Please show all work A thin airfoil has the following Acp = Cp,lowerCp,upper distribution at a = 0°:ΔεAcp = 2² (1-2)2층(1층)a) Compute the lift coefficient (ce), center of pressure (Tcp), aerodynamic center (ac), andstability margin (SM) of the airfoil at a = 0°.b) Determine the formula for Acp(x/c) at an arbitrary a and provide a computer generated plotof Acp(x/c) for a = 5°.c) Determine (numerically is acceptable) a camber line ze(x/c) that is consistent with the givenpressure distribution, and provide a computer-generated plot of the camber line. Also presentthe Bn values up to n = 10.

Answered: A thin airfoil has the following Acp =…

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.
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.
Q1: Find the boundary layer thickness (6) equation, the shear stress (to) and the coefficient of drag (C₁) if the velocity distribution in the laminar boundary layer over the face of a spillway was observed to be: u 7- (C)-()*+C)*. U Then calculate the boundary layer thickness and drag force if the air flows over a sharp edged flat plate 0.25m long and 0.5m wide at a velocity 1 m/s, take the air density 1.23 kg/m³ and the kinematic viscosity is 1.46*10-5 m/s².
Q1: Find the boundary layer thickness (6) equation, the shear stress (to) and the coefficient of drag (Cò ) if the velocity distribution in the laminar boundary layer over the face of a spillway was observed to be: *--(C)- ()*+C). U Then calculate the boundary layer thickness and drag force if the air flows over a sharp edged flat plate 0.25m long and 0.5m wide at a velocity 1 m/s, take the air density 1.23 kg/m³ and the kinematic viscosity is 1.46*10-5 m/s².
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
= 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
the wing of an aircraft haa a taper ratio of 0.63, and mean aerodynamic chord of 2.65ft. find the measurement of the wing's root chord, tip chord and wing area if the wingspan is 30 ft
solve show all steps and assumptions. answer all parts a-d
The following equation may be used to estimate the take-off ground run for an aircraft: Equation has been attached as an image. Calculate the take-off ground run, from a runway at ISA-SL conditions, for a twin engine aircraft for which the following data may be assumed Aircraft lift-off speed 155 knots Max take-off gross weight 220 tonnes Wing planform area (S) 358 m Wing CL (t/o flaps deployed, a = 0) 1.1 Wing span 53.18 m Oswald efficiency factor, e 0.7 KGE = CD(IGE) / Co(OGE) 0.4 Co sum (fuselage, wing, tailplane and nacelle) 0.015 Co for undercarriage 0.021 Co for flaps at taking-off setting 0.0073 Coefficient of rolling friction, u 0.02 Engine thrust (assumed constant) 310 kN per engine It may be assumed that 1knot = 0.51444 m/s It may be assumed that 1knot = 0.51444 m/s
Comment and explain the L/D ratio for the two shapes
reply me within 40 min
1. The given information is as follows: Cd=1.4 (coefficient of drag) ¥=10 ft/sec (velocity) Ƴ=62.4 #/ft3 g=32.3 ft/sec2 a. What is the water pressure in PSF against the obstruction? b. If the pier is 6 ft wide and 14 ft deep, what is the Resultant Pressure on the WHOLE obstruction? c. If the obstruction is fixed into the bottom of the river (mud line), what is the bending moment at the mudline on the obstruction?

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