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.
4
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.
Please show all work!!
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.
Please show all work
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.
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 °C (Note: Take Air Density at 30 °C) Required: - Find Lift and Drag Coefficient at 15 Degree Attack Angle.
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
Qustion. An aircraft is flying in level flight at a speed of 200 km/h through air (density, p = 1.2 kg/m³, and viscosity, μ = 1.6 x 10-5 N-s/m²). The lift coefficient at this speed is 0.4 and the drag coefficient is 0.0065. The mass of the aircraft is 800 kg. The effective lift area of the aircraft NOTE= Do not give wrong answer I will dislike , If not sure skip it
An airplane with a NACA 23012 airfoil cruises at 150 m/s at an altitude of 6000 m. The airfoil has an aspect ratio of 10 with a span of 36 m. Using the airfoil data as in Fig. 3, determine the lift and drag forces. Then determine the power required to overcome drag. Consider the airplane flying at an angle of attack equal to 2.

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