Reinventing Space Conference 2018

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Journal of the British Interplanetary Society

VOLUME 71 NO.11 NOVEMBER 2018 Reinventing Space Conference 2018

THE MARKET FOR A UK LAUNCHER Vadim Zakirov et al THE REMOVEDEBRIS SPACE HARPOON Alexander Hall RAPID CONSTELLATION DEPLOYMENT from the UK Christopher Loghry & Marissa Stender REDESIGN & SPACE QUALIFICATION OF A 3D-PRINTED SATELLITE STRUCTURE with Polietherimide Jonathan Becedas et al THE EPSILON LAUNCH VEHICLE: Status and Future Ryoma Yamashiro & Imoto Takayuki THE NORMS OF BEHAVIOUR IN SPACE: Our space – Whose rules? Lesley Jane Smith

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ISSN 0007-084X PUBLICATION DATE: 31 JANUARY 2019 Submitting papers International Advisory Board to JBIS

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Published by the British Interplanetary Society. Registered Company No: 402498. Registered Charity No: 250556. Printed by Latimer Trend & Company Ltd, Estover Road, Plymouth, PL6 7PY, England. © 2018 British Interplanetary Society. No part of this magazine may be reproduced or transmitted in any form or by any means, electronic or mechanical, including photocopying or recording by any information storage or retrieval system without prior permission from the Publishers. contents VOLUME 71 NO.11 NOVEMBER 2018

399 THE MARKET FOR A UK LAUNCHER Vadim Zakirov et al

406 THE REMOVEDEBRIS SPACE HARPOON Alexander Hall

410 RAPID CONSTELLATION DEPLOYMENT from the UK Christopher Loghry & Marissa Stender

416 REDESIGN & SPACE QUALIFICATION OF A 3D-PRINTED SATELLITE STRUCTURE with Polietherimide Jonathan Becedas et al

426 THE EPSILON LAUNCH VEHICLE: Status and Future Ryoma Yamashiro & Imoto Takayuki

431 THE NORMS OF BEHAVIOUR IN SPACE: Our space – Whose rules? Lesley Jane Smith

OUR MISSION STATEMENT The British Interplanetary Society promotes the exploration and use of space for the benefit of humanity, connecting people to create, educate and inspire, and advance knowledge in all aspects of astronautics.

JBIS Vol 71 No.11 November 2018 397 INTRODUCTION

Introduction RiSpace 2018, London

Now under the stewardship of the BIS, the 2018 RiSpace Conference was held in the convivial surroundings of the Hamilton Place Conference Centre in London over three days 30 October–1 November.

ince its inception, the Re-Inventing Space conference has aimed to provide a forum where novel ideas for space systems can be presented and discussed. Fol- lowing its move to the UK in 2014 under the steward- Sship of the BIS, the conference has increasingly at- tracted contributions from around the world, demonstrating the planet-wide interest in, and dependence upon, the space domain.

This tradition was continued at the 2018 RISpace which was held in the Regency surroundings of Hamilton Place in London. The papers in this issue of JBIS are a selection of the highest-ranked submissions that the BIS received following its call for papers in 2018, chosen to provide a balanced reflection of the event itself. They cover a range of novel concepts and show conclusively that advances in technology present satellite system designers with a continuing challenge to Re-Invent Space.

Details of future conferences in this series can be found at http://rispace.org/

Stuart Eves, Chair of Programme Committee

398 Vol 71 No.11 November 2018 JBIS JBIS VOLUME 71 2018 PAGES 399-405

THE MARKET FOR A UK LAUNCHER

VADIM ZAKIROV, ALAN PERERA-WEBB, RICHARD OSBORNE, CONSTANTINE MILYAEV, GERRY WEBB, Commercial Space Technologies Ltd., 67 Shakespeare Road, Hanwell, London W7 1LU, United Kingdom .

Email [email protected]

After five decades since the launch of Black Arrow, the UK government resurrects the national launcher programme. Selection of the new national commercial launcher is the next critical milestone. While the future UK launcher specs from UK spaceports are still vague, Commercial Space Technologies (CST) Limited continually forecasts its general prospects. The forecast is based on data obtained from SpaceTrak database by Seradata Limited. To ensure commercial success, the new UK launcher must secure a significant share of the commercial launch market. CST’s forecast defines the launch market of customers willing to launch their spacecraft (limited to 225 kg by mass) commercially from the UK vertical launch site to Sun-Synchronous Orbit (SSO) and Polar Orbits (PO) during 2020-2030. It is assumed that the UK launcher has to compete with other small launchers for the market. It is demonstrated that success rests on the implementation of a smart launch pricing strategy and a timely entry to the market. The pricing strategy winning the UK launcher from 32% to 61% of the small launcher market is presented and the assessments for the UK launcher sales are given. The key outcome is compared to the earlier UK government, Euroconsult and other forecasts and found to be less a optimistic one. The discrepancy among the CST’s and the other forecast numbers is explained. Further enhancements in the forecast’s accuracy are possible once the UK launcher and spaceport specs are determined.

Keywords: Spacecraft, Launcher, Market

1 INTRODUCTION Lockheed Martin launcher (a variant of Rocket Lab’s Electron); a small launcher under development by British start-up com- On 9 August, 2018 the UK government issued a press release pany Orbex and the Skyrora XL built by yet another British revealing the country’s plans for competing in a high value company, Skyrora 1, 4, 5, 6]. market to launch an estimated 2,000 small satellites by 2030 [1]. To achieve this goal the country plans on building the first ever While the future UK launcher specs from UK spaceports national commercial spaceport located in Sutherland, north- are still vague, Commercial Space Technologies Limited con- ern Scotland. The selected site offers an opportunity to develop tinually forecasts (with what information is at hand) its general a launch complex for small, vertically launched rockets that fly prospects. The forecasts are presented in this paper. northwards [1, 2]. This offers access to sun-synchronous orbit (SSO – an orbit in high commercial demand from the small 2 FORECAST spacecraft community) as well as other high-inclination orbits. The goal of the forecast is to determine the share of the world In July, 2018 £31.5M of funding was announced to support launches the prospective UK launcher may get in the fierce the Sutherland spaceport [3]. This is considered a key part of competition amongst small launcher operators over the decade the Industrial Strategy to ensure the UK prospers in the com- time frame of 2020–2030. mercial space age. It is expected that the spaceport could be worth 400 jobs to local Scottish economy [1]. Commercial ver- The forecast is based on data obtained from SpaceTrak da- tical and horizontal launch demand is estimated worth a po- tabase by Seradata Limited [7]. Only “commercially launched tential £3.8 billion to the UK economy over the next decade spacecraft” from this database are included as it is assumed and will support further growth of the country’s space sector. that this is most reflective of the customer base a UK launcher operating from UK spaceport would be targeting. This study Selection of small launcher for the job is the next logical step defines “commercially launched spacecraft” as spacecraft that for the plan implementation. At present, out of several launch- were launched from a nation that did not build (or have in- er options, the three publicised candidates (two of which have volvement in the development or operation) of the spacecraft. received funding from the UK Space Agency (UKSA)) are: the This data was however carefully scrutinised and there are a few rare exceptions to this rule that are also included in the dataset, such as spacecraft that were originally planned to launch from This paper was first presented at the 16th Reinventing Space a foreign nation, but ended up launching on a national launch- (RiSpace) Conference, London, 2018. er due to peculiarities in the original launch.

JBIS Vol 71 No.11 November 2018 399 VADIM ZAKIROV et al

Fig.1 Launch data and forecasts for small commercial spacecraft Fig.2 Contemporary figures and pessimistic forecasts for commercial going to SSO and PO during 2008-2030. spacecraft to SSO and PO (split by mass categories) during 2008-2030.

Launches and spacecraft for insertion to SSO and PO are of interest because of the UK spaceport accessible inclinations. The spacecraft for insertion into those orbits are limited by mass to 225 kg, which is based on an average of announced prospective launcher payload capacities [4, 5, 6]. Spacecraft are split in four sub-categories by mass: >0-10 kg; >10-50 kg; >50- 100 kg; and >100-225 kg to identify the most expected customer profile. The dates from historic launch logs are extrapolated into the future, up to the year 2030.

The forecast presents two scenarios: pessimistic and optimistic. The pessimistic scenario assumes that small spacecraft Fig.3 Contemporary figures and pessimistic forecasts for mass of customers remain conservative and stick to existing rideshare commercial spacecraft to SSO and PO (split by mass categories) options. The launch market will keep exploiting rideshare of during 2008-2030. existing launchers. Slow, constant growth is assumed for the time period. The optimistic scenario assumes that new commercial launchers will join the market providing affordable launch services. Steady, constant growth of launch numbers is assumed for the time period.

Another significant output of this study is a prediction of the market potential for commercially supported launches over the future decade timeframe.

All data in this paper is complete up to January 2018, when it was extracted for analysis from the Seradata database.

2.1 Spacecraft market Fig.4 Contemporary figures and optimistic forecasts for Statistical data for the small spacecraft launched to SSO and commercial spacecraft to SSO and PO (split by mass categories) Low Earth Polar orbits during 2000-2017 were obtained from during 2008-2030. the database [7]. Later on this dataset was reduced to data starting from 2008, in order to remove the contributions of retired launchers such as Ariane-4, Cosmos-3M, Tsyklon-3 and the Japanese M-V, which could add inaccuracy to the model.

