Aerodynamics Investigation and Comparison of Airfoils Rocie Benavent Chelseyann Bipat Brandon Gilyard Julian Marcon New York Institute of Technology Fall 2013

2 2 Executive Summary Airfoil design has gone through a major progression within the history of flight, and it has been with that progression that advances have been able to contribute to improved stability, reduction in undesirable parameters, and ability to trade-off between previous and current designs. This project dealt with taking an original airfoil that of the Piper Warrior II PA and comparing it to the NACA 2412 airfoil. Originally, the assigned case was a Piper Archer II that utilized the same airfoil (NACA ) so therefore, another airfoil that was used with the Cessna 172K was chosen (NACA 2412). This was because of the fact that the Cessna 172K has a similar horsepower and passenger capacity and weight, and therefore, is a suitable choice for analyzing the effect of swapping airfoils. Within the scope of Theory of Wing Sections by Abbott and Von Doenhoff, Jane s All the World s Aircraft, and other scholastic journals, the wing comparisons focused on the following: operating range parameters, wing loading requirements, airfoil profiles, lift coefficients and angles of attack, drag polars, design lift coefficients, maximum L/D, and the respective NACA designations. The entire comparison and trade-off analyses processes consisted of calculating both airfoil lift coefficients for the plane speeds obtained from their respective references; determining the range in angle of attack for those speeds; discussing trading off the original airfoil with the alternate; analyzing the changing of angle of incidence, dihedral, and angle of twist; addressing how 3D flow relates to the aforementioned topics; and utilizing Design Foil (an aerodynamics software package) to show effects of increased thickness and camber on the original wing. Ultimately, it is seen that the comparison yields an unfavorable result. It is not wise to replace the original NACA airfoil with the NACA 2412 as a decrease in lift and increase and drag were obtained. Even though the designs are valid for aircraft of a similar variety, the numbers for the sections chosen are not valid.

3 3 Table of Contents Executive Summary...2 Part Introduction/Problem Statement...4 Description of Selected Wings...5 Describe NACA information and all available data for both wings...12 Part Airfoil Lift Coefficients Range in Angles of Attack Replacing the original airfoil with the alternate Change in the Angle of Incidence Effects Change in the Dihedral or Angle of Twist Benefits For 3, show how 3D flow (airfoil vs wing) changes the above results DesignFoil simulation of increased camber and thickness...19 References...22

4 4 Part 1 Introduction/Problem Statement Since the advent of the first plane successfully flown by the Wright brothers, all aviation aircraft have continued to advance in areas such as performance, efficiency, range, and material constituents just to name a few. Airplane design has been a process that examines several parameters, characteristics, and adherence to stringent requirements. These categories have focused on, but have not been limited to: cruise speed, take-off speed, landing speed, stall speed, glide speed; aerodynamic efficiency and altitude limitations, take-off and landing distances, power and wing loading, airfoil profiles, lift coefficients and associate angles of attack, drag polars, Lift to Drag (L/D) maximums, and NACA profile designations. The aforementioned factors, and their relations to all types of aircraft, have had a high influence on the improvements on future designs as a result of comparisons. Wing selection is one of the most important steps in the design of an airplane. After deciding the basic design configurations such as overall weight, size, capacity, flying range comes the choice of an effective airfoil profile for the airplane in process of design. There are now a vast amount of different profiles - NACA series being among the most famous and widely used ones - that differ in many parameters. The first aspect that should be considered when selecting a wing profile is the type of airplane that is being designed. Indeed, different airplane types such as commercial transport jets, fighter jets, propeller-driven airplanes or even gliders will have wings with significant differences in their aerodynamic properties. These properties are affected by basic airfoil parameters such as chord length, thickness, and camber. Following basic airfoil parameters, other factors affect the aerodynamic forces applied to the tridimensional wing. These include angle of attack, angle of incidence, taper ratio, and dihedral angle among others. This project aims to compare two different airfoils in terms of aerodynamic properties. These properties will first be identified from official references and compared for an accurate overview of the problem. Then, analyses will be made and pros and cons discussed about the possibility of replacing the original airfoil with the newer one. Main improvements are sought in increase of lift coefficient and decrease of drag coefficient for better overall aerodynamic efficiency. Further implications of this change of wing profile will also be discussed as well as

