Final Report Summary - REVUS (Reducing the Vulnerability of Space Systems)
Executive Summary:
The population of debris in space is continuously growing, and could become an increasing risk for the survivability of space assets. Unlike large debris, small debris cannot be tracked nor avoided; a collision with a LEO satellite could result in degradation, or even loss of the mission.
The ReVuS project aims to propose design solutions at system and satellite architecture levels to reduce the vulnerability of future LEO satellites to these small-sized debris (0.1 to 5 cm). It is supported both by analyses carried out with adequate tools and models, and by tests of shielding materials.
A vulnerability analysis has been carried out to evaluate the probability of having small debris colliding a satellite, and the effects of such an impact (probability of penetration, and risk of mission failure). The level of vulnerability of a satellite depends on its orbital position and on its configuration
This analysis was done on two reference LEO satellites (optical and SAR) characterized by a different configuration and a different altitude. Particles in the size range 1-5 mm dominate the failure risk (due to debris impact), with significant peak at around 2-3mm. The debris environment varies significantly with the orbit altitude, being much severe at 800 km than at 500 km; highest impact flux comes from the front direction, and the critical satellite units are the external ones or those mounted close to the front side.
Solutions at system level (fractionated satellite, distributed architecture, in orbit spare) cover the full range of debris. They impact the overall mission and system. They have a high efficiency as the probability of losing the mission is quite negligible, but are penalized by the cost and, for some solutions, by the technical maturity. In orbit spare solution could be attractive for mission involving several identical satellites on the same orbit.
Solutions at satellite architecture level encompass the repositioning of critical equipment units (relocation, removal from front side, rotation, hide behind other units), the re-orientation of the satellite through the polar areas, the use of external redundancy, additional margins and the shielding of critical equipment or of areas protecting several critical units.
The shielding concept depends on the location of the equipment in the spacecraft and of its local environment. While defining ways to obtain efficient shielding concepts adapted to equipment configuration, alternative shielding building blocks have been identified for reinforced MLI, reinforced structural panels, intermediate layers, and reinforced equipment box. A preliminary test campaign has allowed comparing alternative building blocks in each family with respect to their capacity of protection, resulting areal density and impact on satellite.
Based on these results, shielding solutions have been elaborated for each of four selected equipment configurations, and tested in order to determine their ballistic limit at 15 km/s. These solutions allow protecting the equipment against particles of at least 3mm (6 mm for the solutions selected for the tank). Some of the enhanced configurations appear lighter than the basic ones. The tests show that stand-offs have the biggest potential, heavy stainless steel mesh are valuable to reinforce the first layer, aramid and Nextel show a good mass-to-shielding ratio for intermediate layer.
An evaluation of all satellite architecture level solutions has been done with respect to a set of criteria, including performances, additional mass, volume, harness and connectors, maturity level, cost aspects. Solutions with the highest reduction of satellite risk of failure (such as external redundancy, shielding of equipment) are penalized by their impact on satellite layout (additional mass, volume, harness) and by cost aspects.
To minimize impacts on the satellite, it is worthwhile to take into account the protection against debris since the early phases of development, and also to rely on a combination of solutions rather than on a single one.
In addition to their protection function, some of the solutions could also improve the reliability of the satellite (e.g. external redundancy), or its performances (additional margins, in-orbit spare). In particular, solutions improving reliability of the satellite would be seen extremely positively by the space insurance.
Each solution is not applicable to all equipment or to all types of satellites. For instance, solutions with additional mass and volume are better suited to large satellites rather than small and compact ones. There is no generic solution applicable to all satellites; indeed, most of the performance evaluations are geometry dependent. The set of solutions will be specific to each satellite design, depending on application, orbital parameters, satellite configuration and needs.
ReVuS brings a palette of solutions with their potential gains and impacts. It improves the quality of the design of European satellite.
Project Context and Objectives:
Context
The population of debris in space is continuously growing, especially as a result of the collisions that have occurred in the past few years. It appears that even if guidelines to minimise the post-mission orbital life are applied by the space community, this population of debris will continue to grow in the years to come. Thus the presence of debris could become an increasing risk for the survivability of LEO space assets. Indeed, the probability for a satellite to collide with orbital debris, although very low, could become non negligible. This probability and the relevant consequences differ according to the size of the debris:
Debris larger than 10 cm can be detected, catalogued and tracked from ground, so that collisions can be predicted and avoidance manoeuvres can be performed.
But smaller debris cannot be tracked, and their presence remains a threat for space assets.
While the very small debris (< 0.1mm) have an energy low enough to be absorbed by the structure materials, the small sized debris (from 0.1 mm to 10 cm) could generate critical damages, such as the loss of a part of the mission, the degradation of the performances of the satellite, or even the loss of the mission in case of collision. Indeed, their energy is high enough to penetrate the satellite structure and they constitute a risk for internal mounted units. The environment models show that a satellite located at an altitude of 800 km will have in the coming years an average of 40 impacts/m2/year with this type of debris, among which 0,04 particle/ m2/year will penetrate inside the satellite.
In order to mitigate this risk, several approaches are currently considered, as shown in Fig 1-1:
• Prevent: the aim is for new vehicles (spacecraft, upper stages) to avoid staying in operational orbits too long after the end of their lifetimes, thus increasing the population of debris. It relies on the use of guidelines, laws, rules and standards to limit the post-mission presence of vehicles in operational orbits
• Avoid: it consists of the satellite performing manoeuvres to avoid a collision with debris. It is currently used by the main satellite operators. It concerns large size debris which can be detected and tracked from the ground
• Remove: it deals with debris removal. Studies are currently done on active debris removal. A first objective is to remove large objects (dead satellites and upper stages).
• Survive: It is related to the small size debris that cannot be tracked. It consists of defining design rules to minimise the effects of debris impacts on the satellite and its mission.
Fig 1-1: Approach for debris risk mitigation
Objectives
The ReVuS project is an answer to the Survive approach.
It aims to propose design solutions at system and satellite architecture levels to reduce the vulnerability of future LEO satellites to these small-sized debris (0.1 to 5 cm). It is supported both by analyses carried out with adequate tools and models, and by tests of shielding materials.
In order to achieve this aim, the ReVuS project had to meet several objectives that drive the logic of the project, illustrated on Figure 1-2.
Fig 1-2: Logic of the ReVuS project
The first objective was to evaluate the probability of having small debris colliding a satellite, and the effects of such an impact (probability of penetration, and risk of mission failure). It was the object of the vulnerability analysis. In order to have quantitative figures, two reference LEO satellites (optical and SAR) characterized by a different configuration and a different altitude (825 km for optical one, 515 km for the SAR one) have been selected. This analysis was based on the use of the impact risk assessment SHIELD3 tool which evaluated the probabilities of penetration of small debris particles in the satellite and in the equipment and determined the failure probability for all equipment parts considered.
The SHIELD3 tool relies on information about the design of the satellite and on the directional distribution of the flux of debris on the satellites computed using the MASTER 2009 environment model.
This analysis provided information on the effect of the orbit altitude, on the flux of particles impacting and penetrating the satellite for different altitudes, on the probability of satellite failure in the two reference cases, on the range of debris that is contribute the most to the risk of failure, on the satellite units that appear critical with respect to debris impact. All these results are needed for the next objectives.
The second objective was to define and analyze the possible solutions at system level. Several solutions have been identified, and four of them have been analyzed and kept. Such solutions have a direct impact not only on the satellite itself, but also on the mission and on the system definition. Evaluation of their efficiency to reduce vulnerability of the satellite, of their maturity level and of their impacts on the mass and cost aspects has been done.
The third objective was to define and analyze the possible solutions at spacecraft architecture level. These solutions encompass the repositioning of critical equipment units (it includes mainly the relocation of equipment in a safer zone, the increase of distance between the critical equipment and the front side, the rotation of equipment, to hide critical equipment behind other units), the re-orientation of the satellite when flying through the polar areas, the use of external redundancy, the additional margins and the shielding of critical equipment or of areas protecting several critical units. These solutions have been evaluated with respect to their accommodation on the spacecraft and the impacts on the spacecraft configuration in terms of layout, mass, environment, thermal aspects, etc.
In order to have qualitative figures, several improved configurations have been defined for the two reference satellites, combining relocation of some critical elements and shielding solutions. The objective was to evaluate the achieved reduction of probability of failure with SHIELD 3 tool and the resulting additional mass.
The fourth objective was to define and test possible shielding material and shielding configurations.
The shielding concept (including two layers or more, various materials, distance between layers) depends on the location of the equipment in the spacecraft and of its local environment (structure, thermal).
Alternative shielding building blocks have first been identified for reinforced MLI, reinforced structural panels, intermediate layers, and reinforced equipment box. The aim was to define solutions with different materials, different thicknesses and different arrangement of the material layers, and to evaluate the performances of these alternative solutions in order to compare them in term of performances (efficiency of protection) and mass areal density) through a preliminary test campaign.
Based on these results, shielding solutions, sized to protect the equipment against particles of at least 3mm, have been elaborated for four selected equipment configurations, and tested in order to determine their ballistic limit at 15 km/s. These solutions have been evaluated with respect to their performances, impacts on satellite configuration (mass, volume, structural constraints), on environment and on cost.
The fifth objective was to evaluate the benefits of all the proposed solutions, their resiliency to debris impact and the interest of combining them.
All these potential solutions have been assessed with respect to a set of criteria including their performances (in terms of reduction of probability of failure of the satellite), additional mass, volume, harness and connectors, maturity level, cost aspects, added value. In particular, the interest of combining several solutions has also been assessed. One objection of this evaluation was also to identify the applicability of each solution to the various types of missions and satellites in order to know what set of solutions could be used for each category of satellite.
The sixth objective was to propose new design rules to increase the robustness of European LEO satellites in case of impact with small debris.
A set of such design rules has been elaborated to assist the satellite manufacturer with the choice and implementation of protection solutions; they will be presented to international organizations involved in the development of guidelines and standards.
The last objective was to disseminate the results of the project. Indeed, a main purpose of the ReVuS project was also to ensure that the results of the work are known in the communities outside the ReVuS consortium. That has been achieved through presentations to various conferences, publication of papers and more particularly a workshop open to public.
Project Results:
The ReVuS project targets the small-sized debris, from 0.1mm to 5 cm. It aims to define design solutions in order to reduce the vulnerability of future LEO satellites to this range of debris.
The project followed a three-step approach, as illustrated in Figure 2-1:
• the vulnerability analysis, to evaluate the effects of a collision of a LEO satellite with small-sized debris, the critical parts of the satellite, and the risk of mission degradation
• the identification and analysis of potential solutions at system level, and at satellite architecture level, with a focus on the shielding concepts and shielding materials
• the resiliency analysis, aiming at evaluating the resiliency of the selected solutions with respect to debris impact and at proposing design rules and standards.
Figure 2-1: Logic of the ReVuS project
VULNERABILITY ANALYSIS
The vulnerability analysis has been carried out to evaluate the probabilities of having small debris impacting the satellite and to determine the effects of such an impact.
As shown on the Figure 2-2, this analysis is based on the use of the impact risk assessment tool SHIELD3 which evaluates the probabilities of penetration of small debris particles in the satellite and in the equipment. In order to have quantitative figures, two reference satellites, an Earth Observation optical satellite and an Earth Observation radar satellite, have been taken into account. They have different configurations and orbit altitudes: the optical satellite has a deployed solar array and is located at 820 km altitude, 98,7 deg inclination while the radar satellite has a rigid body mounted solar array and is located at 515 km altitude, 97,4 deg inclination.
Figure 2-2: Approach for vulnerability analysis
The SHIELD tool needs information about the design of the satellite (geometric data file, material description, information on redundancies, updated damage equations) in order to draw correct conclusions on spacecraft vulnerability while having a significantly reduced complexity of the model.
