Journal of Aerospace Engineering and Technology Volume 3, Issue 2, ISSN: 2231-038X Aerofoil Profile Analysis and Design Optimisation Kondapalli Siva Prasad*, Vommi Krishna, B.B. Ashok Kumar Department of Mechanical Engineering, Anil Neerukonda Institute of Technology & Sciences, Visakhapatnam, India Abstract The aerofoil section is the incarnation of a wing or a lifting surface which is very important in an airplane wing design. While the shape of the aerofoil changes, their aerodynamic characteristics also change. This paper deals with a standard symmetrical aerofoil as reference and the effect of changes in shape due to minor variations in the coordinates. Eight new aerofoil shapes have been produced in this optimisation process. The aerodynamic characteristic results such as the coefficients of lift and drag (Cd, Cl), pressure coefficient (Cp), moment coefficient (Cm) are noted for all eight different profiles, produced from the standard NACA 0012 aerofoil with changes in the chord thickness distance but no change in the maximum thickness in percentage of chord. The modus-operandi used in this optimisation process is the Computational Fluid Dynamics (CFD). Flow changes have been recorded for these aerofoil shapes and the results are arrived for finding the best aerofoil that can be advisable to be used in compressors, turbines, etc. with reduced flutter and maximum life. Keywords: symmetrical aerofoil, aerofoil shape, aerodynamic characteristics, CFD, co-ordinates, NACA 0012 *Author for Correspondence E-mail: kspanits@gmail.com INTRODUCTION With reference to many aerofoil profiles, a method to design new aerofoil shapes has been undertaken in this research work by altering the position of maximum thickness along the aerofoil chord. The 4-digit NACA aerofoil series has been chosen as reference to start with. Experimental investigations have been carried out on the standard NACA 0012 symmetrical aerofoil and the responses to the changes in the aerofoil shapes have been noted. Starting from the pressure variations to the aerodynamic characteristics and structural strength, different shapes give various comparative results. For varying Mach number and Angle of Attack, the results are compared and charted. Shape is an important factor in deciding the aerofoil performance and control. The aerodynamic characteristics cause changes to the aerodynamic forces and pressures on the upper and lower surfaces of the aerofoil. The works done so far in the history of aerofoil have led to the usage of many aerofoil families and shapes. By analyzing the air flow over and around the aerofoil for different flow conditions, the shape can be changed as desired to yield a better output in the form of a smoother flow [1]. HISTORY OF AEROFOIL Aerofoil profiles were designed based on some major needs. One was to meet the requirements of flight and the other was to develop new concepts of slender, sleek and efficient shapes. In the 1800 s, the works on aerofoil started with advancements continuing till today. Keeping birds flight in mind, the flat plate was kept at an angle of incidence to the incoming airstreams and the lift forces were derived. Further, the curvature was applied to the leading edges of the flat plate and hence to avoid retardation of air speed over it. Many aerodynamicists and researchers have paved way to the modified shapes and sizes of the aerofoil have to be used for their specific research tasks and applications. One such research led to the invention of Gliders the inspiration to fly along with the wind. And the control over the gliders needed air flow past their wings from the leading edge, which was satisfactorily achieved at first JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 1
