Consider a 117 m/s airflow over a flat plate that is 1.2 m long perpendicular to the freestream direction. The freestream conditions correspond to standard sea level values. Assuming turbulent flow all throughout, calculate the total skin friction drag if the turbulent boundary layer has a thickness of 1.37 x 10-3 m at the trailing edge. Provide answer in Newtons, and round it off to the nearest thousandths.

Consider a 117 m/s airflow over a flat plate that is 1.2 m long perpendicular to the freestream direction. The freestream conditions correspond to standard sea level values. Assuming turbulent flow all throughout, calculate the total skin friction drag if the turbulent boundary layer has a thickness of 1.37 x 10-3 m at the trailing edge. Provide answer in Newtons, and round it off to the nearest thousandths.

Principles of Heat Transfer (Activate Learning with these NEW titles from Engineering!)

8th Edition

ISBN:9781305387102

Author:Kreith, Frank; Manglik, Raj M.

Publisher:Kreith, Frank; Manglik, Raj M.

Chapter5: Analysis Of Convection Heat Transfer

Section: Chapter Questions

Problem 5.49P

See similar textbooks

Similar questions

Consider a 102 mph airflow over a flat plate with a span of 9.7 ft and a planform area of 60 ft2. The boundary layer transitioned from laminar to turbulent. If the total skin friction drag for the laminar section is 0.858 lbs, determine the location of the transition point. Assume SSLC, Provide answer in feet and round off answer to the nearest thousandths.
Consider an infinite wing with a NACA 1412 airfoil section and a chord length of 3 ft. The wing is at an angle of attack of 5° in an airflow velocity of 100 ft/s at standard sea-level conditions. Calculate the Reynold’s number, lift, drag, and moment about the quarter-chord per unit span. Cl = 0.67, Cmc/4 = -0.025 and Cd = 0.007Detailed solution pls
Assume air at 20°C and atmospheric pressure on a smooth flat plate with a speed free current of 25 m/s (kinematic viscosity of 1.5x10-5 m2/s and density 1.2 kg/m3). The plate is 4 m long and 1.5 m wide. If the boundary layer is initially laminate and then become turbulent boundary layer with a transition Re of 5x10^5: a) What is the drag coefficient for the plate?b) What is the drag force on the entire plate?c) At what distance x from the plate the profile becomes turbulent.d) What is drag force on the turbulent part?e) What is the drag force on the laminar part?
From Table the drag coefficient of a wide plate normalto a stream is approximately 2.0. Let the stream conditionsbe U∞ and p∞. If the average pressure on the front of theplate is approximately equal to the free-stream stagnationpressure, what is the average pressure on the rear?
For flow over a smooth plate, what approximately is the maximum length of the boundary layer if V_o = 9.0 m/s in the irrational uniform flow and the fluid is air? Water?
The airfoil section of the wing of the British Spitfire of World War II fame is an NACA 2213 at the wing root, tapering to an NACA 2205 at the wing tip. The root chord is 8.33 ft. The measured profile drag coefficient of the NACA 2213 airfoil is 0.006 at a Reynolds number of 9 × 106. Consider the Spitfire cruising at an altitude of 19000 ft. Assume that μ varies as the square root of temperature. At this velocity and altitude, assuming completely turbulent flow, estimate the skin-friction drag coefficient for the NACA 2213 airfoil, and compare this with the total profile drag coefficient. Calculate the percentage of the profile drag coefficient that is due to pressure drag. (Round the final answer to three decimal places.) The skin-friction drag coefficient for the NACA 2213 airfoil is .
A wing of 2 percent camber, 5-in chord, and 30-in span istested at a certain angle of attack in a wind tunnel with sealevelstandard air at 200 ft/s and is found to have lift of 30 lbfand drag of 1.5 lbf. Estimate from wing theory (a) the angle ofattack, (b) the minimum drag of the wing and the angle of attackat which it occurs, and (c) the maximum lift-to-drag ratio.
Read the question carefully and give me right solutions with clear calculations. The air flows past a 1m diameter disk oriented normal to upstream flow at 6.8m/s. The pressure distribution on front of the disk is function of the radius, where p(r) = (25 − 100 r2 ) N/m2 , and the pressure acts uniformly on back side with suction pressure of pback = −22 P a. a. Determine the total drag force on the disk. b. Calculate the drag coefficient for the disk . (Density = 1.25 kg/m3 ).
Consider an NACA 2412 airfoil with a 2-m chord in an airstream with avelocity of 50 m/s at standard sea level conditions. If the lift per unit spanis 1353 N/m, what is the angle of attack?
Derive an expression for turbulent boundary layer thickness over a flat plate
Explain briefly in your own words (parts a-d)a) Why doesn’t the surface roughness affect the pressure drop in a laminar pipe flow? b) For bluff bodies, the drag coefficient decreases abruptly when the flow becomes turbulent. Why? c) Consider a spinning ball (in the clockwise direction) moving with a velocity U. Explain how the lift force is formed on the ball and show its direction. Uwd) For flow over streamlined bodies, drag force usually increases when the flow becomes turbulent.Why? e) Consider a pipe with a radius R and length L (extending from (1) to (2)). Inside the pipe there is a solid cylinder with a radius k×R (k  1). The planes (1) and (2), separated by a distance L have pressures P1 and P2 respectively, with P1 > P2. At t = 0, the inner solid cylinder suddenly begins to be pulled at a constant speed U in the negative z direction. Thefluid is an incompressible Newtonian liquid with dynamic viscosity μ and density ρ. In orderto find the solution for the velocity field in…
Consider the NACA 2412 airfoil, data for which is given in 4.10 and 4.11. The data are given for two values of the Reynolds number based on chord length. For the case where Rec = 3.1×106, estimate: (a) the laminar boundary layer thickness at the trailing edge for a chord length of 1.5 m and (b) the net laminar skin-friction drag coefficient for the airfoil.