P A commercial aircraft (with a mass of 71411 kg, wing surface area of 119 m² and wing aspect ratio of 10) produces a maximum of 209459 newtons of thrust. Assuming CLMAX = 2.1 and CDCalculate the Climb Rate (in feet/min) at an initial climb airspeed of 169 knots.-89O 11.793570.55171697.812320.13= 0.06,

The objective of the question is to calculate the climb rate of a commercial aircraft given its…

Answered: A commercial aircraft (with a mass of…

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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#1. Consider a Boeing 747-700 in straight-and-level cruise flight on a standard day. The aircraft has a wing area of 525 m and is flying geometric altitude of he = 10,000 m. At a certain time during the flight, its mass is 3.5 x 105 kg and each of the four engines produces 50 kN of thrust. C was calculated as 0.4. a) Determine the cruise airspeed. b) Determine the total drag coefficient.
A light aircraft is flying at a speed of 60 m/s in ISA conditions for an altitude of 2000 m. The aircraft has a mass of 750 kg and a rectangular planform (zero sweep) wing of 15 m? area, 9 m span and 160 mm thickness. Due to the fuselage, 85% of the wing is exposed to the air. A 20% increment should be applied to the friction factor to account for wing surface roughness, u= 1.750 x10 kg/ms. Calculate the skin friction drag coefficient for the wing. Keep 5 decimals.
The test results of a NACA 2518 shows: Angle of attack 0 deg 8 deg CL 0.045 0.75 If the airfoil is used to construct a wing with a span of 60 ft and a wing area of 450 square feet. Determine the wing lift curve slope. (e = 0.88)
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
An aircraft is equipped with a wing of symmetrical airfoils. The lift curve slope of the total aircraft is estimated to be 0CL/Oa = 5.0265 1/rad. The stall angle of attack (AOA) is 12°. Wing area is 16 m2. Use g = g0 and r = rO. What is the aircraft's mass during a flight on which a stall speed of 50 kt was observed?
If the pressure difference in the wind tunnel experiment is 50 mmH,O, the density of air is 1.2 3 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=28.9 N Select one: a. 2.94715 b. 11.61715 c. 5.83715 d. 14.50715 e. 8.72715
The drag polar for the ASW 27B high performance sailplane is: CD = 0.01 +0.02CL2. What is its maximum lift-to-drag (L/D) ratio? 18 25 35 70
The rectangular wings of a microlight aircraft have a combined spanwise length of 9.144 m and streamwise (direction of air flow) width of 0.9144 m The wings have a coefficient of form drag of 0.01 and a lift coefficient of 0.3 in cruise. When cruising at a speed of 150km/h and altitude of 3,000 feet, the temperature and absolute pressure of the atmospheric air are 5ºC and 90 kPa, respectively. a) Calculate the density of the air at cruising altitude. b) Calculate the form drag on the wings. c) Calculate the Reynolds number of the wings. d) Calculate the skin friction drag on the wings. e) If the fuselage of the microlight is spherical with a diameter of 1m and the total drag of the aircraft is represented by the contributio combined, and assuming the skin friction drag is negligible for the fuselage only, calculate the combined total drag force. fuselage and wings f) Calculate the power required to overcome total drag in cruise, using the value calculated in part e). g) Calculate the…
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
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.
8.) A circular cylinder 4m long and 3m in diameter is in an air stream. The flow velocity is 5 m/s L parallel to the axis of the cylinder. Given that the coefficient of drag, CD on the cylinder is 1.3, and the density of air is 1.225 kg/m3. What is the force on the cylinder? A. 0.090 KN B. 0.14 KN C. 0.24 KN D. 0.91 kN
31) Estimate the drag force on the fuselage shown below for a cruising speed of 200 m/s at 10,000 m. 8 m 5 m W m W = 20 m Hint 1: To calculate the drag force split the fuselage into 4 parts: front hemisphere, cylindrical body, vertical stabilizer, back hemisphere. Model the front and back hemispheres as flow over a sphere. For simplicity treat the cylindrical body and vertical stabilizer as flat plates Hint 2: Use Ca vs Reynolds number graphs for sphere and flat plate. If your Reynolds number is greater/smaller than the Cd vs Reynolds graph range, you can instead use the greatest/smallest number available on the graph.

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