Attempting to assess the target market for any small launch vehicle, in terms of numbers of spacecraft launched (as shown in Fig. 1), can be misleading. The reason for this is that a number of the spacecraft launched in recent years are pico- and na- nosats in the <10kg mass range, most commonly CubeSats of approximately 1–3U in size (the standardised dimensions of a single cube unit being 10x10x10cm). The low mass and mission budgets of these spacecraft means that a bunch of them would have to be clustered together to cover the cost of one rocket launch. Therefore numbers of spacecraft launched (especially Fig.5 Contemporary figures and optimistic forecasts for mass of referring to the 2017 spike in Fig 1.) don’t directly translate to commercial spacecraft to SSO and PO (split by mass categories) number of potential launches. during 2008-2030.

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For more representative analysis spacecraft are split into mass categories and supplemented by total spacecraft mass data in Fig. 2 to Fig. 5.

By comparing the data from Fig 1. to Fig. 5 the following can be noted: • Th e numbers and total mass of small spacecraft launching into SSO and PO have been rising, In both cases this rate has generally increased since 2013, with spacecraft numbers already at their highest in 2017 and total mass in 2017 to inevitably follow suit (overtaking the previous highest peak in 2016); • Th e largest contribution to the commercially launched spacecraft mass budget comes from the mass category: >100 to 225kg; • I n 2017 the contributions to the mass budget of commer- Fig.6 The number of commercial spacecraft looking for launch to cially launched spacecraft of <10kg (primarily CubeSats) SSO and PO (split by mass categories) that could support new small has grown from insignificant to the second largest in re- launchers over 2020-2030. cent years; • Th e contributions of the other spacecraft mass categories to the mass budget are subordinate; • A lthough 189 of nano- spacecraft <10kg have so far been launched in 2017, they represent a rather small mass of 945kg. This, on one hand, requires only one light launcher (such as Vega [8]) to lift all of them up to the orbit, but on the other hand, can employ from 4 to 19 small launchers [4, 5, 6].

The difference between optimistic and pessimistic forecast figures in Figs. 2 and 4 represents the spacecraft (in Fig. 6) available for launch on new launchers joining the market. This future spacecraft market will be driven by increasing -de mand in launching nano-sats. New small launchers entering operations could seed the market and increase the number of nano-sats launched to orbit. Fig.7 Mass of commercial spacecraft for launch to SSO and PO (split by mass categories) by new small launchers during 2020-2030. Fig. 7 shows that nano-satellites (<10 kg) will also dominate by their total mass. Spacecraft in the >100 to 225 kg mass range will take the second mass share launched by new launchers. by just one light launcher, or by 4 to 10 small launchers. The mass share of spacecraft in the >10 to 50 kg mass range will remain rather modest. Although the model predicts no space- Cumulative numbers of commercial spacecraft potentially craft in the >50 to 100 kg mass looking for commercial launch to launch to SSO and PO on new small launchers over 2020- on new launchers, in fact it only means that the share of these 2030 are shown in Figs. 8 and 9. They confirm earlier state- spacecraft will be rather small. ments that: • n ano-size spacecraft will take the largest share of the All spacecraft in Fig. 7 for each year could be lifted to orbit commercial spacecraft market;

Fig.8 Spacecraft for launch to SSO + PO during the decade from Fig.9 Mass of spacecraft for launch to SSO + PO during the decade 2020-2030 split by mass categories. from 2020-2030 split by mass categories.

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• t he second largest share would be the launch of spacecraft in the >100 to 225 kg mass range; • in third would be spacecraft in the >10 to 50 kg mass range; • m eanwhile the share of spacecraft in the >50 to 100 kg mass range remains insignificant.

The forecast demonstrates that spacecraft in the >100 to 225 kg mass range could constitute one of, if not the most significant customer base for UK launcher, because of their notewor- thy mass this could directly translate to numbers of launches. Nano spacecraft, although the most considerable by mass and numbers for UK launcher, would be more of a challenge to manage because of the sheer number of customers involved. Fig.10 Contemporary figures and optimistic forecasts for mass of These customer numbers can however be made more manage- commercial spacecraft to SSO and PO (split by mass categories) able by using CubeSat integrators such as Innovative Solutions during 2008-2030. in Space BV (ISIS) and Astrofein GmbH [9, 10]. TABLE 1 New commercial small spacecraft-dedicated 2.2 Launcher market launchers Launcher Payload Launch cost Reference The forecast for the small launcher launch market is depicted (kg) (US$ million) in Fig.10. It predicts a steady growth of small launcher launch numbers for the foreseeable future, amounting to 76 over the UK launcher - UKLV 225 13.5-4.9 N/A total decade timeframe. Orbital ATK Pegasus 325 40.0 12

The goal of this forecast is to determine a UK launcher pric- Rocket Labs Electron 150 4.9 13-14 ing strategy, leading to successful competition on the interna- Nammo NSLV 50 2.5 15,17 tional stage. Competition from other small launchers has been selected (out of over a hundred small launchers currently un- Virgin Orbit LauncherOne 300 10.0 11,16 der development [11]) based on the developers’ background and in-house competitor analysis. The list of competitors is Vector-R 64 1.5 11 presented in Table 1. VLM-1 150 10.0 11

First, the simplest case of NO competitors in commercial launch market is considered. For this case the UK launch vehicle (UKLV) launch market looks the same as the one in Fig. three launchers. The pricing plan displayed in Table 1 allows 10. Unfortunately, this scenario is unrealistic because the other UKLV to achieve steady growth in the launch market. launcher, the Rocket Labs’ Electron, is already on the market [4, 13, 14]. For this reason, the next comparison includes the Virgin Orbit’s LauncherOne [11] launcher, announced to Electron launcher. The results are shown in Fig. 11. The analy- enter service in first half of 2018 but still grounded as of Octo- sis demonstrates that for successful competition with the Elec- ber 2018, may win a significant share of the market if/when it tron launcher a UKLV does not need to drop its launch price enters service, providing its price is kept as quoted in Table 1. for first 2 years because the market is still unsaturated. This market scenario is illustrated in Fig 13.

Another competitor currently in service is Orbital ATK’s Pe- This September, after about a decade of development, Nam- gasus air launcher. Although Pegasus has a high cost it has built mo’s hybrid rocket motor propulsion was successfully tested in flight heritage since 1990 and is flexible on orbital inclination. sub-orbital flight of Nucleus sounding rocket launched from Figure 12 demonstrates the competition scenario among the Andøya Space Center [17]. Although the further upgrades of

Fig.11 Commercial launch market share for new launchers. UKLV + Fig.12 Commercial launch market share for new launchers. UKLV + 1 competitor. 2 competitors.

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Fig.13 Commercial launch market share for new launchers. UKLV + Fig.14 Commercial launch market share for new launchers. UKLV + 3 competitors. 6 competitors. the sounding rocket to an orbital launcher are planned [17] the development may take a while. Even if the final goal, Nam- mo NSLV [15], is built, it will be too small and expensive to get a noticeable share of the launch market according to our model.

Fig. 15 is a combined plot of UKLV launches extracted from Figs. 10 to 14. The analysis demonstrates that the market share of UKLV launches becomes less sensitive to the number of competitors entering the market when they exceed three.

Fig. 16 presents the pricing plan assumed for the analysis.

Fig. 17 is an illustration of a proposed UKLV price strategy Fig.15 Commercial launch market share for UKLV. that could win it a significant share of the launcher market (re- vealed in Figs. 10 to 14). The launch price may be fairly high at the beginning due to lack of competition and the need to recoup development costs, but this would have to drop as the competition for the launch market intensifies.

3 UKLV MARKET POTENTIAL

Consideration of analysed scenarios in Figs. 11 to 14 suggests the most likely range for the UKLV launches depicted in Fig. 18. The total number of the UKLV launches would be varying from 24 to 46 for 2020-2030 timeframe. The UKLV expected share of the new small launcher launch market is going to be from 32 to 61%.

The total number of spacecraft launched by UKLV during 2020-2030 timeframe is estimated to reach 700. Fig.16 Launch pricing plan used for the analysis.

Fig.17 UKLV launch price plan. Fig.18 UKLV launches range forecast during 2020-2030.

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tive UKLV capacity of 225 kg while the others [18, 19] consider heavier spacecraft.

The CST consultancy team’s assessments are also less optimistic compared to the UK government ones [1]. The difference in the total number of spacecraft to be launched during 2020-2030 (2000 by the UK government vs. 700 by CST) may be explained by higher mass of the spacecraft in the CST’s assessments. The launch sales numbers by CST do not account for all the economy profits by the country’s space sector. CST assessments account for only vertical launches while the UK government accounts the horizontal launches as well.

Because it usually takes about 10 years to develop a new launcher the CST consultancy experts anticipate from 2 to 5 Fig.19 UKLV launch sales range forecast during 2020-2030. year delays in the current plan to conduct the first launch from the UK spaceport in 2020.

