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.
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P M V Subbarao
Mechanical Engineering Department
I I T Delhi
Meeting the Cruising Conditions…
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.
A ratio of LO2:LH2 =6:1
T0 = 3300K.
P0 = 20.4 Mpa
• Specific Impulse is a commonly used measure of performance
For Rocket Engines,and for steady state-engine operation is defined
• At 100% Throttle a RE has the Following performance characteristics
Fvacuum = 2298 kNt
Ispvacuum = 450 sec.
Select a technology : Isp & Fthrust
Estimate the mass flow rate of propellent.
Carryout heat release or combustion calculations and estimate
T0 & p0
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.
T0 = 3300K
Tthroat = 2933.3 K
P0 = 20.4Mpa
Pthroat = 11.32 MPa
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 ?
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!
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 p0 and T0 remain same ?
Calculate mass flow rate possible at sea level.
Will it satisfy the throat condition?