In parallel to the reference satellite models, the directional distribution of the flux of debris on the satellites has been computed using the MASTER 2009 environment model for the “Business-as-usual” (BAU) scenario, which assumes that current practices are to continue in the future.
A period of 10 years following January 1st, 2020 has been considered for the analysis.
The debris environment varies significantly with the orbit altitude, being much severe at 800 km than at 500 km. As shown on Figure 2-3, the highest impact flux comes from the “front right” and “front left” directions in the case of optical satellite. The azimuth distribution is flatter in the case of SAR satellite. Subsequently, the critical satellite units are the external ones or those mounted close to the front side.
An evaluation of the flux of debris impacting, and of the flux of debris penetrating the reference satellites on their different faces has been done for the successive ranges of particles diameter. The results are summarized in Figure 2-4. It appears that:
• On average, debris flux level for SAR satellite is three times below that of the optical satellite.
• Debris below 1mm diameter has a high probability of impact, but only few penetrate the satellite, so that the effects on equipment are very low.
• Debris particles with a diameter in the range (1-10mm) impact and penetrate the satellites 100 times more often than debris particles in the range (10-50mm).
• The risk of being impacted or penetrated by small debris is much higher at 800 km than at 500 km
Figure 2-3: Illustration of the flux distribution for the two reference satellites
Figure 2-4: Flux of particles impact and penetrating the reference satellites
The SHIELD tool has evaluated the probability of penetration of the resulting debris particles in the satellite equipment. This evaluation is illustrated on Figure 2-5. Then, taking into account the redundancy scheme, in particular whether the equipment has an internal or an external redundancy, an evaluation of the probability of no failure of the satellite has been derived. Figure 2-6 illustrates the probability of failure of the reference satellites as a function of the diameter of debris particles. This is only applicable to the two reference satellites as the result depends highly on the size of satellite, its orbit, its layout (deployed or rigid body mounted solar arrays, etc). These figures cannot be taken as general/average numbers. In addition, they depend on the criteria of failure that have been taken into account.
As a main result, the vulnerability analysis highlights that particles in the size range 1-5 mm dominate the satellite failure risk (due to debris impact), with significant peak at around 2-3mm. The impacts of particles with larger diameter, above 1 cm, will result obviously in more dramatic consequences (higher risk of losing the mission), but the probability of such an impact is much lower, as illustrated in Figure 2-4.
However, the penetration of a debris particle in the satellite, and in equipment does not mean the loss of the satellite. Indeed, the potential damages that could result from such penetration could be:
• the degradation of performances of the satellites, resulting from the loss of resources like battery, the degradation of solar cells, radiators, tank leakage, etc;
• the reduction of the satellite reliability (loss of redundant equipment).
• The degradation of the mission (loss of instruments or payload electronic units)
• the loss of the mission, that could result from penetration of debris in the tank (with a risk of explosion or only leakage depending on particle size and impact conditions), or in a non-externally redundant equipment (in the case of internally redundant equipment, the level of failure depends on the internal architecture).
Exposed functional surfaces, which are not protected, such as solar arrays, are in general designed to tolerate a debris impact flux. Such surfaces have not been considered in the evaluation of probability of no failure.
The radar satellite presents a very low vulnerability to small debris, with a high Probability of Non Failure (PNF) due to its cylindrical shape, with axis along velocity vector and rigid body mounted solar arrays, and its low altitude (515 km) outside the zones where the density of debris particles is high.
Figure 2-5: Illustration of the flux of particles penetrating the SAR satellite and probability of penetration in the equipment.
Figure 2-6: Illustration of Probability of failure as a function of debris size for the two reference satellites.
PROPOSED SOLUTIONS
Based on the results of the vulnerability analysis, two main categories of solutions (Figure 2-7) have been defined: solutions at system level, and solutions at satellite architecture level, which includes the shielding solutions.
A list of proposed solutions has been established for each category and is given in Figure 2-8.
Figure 2-7: Categories of solutions.
Figure 2-8: List of proposed solutions.
SYSTEM LEVEL SOLUTIONS
The system level solutions aim at mitigating the risk at system level. Such solutions can take into account the full range of debris size. They have a high efficiency as the probability of losing the mission is quite negligible, but are penalized by the cost and, for some solutions, by the technical maturity.
A candidate solution is the fractionated satellite concept (see Figure 2-9), which consists in sharing some functions of a satellite (computing capability, communications with ground, payloads, etc) on separate modules forming a cluster, based on wireless communications and interconnecting network. It improves the maintainability and the responsiveness to unexpected events. With an adequate distance between the modules, a collision with debris could lead to lose a module, but not the complete mission. In the case of the optical satellite, the probability of losing the mission is 50 times lower that the one of a monolithic satellite, but the probability of losing at least one module is 3 times higher than the probability of losing the monolithic satellite.
Another solution is the distributed system concept (see Figure 2-9), which will adapt the principles of existing terrestrial wireless systems to distributed space system architectures. It consists in a novel generic architecture for fault tolerant distributed on-board computing. The distributed computing system is comprised of multiple nodes. This approach is based on task migration. The computing system can be distributed within the spacecraft, or over several separated modules, linked by wireless communications. This latter case is the system level solution.
A third possible solution is the in-orbit spare. This solution could be attractive for mission involving several identical satellites on the same orbit.
Figure 2-9: Illustration of system level solutions
SPACECRAFT LEVEL SOLUTIONS
At spacecraft architecture level, various types of solutions have been considered and evaluated. They are described in Figure 2-8.
Each solution has been assessed with respect to common limitations and design principles generally used, taking into account implications on design, cost, performances, etc (qualitative aspects).
The solutions implementing shielding are discussed in the next section.
It appeared that most interesting solutions are:
Re-orientation
This solution consists in adapting the orientation of the satellite to expose less critical side during fly-through of critical zones (mainly the poles area). The attitude of the satellite could be slightly modified around yaw or pitch axis, thus exposing an angled surface to the major flux vector.
Equipment compartmentalisation
This solution consists in implementing a protection (layer) inside the equipment that has internal redundancy in order to protect the redundant part (or the nominal part) from the penetration of a debris inside the equipment. As a result, the equipment is modified, with additional layer and boards. It is almost similar to the juxtaposition of two identical equipment units not internally redundant.
Positioning into a less critical area
This solution consists in positioning critical units (those that have a low PNP) in less exposed area, typically in the rear side. Such solution has impacts on the satellite layout (additional free volume) and AIT, unless it is taken into account in the early phase of the design.
Hide behind less fragile equipment
This solution consists in placing critical units (elements with low PNP) behind less fragile (or massive) elements (such as a passive redundant equipment) that will provide a protection. This solution has impacts on satellite configuration and AIT if not taken into account in the early design phases. It could be interesting for harness or external equipment.
Increase distance between equipment and wall
This solution consists in increasing the distance along the debris flux vector between the equipment and the wall (located in front of equipment with respect to main debris flux) in order to distribute fragments on a wider surface (thus reducing surface energy). It can be done without significant impact except on the layout due to the need for sufficient volume for equipment mounted on panels that are perpendicular to the flux vector. For the other equipment, the constraints generated on the mass, structure and thermal dissipation, plus the volume are probably too penalising with respect to the gain.
Rotation of the equipment
This solution consists in slightly rotating the equipment in order not to have a face perpendicular to the flux of debris and thus to reduce the effects of impacts of debris on the equipment box.
Physical segregation of redundant units
This solution consists in segregating physically the redundant units (implemented on two different areas of the satellite) rather than having both units inside the equipment. This solution will need additional volume, harness, switching unit (that has a risk of failure), but the common causes of failure are avoided.
In case of new satellite, the various impacts are reduced as there is no existing layout.
Addition of external redundancy
This solution consists in adding spare (redundant) units in order to increase the PNF of the function when the operational equipment items are critical with respect to debris (low PNP). For instance, to add a battery, or an external sensor.
Distributed architecture
This solution consists in implementing a distributed architecture for the data management system (computers and RTU): functions are shared over several computers, each being able to take over functions of other ones. It is similar to what has been quoted in the system level solutions, but applied to the satellite architecture (with wired links). A reliability and availability analysis shows that the distributed architecture (distributed OBC) has more reliability and availability as compared to centralised system. It can provide more computing performance because of inherent availability of multiple processing units. A prototype of a fault tolerant distributed OBC (Fig 2-10) with three nodes has been tested: the reconfiguration time, for the TDMA communication scheme, is approximately the same regardless of the number of active modes; each task resumes its data state from the current data state rather than initialising as in legacy systems; high data rate protocols are needed for transfer of large task data states; power consumption is higher than in centralised system.
This solution leads to new avionics architecture, and a new satellite layout. As compared to a centralised system (with an internally redundant computer), there will be additional mass, harness, power consumption, AIT activities (especially for tests) and probably higher cost. When starting from the early design phase, the resulting impacts will be low.
Additional margins
This solution consists in considering additional margins in the sizing of resources to take into account the possible impacts of small debris and thus to avoid a degradation of resources below the level required by the mission. Indeed, for resources like solar arrays or radiators, it is not possible to define a protection. The architecture of the solar array shall be such that the loss of power due to a debris impact (hole) on a cell is minimised. In the case of radiators, oversizing shall be avoided not to modify the thermal balance.
SHIELDING CONCEPT
The shielding solutions will have a significant impact on the mass and on the layout of the satellite. Thus, they cannot be sized to protect the satellite or the equipment units against the full range of debris size.
The vulnerability analysis has shown that a ballistic limit of 3 to 4 mm at 15 km/s is necessary to reduce significantly (up to 50%) the probability of failure (due to debris) of the satellite. As illustrated in Figure 2-11, this ballistic limit objective is confirmed for most of the equipment; for the propellant tank, there is a need for a ballistic limit of 5 to 6 mm to reduce significantly the probability of penetration.
Figure 2-11: Illustration of ballistic limit objective for two equipment of the SAR satellite
The assessment of the needed shielding configuration depends on the location of the equipment in the satellite and its local environment (structure, thermal hardware, etc.).
The analysis of the reference satellites, and also of the current and future LEO satellites, has led to identify fifteen basic configurations of equipment.
The shielding performances of each basic configuration have been estimated by computing their ballistic limit using published data. This estimation confirms the need to protect the critical equipment in order to achieve the required objective.
Shielding principles
A typical efficient shielding has a multi-layer configuration characterised by the number of layers, the thickness and material of each layer, the distance between layers and the orientation of the layer with respect to impacting particle. Each layer has a specific function: the bumper layer (the first one) breaks up and melts the projectile, the inner layer traps the secondary debris and the back wall (usually the box wall) provides the last line of defence. A monolithic shield appears inefficient in terms of mass versus protection.
To obtain an efficient shielding adapted as far as possible to a basic configuration, different possibilities can be considered:
• Re-use the existing structural or thermal elements (typically sandwich panels, multi-layer insulation, equipment box wall), improve their efficiency by increasing thickness or changing the distance between the items
• Change the materials
• Addition of intermediate layers
Shield building blocks and preliminary tests
Based on these principles, the following families of shield building blocks have been defined:
• Reinforced MLI: there are 8 shield building blocks within this family
• Reinforced sandwich panel (with Al skin or CFRP skin): there are 13 building blocks in this family
• Intermediate layer: there are 5 building blocks in this family
These shield building blocks have been investigated through the preliminary test campaign, with the objective to evaluate the characteristics of the shield building blocks and make a ranking between the solutions within each family.
Figure 2-12: Shield building blocks test conditions
The tests have been performed at Fraunhofer EMI’s two-stage light-gas guns at 7 km/s. All the samples required for the tests have been manufactured by Tencate Advanced Composites.
The test conditions are illustrated in the Figure 2-12. The shielding components are placed within a set-up that is representative for their occurrence within a spacecraft. The targets are impacted with nominally identical impact conditions above their ballistic limit. Witness plates are placed behind each target. The first witness plate behind the target (WP1) is considered somewhat representative of module walls.