The potential launch sales for the UKLV are shown in Fig. 5 CONCLUSIONS 19. The total sales during 2020-2030 timeframe may reach from US$164M to US$336M. While the UK government steps up its plan to secure the country independent access to space, UKLV selection is the next The above numbers may be achieved if the UKLV launch critical milestone. To ensure success, UKLV must secure a sig- price plan in Fig. 17 is respected. nificant share of the commercial SSO and PO launch market during 2020-2030. Success rests on the implementation of a 4 DISCUSSION smart launch pricing strategy and a timely entry to the market. A sample of such a strategy is presented in this paper, as well as The abovementioned forecast is a sample of the analysis and scenarios where UKLV is 2nd up to 7th to market. According expertise delivered by the CST consultancy team to its cus- to the presented forecast and pricing strategy, UKLV may ac- tomers. For the market forecast, CST used a combination of quire from one third to three fifth of the future small launcher statistical data analysis and proprietary models. The most sig- market share during 2020-2030. nificant challenge for this analysis is the present uncertainty in the UKLV option. This is also the reason why in Fig. 13 and At present, the accuracy of this forecast is challenged by Fig, 14, for 2020, UKLV got no share in the launch market, i.e. uncertainties in UKLV option and its technical specifications in the authors’ opinion it is unlikely that the launcher will enter (it is also dependent on competing launchers staying close to service by that time. their announced launch price). Further analysis is required after UKLV selection to enhance the accuracy of the current The CST consultancy team’s assessments regarding space- assessments. craft and launchers may look conservative in comparison to recent Euroconsult and other assessments [18, 19] due to sig- Acknowledgements nificant differences in their goals. While the others [18, 19] consider access to all orbits, this study is focused on the launch The authors are grateful to Seradata Ltd. for the assistance market for spacecraft to be injected into SSO and PO. Moreo- in obtaining all necessary statistics from SpaceTrak – their ver, the payload mass for this study is limited to the prospec- launcher and satellite database.

REFERENCES

1. UK government Press release, Britain competes for the launch of an 8. Vega, The light launcher, Vega overview, ArianeSpace, Ariane Group estimated 2,000 satellites by 2030, https://www.gov.uk/government/ website, http://www.arianespace.com/vehicle/vega/, (accessed news/britain-competes-for-the-launch-of-an-estimated-2000-satellites- 13.10.2018). by-2030 (accessed 12.10.2018). 9. ISIS Website, https://www.isispace.nl/, (accessed 13.10.2018). 2. J. Foust, Five decades after Black Arrow, a reawakening UK launch 10. Astrofein Website, http://www.astrofein.com/, (accessed 13.10.2018). industry aims for bullseye. 8 August 2018, https://spacenews.com/five- decades-after-black-arrow-a-reawakening-uk-launch-industry-aims- 11. P. B. de Selding, Count ‘em: 101 new commercial smallsat-dedicated for-bullseye/, (accessed 12.10.2018). launch vehicles in development, 8 August 2018, https://www. spaceintelreport.com/count-em-101-new-commercial-smallsat- 3. G. McPherson, the Courier.co.uk, UK ministers to hold talks about an dedicated-launch-vehicles-in-development/, (accessed 13.10.2018). aerospace centre for Tayside, July 18 2018, https://www.thecourier.co.uk/ fp/news/politics/scottish-politics/691128/uk-ministers-to-hold-talks- 12. D. Messier, Orbital ATK’s Small Satellite Launch Vehicles Facing about-an-aerospace-centre-for-tayside, (accessed 12.10.2018). Increased Competition. 9 June 2016, http://www.parabolicarc. com/2016/06/09/orbital-atks-small-satellite-launch-vehicles-facing- 4. Electron RocketLab website, https://www.rocketlabusa.com/electron/, increasing-competition/, (accessed 13.10.2018). (accessed 12.10.2018). 13. A Boyle, Rocket Lab sends Electron rocket into orbit, marking milestone 5. The Orbex company website, https://orbex.space/, (accessed 18.08.18). for Moon mission. 20 January 2018, https://www.geekwire.com/2018/ 6. Skyrora website, Orbital launch vehicle, Skyrora XL, https://www. rocket-lab-sends-electron-rocket-orbit-marking-milestone-moon- skyrora.com/skyroraxl, (accessed 12.10.2018). mission/, (accessed 13.10.2018). 7. Seradata website, The SpaceTrak, https://www.seradata.com/products/ 14. C. Gebhardt, Rocket Lab’s Electron conducts inaugural flight from New spacetrak/ , (accessed 12.10.2018). Zealand. 24 May 2017, https://www.nasaspaceflight.com/2017/05/

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rocket-labs-electron-inaugural-flight-new-zealand/, (accessed 18. D. Werner, Government agencies welcome small rockets with contracts, 13.10.2018). awards and reduced red tape, 2015, 7 August 2018, https://spacenews. 15. A. van Kleef, B.A. Oving, C.J. Verberne, B. Haemmerli, M. Kuhn, com/government-agencies-welcome-small-rockets-with-contracts- I. Müller, I. Petkov, Innovative Small Launcher. 13th Reinventing awards-and-reduced-red-tape/ , (accessed 14.10.2018). Space Conference, 11 November 2015, Oxford, UK, http://www. 19. T. Wekerle, J. B. P.Filho, L. E. V. L. da Costa, L. G. Trabasso, Status and small-launcher.eu/wp-content/uploads/BIS-RS-2015-16.Presentation. Trends of Smallsats and Their Launch Vehicles – An Up-to-date Review, VanKleef.pdf, (accessed 13.10.2018). July/September 2017, http://www.scielo.br/scielo.php?script=sci_ 16. Virgin Orbit website, https://virginorbit.com/, (accessed 13.10.2018). arttext&pid=S2175-91462017000300269 , (accessed 25.08.18); see also the same article in J. Aerosp. Technol. Manag. vol.9 no.3 São José 17. ESA website, Norway takes the lead in hybrid propulsion, 27 September dos Campos July/Sept. 2017, http://dx.doi.org/10.5028/jatm.v9i3.853, 2018, https://www.esa.int/Our_Activities/Space_Transportation/ (accessed 14.10.2018) Norway_takes_the_lead_in_hybrid_propulsion, (accessed 13.10.2018).

Received 8 January 2019 Approved 10 January 2019

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THE RemoveDEBRIS SPACE HARPOON

ALEXANDER HALL Airbus Defence and Space Gunnels Wood Rd, Stevenage, SG1 2AS, United Kingdom email [email protected]

The European Commission FP7 RemoveDEBRIS mission, successfully deployed from the ISS on the 20th June 2018, is the world’s first mission to demonstrate Active Debris Removal (ADR) technologies in orbit. The experiments include a net and a harpoon, as well as other technologies such as visual based navigation and a drag sail to de-orbit the spacecraft. The tethered Harpoon experiment, developed at Airbus in Stevenage, is due to run in February 2019 and will be captured on video with onboard platform cameras. This data will help understand the mechanics of firing a harpoon in space, as well as raise the TRL of the technology for future missions. This harpoon has progressed from the first concept in 2013 through to the flight version used today; it has proved accurate enough for use to capture debris and is currently undergoing final preparations in orbit for activation.

Keywords: Airbus, Space, Debris, Harpoon, RemoveDEBRIS

1 INTRODUCTION

The RemoveDEBRIS mission is a low cost Active Debris Re- moval (ADR) demonstrator that is co-funded by the Europe- an Commission FP7 programme; it is the first ADR mission to demonstrate these key technologies in orbit for capturing space debris. The satellite was launched to the International Space Station (ISS) in April 2018 on a SpaceX Falcon 9 re-supply mission CRS-14, it was then released from the ISS via the NanoRacks Kaber system into an orbit of around 400 km.

The experiments aboard RemoveDEBRIS include a net, a visual based navigation LIDAR, a harpoon and a drag sail; each with the aim to test space debris removal techniques. Prelim- inary results from the mission can be found in other papers [1][2]. This paper outlines the harpoon which was developed by Airbus Defence and Space in Stevenage; the harpoon itself is housed within the Harpoon Target Assembly (HTA) chassis Fig.1 Top Left: inside the Harpoon target assembly (HTA); Bottom which was developed by Surrey Space Centre (SSC). Left: CAD model of HTA with target partially deployed. Right: HTA with CPRF boom fully deployed during fire testing (HTA at 2 REMOVEDEBRIS HARPOON PAYLOAD bottom, target plate at top).

2.1 Harpoon Experiment Overview as well as a partially deployed configuration and the full target The RemoveDEBRIS harpoon is designed to attach itself to de- deployment. bris and provide a tethered link for a chaser satellite to tow the debris out of orbit. The harpoon is composed of three main Several harpoon concepts were analysed before the final de- elements; Deployer, Projectile and Tether. The experiment will sign was chosen, these can be seen in Figure 2. The final chosen deploy a piece of 8 mm aluminium honeycomb panel (similar projectile is manufactured from Titanium and conceals two to that used by satellites) to a distance of 1.5 m, the harpoon barbs beneath a sliding sleeve seen in Figure 3, once the pro- will then be fired at the target and cameras on the satellite plat- jectile has fully penetrated the target the sleeve is pushed back form will record the sequence with high frame rates. Figure 1 and the barbs are deployed. These provide a permanent fixing shows the FM HTA stowed and ready for platform integration, point to the target. A tether, made from Dyneema, is attached to the projectile on a sliding shaft. The tether has to be located at the front of the harpoon during firing to fully fit it into the This paper was presented at the 16th Reinventing Space Conference, deployer, it was found that if the tether remained attached to London, 30 October–1 November 2018. the front of the projectile it caused instability during flight. A

406 Vol 71 No.11 November 2018 JBIS THE RemoveDEBRIS SPACE HARPOON

Fig.2 Harpoon evolution.

