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Spacecraft Contamination Causes

Spacecraft contamination is typically caused by a combination of materials outgassing, thruster operations, venting of gases and liquids, and the generation and release of particulates.

Outgassing from materials is a product of the release of volatile components contained in the materials with vacuum exposure. Nonmetallic materials are typically the most significant sources of outgassing contamination on spacecraft. For example, organic silicone-based materials such as room-temperature vulcanizing silicones outgas a variety of substances – including silicones compounds – that can produce significant contamination impacts on sensitive surfaces.

Thruster operations induce both contamination and erosion on exposed surfaces. Chemical thrusters typically produce a two-phase plume. While the majority of the effluent is in the gas phase, combusted and partially combusted by-products are also present in the liquid phase. Contamination (and erosion) from the liquid phase affects surfaces operating at any temperature. Contamination from the gas phase is limited to surfaces that are cold enough to condense the gaseous species.

Particulate releases also produce contamination. Particulate matter originates from ground contamination (introduced during assembly and storage), from degradation of coatings and materials, and with venting operations.

Contamination Hazard Effects

Contamination can produce hazard effects on spacecraft as contaminant deposits can degrade the performance of spacecraft systems, restrict the spacecraft operational envelope and extravehicular activity operations, and damage mechanisms.

Contamination Hazard Controls

Contamination hazards can be controlled during the design phase of a spacecraft. Analysis tools are used to model contamination and support trade studies. The results of these studies are used during materials selection; design of venting paths; and design, placement, and orientation of vacuum vent and thrusters.

International Space Station Examples

During a typical year of International Space Station operations, induced environments are produced with thruster operations (i.e., station re-boost and attitude control as well as proximity operations of the Space Shuttle Orbiter and Russian, European, and Japanese visiting vehicles) and with the operation of vacuum vents, which vent a variety of gases and, in some cases, liquids to space.

Unique challenges were encountered during the design, assembly, and operation of the International Space Station. Mitigation strategies were developed to protect the station from induced environment effects, and to develop risk mitigation and safing constraints to protect the vehicle and its science utilization capabilities. These mitigation strategies include protecting the crew from exposure to thruster-induced contamination during extravehicular activities as well as protection of vehicle systems from thruster plume effects and operations, vacuum venting of liquids, and optical property degradation. Examples are provided in the following sections that illustrate both the issues and risk mitigation strategies and the safing constraints that were developed to resolve these issues. These strategies are applicable to the development of future long-duration space systems, not only during the design phase but also during assembly and operation of such systems.

References

1. Larin M, Lumpkin F, Stuart P. Modeling unburned propellant droplet distribution and velocities in plumes of small bipropellant thrusters. AIAA Paper 2001–2816. The 39th AIAA Aerospace Sciences Meeting and Exhibit Reno, NV: American Institute of Aeronautics and Astronautics; 2001.

2. Pankop C, Alred J, Boeder P. Mitigation of thruster plume erosion of International Space Station solar array coatings. Journal of Spacecraft and Rockets. 2006;43(3):545–550.

3. Schmidl WD, et al. Mitigating potential water dump particle impact damage to the International Space Station. Journal of Spacecraft and Rockets. 2006;43(3):551–556.

4. Schmidl WD, et al. N-nitrosodimethylamine release from fuel oxidizer reaction product contaminated extravehicular activity suits. Journal of Spacecraft and Rockets. 2006;43(3):557–564.

5. Soares CE, Mikatarian RR. Thruster plume induced contamination measurements from the PIC and SPIFEX flight experiments. International Symposium on Optical Science and Technology. Vol. 4774-20 Bellingham, WA: Society of Photo-Optical Instrumentation Engineers (International Society for Optical Engineering); 2002.

6. Soares C, Mikatarian R, Barsamian H. International Space Station bipropellant plume contamination model AIAA Paper 2002-3016 The 40th AIAA Aerospace Sciences Meeting and Exhibit. Portsmouth, VA: American Institute of Aeronautics and Astronautics; 2002.

Principal Causes of Post-Mission Fragmentations

The analyses of nearly 150 accidental on-orbit explosions have identified the three primary sources for self-induced satellite breakups: propellant systems, pressurant systems, and electrical systems. This is not surprising, since these systems can contain substantial levels of stored energy, even after the vehicle has been decommissioned.

Mitigation Measure Guidelines

In 1995 NASA published the first set of detailed orbital debris mitigation guidelines as NASA Safety Standard 1740.14. One of the principal goals of this standard was to curtail the creation of orbital debris during both the active operations of spacecraft and launch vehicle orbital stages and their post-mission states. To avoid accidental explosions after the completion of mission operations, Guideline 4-2 of the standard stated that “All on-board sources of stored energy will be depleted when they are no longer required for mission operations or post-mission disposal. Depletion will occur as soon as such an operation does not pose an unacceptable risk to the payload.”

The depletion of all stored energy on space vehicles is known as passivation and includes all propellants, pressurants, batteries, momentum wheels, and other devices with stored mechanical or chemical energy. Experience has shown that complete passivation of spacecraft and rocket bodies at end-of-mission prevents future self-induced debris generation events.

To increase the likelihood that such passivation measures are possible at the end of mission, which might be many years after launch, a separate guideline (Guideline 5-2) addressed the vulnerability of critical subsystems to the space environment. Specifically, the guideline stated that “In developing the design of a spacecraft or upper stage, a program should estimate and limit the probability of collisions with small debris of size sufficient to cause loss of control to prevent post-mission disposal.”

Hence, communications and electrical systems must remain capable of performing their passivation tasks. If a disposal maneuver is necessary (see next), then the propulsion system must also be protected.

For the long-term preservation of the near-Earth space environment and its availability to support vital applications, exploration, and science missions, collision risks between resident space objects must be curtailed. It is now generally accepted that portions of the region of low Earth orbit (LEO, i.e., below 2000 km) are already environmentally unstable. In other words, the congestion (or spatial density) of objects in certain altitude regimes will lead to the production of collision-induced debris at a rate faster than the debris can naturally decay to lower altitudes, thus leading to a steady increase in the Earth’s satellite population.

