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A rocket's required mass ratio as a function of effective exhaust velocity ratio. The classical rocket equation, or ideal rocket equation is a mathematical equation that describes the motion of vehicles that follow the basic principle of a rocket: a device that can apply acceleration to itself using thrust by expelling part of its mass with high velocity and can thereby move due to the ...
Absolute cylinder pressure is used to calculate the dynamic compression ratio, using the following formula: = where is a polytropic value for the ratio of specific heats for the combustion gases at the temperatures present (this compensates for the temperature rise caused by compression, as well as heat lost to the cylinder)
The flow in a real exhaust port can easily be sonic with choked flow occurring and even supersonic flow in areas. The very high temperature causes the viscosity of the gas to increase, all of which alters the Reynolds number drastically. Added to the above is the profound effect that downstream elements have on the flow of the exhaust port.
The bypass ratio (BPR) of a turbofan engine is the ratio between the mass flow rate of the bypass stream to the mass flow rate entering the core. [1] A 10:1 bypass ratio, for example, means that 10 kg of air passes through the bypass duct for every 1 kg of air passing through the core.
Calculating the effective exhaust velocity requires averaging the two mass flows as well as accounting for any atmospheric pressure. [12] For air-breathing jet engines, particularly turbofans, the actual exhaust velocity and the effective exhaust velocity are different by orders of magnitude. This happens for several reasons.
The optimal size of a rocket engine nozzle is achieved when the exit pressure equals ambient (atmospheric) pressure, which decreases with increasing altitude. The reason for this is as follows: using a quasi-one-dimensional approximation of the flow, if ambient pressure is higher than the exit pressure, it decreases the net thrust produced by ...
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Beyond this point the nozzle diameter becomes the biggest diameter and starts to incur increasing drag. Nozzles are thus limited to the installation size and the loss in thrust incurred is a trade off with other considerations such as lower drag, less weight. Examples are the F-16 at Mach 2.0 [21] and the XB-70 at Mach 3.0. [22]