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1. For an ideal rocket with a characteristic velocity of c 1220 m/s, a mass flow...
A certain ideal rocket with a nozzle are ratio of 2.3 and a throat area of...
A certain ideal rocket with a nozzle are ratio of 2.3 and a throat area of 5 sq. in. delivers gases at γ = 30 and R = 66 ft-lbf/lbm-⁰R at a chamber pressure of 300 psia and a constant chamber temperature of 5300 ⁰R against a back atmospheric pressure of 10 psia. By means of an appropriate valve arrangement, it is possible to throttle the propellant flow to the thrust chamber. Calculate and plot against pressure the following quantities...
A liquid propellant engine chamber pressure=10 MPa, constant ratio of specific heats = 1.2, a characteristic...
A liquid propellant engine chamber pressure=10 MPa, constant ratio of specific heats = 1.2, a characteristic velocity =1800 m/s. The nozzle has a throat area = 0.015 m2 and an area ratio = 16. Find the mass flux through the throat
Combustion Nozdc Exhaust Fud Exit - V - Velocity m- mass flow rate p- pressure Throat...
Combustion Nozdc Exhaust Fud Exit - V - Velocity m- mass flow rate p- pressure Throat 2.4 kg/s of liquid hydrogen and 24 kg/s of liquid oxygen enter a rocket engine where they are supposed to combust. A large leak in the combustion chamber allows 13 kg/s of the mixture to escape prior to combustion. (Assume that the leak is composed of the same proportion of Hz and Oz as flow into the engine.) The remaining gases combust, and the...
Throat area of 0.25m2 and exit to throat area ratio (Ae/At) of 35.5 are measured to be 1500 kPa a...
throat area of 0.25m2 and exit to throat area ratio (Ae/At) of 35.5 are measured to be 1500 kPa and 3000 K respectively. Assuming that the combustion product is steam ( 1.33, R 461.52 J/kg K), calculate the following: (e) The radius of the throat and exi plane of the nozde if the nozle has a circular cos- section. (b) The nozzle-exit plane conditions, i.e. Mach number, mass flow rate, velocity, pressure, าน temperature, and density. (c) The thrust and...
SP 7. The SpaceX Super Draco reaction control thrusters employ a converging-diverging nozzle to isentropically accelerate...
SP 7. The SpaceX Super Draco reaction control thrusters employ a converging-diverging nozzle to isentropically accelerate the flow of combusted monomethyl hydrazine/nitrogen tetroxide gases to supersonic speeds. The throat area in these engines is At = 0.0345 m2 https://internetprotocol.co/content/images/2020/01/Space-X.png a) Given that the pressure and temperature within the combustion chamber (where velocity 0) are po = 5 MPa and To = 3600 K, respectively, and that the flow exits the nozzle at an exit pressure, Pe 0.5 MPa , find...
Air flows through a converging-diverging nozzle/diffuser. Assuming isentropic flow, air as an ideal gas, and constant...
Air flows through a converging-diverging nozzle/diffuser. Assuming isentropic flow, air as an ideal gas, and constant specific heats determine the state at several locations in the system. Solve using equations rather than with the tables. Note: The specific heat ratio and gas constant for air are given as k=1.4 and R-0287 kJ/kg-K respectively --Given Values-- Inlet Temperature: TI (K) 349 Inlet pressure: Pl (kPa) 460 Inlet Velocity: V1 (m/s) 73 Area at nozzle inlet: Al (cmA2) 8.19 Throat area: A...
Consider the flow through a rocket engine nozzle. In the combustion chamber, the gas which results...
Consider the flow through a rocket engine nozzle. In the combustion chamber, the gas which results from the combustion of the rocket fuel and oxidizer is at a pressure and temperature of 15 atm and 2500K, respectively; the molecular weight and specific heat at constant pressure of the combustion gas are 12 kg/kmol and 4157 J/kg · K, respectively. Assume that the gas flow through the nozzle is an isentropic expansion of calorically perfect gas, with a temperature of 1350K...
Air flows through a converging-diverging nozzle/diffuser. Assuming isentropic flow, air as an ideal gas, and constant...
Air flows through a converging-diverging nozzle/diffuser. Assuming isentropic flow, air as an ideal gas, and constant specific heats determine the state at several locations the system. Solve using equations rather than with the tables. Note: The specific heat ratio and gas constant for air are given as k=1.4 and R=0.287 kJ/kg-K respectively. --Given Values-- Inlet Temperature: T1 (K) = 353 Inlet pressure: Pl (kPa) = 546 Inlet Velocity: V1 (m/s) = 61 Area at nozzle inlet: A1 (cm^2) = 7.24...