Recognizing this effect, the NASA standard set a limit for the residual orbital lifetime of spacecraft and rocket bodies in LEO to a maximum of 25 years following the end of mission (Guideline 6-1). For a spacecraft, end-of-mission normally is defined as when the payload is turned-off or ceases to operate due to a malfunction. Launch vehicle stages in LEO typically reach end-of-mission on the day of launch. If a spacecraft or rocket body is in a sufficiently low orbit that natural orbital decay will lead to re-entry into the Earth’s atmosphere within 25 years, then no specific additional action beyond passivation is necessary. However, if the vehicle is in a higher altitude orbit, then a maneuver into a lower orbit or some other option must be invoked.

Although the geosynchronous orbital regime (GEO) does not exhibit spatial densities that currently result in significant accidental collision risks, the exceptionally long-lived nature of any debris produced in such a collision has led to a general international consensus that GEO satellites should be maneuvered into higher altitude storage orbits at end of mission (1995 NASA standard Guideline 6-2). Since the 1970s the definition of an appropriate disposal orbit for GEO satellites has evolved, and today a commonly held view is that the orbit should be sufficiently high in order to remain at least 200 km above the nominal GEO altitude of 35,786 km for a period of at least 100 years.

The aforementioned measures to mitigate the creation of new orbital debris have been broadly accepted by several other national space agencies and the international aerospace community as a whole. Japan’s first Space Debris Mitigation Standard (NASDA-STD-18A) in 1996 reflected many of the previously established NASA guidelines. The French space agency CNES followed with a similar set of guidelines (MPM-50-00-12) in 1999. The Russian Federation and the People’s Republic of China also produced orbital debris mitigation guidelines in 2000 and 2005, respectively.

The first multi-national set of orbital debris mitigation guidelines was adopted in 2002 by the 11 members of the Inter-Agency Space Debris Coordination Committee (IADC) which was comprised of national space agencies of 10 countries (China, France, Germany, India, Italy, Japan, Russia, Ukraine, United Kingdom, and the United States), as well as the European Space Agency. The IADC guidelines contained the by then widely–accepted recommendations for post-mission vehicle passivation and disposals in orbits which would result in limited-duration stays in the LEO or GEO regions.

The IADC guidelines were presented to the Scientific and Technical Subcommittee of the United Nations Committee on the Peaceful Uses of Outer Space (COPUOS) in early 2003. After several years of deliberations, the subcommittee produced in 2007 a very similar set of mitigation guidelines, which were endorsed later that year by both the full COPUOS and the UN General Assembly.

Mitigation Measure Practices

Some national standards call for the preparation of a preliminary end-of-mission plan early in a space project’s development. Issues such as passivation and disposal are best addressed no later than the project’s Phase A in order to promote cost-effective compliance with all the necessary requirements. A draft end-of-mission plan should be completed before launch and updated during the mission, particularly if critical subsystems degrade or suffer serious malfunctions.

Passivation of liquid or gaseous propulsion systems is normally accomplished by either depletion burns or venting. For example, U.S. Delta launch vehicles execute depletion burns to expend residual propellants, whereas U.S. Atlas launch vehicles choose to vent residual propellants.

Depletion burns have the potential added advantage of being able to assist in reaching acceptable disposal orbits without additional cost. For example, a Delta 2 second stage executed a depletion burn in 1996 to enable a re-entry from a 900 km circular orbit in only 9 months. In 2006 a Delta 4 second stage performed a controlled re-entry from an altitude of 850 km. Similarly, after a highly successful mission of 14 years NASA’s UARS spacecraft conducted a depletion burn in late 2005, accelerating its fall back to Earth by more than 20 years. In 2010 the commercial Orbview-3 spacecraft executed a controlled re-entry with its residual propellants.

Whether propellants are vented or burned, the amount of propellants remaining on the vehicle should be as close to zero as possible. Satellites with less than 10 kg of propellants have fragmented, leaving hazardous orbital debris in their wake. It is recognized that very small amounts of propellants might unavoidably adhere to tank and propellant line walls after passivation. Such small amounts are unlikely to pose future breakup risks.

As noted above in the two examples of the Proton Briz-M stages which shutdown prematurely after experiencing engine malfunctions, unanticipated residual propellants can pose risks of future breakups. Consequently, orbital stages should be designed to automatically vent propellants in such malfunction scenarios.

The pressure in propellant and pressurant tanks should also be as close to zero as possible. Limits such as 25% of maximum expected operating pressure (MEOP) are undesirable due, in part, to the unknown failure modes of tanks, even those of leak-before-burst designs, decades into the future, e.g., in GEO disposal orbits. Moreover, hypervelocity impact laboratory tests have clearly demonstrated that small particle (natural or man-made) impacts on tanks with as little as 1 atmosphere of pressure (i.e., 14.7 psi) can produce significantly greater damage and resultant numbers of debris than tanks which are in a vacuum state. Hence, it is desirable that all tanks be vented to space, even if depletion burns have been conducted.

Designing spacecraft and launch vehicle stage tanks to be vented to space is not difficult and need not reduce the reliability of the vehicle. In fact, many modern spacecraft have implemented this capability. Concerns about accidental venting can easily be addressed with multiple valves in the final vent line. For spacecraft the software commands needed to activate the valves do not need to be loaded into the spacecraft until shortly before the decommissioning process begins.

With current technology designs, passivating some pressurized components would present challenges and little benefits. For example, many batteries and heat pipes are pressurized with no option for relieving that pressure at end of mission. Heat pipes by nature are very ruggedly built and less susceptible to small particle impacts. To date no satellite breakup appears to have been caused by a heat pipe, and thus NASA exempts this component from passivation measures.

If batteries are removed from charging circuits as recommended (or required), the chances for explosion are reduced. Leaving batteries connected to charging circuits but with an electrical load (e.g., a heater) is not desirable since that load could fail months or years later, possibly allowing the battery to overcharge and over-pressurize. Realigning the solar arrays to point away from the Sun is also not a viable permanent solution since spacecraft are subject to numerous perturbations, including small particle impacts, which can upset the attitude of the vehicle.

