| Back pressure | is the pressure applied at the nozzle discharge region.
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| Bernoulli equation | is the result of the energy analysis for the reversible, steady-flow of an incompressible liquid through a device that involves no work interactions (such as a nozzle or a pipe section). For frictionless flow, it states that the sum of the pressure, velocity, and potential energy heads is constant. It is also a form of the conservation of momentum principle for steady-flow control volumes.
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| Bow wave | (see detached oblique shock)
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| Choked flow | occurs in a nozzle when the mass flow reaches a maximum value for the minimum flow area. This happens when the flow properties are those required to increase the fluid velocity to the velocity of sound at the minimum flow area location.
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| Choked Rayleigh flow | occurs in a duct when a fluid can no longer be accelerated by heating above sonic velocity to supersonic velocities.
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| Compressing flow | is a flow that produces an oblique shock.
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| Converging-diverging nozzle | also called Laval nozzle after Carl G. B. de Laval, is a duct in which the flow area first decreases and then increases in the direction of the flow and is used to accelerate gases to supersonic speeds.
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| Critical properties | are the properties of a fluid at a location where the Mach number is unity.
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| Critical ratios | are the ratios of the stagnation to static properties when the Mach number is unity.
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| Deflection angle | (see turning angle)
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| Detached oblique shock or a bow wave | is an oblique shock that has become curved and detached from the nose of a wedge.
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| Dynamic temperature | is the kinetic energy per unit mass divided by the constant pressure specific heat and corresponds to the temperature rise during the stagnation process.
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| Expanding flow | are those flows where supersonic flow is turned in the opposite direction; however, the flow does not turn suddenly, as through a shock, but gradually-each successive Mach wave turns the flow by an infinitesimal amount.
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| Expansion fan | is a continuous expanding region of supersonic flow composed of an infinite number of Mach waves called Prandtl-Meyer expansion waves.
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| Fanno line | is the locus of all states for frictionless adiabatic flow in a constant-area duct plotted on an h-s diagram. These states have the same value of stagnation enthalpy and mass flux (mass flow per unit area).
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| Hypersonic flow | occurs when a flow has a Mach number M > > 1.
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| Isentropic stagnation state | is the stagnation state when the stagnation process is reversible as well as adiabatic (i.e., isentropic). The entropy of a fluid remains constant during an isentropic stagnation process.
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| Laval nozzles | (see converging-diverging nozzles)
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| Mach angle | is the shock angle for Mach waves and is a unique function of the Mach number.
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| Mach number | named after the Austrian physicist Ernst Mach (1838-1916), is the ratio of the actual velocity of the fluid (or an object in still air) to the speed of sound in the same fluid at the same state.
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| Mach wave | is the weakest possible oblique shock at a Mach number.
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| Normal components | are components that are perpendicular to the quantity in question.
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| Normal shock wave | is a shock wave resulting in an abrupt change over a very thin section normal to the direction of flow.
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| Oblique shock | is a complicated shock pattern consisting of inclined shock waves in which some portions of an oblique shock are curved, while other portions are straight.
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| Prandtl-Meyer expansion waves | are the Mach waves that compose a continuous expanding region called an expansion fan.
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| Prandtl-Meyer function | is the angle through which flow must expand, starting with the function value of zero at Ma = 1, in order to reach a supersonic Mach number, Ma > 1.
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| Rayleigh flow | is the steady one-dimensional flow of an ideal gas with constant specific heats through a constant-area duct with heat transfer, but with negligible friction.
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| Rayleigh line | is the locus of all states for frictionless flow in a constant-area duct with heat transfer plotted on an h-s diagram and results from combining the conservation of mass and momentum equations into a single equation.
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| Shock wave | (wave angle) is an abrupt change over a very thin section of flow in which the flow transitions from supersonic to subsonic flow. This abrupt change in the flow causes a sudden drop in velocity to subsonic levels and a sudden increase in pressure. Flow through the shock is highly irreversible; and, thus, it cannot be approximated as isentropic.
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| Sonic flow | occurs when a flow has a Mach number M =1.
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| Sonic speed | (see speed of sound)
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| Shock angle | is the angle at which straight oblique shocks are deflected relative to the oncoming flow as the flow comes upon a body.
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| Speed of sound | (sonic speed) is the speed at which an infinitesimally small pressure wave travels through a medium.
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| Stagnation enthalpy | (total enthalpy) is the sum of the enthalpy and kinetic energy of the flow and represents the total energy of a flowing fluid stream per unit mass. It represents the enthalpy of a fluid when it is brought to rest adiabatically with no work. The stagnation enthalpy equals the static enthalpy when the kinetic energy of the fluid is negligible.
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| Stagnation pressure | is the pressure a fluid attains when brought to rest isentropically. For ideal gases with constant specific heats, the stagnation pressure is related to the static pressure of the fluid through the isentropic process equation relating pressure and temperature.
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| Stagnation properties | are the properties of a fluid at the stagnation state. These properties are called stagnation temperature, stagnation pressure, stagnation density, etc. The stagnation state and the stagnation properties are indicated by the subscript 0.
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| Stagnation temperature | (total temperature) is the temperature an ideal gas will attain when it is brought to rest adiabatically.
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| Static enthalpy | is the ordinary enthalpy of the flow measured at the fluid state.
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| Strong oblique shocks | are straight oblique shocks that have the larger possible values of the shock angles for deflection angles less than the maximum deflection angle.
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| Subsonic flow | occurs when a flow has a Mach number M < 1.
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| Supersonic flow | occurs when a flow has a Mach number M > 1.
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| Supersaturated steam | is steam that exists in the wet region without containing any liquid. This phenomenon would exist due to the supersaturation process.
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| Supersaturation | is the phenomenon owing to steam flowing through a nozzle with the high velocities and exiting the nozzle in the saturated region. Since the residence time of the steam in the nozzle is small, and there may not be sufficient time for the necessary heat transfer and the formation of liquid droplets, the condensation of the steam may be delayed for a little while.
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| Throat | is the smallest flow area of a converging-diverging nozzle.
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| Transonic flow | occurs when a flow has a Mach number M is approximately equal to 1.
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| Total enthalpy | (see stagnation enthalpy)
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| Total temperature | (see stagnation temperature)
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| Turning angle | (deflection angle) is the angle at which straight oblique shocks are deflected as flow comes upon a body, like that produced when a uniform supersonic flow impinges on a slender, two-dimensional wedge.
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| Venturi nozzle | is a duct in which the flow area first decreases and then increases in the direction of the flow and is used strictly for incompressible flow.
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| Wave angle | (see shock angle)
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| Weak oblique shocks | are straight oblique shocks that have the smaller of the possible values of the shock angles for deflection angles less than the maximum deflection angle.
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| Wilson line | is the locus of points where condensation will take place regardless of the initial temperature and pressure as steam flows through a high-velocity nozzle. The Wilson line is often approximated by the 4 percent moisture line on the h-s diagram for steam. Therefore, steam flowing through a high-velocity nozzle is assumed to begin condensation when the 4 percent moisture line is crossed.
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