Please show all work!! Using thin-airfoil theory, design an airfoil camber line that meets the following three requirements:⚫ The lift coefficient is c₁ = 0.5⚫ There is no leading-edge suction peak• The airfoil has 10% static stability, SM = 10%a) Based on the requirements, determine the thin-airfoil theory constants B1 and B2. b) Assume that the camber line is given byZe=a= a²² (1-2²) (1+6²)сDetermine the constants a and b, and the angle of attack a, such that the design requirementsare met. Provide a computer-generated plot of the camber line.c) Plot the difference in pressure coefficient versus x/c,Acp(x/c) = Cp,lower (x/c) - Cp,upper (x/c),and verify numerically, using point values of Acp and trapezoidal integration (trapz) thatce ==1A(z) dz

a) Determining Thin-Airfoil ConstantsB1 and B2 In thin-airfoil theory, the lift coefficient cℓ…

Answered: Using thin-airfoil theory, design an…

Elements Of Electromagnetics

7th Edition

ISBN:9780190698614

Author:Sadiku, Matthew N. O.

Publisher:Sadiku, Matthew N. O.

ChapterMA: Math Assessment

Section: Chapter Questions

Problem 1.1MA

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2. The table below shows experimental data for the shape and pressure distribution on the upper and lower surface of an airfoil at zero angle of attack. X 0 0.25 0.5 0.75 1 Y Fx = Yu 0 0.0952 0.0922 0.0588 0 (Pu dYu dx - Calculate the drag and lift force on the airfoil by numerically evaluating the integrals Fy (P₁-P₂)dx = √ ( P₁ - 1 Pi Y₁ 0 -0.0254 -0.0144 -0.0052 0 dv) Y = Yu (x) dx y = Ye (x) Pu 1.000 -1.640 -0.786 -0.212 0 x Pi 1.000 0.589 0.426 0.322 0 Use finite difference approximations of the derivatives and the composite Simpson's rule to evaluate the integrals. Hint: Use central differences to estimate the derivatives wherever possible since they are more accurate than forward or backward differences.
= 3570 T AC =144 Example: 1.5 ft It is desired to determine the drag force at a given speed on a prototype sailboat hull. A model is placed in a test channel and three cables are used to align its bow on the channel centerline. For a given speed, the tension is 40 lb in cable AB and 60 lb in cable AE. Flow 4 ft Determine the drag force exerted on the hull and the tension in cable AC. Choc ug the h- as the free hod- DApress the
2. The table below shows experimental data for the shape and pressure distribution on the upper and lower surface of an airfoil at zero angle of attack. X 0 0.25 0.5 0.75 1 Yu 0 0.0952 0.0922 0.0588 0 Fx = dYu dx 5 Y₁ 0 -0.0254 -0.0144 -0.0052 0 PL Pd) d: dx Y = Yu (x) Calculate the drag and lift force on the airfoil by numerically evaluating the integrals = S₁² ( P₂₁ = ['(P₁ - F (P₁ - Pu)dx 0 y = Y₂ (x) Pu 1.000 -1.640 -0.786 -0.212 0 Fy x Pi 1.000 0.589 0.426 0.322 0 Use finite difference approximations of the derivatives and the composite Simpson's rule to evaluate the integrals. Hint: Use central differences to estimate the derivatives wherever possible since they are more accurate than forward or backward differences.
A conventional low speed aircraft had the following aerodynamic characteristics: CDo=0.020. e=1.0 (Oswald efficiency). The aircraft was flown to maintain a steady level flight and for minimum thrust required, at a lift coefficient of Cz=0.8. The numerical value of the aspect ratio of the wing is three decimal places). (in
Aerodynamic Model: - Circular Disk Inside Dia.= 40 mm Outside Dia.= 80 mm Angle of Attack= 15 Degree Lift Force= 0.38 N Drag Force=0.40 N Wind Velocity= 20 m/sec Air Temperature= 30 oC (Note: Take Air Density at 30 oC) Required: - Find Lift and Drag Coefficient at 15 Degree Attack Angle.
If the pressure difference in the wind tunnel experiment is 50 mmH,0, the density of air is 1.2 kg/m in wind tunnel experiment, and the airfoil span is 204 mm, and the cord length is 98 mm, what is the lift coefficient if the lift force is L=13.4 N
A NACA 2412 airfoil with a chord of 0.64m is flying in an airstream of standard sea level conditions. The freestream velocity is 70 m/s. Given the lift per unit span is 1,254 N/m. By using the experimental data for NACA 2412 data plot in Figure Q1c, investigatethe angle of attack of the airfoil and the analyze the value of drag per unit spanof the airfoil. Given that at standard sea level, ?=1.789×10-5 kg/m.s.
Consider the following NACA airfoils: 2412, 23012, 23021 a. For each of these airfoils, find the L/D ratio for angle of attack of 0, 4, 8 and 12 b. For each airfoil find the angle of maximum L/D and find the stall angle c. Find which airfoil will provide the maximum lift for any given geometry and flow conditions
Correctly solve it.
1. An airplane weighs 36,000 lb. and has a wing area of 450 fte. The drag equation is C, =0.014 + 0.05C} . It is desired to equip this airplane with turboprop engines with available power such that a maximum speed of 602.6 mph at sea level can be reached. The available power is assumed to be independent of flight speed. Calculate the maximum rate of climb and the speed at which it occurs. Given: W = 36,000 lb S = 450 ft² C, = 0.014+0.05C V = 602.6 mph max THPy = cons tan t THP AV constant Max EHP Point of THP, REQD. Flisht Speed. V Vmax 602.6 mph Required: Max R.C. and Vmax R.C. Horsepower, hp
A closed-loop wind tunnel has a test section of cross-section W =3 ft by H =5 ft and length L =4 m. An aerofoil with 21% thickness to chord ratio and a chord length of 0.52 m is mounted vertically in the test section and spans the entire height. If the measured lift and drag coefficients are C; = 0.325 and Ca =0.0071, what is the lift coefficient after correcting for solid blockage, wake blockage and streamline curvature? Give your answer to 3 decimal places.
4. Write down the lift curve equation (lift coefficient versus angle of attack) according to the thin airfoil theory. What is the amount of the lift curve slope for a thin airfoil ? NACA 2412 airfoil Eg. (4,57) C 2.0 - 1.e 1.2 Lift coefficient 0.8 0.4 -0.4 Eg. (4,64) -0.1 Moment coefficient -0.8 -0.2 -1.2 -0.3 Re - 3.1 X 10 O Re- 8.9 x 10 -0.4 16 a, degrees de; = ? da

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