Aerospace Engineering projects in aerodynamics, finite element analysis, math modelling, and database modelling

Editor's Notes

• #4: function NACA_Clmax() Clmax = [1 1.8 0.9;2 2.1 1.3;3 1.8 0.8;4 1.8 0.8;5 1.9 1.0;6 2.4 1.6;7 2.3 1.4;8 2.5 1.6;... 9 2.1 1.3;10 2.3 1.5;11 2.5 1.6;12 2.3 1.5;13 2.5 1.7;14 1.7 1.2;15 1.6 1.2;16 1.5 1.1;17 1.5 0.9;... 18 1.3 0.8;19 2.7 1.7;20 2.8 1.6;21 2.7 1.5;22 2.7 1.5;23 2.7 1.4;24 1.8 1.2;25 1.7 1.2;26 1.6 1.1;... 27 1.5 0.9;28 1.4 0.8;29 1.8 0.9;30 1.8 1.2;31 1.9 1.1;32 2.1 1.4;33 2.1 1.5;34 2.2 1.5;35 2.4 1.2;... 36 2.5 1.3;37 2.4 1.1;38 2.6 1.3;39 2.7 1.3;40 2.8 1.4;41 2.7 0.9;42 2.8 0.9;43 2.8 1.0;44 2.9 0.8;... 45 2.7 0.7;46 2.7 0.6;47 2.7 0.7;48 1.8 0.6;49 2 1.0;50 1.8 0.4;51 1.8 0.7;52 1.8 0.7;53 2 0.8;... 54 1.8 0.8;55 1.9 0.8;56 2 0.9;57 2 0.9;58 2.2 0.9;59 2.3 1.0;60 2.4 1.0;61 2.5 1.2;62 2.4 0.9;... 63 2.6 1.0;64 2.7 0.9;65 2.7 1.0;66 2.7 0.8;67 2.8 0.9;68 2.8 0.9;69 2.8 0.7;70 2.7 0.8;71 2.8 0.7;... 72 1.9 0.7;73 2.2 0.8;74 2.4 1.0;75 2.4 0.9;76 2.6 1.2;77 1.8 0.9;78 1.8 1.1;79 1.8 1.0;80 2 1.0;... 81 2.1 1.0;82 2.2 1.2;83 2.1 1.0;84 2.3 1.1;85 2.3 1.0;86 2.4 1.3;87 2.5 1.0;88 2.5 1.1;89 2.6 1.2;... 90 1.6 1.4;91 2.6 1.1;92 2.7 1.1;93 2.7 1.2;94 1.5 1.2;95 2.6 1.3;96 1.5 1.2;97 2.7 1.1;98 2.7 1.2;... 99 2.7 1.2;100 1.4 1.1;101 1.8 0.8;102 1.8 1.1;103 1.8 1.0;104 2 1.2;105 2 1.0;106 2.1 0.9;107 2.2 1.0;... 108 2.4 1.0;109 2.5 1.1;110 2.6 1.2;111 2.6 1.1;112 2.7 1.1;113 2.6 1.3;114 2.5 1.0;115 2.5 1.2;116 1.5 1.0;... 117 2.6 1.1;118 2.7 1.2]; selc = zeros(length(Clmax),1); GAV_Clmax_landing = 2.42; GAV_Clmax_clean = 1.47; for n = 1:118 if (Clmax(n,2) >= (GAV_Clmax_landing))&&(Clmax(n,2) <= (GAV_Clmax_landing+0.3))&&... (Clmax(n,3) >= (GAV_Clmax_clean))&&(Clmax(n,3) <= (GAV_Clmax_clean+0.3)) selc(n)=Clmax(n,1); end end %Write all the data to a file. dlmwrite('MAE478_GAV_Airfoil.xls',selc, ';'); warning off MATLAB:xlswrite:AddSheet;
• #5: %@Ct: c_tip %@Cr: c_root %@b: wing span %@c_bar: mean root wing cord %@t: gap or metal thickness between telescoping sections %@n: number of telescoping sections %@aoa_L0: the zero lift angle of attack, ie., airfoil 4412: aoa_L0 = -4 deg %@aoa_Clmax: the angle of attack at Clmax, ie., airfoil 4412: aoa_Clmax(clean) = 14 deg %@theta_rad: the interval of angle under investigation of a wing, 0<=theta_rad<=pi %@m: the number of A coefficients %@pts: the number of points to draw on a curve of C_D,i and CL vs. aoa %Current %config:Ct=1.27;b=10.3;t=0.002;n=4;aoa_L0=-4;aoa_Clmax=14;theta_rad=pi/2;m=20;pts=20;
• #6: *Since a small amount of lift is carried on other parts of the airplane, such as the fuselage, nacelles, and horizontal tail, the value of the airplane lift curve slope is slightly higher than the wing lift curve slope. *An analysis of all these parameters is made in the preliminary design phase of the program, with the objective of having the angle of attack of the fuselage reference plane, or floor line the passenger cabin, between 0 and 2 degrees at the cruise condition. *The maximum lift coefficient in the cruise configuration is dependent on two primary parameters, the spanwise variation of local wing section lift coefficients, as the wing approaches the angle of attack for stall, and the wing airfoil section maximum lift coefficients, which are unique values for each airfoil section. *For performance reasons, airfoil sections with high values of maximum lift coefficient are usually selected to achieve the highest value of airplane maximum lift coefficient. *However, as shown in fig.11.2, the airfoil section maximum lift coefficient values are usually varied across the span, so that the wing spanwise lift distribution will reach values of the airfoil maximum lift coefficients over the inboard portion of the wing, producing local stalling inboard while maintaining some lift coefficient margin to stall over the outer portion of the wing. *This margin in lift coefficient to stall is used to protect against initial local stalling over the outer portion of the wing, which leads to severe roll off and loss of aileron control at the stall. *For this reason, the airplane maximum lift coefficient can never be as high as the airfoil maximum lift coefficients.