Fig.5 Harpoon impact test.

a motor and guiding mechanism. The system has sensors to determine the length of the boom and also provides the ability to retract if necessary. This however is not used for nominal operations in the RemoveDEBRIS mission.

Fig.3 Harpoon barbs deployed. 3 TESTING

3.1 Firing Tests sliding mechanism was therefore introduced to bring the tether to the rear of the harpoon and allow stable flight. Multiple vertical firing tests were performed to characterise the performance of the harpoon, an EQM model was used for The harpoon is fired by a piston propelled by gas. The point most of the testing as repeated firings caused wear on the firing at which the piston releases is determined by a tear-pin within mechanism. Vertical tests minimised the effect of gravity on the deployer casing, this restrains the piston until the pressure the accuracy of the projectile but did affect its speed. A reduced is great enough to break the pin. This method allows the firing firing test campaign was performed on the flight harpoon and speed to be modified by varying the pin size but doesn’t allow showed that the harpoon was accurate to +/- 10 mm to the cen- the speed to be varied whilst in orbit, other methods were de- tre of the target over 1.5 m, this included the size of the har- veloped to vary the speed on-the-fly but incured a considerable poon tip. The target on the HTA is 100 mm x 100 mm which mass increase. The RemoveDEBRIS harpoon pin is sized to fire provides plenty of margin for variations of the target position the projectile at 20 m/s. in space compared to on ground. A harpoon impact test can be seen in Figure 5, the flight harpoon firing accuracy can be seen The gas used to propel the projectile is produced by two in Figure 6. TNO cold gas generators (CGG), one prime and one redundant. They produce gas once an initiator is activated inside the CGG. The fully assembled harpoon, along with the deployer, projectile, CGGs and door holding frangibolt, can be seen in Figure 4.

The harpoon deployment system is housed within the HTA chassis which was manufactured by SSC, this also provides the target for the harpoon to capture. An Oxford Space Sytems (OSS) Astrotube boom is used to deploy the target plate to 1.5 m, the OSS boom uses rolled carbon fibre and is deployed with

Fig.4 FM harpoon ready for HTA integration. Fig.6 Harpoon tests firing accuracy.

JBIS Vol 71 No.11 November 2018 407 ALEXANDER HALL

Fig.7 Left: RemoveDEBRIS observed from the ISS shortly after deployment, Right: Close up of platform.

3.2 Vibration and TVAC Testing moveDEBRIS satellite was successfully launched to the ISS on the 2 April 2018 in a SpaceX Dragon capsule; it was then un- The payload was subjected to environmental tests, representing packed and deployed from the ISS on the 20 June. It is currently the conditions it would face during launch and whilst in or- the largest satellite to be deployed from the ISS, Figure 7 shows bit. The most strenuous of these were the vibration tests which images of the satellite and the HTA can just be seen in the top prove the payload can survive launch to the ISS. However, as left of the satellite. the satellite is launched within a SpaceX Dragon capsule, a human rated capsule, the load requirements were relatively small. So far RemoveDEBRIS has successfully deployed the Net Thermal-Vacuum cycling was performed after vibration. This which captured a spinning cubesat in September. The next ex- test sequence was used as micro-fractures caused by vibration periment to run will be the visual based navigation, followed by will propagate through the structure during thermal cycling, the harpoon in February 2019. The final experiment will be the showing failures not found if TVAC was tested first. Follow- drag sail which will pull the satellite out of orbit. ing environmental tests functionality testing was performed to verify the integrity of the payload. 5 CONCLUSION

3.3 Final Assembly The RemoveDEBRIS harpoon has been developed over 5 years and is now in low Earth orbit. The harpoon is a scalable solu- Once the harpoon accuracy was characterised, the harpoon tion to capturing space debris and is tailored for capturing larg- was integrated into the HTA chassis. A laser was fit to the tip er objects such as malfunctioning satellites or rocket casings. of the harpoon to align it with the target. The whole HTA was It is also capable of capturing tumbling debris which would be attached to a frame and deployed vertically down, by remov- difficult with other grappling method. ing the target the end of the boom could hold its own weight and was in a representative position as to where it would be in During testing it was found that the harpoon design was space. accurate to +/- 10 mm to the centre of the target over 1.5 m. Environmental testing and subsequent functional tests showed 4 LAUNCH it was capable of surviving launch conditions.

The harpoon payload was delivered to Surrey Satellite Technol- In February 2019 the harpoon will be fired at an 8 mm al- ogy Ltd (SSTL) in 2017 to be integrated with the satellite plat- uminium honeycomb target and the test will be recorded by form. Once all payloads were attached the platform underwent high frame rate cameras on-board the satellite. This will be the system end-to-end testing and environmental testing before it first time a harpoon is used in space to capture debris and will was packed in a clam shell for delivery to NanoRacks. The Re- be an insight into the future of ADR technologies.

REFERENCES

1 B. Taylor, G. S. Aglietti. et al. October “RemoveDebris Preliminary 2 J. L. Foreshaw., G. S. Aglietti et al. “RemoveDEBRIS: an in-orbit active Mission Results”, 69th International Astronautical Congress (IAC), debris removal demonstration mission”, Acta Astronautica. 127 448–463 Bremen, Germany. 2018. 2016.

Received 4 December 2018 Approved 13 December 2018

408 Vol 71 No.11 November 2018 JBIS -

Have you got what it takes?

After two years spent successfully steeringJBIS towards its new look, Editor Roger Longstaff is moving on to fresh challenges. The Society is now looking for someone to replace him. This is a part- time position, typically taking two days a week, that would suit someone who is either in part time employment, self-employed or retired but still takes a keen interest in the field of astronautics, and who has a background in related academia, astronautics or the space industry itself. Administrative help will be provided and the position attracts remuneration for each issue published. If you think you might fit the bill, please contact Executive Secretary Gill Norman at [email protected] for more details.

JBIS Vol 71 No.11 November 2018 409 JBIS VOLUME 71 2018 PAGES 410-415

RAPID CONSTELLATION DEPLOYMENT from the UK

CHRISTOPHER LOGHRY1, MARISSA STENDER2, 1Moog Inc, 21339 Nordhoff St, Chatsworth, CA 91311, USA; 2Moog Inc, 2581 Leghorn St, Mountain View, CA 94043, USA email [email protected]

The growing number of small launch vehicles, including ones being planned to be launched from the United Kingdom, will lower the cost of space access in the coming years but challenges remain in utilizing these for small payloads particularly Cubesat. Many of these challenges can be met through the use of a propulsive rideshare adapter or Small Launch Orbital Maneuvering Vehicle (SL-OMV). The SL-OMV is a low mass and low cost propulsive adapter that can be used to distribute Cubesat payloads with different orbital parameters than the primary payload or each Cubesat with a different orbital destination. This is important for both Rideshare launches but single mission dedicated cluster launches that are useful across commercial, civil, and military space applications. The SL-OMV is designed specifically for these small launchers including vehicles as part of the UK Space Agency (UKSA) Spaceflight Programme initiative to have domestic launch capability by the 2020s. The SL-OMV can be used on either vertical or horizontal launch configurations for vehicles having 150 kg capability or greater. As part of the program a unit will be developed and produced in the UK for the inaugural launch of the vertical launch option. This will demonstrate that capability in addition to placing six different CubeSats in various orbits. As the SL-OMV is adaptable to a wide-range of launch applications this can be useful for US and European launches as new spaceports come online in particular in the United Kingdom.

Keywords: CubeSats, Constellation, OMV, Propulsive Adapter, Rideshare, UKSA

1 INTRODUCTION the Small Launch Vehicle (SLV) market and greatly enhances the utility of low cost access to LEO. The SL-OMV and poten- The Moog Small Launch Orbital Maneuvering Vehicle (SL- tial future variants can allow for easier utilization of space and OMV) is a propulsive small satellite launch adapter designed access beyond LEO. for use on small launch vehicles, particularly those of the 300 kg to LEO class. The SL-OMV provides the ability to deploy Recent funding from the UKSA’s Spaceflight Programme payloads at different orbits which is critical for payloads from initiative provides an opportunity for multiple launch vehicles different customers (i.e. rideshare) or payloads from the same and launch sites to be developed reading the UK for domestic customer that require constellation phasing or even a combina- launch beginning in the 2020s. This is seen as a key compliment tion of the two. This furthers the investment made in the small to the existing market and capabilities within the UK and pro- launch vehicles themselves and allows for mixed manifest with vides the last step to having an “end to end” space economy in a variety of customers or end users. the UK.

The SL-OMV is a turnkey solution that can be leveraged on 2 HOW SMALL LAUNCH VEHICLES ARE CHANGING THE several different launch vehicles, both existing, and in develop- LAUNCH MARKET ment. The system is designed to be mass and cost efficient and includes a Green Propulsion system enabling launches from With smaller launch vehicles in development, it is finally an Spaceports and other facilities that do not have hazardous pay- option for a small satellite, or constellation of CubeSats, to load processing capabilities. Moog has been developing the SL- purchase the majority of a launch vehicle’s capacity and bene- OMV based on market input from across the global space com- fit from greater influence over the schedule, drop-off location, munity while leveraging hardware and technologies that are and even the launch environments as compared to a “hitchhik- being developed for this market as well. The design reflects this er” payload. market input in performance, capabilities, modularity, and cost. Many of the new, commercial spacecraft companies require Moog sees the SL-OMV as a complimentary technology to anywhere from one to hundreds of spacecraft to meet their data collection goals and often require constellation refresh- es at a regular rate. Unlike rideshare opportunities, which are This paper was presented at the 16th Reinventing Space Conference, sporadic and vary widely in cost per kilogram, a small launch London, 30 October–1 November 2018. vehicle can offer dedicated launch at a regular cadence.

