Scholarly article on topic 'Effects of Debris Cloud Interaction with Satellites Critical Equipments – Experiments and Modeling'

Effects of Debris Cloud Interaction with Satellites Critical Equipments – Experiments and Modeling Academic research paper on "Earth and related environmental sciences"

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{satellite / damage / "solar cell" / harness / mirror / "debris cloud" / modeling.}

Abstract of research paper on Earth and related environmental sciences, author of scientific article — J.-M. Sibeaud, C. Puillet

Abstract The level of damage imparted on three types of satellites components resulting from hypervelocity impacts of aluminum projectiles onto a neighboring structural part have been assessed experimentally by using a double stage light gas gun. The experimental configurations were designed to simulate debris encounters with a space vehicle during its operational life time around the Earth. The first test was conducted against a set of separate harness strapped down on the inner surface of a carbon facesheet sandwich panel in order to assess the effectiveness of redundant cables spacing in case of unitary debris perforating impact onto the structure. The resistance to damaging of a silicon carbide mirror and a solar panel was then evaluated against debris cloud generated by prior impact of chunky debris with the satellite structure. Potential losses of electrical functional capabilities were then derived. The vulnerability/Survivability Pléiades software was updated accordingly.

Academic research paper on topic "Effects of Debris Cloud Interaction with Satellites Critical Equipments – Experiments and Modeling"

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Procedía Engineering 103 (2015) 561 - 568

Procedía Engineering

www.elsevier.com/locate/procedia

The 13th Hypervelocity Impact Symposium

of Debris Cloud Interaction with Satellites Critical Equipments - Experiments and Modeling -

J-M Sibeauda*, C. Puilletb

aCEA DAM, GRAMAT, F-46500 Gramat - France bCNES Centre Spatial de Toulouse, 18 avenue Edouard Belin, F-31034 Toulouse Cedex, France

Abstract

The level of damage imparted on three types of satellites components resulting from hypervelocity impacts of aluminum projectiles onto a neighboring structural part have been assessed experimentally by using a double stage light gas gun. The experimental configurations were designed to simulate debris encounters with a space vehicle during its operational life time around the Earth. The first test was conducted against a set of separate harness strapped down on the inner surface of a carbon facesheet sandwich panel in order to assess the effectiveness of redundant cables spacing in case of unitary debris perforating impact onto the structure. The resistance to damaging of a silicon carbide mirror and a solar panel was then evaluated against debris cloud generated by prior impact of chunky debris with the satellite structure. Potential losses of electrical functional capabilities were then derived. The vulnerability/Survivability Pléiades software was updated accordingly.

© 2015TheAuthors.Publishedby ElsevierLtd. Thisisanopen accessarticleunderthe CC BY-NC-NDlicense (http://creativecommons.org/licenses/by-nc-nd/4.0/).

Peer-review under responsibility of the Curators of the University of Missouri On behalf of the Missouri University of Science and Technology Keywords: satellite; damage; solar cell; harness; mirror; debris cloud; modeling.

Effects

1. Introduction

The ever growing threat level posed by orbiting debris has led the French official services to improve performance knowledge of space vehicles in terms of survivability against hypervelocity encounters during their operational lifetime. In order to meet this specific need, the CEA Gramat has been developing new capabilities within its global vulnerability software suite Pleiades. This tool, which is devoted to terminal ballistics applications, provides the necessary accommodation structure to the vulnerability modeling of spacecraft orbiting around the Earth. The threats considered are the orbiting debris which may affect the function of satellites as a result of hypervelocity impact onto their external walls and equipment.

On behalf of the procurement agency (DGA) of the French Ministry of Defense and the French Space Agency (CNES), the CEA Gramat carried out experimental and modelling studies aiming at studying the effect of hypervelocity impacts of orbital debris on space vehicles in configurations representative of spacecraft operational life cycle. A work program based on trials on real components was conducted within the frame of the DGA's national project on knowledge database build-up relating to embedded architectures vulnerability to natural and man-made environments, including the direct and indirect effect of hypervelocity impact. The outcome of such program is to make available a modeling tool and its associated components vulnerability database in order to help assessing the risks of satellites loss of functionality during their

* Jean-Marc Sibeaud. Tel.:+33-565 10 5355 ; fax:+33-565 10 5433 E-mail address: jean-marc.sibeaud@cea.fr

1877-7058 © 2015 The Authors. Published by Elsevier Ltd. This is an open access article under the CC BY-NC-ND license (http://creativecommons.Org/licenses/by-nc-nd/4.0/).

