CC BY
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UDC 629.7
*M. Emanuelli, **A. Ronse, *C. Tintori, ***Professor V.I. Trushlyakov *Politecnico di Milan, Milano, Italy **TUDelft, Delft, The Netherlands / Belgium ***Omsk State Technical University, Omsk, Russia
A SPACE DEBRIS REMOVAL MISSION USING THE ORBITAL STAGE
OF LAUNCHERS
Introduction
Past years, several studies have been done to form an image of the current state and future evolution of these regions, showing that space debris mitigation is necessary. Amongst others, this led the IADC to develop a series of (passive) mitigation guidelines adopted in a 2007 UN resolution. These studies evaluating the long-term debris evolution have indicated that the debris density has reached such a high level, that there will be an ongoing increase of the number of debris objects, primarily driven by collision activity. This effect will manifest itself mainly in the Low Earth Orbit (LEO) region, due to a combination of high spatial densities, high relative velocities and a large number of object crossings.
The passive mitigation measures currently used do not suffice as an insurance of a stable debris environment for the next years endangering the future of the space mission not only in LEO. After the UN resolution, the space debris topic started to become an important matter of discussion in many international workshop and conference leading the idea of creating a space debris removal mission. Many space agencies put efforts in this and a number of projects have been started up.
A space debris removal mission is challenging because it requires a high level of expertise in various fields and also new technology not yet well proven in space. Thinking about this, the Omsk State Technical University (OmSTU) concept of using the launcher‘s orbital stage (and its unspent resources) combines more reliable technology to unquestionable economical benefit.
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The OmSTU concept explained in this document could be the answer to the need of space debris mitigation also because it offers the opportunity of collaboration between different agencies and companies, creating a joint project that will ensure the development in a relative short time.
The Space Debris Problem
In order to make the debris removal process as effective as possible, first it is assessed what type of debris should be chosen for removal and in what orbital regions the objects are situated. These choices are made using recent studies made on the current and future debris environment based on NASA‘s long-term orbital debris evolutionary model LEGEND [1].
The current and projected space debris situation: LEO and GEO
As can be seen in Figure 1, 89%of the ~950 operational satellites are either in a low earth orbit (LEO, 300-2000 km altitude) or a geosynchronous orbit (GEO, ~36000 km altitude). For this reason, these two regions form the first focus of a selection. Both have specific characteristics regarding the presence and evolution of space debris, described below.
LEO 46%
Other 3%
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Figure 9: Orbital zones of currently operational satellites. Source: Wright, 2010 [6]
• LEO: In the LEO region, satellites and debris elements are quite widely scattered in terms of altitude, inclination and ascending node. This, in combination with the fact that orbital speeds are considerably higher than in the GEO case, makes both the amount of crossings and the relative velocities of the bodies during these crossings averagely very high. The wide and random distribution of objects also implies that a system of graveyard orbits (as in the GEO case) is not possible. Another critical issue is that manned space-missions are performed at (low) LEO altitudes, making it essential that the risk of collision is minimized to the greatest possible extent. On the other hand, objects in LEO experience a certain amount of atmospheric drag which causes them to gradually spiral down towards the Earth, a process of which the duration depends on the object's altitude, area-to-mass ratio (A/M) and the solar activity. Thus, one could say that, in case no new debris would be added in LEO, on the long term, the region would become clear of debris.
• GEO: Unlike the LEO case, the majority of satellites at GEO altitudes are located in a confined ring in which geosynchronous motion is possible. Another downside is the fact that, due to the large distance, detection is limited to objects larger than ~1 meter. Also, the debris in GEO will orbit the Earth for many decades, as the stabilizing effect of the atmospheric drag is absent. However, as the semi-major axis and thus the circumferential area of geosynchronous orbits is so large, the spatial densities in the GEO band are still two or three orders of magnitude lower than in the most crowded regions of the LEO region [4]. Also, because of the uniform motion of all objects and their high altitude, relative velocities are substantially lower than in the LEO region, leading to less severe collisions. Finally, it should be noted that after their mission lifetime, the satellites can be in-
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jected into a quasi non-decaying graveyard orbit (no drag) and the debris is not hazardous for manned missions.
The discussed pros and cons are summarized in Table 1.
LEO and GEO characteristics concerning space debris
Table 5
LEO (300-2000 km) GEO (-36000 km)
Detectable objects Average relative velocity Debris concentration Number of crossings Drag decay Danger for manned missions >~10 cm ~ 10 km/s Higher Higher Yes Yes >~1 m ~0.5 km/s Lower Lower No No
As described in [8], the combination of a higher debris concentration, a large number of crossings and high relative velocities in the LEO region may lead to an exponential growth of debris objects by a future cascade of collisions. This effect is further addressed in section 2.2. In, [2], a comparison is made between the estimated collisional activities in LEO and GEO for the next 100 years. The results can be seen in Table 2.
Table 6
Collisions by orbit type. The numbers are averages from 30 Monte-Carlo runs using NASA‘s LEGEND simulation software, assuming a continuous 1996-2003 launch cycle and no post-mission
disposal option. Source [2]
Orbit type Number of collisions of 10 cm objects for the next 100 years. Catastrophic/non-catastrophic
LEO-LEO 21.9 12.6/9.3
LEO-GTO 0.6 0.3/0.3
GEO-GEO 1.2 0.4/0.8
Total 23.7 13.3/10.4
It should be noted that the current knowledge of sub-meter objects in GEO is not yet complete. This makes the values for the GEO-GEO collisions in the table presumably too low. However, the table shows the general trend that the probabilities of (catastrophic) collisions in LEO are substantially higher. Therefore, active removal of space debris is more vital for the LEO region than for GEO. The further selection of debris objects and critical orbits shall thus focus on the orbits between $300-2000$ km altitude.
Collisional Cascading
Although the name was introduced only later, the concept of collisional cascading was initially investigated in a 1978 NASA study, see [5]. One of the research‘s main conclusions was that col-lisional breakup of satellites would become a new major source of orbital debris, possibly exponentially increasing over time. The cause of this scenario is the fact that every intact satellite or other large body, has the potential to fragment into numerous smaller pieces due to a collision with a debris object. Many resulting fragments will, on their turn, impose a certain risk for the catastrophic destruction of another large orbital body, and so on. Once a certain debris density has been reached, this effect causes the debris population to continue its growth, even without the addition of any new
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objects. More recent studies have confirmed the fact that the critical spatial density has already been reached at certain LEO altitude regions, see [2], [3] and [4].
Figure 2 shows the result of a simulation using LEGEND, based on a ||no-launches after 20061| scenario. It can be clearly seen that the total number of debris will continue to grow, caused by an ongoing process of collisions. This clearly indicates that in the LEO debris environment, collisional cascading will be a major source of debris accumulation in the next decades. In section2.4, it is determined which orbital altitudes form the main source of this instability effect.