As with tank venting designs, the disconnection of batteries from charging circuits can be done in safe and reliable ways. In contrast to installing propellant and pressurant valves in series, electrical relays can be installed in parallel to ensure that a single malfunction does not prematurely disable the charging system. Again, software commands to activate the relays can be loaded only when they are needed. NASA is employing this technique with increasing frequency.

For many spacecraft and rocket bodies completing their missions in orbits above 600 km, meeting the goal of limiting residual orbital lifetime to no more than 25 years in LEO might require an additional action. For spacecraft in orbits up to 800 km, only a modest change in velocity (delta-V) is necessary to lower perigee such that the vehicle will re-enter within the desired time. The exact magnitude of delta-Vs required will be dependent upon when in the solar cycle the end of mission occurs. Moving from a circular orbit to an elliptical one is always more efficient from an energy standpoint than maneuvering into a lower circular orbit with the same residual orbital lifetime.

Higher altitude spacecraft would necessarily require greater propellant expenditures. One solution might be the use of low thrust, higher efficiency engines, such as ion engines. On the other hand, such operations would require maneuvers over a much long timeframe, perhaps months or even longer, raising additional concerns about spacecraft reliability and ground support costs.

For spacecraft in orbits above 600 km with no maneuvering ability, alternative means must be investigated to comply with the 25-year rule. The deployment of drag-augmentation devices, e.g., an inflatable balloon, might be an option. However, the increased cross-sectional area needed to accelerate drag effects is off-set by a greater collisional cross-section, increasing the possibility of colliding with other resident space objects. Hence, the drag-augmentation device must be designed to minimize the consequences of such a collision.

Another alternative is the use of an electrodynamic tether. Such tethers can be packed into very small volumes during the spacecraft mission. At the conclusion of the mission, the tether can be extended, producing an effective thrust to lower the orbit. One of the advantages of electrodynamic tethers is their ability to quickly lower the orbit of a satellite with an acceptable tether length. Short operational periods and special tether designs can also significantly limit the chances that the tether might be severed by a collision with small particles.

When mission requirements call for spacecraft to operate at moderate to high altitudes within the LEO region, disposal of their launch vehicle orbital stages must also be addressed. One solution is to trade-off the lift capacity of the launch vehicle with the design of the spacecraft, opting for a slightly more massive and more capable spacecraft to be deployed initially at a lower altitude, followed by orbit raising maneuvers by the spacecraft itself. Hence, the orbital stages can more easily fall back to Earth within 25 years. This deployment plan is particularly attractive for missions using solid-propellant upper stages which cannot reignite and maneuver after shutdown.

The deployment of the Iridium and Globalstar U.S. LEO commercial communications networks in the 1990s provides an outstanding example of how mission requirements can be met while still limiting residual orbital lifetimes of the launch vehicle stages. During 1997–1999, 88 Iridium spacecraft were launched using three different launch vehicles (Proton, Delta, and Long March) from three different countries. Although the operational altitude of the Iridium spacecraft was to be 780 km, all spacecraft were released from their launch vehicles at altitudes between 500 and 650 km. After payload release the Proton orbital stages executed controlled re-entries, whereas the Delta and Long March stages maneuvered into lower, short-lived orbits. Of all 26 orbital stages only one was left in a long-lived orbit following a malfunction; all the others remained in orbit less than 1 year.

The Globalstar constellation posed an even greater challenge for both spacecraft and rocket bodies due to its chosen operational altitude of 1415 km. Again, the deployment scheme called for release of the Globalstar spacecraft at intermediate altitudes by both Delta and Soyuz launch vehicles, followed by spacecraft maneuvers to reach the operational orbits. Delta stages then lowered their orbits using depletion burns, and Soyuz IKAR stages were deorbited over the Pacific Ocean. Only two of 19 stages utilized in deploying 52 spacecraft remain in Earth orbit; the others all re-entered within 4 years of launch (most much sooner).

Having been successfully inserted into orbits at an altitude of 1415 km, the Globalstar spacecraft were faced with a new challenge of proper disposal at end of mission. Lowering the perigees of their orbits to ensure remaining lifetimes of less than 25 years would have required significant propellant expenditures. However, moving to higher orbits above 2000 km would require less energy, and that was the solution chosen. For example, in 2010 four Globalstar satellites, all launched in 1999, began months-long treks to disposal orbits near 2000 km or above.

Rocket bodies and mission-related debris abandoned in geosynchronous transfer orbits (GTOs) should also be disposed of in a manner which preserves the future near-Earth space environment. Two principal options exist: the use of GTO perigees at very low altitudes or at altitudes above 2000 km. Both are used in practice. Normally, perigees of 300 km or lower will naturally decay within 25 years, but the effect of solar-lunar gravitational perturbations must be evaluated, since in some cases these perturbations can cause perigees and orbit lifetimes to both increase. Slightly higher perigees can be employed for upper stages which can then lower their orbits when performing depletion burns.

The other extreme is to select perigees above 2000 km to avoid the 25-year orbital lifetime limitation. The multi-national Sea Launch organization has chosen this technique for several of its commercial missions. Such trajectories need not result in burdens on the payloads being launched and typically are energy neutral or slightly advantageous to the payload.

When NASA began using the large Delta 4 and Atlas 5 launch vehicles, another incentive was found to use high perigee transfer orbits. The upper stages of these launch vehicles pose undesirable risks of human casualty on the Earth following re-entry (see next chapter “Re-Entry Operations Safety”). Consequently, by using high GTO perigees, the re-entry of the stages can be delayed for centuries. NASA has used high perigees for multiple GEO missions, including the GOES 13, 14, and 15 missions flown on Delta 4 launch vehicles and the Solar Dynamics Observatory flown on an Atlas 5 launch vehicle.