Aerofoil Profile Analysis and Design Optimisation Prasad et al. and then controlled with manual adjustments during flight. Eminent personalities who started sharing their ideas were Sir. George Cayley, Horatio F Phillips, Otto Lilienthal, Wilbur Wright, Orville Wright etc. Next to their works, changes were made by the National Physics Laboratory (NPL) and NACA in the late 1930 s with common names as the 4-digit and 5-digit series of aerofoil after testing in the virtual wind tunnel at Langley Aeronautics Laboratory. Now there are many aerofoil series suited for individualistic applications and performances. In the 1990 s supercritical aerofoil profiles were meeting the high speed flying conditions and thus the evolution of high speed aerofoil designs took place [2]. NACA 0012 SYMMETRIC AEROFOIL The reason to choose the NACA 0012 symmetric aerofoil is that, it is the first family series of aerofoil in the history and the most common kind used for research purposes in most cases. Symmetrical aerofoil reduces the complexity in design and imparts easiness in the design modifications of the aerofoil. The centre of pressure remains at a constant position as the upper and lower surfaces are identical in a symmetrical aerofoil. This reduces problems of Cp variations with varying angles of attack of air flow over the aerofoil. With changes in the positions of maximum thickness in percentages of chord and along the chord, the following profiles have been named as per their specification criteria. It is to be noted that the amount of maximum thickness is not disturbed in this research content. The eight modified profiles are shown below in Figure 1. Fig. 1: Eight Modified Aerofoil Profiles. JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 2
Journal of Aerospace Engineering and Technology Volume 3, Issue 2, ISSN: 2231-038X TESTING CONDITIONS MADE TO MODIFIED NACA 0012 SYMMETRIC AEROFOIL PROFILES The following changes have been made to the standard reference aerofoil. They are: 1. Mach number ranging from 0.2 M to 0.7 M in increments of one Mach number. 2. Angle of Attacks of values 3, 5, 8 and 12. The results will be evaluated within this low speed subsonic range. So, the best aerofoil profile will show its effectiveness within these ranges of Mach number and angle of attack only. After making such changes to the eight profiles, the tabulated values are compared to find the best aerofoil profile [3]. EFFECT OF SHAPE ON AERODYNAMIC CHARACTERISTICS L/D Ratio The effect of shape on the Lift to drag ratio has been found by the analysis of aerofoil shapes for various conditions of Mach number and Angle of attack and it is tabulated as follows. Some of the tabulated results are shown hereunder. The aerodynamic coefficients are dependent on their body shape (airfoil section chosen), and also on the attitude (angle of attack), Reynolds number, Mach number, surface roughness, and air turbulence [4]. Table 1: Comparison of (L/D) Ratios for NACA 0012 and Eight Modified Aerofoil at AOA=3. Ref. 1 2 3 4 5 6 7 8 M (L/D) in % of Chord at 3 AOA 12 15 20 25 30 35 40 45 50 0.2 21.9 15.5 18.4 18.3 23.9 25.7 30.3 27.3 26.0 0.3 21.0 15.0 17.9 17.9 23.2 25.0 29.4 26.2 25.9 0.4 20.5 14.5 14.5 16.7 22.6 24.4 28.5 26.6 26.0 0.5 20.5 14.0 14.1 15.2 22.0 23.9 27.8 26.5 26.1 0.6 19.5 12.9 13.7 12.9 21.6 23.3 27.1 26.4 12.2 0.7 19.1 7.6 10.2 14.8 20.3 22.8 26.4 2.0 10.4 Examining Graph 1 for the above data (Table 1) shows that there is a good level of high (L/D) ratio for the 40% of chord design and is very much better as compared to NACA 0012, with nearly 10 ordinates distinction. The performance of the other aerofoil designs such as the 45% of chord and the 50% of chord designs are approaching the good performance levels till Mach 0.5 but steadily decrease with increase in the Mach number. Also, 30 and 35% of chord seems to bear out better than the standard NACA 0012 aerofoil at AOA=3. The aerofoil design such as the ones with a 15, 20 and 25% of chord doesn t show good results. The value of the lift coefficient is expected to be the maximum with a minimal subsequent drag coefficient. The standard NACA 0012 aerofoil has a maximum lift coefficient of 0.0194 at 0.7 Mach at AOA=8. Hence, lift to drag ratio plays an important role in the design of an aerofoil [5]. JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 3