Peer-review under responsibility of the Curators of the University of Missouri On behalf of the Missouri University of Science and Technology doi: 10.1016/j.proeng.2015.04.073

operational life cycle due to orbital debris encounters. Given prior knowledge of the debris flux density in orbit (ESA MASTER or NASA ORDEM) it will help as well designers to harden space vehicles in the face of such harsh environment.

2. Experimental configurations

Following previous hypervelocity direct impact experiments carried out on aluminum spaced plates [1] and honeycomb structural panels [2-3], a set of critical components were selected and tested against the effects of debris cloud resulting from prior impact of a chunky projectile onto neighboring structural elements.

The following three types of components were provided by the French Space Agency (CNES) for damage assessment:

- An internal electrical harness strapped on the back face of a honeycomb sandwich panel,

- An external terrestrial observation optical sensor,

- A solar generator panel.

The following subsections describe the three types of components to be investigated which are typical of the most exposed ones to disruption due to hypervelocity expanding fragment clouds. The experiments were carried out with the Persephone double stage light gas gun in operation at CEA Gramat. Down selection of experimental configurations and target set-up are also explained. All target configurations were tested against a 0.49 gram, 7 mm diameter spherical aluminum projectile launched at the nominal velocity of 6 km/s.

2.1. Redundant electrical harness

The vulnerable component investigated was a redundant harness consisting in a set of three pairs of cables attached to the inner surface of a carbon facesheet honeycomb sandwich representing the satellite's structural frame. The target module built-up for the experiment consisted in 25 mm aluminum honeycomb core glued with 0.4 mm carbon composite facesheets. The honeycomb was made of 0.05 millimeter thick aluminum sheets. The front face of the panel was covered with MultiLayer Insulation (MLI). The full target assembly is shown in Figure 1. It comprises a 10 mm witness plate (WP) hold steadily to the sandwich structure 150 mm downrange from the rear carbon facesheet. The pairs of cables were attached horizontally and parallel to each other on the back face of the sandwich. The target was set up at normal incidence. The central pair was placed on the shooting line; the two others were located 38 mm off-axis.

When the incoming orbital debris perforates the sandwich wall under hypervelocity conditions, it will shatter and form an expanding debris cloud that will eventually break off one pair of cables and leave the distant ones undamaged if the separating distance between them is large enough.

Fig. 1. Photograph of target and sketch of Experiment #1 where three groups of cable pairs lying on the back face of a carbon composite honeycomb sandwich are exposed to the debris flux generated by a single spherical projectile impact on a carbon honeycomb sandwich.

2.2. Baffled optical sensor

Of course, due to its orientation in the direction of the Earth, a mirror equipping an optical sensor has low probability to suffer direct impact of centimeter class debris because of the presence of the surrounding baffle whose function is primarily for preventing unexpected light rays to reach the sensor surface. By its shape and dimension, the baffle would protect as well the mirror from direct hit by the orbital debris itself but in the other hand would cause a cloud of smaller debris to form that may eventually hit the mirror and reduce its optical properties.

The objective of the experiment was to examine this type of scenario where a planar Silicon Carbide mirror is exposed to the debris cloud generated by prior hypervelocity impact of 7 mm aluminum spherical projectile onto a 2 mm aluminum alloy plate at 45° obliquity. In order not to expose the mirror to the fragmenting projectile but rather to the lighter fragments coming from the aluminum plate, the mirror was shifted off-axis and oriented correctly with regard to the plate so that to represent more closely the real sensor mounting. The target arrangement is shown in Figure 2.

Fig. 2. Photograph of target and sketch of Experiment #2 where a planar optical silicon carbide mirror is exposed to the debris flux generated by prior impact of a unitary spherical projectile onto a 2 mm aluminum alloy plate under 45° obliquity.

2.3. Solar panel

Due to their large dimensions solar panels may be subject to direct hits by centimeter class orbital debris. Most of the time, such events cause only minor damage to the cells being hit or to their neighbors. The overall solar panel retains therefore a large proportion of its initial power which leaves the satellite's operational capability unaffected. On the other hand, an indirect hit caused by prior impact of the debris onto a neighboring structural element may affect the panel more strongly as a result of a larger foot print due to the subsequent fragment cloud expansion originating from impact.

