IFM Nano - IFM Nano Thruster Module

Objectives

The primary objectives for project IFM Nano are:

  • The adaptation of an established FEEP (field emission electric propulsion) ion thruster design for use on micro and nano satellites (particularly CubeSats). Such spacecraft could use the thruster either for continuous drag compensation in LEO or even as an efficient main drive.  
  • A full characterization and test campaign including functional and environmental tests to verify the design’s specifications and its fitness to operate reliably in orbit.

Challenges

  • Vibrational tests shall demonstrate the fitness of the design to withstand launch stresses.
  • In the present design, only COTS components have been used for thermal sensors and electronics. As it is not known how the radiation and thermal vacuum environment in orbit will affect the lifetime of these parts, irradiation and thermal tests in compliance with ECSS standards have to be performed to verify the suitability of all parts for the intended operation.
  • Spacecraft charging due to the emitted ion current is compensated by electron-emitting cathodes whose lifetime is known to be rather limited. In the absence of other charge neutralization measures on the spacecraft, breakage of the cathode would limit the effective thruster lifetime. Redundancy is used to mitigate this problem.
  • The long-term performance and degradation of the present design are unknown. Detailed lifetime tests are being performed to identify and counter the most important sources of errors.

Benefits

The IFM Nano Thruster addresses the urgent need of a propulsion system for micro- and nano-satellites: its wide range of thrust (1μN to 1mN), the excellent throttlability, and a high specific impulse (up to 5000 s), allow to significantly increase the mission range of such satellites in low orbits. The high Isp on the other hand allows for very high delta-v manoeuvres at a high propellant mass utilization efficiency (80%). The modularity and the small volume (less than 1 dm³ including propellant and electronics) and its light mass (0.8 kg), make the thruster suitable for all small satellites from 1 to 500 kg. The combination of high Isp with medium very well controllable thrust levels in a small and light package makes the IFM Nano thruster a strong competitor for existing colloid, cold gas, or Hall effect thrusters.

Features

  • Design based on a decade-long ESA-approved evolution and development process.
  • Thrust level 1 µN — 1 mN at Isp up to 5000 s, capacity of 10kNs and a power to thrust ratio of <80 W/mN.
  • Monolithic outer structure in a 10x10x10 cm³ package weighting less than 1 kg, including electronics, tank, and thruster.
  • Power interfaces 3.3V and 12V, control via I2C or UART
  • No moving parts, all-solid in OFF state (during launch)
  • Fully tested according to relevant ESCC standards
  • COTS components tested for radiation and thermal compliance 

System Architecture

Mechanically, the thruster comprises of a tank filled with the metal propellant and rigidly connected to the crown emitter. The extractor anode, which creates the top plate, is separated via an isolator from the electronics compartment and the tank. An optional housing encloses all parts but is not required for mechanical stability.

Electrically, the thruster is commanded via the spacecraft bus via I2C or UART. The central command and control module is connected to the independent power supplies for the heater, the two cathodes (two separate supplies), the emitter, and the extractor. Towards the spacecraft, the controller offers full access to all parameters as well as a high-level interface to set thrust levels.

Plan

The project started after a PDR. Before that, prior work was performed to estimate the major risks of the project. All work before kick-off has been reviewed in a PWR. The initial plan contained design, construction and test (mechanical, thermal, lifetime) of the thruster and PPU. After investigation of cathode lifetime limiting processes, and a change of the designated target orbit from LEO to MEO and above, a CDR was implemented to perform radiation testing, checking of components against COTS radiation databases, and selection of a cathode. Actual flight tests are foreseen for follow-up projects

Current status

Most work packages from the original project plan (construction, lifetime, vibration, thermal, and performance test) are finished or about to be finished. An exception is the floating and cathode performance test to be commenced in 2018 and finished before CO in mid 2018. Aspects added in the CDR, namely the COTS database check, radiation testing and cathode selection are being worked on at the moment with TRB 3 coming up in November 2017 and TRB2 together with CO in mid 2018. At the moment, no blocking technical issues are known. An issue of concern is the lifetime of the cathode being 500h only (as specified in the CCN).

Contacts

ESA Contacts

Status date

Friday, November 17, 2017 - 09:03