Examples of the shield building blocks state after the tests are shown in the Figure 2-13 for the Al sandwich panel family. In order to compare the performance of the shielding bricks, the penetration capability of the most damaging fragment impacting the witness plate simulating the module wall (WP1) is estimated. This penetration capability is given in terms of the penetrated areal density of the shield. This number includes the (nominal) areal density of all layers that would have been necessary to stop the impacting particle.
This penetration capability is a measure of the quality of the investigated sample. It describes both the sample’s ability to disperse the fragment cloud over a larger area, and (especially for intermediate layer samples) to decrease a fragment cloud’s energy.
Using this number, the different shield building blocks can be compared against each other.
The figures 2-14 and 2-15 give the results of the preliminary tests for the reinforced MLI and for the Al sandwich panel.
Figure 2-13: Examples of Al sandwich panel shield building blocks status after tests
Figure 2-14: Penetrated areal density plotted vs. sample areal density for the MLI targets. Filled symbols indicate WP1 perforation. Solid line is identity.
Figure 2-15: Penetrated areal density plotted vs. sample areal density for the Al sandwich panel targets. Filled symbols indicate WP1 perforation. Solid line is identity.
From these preliminary tests, it appears that:
• The performances of MLI with respect to shielding can be increased by adding a stainless steel mesh (1.4 1.3 on Fig 2-14), or by adding a reinforced MLI layer with stand-off distance (1.5). These solutions have a mass impact.
• The performances of Al sandwich panel could be improved by substituting the honeycomb with either foam (2.6) or corrugated plate (2.7) that allows a greater dispersion of fragments on subsequent layers, or by introducing an additional Al layer in the middle of the core (2.3) parallel to the face sheets, causing additional shocks in fragments.
• For CFRP sandwich panel, embedding tungsten particles into the face sheets enhances the protective capabilities of the panel. In addition, same enhancements as for Al sandwich panel are available
• Placing an intermediate layer (using if possible Nextel or aramid) drastically reduces the shielding mass required for stopping a certain particle. Spacecraft integration could be an issue.
These solutions have also been evaluated with respect to impact on the satellite (mass areal density, volume, structure and thermal aspects, electrical design, radiation and space environment, AIT/cost, angle of debris flux, minimum areal density including reinforced box) and maturity of the concept. For instance, in the case of reinforced MLI, the application of these criteria leads to prefer the solution 1.3.
Enhanced Shielding configurations and tests
Among the 15 basic configurations, four have been selected for testing, taking into account the need for protection and the frequency of occurrence of the configurations identified for a large panel of spacecraft. They are illustrated on Figure 2-16. For each of them, two shielding concepts have been defined, taking into account the results of the preliminary tests, and considering the coherence with the spacecraft needs.
Figure 2-16: Selected basic configurations and relevant enhanced configurations for testing
The goal of this test campaign was to determine the Ballistic Limit at 15km/s, using tests results at 7km/s and hypervelocity theory equations. 31 tests on 10 configurations have been carried out. In general the enhanced configurations outperform the basic configurations.
For configurations B4 and B7, testing of the enhanced configurations has shown that they require less mass for the same performance than the basic configurations. Some of the enhanced configurations require more spacing than the basic configuration. Two or three tests have been performed for each enhanced configuration in order to derive their ballistic limit curves.
Accommodation of these shielding configurations on the satellite have been evaluated, in terms of areal density (with box) compared to shielding performances, thermal performances (interference with thermal design), mechanical performance and volume requirement.
It results that:
• Stand-offs have the biggest potential, no matter which layer is regarded
• Heavy stainless mesh is valuable to reinforce the first layer (stand-off); More detailed mechanical analyses are needed and a detailed assessment of thermal aspects should be performed.
• Aramid and Nextel show a good mass-to-shielding ratio as Intermediate layers. Applicability in clean room should be investigated.
• Intermediate skins and Tungsten reinforced face sheets are promising for panels. The reference for fuel tank showed good performances. Improvements should however been investigated considering te criticality.
Several solutions of architectural and shielding improvements have been defined for the two reference satellites. In the case of the SAR satellite, it appears that the risk of failure is reduced by 75 to 82% with an additional mass of 2,5 to 3,5% of the dry mass. In the case of the optical satellite, the risk of failure is reduced by 40 to 42% with an additional mass of 0,8% of the dry mass.
ASSESSMENT OF SELECTED SOLUTIONS
Evaluation of solutions
The selected solutions have been assessed and compared with respect to a set of criteria including their performances, their implementation, their impacts on spacecraft design, equipment design, environment, operations, system reliability, AIT, cost, and their technical maturity.
The performance of the solution characterises its efficiency with respect to debris impact. It has been evaluated for the satellite architecture solutions with a generic vulnerability tool. This evaluation has been carried out on the reference optical satellite. The figure 2-17 illustrates the reduction of the reference satellite failure probability when applying the equipment repositioning solutions and the satellite re-orientation solution.
Figure 2-17: Efficiency of selected satellite re-orientation and equipment repositioning solutions in the case of the reference optical satellite. These figures assume a re-orientation all along the orbit.
The efficiency of the solution depends on the equipment on which the solution is applied. For illustration the Figure 2-18 shows the relative efficiency of different solutions applicable to a critical equipment of the optical satellite. Only the enhanced external protection improves the PNP of several equipment items.
In the case of translation of the equipment (to move it away from the front side), the efficiency increases with the distance, but moving the equipment by a few centimeters gives already about half of the expected efficiency.
Figure 2-18: Comparison of the solutions with respect to their efficiencies: case of the CCU of the reference optical satellite
The solutions have also been compared with respect to the other criteria, and in particular with respect to their impacts on the satellite configuration. The Figure 2-19 illustrates the relative position of the selected solutions with respect to both the impacts of the solutions on the satellite layout and the expected efficiency, when the solution is taken into account in early phases of the development. To carry out this relative comparison, the solutions have been applied to the same equipment (re-orientation and additional margins are exceptions).
It appears that that in general the higher the efficiency of the solution, the higher its impacts on the satellite configuration and on cost aspects. Thus the physical segregation of redundancy is an efficient solution, but is also the most complex one in terms of implementation as it requires additional volume, harness and connectors, and generates additional mass and power need. On another hand, hiding an equipment behind another one, or relocating a critical equipment leads to a lower efficiency, but has low impact on the design, even no impact if it is taken into account since the Phase 0/A.
It is then recommended to take into account the protection against debris since the early phases of development to minimize or even avoid impacts on the satellite.
Figure 2-19: Efficiency and complexity of implementation of architecture level solutions (case of implementation
Other axes of comparison of these solutions are their availability and their interest.
The availability of these solutions is linked to their technical maturity and the need for additional development. As illustrated on the Figure 2-20, a major part of the solutions are available at short term and could be taken into account in a Phase 0/A of a project starting now.
Some solutions require a specific development. The equipment compartment assumes a modification of the equipment itself. The shielding of an equipment item has to be designed and tested, even if one of the proposed shielding configurations is used. Likewise the re-enforcement of the satellite wall has also to be tested, even if one of the proposed shielding blocks is used. Finally, the distributed architecture for the satellite has also to be developed and tested.
The fractionated satellite and the distributed architecture among various modules requires technologies development, such as the interlink communication, the control of the cluster, etc.
Figure 2-20: Availability of the selected solutions
Some of the solutions do not only protect the satellite elements against impacts with small debris, but could also improve the reliability of the satellite (e.g. external redundancy, distributed architecture), or improve its performance (additional margins, in orbit spare) as illustrated in Figure 2-21. In particular, solutions improving the reliability of the satellite would be seen positively by the space insurance.
Figure 2-21: Added value of selected solutions with respect to their impact on the satellite
In order to compare the solutions with respect to the various criteria already mentioned, a strength/attractiveness approach has been considered where the strength describes the capabilities of the solution and the attractiveness represents the interest of the solution in terms of easiness of implementation.
The strength includes the technical maturity, the potential added value, the efficiency and the capability to reuse equipment.
The attractiveness includes the capability to minimize the impacts on the satellite layout, the additional harness, the impacts on operations, the additional cost and the impacts on AIT.
The Figure 2-22 gives the evaluation of the selected solutions following this approach.
This approach allows also to define complementary actions to improve the attractiveness or the strength. To increase attractiveness, one could work on the industrialization of some options or reduce the cost. To increase the strength, one could work on the efficiency of a solution, or its road map to raise its TRL.
Figure 2-22: Evaluation of the strength and attractiveness of the satellite architecture solutions
Applicability
The applicability of the selected solutions to the different LEO missions and different types of LEO satellites has been evaluated and the applicability matrix is shown on Figure 2-23.
Each solution is only applicable to some types of mission and satellites, or to some of the equipment. For instance, solutions with additional mass and volume are applicable to large satellites, but not to small and compact ones.
Combination of solutions
This evaluation shows that a single solution does not bring the required significant reduction of satellite probability of failure, or cannot be applied to all critical equipment; in addition, the implementation of a single solution could be limited by its impacts on the satellite. Thus, combining several solutions seems an attractive approach.
The combination of solutions depends on the type and size of the satellite, of its orbit and of the mission needs. Indeed, each solution has its domain of applicability in terms of mission, and cannot be used for all the satellites. The Figure 2-24 illustrates the approach to be followed for defining the combination.
Figure 2-23: Applicability of solutions to satellites
Figure 2-24: Possible combinations of solutions in early phases of the project
PROPOSED GUIDELINES
Subsequently to the definition and evaluation of the solutions, a set of 67 design rules to assist in the choice and implementation of impact protection solutions has been derived. The design rules have been compiled according to the following categories:
• Impact risk assessment procedure
• Levels of impact protection in a space system
• Criteria for evaluating feasibility of protection solutions
• Impact protection limits
• Impact testing
• Shielding materials
• Shielding design
• Spacecraft-level solutions
• Spacecraft subsystems architecture
• Relationship to the phases of a spacecraft programme
• System-level solutions
The rules are considered to be sufficiently generalised that they can be followed during the design of any unmanned LEO space system. Therefore, they will be presented to international organisations involved in the development of guidelines and standards (e.g. IADC and ISO) for consideration and possible adoption.
RECOMMENDATIONS
There is no generic solution for the protection of LEO satellites against small debris; a case by case analysis has to be done for each mission and satellite, as the particular solution is strongly dependant on the satellites configuration (geometry, layout size) and mission characteristics like the orbital parameters.
A palette of solutions is proposed, with a range of applications, advantages and drawbacks.
System level solutions shall be decided early at mission and system level, as solutions such as fractionated architecture will impact the system definition while architecture level solutions shall be considered during the development of the satellite, as it impacts the satellite definition.
It is worth taking into account the needs and solutions for protection of the satellites against debris since the early phases (0, A) of the project. High level numerical simulations could be carried out to evaluate the vulnerability of the preliminary configuration of the satellite with respect to debris and trade–off various solutions of protection.
The simulations allow to identify the most critical equipment of a satellite configuration. Working on the protection of these equipment items will already highly improve the overall survivability.
Moving equipment to a safer place is the simplest solution; solutions to reduce the area of critical surface, such as rotation of equipment or of the satellite shall also be considered.
The use of shielding shall preferably be restricted to the protection against debris of 3 to 4 mm size. It allows to reduce significantly the risk of failure.
The selected solution shall be compatible with the Design to Demise, in order to allow an uncontrolled re-entry after the end of mission of the satellite.
Additional experimental tests should be carried out to supply the Ballistic Limit Equation dedicated to the desired configuration if it is not available, but also to consolidate the extrapolation of the BLE above 7 km/s. To that aim, experimental tests at 10 km/s or more should be necessary.
Innovative shielding concepts using new materials have been defined and tested.