Figure 3 shows the target arrangement that was set up to investigate the solar panel vulnerability to expanding fragments cloud exposure. A similar approach to the one already presented for the optical sensor was considered: a 45° obliquity aluminum alloy plate was used to create the debris cloud. The panels were exposed to the light fragments emitted by this cloud by shifting off the shot line the solar panel itself.

Fig. 3. Photograph of target and sketch of Experiment #3 where a solar panel is exposed to the debris flux generated by prior impact of a unitary spherical projectile onto a 2 mm aluminum alloy plate under 45° obliquity.

3. Experimental results

3.1. Electrical harness on the back face of honeycomb sandwich

Due to flash radiography malfunction, the actual impact velocity could not be accurately determined. However, the launching conditions were exactly the same to the ones relating to the subsequent set of experiments (see following paragraphs). An impact velocity of 5700 ± 20 m/s was therefore assumed. Nevertheless; a preliminary three-dimensional Ouranos hydrocode modeling had been undertaken to predict the axial residual debris cloud velocity which was found to be 5850 m/s, or 97 % of impact velocity, for an impact velocity of 6 km/s, as can be seen on Figure 4. Moreover, the axial velocity of the third ring of fragments appears to be 5.4 km/s.

Applying merely the ratio of velocity of 97 % to the experimental data would therefore provide a value of 5.53 km/s for the main residual fragment velocity.

Fig. 4. Ouranos hydrocode simulation of normal impact on honeycomb sandwich.

Figure 5 shows the residual honeycomb sandwich recovered after the shot, revealing that the central pair of cables was broken off while the lateral redundant pairs were left undamaged. In such situation, the remaining conductors keep on assuring their function of redundancy for the one destroyed.

Moreover, the geometrical reconstruction of the fragmenting cones was possible by analyzing the witness plate, as can be seen also in Figure 5. The witness plate did not reveal any local reduction of the shattering effect due to fragment masking by the harness which means that the residual fragment kinetic energy was in large excess upon exit from the honeycomb module. The inner fragmenting cone of large fragments, essentially from the projectile are scattered within a 118 mm diameter circle corresponding to a 36° apex angle measured from the spherical projectile center at the point of impact. The maximum extension of fragment impacts was on a diameter of 200 mm corresponding to a 60° apex cone aperture. The maximum angular extension of debris fragments for which there would not have been interaction with the lateral pair of cables appears to be close to 80°.

Lastly, it should be noted that the projectile experienced a rather large deviation, of nearly 7° from its initial path, due to the driving process created by each individual hexagonal cell wall, considering the fact that the point of impact was not centered on one of them.

Fig. 5. Post shot target #1 with estimated hit velocity of 5.7 km/s showing damaged harness and fragmenting cloud extension on witness plate.

3.2. Optical sensor

The impact velocity was found to be 5699 m/s and the projectile integrity before impact was then confirmed. Taking advantage of the separation distance from the aluminum shattering plate and the optical component itself, one pair of X-ray channels were located on orthogonal axis, in order to allow the determination of the post perforation debris cloud expansion dynamics. Figure 6 shows the vertical axis X-Ray view together with the post shot aspect of the mirror surface. A debris cloud axial velocity of 3569 m/s was therefore derived. At this stage, it is worth noting that the perforation hole was 18.5mm height and 15.5mm in the aluminum shattering plate. Consequently, 4.9 grams of aluminum was expelled from it, more than ten times the projectile weight (0.47g).

The debris cloud appears like a unique layer of relatively homogenous mass density, except at the leading edge where heavy projectile fragments keep on flying in the direction of the shot line. Only a small proportion of light fragments were able to reach the mirror surface as expected. However, the mirror breaking-off near its bottom edge was caused by the impact of heavy fragments on the lower part of the retainer plate surrounding the mirror and not by the light fragment impacts.

Analysis of the mirror surface (see Figure 6) shows white markings from 2 to 10 mm long corresponding to the impact of each of these individual light fragments. The deposit of matter on the mirror seems to follow some natural geometrical rule of under angles varying from 14° to 25° from top to bottom of the reflecting surface assuming the fragments have typical dimensions of 1-2 mm. Surface prints not deeper than 2 ^m at each white marking was evidenced by the CNES from specific analysis. The number of impacts on the surface of the mirror varies from 86 to 100 depending on the count of micro impacts. The fragment impact density on the reflecting surface is therefore between 1/cm2 to 1.2/cm2. The loss of reflecting power of the mirror will be addressed accordingly by the CNES.