5 5 other possibility modifications such as a change in angles of incidence and of twist and dihedral angle. Finally, educational airfoil simulation software such as DesignFoil will be used to identify the results of additional modifications to the wing profile such as increased camber and thickness. Description of selected wings: A - List Operating Range Parameters available from Reference 2 a. Speeds - cruise, take-off, landing, stall (without flaps),best glide b. Aerodynamic efficiency (lift to drag ratio) c. Altitude limitation d. Take off/landing distance minimums B - List Wing Requirements from Reference 2 a. Power loading b. Wing loading Answers to parts A and B are presented below in a Table format: 2

6 6 C - Identify Characteristics from Reference 1 a. Draw airfoil profiles NACA NACA 2412

7 b. Lift coefficients vs. angle of attack 7

8 8

9 c. Drag polar 9

10 10

11 11 d. Design lift coefficient From the NACA 6-series notation described below, the design lift coefficient is indicated by the first digit after the hyphen, in tenths. The wing used is NACA , therefore 4/10 = 0.4. Design lift coefficient = 0.4 e. Maximum L/D From the drag polar graph for NACA , this occurs at the point C l =0.6, and C d = The max L/D at this point is From the drag polar graph for NACA 2412, this occurs at the point C l =1.1, and C d = The max L/D at this point is This is one of the most important aspect of considering whether or not the airfoil is suitable because this is the lift coefficient at the best glide speed, which is when it is most important to evaluate whether or not the airfoil will be effective at all. It is also important to note that the best glide speed for both aircrafts are taken from the pilot s operating handbook. It should also be noted that because of this, this includes the lift and drag included with the empennage and fuselage of the airplane. f. NACA or other profile designation - explain NACA This new series shows an improvement over 1-series airfoils with emphasis on maximizing laminar flow. The airfoil is described using six digits in the following sequence: 1. The number "6" indicating the series. 2. One digit describing the distance of the minimum pressure area in tens of percent of chord. 3. The subscript digit gives the range of lift coefficient in tenths above and below the design lift coefficient in which favorable pressure gradients exist on both surfaces 4. A hyphen. 5. One digit describing the design lift coefficient in tenths. 6. Two digits describing the maximum thickness as percent of chord.

12 12 NACA The NACA four-digit wing sections define the profile by: 1. First digit describing maximum camber as percentage of the chord. 2. Second digit describing the distance of maximum camber from the airfoil leading edge in tens of percent of the chord. 3. Last two digits describing maximum thickness of the airfoil as percent of the chord. Describe NACA information and all available data for both wings Piper Warrior II PA Wing and Power Loading The Piper Warrior II, which features tapered wings, is quite the capable aircraft. The wing requirements, as stated in Reference 2, for both maximum wing and maximum power loading are kg/m 2 or lb/sq ft, and 8.86 kg/kw or lb/hp, respectively. Other Wing Data The associated areas for wings, ailerons, trailing-edge flaps, fin, rudder, and tailplane, of the Warrior II are: 15.8 m 2, 1.23 m 2, 1.36 m 2, 0.69 m 2, 0.38 m 2, and 2.46 m 2, respectively. The wing span, as stated in Reference 2, is m (35 ft); the wing chord at the root and tip are 1.60 and 1.07 m respectively; the aspect ratio is 7.24; the overall length and height are 7.25 m and 2.22 m respectively; the tailplane span is 3.96; the wheel track and wheelbase are 3.05 m and 2.03 m respectively; the propeller diameter and ground clearance are 1.88 m and 0.21 m respectively; and the cabin door height and width are 0.89 and 0.91 m respectively. Maximum Speeds The maximum take-off speed and optimum power cruising speed at 2,745 rpms of the Warrior is 127 knots; the best power cruising speeds at 65 percent and 55 percent, are 118 knots and 107 knots at 12,500 ft, respectively. Length Attributes/Capabilities In reference to Abbott and Von Doenhoff, the absolute ceiling and landing run for the Warrior II is 4,600m and 181m respectively.