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Figure 10: Future model of amount of large debris objects in the LEO region, based on a "no-launches after 2006" scenario. Source: [7]
LEO debris size and characteristics
Generally speaking, space debris in LEO can be divided into three categories in terms of size, risks and tracking capabilities, as can be seen in Table 3.
Characteristics of different sized debris in LEO. Data based on [3] and [6] Table 7
Size Potential risk upon collision Defection Number Mass fraction
> 10 cm Complete destruction Can be tracked ~14000 >95%
1-10 cm Partial/total destruction Only partially tracked -370000 <5%
<1 cm Damage, can be shielded Not tracked >10 million
An important fact is that although the number of debris objects is many times higher for the small-sized debris, nearly all the mass of the LEO debris is concentrated in the large objects. One can also conclude that at current time, the most dangerous debris for operating satellites are the many untraceable 1-10 cm objects which could partially or totally destroy the satellite. Being accurately tracked, the larger objects can be accounted for by performing a preliminary avoidance ma-
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neuver (it should be noted that not all operable satellite have active means for undertaking such action). However, several studies have indicated that, on the long term, the large >10 cm objects will play a more critical role. As they contain such a majority of the mass, they can form the source of giant amounts of new, smaller, debris when they burst into piece upon high speed [5] or other causes of defragmentation. It are these potential collisions which form the danger for a cascade of debris objects as depicted in Figure 2.
This mission aims at creating a sustainable solution for the space debris problem. In the first place, this involves the avoidance of any instability in the space debris environment as described in
[3] and Section 2.2. To do so, the debris removal process of this project will focus on the large debris objects in the LEO region.
In [3], a study is done on the type of objects which are predicted to be involved in future collisions. The results are tabulated in Table 4 and as shown, a distinction is made between intact objects and fragments. Notice that fragment-fragment collisions occur very seldom and intact-fragment collisions happen about twice as much as intact-intact hits. However, as the latter are generally much more severe and contain more mass in the process, both types of collisions have a similar contribution in the future debris population.
Table 8
Collisions by object type. The numbers are averages from 150 Monte-Carlo runs using NASA‘s LEGEND simulation software, assuming no new launches after December 2005 (bottom line prediction). Source: [3]
Object type Number of collisions for the next 200 years. Catastrophic/non-catastrophic
Intact-intact 4.9 4.8/0.1
Intact-fragment 10.8 4.5/6.3
Fragment-fragment 1.4 1.0/0.4
Total 17.1 10.3/6.8
It can be concluded that the most effective method for active debris removal, is one in which the most problematic intact objects are targeted, as in that way the risk of both intact-intact and in-tact-fragment collisions is diminished. This also has several other advantages, as intact objects have accurately known size, mass and shape characteristics, decreasing the number of assumptions and simplifying disposal. The amount of objects and the respective orbits for which removal is the most crucial and effective, are assessed in the next sections.
Critical LEO orbits
As concluded in the previous sections, the most crucial debris to remove for ensuring a longterm stable state of the future space environment, are the >10 cm objects in the LEO region. The
remaining question is which orbits in LEO are the most in need of active debris removal. In a study done in 1991 on the collisional cascading effect, it was concluded that the 900-1000km region had passed the critical spatial density, with the 800-900km at close risk [4]. More recently, studies were performed for predicting the collisional activity in the next centuries using NASA‘s LEGEND model based on the past and current debris environment, see [2] and [3]. These studies will form the basis of further orbital selection.
Altitude
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In Figure 2.3, a graph is depicted describing the normalized distribution of predicted impact occurrence as a function of altitude. It shows the percentage of catastrophic impacts per altitude band in the next 200 years and was acquired using the LEGEND debris evolution model. Note that the simulation took into account a ||no-future launches! scenario. This is not a very realistic assumption, but it forms a reliable bottom line of the future collisional environment.
A peak concentration of collisions can be seen between 900 and 1000 km altitude. These collisions will, even in a no-future-launches scenario, produce a continuous growth in debris at that region in the next 200 years. This figure also indicates that the collisional cascading effect described in section 2.2, is dominated by the collisions in the 900-1000 km altitude region.
The debris model used for this study was based on the 2005 LEO state, i.e. before the satellite destruction of the Fengyun 1c weather satellite in 2007 and the hypervelocity collision of the Ir i-dium 33 and Cosmos 2251 satellites in 2009. Extensive studies were done on the short- and longterm consequences of these breakup events, see [8] and [9]. In these studies it is concluded that, although the events resulted in a drastic increase in fragments at different alt itudes, the long-term effect on the collisional activity in LEO is limited. The explanation for this is that the large area-to-mass ration (A/M) of most fragments, results in a rather rapid decay. However, the cumulative effect of both fragmentations does produce a significant shift in the collisional probability at the altitude bands containing the highest fragment concentration of each event (Fengyun 1c: ~850km, Ir i-dium-Cosmos: ~800km) in the next 50 years. Therefore, the critical altitude band which forms the prior region for this project‘s active debris removal is extended to 800-1000 km.
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Figure 11: Normalized distribution of predicted catastrophic collisions
as a function of altitude. Source: [3]
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Figure 12: Spatial density distributions, for objects 10 cm and larger, at the end of 2005 and 2205. Source: [3]
Inclination
When looking at the different inclinations at which the current (detectable) LEO debris is orbiting, one can notice that especially the higher inclinations (60°-110°) are crowded (see Figure 5). This is a direct result of the high number of past and present satellites which use these zones to fulfill their mission goals. Note that the peak between 90° and 100° is due to the high amount of fragments of Fengyun 1c at the mean inclination of ~99°. In the 100-year collisional simulation performed in [2], it was also studied at which inclinations the predicted collisions occur. The results can be seen in Figure 6 and confirm the trend that the main risk manifests itself in the 60°-110° inclination range. Efforts for active removal of debris should thus focus on objects in those orbits.
Figure 13: Amount of detected LEO objects per inclination band for 800-1000km altitudes. Results are based on SSN data of October 2010 [10] and [11]
Figure 14: Inclination distribution for objects involved in future LEO collisions. Source: [2] Other orbital elements
Unlike the semi-major axis and inclination, no particular trend can be seen in the right ascension of the ascending node (RAAN) of the current debris environment and the future collisions. This is because the Earth RAAN is a permanently evolving parameter -resulting from the oblateness of the Earth, described by following equation:
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Here, J2 describes the gravitational perturbation due to the flattening of the Earth, Re is the Earth's radius, a the orbit‘s semi-major axis, e the eccentricity and i the inclination of the orbital plane. As a result, only slight differences in the semi-major axis of two objects cause the ascending to develop in varying ways. This effect, together with the different periodicity at various altitudes, causes debris clouds from collisions or other fragmentations to evolve towards a full-enclosure of the Earth over time. A perfect example of this effect is the Iridium-Cosmos collision described earlier. The post-collision orbital paths of the different debris objects are shown in Figure 7. Note that
after six months, the debris has formed a cloud nearly enclosing the whole Earth.
