Accession Number:

ADA190514

Title:

Composite Repair of Cracked Aluminum Alloy Aircraft Structure

Descriptive Note:

Final rept. Oct 1981-Apr 1984

Corporate Author:

AIR FORCE WRIGHT AERONAUTICAL LABS WRIGHT-PATTERSON AFB OH

Report Date:

1987-09-01

Pagination or Media Count:

40.0

Abstract:

A bonded composite patch repair of fatigue-cracked aluminum on aircraft has advantages over a standard bolted metal patch repair, such as no severe stress concentrations no bolt holes, fatigue-resistant patch, thinner patch, simple molding techniques, a sealed interface to help prevent corrosion, and usually no inspection NDI problems. The objective of this program was to determine the effect of composite patches on stress intensity and crack growth characteristics of aluminum. This was accomplished by studying metal thickness and patch parameters area, thickness, and ply orientation effects on crack growth rate of the composite patchaluminum specimen. Both room temperature and elevated temperature 250 F curing adhesives were studied. The testing procedure consists of edge cracking a 4-inch x 18-inch 2024-t3 aluminum specimen to a length of between 03 and 05 inch. The aluminum is then prepared for bonding, normally using the phosphoric acid non-tank anodize PANTA method, primed, and patched. The specimen then cycled to failure. Both constant amplitude and flight spectrum loading were used. Patch material for most specimens was 55214 boronepoxy. Results have shown thickness of the metal being repaired to be the most significant factor in the repair process. There was also a significant difference in results between constant amplitude and spectrum tests. 18-inch thick aluminum constant amplitude tests showed lifetime extensions of about 15 times, while 116-inch thick and 18-inch thick spectrum loaded-specimens showed extensions of about 15 and 7 times, respectively.

Subject Categories:

  • Aircraft

Distribution Statement:

APPROVED FOR PUBLIC RELEASE