Shreeya Khanal

Find the drag force difference on a flat plate of size 1.5 m x 1.5 m when the plate is moving at a speed of 5 m/s normal to its plate first in water and second in air of density 1.24 kg/m³. Co-efficient of drag is given as 1.10.

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Drag Force Calculation in Water and Air Problem Statement Find the drag force difference on a flat plate of size […]

Find the drag force difference on a flat plate of size 1.5 m x 1.5 m when the plate is moving at a speed of 5 m/s normal to its plate first in water and second in air of density 1.24 kg/m³. Co-efficient of drag is given as 1.10. Read More »

A flat plate 2 m x 2 m moves at 40 km/hour in stationary air of density 1.25 kg/m³. If the co-efficient of drag and lift are 0.2 and 0.8 respectively, find : (i) the lift force, (ii) the drag force, (iii) the resultant force, and (iv) the power required to keep the plate in motion.

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Aerodynamic Forces on a Moving Plate Problem Statement A flat plate 2 m x 2 m moves at 40 km/hour

A flat plate 2 m x 2 m moves at 40 km/hour in stationary air of density 1.25 kg/m³. If the co-efficient of drag and lift are 0.2 and 0.8 respectively, find : (i) the lift force, (ii) the drag force, (iii) the resultant force, and (iv) the power required to keep the plate in motion. Read More »

Find the Mach number when an aeroplane is flying at 1000 km/hour through still air having pressure of 7 N/cm² and temperature of -5°C. Take R = 287.14 J/kg K. Calculate the pressure and temperature of air at stagnation point. Take k = 1.4.

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Aeroplane Stagnation Properties Calculation Problem Statement Find the Mach number when an aeroplane is flying at 1000 km/hour through still

Find the Mach number when an aeroplane is flying at 1000 km/hour through still air having pressure of 7 N/cm² and temperature of -5°C. Take R = 287.14 J/kg K. Calculate the pressure and temperature of air at stagnation point. Take k = 1.4. Read More »

Calculate the numerical factor by which the actual pressure difference shown by the gauge of a pitot-tube should be multiplied to allow for compressibility when the value of the Mach number is 0.7. Take k = 1.4.

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Pitot-Tube Compressibility Factor Problem Statement Calculate the numerical factor by which the actual pressure difference shown by the gauge of

Calculate the numerical factor by which the actual pressure difference shown by the gauge of a pitot-tube should be multiplied to allow for compressibility when the value of the Mach number is 0.7. Take k = 1.4. Read More »

Find the mass rate of flow of air through a venturimeter having inlet diameter as 400 mm and throat diameter 200 mm. The pressure at the inlet of the venturimeter is 27.468 N/cm² (abs.) and temperature of air at inlet is 20°C. The pressure at the throat is given as 25.506 N/cm² absolute. Take k = 1.4 and R = 287 J/kg K.

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Mass Flow Rate Through a Venturimeter Problem Statement Find the mass rate of flow of air through a venturimeter having

Find the mass rate of flow of air through a venturimeter having inlet diameter as 400 mm and throat diameter 200 mm. The pressure at the inlet of the venturimeter is 27.468 N/cm² (abs.) and temperature of air at inlet is 20°C. The pressure at the throat is given as 25.506 N/cm² absolute. Take k = 1.4 and R = 287 J/kg K. Read More »

A nozzle of diameter 20 mm is fitted to a large tank which contains air at 20°C. The air flows from the tank into atmosphere. For adiabatic flow, find the mass rate of flow of air through the nozzle when pressure of air in tank is (i) 5.886 N/cm² (gauge) and (ii) 29.43 N/cm² (gauge). Take k = 1.4 and R = 287 J/kg K and atmospheric pressure = 9.81 N/cm².