Figure 15: Evolution of debris cloud from Iridium 33 and Cosmos 2251 satellite collision in 2009. Image (a) shows the situation at the time of the collision, (b) depicts the orbits of the various debris
objects 6 months later. Source: [12]
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2
It may be concluded from the previous sections that, in order to avoid any collisional instabili-
ty of our future space environment, active debris removal will be necessary for certain orbital regions. Based on studies of the present debris environment and simulations of the future collisional activity, it was decided that the most effective way to ensure the long-term absence of an exponential debris growth, is to focus on the removal of large, intact objects at 60°-110°'inclined, circular orbits with an altitude between 800 and 1000 km. The amount of debris which should be removed, as well as the selection of candidate objects, is a subject of subsequent sections.
Figure 16: Amount of detected LEO objects a function of eccentricity. Results are based on SSN data of October 2010 [10] and [11]
Removal Frequency
Another primary question in the design of a debris removal system is how frequently a debris object should be collected and disposed. This factor will mainly drive the amount of launches on which the system will be implemented. Also, if this factor is lower, the amount of launches to choose from is higher, due to which an ‘ideal launch‘ can be selected more easily every time, bringing the initial orbit of the orbital stage as close as possible to a target debris object. Consequently, this limits the amount of AV and thus propellant necessary for the far-guidance phase. Higher removal frequencies will force system implementation on more launches, not always as ideally close to a target object, thus increasing propulsive needs. In a study performed by NASA, again using their LEGEND debris evolutionary model, it was investigated what effect active debris removal had on the future collision activity and thus debris instability [13]. The used models assume an ongoing removal program starting in 2020, compliance to the UN post-mission-disposal guidelines (removal within 25 years) and a persistent repetition of the 1999-2006 launch cycle. Different cases were analyzed; a no-removal scenario , leading to an exponential growth in debris (as previously described), active removal of two objects per year and active removal of five objects per year. The result is depicted in Figure 9. It can be seen that the latter case (active removal of five objects per year) leads to a stable environment. This will thus be the target frequency of the debris removal mission described in this report.
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Figure 17: Comparison of three different scenarios. From top to bottom: post mission disposal (PMD) only, PMD and ADR of two objects per year, and PMD and ADR of five objects per year,
respectively. Source: [13]
Concept Establishment
Phase one - Development of de-orbiting means for rocket bodies
In the so-called ||phase 1|| of the space debris removal project, OmSTU (in a cooperative connection with the polytechnic university of Milan) is currently addressing the design of an effective method for the de-orbiting of launch vehicle separating parts and orbital rocket bodies after payload deployment. The upper stages of current satellite missions normally stay in quasi the same orbit as the payload, while being of no use anymore. Moreover, as explained in previous section, their high mass and, in some cases, high collisional risk make them prominent targets for disposal. Initially, different removal concepts were investigated:
• Passive systems, such as the use of inflatable elements for the increase of area-to-mass ratio. However, this technique had major downsides because of its vulnerability, insufficient practical experience and the fact that this technique makes use of the drag effect, which is extremely small at altitudes larger than 750 kilometers.
• Tether systems, in which a long, deployable tether is used for gradually decreasing the orbital altitude. This technique was also found to be very new, missing solid practical experience. Moreover, the deployment and control of the tether attitude and the potentially high mass were considered large obstacles.
• Active on-board de-orbiting system (ADS), in which a small engine is installed on the rocket body, which can provide the necessary AV impulse to remove it from its orbit. This technique showed the most promising results, due to the extensive practical experience in different types of rocket engines, its high power and thus rapid disposal potential, the high reliability together with a relatively low mass.
After studying the ADS possibilities, the design of a new generation gas propulsion engine was initiated, utilizing the liquid propellant residues in the rocket body‘s tanks. If used on specific parts of launch vehicles, this technology could lead to following advantages:
• Disposal of the rocket body after payload separation.
• Controlled entry into Earth‘s atmosphere.
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• Controlled descent of first stage(s) of launch vehicle onto preset target locations.
• Full expenditure of (potentially poisonous) rocket propellant.
• Possibility of straightforward upgrading of current launch vehicles.
• Low development cost in comparison to other debris removal alternatives
A conceptual drawing of the use of the ADS on a rocket body is shown in figure 32a. It was
traced that in current satellite launches up to ~5% of the liquid propellant and ~40% of the electric capacity of the orbital stages are unspent. These energy sources can be used for a de-orbiting maneuver of the rocket stage after payload deployment. Following characteristics were identified as key elements in the design:
• Maintain engine operation in the conditions of zero-gravity and an uncertain position and state of liquid components in tanks.
• Ensure complete expenditure of propellant
• Maximize specific impulse
• Enable multiple restarts
• Keep development and integration costs within 10% of total launch vehicle cost.
In order to ensure the operation in zero-gravity conditions, a concept was established in which the residual liquids are gasified using a heat carrier before being used for propulsive means. Although for the time being mainly focus is lain on the design of the gasification and propulsive subsystems, the on-board de-orbiting system will also include a telemetry and control device. As depicted in figure 32 (b), the design process of the gasification-propulsion system is split up in a number of steps to ensure proper functionality. The study has already led to extensive research concerning the treatment of the residual liquids (gasification using heat-carrier and acoustic effects, additive addition) and the detailed study of the combustion process.
Figure 18: Conceptual drawing of the ADS integration on separated rocket body. (b) Flow diagram for the design process of gasification-propulsion system. Source [39]
Based on rough mass and financial budget estimates, the concept seems quite promising when compared to other debris removal systems. The dry mass was computed to amount around 0.5%-0.7% of the upper stage‘s rocket body‘s dry mass. Moreover, the development and upgrade costs of current rocket bodies would not exceed 7% of the total stage costs.
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Phase two - Removal of rocket bodies from nearby orbits
During the development and study of the active de-orbiting system described in the previous section, it was reasoned that the accumulated kinetic energy of the rocket body, together with the power and potential of the ADS, might as well be used for the collection and removal of other debris objects which are orbiting nearby. Focus will be lain on the objects identified in the background section, being rocket bodies at altitudes of 800-1000 km and inclinations of 60°-110°. The current proposal uses a space micro-tug (SMT) for the detailed maneuvering once the debris object is approached close enough. The SMT is connected to the main module using a tether system. Furthermore, the SMT will contain a docking system for acquiring a connection to the debris object (rocket body). A schematic sketch of the design is shown in figure 33.
Thus, the full system will consist of the following:
• Separated part of orbital stage: This element consists of the orbital stage‘s booster equipped with the autonomous de-orbiting system, together with subsystems providing navigation, control and telemetry means.
• Space Micro Tug: The SMT will include a Cartesian propulsion system, providing accurate guidance impulses in all degrees of freedom.