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2.1 More Options for Small Satellites

The rapid growth of small satellite industry has led to an equally rapid growth in the Small Launch Vehicle market. SLV systems in development range from 4-5 kg capacity up to 500 kg capacity. There are over 50 systems in development, with at least 34 of them still active in the last 2 years, in addition to existing systems like Pegasus XL and Minotaur [1,2 3]. The lower capacity SLV are aimed towards Cubesat-class payloads where one launch may carry only one or two payloads. Many of the vehicles have a 200 kg to 400 kg throw mass and even with an “ESPA Class” payload of 180 kg there is additional capacity for small payloads such as Cubesats. This capacity also allows for a cluster of small payloads which enables an entire constellation to be launched at once. Some of these launch vehicles are planning for launch rates of 20-30 per year with Virgin Orbit’s LauncherOne [4] or even up to once a week with Rocket Lab’s Electron [5].

This capability is important as estimates show up to 2600 small payloads between 1-50 kg will be launched from 2018- 2022 [6]. Cubesats are transitioning from technology demonstration platforms to critical assets as part of commercial business models and even military applications. From 2013-2017 38% of the launched small payloads were for scientific or technology purposes whereas the forecast from 2018-2022 is showing over 70% will be for Earth Observation, Remote Sensing, and Communications. The largest class of these payloads will Fig.1 SL-OMV in a Ø1067 mm Diameter Payload Fairing. be in the 3U or 6U range [6].

2.2 Role of the SL-OMV The UKSA announced several other grants and initiatives related to launch including multiple other launch sites and Although the luxury of being the primary customer on a launch vehicles [8]. The SL-OMV is designed to compatible launch affords many privileges, small launch customers com- with several launch vehicles including both horizontal and peting in the new space market cannot afford to waste launch vertical launch. Currently the SL-OMV is capable of being capacity. New ways to optimize the capability of an SLV is launched on each of the options as part of this grant funding required. ensuring maximum utility of the SL-OMV. Additionally the SL-OMV design is being assessed to maximize the supply chain For this reason, the SL-OMV has been designed to be com- and domestic capabilities from within the UK making this a patible with the new class of Small Launchers and existing truly indigenous capability that can support both UK and com- small launch systems such as the Pegasus and the Minotaur mercial missions. family. Only the diameter of the adapter and propulsion tank attachments need to change in order to interface with LVs from 4 SL-OMV CAPABILITY SUMMARY 610 mm to 986 mm. The key characteristics of the SL-OMV system are a light- For SLVs such as Rocket Lab’s Electron, the SL-OMV would weight, composite cylinder adapter, a composite propellant be the primary passenger (Fig. 1). Whereas on vehicles, such tank, a fixed solar array and the flexibility to mount multiple as Virgin Orbit LauncherOne, Orbex Prime, Firefly Alpha, kinds of payload adapters. Northrop Grumman Innovation Systems Pegasus XL or a Mi- notaur vehicle the SL-OMV could be placed under a primary The key vehicle properties are highlighted here, with further passenger in a similar fashion to the ESPA ring on the EELV- detail on the vehicle’s subsystems and mission scenarios pre- class rockets. A stacked configuration is also possible within sented later in the paper. larger fairings such as the Minotaur I. 4.1 Dimensions and Mass 3 SL-OMV AND LAUNCH FROM THE UNITED KINGDOM The SL-OMV structure has a nominal height of 610 mm and Moog in partnership with Lockheed Martin and the UKSA, a diameter of 610 mm as well. With six payload dispensers at- announced the first scheduled launch of its SL-OMV on July tached, the vehicle measures ~1020 mm across. The nominal 17, 2018 as part of the UK Spaceflight Programme [7]. The SL- diameter can be expanded to 986 mm to accommodate a vari- OMV’s maiden voyage will launch in the early 2020s from the ety of vehicles and standard separation systems. UK’s first commercial spaceport at the Sutherland site in Mel- ness, Scotland. Lockheed Martin, the prime contractor for UK- The limited mass of the new small launch vehicles requires SA’s Spaceflight Programme, is the launch provider responsible that the SL-OMV be as light as possible. In total, the wet mass for payload integration and launch service. The Moog opera- of the vehicle is approximately 67 kg, including approximate- tion in Reading, UK will perform final integration and test of ly 17 kg of green propellant. With a full complement of 6x6U the SL-OMV while leveraging the local supply chain. CubeSats and dispensers (~90 kg), the vehicle is approximately

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154 kg (varies based on payload and dispenser masses). Fig. 2 shows how these values breakdown over the whole system.

4.2 Capabilities

The SL-OMV is capable of on-orbit operations nominally up to 12 months (but it can be more), allowing the vehicle to deploy satellites over longer periods of time than a typical upper stage of just a few hours or stay on-orbit to act as a hosted payload platform once payloads have been deployed.

The propulsive capability of the SL-OMV varies based on the satellite and dispenser mass. Example configurations and corresponding delta-V values are shown in Table 1. Fig.2 SL-OMV Mass Breakdown. 4.3 Payload Interface Options

The composite structure is capable of mounting a range of TABLE 1 SL-OMV Delta-V Capability small satellite dispensers and adapters. On top of the ring, a Payload Delta-V payload can mount directly to the “24 inch” diameter SL-OMV, or an adapter cone or plate can be used to mount a 300 mm or 12 x 3U CubeSats 260 m/s 381 mm Motorized Light Band (MLB). 6 x 6U CubeSats 115 m/s Alternative configurations are possible where payloads are mounted on top of the SL-OMV and the solar array is moved to the side as shown in Fig. 3. This allows for larger 12U form factors while staying within small payload fairings or 8x12Us oped by Tyvak under SBIR Phase II funding (Fig. 4). The avion- can be installed with MLBs. ics suite performs: Vehicle State Estimation, Vehicle Guidance Navigation and Control, Redundant GPS, Redundant MEMS Many different payload configurations can be achieved by IMU and an S-Band transmitter. This SLV avionics suite lev- swapping dispensers. The Planetary Systems Corp. (PSC) 6U erages Tyvak’s extensive spacecraft flight heritage avionics but adapter [9] was shown in Fig. 1 and was incorporated into the modified for use on the same Small Launchers that the SL- initial design due to the need to bound the design for a rea- OMV would be deployed on. This synergy creates a common sonable maximum mass. Other ‘skeletonized’ designs like the ‘ecosystem’ across the launch vehicle and spacecraft industries Tyvak RailPods with extremely lightweight materials are used that can be used to reduce overall system implementation to optimize the payload carrying capability and a likely better costs. This also creates an environment for the different specific option [10]. hardware solutions that can have applications across a whole mission (e.g. launch vehicle, SL OMV, and spacecraft). All of 5 SL-OMV SYSTEM DETAILS this reduces cost and can reduce the amount of development time and risk for new systems. The SL-OMV system is very similar to a small satellite, but with a much greater emphasis on low cost and low mass hardware, 5.2 Power often derived from the growing expertise within the Cube- Sat community. The overall system must be cost competitive A further advantage of utilized MicroSat avionics and sensors enough relative to the cost of dedicated launch vehicles and is the low power draw of the components. This feeds directly mass efficient due to the 200-400 kg to LEO capacity of the into the sizing of the solar array and the ability to fit it within target launch vehicles. the small area available on the top of the SL OMV. This fixed array simplifies the overall power system design by eliminating 5.1 C&DH and EPS Systems the need for deployment systems or actuators (see Fig. 5). In the event a fixed array is not enough power for the required The core technology to for this vehicle is the platform inde- mission a deployable array option based on deployable Cube- pendent small launch vehicle avionics currently being devel- Sat solar arrays is available.

Fig.3 SL-OMV with Payloads on Top (left) and 12Us on the Side Fig.4 Tyvak’s Micro-Avionics Multi-Purpose Platform (right). (MicroAMPP).

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kg. The carbon fiber composite system provides other advan- tages such as a more stable thermal environment and micro- meteorite protection and is manufactured more easily from existing capabilities. The additively manufactured variant may be applicable for passive adapter applications (similar to ESPA).

