Low velocity injector manifold for hypergolic rocket engine

• US 9,404,441 B2
• Filed: 08/18/2008
• Issued: 08/02/2016
• Est. Priority Date: 08/18/2008
• Status: Active Grant

• Current Assignee: Aerojet Rocketdyne Of De, Inc. (L3Harris Technologies, Inc.)
• Original Assignee: Aerojet Rocketdyne Of De, Inc. (L3Harris Technologies, Inc.)
• Inventors: Brown, William S., Walczuk, Thomas M., Noble, Rodney, Dodd, Frederick
• Primary Examiner(s): MEADE, LORNE EDWARD
• Application Number:
  US12/193,679
• Publication Number:
  US 20100037590A1
• Time in Patent Office: 2,906 Days
• Field of Search: 602/001, 602/11, 602/15, 602/57, 602/58, 607/25, 607/39, 607/42, 181/213, 181/220, 431/114
• US Class Current: 1/1
• CPC Class Codes: F02K 9/44 Feeding propellants
• CPC Class Codes: F02K 9/52 Injectors
• CPC Class Codes: F05D 2250/232 conical

Abstract

A fuel manifold for a thrust chamber assembly includes a main fuel chamber which is generally frustro-conical in shape. The main fuel chamber provides a resonance frequency that is different than an acoustic resonance frequency of a combustion chamber.

Claims

1. A rocket engine comprising:

a combustion chamber having an acoustic resonance frequency; and

a fuel manifold having a resonance frequency that is different than said acoustic resonance frequency, said fuel manifold including a main fuel chamber that is frustro-conical in shape, said main fuel chamber being directly coupled to the combustion chamber.

2. The rocket engine as recited in claim 1, wherein the rocket engine is a hypergolic rocket engine, the fuel manifold having a plurality of fuel injector passages for communicating fuel into the combustion chamber, further including an oxidizer manifold having an inlet and a plurality of oxidizer injector passages for communicating oxidizer into the combustion chamber.

3. The rocket engine as recited in claim 1, wherein said fuel manifold includes an inlet passage with multiple outlet lobes in fluid communication with said main fuel chamber, each of said multiple outlet lobes disposed at a discrete vector to uniformly distribute a fuel through the fuel manifold.

4. The rocket engine as recited in claim 1, wherein said fuel manifold further includes an inlet passage and a plurality of fuel injector passages that extend from said main fuel chamber to said combustion chamber, said plurality of fuel injector passages each having a uniform cross-sectional area between said main fuel chamber and said combustion chamber.

5. The rocket engine as recited in claim 1, wherein said fuel manifold further includes a plurality of fuel injector passages extending from said main fuel chamber to the combustion chamber and a single common inlet passage leading into said main fuel chamber.

6. The rocket engine as recited in claim 5, wherein said single common inlet passage splits into a plurality of passages that open into said main fuel chamber, said plurality of passages being open to each other.

7. The rocket engine as recited in claim 1, wherein said fuel manifold includes a single common inlet passage and a plurality of fuel injector passages that extend from said main fuel chamber to said combustion chamber, the single common inlet passage intersecting a conical outer wall of said main fuel chamber.

8. The rocket engine as recited in claim 1, wherein said main fuel chamber includes a conical outer wall and a conical inner wall.

9. The rocket engine as recited in claim 1, wherein said fuel manifold is defined annularly about a central axis and includes a single common inlet passage and a plurality of fuel injector passages that extend from said main fuel chamber to said combustion chamber, and said single common inlet passage is displaced from said axis.