Page 240 - MEMS and Microstructures in Aerospace Applications

Osiander / MEMS and microstructures in Aerospace applications DK3181_c011 Final Proof page 231 1.9.2005 12:31pm

Micropropulsion Technologies 231

pulsed plasma thrusts (mPPT) have been shown to be good candidates for many
missions requiring approximately mN-s to mN-s impulse bits; however, these
devices are pulsed, and shot-to-shot variation can sometimes be significant.
Besides performance, another significant parameter is the system mass. Some
of these technologies can benefit from the use of MEMS, which enables reduction
of the mass of the thruster itself. Nevertheless, the thruster itself is only one part of a
complete propulsion system, and in many cases, a small thruster requires additional
overhead mass like PPU, tanks, valves, etc. to function properly. This prompts the
question: How good is a MEMS thruster with a total mass of a few grams, when the
PPU mass cannot be accommodated within the spacecraft budget?
Also consider that the mass of a propulsion system consists of the dry mass and
the amount of propellant that needs to be carried. Mission parameters that define the
requirements for propulsion systems include total D-V, required payload or struc-
ture of the spacecraft, and time allocated for the mission.
The amount of propellant needed depends on the D-V requirements and the
exhaust velocity of the propulsion system, which has been expressed by Tsiolk-
ovsky in the famous rocket equation as shown in Equation (11.1): 5

M 0
DV ¼ v e ln (11:1)
M 0 M P
with M 0 and M P being the initial mass of the spacecraft and the amount of propellant
needed, respectively, and v e describing the exit velocity. From this equation it is
obvious that for a given D-V and spacecraft mass, the amount of propellant required
depends on the propellant velocity. The higher the velocity, the less the propellant
needed. Electric propulsion (EP) systems have been shown to provide high exit
velocities ranging from 10,000 up to 100,000 m/sec, whereas chemical propulsion
systems are usually limited to exhaust velocities between 500 and 3000 m/sec.
Therefore, at first glance, the choice seems obvious.
Apart from the propellant, both classes systems include additional mass over-
head. In the case of chemical systems, this will include tanks and valves. In the case
of EP systems a PPU is needed. The mass of a PPU has been shown to be a function
of the average power they can handle, thereby defining a specific mass a, which
commonly scales as 30 g/W. With EP thrust-to-power ratios averaging approxi-
mately 10 mN/W, the importance of taking the PPU mass into account becomes
obvious. Looking at an example it can be shown how a chemical system can be
more advantageous than an EP system despite its much lower exhaust velocity.
Assuming a total spacecraft mass of 5 kg, the amount of propellant needed for a
DV of 300 m/sec can be calculated to be 15 g for a v e of 100,000 m/sec and 696 g for
a v e of 2,000 m/sec. The average thrust T needed depends on the duration of the
mission Dt, as shown in Equation (11.2).

M P v e
T ¼ (11:2)
Dt
For an EP system the mass of the power supply is given by Equation (11.3),

© 2006 by Taylor & Francis Group, LLC