Thrust augmentation in plug nozzles and expansion-deflection nozzles
View Patent ↗Thrust augmentation in a rocket nozzle is achieved by incorporating injectors for the introduction of unburnt fuel and oxidizer into a nozzle to combust in the nozzle and thereby supplement the primary thrust that is supplied by fuel and oxidizer that are combusted prior to entry into the nozzle. These secondary injectors are incorporated into the design of expansion-deflection nozzles and plug nozzles. In expansion-deflection nozzles, the injectors are either in the flow deflector itself or in the wall of the divergent section of the nozzle. In plug nozzles, the injectors are either in a shell of the nozzle surrounding the forward end of the centerbody or in the centerbody itself.
1. A propulsion system comprising:
a rocket nozzle comprising a combustion zone, a throat, and a divergent section disposed along a longitudinal axis to define a serially axial direction of flow extending from said combustion zone, through said throat, and then through said divergent section;
a flow deflector disposed in said divergent section, said flow deflector having a flared end flaring outward along said direction of flow to deflect combustion gas emerging from said throat away from said longitudinal axis; and
separate flow passages extending through said flow deflector, said flow passages parallel to said longitudinal axis and terminating at said flared end to separately inject said fluid fuel and oxidizer into said core.
2. The propulsion system of claim 1 wherein said rocket nozzle is an overexpanded nozzle.
3. The propulsion system of claim 1 wherein said rocket nozzle and said flow deflector are axisymmetrical about said longitudinal axis.
4. The propulsion system of claim 1 wherein said rocket nozzle has an area ratio of at least about 25:1.
5. The propulsion system of claim 1 wherein said rocket nozzle has an area ratio within the range of from about 25:1 to about 150:1.
6. The propulsion system of claim 1 wherein said rocket nozzle has an area ratio within the range of from about 65:1 to about 85:1.
7. The propulsion system of claim 1 wherein said combustion gas deflected by said flow deflector is defined as a primary combustion gas, said rocket nozzle creating a volumetric flow rate of said primary combustion gas through said divergent section, and said passages within said flow deflector are sufficiently large relative to said rocket nozzle to produce, upon combustion of said fuel and oxidizer in said divergent section, a secondary combustion gas at a volumetric flow rate equal to from about 25% to about 75% of said volumetric flow rate of said primary combustion gas.
8. The propulsion system of claim 1 wherein said flow deflector has an aft face that is planar.
9. The propulsion system of claim 1 wherein said rocket nozzle and said flow deflector are axisymmetrical about said longitudinal axis, and said flow deflector has an aft face that is planar and perpendicular to said longitudinal axis.
10. The propulsion system of claim 1 wherein flow deflector is mounted to an aft end of a shaft passing through said combustion zone to define an annular passage within said combustion zone around said shaft, said annular passage having a cross section that constitutes from about 20% to about 90% of the combined cross sections of said annular passage and said shaft.
11. The propulsion system of claim 10 wherein said annular passage has a cross section that constitutes from about 20% to about 50% of the combined cross sections of said annular passage and said flow deflector.