• Tether system: The tether system will connect the separated part to the SMT (and thus the debris object). Of course, the tether will have a certain maximum length, which specifies how close the debris object should be approached before the SMT can be used.
• Docking system: This system will enable the coupling between the maneuverable SMT and the debris rocket body which will be removed from its orbit. It has already been reasoned that the target rocket body nozzle would form a very good candidate for docking, but this theory has not been worked out yet.
• Control, avionics and navigation system: These systems will have to enable:
• Gasification system: The phase-1 gasification-propulsive system will be adapted in following ways:
• Multiple activation heat carrier supply system (up to 4 times)
• Maintenance and monitoring of amount and ration of propellant components
• Improvement of efficiency of gasification system and specific impulse of gas rocket engine (GRE) (fuel additives and ultra sound treatment of supplied heat carrier).
The approach of the debris object will be performed in two stages, the far-guidance phase and the close-guidance phase. During the far-guidance phase, the rocket body‘s ADS is used for bringing the vehicle in the rough vicinity of the object (R/B) which is to be removed. The close-guidance
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phase is initiated thereafter and during it, the SMT will maneuver towards the target object. When the approach is complete, the docking system shall perform the coupling actions between both units. Finally, the whole system is de-orbited and transported to the atmosphere. This final maneuver is again done by the ADS. A functional flow of the full disposal process, together with the subsystems involved, is shown in Figure 34.
Orbital transfer
Using the USSTRATCOM TLE database, containing the Keplerian elements of all detectable debris objects, a list was made of all candidate debris objects relevant for disposal. Filtering was based on object size and type as well as the most populated orbital regions. The target objects amount to 687, 402 of which are intact satellites, the rest being rocket bodies. When the rocket bodies are counted per type, the distribution shown in Figure 35 can be deduced. Notice that the bulk of the candidate rocket bodies are the Omsk-made Kosmos-3M upper stages.
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■ Kosmos
■ Soyuz
■ Tsyklon-3
■ Zenit
■ Dnepr
■ THOR Burner 2A
■ Scout
■ Other
Figure 21: Types of rocket bodies among candidate disposal objects in critical region.
Based on the launch scenario of the past 5 years, it was studied how many AV would be necessary to remove five objects per year. The transfer would bring the upper stage of the launcher from its initial orbit (the orbit in which the primary payload was deployed) to that of a candidate debris object. Also for each debris object, it was studied how much AV is required to transfer it to a disposal orbit in which it is ensured to re-enter in the atmosphere within 25 years. The results are depicted in Table 5. It can be seen that the maximum AVRV (rendezvous) amounts to ~270 m/s, whereas the AVD (disposal) does not exceed ~127 m/s. Both values are kept separate, because the
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mass of the rendezvous module is less than the mass of the full disposal ensemble. This alters the propellant needs considerably for both cases. Ideally, a factor should be introduced which defines how much more severe an increase in AVD weighs in comparison with an increase of AVRV. This relation depends on the mass of the collector and that of the debris object itself. The sixth column shows the time between both transfer burns. This value can be used in the cyclograms of the engine design.
Top 25 minimum energy transfers from simulation considering 2006-2010 launch cycle and a two-
burn transfer
Table 9
Collector Target Object Launch AVlflV' [m/s] AV2r\ [nVs] Ati2 [s] &VD [m/s]
SL-8 SL-8 11-9-2007 23.3 5.5 3135 120.9
SL-8 SL-8 21-7-2009 30.4 29.2 3130 121.3
CZ-2C SAF1R 2 11 -4-2007 56.7 4.5 3035 84.7
SL-8 COSMOS 726 27-4-2010 60.5 15.6 3145 120.3
CZ-2C ADEOS 2 12-11-2009 75.5 30.9 3000 82.9
PSLV FEDSAT 10-1-2007 53.2 73.3 2965 79.9
CZA NOAA 13 19-9-2007 76.9 75.2 3010 93.4
CZ-4B PSLV 25-10-2008 54.1 133.0 2940 81.8
H-2A OPS 8579 (DMSP 5B F5) 24-1-2006 101.8 92.0 2965 69.2
PEGASUS OPS 6073 (DMSP 4A F2) 254-2007 104.3 101.3 2970 73.5
CZ-4C AAU CUBESAT 9-8-2010 66.4 141.1 2910 85.7
SL-19 NOAA 9 28-7-2006 130.1 79.3 2970 89.4
PSLV ROCSAT 2 12-7-2010 95.9 121.3| 2990 104.7
SL-24 SL-3 17-4-2007 86.2 131.6 3000 99.6
SL-24 IRAS 29-7-2009 148.9 71.8 3020 101.4
SL-8 OPS 1127 (DMSP 4B F3) 19-12-2006 118.5 107.2 2930 72.9
CZ-4C WEOS 11-11-2007 118.8 107.5 2930 79.3
SL-14 INTERCOSMOS 22 30-1-2009 99.4 135.5 2950 77.9
SL-24 COSMOS 1736 28-6-2007 119.4 124.8 3000 112.6
CZ4 OPS 3367 A 23-10-2006 101.7 147.3 2950 78.5
SL-16 GGSE4 29-6-2007 76.1 173.7 3075 107.2
CZ-2D OSCAR 8 15-6-2010 106.2 146.9 2985 105.6
CZ-4C SPOT 5 15-12-2009 95.4 159.9 3165 88.1
CZ-2C THOR BURNER 2 22-4-2009 105.1 159.3 2915 72.1
SL-19 COSMOS 1531 8-9-2010 188 81.0 3245 126.4
Debris rendezvous
Based on an analysis of previous rendezvous missions (HTV, ATV, Soyuz, XSS-11 & Dart), a preliminary study was done concerning the potential rendezvous sequence and technology.
In terms of sensors, it could be concluded that a scanning laser system will be very useful for the detection, tracking and inspection of the target debris object. For absolute navigation in the LEO region, GPS or GLONASS can be used, together with tracking from ground stations. The use of relative GPS (RGPS) is not possible, as the target object should be assumed fully passive.
The trajectory will consist of roughly the same phases as a usual rendezvous mission, except that instead of starting with a launch directly into the orbital plane, a transfer is performed from the initial orbital plane (that in which the upper stage‘s payload was injected) to that of the debris object. It may be desirable to perform a certain amount of the phasing sequence before doing the transfer. A second difference with the traditional rendezvous mission is that there is no departure phase. Once the SMT completes the docking phase, both bodies will stay connected until their burn-up in
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the atmosphere. Note that the system then constitutes of two main bodies which are connected with a tether, making the de-orbit maneuver rather complex. A separate study is required for this. Before starting the close approach, most probably an inspection procedure should be carried out to analyze the exact condition and rotation of the target body. In that way the best capture point can be chosen. Depending on the target this could be e.g. a nozzle, a protruding feature suitable for grappling or a surface suitable for a harpoon system to capture. This depends on the used docking system. Finally, the trajectory will be largely influenced by the division of tasks over the ADS and the SMT. As for now, orbital accuracy requirements and the necessity for relative navigation during the SMT‘s functioning, make it most probable that the task division is as depicted in Figure 36.
