Research
Introduction
Though not as widely used on CubeSats, a large amount of smallsat utilize a propulsion system for reasons such as orbit raising, collision avoidance, and end-of-life disposal. The flexibility offered allows for spacecraft to not be limited to the orbit a launch vehicle can provide.
When selecting a propulsion system for a mission, one important choice often made is between utilizing a chemical propulsion system or an electric propulsion system. Chemical propulsion systems release the chemical energy of their propellants through chemical reactions to generate relatively high amounts of thrust but at a low specific impulse.
On the other hand, electric propulsion systems use an external source of electrical power to accelerate their propellant to much higher exhaust velocities than chemical propulsion, thus attaining a higher specific impulse, but at the cost of significantly lower thrust. Each of these types of systems has its uses, with chemical being good at performing maneuvers quickly, and electric propulsion being able to perform much higher Δv maneuvers with a similar mass of propellant, but at the cost of taking much more time to do so.
Figure 1: Dual-mode propulsion concepts
One solution that allows for the best of both worlds, is a dual-mode propulsion system. This concept combines the high thrust of a chemical system and the high specific impulse of an electric system by employing both kinds of thrusters fed by a common propellant tank. Using this, mission planners no longer must compromise as much on the maneuvers that can be performed during a mission.
The APEx mission will verify the performance of a novel pulsed magnetoplasmadynamic (PMPD) thruster in space. APEx is designed around a purely electric propulsion system, but this system is well suited to be integrated into a dual mode propulsion system in the future, mainly due to its use of
a liquid chemical monopropellant called ASCENT.
ASCENT is an alternative to hydrazine designed by the AFRL to have reduced toxicity, higher specific impulse as a chemical propellant, and a low vapor pressure. This allows for the use of a porous fritted glass disk for passive propellant injection via capillary action. It holds ASCENT in a thin layer on the surface to be vaporized during a discharge of the thruster. This avoids the late-discharge vaporization problems of solid propellants, and the injection valve timing issues present with gaseous propellants.
Figure 2: PMPD thruster cutaway view
Another major design feature of the PMPD thruster is its coaxial thruster geometry which allows for reduced discharge voltages when compared to rail based Pulsed Plasma Thrusters (PPTs), which leads to a smaller and more efficient power processing unit. This resembles the geometry of magnetoplasmadynamic (MPD) thrusters and is where our thruster’s name is derived from.
Figure 3: PMPD thruster and propellant system
Figure 1: Dual-mode propulsion concepts
One solution that allows for the best of both worlds, is a dual-mode propulsion system. This concept combines the high thrust of a chemical system and the high specific impulse of an electric system by employing both kinds of thrusters fed by a common propellant tank. Using this, mission planners no longer must compromise as much on the maneuvers that can be performed during a mission.
The APEx mission will verify the performance of a novel pulsed magnetoplasmadynamic (PMPD) thruster in space. APEx is designed around a purely electric propulsion system, but this system is well suited to be integrated into a dual mode propulsion system in the future, mainly due to its use of
a liquid chemical monopropellant called ASCENT.
ASCENT is an alternative to hydrazine designed by the AFRL to have reduced toxicity, higher specific impulse as a chemical propellant, and a low vapor pressure. This allows for the use of a porous fritted glass disk for passive propellant injection via capillary action. It holds ASCENT in a thin layer on the surface to be vaporized during a discharge of the thruster. This avoids the late-discharge vaporization problems of solid propellants, and the injection valve timing issues present with gaseous propellants.
Figure 2: PMPD thruster cutaway view
Another major design feature of the PMPD thruster is its coaxial thruster geometry which allows for reduced discharge voltages when compared to rail based Pulsed Plasma Thrusters (PPTs), which leads to a smaller and more efficient power processing unit. This resembles the geometry of magnetoplasmadynamic (MPD) thrusters and is where our thruster’s name is derived from.
Figure 3: PMPD thruster and propellant system
Mission Objectives
MO.1
Operate the PMPD propulsion system in LEO to execute an orbit-raising or lowering maneuver.
MO.2
Determine the key performance metrics of the PMPD system in LEO including impulse bit, specific impulse, and total efficiency of power to thrust conversion.
MO.3
Continuously record the propulsion system health metrics during thruster operation including average power consumption, propellant injector temperature, propellant tank level, and current/voltage waveforms for the discharge and igniter circuits.
Overall CONOPS and Experiment Plan
FE.1 (1.5 hours)
Thruster checkouts
* Fire single prograde pulse
* Measure thruster health metrics (voltage and current waveforms)
* Downlink results + wait to proceed
FE.2 (24 hours)
Short-duration performance characterization
Figure 4: Overall Mission CONOPS
- Repeatedly pulse thruster and coast to recharge
- Gather data to allow for performance metrics (impulse bit, specific impulse, total efficiency)
Figure 4: Overall Mission CONOPS
Mission Modes
Deployment/Setup
* Power on immediately once separating from launch vehicle
* Wait 30 minutes using internal timer after deployment for antenna deployment
* Attempt beaconing a small data packet after 45 minutes, repeating once every 6 minutes
* Detumble until reaching angular velocity under 1 deg/s followed by complete rotations on all axes to complete ADCS test
* Enter standby mode
Standby
* Actively point solar panels towards sun
* Transmit data for 1 minute every 10 minutes on incident power generation and battery level
* Log system telemetry every minute to onboard memory
* Change mode after command from ground
Downlink
* Occurs approximately once a month when data logs are approaching full
* After ground command, transmit when above Purdue’s ground station
* After packets are acknowledged from ground, data is removed from satellite storage
* Process is estimated to take roughly 3 days
* Return to Standby when complete
Safe Mode
* Triggered by passing below 42% battery charge, or if software errors are detected
* Satellite transmits packet to ground indicating critical condition before entering safe mode
* Satellite stops regular beacon packets and disables ADCS system to conserve power
* Once above 48% power, satellite will wait for ground command to exit safe mode
End of Life
* Spacecraft enters end of life on receiving command from ground station
* Spacecraft passivated by placing batteries in fully discharged state
* Satellite shuts down all systems and passively deorbits into Earth’s atmosphere
CubeSat Design
The APEx mission satellite is still in the early stages of design and only rough layouts of the spacecraft have been created. A more complete CAD model is in development currently but is still pending certain design decisions. It will consist of a 6U satellite bus with the PMPD payload mounted to direct thrust through the spacecraft’s center of mass. This system, with the thruster and its propellant system is currently estimated to require 2U of volume. A GPS patch antenna will provide position telemetry for determining the impulse bit of the thruster, voltage and current sensors in the thruster power supply will monitor thruster power usage, and a combination of a microflow meter and thermal propellant gauging technique are planned to collect propellant data. A 6U fixed solar array will provide power with an 80 W*hr. battery for storage, and four deployable UHF turnstile antennas will provide a link to the Purdue ground station.
Figure 5: Preliminary spacecraft geometry