5.5 Propulsion

A major element of the SL-OMV is its on-board propulsion system. Moog is leveraging a flexible propulsion system designed to be compatible with Hydrazine but baselining the green propellant LMP-103S that has been successfully deployed on the Planet Skysat-C constellation. The ability to use a green propellant as part of the baseline design is a critical enabler to this particular system as many of the target launch vehicles will not include Hydrazine processing as part of their standard service. A simple blowdown system with six 1N engines are used to provide 3 Degrees of Freedom (3DOF) control (see Fig. 6). Fig.5 Top View of SL-OMV. Launch vehicle and spacecraft propulsion system contain- The MicroAMPP avionics includes a battery that is sized to ing hydrazine or other hazardous propellants require addition- accommodate the rest of the SL-OMV. The battery is sized to al design, test, qualification, and operational considerations provide enough energy to perform a typical transfer maneuver to ensure hardware and personnel safety. In many cases this on battery power alone (e.g. 100 km orbit change). Upon com- drives costlier system designs. Green propellants allow for a pletion of the burn, the SL OMV would orient itself sun-point- “single-fault tolerant” designs that can reduce overall life cycle ing and recharge the battery. This process can be repeated as costs including reduced component count [11], system testing, needed for subsequent burns and maneuvers. and lower cost manufacturing techniques such as flared tub- ing and mechanical fittings instead of all-welded systems that The MicroSat avionics are designed to have very low power have 100% radiographic (X-Ray) inspection requirements for draw and the largest power draw during operation is that of the hypergolic systems [12]. thrusters while firing. 6 SL-OMV RIDESHARE MISSION EXAMPLES 5.3 Guidance, Navigation, and Control The mission examples presented here are relevant to both The baseline SL-OMV GNC subsystem is designed for on-orbit commercial and scientific pursuits within the quickly evolving maneuvering and payload deployments. The basic missions of small satellite community and SLV market. For this reason, the this CubeSat tug can be completed with a minimally complex SL-OMV missions vary from one day to over 12 months. system consisting of thrusters, a startracker and an IMU. 6.1 CubeSat Constellation Deployment For extended missions with fine pointing requirements, reaction wheels can be added and a second star tracker may be One of the primary uses envisioned for the SL-OMV is the de- included depending on the agility required. ployment of CubeSats into a constellation. Depending on the mission goals, the SL-OMV is capable of dispersing smallsats 5.4 Structure around a single orbital plane or delivering them to entirely new orbits including different altitudes or varying Right Angle of Moog is developing the composite cylinder structure that will the Ascending Node (RAAN). support the SL-OMV payload elements and any additional payloads located on top of the adapter. The composite structure will leverage a previous Moog program which developed a composite dual-launch adapter for the Minotaur IV launch vehicles called CASPAR. Moog has also assessed using Additive Manufacturing as an option for rapid configuration changes.

Moog and Tyvak have engaged with various SLV providers and USG satellite users to understand the range of potential primary payload masses and launch vehicle environments. The structure will use traditional carbon fiber elements and manufacturing techniques, reducing development risk and costs. These structures are commonplace in space applications including many launch vehicle structures being made from similar materials. This reduces both the mass and perceived risk to SLV providers.

An additively manufactured variant using Titanium was developed and analyzed and was approximately 30%-40% heavier than the carbon fiber variant, but in absolute terms this was ~3 Fig.6 Bottom View of SL-OMV (including six thruster locations).

JBIS Vol 71 No.11 November 2018 413 CHRISTOPHER LOGHRY, MARISSA STENDER

For a commercial constellation with multiple launches, the would allow for the same engines to be used for attitude con- SL-OMV can be used to quickly deliver the small spacecraft trol but a higher thrust and higher performance “main engine” payloads around a single orbital plane. This scenario is par- could be used for higher energy transfers. This is a potential ticularly relevant for a rideshare mission in which an SL OMV, growth path for the SL-OMV as uses cases for CubeSats and holding up to 12x3U satellites, could be launched below the Small Launch Vehicles move beyond LEO. primary small satellite. 6.3 On-Orbit Sentinel For a science mission that requires satellites in multiple planes at the same altitude and inclination (eg. multiple One application for a SL-OMV is existing on-orbit assets that RAANs), a stack of SL-OMVs can be utilized. can be called upon and activated when needed. This could be as part of a dedicated SL-OMV with a full complement of pay- Examples of both mission types are presented below. loads or potentially left onboard after some of the payloads have been deployed. Deployment around a single orbital plane - In this scenario the goal is to space multiple spacecraft around a single plane to This scenario has civil, commercial, and military applica- enable full constellation use within a few days of launch. tions. Some science missions may utilize payloads that are in low orbits and decay rapidly or at unpredictable intervals. Here Constellation Description: an SL-OMV could remain on orbit at a higher altitude and re- • Number of Spacecraft: 6 plenish payloads immediately as their predecessors deorbit. • Altitude: 450 km This could be useful for deploying payloads rapidly to observe • Spacing between satellites: 60 degrees transient events such as solar wind or coronal mass ejection (CME) effects on the upper atmosphere or in preparation for The role of the SL-OMV is to change the true anomaly by a large weather event with CubeSats that are equipped with maneuvering to a lower orbit and then returning to the goal or- weather payloads. bit. Fuel use is minimized by spending approximately one day at the lower orbit. The entire process can be sped up by making Commercial users may find having on-orbit spares useful. a larger change in altitude and using more propellant. This is important when continuous service is needed or po- An example vehicle configuration would consist of 6x6U tential constellation augmentation needed for increased data CubeSat dispensers attached to the adapter and all CubeSats gathering or transmission. Spacecraft insurance companies would be deployed evenly around a plane (i.e. 60° spacing in may have on-orbit assets ready to perform inspections as part their orbit) in less than 10 days. of an insurance claim. The SL-OMV could be used to maneuver towards a failed asset and deploy payloads. Several cubesats 6.2 Small Launch Vehicle Insertion Stage could be used for inspection of a spacecraft using optical or potential radar inspection. The SL-OMV could be configured The SL-OMV can provide an insertion stage capability for the to have inspection hardware as well. small launchers (both new and existing) as well increasing the capability of the vehicles using a “bolt on” stage. The SL-OMV Military applications could include on-orbit assets ready to is designed with the future small launch vehicle systems in reconstitute (or create) a constellation in the event a primary mind including using low cost platform avionics and compos- asset is destroyed or otherwise rendered inoperable. This is ites, composites for low mass structures, and green propellant important in overhead reconnaissance and tactical commu- for spaceport operations and especially for European launches nications. A SL-OMV could be positioned near critical LEO where REACH legislation may limit traditional propellants like assets that it or its payloads could be “sacrificed” in the event Hydrazine. a ground-based or spaced-based attack. The SL-OMV could potentially deploy payloads configured as countermeasures In this scenario the SL-OMV acts as essentially a liquid up- (similar to flares and chaff used on airplanes) as part of a -de per stage with unlimited restart capability. This can be useful fense barrier. SL OMVs could be treated similar to the support when the SLV upper stage is solid propellant based or a liquid ships in a carrier battle group. The flexibility of different pay- upper stage that has a limited number of restart capabilities. loads that can be supported using a common platform reduces The SL-OMV can be used when the payload(s) requires finer the cost to implement a system. The potentially large number orbit insertion accuracy than the launch vehicle can provide of units that could be on-orbit provide a disaggregated system or potentially perform a transfer to a higher altitude than the and potentially an overwhelming number of small targets for launch vehicle is designed for. an enemy.

The Minotaur and Pegasus XL launch vehicles use a simi- 6.4 Deployment to Multiple Planes lar concept called the Hydrazine Auxiliary Propulsion System (HAPS) [13] that is proprietary to those launch vehicles. The Satellite constellations that require frequent revisits, such as SL-OMV is widely available and leveraging a green propellant earth observation systems or weather monitoring sensors, of- allows for a wider range of applications. ten desire to have satellites placed in multiple orbital planes at the same altitude and inclination. The challenge with these This example mission is also relevant for launch configura- types of constellations is the amount of time and propellant tions that need to drop satellites at multiple altitudes without that is required to achieve the correct orbital positioning. And purchasing (and waiting for) a second launch. the smaller the satellites become, the harder it is to physically fit the propulsion capability within the limited volume of a small The SL-OMV can benefit from higher performance green satellite. Additionally, small satellites are less likely to be able propellant blends including leveraging the existing LMP 103S to afford an individual launch, which is the way much larger propellant and High Test Peroxide (HTP) as a fuel [14]. This constellations (e.g. GPS) reach many different orbits.

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The solution to reaching multiple orbital planes from a sin- (both altitude and inclination) to achieve the differential preces- gle launch is by utilizing differential orbital precession rates to sion with respect to the reference orbit. One SL-OMV (#1) will achieve the desired RAAN spacing between multiple orbits. maneuver to a lower orbit and the other SL-OMV will maneu- An example of this type of constellation is the MIT TROPICS ver to a higher orbit. This allows one OMV to move to a RAAN mission [15]. The mission is described as: “The Time-Resolved -120 degrees to the reference orbit while the other moves to an Observations of Precipitation structure and storm Intensity orbit with a RAAN +120 degrees to the reference orbit, respec- with a Constellation of Smallsats (TROPICS) mission will pro- tively. Both will loiter approximately seven months at their new vide rapid-refresh microwave measurements over the tropics orbit before returning to the reference orbit and deploying four that can be used to observe the thermodynamics of the trop- Cubesats each. Compared to the cost and time of waiting for osphere and precipitation structure for storm systems at the multiple launches, this scenario shows that an SL-OMV is ca- mesoscale and synoptic scale over the entire storm lifecycle. pable of positioning a global constellation within seven months. TROPICS comprises a constellation of CubeSats in three low- Earth orbital planes.” 7 CONCLUSIONS

The parameters of the MIT Tropics mission are listed below The launch vehicle industry is slowly catching up to the evolv- for use in this example scenario: ing market of small satellites. SLVs are on the verge of providing regular access to space for spacecraft in the 200-400 kg class • Number of Spacecraft: 12 of vehicle. Rather than take a breath and wait years or even • Altitude: 550 km decades for the launch market to meet the specific needs of • Inclination: 30 degrees CubeSats, an interim solution can be created in parallel with • Number of Planes: 3 the emerging SLV market. Moog’s proposed solution, the SL- • Spacing between Planes: 120 degrees OMV, can enable on-orbit placement of large commercial and government constellations that have recently become financial- To deliver satellites to three separate planes from a single ly viable due to the small satellite revolution. launch, two SL-OMVs are required. SL-OMV #1 has two roles: drop-off four satellites in the initial (reference) orbit and then The recent UKSA Spaceflight Programme grant funding will complete multiple maneuvers over a seven month timeframe further accelerate this and make it possible for the SL OMV to prior to dropping off four more satellites in a second plane. SL- be demonstrated in orbit having been launched from the UK. OMV #2 has only four satellites to drop-off in the third orbital As the SL-OMV is compatible with multiple launch options plane after a mission lifetime of approximately seven months. that could originate from the UK there are several opportuni- Each SL-OMV is required to make two large orbit changes ties to leverage this innovative technology.