The proper disposal near GEO of spacecraft and rocket bodies is also vital to preserving that unique orbital regime. The majority of operators of GEO satellites have accepted the recommendations of the global aerospace community to remove spacecraft from the GEO regime at the end of mission. Failure to do so will result in derelict spacecraft drifting along the geosynchronous arc, posing collision risks to the highly valuable operational spacecraft population, which now numbers almost 400. Due to gravitational perturbations, many of the wandering spacecraft will oscillate around one or both of two gravity wells near 75 E and 105 W longitude. Operational spacecraft near these locations are subject to increased collision risks.

During 2010, 15 of the 18 spacecraft reaching end of mission were maneuvered into higher disposal orbits, although a few of the disposal orbits did not fully meet international recommendations. It should be recognized that many of the GEO spacecraft now concluding their missions were launched prior to the adoption of more recent disposal recommendations, e.g., of the United Nations. Some of these spacecraft were designed when the recommended disposal orbits were as low as 150 km above GEO.

Today, few launch vehicle orbital stages enter nearly circular orbits near GEO. The Proton Block-DM and Briz (aka Breeze) stages and the Atlas 5 Centaur stage are occasional exceptions. For such missions two options exist. The first option is to insert the stage into an orbit which is compliant with GEO spacecraft disposal recommendations and then maneuver the payload down into an operational geosynchronous orbit. An alternative is for the stage to place the payload into an orbit very close to GEO and then for the stage to maneuver to a higher disposal orbit.

Overall, compliance with end-of-mission recommendations designed to preserve the future near-Earth environment appears to be improving with each passing year as spacecraft and launch vehicle manufacturers and operators become better acquainted with the recommendations. Certainly, such measures are in the long-term best interests of not only the aerospace community, but also the much larger population, which relies on uninterrupted space services.

Threats to the Safety of Assets in LEO

There are three major groups of man-made debris objects threatening the safety of assets in LEO: intact satellites and rocket bodies, mission-related components and large debris fragments, and hundreds of thousands of small fragments (“shrapnel”) generated in explosions and collisions (Table 8.9.1).

Tracked debris can be avoided by active maneuvering, but the conjunctions are becoming more frequent and the task of avoidance is getting more and more complicated as the catalog grows. Figure 8.9.1 illustrates how many tracked debris objects cross the orbital plane of a satellite at a 50° inclination. This pattern is repeated every orbit with somewhat different phasing, and the number of conjunctions accumulates rapidly with time. The snapshot in Figure 8.9.1 was taken after the Cosmos–Iridium collision, and the two streams of fragments are clearly distinguishable around 780 km.

The need for avoidance maneuvering has increased dramatically after the Fengyun-1C and Cosmos–Iridium events. Most of the fragments produced in these collisions are still in orbit, and every day they create hundreds of conjunctions with satellites. Table 8.9.2 shows that the fragments of Cosmos and Iridium were the primary cause of collision avoidance for the NASA Earth-observing satellites in 2010 (USA SDE, 2011). Also, the International Space Station had to maneuver on April 2, 2011, to avoid a fragment of Cosmos 2251 (ISS ADD, 2011).

Small fragments in the centimeter range (“shrapnel”) are currently untracked and, therefore, impossible to avoid, but they can disable or seriously damage operational satellites. In a recent hypervelocity impact test performed by ESA (Hypervelocity impacts, ESA), a 1.2 cm aluminum ball traveling at 6.8 km/s made a large crater in an aluminum plate 18 cm thick (Figure 8.9.2). Satellites cannot be currently shielded against such impacts.

For each tracked debris object, there are estimated 30–50 untracked fragments in the centimeter range presenting real danger to operational satellites. Due to the large numbers, this “shrapnel” is the primary threat to operational satellites. New observation techniques will allow tracking of smaller fragments in the future, and the chart of the cataloged LEO population is set to explode upward, while the map of the cross-traffic shown in Figure 8.9.1 will become much denser.

Catastrophic Collisions

Collisions between intact objects (inactive or operational) in low Earth orbits will be catastrophic in most cases, resulting in complete disintegration of the objects. As debris mitigation practices improve, they can become the dominant source of debris pollution in LEO, producing hundreds of thousands of fragments in the centimeter range that are hard to track, but could be lethal for operational spacecraft. Note that because of high kinetic energy, even a relatively small object, such as a 3U CubeSat, can smash a large object, such as a rocket stage, into pieces at orbital velocities (Figure 8.9.3).

Since the beginning of the space age, the annual probability of a catastrophic collision in LEO has been growing approximately quadratically with time, as shown in Figure 8.9.4. This is because it is roughly proportional to the second power of the number of intact objects in the catalog, and the catalog has been growing nearly linearly with time. In 50 years of space flight, this probability has been accumulating to the point where a catastrophic collision has become very likely, and it finally happened. This probability will continue to grow, unless the catalog is substantially reduced by active debris removal.

The consequences of future collisions can be measured by the yield of fragments. The average statistically expected yield in terms of mass can be calculated as

(1)

where are the masses of the objects, are the probabilities of a collision between objects k and n, is the overall probability of a catastrophic collision,

(2)

and is the probability of a catastrophic collision involving object k,

(3)

The values are rather small, and only linear terms are retained in the cumulative probability calculation. The collision probabilities can be calculated using a method developed by Kessler, (1981) or other methods. Objects that successfully maneuver to avoid all tracked objects do not contribute to the totals in eq. (1). The whole LEO debris cloud is dynamic, and the probabilities vary with time, as the orbits and population change. For example, after the Cosmos–Iridium collision, the corresponding terms dropped out of the probability matrix.

In the current LEO debris field, the average statistically expected yield of fragments (eq. (1)) is about 2.7 tons, which is more than the yield of the Fengyun-1C and Cosmos–Iridium collisions combined. Figure 8.9.5 shows the distribution of the expected fragment yields by 1-ton ranges in the current LEO debris field. The last bar covers the range from 10 to 17 tons. We see that the Cosmos–Iridium collision was on the small side. Over 60% of the catastrophic collisions will yield more than 2 tons of fragments. Collisions involving Zenit upper stages can yield up to 17 tons of fragments, but they are much less likely.

Judging by the number of small fragments produced in the Fengyun-1C and Cosmos–Iridium events, we can expect on the order of half a million “shrapnel” pieces in the centimeter range to be released in an average catastrophic collision. Figure 8.9.6 illustrates the aftermath of an average catastrophic collision.