Design rules have been elaborated to increase the robustness of European satellites in the growing population of small debris
Potential Impact:
Potential impacts and beneficiaries
The ReVuS project has allowed to propose design solutions to reduce the vulnerability of the future LEO satellites, whatever their size, mission and orbit, against impacts of small-sized debris, that means debris which cannot be detected and tracked. Through its results, it significantly contributes to the European capacity to protect space assets from space debris.
Potential impacts
The main benefits of the ReVuS project are:
• The evaluation of the respective contribution of each range of debris size (between 0.1 mm to 5 cm) to the risk of failure of a satellite, and the highlighting of the 3 to 5 mm range as the main contributor. This is a major impact for the selection of a solution to protect the satellite and in particular for the design of shielding configuration
• The assessment of a set of promising shield building blocks, with their tested performances. As a result, the project indicates what type of building block is more adapted to the different layers, and gives their performance in terms of ballistic limit versus the overall areal mass, and reports whether it is used in a shielding configuration or as re-enforcement of existing MLI or sandwich panel
• A set of tested enhanced shielding configurations for several basic equipment configurations. The best enhanced shielding solution requires only 1/4 of the mass when compared to a corresponding state-of-the-art shielding solution. The project gives quantitative evaluation of these enhanced shielding configurations in terms of ballistic limit curve, expected reduction of satellite vulnerability, and impacts on the satellite (mass, environment, etc).
• A palette of satellite architecture solutions, with their advantages and drawbacks, to be used on existing (in development) or new design of LEO satellites. This palette of solutions includes also an applicability matrix, as function of orbital parameters and type (size) of satellite, as there is no solution applicable to all satellites.
• Design rules for taking into account protection against small debris in the design of future LEO satellites
• A proposed approach to evaluate the vulnerability of satellite since the early phases of the design and to support the definition and selection of adequate solution, in particular the combination of solutions.
• A new collision risk algorithm, implemented within a space debris evolutionary model that allows for the assessment of small debris impact risks and potential failures of active spacecraft.
Beneficiaries
The definition of possible solutions and design rules to reduce vulnerability of satellites against small debris is a main benefit for the next generation of LEO satellites, whatever the user is. It will contribute to a more efficient design, a higher quality and enhance the competitiveness of European space systems.
These results of the ReVuS project will benefit to the space community, institutional, industrial as well as scientific.
Industrials, satellite and space system manufacturers will benefit from approach to evaluate the vulnerability of future satellite to small debris, from a palette of solutions and design rules to reduce this vulnerability, and from a methodology to define the adequate solution. The existing set of shield building blocks and shielding configuration with their performances will provide them either with an existing adequate solution, or with guides to define the adequate solution.
Research institutes and universities have used their scientific findings and models to improve the design of space system. They have also improved their models of environment for the evaluation of impacts of small debris on a satellite and have developed architectures to reduce the vulnerability of satellites. The advancements generated during the project will be used as a baseline for future research. Research institutes and university will benefit from a reduced amount of work required to attain a certain research level, facilitating further advancements in research.
The users of space system (mainly operator) could now have a clearer picture of the risk of degrading or losing the mission following an impact with small debris over the expected lifetime, of the potential solutions to reduce this risk and of their relevant impacts on the satellite, and thus has a support enabling adhoc decision with respect to problem of small debris.
The insurers will have also a clearer view on the risk of degrading or losing the satellite, of the interest, risk and possibly positive effect on satellite reliability of the possible solutions.
The institutional players will have a starting point directly applicable to on-going and future projects. In particularly, the proposed ReVuS solutions could be used to reduce the risk of losing the post mission disposal capability of the future satellites, which is a key contributor to the stabilization of the population of debris.
In addition to the space community, ReVuS will also benefit to European citizen. Indeed, Europe is using Low Earth Orbit satellites to provide services like earth observation, resources monitoring, defence and security systems that at the end benefit to citizens. Reducing the satellite vulnerability to small debris allows to improve the maintenance of the continuity of these services to the citizens.
On societal point of view, ReVuS has allowed to gather in the same study research laboratories, universities, manufacturers, industrials and users (insurer) for the complex problem of small debris remediation. In particular, the team was not restricted to the space community, but involves also an insurer, who gives a complementary angle of view of the debris problem, and a manufacturer of shielding devices for civil and military applications.
Dissemination of results
The results of ReVuS have been continuously disseminated all along the project. This dissemination has been carried out through the ReVuS site, conferences, publications and a dedicated workshop last October, at the end of the project.
Lecturers have been done at adequate conferences and symposium all along the 34 months of the project, in order to disseminate interesting results as function of the progress of the project, and to make the project known by the space community, and beyond.
To that aim, four types of presentation have been targeted, according to the content and audience of the conferences: a general presentation, describing the overall ReVuS objectives and results at successive steps of progress, presentations focused on technical subjects to show in more details some main results, and short (overview) presentation of the project to space and non-space audience.
• General presentation of the project, describing the overall ReVuS objectives and results at successive steps of progress:
- Conferences covering all space subjects: IAC 2012 (October 2012, at mid project) - 2nd FP7 Space Conference, November 2012 - IAC 2013 (October 2013, close to the end)
- Conferences on space security (IAASS, October 2011 and May 2013), on sustainability of space activities (ISU, February 2012), on space debris (European conference on space debris, April 2013)
- Forums on space risk (World Space Risk Forum, June 2012)
• Presentations focused on technical subjects to show in more details some main results:
- Distributed architecture (NASA/ESA Conference on Adaptive Hardware and Systems, June 2012 - Industrial Showcase event of Department of Engineering, University of Leicester, February 2013 – Lecture on Microelectronics and Embedded Systems for Space Applications, June 2012 - IEEE Aerospace Conference, March 2014)
- Shielding material and tests (12th International Symposium on Materials in the Space Environment, September 2012 - 6th European conference on space debris, April 2013 - IAC 2013, October 2013 – AAAF conference on dynamic behavior of materials and structures, January 2014)
- Vulnerability analysis (IAC 2012 ,October 2012 - 6th European conference on space debris, April 2013)
• Short general presentations (overview) of the project to space and non-space audience:
- Space Situational Awareness (SSA) Seminar, organized within Polish Presidency of the Council of EU, September 2011
- NordicBalSat Conference Broadening the Base of Europe’s Space Community, February 2012
- EISC Seminar, May 2012
- Science Pique-Nique, May 2012
- ILA, Berlin Air Show, September 2012
- WIRE (Week of Innovative Regions in Europe), June 2013
- SSA conference, November 2013
• Participation to specific working groups on space debris in order to present the ReVuS results, or to discuss the utilization of some of the project results:
- Overview of ReVuS results and progress at the CNES synthesis group of debris (CNES June 2011, June 2012, June 2013)
- Discussion (PHS, EMI) on the results of the first phase of the vulnerability analysis and of the shielding aspects at a joint session of the IADC WG3 (Protection Working Group) and WG4 (Mitigation Working Group) meetings in Montreal, Canada during 22 – 25 May 2012.
- Presentation (PHS) of the preliminary results of the design rules activity at the ISO TC20/SC14/WG7 (Orbital Debris Working Group) meeting in Guildford, UK on 7/8 November 2013.
- Presentation of the final results of the design rules activity at the IADC WG3 (Protection Working Group) meeting in Beijing, China on 12-15 May 2014 and at the ISO TC20/SC14/WG7 (Orbital Debris Working Group) meeting in Tokyo, Japan on 26-28 May 2014.
- Discussion (EMI) on the results of the shielding aspects at the 31st IADC meeting, 17-19 April 2013 in Darmstadt, Germany
• Exhibition of space debris modelling to non-space audience:
- Exhibit run by Southampton at the University of Southampton’s award winning science and engineering day, 2012 and 2013, with visitor numbers in excess of 2000.
In addition to the conferences, a ReVuS workshop has been organized last October, close to the end of the project, over a full day. The objective was to present in detail all the results of ReVuS to institutions (agencies), academia and industry. This workshop gathered up to 40 people and has been a main contributor to the dissemination of the ReVuS project.
Figure 3-1: Dissemination of ReVuS results
Exploitation of results
Most of the ReVuS results are public and have been disseminated all along the project, as discussed above. There is no patent and thus no restricted exploitation of the data.
ReVuS proposes an approach to evaluate the level of risk due to small debris (and thus whether there is a need for protection or not) and to identify the solutions for reducing the vulnerability that could be applied to a satellite in development. This approach could be used by the manufacturers or even by the operators of future LEO satellites.
The vulnerability analysis, aiming at evaluating the probability of failure of the satellite configuration, and at identifying the critical elements, has been described in the project and can be reused by manufacturer or insurer. There are existing tools to carry out such an analysis.
Although this analysis has been done on reference satellites, some results can be directly exploited:
• the debris with size in the range 3 to 5 mm appear as the main contributor to the satellite probability of failure, for the two reference satellites which are different both on orbit and configuration. In both cases, it has been shown that protecting the satellite against this range decrease significantly the probability of failure. This size of debris could be taken into account in the preliminary development phase of a satellite, in particular for the sizing of shielding protection.
• The level and characteristics (direction, repartition as function of true latitude) of the flux of particles has been given for different altitudes. These figures can be used for preliminary trade-offs (short evaluation of preliminary configurations of satellite, candidate orbits, etc).
Preliminary results have been presented at the IADC working groups.
ReVuS proposes a palette of solutions at architecture level to reduce the vulnerability and gives an applicability matrix. This matrix is a key as each solution is only applicable to a certain range of satellite. Thus, depending on the orbit altitude and type of satellite, only a set of solutions can be considered. In addition, these solutions have been evaluated for a reference satellite, so that, even if the performance of the solution depends on the satellite configuration, the user will have in hands an applicable set of solutions with their relative performances and impacts.
Among the candidate solutions, ReVuS has identified solutions with added value: these solutions could increase the performances of the satellite or its reliability. Such solutions are of interest for different users: the manufacturer, because they participate to the performances (margin on resources) or the overall reliability (for instance external redundancy) of the satellite; for the operator, as it could expect a higher performance at the beginning of utilization of the satellite; for the insurer who could reduce the premium if the reliability is higher.
The distributed architecture solution proposes a new computing architecture inside a satellite, at the opposite of the usual centralized computer, with advantages in terms of reliability. Results of tests are given in the project. However, additional development and verification will be needed in parallel to the development of satellite.
ReVuS makes available results of the test of a large set of shielding building blocks. These building blocks are enhancements of protective elements like MLI, sandwich panels, intermediate layer. They use different materials with various thicknesses. These building blocks have been evaluated in terms of performances and impacts on the satellite. All these data could be used by the satellite manufacturer to re-enforce external sides of the satellite, or define shielding configuration. But they are also available for non-space user that could be interested by the performance of some material for the protection of ground vehicle, aircraft, people against particles of lower velocity than debris.
ReVuS has defined a dozen of equipment configurations, taking into account their relative location in the satellite and their mechanical and thermal environment. For four of them, ReVuS proposes a set of two shielding configurations, each relying on the shield building blocks. They have been designed to protect the equipment against debris of size up to 3 mm. They have been tested, and ReVuS makes available the results of their tests, their ballistic limit curve and their characteristics. The user can benefit either as an existing product (re-use the shielding configuration as is while adapting the surface to the equipment) or as a set of data to support the definition of relevant shielding configuration. These data are valuable support which gives ideas of mass, size to protect equipment against debris of 3 mm.
As a synthesis of all the achieved results, design rules to protect LEO satellites against small debris have been generated. They could be used by manufacturer of future LEO satellites. They could be of interest for international organizations involved in the development of guidelines and standards (e.g. IADC and ISO) for consideration and possible adoption. Preliminary results have been presented at the Orbital Debris Working Group of ISO in last November. The final design rules will be presented to the adequate Working Groups of IADC and ISO in May 2014. This is the first identified exploitation of the ReVuS results.