Fig. 6. Post shot target #2 with damaged reflecting surface of mirror by light debris cloud generated by prior impact at 5.699 km/s - X-Ray visualizing of debris cloud 24.9 ^s after impact on shattering aluminum plate.

3.3. Solar panel

The impact velocity was 5700 m/s from analysis of pre impact X-Ray pictures. The dual X-Ray debris picture taken at 24.95 ^s and 32.92 ^s after impact of the projectile on the aluminum shattering plate provided an axial debris cloud velocity of 3722 m/s. The Horizontal axis X-Ray taken at 24.95 ^s is shown in Figure 7 as well as the damaged solar panel which was impacted by the light fragments. As in experiment #2, the mass of aluminum torn from the 2 mm plate is approximately ten times the projectile mass (5.3 grams).

The surface of the coupon appears to be damaged by penetrating and non-penetrating impacts. The impact points are associated with an elongated white deposit that can reach 40 mm long, which is much longer than expected based on geometric reconstruction and of complete fragment erosion. These deposits resulted probably from the deposition of material torn from the solar cells themselves. Further analysis addressed by the CNES has highlighted the loss of power output resulting from the panel obscuration.

Fig. 7. Post shot target #3 with damaged surface of solar panel by light debris cloud generated by prior impact at 5.699 km/s - X-Ray visualizing of debris cloud 24.9 ^s after impact on shattering aluminum plate.

4. Pléiades modeling of the satellite damaging in orbit

4.1. Development context

Pléiades is the software tool developed by CEA Gramat for dealing with vulnerability and survivability issues of weapon systems in the context of operational scenarios [4]. The Pléiades structure offers the necessary services for hosting newly developed interaction phenomena. This possibility is illustrated in the next sub-section for satellites vulnerability to orbital debris interaction at hypervelocity.

The subject of satellite survivability was primarily initially on behalf of the French Space Agency (CNES) in the context of an ever growing orbital debris population which may cause premature loss of missions due to the disruption of embedded functionalities. The CEA Gramat was tasked to work on this subject in order to build-up an engineering tool that would be able to help anticipating the consequences of hypervelocity impacts of debris on satellites structure and therefore help identifying design guidelines aimed at improving the survivability of missions. CEA Gramat responded to this request by proposing the Pléiades software as the accommodating structure for new model development.

The model of wall perforation and creation of debris cloud resulting from hypervelocity interaction with orbital debris has already been presented in [3] and will not be recalled here. This model is based on Christiansen's Ballistic Limit Equation (BLE) [5] providing the critical projectile diameter for perforating a wall. The equation is valid for homogenous plates, spaced plates, honeycomb sandwich structures with aluminum or carbon composite facesheets, as well as for a combination of these. It requires the knowledge of physical and geometrical characteristics of all subcomponents (Honeycomb definition and thickness, facesheet material and thickness...). Other investigators' developments are considered as well for implementation as they rely on a large amount of test data on various types of spaced plates structures, see for example [6-8] to quote a few of the numerous articles published in the area.

Vulnerability models for the three types of components studied have been established and implemented under the form of threshold damage, based on the phenomenology of hypervelocity interaction described previously. The damage producing mechanism is brought here by the expanding fragment cloud resulting from prior impact of single large debris with the external structure of the satellite. Pléiades susceptibility calculations of critical components to the impact by the debris flux are illustrated in case of a generic unitary hypervelocity encounters.

4.2. Susceptibility modeling of harness

A Pléiades simulation was performed with the same honeycomb structure as the one presented earlier except for the line of impact aiming one of the lateral pair of cables. For completion, the maximum dimension of the elementary hexagonal cell of honeycomb structure was 4 mm and the delamination threshold resistance of carbon composite facesheets was 50 MPa.

The harness vulnerability is simply modeled by considering it is out of order when exposed to the debris flus of an impact generated fragment cloud. The Pléiades model is exercised in a generic impact scenario on the satellite for which the impact conditions are exactly the same as in experiment#1. Figure 8 shows the Pléiades result in terms of fragmenting cloud interaction with the harness and as well with the satellites internal components. Of course, further calculations can be performed in terms of functional damage with the help of the fault tree update.

Moreover, most internal components appear to be in the form of aluminum boxes for which a specific damage model based on the areal kinetic energy density of the fragment cloud downrange. This areal density decreases with the cloud expansion. The vulnerability of an aluminum box is therefore a function of its distance to the external wall structure. Whereas an aluminum box placed immediately behind the satellite's wall shall be destroyed, it may survive the impact if located closer to the satellite's core.

