Deflectable distributed aerospike rocket nozzle
A rocket engine nozzle includes an aerospike having a plurality of adjustable airfoil vanes distributed around a central longitudinal axis of a rocket engine combustion chamber. The aerospike is integrated on an exit plane at an exit end of the combustion chamber. The adjustable airfoil vanes and an inner perimeter of the combustion chamber define a plurality of apertures which choke an exhaust exiting the combustion chamber and cause the exhaust to expand supersonically along the adjustable airfoil vanes, creating a supersonic jet. An actuator is configured to adjust a position of each of the adjustable airfoil vane relative to each other so as to direct the exhaust exiting the rocket engine combustion chamber as the exhaust expands supersonically over the airfoil vanes without causing a shockwave to be imparted on the supersonic jet that is created. Accordingly, performance of the rocket engine is improved over conventional systems.
1. A rocket engine nozzle comprising:
an aerospike including a plurality of adjustable airfoil vanes disposed at an exit end of a rocket engine combustion chamber and extending across a circular exit opening of the rocket engine combustion chamber at the exit end of the rocket engine combustion chamber from a first point on an inner perimeter of the rocket engine combustion chamber at the exit end to a second point on the inner perimeter of the rocket engine combustion chamber at the exit end, the plurality of adjustable airfoil vanes being distributed around a central longitudinal axis, the plurality of adjustable airfoil vanes and the inner perimeter of the rocket engine combustion chamber at the exit end defining a plurality of apertures between adjacent adjustable airfoil vanes at the exit opening, the plurality of apertures being configured to choke an exhaust exiting the rocket engine combustion chamber and cause the exhaust to expand supersonically along the plurality of adjustable airfoil vanes to create a supersonic jet; and
an actuator configured to adjust a position of the plurality of adjustable airfoil vanes relative to each other to thereby direct the exhaust exiting the rocket engine combustion chamber as the exhaust expands supersonically along the plurality of adjustable airfoil vanes to thereby direct the supersonic jet and vector a thrust of the supersonic jet.
2. The rocket engine nozzle according to claim 1 , wherein the actuator is configured to rotate one or more of the plurality of adjustable airfoil vanes around a respective radial axis on which the one or more of the plurality of adjustable airfoil vane is positioned.
3. The rocket ending nozzle according to claim 1 , wherein the actuator includes one or more actuation hinges.
4. The rocket engine nozzle according to claim 1 , wherein the actuator is configured to translate one or more of the plurality of adjustable airfoil vanes linearly along a respective oblique axis, the respective oblique axis being perpendicular to a respective radial axis on which the one or more adjustable airfoil vane is positioned.
5. The rocket engine nozzle according to claim 1 , wherein the actuator includes one or more linear actuators.
6. The rocket engine nozzle according to claim 1 , wherein the plurality of adjustable airfoil vanes include four adjustable airfoil vanes.
7. The rocket engine nozzle according to claim 1 , wherein the actuator is configured to adjust each of the plurality of adjustable airfoil vanes independently from each other.
8. The rocket engine nozzle according to claim 1 , wherein the actuator is configured to adjust one or more of the plurality of airfoil vanes in a coordinated manner.
9. A rocket engine comprising:
a rocket engine combustion chamber; and
a rocket engine nozzle including:
an aerospike including a plurality of adjustable airfoil vanes disposed at an exit end of the rocket engine combustion chamber and extending across a circular exit opening of the rocket engine combustion chamber at the exit end of the rocket engine combustion chamber from a first point on an inner perimeter of the rocket engine combustion chamber at the exit end to a second point on the inner perimeter of the rocket engine combustion chamber at the exit end, the plurality of adjustable airfoil vanes being distributed around a central longitudinal axis, the plurality of adjustable airfoil vanes and the inner perimeter of the rocket engine combustion chamber at the exit end defining a plurality of apertures between adjacent adjustable airfoil vanes at the exit opening, the plurality of apertures being configured to choke an exhaust exiting the rocket engine combustion chamber and cause the exhaust to expand supersonically along the plurality of adjustable airfoil vanes to create a supersonic jet; and
an actuator configured to adjust a position of the plurality of adjustable airfoil vanes relative to each other to thereby direct the exhaust exiting the rocket engine combustion chamber as the exhaust expands supersonically along the plurality of adjustable airfoil vanes to thereby direct the supersonic jet and vector a thrust of the supersonic jet.
10. The rocket engine according to claim 9 , wherein the actuator is configured to rotate one or more of the plurality of adjustable airfoil vanes around a respective radial axis on which the one or more of the plurality of adjustable airfoil vane is positioned.
11. The rocket engine according to claim 9 , wherein the actuator includes one or more actuation hinges.
12. The rocket engine according to claim 9 , wherein the actuator is configured to translate one or more of the plurality of adjustable airfoil vanes linearly along a respective oblique axis, the respective oblique axis being perpendicular to a respective radial axis on which the one or more adjustable airfoil vane is positioned.
13. The rocket engine according to claim 9 , wherein the actuator includes one or more linear actuators.
14. The rocket engine according to claim 9 , wherein the plurality of adjustable airfoil vanes include four adjustable airfoil vanes.
15. The rocket engine according to claim 9 , wherein the actuator is configured to adjust each of the plurality of adjustable airfoil vanes independently from each other.
16. The rocket engine according to claim 9 , wherein the actuator is configured to adjust one or more of the plurality of airfoil vanes in a coordinated manner.
17. A method of operating a rocket propulsion system, the method comprising:
providing a rocket engine, the rocket engine including a rocket engine combustion chamber and a rocket engine nozzle, the rocket engine nozzle including an aerospike including a plurality of adjustable airfoil vanes disposed at an exit end of the rocket engine combustion chamber and extending across a circular exit opening of the rocket engine combustion chamber at the exit end of the rocket engine combustion chamber from a first point on an inner perimeter of the rocket engine combustion chamber at the exit end to a second point on the inner perimeter of the rocket engine combustion chamber at the exit end, the plurality of adjustable airfoil vanes being distributed around a central longitudinal axis, the plurality of adjustable airfoil vanes and the inner perimeter of the rocket engine combustion chamber at the exit end defining a plurality of apertures between adjacent adjustable airfoil vanes at the exit opening, the plurality of apertures being configured to choke an exhaust exiting the rocket engine combustion chamber and cause the exhaust to expand supersonically along the plurality of adjustable airfoil vanes to create a supersonic jet; an actuator configured to adjust a position of the plurality of adjustable airfoil vanes relative to each other to thereby direct the exhaust exiting the rocket engine combustion chamber as the exhaust expands supersonically along the plurality of adjustable airfoil vanes to thereby direct the supersonic jet and vector a thrust of the supersonic jet, the method further comprising adjusting a position of one or more of the plurality of adjustable airfoil vanes relative to each other; and operating the rocket engine such that the exhaust exits the rocket engine combustion chamber at the exit plane and the exhaust is choked by the plurality of apertures and expands supersonically along the plurality of airfoil vanes in their respective positions thereby directing the supersonic jet and vectoring the thrust of the supersonic jet.
18. The method according to claim 17 , wherein the adjusting includes rotating the one or more of the plurality of adjustable airfoil vanes around a respective radial axis on which the one or more adjustable airfoil vane is positioned.
19. The method according to claim 17 , wherein the adjusting includes translating the one or more of the plurality of adjustable airfoil vanes linearly along a respective oblique axis, the respective oblique axis being perpendicular to a respective radial axis on which the one or more adjustable airfoil vane is positioned.