List of Websites:
htpp://www.revus-project.eu
Contact points:
Claude Cougnet: claude.cougnet@astrium.eads.net
Krystyna Macioszek: Krystyna.MACIOSZEK@astripolska.pl
The population of debris in space is continuously growing, and could become an increasing risk for the survivability of space assets. Unlike large debris, small debris cannot be tracked nor avoided; a collision with a LEO satellite could result in degradation, or even loss of the mission.
The ReVuS project aims to propose design solutions at system and satellite architecture levels to reduce the vulnerability of future LEO satellites to these small-sized debris (0.1 to 5 cm). It is supported both by analyses carried out with adequate tools and models, and by tests of shielding materials.
A vulnerability analysis has been carried out to evaluate the probability of having small debris colliding a satellite, and the effects of such an impact (probability of penetration, and risk of mission failure). The level of vulnerability of a satellite depends on its orbital position and on its configuration
This analysis was done on two reference LEO satellites (optical and SAR) characterized by a different configuration and a different altitude. Particles in the size range 1-5 mm dominate the failure risk (due to debris impact), with significant peak at around 2-3mm. The debris environment varies significantly with the orbit altitude, being much severe at 800 km than at 500 km; highest impact flux comes from the front direction, and the critical satellite units are the external ones or those mounted close to the front side.
Solutions at system level (fractionated satellite, distributed architecture, in orbit spare) cover the full range of debris. They impact the overall mission and system. They have a high efficiency as the probability of losing the mission is quite negligible, but are penalized by the cost and, for some solutions, by the technical maturity. In orbit spare solution could be attractive for mission involving several identical satellites on the same orbit.
Solutions at satellite architecture level encompass the repositioning of critical equipment units (relocation, removal from front side, rotation, hide behind other units), the re-orientation of the satellite through the polar areas, the use of external redundancy, additional margins and the shielding of critical equipment or of areas protecting several critical units.
The shielding concept depends on the location of the equipment in the spacecraft and of its local environment. While defining ways to obtain efficient shielding concepts adapted to equipment configuration, alternative shielding building blocks have been identified for reinforced MLI, reinforced structural panels, intermediate layers, and reinforced equipment box. A preliminary test campaign has allowed comparing alternative building blocks in each family with respect to their capacity of protection, resulting areal density and impact on satellite.
Based on these results, shielding solutions have been elaborated for each of four selected equipment configurations, and tested in order to determine their ballistic limit at 15 km/s. These solutions allow protecting the equipment against particles of at least 3mm (6 mm for the solutions selected for the tank). Some of the enhanced configurations appear lighter than the basic ones. The tests show that stand-offs have the biggest potential, heavy stainless steel mesh are valuable to reinforce the first layer, aramid and Nextel show a good mass-to-shielding ratio for intermediate layer.
An evaluation of all satellite architecture level solutions has been done with respect to a set of criteria, including performances, additional mass, volume, harness and connectors, maturity level, cost aspects. Solutions with the highest reduction of satellite risk of failure (such as external redundancy, shielding of equipment) are penalized by their impact on satellite layout (additional mass, volume, harness) and by cost aspects.
To minimize impacts on the satellite, it is worthwhile to take into account the protection against debris since the early phases of development, and also to rely on a combination of solutions rather than on a single one.
In addition to their protection function, some of the solutions could also improve the reliability of the satellite (e.g. external redundancy), or its performances (additional margins, in-orbit spare). In particular, solutions improving reliability of the satellite would be seen extremely positively by the space insurance.
Each solution is not applicable to all equipment or to all types of satellites. For instance, solutions with additional mass and volume are better suited to large satellites rather than small and compact ones. There is no generic solution applicable to all satellites; indeed, most of the performance evaluations are geometry dependent. The set of solutions will be specific to each satellite design, depending on application, orbital parameters, satellite configuration and needs.
ReVuS brings a palette of solutions with their potential gains and impacts. It improves the quality of the design of European satellite.
Project Context and Objectives:
Context
The population of debris in space is continuously growing, especially as a result of the collisions that have occurred in the past few years. It appears that even if guidelines to minimise the post-mission orbital life are applied by the space community, this population of debris will continue to grow in the years to come. Thus the presence of debris could become an increasing risk for the survivability of LEO space assets. Indeed, the probability for a satellite to collide with orbital debris, although very low, could become non negligible. This probability and the relevant consequences differ according to the size of the debris:
Debris larger than 10 cm can be detected, catalogued and tracked from ground, so that collisions can be predicted and avoidance manoeuvres can be performed.
But smaller debris cannot be tracked, and their presence remains a threat for space assets.
While the very small debris (< 0.1mm) have an energy low enough to be absorbed by the structure materials, the small sized debris (from 0.1 mm to 10 cm) could generate critical damages, such as the loss of a part of the mission, the degradation of the performances of the satellite, or even the loss of the mission in case of collision. Indeed, their energy is high enough to penetrate the satellite structure and they constitute a risk for internal mounted units. The environment models show that a satellite located at an altitude of 800 km will have in the coming years an average of 40 impacts/m2/year with this type of debris, among which 0,04 particle/ m2/year will penetrate inside the satellite.
In order to mitigate this risk, several approaches are currently considered, as shown in Fig 1-1:
• Prevent: the aim is for new vehicles (spacecraft, upper stages) to avoid staying in operational orbits too long after the end of their lifetimes, thus increasing the population of debris. It relies on the use of guidelines, laws, rules and standards to limit the post-mission presence of vehicles in operational orbits
• Avoid: it consists of the satellite performing manoeuvres to avoid a collision with debris. It is currently used by the main satellite operators. It concerns large size debris which can be detected and tracked from the ground
• Remove: it deals with debris removal. Studies are currently done on active debris removal. A first objective is to remove large objects (dead satellites and upper stages).
• Survive: It is related to the small size debris that cannot be tracked. It consists of defining design rules to minimise the effects of debris impacts on the satellite and its mission.
Fig 1-1: Approach for debris risk mitigation
Objectives
The ReVuS project is an answer to the Survive approach.
It aims to propose design solutions at system and satellite architecture levels to reduce the vulnerability of future LEO satellites to these small-sized debris (0.1 to 5 cm). It is supported both by analyses carried out with adequate tools and models, and by tests of shielding materials.
In order to achieve this aim, the ReVuS project had to meet several objectives that drive the logic of the project, illustrated on Figure 1-2.
Fig 1-2: Logic of the ReVuS project
The first objective was to evaluate the probability of having small debris colliding a satellite, and the effects of such an impact (probability of penetration, and risk of mission failure). It was the object of the vulnerability analysis. In order to have quantitative figures, two reference LEO satellites (optical and SAR) characterized by a different configuration and a different altitude (825 km for optical one, 515 km for the SAR one) have been selected. This analysis was based on the use of the impact risk assessment SHIELD3 tool which evaluated the probabilities of penetration of small debris particles in the satellite and in the equipment and determined the failure probability for all equipment parts considered.
The SHIELD3 tool relies on information about the design of the satellite and on the directional distribution of the flux of debris on the satellites computed using the MASTER 2009 environment model.
This analysis provided information on the effect of the orbit altitude, on the flux of particles impacting and penetrating the satellite for different altitudes, on the probability of satellite failure in the two reference cases, on the range of debris that is contribute the most to the risk of failure, on the satellite units that appear critical with respect to debris impact. All these results are needed for the next objectives.
The second objective was to define and analyze the possible solutions at system level. Several solutions have been identified, and four of them have been analyzed and kept. Such solutions have a direct impact not only on the satellite itself, but also on the mission and on the system definition. Evaluation of their efficiency to reduce vulnerability of the satellite, of their maturity level and of their impacts on the mass and cost aspects has been done.
The third objective was to define and analyze the possible solutions at spacecraft architecture level. These solutions encompass the repositioning of critical equipment units (it includes mainly the relocation of equipment in a safer zone, the increase of distance between the critical equipment and the front side, the rotation of equipment, to hide critical equipment behind other units), the re-orientation of the satellite when flying through the polar areas, the use of external redundancy, the additional margins and the shielding of critical equipment or of areas protecting several critical units. These solutions have been evaluated with respect to their accommodation on the spacecraft and the impacts on the spacecraft configuration in terms of layout, mass, environment, thermal aspects, etc.
In order to have qualitative figures, several improved configurations have been defined for the two reference satellites, combining relocation of some critical elements and shielding solutions. The objective was to evaluate the achieved reduction of probability of failure with SHIELD 3 tool and the resulting additional mass.
The fourth objective was to define and test possible shielding material and shielding configurations.
The shielding concept (including two layers or more, various materials, distance between layers) depends on the location of the equipment in the spacecraft and of its local environment (structure, thermal).
Alternative shielding building blocks have first been identified for reinforced MLI, reinforced structural panels, intermediate layers, and reinforced equipment box. The aim was to define solutions with different materials, different thicknesses and different arrangement of the material layers, and to evaluate the performances of these alternative solutions in order to compare them in term of performances (efficiency of protection) and mass areal density) through a preliminary test campaign.
Based on these results, shielding solutions, sized to protect the equipment against particles of at least 3mm, have been elaborated for four selected equipment configurations, and tested in order to determine their ballistic limit at 15 km/s. These solutions have been evaluated with respect to their performances, impacts on satellite configuration (mass, volume, structural constraints), on environment and on cost.
The fifth objective was to evaluate the benefits of all the proposed solutions, their resiliency to debris impact and the interest of combining them.
All these potential solutions have been assessed with respect to a set of criteria including their performances (in terms of reduction of probability of failure of the satellite), additional mass, volume, harness and connectors, maturity level, cost aspects, added value. In particular, the interest of combining several solutions has also been assessed. One objection of this evaluation was also to identify the applicability of each solution to the various types of missions and satellites in order to know what set of solutions could be used for each category of satellite.
The sixth objective was to propose new design rules to increase the robustness of European LEO satellites in case of impact with small debris.
A set of such design rules has been elaborated to assist the satellite manufacturer with the choice and implementation of protection solutions; they will be presented to international organizations involved in the development of guidelines and standards.
The last objective was to disseminate the results of the project. Indeed, a main purpose of the ReVuS project was also to ensure that the results of the work are known in the communities outside the ReVuS consortium. That has been achieved through presentations to various conferences, publication of papers and more particularly a workshop open to public.
Project Results:
The ReVuS project targets the small-sized debris, from 0.1mm to 5 cm. It aims to define design solutions in order to reduce the vulnerability of future LEO satellites to this range of debris.
The project followed a three-step approach, as illustrated in Figure 2-1:
• the vulnerability analysis, to evaluate the effects of a collision of a LEO satellite with small-sized debris, the critical parts of the satellite, and the risk of mission degradation
• the identification and analysis of potential solutions at system level, and at satellite architecture level, with a focus on the shielding concepts and shielding materials
• the resiliency analysis, aiming at evaluating the resiliency of the selected solutions with respect to debris impact and at proposing design rules and standards.
Figure 2-1: Logic of the ReVuS project
VULNERABILITY ANALYSIS
The vulnerability analysis has been carried out to evaluate the probabilities of having small debris impacting the satellite and to determine the effects of such an impact.
As shown on the Figure 2-2, this analysis is based on the use of the impact risk assessment tool SHIELD3 which evaluates the probabilities of penetration of small debris particles in the satellite and in the equipment. In order to have quantitative figures, two reference satellites, an Earth Observation optical satellite and an Earth Observation radar satellite, have been taken into account. They have different configurations and orbit altitudes: the optical satellite has a deployed solar array and is located at 820 km altitude, 98,7 deg inclination while the radar satellite has a rigid body mounted solar array and is located at 515 km altitude, 97,4 deg inclination.
Figure 2-2: Approach for vulnerability analysis
The SHIELD tool needs information about the design of the satellite (geometric data file, material description, information on redundancies, updated damage equations) in order to draw correct conclusions on spacecraft vulnerability while having a significantly reduced complexity of the model.
In parallel to the reference satellite models, the directional distribution of the flux of debris on the satellites has been computed using the MASTER 2009 environment model for the “Business-as-usual” (BAU) scenario, which assumes that current practices are to continue in the future.