Figure 22: Rendezvous phases and task division over ADS and SMT
Debris docking
A docking system will be transported onboard the SMT towards the debris object. Its design will be mainly determined by the characteristics of the target debris object. The target‘s dimensions, mass, shape, attitude and angular motion all play a role in its design. Several docking methods have been previously tested in space missions.
As described in previous sections, there is one orbital region (800-1000 km) requiring a removal program the most stringently. Many debris objects which are primary candidates for removal are of the same or very similar designs. E.g. the many Kosmos-3M rocket bodies are one of the main contributors of the hazardous environment. These two aspects lead to the fact that rendezvous and docking systems can be designed in a very general manner. Several concepts are being studied:
• Robotic arm
• Docking through propulsive nozzle
• Harpoon
• Net
Propulsive subsystem
As seen in the previous chapters, the system composed by the orbital stage of the rocket and by the SMT should be able to perform a transfer orbit to reach the designed space debris and, after the part of rendezvous and docking, it should be able to de-orbit itself to burn in the atmosphere.
To accomplish this mission, the orbital stage should carry an Autonomous De-Orbiting System (ADS) that would make the orbital stage, after the deploying of the payload in the right orbit,
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de facto, a new satellite and not a debris as usually the orbital stage of rockets are after the dismissal.
The choice of the proper ADS isn‘t easy and mostly can depend on the rocket used, but it‘s possible to make some general assumptions to define common criteria.
To do this, first of all, it has to be determined to type of the mission. The ADS system could perform the transfer orbit till the space debris, but also a simple de-orbiting of the 2nd stage without waiting the natural lower of the orbit perigee due to drag or other factors and so fasten the re-entry in the atmosphere.
In addition, the criteria chosen to define the suitable engine should be applied to all the engines to have a common way to discriminate among them. So it were chosen mass, length of the engine, thrust, power input and TRL as criteria for the first creaming off. Of course other parameters, like volumes and propellants could enter in the criteria, but they aren‘t the main important at this level of study, even if they will be discussed further in this document.
Choice of the rocket
The ADOS system concept should be able to work with many different kind of rockets. The feasibility of the project was studied initially considering a COSMOS 3M 2nd stage as example, but due the time of retirement of this rocket, also the RUS-M was considered. As previously said, the
considerations done in this work, can be applied, with some modifications, to all kind of rockets so, going on this chapter, it will not be considered a definitive launcher, but it will be taken only an indicative mass.
Criteria of choice
Mass, length of engine, thrust, power input and TRL, as said, are the main criteria chosen for the definition of the suitable engine.
Mass is obviously chosen for the implication in the lift-off phase and it should be as less as possible to not interfere with the mass of the rocket payload. In fact it must be remembered that the rockets should be able to carry in orbit its payload and the space debris mitigation mission should be considered only a parallel mission. The mass criteria are without doubt the most important because its connection with the mission cost.
Length is chosen because the engines should be placed in the space between the rocket and the fairing and even if it‘s possible to manage the volume in different shape (ex. Tank), combustion chamber and nozzle cannot assume fancy shape and the length becomes important parameters in the analysis. As for mass, also the length of the engine can influence the fairing and one of the designers main goal it would be to fit the engine without losing precious space for the payload.
The thrust is a value needed for the calculating the time of transfer and de-orbiting and its importance is valuable in determining the quality of the removal mission. High Thrust can permit faster de-orbit.
The power input of the engine can give important information about the effectiveness. In fact, it has to be considered that the orbital stage can use only the unspent power and the engines should work only with that power. Too much power used for an engine can cause the need of carrying more batteries and so using carrying more mass, connecting again at the first criteria.
Technology Readiness Level (TRL) gives an idea about the time need to develop a technology for the space application. Even if the project is very ambitious, at the same time it should be co m-pleted in few years to improve the space environment significantly. A correct balance between new technologies and reliability is the right approach to have a successful mission in short time.
Preliminary evaluation
The preliminary considerations have brought to not consider in the evaluation the monopropellant engines for the lack of thrust necessary to perform the orbital transfer with masses on the order of at least more than 1 tons. Electrical engines weren‘t taken in consideration because the
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concept of the mission doesn‘t permit the use of the power needed by an electrical engine. So, the analysis was focalized only on chemicals engines: liquid bi-propellant engine, hybrid engine, solid rocket engine and gas rocket engine.
Here it is a table about typical mono-propellant engine. As said previously, the thrust is too low in fact mono-propellant engines are used to attitude control no to transfer orbit.
Examples of Liquid Mono-propellant engines and characteristics
Table 10
Name MR-1G7 MR-15 OMV DOK-5G MR-1G4
Company Aerojet Northrop Grumman Isayev Aerojet
Mass 0.89 1.13 1.10 1.86
Engine
[Kg]
Fuel/Oxy Hydrazine Hydrazine Hydrazine Hydrazine
Restart 7005 pulses 100,000 cycles 40000 1742 pulses
Thrust [N] 257 89 50 441
Isp [s] 236 225 229 239
Tb [s] 1200 2000 1500
Spacecraft and upper stage attitude control and AV corrections
Attitude control and velocity corrections: Voyager, Magellan, DMSP, Tiros N, Landsat.
To be more specific, a mono-propellant engine is not an acceptable hypothesis for the ADS because of the poor thrust. In fact we have to consider that the separating part has a lot useless mass (2nd stage structure and main engine, unspent propellant in the tank...) and according to the mission phases, the separating part has to connect also with the space debris. Even if a Mono-propellant engine is capable of multi-ignition, the specific impulse and the thrust aren‘t enough to perform the maneuver requested in relatively short time. It‘s not possible also to consider using only a small satellite with obviously a mass less than the 2nd stage mass of a rocket because, in the end, for the mission‘s purpose even this satellite would see its final mass increasing due to the grabbing of a space debris.
Going on with the preliminary considerations, it was made a research of existing and suitable engines (Liquid and Hybrid) to have a comparison with real engines. In the table below it‘s possible to observe the list of possible chemical engine for the mission‘s purpose used in space.
Examples of Liquid Bi-propellant engines and characteristics Table 11
Name R-42B TR-3GS TR-312- 1GGMN 11D458
Company Aerojet Northrop Grumman Northrop Grumman Isayev
Mass Engine [Kg] 4.53 4.76 6.03
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Fuel/Oxy N2O4/MMH N2O4/N2H4 N2O4/MMH N2O4/UDMH
Restart 134
Thrust[N] 890 454 502 392
Isp [s] 303 322 325 275
Tb [s] 3940 3000 3000 3200
Dual Mode Dual Mode Small attitude control thruster in Briz upper stage
The specific impulse and the thrust are evidently higher. These engines represent a criterion of evaluation for further calculations.