ΔCp Aerofoil Profile Analysis and Design Optimisation Prasad et al. Graph 1: (L/D) vs. Mach number at AOA=3. Similarly, tabulated results are available for the same range of Mach numbers and angle of attacks for 5, 8 and 12. The graphs are plotted for comparing the tables easily in a sketched format. Pressure Variations The pressure variations on the aerofoil determines the lift and drag coefficients on the aerofoil for various speeds. This ultimately decides the (L/D) ratio. Analysis has been carried out in FLUENT and GAMBIT to find out the best pressure gradient of the aerofoil profile among the eight shapes. The better the pressure variation, the better will be the aerodynamic performance of the aerofoil. The pressure coefficient is the difference between the local static pressure and the free stream pressure. The aerofoil geometry and the pressure are directly related to each other. The pressure differential is now found out in this process. The changes in pressure above and below the aerofoil will change the lift values of the aerofoil. This is a major design criterion in the wing design of an aircraft. The lesser the difference in pressure the better will be the performance for a symmetric aerofoil. To prove this a comparative chart has been plotted (Graph 2). 200 180 160 140 120 100 80 60 40 20 0 0.2 0.3 0.4 0.5 0.6 0.7 Mach number NACA 0012 Design 1 Design 2 Design 3 Design 4 Design 5 Design 6 Design 7 Design 8 Graph 2: Differential Cp at AOA=8. JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 4
Journal of Aerospace Engineering and Technology Volume 3, Issue 2, ISSN: 2231-038X Also are the flow analysis results that show the variations of pressure on the aerofoil surface for all the conditions of Mach number and AOA. The data collected are for Mach number range: 0.2 to 0.7 and Angle Of Attack: 3, 5, 8, 12 whose pressure coefficient graph with the position of chord is to be found by pressure plots and vector plots as the output to study the aerofoil characteristics [6]. The co-ordinates of NACA 0012 aerofoil are taken from Design FOIL and then the changes are made for the analysis to proceed. The Figures 2, 3 and 4 below are for 15% of chord at AOA=3 and 0.2 M. The procedure is extended to all combinations of the specified Mach number and AOA and for the modified 20, 25, 30, 35, 40, 45 and 50% of chord aerofoil profiles. Fig. 2: NACA 0012 15% of chord at AOA=3, 0.2M. Fig. 3: Pressure Distribution over NACA 0012 15% of Chord at AOA=3, 0.2 M. JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 5
Aerofoil Profile Analysis and Design Optimisation Prasad et al. Fig. 4: Velocity Distribution over NACA 0012 15% of Chord at AOA=3, 0.2 M. Coefficient of Moment The lesser the moment coefficient the lesser will be the nose-over pitching moment which is not desirable during flight due to control difficulties. The values are plotted along with the Cl and Cd values to compare for the lesser Cm and better lift. Table 2: NACA 0012 30% of Chord at AOA=8. M Cl Cd Cm L/D 0.2 0.632204 0.04175328-0.01209 15.14142122 0.3 0.63336596 0.0433502-0.01324 14.6104507 0.4 0.63613842 0.0456085-0.015644 13.94780403 0.5 0.64872602 0.05077-0.0186 12.77774316 0.6 0.66674428 0.05699-0.0283475 11.69932058 0.7 0.7633115 0.09304-0.02133 8.204121883 And so, the values are compared to find the best aerofoil profile with a lesser Cm. The other values are also noted for the other modified designs as shown in Table 2. MODAL ANALYSIS OF PROFILES Modal analysis of an aerofoil is carried out to find the natural characteristics or the mode shapes and vibration frequencies that are responsible for the vibratory oscillations of the aerofoil. Here we have chosen the self-weight condition of load only. No extra loads are added to its self-weight. This is mainly to identify the flutter characteristic of aerofoil profiles and the means of noise related problems during flight. JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 6