A period of 10 years following January 1st, 2020 has been considered for the analysis.
The debris environment varies significantly with the orbit altitude, being much severe at 800 km than at 500 km. As shown on Figure 2-3, the highest impact flux comes from the “front right” and “front left” directions in the case of optical satellite. The azimuth distribution is flatter in the case of SAR satellite. Subsequently, the critical satellite units are the external ones or those mounted close to the front side.
An evaluation of the flux of debris impacting, and of the flux of debris penetrating the reference satellites on their different faces has been done for the successive ranges of particles diameter. The results are summarized in Figure 2-4. It appears that:
• On average, debris flux level for SAR satellite is three times below that of the optical satellite.
• Debris below 1mm diameter has a high probability of impact, but only few penetrate the satellite, so that the effects on equipment are very low.
• Debris particles with a diameter in the range (1-10mm) impact and penetrate the satellites 100 times more often than debris particles in the range (10-50mm).
• The risk of being impacted or penetrated by small debris is much higher at 800 km than at 500 km
Figure 2-3: Illustration of the flux distribution for the two reference satellites
Figure 2-4: Flux of particles impact and penetrating the reference satellites
The SHIELD tool has evaluated the probability of penetration of the resulting debris particles in the satellite equipment. This evaluation is illustrated on Figure 2-5. Then, taking into account the redundancy scheme, in particular whether the equipment has an internal or an external redundancy, an evaluation of the probability of no failure of the satellite has been derived. Figure 2-6 illustrates the probability of failure of the reference satellites as a function of the diameter of debris particles. This is only applicable to the two reference satellites as the result depends highly on the size of satellite, its orbit, its layout (deployed or rigid body mounted solar arrays, etc). These figures cannot be taken as general/average numbers. In addition, they depend on the criteria of failure that have been taken into account.
As a main result, the vulnerability analysis highlights that particles in the size range 1-5 mm dominate the satellite failure risk (due to debris impact), with significant peak at around 2-3mm. The impacts of particles with larger diameter, above 1 cm, will result obviously in more dramatic consequences (higher risk of losing the mission), but the probability of such an impact is much lower, as illustrated in Figure 2-4.
However, the penetration of a debris particle in the satellite, and in equipment does not mean the loss of the satellite. Indeed, the potential damages that could result from such penetration could be:
• the degradation of performances of the satellites, resulting from the loss of resources like battery, the degradation of solar cells, radiators, tank leakage, etc;
• the reduction of the satellite reliability (loss of redundant equipment).
• The degradation of the mission (loss of instruments or payload electronic units)
• the loss of the mission, that could result from penetration of debris in the tank (with a risk of explosion or only leakage depending on particle size and impact conditions), or in a non-externally redundant equipment (in the case of internally redundant equipment, the level of failure depends on the internal architecture).
Exposed functional surfaces, which are not protected, such as solar arrays, are in general designed to tolerate a debris impact flux. Such surfaces have not been considered in the evaluation of probability of no failure.
The radar satellite presents a very low vulnerability to small debris, with a high Probability of Non Failure (PNF) due to its cylindrical shape, with axis along velocity vector and rigid body mounted solar arrays, and its low altitude (515 km) outside the zones where the density of debris particles is high.
Figure 2-5: Illustration of the flux of particles penetrating the SAR satellite and probability of penetration in the equipment.
Figure 2-6: Illustration of Probability of failure as a function of debris size for the two reference satellites.
PROPOSED SOLUTIONS
Based on the results of the vulnerability analysis, two main categories of solutions (Figure 2-7) have been defined: solutions at system level, and solutions at satellite architecture level, which includes the shielding solutions.
A list of proposed solutions has been established for each category and is given in Figure 2-8.
Figure 2-7: Categories of solutions.
Figure 2-8: List of proposed solutions.
SYSTEM LEVEL SOLUTIONS
The system level solutions aim at mitigating the risk at system level. Such solutions can take into account the full range of debris size. They have a high efficiency as the probability of losing the mission is quite negligible, but are penalized by the cost and, for some solutions, by the technical maturity.
A candidate solution is the fractionated satellite concept (see Figure 2-9), which consists in sharing some functions of a satellite (computing capability, communications with ground, payloads, etc) on separate modules forming a cluster, based on wireless communications and interconnecting network. It improves the maintainability and the responsiveness to unexpected events. With an adequate distance between the modules, a collision with debris could lead to lose a module, but not the complete mission. In the case of the optical satellite, the probability of losing the mission is 50 times lower that the one of a monolithic satellite, but the probability of losing at least one module is 3 times higher than the probability of losing the monolithic satellite.
Another solution is the distributed system concept (see Figure 2-9), which will adapt the principles of existing terrestrial wireless systems to distributed space system architectures. It consists in a novel generic architecture for fault tolerant distributed on-board computing. The distributed computing system is comprised of multiple nodes. This approach is based on task migration. The computing system can be distributed within the spacecraft, or over several separated modules, linked by wireless communications. This latter case is the system level solution.
A third possible solution is the in-orbit spare. This solution could be attractive for mission involving several identical satellites on the same orbit.
Figure 2-9: Illustration of system level solutions
SPACECRAFT LEVEL SOLUTIONS
At spacecraft architecture level, various types of solutions have been considered and evaluated. They are described in Figure 2-8.
Each solution has been assessed with respect to common limitations and design principles generally used, taking into account implications on design, cost, performances, etc (qualitative aspects).
The solutions implementing shielding are discussed in the next section.
It appeared that most interesting solutions are:
Re-orientation
This solution consists in adapting the orientation of the satellite to expose less critical side during fly-through of critical zones (mainly the poles area). The attitude of the satellite could be slightly modified around yaw or pitch axis, thus exposing an angled surface to the major flux vector.
Equipment compartmentalisation
This solution consists in implementing a protection (layer) inside the equipment that has internal redundancy in order to protect the redundant part (or the nominal part) from the penetration of a debris inside the equipment. As a result, the equipment is modified, with additional layer and boards. It is almost similar to the juxtaposition of two identical equipment units not internally redundant.
Positioning into a less critical area
This solution consists in positioning critical units (those that have a low PNP) in less exposed area, typically in the rear side. Such solution has impacts on the satellite layout (additional free volume) and AIT, unless it is taken into account in the early phase of the design.
Hide behind less fragile equipment
This solution consists in placing critical units (elements with low PNP) behind less fragile (or massive) elements (such as a passive redundant equipment) that will provide a protection. This solution has impacts on satellite configuration and AIT if not taken into account in the early design phases. It could be interesting for harness or external equipment.
Increase distance between equipment and wall
This solution consists in increasing the distance along the debris flux vector between the equipment and the wall (located in front of equipment with respect to main debris flux) in order to distribute fragments on a wider surface (thus reducing surface energy). It can be done without significant impact except on the layout due to the need for sufficient volume for equipment mounted on panels that are perpendicular to the flux vector. For the other equipment, the constraints generated on the mass, structure and thermal dissipation, plus the volume are probably too penalising with respect to the gain.
Rotation of the equipment
This solution consists in slightly rotating the equipment in order not to have a face perpendicular to the flux of debris and thus to reduce the effects of impacts of debris on the equipment box.
Physical segregation of redundant units
This solution consists in segregating physically the redundant units (implemented on two different areas of the satellite) rather than having both units inside the equipment. This solution will need additional volume, harness, switching unit (that has a risk of failure), but the common causes of failure are avoided.
In case of new satellite, the various impacts are reduced as there is no existing layout.
Addition of external redundancy
This solution consists in adding spare (redundant) units in order to increase the PNF of the function when the operational equipment items are critical with respect to debris (low PNP). For instance, to add a battery, or an external sensor.
Distributed architecture
This solution consists in implementing a distributed architecture for the data management system (computers and RTU): functions are shared over several computers, each being able to take over functions of other ones. It is similar to what has been quoted in the system level solutions, but applied to the satellite architecture (with wired links). A reliability and availability analysis shows that the distributed architecture (distributed OBC) has more reliability and availability as compared to centralised system. It can provide more computing performance because of inherent availability of multiple processing units. A prototype of a fault tolerant distributed OBC (Fig 2-10) with three nodes has been tested: the reconfiguration time, for the TDMA communication scheme, is approximately the same regardless of the number of active modes; each task resumes its data state from the current data state rather than initialising as in legacy systems; high data rate protocols are needed for transfer of large task data states; power consumption is higher than in centralised system.
This solution leads to new avionics architecture, and a new satellite layout. As compared to a centralised system (with an internally redundant computer), there will be additional mass, harness, power consumption, AIT activities (especially for tests) and probably higher cost. When starting from the early design phase, the resulting impacts will be low.
Additional margins
This solution consists in considering additional margins in the sizing of resources to take into account the possible impacts of small debris and thus to avoid a degradation of resources below the level required by the mission. Indeed, for resources like solar arrays or radiators, it is not possible to define a protection. The architecture of the solar array shall be such that the loss of power due to a debris impact (hole) on a cell is minimised. In the case of radiators, oversizing shall be avoided not to modify the thermal balance.
SHIELDING CONCEPT
The shielding solutions will have a significant impact on the mass and on the layout of the satellite. Thus, they cannot be sized to protect the satellite or the equipment units against the full range of debris size.
The vulnerability analysis has shown that a ballistic limit of 3 to 4 mm at 15 km/s is necessary to reduce significantly (up to 50%) the probability of failure (due to debris) of the satellite. As illustrated in Figure 2-11, this ballistic limit objective is confirmed for most of the equipment; for the propellant tank, there is a need for a ballistic limit of 5 to 6 mm to reduce significantly the probability of penetration.
Figure 2-11: Illustration of ballistic limit objective for two equipment of the SAR satellite
The assessment of the needed shielding configuration depends on the location of the equipment in the satellite and its local environment (structure, thermal hardware, etc.).
The analysis of the reference satellites, and also of the current and future LEO satellites, has led to identify fifteen basic configurations of equipment.
The shielding performances of each basic configuration have been estimated by computing their ballistic limit using published data. This estimation confirms the need to protect the critical equipment in order to achieve the required objective.
Shielding principles
A typical efficient shielding has a multi-layer configuration characterised by the number of layers, the thickness and material of each layer, the distance between layers and the orientation of the layer with respect to impacting particle. Each layer has a specific function: the bumper layer (the first one) breaks up and melts the projectile, the inner layer traps the secondary debris and the back wall (usually the box wall) provides the last line of defence. A monolithic shield appears inefficient in terms of mass versus protection.
To obtain an efficient shielding adapted as far as possible to a basic configuration, different possibilities can be considered:
• Re-use the existing structural or thermal elements (typically sandwich panels, multi-layer insulation, equipment box wall), improve their efficiency by increasing thickness or changing the distance between the items
• Change the materials
• Addition of intermediate layers
Shield building blocks and preliminary tests
Based on these principles, the following families of shield building blocks have been defined:
• Reinforced MLI: there are 8 shield building blocks within this family
• Reinforced sandwich panel (with Al skin or CFRP skin): there are 13 building blocks in this family
• Intermediate layer: there are 5 building blocks in this family
These shield building blocks have been investigated through the preliminary test campaign, with the objective to evaluate the characteristics of the shield building blocks and make a ranking between the solutions within each family.
Figure 2-12: Shield building blocks test conditions
The tests have been performed at Fraunhofer EMI’s two-stage light-gas guns at 7 km/s. All the samples required for the tests have been manufactured by Tencate Advanced Composites.
The test conditions are illustrated in the Figure 2-12. The shielding components are placed within a set-up that is representative for their occurrence within a spacecraft. The targets are impacted with nominally identical impact conditions above their ballistic limit. Witness plates are placed behind each target. The first witness plate behind the target (WP1) is considered somewhat representative of module walls.
