You are asked by your college crew to estimate the skin friction drag on their eight-seat racing shell. The hull of the shell may be approximated as half a circular cylinder with 457 mm diameter and 7.32 m length. The speed of the shell through the water is 6.71 m/s. Estimate the location of transition from laminar to turbulent flow in the boundary layer on the hull of the shell. Calculate the thickness of the turbulent boundary layer at the rear of the hull. Determine the total skin friction drag on the hull under the given conditions.
You are asked by your college crew to estimate the skin friction drag on their eight-seat racing shell. The hull of the shell may be approximated as half a circular cylinder with 457 mm diameter and 7.32 m length. The speed of the shell through the water is 6.71 m/s. Estimate the location of transition from laminar to turbulent flow in the boundary layer on the hull of the shell. Calculate the thickness of the turbulent boundary layer at the rear of the hull. Determine the total skin friction drag on the hull under the given conditions.
You are asked by your college crew to estimate the skin friction drag on their eight-seat racing shell. The hull of the shell may be approximated as half a circular cylinder with 457 mm diameter and 7.32 m length. The speed of the shell through the water is 6.71 m/s. Estimate the location of transition from laminar to turbulent flow in the boundary layer on the hull of the shell. Calculate the thickness of the turbulent boundary layer at the rear of the hull. Determine the total skin friction drag on the hull under the given conditions.
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 .
Estimate the drag force on the fuselage shown below for a cruising speed of 210 m/s at 10,000m.
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 Cd 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.
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 what velocity is it flying for the root chord Reynolds number to be 9 × 106? (Round the final answer to the nearest whole number.)
The velocity at which the spitfire is flying for the root chord Reynolds number to be 9 × 106 is ..........ft/s.
Chapter 9 Solutions
Fox and McDonald's Introduction to Fluid Mechanics
DeGarmo's Materials and Processes in Manufacturing
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