Smaller fragments have a tendency to spread wider than larger fragments, and they have a potential to indiscriminately threaten the safety of all operational satellites in LEO, not only the immediate vicinity of the collision altitude. The air drag will cause a gradual decay of the debris population, but fragments produced in high-altitude collisions and fragments ejected into trajectories with high apogees can persist for many decades after the collisions.

Debris Ranking for Removal

It is now recognized that a single catastrophic collision between intact objects in LEO can negate many years of debris mitigation efforts. The amount of “shrapnel” produced in the Fengyun-1C and Cosmos–Iridium events was comparable to the accumulation of explosion fragments over 50 years of spaceflight. These fragments are currently untracked and impossible to avoid, but they can disable or seriously damage operational satellites.

In order to prevent LEO pollution with fragments produced in catastrophic collisions, large debris objects, the primary source of future “shrapnel,” should be removed from densely populated regions in LEO (Klinkrad, 2006; 2010; Liou 2011; Krag and Virgili, 2011; Levin et al., 2012). The collision-generated debris potential associated with groups of large objects can be estimated by the statistically expected cumulative yield of fragments generated in catastrophic collisions (Levin et al., 2012)

(4)

where is the mass of object k, is the probability of a collision between objects k and n, and is the probability of a catastrophic collision involving object k.

An important feature of the probabilities involved in calculation (4) is their sensitivity to “inclination pairing” observed when is approaching . Figure 8.9.7 illustrates this notion by plotting typical multipliers resulting from “inclination pairing,” as described in (Carroll, 2009). For objects at , the multiplier peaks at (a), while for objects at , it peaks at (b). This happens because the orbits at and precess in the opposite directions, and when they become nearly coplanar, the objects move head-on, greatly increasing the probability of collision, as illustrated in Figure 8.9.8.

Using formula (4), we can evaluate collision-generated debris potential of selected groups of debris objects and analyze its cumulative distributions by location and ownership. Figure 8.9.9(a) shows the cumulative distribution of the collision-generated debris potential by inclination bins compared to the distribution of the number of operational satellites in LEO (b). A comprehensive database of operational satellites is available online (IDOS).

Three clusters stand out, the 71–74°, 81–83°, and the Sun-sync clusters. The Sun-sync cluster, populated with most operational spacecraft, is “inclination paired,” as explained above, with the 81–83° cluster, populated mostly by old upper stages, and the two clusters represent elevated collision threats to each other. Removing just the old upper stages from the 71–74°, 81–83°, and the Sun-sync clusters would make a huge difference. The overall collision-generated debris potential (eq. (4)) would drop by a factor of four.

Figure 8.9.10 shows the impact of removing large debris (Levin et al., 2012). We see that only removal of hundreds of tons of large debris can make a noticeable difference in the collision-generated debris potential in LEO. When many objects are removed, the statistically expected frequency of catastrophic collisions will drop drastically as well. According to formula (4), the largest fragment yields come from the objects with high values of , where is the mass of the object k, and is the probability of a catastrophic collision involving this object. Such objects should be considered for early removal.

There is a range of opinions on how many objects should be removed annually. Conservative suggestions call for removal of five to ten large objects per year (Liou, 2011) and rely on a very high post-mission disposal compliance. Such campaigns would be long-term in nature, and many of the 2200 dead satellites and rocket stages currently in low Earth orbits will persist for very long periods of time. During that time, catastrophic collisions and production of “shrapnel” will continue. An alternative could be a short-term wholesale debris removal campaign (Levin et al., 2012), in which a couple of hundred objects are removed per year, and LEO could be mostly free of large debris objects in a decade or so. In this scenario, there will be no more catastrophic collisions, and no more “shrapnel” will be produced.

Debris Removal with Drag Devices

Drag devices (such as inflatables, solar sails, and electrodynamic drag tethers) may be suitable for some newly launched objects, but they are not well suited for removal of large numbers of legacy debris objects. First of all, they have to be somehow delivered and attached to derelict objects, which may require a lot of propellant and many delivery vehicles. Second, spiraling down in large numbers, they will introduce excessive mutual collision risks if used en masse because of their large collision areas. For example, if passive electrodynamic drag tethers are attached to a few dozen large debris objects, a collision during their slow deorbit process will become very likely (Levin et al., 2012).

Debris Removal with Orbital Tugs

Various mission scenarios of debris removal with rocket-propelled orbital tugs are being considered. The tugs can either move debris objects directly, or they can deliver and attach solid propellant deorbiting kits (Bonnal & Bultel, 2009; Castronuovo, 2011).

To understand requirements and general relations, let us consider a simplified problem of moving K debris objects from circular orbits at altitudes to circular orbits at a lower altitude . The migration will be attempted by N tugs with a dry mass and fuel capacity . Let us disregard the penalties for the inclination and node changes and assume that all tugs are placed in orbits with the same inclination as their targets, and that the differential nodal regression is used to match the nodes. Let us also assume that the tugs have low thrust engines with a specific impulse . Each tug will start at an altitude and spiral up to its next target at an altitude , capture the target, and then spiral back down to the altitude , where the debris will be released for natural decay. We will approximate the delta-V for the transfer between the two orbits as

where ω is a fixed orbital angular rate. The amount of fuel consumed on this round trip will be approximated as

where and are the average masses of the tug on the way up and down, and g is gravity. Statistically, with many trips, the tugs will be carrying half of the fuel on average, and we will use an average value of instead of for the purpose of summation. Then, the total amount of fuel consumed in the process of migration will be

(5)

where

is the total mass of the debris objects, and and are their simple and weighted altitude averages,

Substituting into eq. (5), we find the total mass of fuel, and then, the total mass of the tugs with fuel

(6)

The total mass of the tugs is the lowest when the number of tugs is equal to

(7)

where

Here, is the average mass of the debris objects. The total amount of fuel and the total mass of the tugs with fuel are equal to

(8)

while the amount of fuel per tug is equal to

(9)

and the average number of debris objects removed per tug is equal to

(10)

For bi-propellant, it is possible to deorbit debris by lowering the perigee to some altitude that guarantees quick re-entry. In this case, the required delta-V is approximated as

and formulas (7)–(10) apply with

With a given ψ, the number of tugs is proportional to the number of debris objects, while the masses of fuel and the tugs are proportional to the total mass of the debris, and all values are inversely proportional to the specific impulse. The number of tugs grows with ψ, but their total mass drops. Because ψ is inversely proportional to the dry mass , making as small as possible reduces the total mass. However, it also reduces the amount of fuel per tug, according to eq. (9), which may become insufficient for moving large objects.