Examples of the shield building blocks state after the tests are shown in the Figure 2-13 for the Al sandwich panel family. In order to compare the performance of the shielding bricks, the penetration capability of the most damaging fragment impacting the witness plate simulating the module wall (WP1) is estimated. This penetration capability is given in terms of the penetrated areal density of the shield. This number includes the (nominal) areal density of all layers that would have been necessary to stop the impacting particle.
This penetration capability is a measure of the quality of the investigated sample. It describes both the sample’s ability to disperse the fragment cloud over a larger area, and (especially for intermediate layer samples) to decrease a fragment cloud’s energy.
Using this number, the different shield building blocks can be compared against each other.
The figures 2-14 and 2-15 give the results of the preliminary tests for the reinforced MLI and for the Al sandwich panel.
Figure 2-13: Examples of Al sandwich panel shield building blocks status after tests
Figure 2-14: Penetrated areal density plotted vs. sample areal density for the MLI targets. Filled symbols indicate WP1 perforation. Solid line is identity.
Figure 2-15: Penetrated areal density plotted vs. sample areal density for the Al sandwich panel targets. Filled symbols indicate WP1 perforation. Solid line is identity.
From these preliminary tests, it appears that:
• The performances of MLI with respect to shielding can be increased by adding a stainless steel mesh (1.4 1.3 on Fig 2-14), or by adding a reinforced MLI layer with stand-off distance (1.5). These solutions have a mass impact.
• The performances of Al sandwich panel could be improved by substituting the honeycomb with either foam (2.6) or corrugated plate (2.7) that allows a greater dispersion of fragments on subsequent layers, or by introducing an additional Al layer in the middle of the core (2.3) parallel to the face sheets, causing additional shocks in fragments.
• For CFRP sandwich panel, embedding tungsten particles into the face sheets enhances the protective capabilities of the panel. In addition, same enhancements as for Al sandwich panel are available
• Placing an intermediate layer (using if possible Nextel or aramid) drastically reduces the shielding mass required for stopping a certain particle. Spacecraft integration could be an issue.
These solutions have also been evaluated with respect to impact on the satellite (mass areal density, volume, structure and thermal aspects, electrical design, radiation and space environment, AIT/cost, angle of debris flux, minimum areal density including reinforced box) and maturity of the concept. For instance, in the case of reinforced MLI, the application of these criteria leads to prefer the solution 1.3.
Enhanced Shielding configurations and tests
Among the 15 basic configurations, four have been selected for testing, taking into account the need for protection and the frequency of occurrence of the configurations identified for a large panel of spacecraft. They are illustrated on Figure 2-16. For each of them, two shielding concepts have been defined, taking into account the results of the preliminary tests, and considering the coherence with the spacecraft needs.
Figure 2-16: Selected basic configurations and relevant enhanced configurations for testing
The goal of this test campaign was to determine the Ballistic Limit at 15km/s, using tests results at 7km/s and hypervelocity theory equations. 31 tests on 10 configurations have been carried out. In general the enhanced configurations outperform the basic configurations.
For configurations B4 and B7, testing of the enhanced configurations has shown that they require less mass for the same performance than the basic configurations. Some of the enhanced configurations require more spacing than the basic configuration. Two or three tests have been performed for each enhanced configuration in order to derive their ballistic limit curves.
Accommodation of these shielding configurations on the satellite have been evaluated, in terms of areal density (with box) compared to shielding performances, thermal performances (interference with thermal design), mechanical performance and volume requirement.
It results that:
• Stand-offs have the biggest potential, no matter which layer is regarded
• Heavy stainless mesh is valuable to reinforce the first layer (stand-off); More detailed mechanical analyses are needed and a detailed assessment of thermal aspects should be performed.
• Aramid and Nextel show a good mass-to-shielding ratio as Intermediate layers. Applicability in clean room should be investigated.
• Intermediate skins and Tungsten reinforced face sheets are promising for panels. The reference for fuel tank showed good performances. Improvements should however been investigated considering te criticality.
Several solutions of architectural and shielding improvements have been defined for the two reference satellites. In the case of the SAR satellite, it appears that the risk of failure is reduced by 75 to 82% with an additional mass of 2,5 to 3,5% of the dry mass. In the case of the optical satellite, the risk of failure is reduced by 40 to 42% with an additional mass of 0,8% of the dry mass.
ASSESSMENT OF SELECTED SOLUTIONS
Evaluation of solutions
The selected solutions have been assessed and compared with respect to a set of criteria including their performances, their implementation, their impacts on spacecraft design, equipment design, environment, operations, system reliability, AIT, cost, and their technical maturity.
The performance of the solution characterises its efficiency with respect to debris impact. It has been evaluated for the satellite architecture solutions with a generic vulnerability tool. This evaluation has been carried out on the reference optical satellite. The figure 2-17 illustrates the reduction of the reference satellite failure probability when applying the equipment repositioning solutions and the satellite re-orientation solution.
Figure 2-17: Efficiency of selected satellite re-orientation and equipment repositioning solutions in the case of the reference optical satellite. These figures assume a re-orientation all along the orbit.
The efficiency of the solution depends on the equipment on which the solution is applied. For illustration the Figure 2-18 shows the relative efficiency of different solutions applicable to a critical equipment of the optical satellite. Only the enhanced external protection improves the PNP of several equipment items.
In the case of translation of the equipment (to move it away from the front side), the efficiency increases with the distance, but moving the equipment by a few centimeters gives already about half of the expected efficiency.
Figure 2-18: Comparison of the solutions with respect to their efficiencies: case of the CCU of the reference optical satellite
The solutions have also been compared with respect to the other criteria, and in particular with respect to their impacts on the satellite configuration. The Figure 2-19 illustrates the relative position of the selected solutions with respect to both the impacts of the solutions on the satellite layout and the expected efficiency, when the solution is taken into account in early phases of the development. To carry out this relative comparison, the solutions have been applied to the same equipment (re-orientation and additional margins are exceptions).
It appears that that in general the higher the efficiency of the solution, the higher its impacts on the satellite configuration and on cost aspects. Thus the physical segregation of redundancy is an efficient solution, but is also the most complex one in terms of implementation as it requires additional volume, harness and connectors, and generates additional mass and power need. On another hand, hiding an equipment behind another one, or relocating a critical equipment leads to a lower efficiency, but has low impact on the design, even no impact if it is taken into account since the Phase 0/A.
It is then recommended to take into account the protection against debris since the early phases of development to minimize or even avoid impacts on the satellite.
Figure 2-19: Efficiency and complexity of implementation of architecture level solutions (case of implementation
Other axes of comparison of these solutions are their availability and their interest.
The availability of these solutions is linked to their technical maturity and the need for additional development. As illustrated on the Figure 2-20, a major part of the solutions are available at short term and could be taken into account in a Phase 0/A of a project starting now.
Some solutions require a specific development. The equipment compartment assumes a modification of the equipment itself. The shielding of an equipment item has to be designed and tested, even if one of the proposed shielding configurations is used. Likewise the re-enforcement of the satellite wall has also to be tested, even if one of the proposed shielding blocks is used. Finally, the distributed architecture for the satellite has also to be developed and tested.
The fractionated satellite and the distributed architecture among various modules requires technologies development, such as the interlink communication, the control of the cluster, etc.
Figure 2-20: Availability of the selected solutions
Some of the solutions do not only protect the satellite elements against impacts with small debris, but could also improve the reliability of the satellite (e.g. external redundancy, distributed architecture), or improve its performance (additional margins, in orbit spare) as illustrated in Figure 2-21. In particular, solutions improving the reliability of the satellite would be seen positively by the space insurance.
Figure 2-21: Added value of selected solutions with respect to their impact on the satellite
In order to compare the solutions with respect to the various criteria already mentioned, a strength/attractiveness approach has been considered where the strength describes the capabilities of the solution and the attractiveness represents the interest of the solution in terms of easiness of implementation.
The strength includes the technical maturity, the potential added value, the efficiency and the capability to reuse equipment.
The attractiveness includes the capability to minimize the impacts on the satellite layout, the additional harness, the impacts on operations, the additional cost and the impacts on AIT.
The Figure 2-22 gives the evaluation of the selected solutions following this approach.
This approach allows also to define complementary actions to improve the attractiveness or the strength. To increase attractiveness, one could work on the industrialization of some options or reduce the cost. To increase the strength, one could work on the efficiency of a solution, or its road map to raise its TRL.
Figure 2-22: Evaluation of the strength and attractiveness of the satellite architecture solutions
Applicability
The applicability of the selected solutions to the different LEO missions and different types of LEO satellites has been evaluated and the applicability matrix is shown on Figure 2-23.
Each solution is only applicable to some types of mission and satellites, or to some of the equipment. For instance, solutions with additional mass and volume are applicable to large satellites, but not to small and compact ones.
Combination of solutions
This evaluation shows that a single solution does not bring the required significant reduction of satellite probability of failure, or cannot be applied to all critical equipment; in addition, the implementation of a single solution could be limited by its impacts on the satellite. Thus, combining several solutions seems an attractive approach.
The combination of solutions depends on the type and size of the satellite, of its orbit and of the mission needs. Indeed, each solution has its domain of applicability in terms of mission, and cannot be used for all the satellites. The Figure 2-24 illustrates the approach to be followed for defining the combination.
Figure 2-23: Applicability of solutions to satellites
Figure 2-24: Possible combinations of solutions in early phases of the project
PROPOSED GUIDELINES
Subsequently to the definition and evaluation of the solutions, a set of 67 design rules to assist in the choice and implementation of impact protection solutions has been derived. The design rules have been compiled according to the following categories:
• Impact risk assessment procedure
• Levels of impact protection in a space system
• Criteria for evaluating feasibility of protection solutions
• Impact protection limits
• Impact testing
• Shielding materials
• Shielding design
• Spacecraft-level solutions
• Spacecraft subsystems architecture
• Relationship to the phases of a spacecraft programme
• System-level solutions
The rules are considered to be sufficiently generalised that they can be followed during the design of any unmanned LEO space system. Therefore, they will be presented to international organisations involved in the development of guidelines and standards (e.g. IADC and ISO) for consideration and possible adoption.
RECOMMENDATIONS
There is no generic solution for the protection of LEO satellites against small debris; a case by case analysis has to be done for each mission and satellite, as the particular solution is strongly dependant on the satellites configuration (geometry, layout size) and mission characteristics like the orbital parameters.
A palette of solutions is proposed, with a range of applications, advantages and drawbacks.
System level solutions shall be decided early at mission and system level, as solutions such as fractionated architecture will impact the system definition while architecture level solutions shall be considered during the development of the satellite, as it impacts the satellite definition.
It is worth taking into account the needs and solutions for protection of the satellites against debris since the early phases (0, A) of the project. High level numerical simulations could be carried out to evaluate the vulnerability of the preliminary configuration of the satellite with respect to debris and trade–off various solutions of protection.
The simulations allow to identify the most critical equipment of a satellite configuration. Working on the protection of these equipment items will already highly improve the overall survivability.
Moving equipment to a safer place is the simplest solution; solutions to reduce the area of critical surface, such as rotation of equipment or of the satellite shall also be considered.
The use of shielding shall preferably be restricted to the protection against debris of 3 to 4 mm size. It allows to reduce significantly the risk of failure.
The selected solution shall be compatible with the Design to Demise, in order to allow an uncontrolled re-entry after the end of mission of the satellite.
Additional experimental tests should be carried out to supply the Ballistic Limit Equation dedicated to the desired configuration if it is not available, but also to consolidate the extrapolation of the BLE above 7 km/s. To that aim, experimental tests at 10 km/s or more should be necessary.
Innovative shielding concepts using new materials have been defined and tested.
Design rules have been elaborated to increase the robustness of European satellites in the growing population of small debris
Potential Impact:
Potential impacts and beneficiaries
The ReVuS project has allowed to propose design solutions to reduce the vulnerability of the future LEO satellites, whatever their size, mission and orbit, against impacts of small-sized debris, that means debris which cannot be detected and tracked. Through its results, it significantly contributes to the European capacity to protect space assets from space debris.
