I did some reading on this subject while searching for the mechanical properties of composite materials. It turns out the allowable stress/strain you can use in your airplane is FAR LESS than either the strength of fibre or the laminates. Thus the approach of "make your laminate, load it with sandbags, and see when it breaks" is incorrect and dangerous in airplane design.
The degrading factors affecting the allowable strengths are humidity, temperature, openings and impact damage. Unlike metal, fatigue is usually not a limiting factor for composite. This explains why some certified composite airplanes claims unlimited life.
The worst case for tensile strength usually is 1/4 inch opening at cool temperature in dry atmosphere; the worst case for compressive strength usually is impact damage at hight temperature in wet atmosphere. The carbon fibre composite starts delamination at about 3rd or 4th layers under the surface. This delamination degrades the compressive strength signficantly but is not visible. To account for tool drop, runway gravels, hails and damage during production, the compressive strength with barely visible impact damage should be used.
These tests are expensive and the data are usually propriatory. AGATE and NCAMP attemp to share such data publicly. Unfortunately, some so-called "design allables" (both A-basis and B-basis) published by AGATE does NOT use sample with openings and/or impact damages, render them useless.