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Adiabatic Nozzle Flow Calculation Problem Statement A nozzle of diameter 20 mm is fitted to a large tank which contains

A nozzle of diameter 20 mm is fitted to a large tank which contains air at 20°C. The air flows from the tank into atmosphere. For adiabatic flow, find the mass rate of flow of air through the nozzle when pressure of air in tank is (i) 5.886 N/cm² (gauge) and (ii) 29.43 N/cm² (gauge). Take k = 1.4 and R = 287 J/kg K and atmospheric pressure = 9.81 N/cm². Read More »

Find the velocity of air flowing at the outlet of a nozzle, fitted to a large vessel which contains air at a pressure of 294.3 N/cm² (abs.) and at a temperature of 30°C. The pressure at the outlet of the nozzle is 137.34 N/cm² (abs.) Take k = 1.4 and R = 287 J/kg K.

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Mass Flow Rate Through a Nozzle (Adiabatic Flow) Problem Statement Find the velocity of air flowing at the outlet of

Find the velocity of air flowing at the outlet of a nozzle, fitted to a large vessel which contains air at a pressure of 294.3 N/cm² (abs.) and at a temperature of 30°C. The pressure at the outlet of the nozzle is 137.34 N/cm² (abs.) Take k = 1.4 and R = 287 J/kg K. Read More »

A gas is flowing through a horizontal pipe of cross-sectional area of 30 cm². At a point the pressure is 30 N per cm² (gauge) and temperature 20°C. At another section the area of cross-section is 15 cm² and pressure is 25 N/cm² gauge. If the mass rate of flow of gas is 0.15 kg/s, find the velocities of the gas at these two sections, assuming an isothermal change. 

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Gas Flow Velocity Calculation Problem Statement A gas is flowing through a horizontal pipe of cross-sectional area of 30 cm².

A gas is flowing through a horizontal pipe of cross-sectional area of 30 cm². At a point the pressure is 30 N per cm² (gauge) and temperature 20°C. At another section the area of cross-section is 15 cm² and pressure is 25 N/cm² gauge. If the mass rate of flow of gas is 0.15 kg/s, find the velocities of the gas at these two sections, assuming an isothermal change.  Read More »

Find the Mach number when an aeroplane is flying at 900 km/hour through still air having a pressure of 8.0 N/cm² and temperature -15°C. Take k = 1.4 and R = 287 J/kg K. Calculate the pressure, temperature and density of air at the stagnation point on the nose of the plane.

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Aeroplane Stagnation Properties Calculation Problem Statement Find the Mach number when an aeroplane is flying at 900 km/hour through still

Find the Mach number when an aeroplane is flying at 900 km/hour through still air having a pressure of 8.0 N/cm² and temperature -15°C. Take k = 1.4 and R = 287 J/kg K. Calculate the pressure, temperature and density of air at the stagnation point on the nose of the plane. Read More »

A projectile travels in air of pressure 8.829 N/cm² at -10°C at a speed of 1200 km/hour. Find the Mach number and the Mach angle. Take k = 1.4 and R = 287 J/kg K.

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Mach Number and Mach Angle Calculation Problem Statement A projectile travels in air of pressure 8.829 N/cm² at -10°C at

A projectile travels in air of pressure 8.829 N/cm² at -10°C at a speed of 1200 km/hour. Find the Mach number and the Mach angle. Take k = 1.4 and R = 287 J/kg K. Read More »

A projectile is travelling in air having pressure and temperature as 8.829 N/cm² and – 5°C. If the Mach angle is 30°, find the velocity of the projectile. Take k = 1.4 and R = 287 J/kg K.

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Projectile Velocity from Mach Angle Problem Statement A projectile is travelling in air having pressure and temperature as 8.829 N/cm²

A projectile is travelling in air having pressure and temperature as 8.829 N/cm² and – 5°C. If the Mach angle is 30°, find the velocity of the projectile. Take k = 1.4 and R = 287 J/kg K. Read More »

An aeroplane is flying at an height of 20 km, where the temperature is – 40°C. The speed of the plane is corresponding to M = 1.8. Assuming k = 1.4 and R = 287 J/kg K, find the speed of the plane.

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Aircraft Speed Calculation from Mach Number Problem Statement An aeroplane is flying at an height of 20 km, where the

An aeroplane is flying at an height of 20 km, where the temperature is – 40°C. The speed of the plane is corresponding to M = 1.8. Assuming k = 1.4 and R = 287 J/kg K, find the speed of the plane. Read More »

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