Consider an NACA 23012 airfoil. The mean camber line for this airfoil is given by= 2.65950.6075A +0.1147for 0<< 0.2025= 0.02208 (1–) for 0.2025 << 1.0Calculate (a) the angle of attack at zero lift, (b) the lift coefficient when a = 4°, (c) themoment coefficient about the quarter chord, and (d) the location of the center of pressurein terms of xep/c, when a = 4°. Compare the results with experimental data.and

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Answered: Consider an NACA 23012 airfoil. The…

Elements Of Electromagnetics

7th Edition

ISBN:9780190698614

Author:Sadiku, Matthew N. O.

Publisher:Sadiku, Matthew N. O.

ChapterMA: Math Assessment

Section: Chapter Questions

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A straight-tapered flying-wing aircraft has the following characteristics: span, 20 ft; area, 100 ft2 ; root chord, 6 ft; tip chord, 4 ft; leading-edge sweep angle, 40 deg; sweep angle of line of maximum thickness, 24 deg; and Airfoil NACA 2412. Determine the wing's lift-curve slope and induced drag k value. Assume the flight Mach number is 0.25 in standard sea-level conditions.
The lift curve for the 4 digit NACA 2421 airfoil is shown in the Figure. Consider a wing with AR = 8, sweep ? = 15 ???, and airplane flies at the Mach number equal to 0.3. The Oswald efficiency span number ?1 = 0.95 Calculate the lift slope for the finite wing. Clearly show the formula and explanations in the solution.
Consider the same Lockheed F-104 supersonic fighter shown, with the same flight conditions of Mach 2 at an altitude of 11 km. For these conditions the wing angle of attack is α = 0.035 rad = 1.98◦. Assume the chord length of the airfoil is 2.2 m, which is approximately the mean chord length for the wing. Also, assume fully turbulent flow over the airfoil. Calculate: (a) the airfoil skin friction drag coefficient, and (b) the airfoil wave-drag coefficient. Compare the two values of drag.
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.
A Savonius rotor can be approximatedby the two open half-tubes in Fig mounted ona central axis. If the drag of each tube is similar to thatin Table 7.2, derive an approximate formula for therotation rate Ω, as a function of U, D, L, and the fluidproperties (ρ, μ).
A NACA 2412 airfoil with a 4.2 ft chord and 7.5 ft span, which has a rectangular planform, isbeing tested in a wind tunnel at a standard sea level conditions at test-section velocity 70 knotsat an angle of attack of 10 deg.A) What is the airfoil’s maximum thickness, maximum camber, and its location?B) Does this airfoil have negative or positive camber? Why?C) Calculate the operating Reynolds number, Re? Determine whether the flow around the airfoilis laminar or turbulent?D) What are the values of ??, ??, and ?? ?.?. at the above conditions?E) How much lift, drag, and pitching moment about the aerodynamics center are the airfoilgenerating?
A flying wing weighing 20,000 N has a NACA 65‐418 airfoil section, an area of 30 m2, and an aspect ratio of 5.0. Determine the angle of attack for a sea‐level flight at 60 m/s. What is the lift‐to drag ratio? Assume that the nominal airfoil lift curve slope a0 = 0.106/deg̊, the profile drag CD0 = 0.0043 and e = 0.92.
The measured lift slope for the NACA 23012 airfoil is 0.1080 degree−1,and αL=0 = −1.3◦. Consider a finite wing using this airfoil, with AR = 8and taper ratio = 0.8. Assume that δ = τ . Calculate the lift and induceddrag coefficients for this wing at a geometric angle of attack = 7◦.
Air flows at Ma = 2.5 past a half-wedge airfoil whoseangles are 4°, as in Fig. Compute the lift and dragcoefficient at α equal to (a) 0° and (b) 6°.
Determine the velocity required for take off for an aircraft with the following characteristics.Aircraft mass: 7200 kgWingspan: 15 mWing shape: rectangularWing chord length: 1.8 mWing airfoil: NACA 4415Wing angle of attack: 7× 0.5 + 5 degreesHint: Check NACA charts for required parameters.
Consider turbulent fl ow past a sharp, smooth fl at plate ofwidth b and length L. What percentage of the friction dragon the plate is carried by the rear half of the plate?
a)What is the impact of increasing Reynolds number on skin friction and pressure drags over an airfoil? What can be happened for separation in this case? b) What is an adverse pressure gradient and where does it occur on an airfoil (show that on a sketch)? c) Why lift-to-drag ratio is an important parameter for an aircraft? d)How can changing in altitude affect the aircraft power required, PR? Show thatmathematically and graphically?

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