Figure 23: TR-312-100YN. Source: [31]
Figure 24: T R-308. Source: [32]
In the table below two hybrid engines are shown, one for small satellite and the other one for upper stages, but never developed completely.
The hybrid hypothesis it‘s not supported, until now, by actual engine capable of what it is needed in the mission. The analysis will take place in any case also for the hybrid engine, to under-
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stand if it‘s convenient a future work on it but there are no means of comparison but obviously the TRL will not be high.
Examples of hybrid engines and characteristics
Table 12
Name VFP Hybrid rocket MTV Motor
Company Surrey SpaceDev
Mass Engine [Kg] 7.94 or more ...
Fuel/Oxy HTPB+20% alum/H2O2 Nano- HTPB/N20
Restart
Thrust [N] 189
Isp [s] 314
Tb [s]
Upper stages. Small hybrid rocket motor designed for use in the Maneuver and Transfer Vehicle, an upper stage orbital transfer motor.
Even if the solid propulsion too will be analyzed in this study, it‘s not a suitable option for the transfer orbit as thought until now, because of it needs multiple ignition and a solid rocket engine can provide only one ignition.
Analysis of engines for De-orbiting
First of all, it will be shown the result for the de-orbiting only of the 2nd stage, without any space debris attached. The versatility of the project consists in its capability to adapt, not only at different rockets, but also to different purpose. In fact, it‘s possible to imagine a case in which, due to the payload mass request, it couldn‘t be possible to add additional mass for the propellant needed for the transfer orbit and only the de-orbiting phase of the 2nd stage could be successfully accomplished.
The engines‘ parameters were calculated considering a the AV for the de-orbiting of 200 m/s and the mass of the 2nd stage of 6000 kg. These are average value, needed to prove the feasibility. Further study can define them deeply. The configuration of the ADOS engine is made by 4 engines
and 4 nozzles to control the attitude without gimbals like in the image below. The analyses are carried out considering the criteria for every type of engines.
Solid Propulsion
A solid propulsion engine cannot be used for the orbit transfer, because of the impossibility of several ignitions but it can work for a de-orbiting mission as stated by CNES studies showed in this study. Low TRL, high specific impulse and reliability are the main factors of this choice, on the other hand, due to the single ignition, the engine should be aligned in the proper direction to perform the de-orbiting.
The average value of specific impulse for a solid engine is 290 s and the fuel chosen is a standard mixture of 21% Al, 57% AP, 12% HMX and 10% HTPB.
With a 4 nozzles configuration the total thrust F is 2308N, the mass for every engine is 323kg giving a total mass for the system of about 1292 kg. The length of the one engine is 0.4487m, considering both combustion chamber and nozzle. The energy input for this kind of engine is 1641.4kW for each engine [35][36] [37].
Class Star 17: 100 kg, L 1000, 4 400 Class 3tar 27B; 450 k9, L 1300, 4 700
Figure 25: Examples of suitable solid engine
Liquid Propulsion
A liquid propulsion engine can be analyzed in two ways: using an autonomous tank system or receiving fuel and oxidizer from the main tanks of the rocket. The second option presents some technical issue because it should require some device to lead fuel and oxidizer to the 4 combustion chambers of the de-orbiting system so it‘s not so immediate.
Thrust = F = m Ve + (pe-p0) Ae
Figure 26: Liquid engine scheme. Source: [33]
For the autonomous tanks design, the total thrust with an average specific impulse of 330 s is 8829 N considering as propellant a mixture of RP-1 and LOX. The propellant mass in tank it will be at least 430 kg, while the total engine‘s mass is 800kg, including 4 combustion chambers and nozzles. The length of a single engine is 0.4852 m and the energy input for this kind of engine is 2339.6 kW for each engine.
Using the rocket feed system, fuel and oxidizer are provided by the main tanks. So the mass will be only the sum of combustion chambers‘ and nozzles‘ masses and it will be about 8 kg. The
length of the system is about 1.1 m giving a thrust of 4709 N. The energy input for this kind of engine is 1524kW for each engine [35][36] [37].
Hybrid Propulsion
The hybrid propulsion combines both solid and liquid solution.
Injector
Figure 27: Hybrid engine scheme
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The mass of the system is composed by the combustion chambers where the fuel is in solid state and by the tank of oxidizer for a total of 1047kg. The propellant fraction of the system is equal to 462 kg. For a specific impulse of 280 s the total thrust given by the engines is 2545 N and the length of the engine is 0.524m. The energy input for this kind of engine is 4859kW for each engine
[35][36] [37].
Gas Rocket Engine
The gas rocket engine uses the gas generated from the gasification of both fuel and oxidizer in the tank of the rocket to generate del AV needed for the propulsion [38]. The gasification process will be explained further in this work. The mass of the system will be obviously only the mass of the 4 combustion chambers and nozzles equal to 9.3 kg. The length of the system is 1.3m while the thrust generated is 8010 N. The energy input for this kind of engine is 7364 kW for each engine
Figure 28: Gas rocket Engine scheme. Source: [34]
Confrontation (De-orbiting)
Mass (Autonomous Propellant)
1200
1000
800
600
400
200
Engine Liquid Bi-Propepellant Engine (autonomus tank)
205
Mass (Rocket Propellant)
M
0
Solid Hybrid Engine
9,5
8.5 8
7.5
7
Liquid Bi-Propellant Engine (shared tank)
Gas Rocket Engine
Figure 29: Mass confrontation
1,4
1,2
m
C
V
-J
1
0,8
0,6
0,4
0,2
0
i
ii
Lenght
Solid Engine Liquid Bi-
Propepellant Engine (autonomus tank)
Hybrid Engine Liquid BiPropellant Engine (shared tank)
Gas Rocket Engine
I
■M
3
a
j!
0
is
o
&
8000
7000
6000
5000
4000
Thrust [N]
Solid
3000
2000
1000
Power Input
206
10000
9000
8000
7000
6000
5000
4000
3000
2000
1000
Engine Liquid Bi-
Propepellant Engine (autonomus tank)
Thrust
Hybrid Engine Liquid BiPropellant Engine (shared tank)
Gas Rocket Engine
Figure 30: Length, Power and Thrust confrontation
The mass criteria cannot be, all alone, the unique criteria of choice, because the mass of the engine is significantly reduced using the native rocket feed system with the unspent propellant. Considering the length of the system (combustion chamber and nozzle) the tank autonomous systems can be considered the better choice. The power input is huge in the case of the Gas Rocket Engine but also provide a thrust much more elevated than the other system comparable only with the Liquid Bi-Propulsion Engine.
