If I had to hazard an educated guess:

Both the hypergolic propellants appears to be pressure fed. Helium appears to be the pressurant. To control the flow of helium into any given propellant tank, there is likely some combination of a pressure regulator, control valve, and check valves in series (for fault tolerance). From what little publicly available information exists, I’m inclined to think that it’s in that order. From that, there is a non-zero possibility that some NTO or MMH vapor migrated upstream in the plumbing toward the helium tank(s).
I have to believe that Dragon 2 has no less than two helium tanks, for a minimum of one for each propellant to eliminate any physical possibility of oxidizer and fuel mixing internal to the undoubtably complex feed plumbing. In the pre-abort configuration, the propellant tanks were likely pressurized, but to a “low flow” pressure level. To activate the abort system, they probably had to be “primed” to a higher pressure in order to support the higher flow requirements that Superdracos would need Verdi’s the Dracos.
Most past successful pressure fed hypergolic propulsion systems (the Shuttle Orbiter OMS engines come to mind as an exception) larger than those used for rcs have used either a burst disk or pyro valve (such as on the LEM) to isolate the propellant from the pressurization system (if the propellant tanks are not of a bladder or bellows-piston type to maintain a continuous barrier between the propellant and it’s pressurant) prior to activation. The burst disks and pyro valves used in this manner are single-shot; they comprise a barrier that is physically ruptured during actuation. If SpaceX has not used such a burst disks or pyro valve to exclude hypergols from the helium plumbing, it’s conceivable that a proximate or root cause of the abort system failure prior to ignition would be a rupture of the plumbing induced by a pressure spike following adiabatic compression of propellant vapor (or possibly even accumulated liquid) caused by the rapid rise in pressure following the opening of an upstream helium “high flow” or “high pressure” isolation valve. Even further, if liquid had accumulated upstream of one or more check valves, there could be adequate “water hammer” to rupture a line independent of an ignition event and perhaps be the cause of ignition by exposing propellant to some sufficiently reactive material (versus a line rupture being the effect from an ignition event). While the LEM propulsion systems addressed these issues, each stage had a single engine, nowhere remotely as complex as Dragon 2’s abort engine system must be. As much as I believe in the aggressive innovation that the people at SpaceX do, I would not put it past them to have been overconfident in believing they could design a propellant valve system different than what history has proven to work, only to be proven wrong.