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 for
300, 200, and 100 psia chamber pressure:
- Pressure ratio between chamber and atmosphere
- Effective exhaust velocity for area ratio involved
- Ideal exhaust velocity for optimum and actual area ratio
- Propellant flow
- Thrust
- Specific impulse
- Exit pressure
- Exit temperature
Homework Answers
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A certain ideal rocket with a nozzle are ratio of 2.3 and a
throat area of...
The throat area of a rocket nozzle is 5 cm2. The exit to throat area ratio...
The throat area of a rocket nozzle is 5 cm2. The exit to throat area ratio is 2.9. The chamber pressure is 20 atm and the chamber temperature is 2800 k. what is the maximum possible mass flow rate through the nozzle. assume one dimensional isentropic flow inside the nozzle . If the ambient pressure at the exit is p =0.1 atm , is the flow over expanded or underexpanded?
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...
1. For an ideal rocket with a characteristic velocity of c 1220 m/s, a mass flow...
1. For an ideal rocket with a characteristic velocity of c 1220 m/s, a mass flow rate of 73 kg/s, a thrust coefficient of 1.5 and a nozzle throat area (A0.0248 m2), compute a. The effective exhaust velocity, c b. The thrust, F c. The chamber pressure, pc d. And the specific impulse, Isp
Problem 2.3. An ideal ramjet is to fly at 20,000 ft with a Mach number of...
Problem 2.3. An ideal ramjet is to fly at 20,000 ft with a Mach number of 3.5. The burner exit total temperature is to be 3200 °?? and the engine will use 145 lbm/s of air. The heating value of the fuel is 18,500 Btu/lbm. What is the diameter of the rounded exit, thrust, dimensionless thrust, and TSFC at this condition? (Assume that the temperature is 447.38°??, the static pressure is 6.747161 psia, and the specific heat ratio is 1.4...
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...
At launch, the space shuttle main engine (SSME) has 1030 lbm/s gas leaving the combustion chamber...
At launch, the space shuttle main engine (SSME) has 1030 lbm/s gas leaving the combustion chamber at Pt 3000 psia ani Tt 7350 R. The exit area of the SSME's nozzle is 77 times the throat area. If the flow through the nozzle is considered to reversible and adiabatic (isentropic) with Rgc 3800 ft s.°R ani γ= 1.25, find the area of the nozzle throat (in2) and the exit Mad number. Hint: Use the MFP to get the throat area....
The Mach number at the throat of a nozzle with exit-to-throat area ratio of 2.637 is...
The Mach number at the throat of a nozzle with exit-to-throat area ratio of 2.637 is 0.56. Calculate the Mach number at the exit and the plenum pressure if the back pressure is 2atm.
[15 pts] Consider a converging diverging nozzle with an exit-to-throat area ratio of Ae/At = 1.25...
[15 pts] Consider a converging diverging nozzle with an exit-to-throat area ratio of Ae/At = 1.25 as shown below. The stagnation pressure upstream of the throat is 8.5 atm and the stagnation temperature is 1000 K. (a) Assume the air is expanded isentropically to supersonic speed at the exit. Determine the following properties at the nozzle exit: Me, Pe, Te, Pe, ue, Poe, Toe (b) If the area ratio in the subsonic part of the converging diverging nozzle, A1/A is...
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)....
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...
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...
1. For an ideal rocket with a characteristic velocity of c 1220 m/s, a mass flow rate of 73 kg/s, a thrust coefficient of 1.5 and a nozzle throat area (A0.0248 m2), compute a. The effective exhaust velocity, c b. The thrust, F c. The chamber pressure, pc d. And the specific impulse, Isp
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...
At launch, the space shuttle main engine (SSME) has 1030 lbm/s gas leaving the combustion chamber at Pt 3000 psia ani Tt 7350 R. The exit area of the SSME's nozzle is 77 times the throat area. If the flow through the nozzle is considered to reversible and adiabatic (isentropic) with Rgc 3800 ft s.°R ani γ= 1.25, find the area of the nozzle throat (in2) and the exit Mad number. Hint: Use the MFP to get the throat area....
The Mach number at the throat of a nozzle with exit-to-throat area ratio of 2.637 is 0.56. Calculate the Mach number at the exit and the plenum pressure if the back pressure is 2atm.
[15 pts] Consider a converging diverging nozzle with an exit-to-throat area ratio of Ae/At = 1.25 as shown below. The stagnation pressure upstream of the throat is 8.5 atm and the stagnation temperature is 1000 K. (a) Assume the air is expanded isentropically to supersonic speed at the exit. Determine the following properties at the nozzle exit: Me, Pe, Te, Pe, ue, Poe, Toe (b) If the area ratio in the subsonic part of the converging diverging nozzle, A1/A 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)....
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...
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