REFERENCES

1. C. Niederstrasser, W. Frick. 2015. “Small Launch Vehicles – A 8. https://www.gov.uk/government/news/lockheed-martin-and-orbex-to- 2015 State of the Industry Survey,” Proceedings of the AIAA/ launch-uk-into-new-space-age USU Conference on Small Satellites Technical Session II: Launch, 9. Planetary Systems Corporation (PSC) Containerized Satellite Dispenser SSC15-II-7. http://digitalcommons.usu.edu/cgi/viewcontent. (CSD) data sheet web link: http://www.planetarysystemscorp.com/web/ cgi?article=3176&context=smallsat wp-content/uploads/2016/08/2002337D-CSD-Data-Sheet.pdf 2. Parabolic Arc news website: http://www.parabolicarc.com/2016/10/03/ 10. Tyvak RailPOD Users Guide, http://www.tyvak.com/wp-content/ plethora-small-sat-launchers/ uploads/2017/03/RailPOD_MkII_UsersGuide_TK-RPUG-Rev2_1.pdf 3. C. Niederstrasser. 2018. “Small Launch Vehicles – A 2018 11. R. Masse, et al. 2015. “Enabling High Performance Green Propulsion State of the Industry Survey,” Proceedings of the AIAA/USU for SmallSats,” Proceedings of the AIAA/USU Conference on Conference on Small Satellites Technical Session IX: Space Access, Small Satellites Technical Session XI: Advanced Technologies III, SSC18-IX-01. https://digitalcommons.usu.edu/cgi/viewcontent. SSC15-XI-6. http://digitalcommons.usu.edu/cgi/viewcontent. cgi?article=4118&context=smallsat cgi?article=3240&context=smallsat 4. A.C. Charania, et al. 2016. “LauncherOne: Virgin Galactic’s Dedicated 12. Air Force Space Command Manual (AFPSCMAN) 91 710 Volume 3, Launch Vehicle for Small Satellites,” Proceedings of the AIAA/ Chapter 12 USU Conference on Small Satellites Technical Session II: Launch SSC16-II-2. http://digitalcommons.usu.edu/cgi/viewcontent. 13. Orbital ATK Pegasus Users Guide web link: https://www.orbitalatk. cgi?article=3337&context=smallsat com/flight-systems/space-launch-vehicles/pegasus/docs/Pegasus_ UsersGuide.pdf 5. Spaceflight Now news website: https://spaceflightnow.com/2017/12/07/ rocket-lab-pushes-back-second-electron-launch-by-one-day/ 14. US Patent US20160108855A1, Inventor Kjell Anflo, Application 2016- 04-21 https://patents.google.com/patent/US20160108855A1/ 6. SpaceWorks Forecast website: http://www.spaceworkscommercial.com/ 15. MIT/Lincoln Labs TROPICS Mission web page: https://tropics.ll.mit. 7. http://www.moog.com/content/sites/global/en/news/operating-group- edu/CMS/tropics/pdf/nasaTropicsFactSheet.pdf news/2018-2017/SLO-MV-Announcement-07172018.html

Received 8 January 2019 Approved 11 January 2019

JBIS Vol 71 No.11 November 2018 415 JBIS VOLUME 71 2018 PAGES 416-425

REDESIGN & SPACE QUALIFICATION OF A 3D-PRINTED SATELLITE STRUCTURE with Polietherimide

JONATHAN BECEDAS1, ANDRÉS CAPARRÓS1, PABLO MORILLO1, GERARDO RODRÍGUEZ FLORES1 and JOSÉ MANUEL URTEAGA2 1Elecnor Deimos Satellite Systems C/ Francia 9, Polígono Industrial La Nava III, Puertollano, Spain; 2INTA (Instituto Nacional de Técnica Aeroespacial), Crta. Ajalvir km 4,4, 28850 Torrejón de Ardoz (Madrid), Spain email [email protected]

Mass is one of the main constraints when designing a satellite since it directly affects to the costs of the overall mission and in particular to the cost of the launch: one of the most accepted estimations is that the cost of putting one kilogram into a Low Earth Orbit is about $20,000, although the cost for small satellites can be higher: about $50,000 per kilogram. This is notorious, because in the case of small satellites, the mass of the structure represents between 30% and 40% the mass of the whole satellite; considering the use of aluminium alloys, which are commonly used because they are light and present good mechanical behaviour. However, additive manufacturing advances in the last years can enable the use of other materials such as thermoplastics with good mechanical behaviour, resistance to space environmental conditions and half the density of the aluminium, which would drastically reduce the mass of the structure. In this work a small 8U satellite platform is presented and evaluated to demonstrate the feasibility of applying additive manufacturing with polietherimide ULTEM in the whole functional structure of small satellites. The satellite structure was redesigned based on the results of preliminary vibration tests carried out during the first quarter of 2018. Then the new design was manufactured and space qualified by passing a final test campaign with thermal-vacuum tests. The results demonstrated that the new structure with 3D printed polietherimide can be used in the space flight of the next generation of satellites. This study is part of the H2020 European Research Project ReDSHIFT (Project ID 687500).

Keywords: Polyetherimide, ULTEM, Vibration tests, Thermal-vacuum tests, Space qualification, Additive manufacturing, 3D printing, FDM, Satellite structure, Small Satellites, CubeSat

1 INTRODUCTION the brand ULTEM 9085 and ULTEM 1010 are already qualified for space applications [4]. One of the main constraints to launch a satellite is its mass. Although it can vary, putting one kilogram into a Low Earth The entrance of these new types of materials has been ena- Orbit can have (at the date of this publication) an associated bled by the development of the additive manufacturing tech- cost between $40,000 and $60,000 for small satellites under 100 nology, which also provides benefits such as short time man- kg [1]. ufacturing time, weight reduction while ensuring structural properties, less environmental impact by optimizing the use of This is important since, in small satellites, the mass of the material and power consumption, process speed optimization structure is between 30% and 40% of the total mass of the satel- and use of “design for need” paradigm instead of “design for lite because metallic aluminium alloys are commonly used for manufacturing” paradigm among others. that purpose [2]. In this work, the design and space qualification of an 8U However, the use of new materials with good mechanical CubeSat structure 3D printed with PEI ULTEM are exposed. and thermal behaviour for aerospace applications is appearing with the capability to highly reduce the mass of the structure The work is based on preliminary studies and development by maintaining mechanical and thermal functionality. This is of the research team: in [5] the use of additive manufacturing the case of thermoplastics such as polietherimide (PEI), with applied to small satellites was analysed. Later, in [6], the authors resistance to high and low temperatures, high glass transition proposed a design of a small satellite, which was 3D printed in temperature, high resistance to mechanical loads and half the PEI ULTEM. A satellite structural model with the 3D printed density of the aluminium [3]. Besides, polietherimide under structure was mechanically qualified to fly in a Polar Satellite Launch Vehicle (PSLV). The structure was tested under qua- si-static load (QSL), sine load, random load and shock load. This paper was presented at the 16th Reinventing Space Conference, The structure presented higher flexibility and reduced mass London, 30 October–1 November 2018. compared with a similar structure manufactured on alumini-

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Fig.1 Structural model design. um alloy AA-6082 T651 with a classical CNC milling process • Th e X and Y panels were also reinforced in the corner and also qualified to fly in a PSLV. near the bolt which transmit the load to the structure from the –Z panel. The structure here presented is a design retrofit of that pre- • Th e bracket which was the interface between the primary sented in [6]. It was redesigned to improve integrability, versa- structure and the star trackers was removed. Its function- tility and eliminate areas with large displacements. This time ality was substituted by four stiffeners, which are part of it was tested in thermal vacuum conditions to emulate the the same body of the central tray part. space environment in order to completely qualify it to be used • Th e GPS antenna and the star trackers were relocated to in space applications. The results obtained in this new testing improve versatility. campaign carried out with a thermal vacuum chamber are presented and analysed. 2.2 Experiment description

