Okay. This took me a bit because I'll admit I'm in a bit over my head here. I'm not a professional mathematician or physicist and it's been 25 years since I've had a course in either, and orbital mechanics is way beyond anything I did then. But here goes:
If I missed an orbital altitude I apologize, but I started with 2000 km. The orbital velocity at 2000 km is 6900 m/s.
v=sqrt(3.986*10^14 /8.371*10^6) - this is G*Mass of Earth/ Radius of orbit + Radius of Earth.
You wanted to change to orbit by 200 km, but weren't very specific, so since I was in over my head anyway, I decided on the simplest maneuver, a Hohmann Transfer.
The orbital velocity of the desired orbit is 6800 m/s, using the same equation above, but with an additional 200 km added to the radius.
The Hohmann Transfer uses an elliptical orbit that shares a perigee with the first orbit and an apogee with the second orbit.
We need the period of the transfer orbit to calculate the correct velocity, and the orbital period of the transfer orbit (Pt)=sqrt(4pi^2*a^3/G*Mass of Earth) where a=the semi major axis of the ellipse we're using, in this case, 8471 km.
Plug in the numbers and the period of the transfer orbit is 7759.15 seconds.
The velocity we need to reach at perigee, where we enter the transfer orbit is calculated vp=(2pi*a/Pt)*sqrt((2a/radius of the first orbit)-1
After running numbers, the velocity is 6941.10 m/s.
The difference between the current orbital velocity (6900 m/s) and the transfer velocity at perigee (6941.10 m/s) is called the delta V, and is the change in velocity that we need to achieve the transfer orbit.
The velocity of the transfer orbit at apogee (using the same velocity equation above but with radius of the final orbit instead of the first one) is 6779.13 m/s.
The difference between the target orbit velocity (6800 m/s) and the transfer orbit velocity at apogee (6779.13 m/s) is the delta V needed to adjust from the transfer orbit into the final orbit.
The first delta V is 41.10 m/s, and the second is 20.87 m/s.
Delta V is proportional to the logarithm of the fueled to empty mass ratio and to the specific impulse of the engine.
This was more than I wanted to do, so I used a Delta V calculator at
http://www.strout.net/info/science/delta-v/intro.html to do the work for me.
I wasn't sure of the specific impulse of satellite thrusters, so a little digging led me to
https://engineering.purdue.edu/~propulsi/propulsion/rockets/satellites.htmlI used your 500kg for the mass of the satellite, ballparked 300 sec as reasonably within range from the thruster specs on Purdue's page, and then reduced mass for the post-burn entry until I got the delta-v I needed.
The result was just under 7 kg of fuel, for the first maneuver and just over 3.5 kg for the second, for a total fuel consumption of 10.5 kg of fuel. I should probably have adjusted the satellite's mass down to account for the missing fuel after the first burn, but it probably won't amount to too much.
10.5 kg seemed high to me, but in actuality, the approximately 62 m/s delta-v we needed for our maneuver is a bit more than a year's worth of typical station keeping for a satellite, depending on its orbit. 10-11 kg every year doesn't seem that high to me as a lay person, but I admit I'm well out of my depth here.
If I made any errors, I apologize and will happily take any criticism and make corrections as necessary. Like I said earlier, I'm in over my head and did the best I could in my first attempt at orbital maneuvers. Hopefully I get a passing grade.