Analysis, design, fabrication, and testing were performed to create a new joint design for potential use in attaching a domestically available carbon-carbon (C–C) nozzle extension to the turbine exhaust manifold of a J-2X engine. Various attachment methods were investigated for a C–C-to-metallic joint, including the use of higher-thermal-expansion ceramic matrix composites both mechanically attached and also integrally fabricated to the C–C nozzle extension. The goal was to determine the advantages and disadvantages of different material and joint systems in order to converge on a design for a domestic joint and nozzle extension design that resulted in all positive margins of safety.

Composite nozzle extensions (NEs) have been successfully integrated with existing engines in the past. There is, however, still a great need for improved methods to design and manufacture these components without having to rely on costly foreign suppliers. This metallic-to-refractory composite joint also poses a set of design challenges mainly associated with the minimal expansion mismatch between the metal and composite materials. The program objective was to design, fabricate, and test a subcomponent ceramic matrix composite (CMC) joint that would successfully join a domestically available two-dimensional C–C nozzle extension to the metallic turbine exhaust gas (TEG) manifold of the J-2X engine via a higher coefficient of thermal expansion CMC material at the metal-to-composite interface.

When two materials with significantly different coefficients of thermal expansion (CTEs) are bonded or mechanically attached, the material with the lower CTE will experience high in-plane tensile and very high interlaminar stresses, which are caused by the greater thermal expansion of the higher CTE material.

The integrally fabricated hybrid nozzle extension subcomponent is composed of two materials that are integrally fabricated, resulting in a single part. The subcomponent consists of a C–C NE extension of T-300 carbon fabric mat layered with commercial grade Nicalon SiC fabric over a transition region 2.0 in. (≈5 cm) in length, at the center of its 12-in. (≈30-cm) axial length. The SiC fabric makes up the transition flange region. The component was co-processed and experienced chemical vapor deposition of a carbon matrix, resulting in a single subcomponent part.

This innovation resulted in a design concept and subcomponent that mitigates the stress associated with the CTE mismatch between the CMC transition flange and the metallic TEG manifold.

This work was done by Leslie Weller and Brian Sullivan of Materials Research and Design, Inc. for Marshall Space Flight Center. For more information, contact Ronald C. Darty, Licensing Executive in the MSFC Technology Transfer Office, at This email address is being protected from spambots. You need JavaScript enabled to view it.. Refer to MFS-33162-1.


NASA Tech Briefs Magazine

This article first appeared in the November, 2014 issue of NASA Tech Briefs Magazine.

Read more articles from this issue here.

Read more articles from the archives here.