Distributed airfoil aerospike rocket nozzle

Patent metadata

Field	Value
Publication number	US-11512669-B2
Application number	US-202016910342-A
Country	US
Kind code	B2
Filing date	Jun 24, 2020
Priority date	Jun 24, 2020
Publication date	Nov 29, 2022
Grant date	Nov 29, 2022

Abstract

A rocket engine nozzle manufacturable and applicable to tactical missile designs includes an aerospike having a plurality of airfoil fins 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 airfoil fins and an inner perimeter of the combustion chamber define a plurality of apertures which choke an airflow exiting the combustion chamber and cause the airflow to expand supersonically along the airfoil fins. The aerospike rocket engine nozzle requires less machine precision and achieves packing benefits over conventional bell and aerospike nozzle geometries. The configuration of the aerospike rocket engine nozzle also removes the producibility and heating constraints typically encountered with conventional aerospike nozzles in tactical missile applications while improving thrust performance of the rocket engine across a wide range of altitudes.

First claim

What is claimed is:

1. A rocket engine comprising:
a rocket engine combustion chamber, and
an aerospike rocket engine nozzle including:
a plurality of airfoil fins disposed at an exit end of a rocket engine combustion chamber and extending across an exit opening of the rocket engine combustion chamber, the plurality of airfoil fins being distributed around a central longitudinal axis; and
a central airfoil hub from which each of the plurality of airfoil fins extend radially outward, wherein a maximum length of the central airfoil hub in a longitudinal direction is less than or equal to a maximum length of the plurality of airfoil fins in the longitudinal direction;
wherein the plurality of airfoil fins the central airfoil hub and the inner perimeter of the rocket engine combustion chamber define a plurality of apertures between adjacent airfoil fins at the exit opening, the plurality of apertures being configured to choke an airflow exiting the rocket engine combustion chamber and cause the airflow to expand supersonically along the plurality of airfoil fins to create thrust.

2. The rocket engine according to claim 1 , wherein the rocket engine combustion chamber is cylindrical.

3. The rocket engine according to claim 1 , wherein the plurality of airfoil fins include four airfoil fins.

4. The rocket engine according to claim 1 , wherein each of the plurality of airfoil fins are fixed to the exit end of the rocket engine combustion chamber at the exit opening of the rocket engine combustion chamber.

5. A method of operating a rocket propulsion system, the method comprising:
providing a rocket engine including a rocket engine combustion chamber and an aerospike rocket engine nozzle, the aerospike rocket engine nozzle including:
a plurality of airfoil fins disposed at an exit end of the rocket engine combustion chamber and extending across an exit opening of the rocket engine combustion chamber, the plurality of airfoil fins being distributed around a central longitudinal axis,
a central airfoil hub from which each of the plurality of airfoil fins extend radially outward, wherein a maximum length of the central airfoil hub in a longitudinal direction is less than or equal to a maximum length of the plurality of airfoil fins in the longitudinal direction;
wherein the plurality of airfoil fins, the central airfoil hub and an inner perimeter of the rocket engine combustion chamber define a plurality of apertures between adjacent airfoil fins at the exit opening, and

operating the rocket engine such that an airflow exits the rocket engine combustion chamber through the exit opening and the plurality of apertures choke the airflow exiting the rocket engine combustion chamber through the exit opening, causing the airflow to expand supersonically along the plurality of airfoil fins to create thrust.

6. The method of operating a rocket propulsion system according to claim 5 , wherein the plurality of airfoil fins include four airfoil fins.

7. An aerospike rocket engine nozzle comprising:
a plurality of airfoil fins disposed at an exit end of a rocket engine combustion chamber and extending across an exit opening of the rocket engine combustion chamber, the plurality of airfoil fins being distributed around a central longitudinal axis; and
a central airfoil hub from which each of the plurality of airfoil fins extend radially outward, wherein a maximum length of the central airfoil hub in a longitudinal direction is less than or equal to a maximum length of the plurality of airfoil fins in the longitudinal direction;
wherein the plurality of airfoil fins, the central airfoil hub and the inner perimeter of the rocket engine combustion chamber define a plurality of apertures between adjacent airfoil fins at the exit opening, the plurality of apertures being configured to choke an airflow exiting the rocket engine combustion chamber and cause the airflow to expand supersonically along the plurality of airfoil fins to create thrust.

8. The aerospike rocket engine nozzle according to claim 7 , wherein the plurality of airfoil fins include four airfoil fins.

9. The aerospike rocket engine nozzle according to claim 7 , wherein each of the plurality of airfoil fins are fixed to the exit end of the rocket engine combustion chamber at the exit opening of the rocket engine combustion chamber.

Assignees

• Raytheon Co

Inventors

• Johnson Daniel K

Classifications

• F02K9/97 Primary

  Rocket nozzles (thrust or thrust vector control F02K9/80) · CPC title

• F05D2240/1281

  Plug nozzles · CPC title

• F02K9/30 Primary

  with the propulsion gases exhausting through a plurality of nozzles · CPC title