It's not really mentioned in the article but as a starting point they received all the work done for Fastrac [1], a NASA program to build small, cheap, expendable rockets - something like the Falcon 1. That got cancelled as NASA projects are wont to do but it made sense to use that design as a baseline. A clean sheet engine design is very complicated indeed - it's a very high dimensional, highly coupled optimisation problem, so a known good starting point is a very valuable thing to have.

Indeed the turbopump, which is one of the hardest bits of a rocket engine, is made by Barber Nichols, who made the turbopump for Fastrac and who make the turbopumps for the Merlin engines [2]. I've heard rocket engineers describe rocket engines as turbopumps with some extra plumbing. Perhaps a slight exaggeration but not far off, especially since their design is so tightly coupled and dependent upon overall engine parameters. The degree of coupling depends on the topology of the plumbing, or the 'cycle' of the engine, which I'll try and explain:

There are three kinds of rocket engine cycle (well, there are maybe more but these are the three that have been flown historically). The Expander Cycle, the Staged Combustion Cycle, and the Gas Generator cycle. I'll mention the last two.

Merlin, as the article mentions, is an example of a Gas Generator cycle [3]. In this cycle, you take off a little bit of fuel and oxidiser to burn outside the main combustion chamber, to generate some hot energetic gases that you can exhaust over a turbine. This spins the turbine up, which is connected to a shaft with a compressor on the other end. The compressor increases the pressure of the propellents so that they can be injected into the main combustion chamber. This assembly (turbine, shaft, compressor) is called the turbopump. It's necessary because the engines require very high flow rates to get the thrust they need, and that has to be at a high pressure - higher than the pressure of the combusting gases inside the combustion chamber, else you wouldn't be able to inject it!

Back to the bleed-off to drive the turbine. You usually don't want a perfect stoichiometric mix of fuel and oxidiser for this, or even close, because it generates extraordinary hot gases that no turbine would last long in (The turbines are spinning at many tens of thousands of RPM usually so would be subject to much higher forces than the actively cooled walls of the main combustion chamber). For this reason you usually have a large imbalance of one propellent to the other to keep the temperature down. Usually you run with excess fuel, or 'fuel-rich', as the opposite - oxidiser rich - means you have hot oxidising gases which are harder on the metallurgy. I do know of some russian exceptions to this, though, where fuel rich would have left sooty deposits in the plumbing (The materials science employed in the turbines was apparently so witchcraft that when the US got intelligence of oxidiser-rich turbine precombustors, they thought is was deliberate counterintelligence from the russians to get them to waste billions researching the impossible). The gas generator cycle, as the article mentions, dumps this turbine exhaust overboard separately. The problem with this is that there's a load of uncombusted fuel in this exhaust, which you're just wasting, and this hits your rocket performance - the Specific Impulse ( I_{sp} ), as you're not getting as much bang out of a given mass of fuel as you could.

The answer to this is the Staged Combustion Cycle [4], where you also inject the exhaust of the turbine into the combustion chamber to finish off combustion. The performance of these engines is higher but the thermodynamic balance to design a working system is a greater challenge, and some of the engineering is a bit harder too. Staged Combustion engines are mostly russian, although the Space Shuttle Main Engines are a US-design example of Staged combustion.

SpaceX have been gradually and incrementally improving the Merlin's away from their simpler beginnings, and it's been pleasing to watch as an interested outsider. To bring it back to the OP question, "are all engines of that caliber that size or is this one special?", Merlin didn't particularly stand out in terms of power density in the early days, although it's been improving and improving. Now there is the Merlin 1D [5], which claims to have the highest thrust to weight of any rocket engine every made. One should take these claims with a pinch of salt as what counts as 'engine' and what counts as 'plumbing' and what counts as fuel tank is sort of open to debate and you can do some creative accounting to make your numbers look better. However, it's an impressive achievement regardless.

The metric that doesn't lie, from a performance point of view, is the mass fraction of the rocket - that is the fraction of the all-up, fuelled-up mass of the rocket on the pad that makes it into orbit. The higher the better - i.e. you can take bigger payloads for a given size rocket. Note this is just from a performance point of view, not an economics point of view.

However, increasing mass fraction will be important to SpaceXs staged aim of re-usable rockets. That's because as well as the payload, each rocket stage as to also carry the fuel it needs to land. I believe Falcon 9 can launch about 2% of its pad mass into orbit (i.e. the payload can be 2% of the total mass), and Musk reckons if that could be increased to about 4 or 5%, there's be enough margin to carry enough landing fuel and extra landing hardware like legs.

So my bet, just for fun (I'm not connected with SpaceX, this is just sideline speculation for the sake of interest) is that you might start to see development of a Staged combustion engine instead of gas generator, and a switch to methane fuel which, with LOX, has a slightly higher specific impulse than Lox/Kerosene which they currently use. Maybe 20 seconds extra (seconds being the unit of specific impulse), which is maybe 5-7% more than they might be seeing now, which is worth having. Methane is nice because it has a similar density to kerosene. Lox/LH2 has almost 50% better performance than LOX/Kerosene in theory but LH2 is of a very much lower density, so the tanks must be much bigger (the illustrates my earlier point about the slipperyness of engine-only thrust to weight as a metric - how much bigger and heavier are the tanks?).

[Edit: of course instead of building a higher performance engine you could just built a much bigger rocket with the same mass fraction and have the cargo be a smaller percentage of launch mass. probably cheaper than developing a staged combustion engine. I suspect the rocket science equivalent of more servers vs a rewrite in C]

This post has ended up being a bit longer than I thought it would. Hopefully of some interest if you're new to the subject.

[1] http://en.wikipedia.org/wiki/Fastrac_(engine)
[2] http://www.barber-nichols.com/products/rocket-engine-turbopu...
[3] http://en.wikipedia.org/wiki/Gas-generator_cycle_(rocket)
[4] http://en.wikipedia.org/wiki/Staged_combustion_cycle_(rocket...
[5] http://en.wikipedia.org/wiki/Merlin_(rocket_engine_family)#M...