German Aerospace Center (DLR)
in the Helmholtz Association
DLR SART Web Portal
http://www.la.dlr.de/ra/sart/projects/eol/eol.php.en
Date: 18-Mar-2010
German Aerospace Center (DLR)
in the Helmholtz Association
DLR SART Web Portal
http://www.la.dlr.de/ra/sart/projects/eol/eol.php.en
Date: 18-Mar-2010
| Background | Objective | Propulsion Systems | Propulsion System Requirements |
| Assessment Criterion | Assessment Example | Publications |
The
historical practice of abandoning spacecraft and upper stages at the end of
mission life has allowed roughly 2 million kg of debris to accumulate in orbit.
The uncontrolled growth of the space debris population has to
be avoided in order to enable safe operations in space for the future. If no
countermeasures are taken, the number of debris particles will grow with a growth
rate in the order of 5% per year. Due to the very high relative velocities in the
order of km/s, even very small particles in the millimeter size range can destroy
spacecraft subsystems and thus eventually lead to the loss of the complete spacecraft.
In principle two categories of measures to reduce debris growth can be distinguished:
Simulations have shown, that a real reduction of the debris population can only be achieved with very far-reaching measures. The only effective way to limit the growth of the orbiting debris is to systematically remove satellites and rocket upper stages at the end of their mission life from the near Earth space. The natural cleaning of the Earth environment by the trajectory disturbances (substantially air drag) will no longer be sufficient in the future to keep the Earth environment in an acceptable status for the operation of manned and unmanned spacecraft.
The objective of this study was to provide an overview and assessment of propulsion-related methods to de-orbit spacecraft in LEO, or spacecraft which pass through LEO, to assess their applicability to different spacecraft-mission combinations and to establish a know-how basis on end-of-life de-orbit strategies. A major task was the identification of the most suitable propulsion system to perform the de-orbit maneuver for the different classes of spacecraft with mass ranges from below 10 kg to more than 2000 kg. A set of evaluation criteria was established for this purpose.
This study was performed in cooperation with OHB System and HTG under ESA contract No. 5316/NL/CK.
Propulsion systems investigated in this study are chemical propulsion systems (including cold gas) as well as solar-electric propulsion systems. The following table gives a short overview over principal characteristics of spacecraft propulsion systems. A more detailed description is given in the respective chapters.
| Advantages | Disadvantages | |
|---|---|---|
| Cold Gas | Simple Low system cost Reliable Safe |
Extremely low Isp Moderate impulse capability Low density High pressure |
| Mono Propellant | Wide thrust range Modulable Proven |
Low Isp (mostly) toxic fuels |
| Bi-Propellant (storable) |
Wide thrust range Modulable Proven |
Complex Costly Heavy Toxic |
| Solid Propulsion | Simple Reliable Low cost High density Low structural index |
One thruster per burn Total Impulse fix Currently not qualified for long-term space application |
| Hybrid Propulsion | Simple Modulable Low cost Reliable |
Not qualified Lack of suitable oxidizer for long-term mission |
| Electrical Propulsion | Very high Isp | Low thrust Complex Long maneuver time Power consumption |
The choice of a propulsion system for the de-orbit function is a complex task as design choices affect many parameters. Any potential system has to fulfil a number of minimum requirements, which are given by physical or operational requirements. The relevant parameters are listed subsequently. The impact differs considerably depending if a propulsion system for nominal mission life is already foreseen for the satellite and solely the performance has to be adapted or if a propulsion system has to be added.
The selection of the most suitable de-orbiting method is a task influenced by many parameters. On the one hand, the impact of the de-orbiting method on the overall spacecraft budgets (e.g. cost, mass, volume) shall be as small as possible, on the other hand, the maneuver has to be performed with a sufficient probability of success, in order to get the desired LEO-cleaning effect. For the assessment of the different de-orbiting methods, the following set of evaluation criteria was defined:
The IRS-1C S/C belongs to the medium-sized spacecraft and it is designed for Earth observation purposes. It has a nominal mass of 1250 kg without de-orbit device. The spacecraft is 3-axes stabilized and uses a hydrazine OCS.
The integration of a de-orbit function can be realized in two ways. Either the spacecraft OCS is adapted to the extended requirements, comprising the nominal mission and the de-orbit task ('Dual-use'), or a dedicated de-orbit OCS is installed in addition ('Add-on'). However the first option typically is not available for small spacecrafts as they oftentimes can perform their nominal mission without any OCS.
When assessing add-on systems, thrusters, PCU's if applicable, a propellant management system, tubing, cabling, (delta-) tank, and possibly additional power generating equipment is added to the S/C. In contrast, when assessing dual-use systems, many of the above listed items are readily available on the spacecraft and have solely to be adapted to the augmented performance requirements. In this case, only the adaptations have to be assessed.
For the comparative assessment of the de-orbiting concepts, two maneuvers generating a total ΔV of 100m/s respectively 200m/s were defined. This ΔV must be generated by the propulsion systems either in one single-burn maneuver or in a number of finite burns close to the apogee of the initial orbit. A ΔV=100 m/s is for example sufficient to perform a controlled de-orbit from an initial 400 km orbit. It is also sufficient to send a spacecraft from an initial altitude of up to 850 km to an orbit with a remaining lifetime of 25 years.
The impact of the implementation of an add-on de-orbit function on the IRS-1C for a de-orbit burn of ΔV=100 m/s is shown in the following Figure.
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