Figure 8.9.11 shows the total mass and the number of tugs optimally required to remove all LEO debris over 2 kg as a function of the specific impulse of the propulsion system. The tug dry mass is set to kg. The dots on the left represent the bi-propellant solutions with km, while the lines represent the high-specific impulse (Isp) solutions with km. We see that low-Isp systems would require excessive mass and number of tugs, while cost-effective solutions would require Isp’s much higher than currently available. Some economy promised by bi-propellant fuel depots is not that impressive for this particular task and is offset by higher complexity of the design and operation.

Electrodynamic Propulsion for Debris Removal

Electrodynamic propulsion is propellantless and can meet the high delta-V requirements for removal of large numbers of debris objects from LEO. The electrodynamic thrust is the Ampere force acting on a conductor in the geomagnetic field (Figure 8.9.12). Electrons are collected from the ambient plasma on one end and emitted back into the plasma from the other end. The current loop is closed through the ionosphere. Electron collection can be achieved by biasing bare metal surfaces, while the most efficient electron emission devices in the ampere range are hollow cathodes. They spend a small amount of xenon. Taking into account this expenditure, the equivalent Isp of the electrodynamic system described later in this section is on the order of 200,000 s. This places it far on the right on the performance charts in Figure 8.9.11 and makes it a suitable candidate for wholesale debris removal in terms of performance.

The electrodynamic propulsion technology is not new – it has been in development for over 25 years. The first demonstration in orbit took place in 1993 during the Plasma Motor Generator (PMG) experiment by NASA Johnson Space Center. It used insulated copper wire and hollow cathodes on both ends for electron emission and collection. PMG was the second in the series of four successful flights with J. Carroll’s tethers and deployers (Tether Application Inc.) that included also SEDS-1 in 1993, SEDS-2 in 1994, and TiPS in 1996 (Cosmo & Lorenzini, December 1997).

In 1996, TSS-1R demonstrated effective bare surface electron collection, ionospheric circuit closing with emission nearly 20 km away from the collection area, and revealed the arcing problem at high voltages.

In 1998–2002, NASA Marshall Space Flight Center designed and built the Propulsive Small Expendable Deployer System (ProSEDS) to demonstrate electrodynamic deorbit of a 1-ton Delta II upper stage. J. Carroll designed the tether and the deployer. The system involved 500 m of insulated wire, 4.5 km of bare aluminum wire, and a 10-km non-conducting “pilot” tether, with a 20-kg counterweight at the end. The flight was delayed and then canceled due to changing perspectives on risks to ISS after the Columbia accident. It would have been the first debris removal mission with an electrodynamic tether.

In 1999–2000, J. Carroll designed and built the Mir Electrodynamic Tether System (METS) (Tether Applications Inc.; Levin, 2007). The tether consisted of a 6-km insulated wire, 1-km bare aluminum tape for electron collection, and a 0.5-km pilot tether. A spare 200-kg Manned Maneuvering Unit was to be attached in orbit as a counterweight. The system would draw 2 kW of power from Mir and produce 0.2 N of average thrust along track to keep Mir in orbit without fuel re-supply, allowing the newly formed MirCorp to open Mir to commercial space tourists. Unfortunately, the decision was made to deorbit Mir before there was a chance to test METS.

In 2001, it was suggested that rotation will improve stability and widen the range of angles with the geomagnetic field (Levin, 2007), and J. Carroll designed the first rotating electrodynamic tether system. It featured multiple distributed power nodes to reduce voltages and prevent arcing (a lesson learned from TSS-1R) and bare aluminum tapes for the full tether length to improve performance at low plasma densities, provide more flexibility of control, and greatly increase tether survivability.

Electron collection with a bare aluminum tape was tested by JAXA in August 2010, when a tape 133 m long and 25 mm wide was deployed from a suborbital rocket (Fujii & et al, August 8-11, 2011). JAXA is considering active debris removal with electrodynamic tethers (Kawamoto et al., December 8-10, 2009).

In recent years, the Naval Research Laboratory took the initiative of advancing the electrodynamic propulsion technology and is currently building a 3U CubeSat for the Tether Electrodynamics Propulsion CubeSat Experiment (TEPCE) (Coffey et al., January 12-14, 2010; NRL Programs, TEPCE). The 1.5U end-bodies will be energetically separated by a stacer spring, deploying a 1-km conductive tether. The system will demonstrate electron emission and collection, electrodynamic propulsion, tether libration control, orbit determination and navigation. In addition, it will collect plasma measurements over a wide range of ionospheric conditions.

As the technology matures, it can be integrated into a debris removal system. A candidate design layout of an electrodynamic “garbage truck” is shown in Figure 8.9.13. It includes two end-bodies with controllers and electron emitters, and multiple power nodes with solar arrays, all connected sequentially with reinforced bare aluminum tapes. The tapes are 1 km long, 30 mm wide, and 0.04 mm thick, and serve both as conductors and electron collectors. A 10-section vehicle weighs only 100 kg, including 41 kg of tape, 3 kg in each power node, and 16 kg at each end. Two vehicles can fit into one ESPA slot (Levin et al., 2012).

The vehicle is designed to be highly redundant and survivable. Propellantless thrust and on-board global positioning system (GPS) allow avoidance of all tracked objects by wide margins. The tapes are immune to punctures by small particles in the millimeter range and smaller. The probability of a tape cut by untracked debris in the centimeter range is much lower than a typical probability of failure of the spacecraft avionics. Even if the tape is cut, both parts remain controllable and can deorbit themselves in days, avoiding all other objects.

