Patent ID: 11970995
Assignee: VENUS AEROSPACE CORP
Field: Engines, pumps, turbines (Mechanical engineering)
Classification: CPC F | IPC F

Claim 0:
1. A rocket engine system comprising:
a heating system configured to heat an oxidizer;
a combustion section having a flow path from an upstream inlet section through a restricted throat section to a downstream outlet section, the combustion section configured to accelerate the heated oxidizer as an oxidizer stream within the flow path in response to flow dynamics to supersonic speed;
a fuel system configured to introduce a fuel into the flow path to mix supersonically with the heated oxidizer to define a combined fuel and oxidizer stream at a first supersonic speed;
a propellant introduction system configured to introduce the fuel and the oxidizer, the propellant introduction system comprises a central feed nozzle centrally disposed within the inlet section of the combustion section having an annulus area generally surrounding the central feed nozzle;
wherein the combined fuel and oxidizer stream undergoes a deflagration to detonation transition in the combustion section defined by an oblique shock wave and a normal shock wave;
wherein the combustion section is configured such that the oblique shock wave and the normal shock wave interact to achieve a standing detonation wave generally at an upstream portion of the restricted throat section, the standing detonation wave configured to prevent combustion from moving upstream, wherein combustion exits the downstream outlet section to provide thrust; and
wherein the central feed nozzle comprises a convergent section and a divergent section together configured to drive the oxidizer to the supersonic speed.