Design of Supersonic Intake / Nozzle P M V Subbarao Associate Professor Mechanical Engineering Department I I T Delhi Meeting the Cruising Conditions…

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Presentation transcript:

Design of Supersonic Intake / Nozzle P M V Subbarao Associate Professor Mechanical Engineering Department I I T Delhi Meeting the Cruising Conditions…

Design Analysis For a known value of Mach number, it is easy to calculate area ratio. Throat area sizing is the first step in the design. If we know the details of the resource/requirements, we can calculate the size of throat.

Cryogenic Rocket Engines A ratio of LO 2 :LH 2 =6:1 T 0 = 3300  K. P 0 = 20.4 Mpa

Specifications of A Rocket Engine Specific Impulse is a commonly used measure of performance For Rocket Engines,and for steady state-engine operation is defined As: At 100% Throttle a RE has the Following performance characteristics F vacuum = 2298 kNt I sp vacuum =450 sec.

Specific impulse of various propulsion technologies Engine "Ve" effective exhaust velocity (m/s, N·s/kg) Specific impulse (s) Energy per kg (MJ/kg) Turbofan jet engine Solid rocket Bipropellant liquid rocket Plasma Rocket VASIMR

The Variable Specific Impulse Magnetoplasma Rocket

Design Procedure Select a technology : I sp & F thrust Estimate the mass flow rate of propellent. Carryout heat release or combustion calculations and estimate T 0 & p 0

Compute properties of gas at each location. Terminate the design when local static pressure is almost zero. This is exit of the nozzle. Compute Maximum Mach number at the exit. This Mach number will generate the required thrust.

Plot Flow Properties Along Nozzle Length A/A *

Mach Number

Temperature T 0 = 3300  K T throat =  K

Pressure P 0 = 20.4Mpa P throat = MPa

Any Doubts !!! The maximum number corresponding to an almost zero static pressure of the gas. This design is meant to work only in Vacuum !!! What is its performance while launching ??? What is the thrust at sea level ? Will the nozzle exit flow be a supersonic ?

SEA Level Performance Ambient Pressure is maximum at Sea level. The design conditions are vacuum. Will the mass flow rate be same ? How to Calculate the corresponding Mass flow rate of propellant ? Will p 0 and T 0 remain same ? What happens if it is not possible to obtain the design mass flow rate ? One needs to know the Mach number distribution for a given geometric design!

Will it satisfy the throat condition? Find the Maximum Mach number at sea level Calculate mass flow rate possible at sea level.