The entire structure rotates slowly at 6–8 revolutions per orbit. This gives stability and wider range of angles between the conductor and the geomagnetic field, especially in high inclination orbits. All orbital elements, as well as the tether orientation and vibration, can be controlled by varying and reversing the currents in different sections of the conductor (Levin, 2007). Each inboard node produces ~800 W of power, and the vehicle can make large orbit changes in a fairly short time. In deboost mode, additional energy can be extracted from the orbital motion through the emf, substantially increasing the thrust capability. Being propellantless, the vehicle is not limited by the Tsiolkovsky rocket equation and can produce enormous delta-Vs of hundreds of km/s over its operational lifetime.

Figure 8.9.14 compares altitude rates of the rotating system (a) and a conventional vertically oriented system (b) dragging down a 1-ton debris object. The difference is drastic, especially at high inclinations, where most of the large debris is concentrated. There are two reasons. First, for a vertical electrodynamic system to remain stabilized and controllable the thrust must be an order of magnitude less than the tension produced by the gravity gradient while the rotating system can apply all the thrust it is able to produce electrically. Second, a vertical system cannot produce much of a thrust along-track in near-polar orbits because of the orientation of the geomagnetic lines, while the rotating system can spin normal to the orbital plane and get very good “traction” with the geomagnetic field.

Theoretically, a dozen such vehicles launched on one ESPA ring (two per slot) could remove all intact debris objects from LEO (approximately 2200 objects totaling 2000 tons) in about 7 years, and it would take only four vehicles and 7 years to remove all upper stages from the 71–74°, 81–83°, and Sun-sync clusters, reducing the collision-generated debris potential in LEO by a factor of four. This is very promising, but the electrodynamic propulsion technology needs time to mature and go through flight testing before it can be used operationally.

Debris Capture

Even though it may seem straightforward, debris capture is challenging and riddled with engineering problems (Kaplan & et al, September 2010). In the most frequently considered capture scheme, the remover spacecraft rendezvous with a debris object, moves into its close proximity, extends a robotic arm (or arms), and grabs onto a protruding feature, such as the rim of the thruster nozzle. The underlying technologies have been in development and operation for quite some time. Robotic arms have flown on many Shuttle missions since 1981, and the International Space Station has used a robotic arm since 2001. Autonomous robotic systems are being developed for satellite servicing (Henshaw, 2009). Key remote servicing functions have been demonstrated during the Orbital Express mission in 2007. It is believed that future systems of this class should be adequate for debris capture (Teti, November 11-12, 2011). Also, a new generation of miniature capture systems that can be mounted even on CubeSats is emerging (Cleaning up Earth’s orbit).

Most old debris objects are expected to be either gravitationally stabilized or rotate slowly, due to eddy-current damping in their aluminum structure (Williams & Meadows, 1978; Bonnal, June 10, 2011; Praly et al., July 4-8, 2011). Some debris, however, will tumble fast enough to complicate grappling. The remover spacecraft would have to approach and literally “land” on the tumbling object, avoiding appendages, and attach itself quickly, or the object will swing the “lander” away. This is risky, particularly for electric tugs with large solar arrays. One of the suggestions is to use a brush contactor to detumble the object before capture (Nishida & Kawamoto, February 2011).

A radically different idea is to avoid grappling altogether and steer the debris object from a distance using an ion beam. The concept is called the Ion Beam Shepherd (IBS) (Bombardelli & Pelaez, June 2011). It is illustrated in Figure 8.9.15. A spacecraft with electric propulsion will rendezvous and fly in formation with a debris object and direct a low-divergence, high-specific-impulse ion beam towards the object to exert a decelerating force . The reaction force will push the spacecraft away from the target, but the main engine will compensate for it by thrusting in the opposite direction (). The onboard control system will maintain the distance to the debris, and the invisibly linked pair will spiral down to a disposal orbit, where the debris will be left for natural decay, while the remover will proceed to the next target.

Ion engines are the leading candidates for this kind of missions (Bombardelli & Pelaez, June 2011). They have a high specific impulse of ~3000 s, high efficiency (~65–80%), and a low-divergence plasma plume of less than . With a power supply of 2–3 kW, they provide thrust ~0.1 N. It takes approximately as much fuel to maintain the ion beam as to thrust. Therefore, the effective specific impulse will be ~1500 s when paired with a large debris object.

The electrodynamic system shown in Figure 8.9.13 has an advantage of being propellantless, but it requires different capture techniques (Carroll, December 2, 2002). For debris removal, the payload managers at each end carry many large, lightweight nets (~50 g each). To catch a debris object, the tip velocity is matched with the target, a spin-stabilized net is extended from the payload manager by the centrifugal force at the end, the target is approached at a few meters per second under manual control from the ground, and the net is enclosed around the target. The multi-newton tension in the net induced by the rotation of the entire system rapidly synchronizes the object’s rotation with the tether rotation without any special control, even if the object is tumbling up to 1–2 rpm, and continues to hold the object firmly in the net during deorbit. Figure 8.9.16 illustrates schematically how the restoring torque is applied to the debris object. The tether tension is induced by the rotation of the entire system, while the tension in the net straps reacts to the rotation of the debris object.

Nets have an advantage of being indifferent to the shapes and sizes of the objects, but they are still in a prototype stage. An alternative technique would be for the electrodynamic system to carry a small autonomous robotic unit that is released in close proximity to the debris, approaches and grabs onto a suitable debris feature, extended a mating interface, and is recaptured from the payload manager (Carroll, December 2, 2002). Such symbiosis between robotic capture and electrodynamic propulsion would be mutually beneficial.

Disposal of Debris

The naturally occurring clean-up process in LEO is due to the air drag. It slowly drives debris to re-entry in the atmosphere. Most debris removal schemes are simply assisting this natural decay process by moving debris objects to lower altitudes with much shorter orbital life. Altitudes below the ISS may be preferred for the reasons of ISS safety.