13 13 Part 2 1. Airfoil Lift Coefficients 2. Range in Angles of Attack

14 14 Piper wing NACA From the drag polar graph for NACA , this occurs at the point C l =0.6, and C d = The max L/D at this point is For NACA airfoil, angle of attack = 2 degrees. TO min will be achievable by using less than max take off weight. Highest angle of stall 16 degrees Angle of cruise 0 degrees Range min Cl (cruise) = , stall cd =0.02 Best glide for Cl = 0.75, Cd =0.008, AOA = 4 deg Cessna wing NACA 2412 At least roughness CL = 1.7 at stall at 16 degrees AOA Cruise AOA 0 degrees, min Cl = Stall Cd = Best glide - for Cl = 0.75, Cd=0.0075, AOA= 5 deg Comparison At stall, there is less drag associated with the NACA 2412, but at cruise speed, there is more drag with NACA 2412 than with NACA The angles of attack are at the same range (0-16 degrees), and the drag coefficients at stall are approximately equal (0.02). 3. Replacing the original airfoil with the alternate a. Comparison between lift coefficients or reduction in drag coefficients at speeds: For NACA , lift and drag coefficients are as follows: NACA NACA 2412 Cl Cd Cl Cd Cruise Best Glide Stall TO Max Land min Land max

15 15 It can be seen that replacing the NACA airfoil with the NACA 2412 airfoil results in a decrease in lift, and an increase in drag. This shows that the NACA 2412 airfoil is not an ideal replacement for the NACA , even though that both wings are ideally designed for similar aircraft type. b. Compare the angle of attack with the original airfoil at speeds. As seen from the table below, the effective angles of attack are slightly different due to the fact that aspect ratios are different. This angle of attack was found by converting from the geometric angle of attack and induced angle of attack by using the following formulas: AOA i = C L π AR 180 π AOA eff = AOA AOA i C l = dc l daoa (AOA eff AOA 0l ) Where dc_l/daoa was calculated as the slope from the lift coefficient vs angle of attack graphs given previously. NACA AOA eff AOA eff Cruise Best Glide Stall TO Max Land min Land max c. Compare maximum L/D for both airfoils. From the drag polar graphs, the maximum L/D for the airfoils are found from the drag polar graphs, at the point of most lift and least drag. From the drag polar graph for NACA , this occurs at the point C l =0.6, and C d = The max L/D at this point is From the drag polar graph for NACA 2412, this occurs at the point Cl=1.1, and Cd= The max L/D at this point is

16 16 d. Discuss implications of airfoil cross sectional area change. The most important effect that increased camber has on the original airfoil is increased lift at almost all angles of attack. The only exception is the NACA 7412, which features a maximum camber of 7% the chord length. The reason might be that transition between laminar and turbulent flow over the airfoil occurs so early that, although not at stall yet, it is not able to produce more lift at high angles of attack. Another effect, which is negligible at small camber but becomes important at large camber, is an increased drag at low lift coefficient and therefore low angles of attack. This means that, if only little is required, the drag will be more important because of the increased camber. In other words, increased camber should be used on airfoils only when high lift is required. A last effect is the increased moment coefficient, which is to be taken in account when designing the wing internal structure. 4. Change in the Angle of Incidence Effects The angle of incidence, unlike its close relative the angle of attack is the angle that is formed between the wing s chord line and the lengthwise axis of the aircraft plane or that of the fuselage. Suffice it to say, that the angle of attack is representative of the angle between the chord of the airfoil/wing and that of the airflow direction. A change in the angle of incidence (which is usually a fixed parameter) would most certainly change the angle of attack. This would resultantly adjust the lift that is generated by the airfoil as well as the coefficient of lift. 5. Change in the Dihedral or Angle of Twist Benefits On the other hand, the involvement of the dihedral or angle of twist consists of a bit more analysis in relation to the benefits therein. It is to be duly noted that there is a distinctive difference between a dihedral and angle of twist. But, it is also appropriate to know that there exists both aerodynamic and geometric twist. Geometric twist relates to the physical change in the airfoil s angle of incident from the root, and the notion of the airfoil washing out as one would move away from the fuselage. Similarly, the aerodynamic twist represents the angle amongst the angle of zero lift for an airfoil section as well as the angle of zero lift at the

17 17 respective root. Thus, changing the angle of twist allows for the outer section of the wing avoid imminent stall. This is simply due to the reduction in angle of attack in the respective outboard region. Whereas the twist is in and of itself a beneficial parameter, the dihedral purports a similar tactic. This tactic, however, is quite dependent on center of gravity, wing sweep, and any other component that affects the roll stability. If an aircraft is pictured from its front or nose, the dihedral would be, the upward angle of the wing from the vertical. For instance, as Jackson s Wing Twist and Dihedral, If each wing is angled 5 up from the horizontal, then the wind is said to have 5 of dihedral (Jackson 3). The dihedral simply provides stability and an increase in lift for an airfoil/aircraft that may be experiencing yawing or rolling moments. 6. For 3, show how 3D flow (airfoil vs wing) changes the above results Piper: NACA AOA (deg) AOA_eff (deg) C_L C_l Cruise Best Glide Stall TO max Land min Land Max

