1 / 15

Design of Supersonic Intake / Nozzle - PowerPoint PPT Presentation

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.

I am the owner, or an agent authorized to act on behalf of the owner, of the copyrighted work described.

PowerPoint Slideshow about 'Design of Supersonic Intake / Nozzle' - onslow

An Image/Link below is provided (as is) to download presentation

Download Policy: Content on the Website is provided to you AS IS for your information and personal use and may not be sold / licensed / shared on other websites without getting consent from its author.While downloading, if for some reason you are not able to download a presentation, the publisher may have deleted the file from their server.

- - - - - - - - - - - - - - - - - - - - - - - - - - E N D - - - - - - - - - - - - - - - - - - - - - - - - - -
Presentation Transcript

Design of Supersonic Intake / Nozzle

P M V Subbarao

Associate Professor

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

As:

• 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 = 3300K

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 ?

SEA Level Performance

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?