Finally, in the analysis it should be considered also the TRL. The hybrid engine, as previously stated, has an high TRL, not supported by evidence in result that can justify further study. The limit of the solid engine has been already discussed; even if it‘s maybe the cheapest solution on the market, it requires much more accuracy in the direction of thrust because the single ignition doesn‘t allow errors. The gas rocket engine has still an high TRL but it‘s promising because of the application on almost every kind of rockets. The liquid bi-propellant engine with the feed rocket system has as problem, the difficulty of retrieving the fuel and the oxidizer from the tanks. Membranes or an inverse pressurization system could be used to accomplish the feeding of the engines.
The liquid bi-propellant engine with autonomous tanks and the gas rocket engine are the best solution according the criteria. The choice depends on the time needed to develop the engines and make them fully functional and reliable.
Analysis of Engines for Transfer Orbit and De-orbiting
The engines‘ parameters were calculated considering a the total AV for the mission of 400 m/s, 270 m/s for the transfer orbit and 130 m/s for the final de-orbiting the mass of the 2nd stage is still 6000 kg but it has to be added to the mass of space debris of 1400 kg (average dry mass of Cosmos 3M).
Engines' comparison
Table 13
LE - Autonomous tank LE - Shared tank Hybrid Engine G.R.E.
Mass [kg] 1305.2 51.82 48634 145
Length [m] 0.9235 1.8732 1.0109 4.1271
Power Input** [kW] 3239.7 128580 48503 128580
Thrust [N] 221840 39718 21778 39719
(*) Value for each engine
For the combined option of transfer orbit and de-orbiting obviously the AV is higher and so the criteria chosen increase. Another choice to decrease the total AV could be use and Electrodynamic Tether (EDT) to save propellant and mass. Using and EDT the dimension and the mass of the engines chosen for the 2nd stage could be almost the same of the de-orbiting case.
Electrodynamic Tether
The Electrodynamics tether (EDT) could be used to not consider the AV for de-orbiting, saving a considerable amount of propellant and so mass and volume from the separating part.
The University of Madrid studied the problem and trough numerical simulation jumped to the conclusion that a 20 km EDT can absolve to the job of de-orbiting an orbital stage in 25 years, as stated in the international agreements and laws [9]. An electrodynamics tether, as explained previously, is attached to an object, the tether being oriented at an angle to the local vertical between the object and a planet with a magnetic field. When the tether cuts the planet's magnetic field, it generates a current, and thereby converts some of the orbiting body's kinetic energy to electrical energy. As a result of this process, an electrodynamics force acts on the tether and attached object, slowing their orbital motion. The tether's far end can be left bare, making electrical contact with the ionosphere. Functionally, electrons flow from the space plasma into the conductive tether, are passed through a resistive load in a control unit and are emitted into the space plasma by an electron emitter as free electrons [16].
The EDT could be used in the mission provided always that the power consumption wouldn‘t be too high because the power is given by the unspent resources in the orbital stage‘s battery.
Gasification of unspent liquid propellant
In order to reduce the mass for a de-orbiting mission, a method to use the unspent liquid propellant of the rocket stages is studied.
Since at the end of the rocket‘s mission remains in the tanks about the 3% of the beginning propellant mass, we want to exploit completely or at least the more as possible this remaining resource to accomplish maneuvers. To obtain a sufficient thrust that forced this stage in a faster reentry, we want to use the evaporation of this residual propellant.
Gasification process could be used for different rockets, different stages of rockets and different maneuvers such as:
• Orbital maneuver
• De-orbiting
• Safe reentry
Orbital maneuver for example to reach debris and grab them (Figure 1: AVi and AV2) , deorbiting maneuver to force the reentry of debris (Figure 45: AV3) and a safe reentry maneuver needed for the rocket‘s first stage to avoid cities (Figure 46: AV2).
AV.
Figure 31: Orbital maneuver and de-orbiting
Figure 32: Safe re entry
In any case, gasification is made by a system which introduces in tanks warm gases, with temperature in a range of 800K - 1400K, that heat the gases and liquids present in the tanks and by heat exchange from gas and liquid we obtain the evaporation of residual propellants. A pressure relief valve is installed in the tank to let the gas go out when the pressure reach the maximum level.
Important parameters
Mass of the warm gases, time of gasification, and power input and are the main important parameters for this process.
Mass of the warm gases should be as less as possible to not interfere with the mass of the rocket payload but enough to complete the gasification process.
Time of gasification is really important for the safe reentry phase because the 1st stage need between 400 and 600 seconds to reach the ground. So the gasification must be fast enough to allow a sufficient thrust for cities avoidance. For the other phases (orbital maneuver and de-orbiting) this parameter is not so important but it‘s important to keep the gasified propellant at a high temperature.
The power input for the gasification process (as for the engine) can give important information about the effectiveness. In fact, it has to be considered that the orbital stage can use only the unspent power and the gasification process is only one part of the mission.
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Method
The mathematical model used to study this process is based on this two equations [26][27]:
- the first one to study the temperature of the gas inside the tank
ИиИи
ЖЗе
^ — 000 + ^00001+ ^0000 2 00 0000
- and the second is for studying the liquid propellant temperature:
= 0
0000 1
s jnsi
dis
andTOe gasified propellant ~ann rneneat now due mThe*wssfrow~g©mg“out
Because we are in space, in case of orbital maneuver and de-orbiting, or we are free falling, in
case of reentry of the rocket‘s first stage, the liquid propellants will be spread all over the tank as in
Ошибка! Источник ссылки не найден.
Figure 33: Liquid propellant in the tank
The thermal study was done for these three different cases:
Figure 34: Study cases
Mass [kg]
Results
The following graphs show the results of the gasification of propellant RP-1 using as worm gas a mixture of Oxygen and RP-1 gasified:
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Propellant mass in tank with Tjn=1400[K]
Time [s]
Figure 35: Trend of the propellant mass
Temperature [K]
Time [s]
Figure 36: Trend of propellant temperature 211
Figure 37: Trend of tank temperature
Figure 38: Heat needed for gasification 212
Future studies
A future study for this process will be the effects in temperature and velocity of gasification using a permeable membrane which limits the liquid propellant in a known area of the tank.
Figure 39: Tank with permeable membrane
Economical potential
With today's annual launch rates of 60 to 70 and with future break-ups continuing to occur at mean historic rates of four to five per year, the number of objects in space will steadily increase.
As a consequence of the rising object count, the probability for catastrophic collisions will also grow in a progressive manner (doubling the number of objects will increase the collision risk approximately four-fold). Nowadays, insurance is playing big role in the space projects, because it secures the financial stability of the companies and it is considered to be the second main cost in the projects. It is important to notice, that if the number of collision risk will increase, the cost of insurances will increase too affecting the possibilities of cheaper launches that are the key of a space environment easily open to everyone. In addition with the increasing of the problem, it could be the possibility of an —Eco-tax|| to pay for the cleaning of orbits.