Journal of Aerospace Engineering and Technology Volume 3, Issue 2, ISSN: 2231-038X Also, the overall structural performance of the aerofoil can be estimated by knowing the modes of failure of the system due to flutter and other problems caused by the aerofoil shape or profile [7]. Fig. 5: NACA 0012 1 ST SET 40% of Chord. Fig. 6: NACA 0012 2 ND SET 40% of Chord. JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 7
Aerofoil Profile Analysis and Design Optimisation Prasad et al. Fig. 7: NACA 0012 3 RD SET 40% of Chord. Here, the Figures 5, 6 and 7 are corresponding to the modal analysis results for 40% of chord profile. It is noted that the upper surface seems to be less intensely stressed, but the lower surface has been in complete tension and bending has resulted at 3/4ths of the chord. Also the bending moment ranges from 0.09 to 0.11 N/m². The results for the other profiles are also analyzed for the number of modes of failure. Table 3 below shows that the 15% of Chord design has been having the least natural frequency of vibration as that of NACA 0012 for its self weight loading only. If loads are applied, the results may vary. Thus, Flutter effects will be the least for Design 1 aerofoil profile and is better than other profiles. The resonance in the aerofoil can be found and the time and load of this failure at the ultimate limit can be estimated. Table 3: Natural Frequencies in Hertz for NACA 0012 Aerofoil with Maximum Thickness at Various Locations. MODE % of Chord Ref. 15 20 25 30 35 40 45 50 1 18 18 19 19 19 20 19 19 19 2 94 89 92 96 96 97 97 99 101 3 115 124 129 133 133 134 134 130 128 JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 8
Natural frequency Journal of Aerospace Engineering and Technology Volume 3, Issue 2, ISSN: 2231-038X 130 110 90 70 50 30 NACA 0012 15% OF CHORD 20% OF CHORD 25% OF CHORD 30% OF CHORD 35% OF CHORD 40% OF CHORD 45% OF CHORD 50% OF CHORD 10 1 2 3 Modes Graph 3: Natural Frequencies in Hertz for NACA 0012 Aerofoil with Maximum Thickness at Various Locations. For easy verification, Graph 3 is also plotted for the same data in Table 3. 3-D ANALYSIS OF AN AEROFOIL The 3D model that has proved best in the analysis done so far has to be designed in ANSYS. The wing model when analyzed would yield the pressure variations along the length of the wing and so it is useful for the mentioning of pressure on the wing surface instead of just on the aerofoil surface (Figures 8 and 9). Fig. 8: Cp Variation for Design 6 - NACA 0012 40%C Wing at AOA=5, M=0.3. JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 9
Aerofoil Profile Analysis and Design Optimisation Prasad et al. Fig. 9: Pressure Plot for Design 6 - NACA 0012 40%C Wing at AOA=5, M=0.3. CONCLUSIONS The best aerofoil design for Cl and Cd at 3 AOA is Design 6 (40% of Chord); at 5 AOA is Design 5 (35% of Chord); at 8 AOA is Design 5 (35% of Chord); at 12 AOA is Design 1 (15% of Chord). The best aerofoil design for (L/D) ratio at 3 AOA is Design 6 (40% of Chord); at 5 AOA is Design 6 (40% of Chord); at 8 AOA is Design 5 (35% of Chord); at 12 AOA is Design 2 (20% of Chord). These results are proved better than the standard NACA 0012 aerofoil. The Design 6 with a 40% of Chord has been the best suitable aerofoil for the use in turbines and compressors as it has the least Cm, High (L/D) ratio and a better pressure gradient. REFERENCES 1. Alan Pope, Basic Wing & Airfoil Theory, McGraw-Hill Book Company: New York, 1951. 2. National Advisory Committee for Aeronautics: Aerodynamic Characteristics of Aerofoils, NACA Rep. 93, 1920. 3. Abbott, Ira H.; Von Doenhoff, Albert E. Theory of Wing Sections. Dover Publ., Inc., 1959. 4. Somers, Dan M.: Effects of Airfoil Thickness and Maximum Lift Coefficient on Roughness Sensitivity. Airfoils, Inc., State College, Pennsylvania, 1998. 5. John J.Bertin, Aerodynamics for Engineers, 4 th edition, Prentice Hall: Upper Saddle River, New Jersey, 2009. 6. Anderson, J, Fundamentals of Aerodynamics, 2nd Edition, McGraw-Hill: New York, 1991. 7. M.L.J. Verhees DCT 2004.120, Experimental Modal Analysis Of A Turbine Blade, Technische Universiteit Eindhoven, Department Mechanical Engineering, Dynamics and Control Group, Eindhoven, December, 2004. JoAET (2013) 1-10 STM Journals 2013. All Rights Reserved Page 10