Liquid-fuel rocket engines that utilize vortex flow fields to keep combustion-chamber walls cool have been investigated in computational simulations and experiments. In an engine of this type, the vortex flow establishes radial gradients of pressure and density that cause the lower-density hot combustion products to be confined near the central axis of the combustion chamber while cold gases yet to be burned are centrifuged to the combustion-chamber wall. Keeping combustion-chamber walls cool in this manner reduces or eliminates damage by heat and oxidation, thereby (1) eliminating the need for leak-prone cooling passages, (2) extending the operational lifetimes of combustion chambers, and (3) creating an opportunity to fabricate combustion-chamber walls more easily, using relatively inexpensive materials (possibly even lightweight composite materials) instead of expensive alloys.

For the most part, the combustion chamber in an engine of this type has a conventional appearance: it includes a head-end dome, a barrel section, and a section that converges from the barrel to a throat that opens into an expansion bell. A vortex flow field with a co-swirling, counter-flowing character is produced in the following way: One of the propellants (typically, oxygen at supercritical pressure) is injected circumferentially tangential through ports just forward of the junction between the barrel and the section that converges to the throat. The other propellant (typically, liquid hydrogen) is injected either from the head end or through a porous liner in the chamber wall.

The circumferential component of the injection flow forms a free vortex that spirals forward along the wall of the barrel to head end, where it turns inward to form a second vortex, concentrated along the axis, that flows out of the chamber at the aft end. If the fuel is fed through the wall, the upwelling oxygen burns the incoming fuel and carries it forward and into the axial vortex. If the fuel is injected from the head end, then it burns only along the axis. In both fuel-injection schemes, the radial gradients of pressure and density prevent hot combustion products from migrating out to the wall. Instead, the hot gases are buoyed toward the axis. Thus, the wall stays cool.

The propellant injectors can be relatively simple because unlike in other rocket engines, there is no need to atomize or mix the propellant fluids immediately upon injection. All of the necessary atomization and mixing is effected by the coaxial vortex flow. Yet another advantage of engines of this type is that they appear to be immune to combustion instability.

This work was done by William H. Knuth, Martin Chiaverini, and Daniel J. Gramer of Orbital Technologies Corp. for Marshall Space Flight Center. For further information, please contact the company at This email address is being protected from spambots. You need JavaScript enabled to view it..

In accordance with Public Law 96-517, the contractor has elected to retain title to this invention. Inquiries concerning rights for its commercial use should be addressed to

Orbital Technologies Corporation
1212 Fourier Drive
Madison, WI 53717

Refer to MFS-31477, volume and number of this NASA Tech Briefs issue, and the page number.