Fig. 8. Pléiades simulation of a debris impact on a carbon composite honeycomb structure with three pairs of cables strapped on its back face. The debris trajectory is shown as well as the resulting fragment cloud.

4.3. Susceptibility modeling of optical sensor and solar panel

The proposed model is based on the count of elementary impacts on the mirror surface and on the assessment of the total area of the elementary deposits created by the secondary fragments impacts, for a given fragment density per unit solid angle. The loss of reflecting capability is defined by a threshold of damaged area from the secondary cloud impact. The obscured mirror fraction is calculated with the area of each individual secondary fragment and its local angle of impact on the surface. A generic example of this kind of scenario of impact on an external optical sensor is illustrated in Figure 9 for the case of a large debris impact on the baffle surrounding the mirror. The secondary debris cloud generated by the calculation is clearly visible on the figure as well as the resulting impacts on the mirror itself. The outcome of the interaction depends on the threshold value of the obscured mirror fraction.

Figure 10 provides a Pléiades calculation example of a single high velocity debris impact on a structural member of the satellite. The resulting cloud generated by this impact interferes with the solar panel and consequently a level of damage expressed in terms of loss of electrical power output is derived.

trajectory wrt satellite

Fig. 9. Left: Pléiades simulation of a debris impact on an optical sensor - the initial impact on the baffle generates fragments further interacting with the internal mirror; Right: Example of Pléiades simulation of a debris impact on a solar panel.

5. Conclusion

The Pléiades software now takes into account the plate perforation modeling previously developed for the CNES, as well as the generic vulnerability model developed in this study for three types of critical components. Susceptibility calculations have been proven to be effective for any kind of impact scenarios and for a generic satellite. Vulnerability and survivability of satellites in orbit require now vulnerability threshold to the type of stimuli involved in case of hypervelocity encounters in orbit. By formulating assumptions on components vulnerability and functional losses, and by varying them at will, the analyst can quickly assess the weaknesses of a particular embedded architecture and propose improved variants less vulnerable in orbit. Simulations of operational life of satellites considering the evolving spatial environmental factors have also to be carried out.

Acknowledgements

This work was funded by the French Direction Générale de l'Armement (DGA). The authors express their gratitude to Christian Prieur (CEA) who conducted the Persephone 2-stage light gas gun experiments, and would like also to thank Christiane Maurice and Jean-Louis Domingues-Vinhas (CEA) for performing the calculations, Jean-Michel Desmarres (CNES) for SiC mirror measurements and Etienne Rapp (CNES) for evaluating the solar panel electrical damages.

References

[1] Sibeaud JM, Héreil PL, Albouys V. Hypervelocity Impact on Spaced Target Structures: Experimental and Ouranos Simulation Achievements.

International Journal of Impact Engineering. 2003; 29: 647-658.

[2] Sibeaud JM, Prieur C, Puillet C. Hypervelocity Impact on Honeycomb Target Structures: Experimental Part. Proceedings of the 4th European

Conference on Space Debris, Darmstadt, Germany, 18-20 April 2005.

[3] Sibeaud JM, Thamié L, Puillet Hypervelocity Impact on Honeycomb Target Structures : Experiments and Modeling. International Journal of Impact

Engineering. 2008; 35: 1799-1807.

[4] Sibeaud JM. "The Pléiades Vulnerability/Lethality analysis suite". Proceedings of the 14th International Symposium on Interaction of the Effects of

Munitions with Structures (ISIEMS), 2011.

[5] Christiansen EL. Design and Performance Equations for Advanced Meteoroid and Debris Shields. International Journal of Impact Engineering.

1993;14:145-156.

[6] Schäfer FK, Ryan S, Lambert M, Putzar R. Ballistic Limit Equation for Equipment placed behind Satellite Structure Walls. International Journal of

Impact Engineering. 2008;35:1784-1791.

[7] Ryan S, Schäfer FK, Destefanis R, Lambert M. A Ballistic Limit Equation for Hypervelocity Impacts on Composite Honeycomb Sandwich Panel

Satellites Structures. International Journal of Impact Engineering. 2008;41:1152-1166.

[8] W.P. Schonberg, F. Schäfer, R. Putzar. Predicting the Perforation Response of Honeycomb Sandwich Panels Using Ballistic Limit Equations. Journal

of Spacecraft and Rockets. Vol. 46, Number 5, Pages 976-981.