Design and experimental validation of a high flow rate pyrogen igniter for cryogenic engine testing

Ignition is an important design consideration for all types of non-hypergolic propellant combinations used in rocket engines. Among the two proven ignition concepts, a generally classified under chemical or electricalcategories, a conventional approach of a solid propellant-based pyrogen igniter has been successfully applied in ISRO cryogenic (LOX/ LH2) engines. Towards development of a new cryogenic engine, a “work-horse” igniter was selected from an ongoing programme for thrust chamber and gas generator ignition.

For this engine, to reduce the development cycle time, a ground test campaign of the thrust chamber was formulated using the facility pressure head in contrast to the pump-fed flight engine. This presented a difference in ignition conditions and a new igniter with a higher mass flow rate compared to the flight igniter was developed as a fall-back option. This igniter incorporated a HTPB based propellant grain triggered using electrical initiators. After detailed analysis, the system was realised and igniter hot-fire tests were carried out toconfirm adequacy. This paper presents the igniter design methodology, ballistic parameters and details of the development programme.