The cleaning of the LEO must be seen as a future investment. Without an appropriate clean-
ing, as said, the number of collision will dramatically increase, many expensive satellites will be lost without any refund by insurance companies and in the end this will affect the cost of every service based on satellite technology: GPS, weather forecast, telecommunication etc. The big earning in the investment will be to not lose more money in the future.
Private investors interested in more reliable future, insurance companies interested in keeping the risk of collision low, worldwide Eco-tax could be the possible routes of financing.
But such project, using the orbital stage of a launcher after the deploying of the payload, permits to save the amount of money due to the launch like it would be piggyback mission. This particular aspect cannot be ignored, because until now, the projects proposed to the international attention required a specific launch and this lead to large amount of costs only to buy the service of a launcher. Using the orbital stage, already sent in orbit for its commercial or scientific purpose can decrease the cost of the space debris removal mission.
In the table 12 it‘s possible to see an overview of the main different concepts with also the cost of the mission, where indicated. The OmSTU project, after the costs of research and development, due to the —piggyback! solution hasn‘t the fixed voice of the launch‘s expense as the others projects have.
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Overview of the main projects and costs
Table 14:
OmSTU CNES RKK Energia Jaxa
General design philosophy of disposal method Reusibility of disposal system Possibility of installation on ex- Based on the gasification of residual propellants in the separated stage's tanks Autonomous propulsion system based on , solid propellants. Autonomous propulsion system based on nuclear propellants. Autonomous tem based on liq-
No Yes Yes Yes
isting separating parts, without revising the rocket body Number of re- No No No No
moved bodies one single system run Critical technology in the design of disposal method. Initially only one object. At a later stage, collection and disposal of additional space debris object Up to 15 Undefined Up to 1
Gasification of propellant residues. Study of process and improvement methods. Multiple autonomous docking functionality Handling of noncooperative debris. Development of a nuclear rocket engine Multiple autonomous docking functionality Handling of noncooperative debris. Multiple autonomous docking functionality Handling of noncooperative debris.
Estimated development costs, including demonstration experiment (million $) 10 380 Undefined Undefined
Projected development period 2-3 years 3 years Undefined Testing microsatellite in 2016
before testing
Sustainability
It has been estimated that as many as 100 satellites have broken up while in orbit, sometimes due to explosions of propulsion systems, and at other times due to impact with other space debris. The result is a an estimated 40,000 to 80,000 pieces of debris in orbit around the Earth. The main
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cause of on-orbit explosions is related to residual fuel that remains in tanks or fuel lines once a rocket stage or satellite is discarded in Earth orbit. Over time, the harsh space environment can deteriorate the mechanical integrity of external and internal parts, leading to leaks and/or mixing of fuel components, which could trigger self-ignition. The resulting explosion can destroy the source object and spread its mass across numerous fragments with a wide spectrum of masses and imparted velocities. The aim of this project is perfect to solve this problem because first of all we will remove debris avoiding the possibility of crashes and second with the gasification of unspent liquid propellant we can use a source of energy already available without waste of propellants avoiding possible explosions. In addition at all these active removal process, the gasification process uses clean gases to warm up the liquid propellant and gasify it, for example O2 and H2O2 avoiding pollutant products that can be dangerous during a re-entry in the atmosphere.
Figure 40: Satellite explosion due to the propulsive system.
International Scope
Space Debris, an international issue
Based on the projects and studies presented at recent seminars hosted by NASA (Virginia, 2009), ISTC (2010, Moscow), SWF (Beijing, 2010) and CNES (Paris, 2010), an analysis could be performed of the various existing developments in the fields of large debris disposal. This can lead to a possible integration plan for the development of a joint project involving various partners. The principle of the aggregation of various existing national developments into an overall international joint project for the development of an active debris removal program is based on two =pillars‘. Firstly, there is the fact that debris mitigation is an issue affecting space agencies and space related companies worldwide. The assurance of a stable near-Earth space environment is vital for future space projects as a whole and therefore should be pursued internationally. Secondly, the vast majority of feasible debris disposal methods are characterized by a principle of modularity, in which various reasonably independent aspects and tools (i.e. debris detection, launch, approach, docking and de-orbiting) together form the functionality of the system as a whole. The combination of these two factors makes the project very attractive for a joint project involving various partners worldwide,
each using its fields of specialization to contribute to a successful design. Separate organizations should assess the project management, the composition of the groups of experts and the collaboration with organizations involved in rocket design and space activities, including the UN, IADC and national space agencies.
International partners and proposed task division
As discussed previously, the design of a debris removal system forms a good candidate for a joint project between different partners. A study was performed about which academic or industrial
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institutes could be assigned the design of different subsystems. The fact that these entities have previously developed very similar products as required for this debris removal mission, can act very advantageous in terms of money, time and the seek for specialists. A rough overview of the design process and respective stakeholders is depicted in Figure 58. Various contacts have already been established with some of the depicted entities and potential partners. The mission concept has been established, together with the target requirements. The detailed design of the subsystems will be the next step in the process and will require different institutes to share work and insights.
Phase 1: Concept and Requirements Establishment
Debris Removal Concept OmSTU
Phase 2: Subsystems Detailed Design
ADS
Docking
System
OmSTU Sapienza / DLR
Space Micro Tug
CNES / JAXA
Tether System
TU Madrid / Tethers Unlimited
Subsystem production and integration
Phase 4: Testing
Full System Flight Test
Phase 3: Manufacturing and Assembly
Industrial partner(s)
Samara Space Center
Figure 41: Schematic overview of design process. Estimated time: 3-5 years
Conclusions
In conclusion this project has the possibility to change the future not only of our space environment, cleaning the LEO orbit from space debris, reducing the risk of new debris and using clean propellant, but also of the space business, reducing the cost of insurances and keeping the possibilities of launches for satellites in LEO orbit, as explained in chapters 6 and 7. The results achieved show that it is possible to bring to Earth the space debris using different grabbing and propulsion systems and so reduce their number and the risk of collisions or explosions. Because of different parts of the project have already an advanced level of theoretical development, it is possible to assume a mission of active debris removal in the near future so as to follow the trend expected from the NASA‘s LEGEND model, it will be necessary. The active space debris removal could be done for every orbit (LEO or GEO) but due to the limited availability of liquid propellant for the gasification process, it is possible to use this technique only for LEO orbits, which are the most critical, because they require a lower Av for the orbital maneuver and de-orbiting. Because of the possibility of using different kind of rockets (Rus-M, Kosmos-3M, Soyuz, Ariane, Delta, Atlas...) the ratio of space debris removed could be more than five per year (the minimum request by NASA‘s LEGEND model) and the total number of object in space could be reduced in a short time. Of course the design should be adapted to every launcher, but it‘s a hypothetically practicable solution. But to permit the developing of the project, it is necessary a join collaboration among universities and industries. We truly hope that this project could become the —ignition! to attract new academic and industrial partners to design a fully working mission before the 2020.
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