2. An aircraft with a single turbojet engine (with an inlet area of 1 m2) is...
2. An aircraft with a single turbojet engine (with an inlet area of 1 m2) is flying at cruising condition with a flight Mach number of 0.7. The ambient temperature and pressure are 250 K and 100 kPa, respectively. The engine compressor pressure ratio is 12, and the turbine inlet temperature is 1200 K. Assume all mechanical components are operating at isentropic condition and the specific heat can be considered a constant (throughout the entire engine) of 1 kJ/(kg K)....
Air flows through a converging-diverging nozzle/diffuser. Assuming isentropic flow, air as an ideal gas, and constant...
Air flows through a converging-diverging nozzle/diffuser. Assuming isentropic flow, air as an ideal gas, and constant specific heats determine the state at several locations in the system. Solve using equations rather than with the tables. Note: The specific heat ratio and gas constant for air are given as k=1.4 and R=0.287 kJ/kg-K respectively. --Given Values-- Inlet Temperature: T1 (K) = 321 Inlet pressure: P1 (kPa) = 588 Inlet Velocity: V1 (m/s) = 97 Area at nozzle inlet: A1 (cm^2) =...
Combustion Nozdc Exhaust Fud Exit - V - Velocity m- mass flow rate p- pressure Throat 2.4 kg/s of liquid hydrogen and 24 kg/s of liquid oxygen enter a rocket engine where they are supposed to combust. A large leak in the combustion chamber allows 13 kg/s of the mixture to escape prior to combustion. (Assume that the leak is composed of the same proportion of Hz and Oz as flow into the engine.) The remaining gases combust, and the...
throat area of 0.25m2 and exit to throat area ratio (Ae/At) of 35.5 are measured to be 1500 kPa and 3000 K respectively. Assuming that the combustion product is steam ( 1.33, R 461.52 J/kg K), calculate the following: (e) The radius of the throat and exi plane of the nozde if the nozle has a circular cos- section. (b) The nozzle-exit plane conditions, i.e. Mach number, mass flow rate, velocity, pressure, าน temperature, and density. (c) The thrust and...
SP 7. The SpaceX Super Draco reaction control thrusters employ a converging-diverging nozzle to isentropically accelerate the flow of combusted monomethyl hydrazine/nitrogen tetroxide gases to supersonic speeds. The throat area in these engines is At = 0.0345 m2 https://internetprotocol.co/content/images/2020/01/Space-X.png a) Given that the pressure and temperature within the combustion chamber (where velocity 0) are po = 5 MPa and To = 3600 K, respectively, and that the flow exits the nozzle at an exit pressure, Pe 0.5 MPa , find...
Air flows through a converging-diverging nozzle/diffuser. Assuming isentropic flow, air as an ideal gas, and constant specific heats determine the state at several locations in the system. Solve using equations rather than with the tables. Note: The specific heat ratio and gas constant for air are given as k=1.4 and R-0287 kJ/kg-K respectively --Given Values-- Inlet Temperature: TI (K) 349 Inlet pressure: Pl (kPa) 460 Inlet Velocity: V1 (m/s) 73 Area at nozzle inlet: Al (cmA2) 8.19 Throat area: A...
Consider the flow through a rocket engine nozzle. In the combustion chamber, the gas which results from the combustion of the rocket fuel and oxidizer is at a pressure and temperature of 15 atm and 2500K, respectively; the molecular weight and specific heat at constant pressure of the combustion gas are 12 kg/kmol and 4157 J/kg · K, respectively. Assume that the gas flow through the nozzle is an isentropic expansion of calorically perfect gas, with a temperature of 1350K...
Air flows through a converging-diverging nozzle/diffuser. Assuming isentropic flow, air as an ideal gas, and constant specific heats determine the state at several locations the system. Solve using equations rather than with the tables. Note: The specific heat ratio and gas constant for air are given as k=1.4 and R=0.287 kJ/kg-K respectively. --Given Values-- Inlet Temperature: T1 (K) = 353 Inlet pressure: Pl (kPa) = 546 Inlet Velocity: V1 (m/s) = 61 Area at nozzle inlet: A1 (cm^2) = 7.24...
2. An aircraft with a single turbojet engine (with an inlet area of 1 m2) is flying at cruising condition with a flight Mach number of 0.7. The ambient temperature and pressure are 250 K and 100 kPa, respectively. The engine compressor pressure ratio is 12, and the turbine inlet temperature is 1200 K. Assume all mechanical components are operating at isentropic condition and the specific heat can be considered a constant (throughout the entire engine) of 1 kJ/(kg K)....
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