Week 2 - Converging diverging nozzle

Converge-Divergent Nozzle

Consider the converging-diverging nozzle below with the dimensions.

For the following boundary conditions,

• At Inlet, Pressure = 1atm and Temperature = 300K. The outlet pressure= 0.027 atm.

• In the operating conditions, consider the reference pressure to be zero.

Answer

The given 2D diagram of a convergent-divergent nozzle can be drawn in space-claim.

to do this, open space claim and draw as mentioned.

Using a Pull surface tool, a surface can be generated out of the diagram.

Meshing..

Here, for an accurate result, i used 5.0mm cell size. which give more refined results

Named Selections

Case Setup

Solver Type - Pressure Based

2D Space - Axisymmetry

Time - Steady

Viscous - Inviscid

Operating Conditions pressure = 0

Inlet Pressure

Outlet Pressure

Results

Scaled average residuals

Pressure plot across the length of the nozzle

Velocity plot across the length of the plot

RESULTS, CFD POST

Pressure Contour

Velocity Contour

Plot showing Velocity Stream lines

• As seen from the contours the velocity gets increased along the length of nozzle
• Temperature and pressure is higher at inlet
• As mentioned in introduction, flow becomes supersonic at a distance of 500 mm from inlet this can be varified by looking the mach number plot as it shows the mach nuimber as 1 at 0.5 metres from inlet

A nozzle is a relatively simple device, just a specially shaped tube through which hot gases flow. Ramjets and rockets typically use a fixed convergent section followed by a fixed divergent section for the design of the nozzle. This nozzle configuration is called a convergent-divergent, or CD, nozzle. In a CD nozzle, the hot exhaust leaves the combustion chamber and converges down to the minimum area, or throat, of the nozzle. The throat size is chosen to choke the flow and set the mass flow rate through the system. The flow in the throat is sonic which means the Mach number is equal to one in the throat. Downstream of the throat, the geometry diverges and the flow is isentropically expanded to a supersonic Mach number that depends on the area ratio of the exit to the throat. The expansion of a supersonic flow causes the static pressure and temperature to decrease from the throat to the exit, so the amount of the expansion also determines the exit pressure and temperature. The exit temperature determines the exit speed of sound, which determines the exit velocity. The exit velocity, pressure, and mass flow through the nozzle determines the amount of thrust produced by the nozzle.

Ramjets, scramjets, and rockets all use nozzles to accelerate hot exhaust to produce thrust as described by Newron's third law of motion. The amount of thrust produced by the engine depends on the mass flow rate through the engine, the exit velocity of the flow, and the pressure at the exit of the engine. The value of these three flow variables are all determined by the nozzle design.

Thank You.