• #7: Single slotted flaps are standard for personal/utility aircraft. Single slotted flap chords are usually in the range of 25% and 30% of the chord, and have a max deflection of 35 degrees.
• #11: Ref: Loftin’s
• #12: SSLW: second segment limited weight
• #16: Characteristic lines represent left-and right-running mach waves Characteristic mesh consists of intersecting mach waves crossing the flow field. Local mach angle (mu) is a function of mach number, which is function of x & y
• #18: Characteristic lines represent left-and right-running mach waves Characteristic mesh consists of points of intersection of mach waves crossing the flow field. Local mach angle (mu) is a function of mach number, which is function of x & y
• #21: We have obtained an accurate graphical result as shown above. Further optimization will be done at the analysis to the solution of the boundary points.
• #34: *The maximum lift coefficient in the cruise configuration is dependent on two primary parameters, the spanwise variation of local wing section lift coefficients, as the wing approaches the angle of attack for stall, and the wing airfoil section maximum lift coefficients, which are unique values for each airfoil section. *For performance reasons, airfoil sections with high values of maximum lift coefficient are usually selected to achieve the highest value of airplane maximum lift coefficient. *However, as shown in fig.11.2, the airfoil section maximum lift coefficient values are usually varied across the span, so that the wing spanwise lift distribution will reach values of the airfoil maximum lift coefficients over the inboard portion of the wing, producing local stalling inboard while maintaining some lift cofficient margin to stall over the outer portion of the wing. *This margin in lift coefficient to stall is used to protect against initial local stalling over the outer portion of the wing, which leads to severe roll off and loss of aileron control at the stall. *For this reason, the airplane maximum lift coefficient can never be as high as the airfoil maximum lift coefficients. *This margin in lift coefficient to stall is used to protect against initial local stalling over the outer portion of the wing, which leads to severe roll off and loss of aileron control at the stall. *For this reason, the airplane maximum lift coefficient can never be as high as the airfoil maximum lift coefficients.
• #35: D/q has the units of area, and may also be written as f, the equivalent flat plate area (or equivalent parasite drag area). Most components (wing, tail, fuselage, etc.) area aerodynamically smooth, so the drag is mostly due to skin friction, and drag is calculated using skin friction coefficient. Some components, such as deployed landing gear and flaps, and flap hinges are bluff, so drag is mostly due to separation. For these cases, drag is often calculated based directly on the equivalent flat plate area f, and skin friction drag is not calculated. Drag of bluff components may be stated as f/(unit area), in which case this coefficient must be multiplied by the component frontal area (normal to the flow) to obtain the value of f. The skin friction coeff. decreases with increasing RN.
• #39: The higher of the two load factors must be used for structural design.