As discussed earlier, debris removal campaigns are much more efficient when they utilize propulsion with high specific impulse. Electric and electrodynamic propulsion systems are the leading candidates at this time; however, they cannot provide controlled re-entry for large debris objects because of low thrust levels. Uncontrolled re-entry may be acceptable for objects that are expected to burn up in the atmosphere, but it raises concerns for objects whose parts may survive re-entry (Bonnal, June 10, 2011). Controlled re-entry can be achieved with chemical propulsion, but if it is used to move large debris all the way from their original orbits to re-entry, the removal campaign becomes much less efficient.

One way to solve this problem would be to use electric or electrodynamic propulsion to move the debris below ISS and release it there with a small re-entry package. This package could contain either a chemical thruster with a minimal amount of fuel or a small device for aerodynamic steering during re-entry.

An alternative way would be to avoid re-entries altogether and move large debris objects from crowded regions in LEO to intermediate altitudes for storage and possible future utilization. The objects, however, should not be left in random orbits, because this will create another congestion. Instead, they could be gathered and assembled in several maneuverable collections (Carroll et al., October 30–31, 2010). Each collection can be propelled electrodynamically without fuel expenditure for the purposes of collision avoidance and orbit maintenance. When technology is developed, the collections may be reprocessed into construction materials (Carroll et al., October 30–31, 2010). Note that there are 1000 tons of old upper stages in LEO, and much of their mass is aluminum alloy. It took a lot of money to put them in orbit, and they still have some residual value. This value should not be overlooked.

The task or debris relocation to intermediate altitudes is much less demanding than reentering the objects. Let us consider orbital tugs propelled by ion engines with a specific impulse of 3000 s. Using formulas (7) and (8), we estimate that it would take 21 tugs with a total launch mass of 11 tons to move all old upper stages currently orbiting above 600 km in the 71–74°, 81–83°, and the Sun-sync clusters to collections around 600 km. Alternatively, electrodynamic vehicles can do this job with a total launch mass ~400 kg. This campaign will greatly improve the situation in LEO, reducing the probability of a catastrophic collision (eq. (2)) by a factor of nearly three and reducing the collision-generated debris potential (eq. (4)) by more than 70%.

Placing three collections in each cluster and spacing them at nodal distances, a full nodal sweep utilizing differential nodal regression could be complete in less than 7 years. This way the tugs will spend the least amount of fuel on orbital plane changes, as assumed in formulas (7) and (8). On average, around 100 tons of old upper stages will be accumulated per collection. An electrodynamic propulsion system of the kind needed for such collections has already been developed and built once to maintain the orbit of the Mir station (Tether Applications Inc.; Levin, 2007) (it was not flown because of the decision to deorbit Mir). Mir had a comparable mass of 136 tons at the time and experienced a much higher air drag than the collections would.

Debris Removal with Lasers

There is one debris removal technology in development that does not require launches at all. It involves lasers operating from the ground (Phipps & et al, May 2012). If a low-intensity laser is pointed at an orbiting object, the light pressure from the laser beam can slightly change the velocity of the object. However, this effect is too weak for debris removal. A focused beam from a high-intensity laser can cause ablation of the surface of the object, when some of the surface material is melted, vaporized, and ejected from the surface in high-velocity jets. The reaction force from the jets acting on the object can be four to five orders of magnitude higher that the light pressure itself and can be used in debris removal. Pulsed ablation is more efficient than continuous heating of the target, but it requires multi-kilo-Joule pulses and mirrors larger than 10 m in diameter to focus a beam of desired intensity on the target. Recent advances in lasers and telescopes could make such systems feasible.

Objects under 1 kg could be re-entered in a single pass from low orbits by about a thousand 5-nanosecond pulses. But larger targets would require multiple passes, during which they will be pulsed thousands of times, gradually lowering to re-entry. Targeted re-entry for objects that may not burn in the atmosphere is believed to be possible (Phipps & et al, May 2012).

The lasers would have to be combined with telescopes for target acquisition and accurate pointing. Due to the finite speed of light, the lasers will be shooting at positions up to 50 m ahead of the last observed positions of the targets at ranges up to 1000 km. Accurate prediction and execution will be crucial, and a large array of safety measures would have to be put in place to support these operations.

Debris Removal Service

When technical and legal hurdles are overcome, debris removal can be offered as a routine commercial service to launch and spacecraft operators to keep LEO clear of dead satellites and spent stages. In order to make economic sense to them, debris removal should cost much less than a typical launch cost per kilogram. Early estimates indicate that selective removal with rocket-propelled tugs may end up being expensive (Bonnal & Bultel, December 8–10, 2009). It may be undertaken by governments, but may not appeal to commercial operators. Electrodynamic propulsion has potential to drive the cost down toward perhaps $400 per kg at the current prices (Levin et al, 2012). This is where wholesale removal of large debris makes sense not only in terms of preventing LEO pollution, but also economically by offering service at wholesale prices. Ground-based lasers could also drive debris removal costs toward and below $1000 per kg (Phipps & et al, May 2012).

Conclusions

Catastrophic collisions between large objects in LEO will produce hundreds of thousands of debris fragments in the centimeter range (“shrapnel”). The fragments of these sizes are currently untracked and impossible to avoid, but they can disable or seriously damage operational satellites. Statistically, the fragment yield of an average catastrophic collision will be comparable to the yield of the Fengyun-1C and Cosmos–Iridium events combined. To prevent further pollution of LEO with collision fragments, a substantial number of large debris objects, the primary source of future “shrapnel,” should be removed from densely populated regions in LEO.

The task of debris removal is challenging. Selective removal can be performed by rocket-propelled orbital tugs equipped with robotic capture units. Wholesale removal, however, is too demanding for the rockets because of the sheer amount of fuel and vehicles required. Electrodynamic propulsion can assist in this task. It is propellantless and capable of delivering very large delta-Vs over the operational lifetime of the electrodynamic vehicles. Such vehicles can be combined with robotic capture units and deorbit packages for targeted re-entry of large debris. Wholesale removal of small and possibly large debris could also be achieved using ground-based lasers. These technologies still have to mature before they can be used operationally.

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