18 18 Cessna: NACA 2412 AOA (deg) AOA_eff (deg) C_L C_l Cruise Best glide Stall TO max Land min Land max As it can be seen, the effective angle of attack can, can some cases, be much lower than the actual angle of attack. This has 2 important consequences. The first and maybe most important effect is that there is a large loss of lift efficiency. Indeed, the infinite wing section lift coefficient is lower than the finite actual wing lift coefficient. This loss can go from 15% to almost 50%. This is greatly influenced by the aspect ratio. The larger the aspect ratio, the closer the finite wing gets to an infinite wing, and the smaller the losses. These losses are due to the downwash vortices at the tip of the wing. This is the reason why most of the modern commercial jets are equipped with wingtip devices to reduce these vortices and therefore lift losses. Another consequence of this decrease in effect angle of attack is the faster reach of stall. Indeed, finite wings need to have higher angle of attack in order to achieve the same lift coefficient as for an infinite wing. Because of this larger angle of attack, stall angle is reached sooner while the effective angle of attack remains relatively low. Stall is therefore reached faster and a large part of the effective angle of attack range remains unused.

19 19 7. DesignFoil simulation of increased camber and thickness For this part of the project, DesignFoil was chosen because of its large library featuring most of the existing and used airfoils. It has also an automatic plotter for NACA 4-digits series that can come in very handy. As for the existing airfoils from the data, they can also, to some extent only obviously, be modified.the software can simulate airflow over the airfoil at various angles of attack, plot pressure coefficients over the surface, and plot the typical Lift Coefficient versus Angle of Attack and Drag Polar graphs. A lot of settings are customizable so that it makes it very easy to compare different airfoils or a same airfoil with different modifications. For the case of increased camber, the NACA 65(2)-415 airfoil could not be used because the camber line cannot be modified from the database plot. The alternate NACA 2412 airfoil was used instead. As a reminder, the first digit of the 4-digits index indicates the maximum camber in percent of the chord length. This way, it was easy to modify the geometry of the airfoil by simply modify the first digit of the series and letting the software plot the section. The original airfoil features a maximum camber of 2% of chord length and it was compared to airfoil with modified maximum camber in increment of 1%.

20 20 The most important effect that increased camber has on the original airfoil is increased lift at almost all angles of attack. The only exception is the NACA 7412, which features a maximum camber of 7% the chord length. The reason might be that transition between laminar and turbulent flow over the airfoil occurs so early that, although not at stall yet, it is not able to produce more lift at high angles of attack. Another effect, which is negligible at small camber but becomes important at large camber, is an increased drag at low lift coefficient and therefore low angles of attack. This means that, if only little is required, the drag will be more important because of the increased camber. In other words, increased camber should be used on airfoils only when high lift is required. A last effect is the increased moment coefficient, which is to be taken in account when designing the wing internal structure. For the increased thickness, the original NACA 65(2)-415 could be used because of the thickness modification feature of DesignFoil. The original airfoil which possesses a maximum thickness of 15% of the chord was compared with airfoils with modified maximum thickness in

21 21 increments of 2%, i.e. the first airfoil of the chart has 15% of the chord maximum thickness, the second has 17%, the third has 19%, and so on. The most important consequence of increasing maximum thickness is an important increase in moment coefficient. This can lead to important changes in the wing internal structures. Also, at low angles of attack, which represent mainly the region that is not close to stall angle, lift coefficient is slightly increased by the increased thickness. Drag does not get modified except that it gets shifted toward higher lift coefficients.

22 22 References Avallone, Eugene A., Theodore Baumeister III, Ali Sadegh. Marks Standard Handbook for Mechanical Engineers, Eleventh Edition. NY: McGraw-Hill Professional Publishing, Book. Camtrell, Paul. Angle of Attack-Helicopter Aviation. Web. 10 December Jackson, Doug. Wing Twist and Dihedral. 2 December Web. 10 December 2013.

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