1 / 15

Design of Supersonic Intake / Nozzle

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.

onslow
Download Presentation

Design of Supersonic Intake / Nozzle

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. Content is provided to you AS IS for your information and personal use only. Download presentation by click this link. While downloading, if for some reason you are not able to download a presentation, the publisher may have deleted the file from their server. During download, if you can't get a presentation, the file might be deleted by the publisher.

E N D

Presentation Transcript

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

  2. 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.

  3. Cryogenic Rocket Engines A ratio of LO2:LH2 =6:1 T0 = 3300K. P0 = 20.4 Mpa

  4. 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 Fvacuum = 2298 kNt Ispvacuum = 450 sec.

  5. Specific impulse of various propulsion technologies

  6. The Variable Specific Impulse Magnetoplasma Rocket

  7. Design Procedure Select a technology : Isp & Fthrust Estimate the mass flow rate of propellent. Carryout heat release or combustion calculations and estimate T0 & p0

  8. 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.

  9. Plot Flow Properties Along Nozzle Length • A/A*

  10. • Mach Number

  11. • Temperature T0 = 3300K Tthroat = 2933.3 K

  12. • Pressure P0 = 20.4Mpa Pthroat = 11.32 MPa

  13. 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 ?

  14. 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 ?

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

More Related