Potential impacts
The main benefits of the ReVuS project are:
• The evaluation of the respective contribution of each range of debris size (between 0.1 mm to 5 cm) to the risk of failure of a satellite, and the highlighting of the 3 to 5 mm range as the main contributor. This is a major impact for the selection of a solution to protect the satellite and in particular for the design of shielding configuration
• The assessment of a set of promising shield building blocks, with their tested performances. As a result, the project indicates what type of building block is more adapted to the different layers, and gives their performance in terms of ballistic limit versus the overall areal mass, and reports whether it is used in a shielding configuration or as re-enforcement of existing MLI or sandwich panel
• A set of tested enhanced shielding configurations for several basic equipment configurations. The best enhanced shielding solution requires only 1/4 of the mass when compared to a corresponding state-of-the-art shielding solution. The project gives quantitative evaluation of these enhanced shielding configurations in terms of ballistic limit curve, expected reduction of satellite vulnerability, and impacts on the satellite (mass, environment, etc).
• A palette of satellite architecture solutions, with their advantages and drawbacks, to be used on existing (in development) or new design of LEO satellites. This palette of solutions includes also an applicability matrix, as function of orbital parameters and type (size) of satellite, as there is no solution applicable to all satellites.
• Design rules for taking into account protection against small debris in the design of future LEO satellites
• A proposed approach to evaluate the vulnerability of satellite since the early phases of the design and to support the definition and selection of adequate solution, in particular the combination of solutions.
• A new collision risk algorithm, implemented within a space debris evolutionary model that allows for the assessment of small debris impact risks and potential failures of active spacecraft.
Beneficiaries
The definition of possible solutions and design rules to reduce vulnerability of satellites against small debris is a main benefit for the next generation of LEO satellites, whatever the user is. It will contribute to a more efficient design, a higher quality and enhance the competitiveness of European space systems.
These results of the ReVuS project will benefit to the space community, institutional, industrial as well as scientific.
Industrials, satellite and space system manufacturers will benefit from approach to evaluate the vulnerability of future satellite to small debris, from a palette of solutions and design rules to reduce this vulnerability, and from a methodology to define the adequate solution. The existing set of shield building blocks and shielding configuration with their performances will provide them either with an existing adequate solution, or with guides to define the adequate solution.
Research institutes and universities have used their scientific findings and models to improve the design of space system. They have also improved their models of environment for the evaluation of impacts of small debris on a satellite and have developed architectures to reduce the vulnerability of satellites. The advancements generated during the project will be used as a baseline for future research. Research institutes and university will benefit from a reduced amount of work required to attain a certain research level, facilitating further advancements in research.
The users of space system (mainly operator) could now have a clearer picture of the risk of degrading or losing the mission following an impact with small debris over the expected lifetime, of the potential solutions to reduce this risk and of their relevant impacts on the satellite, and thus has a support enabling adhoc decision with respect to problem of small debris.
The insurers will have also a clearer view on the risk of degrading or losing the satellite, of the interest, risk and possibly positive effect on satellite reliability of the possible solutions.
The institutional players will have a starting point directly applicable to on-going and future projects. In particularly, the proposed ReVuS solutions could be used to reduce the risk of losing the post mission disposal capability of the future satellites, which is a key contributor to the stabilization of the population of debris.
In addition to the space community, ReVuS will also benefit to European citizen. Indeed, Europe is using Low Earth Orbit satellites to provide services like earth observation, resources monitoring, defence and security systems that at the end benefit to citizens. Reducing the satellite vulnerability to small debris allows to improve the maintenance of the continuity of these services to the citizens.
On societal point of view, ReVuS has allowed to gather in the same study research laboratories, universities, manufacturers, industrials and users (insurer) for the complex problem of small debris remediation. In particular, the team was not restricted to the space community, but involves also an insurer, who gives a complementary angle of view of the debris problem, and a manufacturer of shielding devices for civil and military applications.
Dissemination of results
The results of ReVuS have been continuously disseminated all along the project. This dissemination has been carried out through the ReVuS site, conferences, publications and a dedicated workshop last October, at the end of the project.
Lecturers have been done at adequate conferences and symposium all along the 34 months of the project, in order to disseminate interesting results as function of the progress of the project, and to make the project known by the space community, and beyond.
To that aim, four types of presentation have been targeted, according to the content and audience of the conferences: a general presentation, describing the overall ReVuS objectives and results at successive steps of progress, presentations focused on technical subjects to show in more details some main results, and short (overview) presentation of the project to space and non-space audience.
• General presentation of the project, describing the overall ReVuS objectives and results at successive steps of progress:
- Conferences covering all space subjects: IAC 2012 (October 2012, at mid project) - 2nd FP7 Space Conference, November 2012 - IAC 2013 (October 2013, close to the end)
- Conferences on space security (IAASS, October 2011 and May 2013), on sustainability of space activities (ISU, February 2012), on space debris (European conference on space debris, April 2013)
- Forums on space risk (World Space Risk Forum, June 2012)
• Presentations focused on technical subjects to show in more details some main results:
- Distributed architecture (NASA/ESA Conference on Adaptive Hardware and Systems, June 2012 - Industrial Showcase event of Department of Engineering, University of Leicester, February 2013 – Lecture on Microelectronics and Embedded Systems for Space Applications, June 2012 - IEEE Aerospace Conference, March 2014)
- Shielding material and tests (12th International Symposium on Materials in the Space Environment, September 2012 - 6th European conference on space debris, April 2013 - IAC 2013, October 2013 – AAAF conference on dynamic behavior of materials and structures, January 2014)
- Vulnerability analysis (IAC 2012 ,October 2012 - 6th European conference on space debris, April 2013)
• Short general presentations (overview) of the project to space and non-space audience:
- Space Situational Awareness (SSA) Seminar, organized within Polish Presidency of the Council of EU, September 2011
- NordicBalSat Conference Broadening the Base of Europe’s Space Community, February 2012
- EISC Seminar, May 2012
- Science Pique-Nique, May 2012
- ILA, Berlin Air Show, September 2012
- WIRE (Week of Innovative Regions in Europe), June 2013
- SSA conference, November 2013
• Participation to specific working groups on space debris in order to present the ReVuS results, or to discuss the utilization of some of the project results:
- Overview of ReVuS results and progress at the CNES synthesis group of debris (CNES June 2011, June 2012, June 2013)
- Discussion (PHS, EMI) on the results of the first phase of the vulnerability analysis and of the shielding aspects at a joint session of the IADC WG3 (Protection Working Group) and WG4 (Mitigation Working Group) meetings in Montreal, Canada during 22 – 25 May 2012.
- Presentation (PHS) of the preliminary results of the design rules activity at the ISO TC20/SC14/WG7 (Orbital Debris Working Group) meeting in Guildford, UK on 7/8 November 2013.
- Presentation of the final results of the design rules activity at the IADC WG3 (Protection Working Group) meeting in Beijing, China on 12-15 May 2014 and at the ISO TC20/SC14/WG7 (Orbital Debris Working Group) meeting in Tokyo, Japan on 26-28 May 2014.
- Discussion (EMI) on the results of the shielding aspects at the 31st IADC meeting, 17-19 April 2013 in Darmstadt, Germany
• Exhibition of space debris modelling to non-space audience:
- Exhibit run by Southampton at the University of Southampton’s award winning science and engineering day, 2012 and 2013, with visitor numbers in excess of 2000.
In addition to the conferences, a ReVuS workshop has been organized last October, close to the end of the project, over a full day. The objective was to present in detail all the results of ReVuS to institutions (agencies), academia and industry. This workshop gathered up to 40 people and has been a main contributor to the dissemination of the ReVuS project.
Figure 3-1: Dissemination of ReVuS results
Exploitation of results
Most of the ReVuS results are public and have been disseminated all along the project, as discussed above. There is no patent and thus no restricted exploitation of the data.
ReVuS proposes an approach to evaluate the level of risk due to small debris (and thus whether there is a need for protection or not) and to identify the solutions for reducing the vulnerability that could be applied to a satellite in development. This approach could be used by the manufacturers or even by the operators of future LEO satellites.
The vulnerability analysis, aiming at evaluating the probability of failure of the satellite configuration, and at identifying the critical elements, has been described in the project and can be reused by manufacturer or insurer. There are existing tools to carry out such an analysis.
Although this analysis has been done on reference satellites, some results can be directly exploited:
• the debris with size in the range 3 to 5 mm appear as the main contributor to the satellite probability of failure, for the two reference satellites which are different both on orbit and configuration. In both cases, it has been shown that protecting the satellite against this range decrease significantly the probability of failure. This size of debris could be taken into account in the preliminary development phase of a satellite, in particular for the sizing of shielding protection.
• The level and characteristics (direction, repartition as function of true latitude) of the flux of particles has been given for different altitudes. These figures can be used for preliminary trade-offs (short evaluation of preliminary configurations of satellite, candidate orbits, etc).
Preliminary results have been presented at the IADC working groups.
ReVuS proposes a palette of solutions at architecture level to reduce the vulnerability and gives an applicability matrix. This matrix is a key as each solution is only applicable to a certain range of satellite. Thus, depending on the orbit altitude and type of satellite, only a set of solutions can be considered. In addition, these solutions have been evaluated for a reference satellite, so that, even if the performance of the solution depends on the satellite configuration, the user will have in hands an applicable set of solutions with their relative performances and impacts.
Among the candidate solutions, ReVuS has identified solutions with added value: these solutions could increase the performances of the satellite or its reliability. Such solutions are of interest for different users: the manufacturer, because they participate to the performances (margin on resources) or the overall reliability (for instance external redundancy) of the satellite; for the operator, as it could expect a higher performance at the beginning of utilization of the satellite; for the insurer who could reduce the premium if the reliability is higher.
The distributed architecture solution proposes a new computing architecture inside a satellite, at the opposite of the usual centralized computer, with advantages in terms of reliability. Results of tests are given in the project. However, additional development and verification will be needed in parallel to the development of satellite.
ReVuS makes available results of the test of a large set of shielding building blocks. These building blocks are enhancements of protective elements like MLI, sandwich panels, intermediate layer. They use different materials with various thicknesses. These building blocks have been evaluated in terms of performances and impacts on the satellite. All these data could be used by the satellite manufacturer to re-enforce external sides of the satellite, or define shielding configuration. But they are also available for non-space user that could be interested by the performance of some material for the protection of ground vehicle, aircraft, people against particles of lower velocity than debris.
ReVuS has defined a dozen of equipment configurations, taking into account their relative location in the satellite and their mechanical and thermal environment. For four of them, ReVuS proposes a set of two shielding configurations, each relying on the shield building blocks. They have been designed to protect the equipment against debris of size up to 3 mm. They have been tested, and ReVuS makes available the results of their tests, their ballistic limit curve and their characteristics. The user can benefit either as an existing product (re-use the shielding configuration as is while adapting the surface to the equipment) or as a set of data to support the definition of relevant shielding configuration. These data are valuable support which gives ideas of mass, size to protect equipment against debris of 3 mm.
As a synthesis of all the achieved results, design rules to protect LEO satellites against small debris have been generated. They could be used by manufacturer of future LEO satellites. They could be of interest for international organizations involved in the development of guidelines and standards (e.g. IADC and ISO) for consideration and possible adoption. Preliminary results have been presented at the Orbital Debris Working Group of ISO in last November. The final design rules will be presented to the adequate Working Groups of IADC and ISO in May 2014. This is the first identified exploitation of the ReVuS results.
List of Websites:
htpp://www.revus-project.eu
Contact points:
Claude Cougnet: claude.cougnet@astrium.eads.net
Krystyna Macioszek: Krystyna.MACIOSZEK@astripolska.pl