2 MATERIAL AND METHODS The objective of this experiment is to validate the use of additive manufacturing technology with PEI to be implemented in The qualification of the structural design presented in this paper functional structures of satellites. The experiment is based on the involves the thermal vacuum tests in which the structural mod- thermal vacuum qualification of an 8U CubeSat structural model behaviour was analysed under space environment emulation. el which was 3D printed in PEI ULTEM™ 9085. The structural model was constituted by the structural subsystem of the satellite The mechanical vibration tests are not included because the (functional structure) and by dummy masses, which were me- proposed design was improved with respect to the original one, chanically equivalent to the real components of the satellite. The which was qualified to fly in a PSLV launcher [6]. The current dummy masses where manufactured in aluminium alloy AA- structure is improved to resist NASA GEVS qualification levels, 6082 T651 through a classical CNC milling process. which are more demanding. 2.2.1 Testing specimen description In this section, the design of the structural model is introduced, and the experiment and the methodology applied in the The test specimen used in this test campaign was an 8U Cube- experimentation are detailed. Sat structural model. The structure was made through a 3D printing process (Fused Deposition Modelling) in which PEI 2.1 Design of the structure ULTEM™ 9085 was used. The dummy masses were made of aluminium alloy AA-6082 T651 with classical CNC milling. The satellite structure was a redesign of the structure present- The density of the PEI ULTEM™ 9085 was 1.34 g/cm3, while the ed in [6]. It was an 8U CubeSat in which modifications were density of the aluminium alloy was 2.70 g/cm3. made based on lessons learned from the testing campaign in the previous paper. The redesign is shown in Fig. 1. The follow- The structural subsystem of the 8U CubeSat consisted of six ing changes took place: panels (cube faces) and a central avionics tray providing me- • I n the new design, the walls of the panels with width 1mm chanical interface for all the equipment on board the satellite. were removed to avoid having large displacements to stat- The brackets to integrate the two star trackers are part of the tray. ic loads in the structure. • S ymmetry was also introduced in X and Y panels to re- The aluminium dummy masses were mechanically equiva- duce complexity in design, analysis, manufacturing, inte- lent to the following components of a medium resolution Earth gration and versatility. Observation mission: an Astro Digital Data & Power Module • Th e –Z panel was reinforced in width around the bolts (DPM), three Sinclair reaction Wheels (RWs), a Sensonor which transmit the load to the structure. Gyrometer (GYR), two BST Star Trackers (STs), a Syrlinks

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X-Band Transmitter Unit (XTU), a VACCO Propulsion Sys- TABLE 1 Mechanical properties of mass dummies tem (PROP), an ANTCOM GPS antenna (GPS), an SCS Space Tag Mass (gr) X CoG (mm) Y CoG (mm) Z CoG (mm) Gecko imager payload (CAM), an EnduroSat X-Band antenna and an Astro Digital UHF Antenna. DPM 1796 -14.96 52.15 40.33 CAM 473 -48.70 -49.65 29.36 The internal distribution of the components in the internal GPS 41 0.00 0.00 113.00 volume of the 8U structure was done by considering the centre ST1 233 -59.65 -67.28 -45.00 of gravity (CoG) requirements imposed by the CubeSat Stand- ard [7], adapted to an 8U form factor. All the dummies were ST2 233 67.28 -59.65 -45.00 manufactured in aluminium and anodized to be preserved RW1 171 -24.13 58.90 -30.05 from corrosion. The total mass of mock-ups was 5.2 kg. The RW2 171 -73.20 10.32 -30.05 structure in ULTEM™ had a mass of 1.38 kg. Table 1 shows the list of mock-ups with the most relevant mechanical properties: RW3 171 -73.25 59.25 -25.07 mass and coordinates of the centre of gravity with respect to GYR 51 -0.39 -51.65 -15.30 the geometric centre of the satellite structure. PROP 1181 52.60 -47.65 42.07 2.2.2 Testing platform description XTU 433 50.02 67.25 -50.06

The thermal vacuum test was performed at INTA facilities. Ta- TABLE 2 Technical specifications INTA’s thermal vacuum ble 2 shows the technical specifications of the thermal vacuum chamber chamber used in the experimentation. Item Specification In order to perform the Thermal Control of the specimen, Shroud temperature range -175ºC ~ +160 ºC a Thermal Special Check-Out Equipment SCOE developed by Degree of vacuum Less than 10-5 mbar Elecnor Deimos Satellite Systems was used. The equipment Max. heat load 2 kW could control 25 independent power lines with up to 3 thermocouples per channel. Shape Mail box Useful test volume 4 x 4 x 4 meters 2.2.3 Instrumentation of the testing specimen Sensors Thermocouples (Type T) Number of channels 400 Three test heaters K010150C3, 3 test heaters K010050C3 and 4 test heaters K010030C5 manufactured by 3M were used to con- Available feedthroughs 10 DN 250 ISO-K, 2 DN 500 ISO-K trol the internal temperature of the dummy masses, emulating Operational mode GN2 the real equipment of the satellite in the space environment in Data acquisition and control 40 Control Channels an orbit of 600 km. In Table 3 the characteristics of the heaters are depicted. They were named as the electronic equipment Mass spectrometer Range: 0 – 200 M/e they were emulating and placed over them: PROP for the pro- Cleanliness ISO 14644 (8 class) pulsion subsystem, CAM for the camera (or imager); DPM for the Data & Power Module; RW for the three reaction wheels; XTU for the X-Band Transmitter Unit; ST1 and ST2 for the 1 to observe the stabilization temperature of the equipment, Star Trackers 1 and 2 respectively; and SPX, SPY and SPZ for Stage 2 to analyse the power required to maintain the equip- the Solar Panel X, Y and Z respectively. To facilitate the installa- ment temperature between acceptable thresholds when per- tion of the heaters emulating sunlight conditions, three panels forming in space, and Stage 3 to monitor the system tempera- made of polietherimide were assembled to the structure. The ture profile when the system is performing in orbit. heaters were defined by their maximum voltage “V”, maximum current “I”, maximum Power “P”, impedance “Z” and length Stage 1: temperature stabilization “L”. In all of them, the width was 25 mm. The GPS and UHF antennas were not considered because of their low power and The purpose of this stage is to check the stabilization temperature low heating generation and transfer as each of them consumes less than 1 W under work load. TABLE 3 Test heaters characteristics Thermocouples were mounted over the Test Heaters to Heater Name V I P Z L monitor and control them. Three thermocouples were attached and Location (volts) (amps) (watts) (ohms) (mm) to every component to monitor the temperature evolution. PROP 120 0.208 25 576 127 One of the thermocouples was directly attached over the Heat- CAM 28 0.54 15 52.26 76 er. The other two were located to measure the temperature of DPM 120 0.208 25 576 127 the equipment. The temperature was determined as the arith- metic mean of the Control Thermocouples. RW 28 0.54 15 52.26 76 XTU 120 0.208 25 576 127 Additionally, control thermocouples were installed in differ- ST1 28 0.54 15 52.26 76 ent points of the structure, their average is represented as RST- COLD in the results section. ST2 28 0.54 15 52.26 76 SPX 120 0.625 75 192 381 2.2.4 Test specifications and setup SPY 120 0.625 75 192 381 The test was designed to be carried out in three stages: Stage SPZ 120 0.625 75 192 381

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Fig.2 Structural-thermal model during its instrumentation with heaters and thermocouples inside the cleanroom. of the equipment when it is in cold (eclipse) and hot (sunlight) TABLE 4 Nominal power consumption conditions. To do that, the nominal power dissipation of the Heater Name and Number Power Consumption equipment was set in the heaters of every dummy mass, Table 4. Location (watts) Data & Power Module 1 13.5 The temperature of the shroud was set to -150ºC and the thermal cycling was conducted 3 external heaters emulating Reaction Wheels 3 1.8 the solar power (1360 W/m2). Gyrometer Sensonor STIM210 1 1.5 Star Trackers BST ST200-S 2 0.67 The first step was to reach high vacuum conditions and cold temperature stabilization inside the chamber. Cold tempera- X-Band Transmitter EWC27 XTU 1 10 ture stabilization was considered when the thermal variation at each channel was lower than 3ºC/h. Propulsion module VACCO 1U 1 10 Payload imager SCS Gecko 1 3.5 The cycle started in high vacuum conditions (10-5 mbar) and cold case, i.e. without power in the external heaters and nominal power dissipation in all the internal components of the structural model. Then wait for stabilization, which was powered on (27.2 W/heater) until the system stabilizes in the considered when temperature variations were less than 3ºC/ hot case. hour. Once the stabilization was reached, the temperature of the structure was registered. Fig. 3 depicts the temperature profile of the shroud (blue line) in time and the external heaters power evolution in time After that, the external heaters emulating sunlight were (red line).

Fig.3 Stage 1 temperature and power profiles.

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Fig.4 Stage 2 Temperature and power profiles.

Stage 2: power analysis equilibrium, using the external heaters when sunlight and setting the equipment nominal thermal dissipation. Fig. 5 shows The purpose of this Stage 2 was to analyse the power required the power profiles of one complete orbit. The external heaters to maintain the equipment temperature between acceptable emulating sunlight were powered on with 27.2 W/heater when thresholds when the satellite is performing in orbit. The pow- sunlight and powered off when eclipse (brown line). The DPM er required to maintain the acceptable temperature range was was always set to dissipating nominal power at 13.5W (green calculated as the sum of the nominal dissipated power of the line). The rest of the equipment was configured to maintain satellite equipment and the additional required power to be in- the target temperature with PID automatic control mode, The jected or dissipated by the thermal control system. camera heater was set to emulate one image acquisition from minute 30 to minute 40 of the start of the orbit (purple line) but The degree of vacuum in the TVAC chamber was maintained considering 10W as peak power to include additional power lower than 10-5 mbar during the whole vacuum test duration. dissipation associated to the data processing. The XTU heater was set to emulate data download from minute 40 to minute Fig. 4 shows the temperature and power profiles of this stage 50 of the start of every orbit (red line). The propulsion mod- 2 experiment. The temperature of the shroud was maintained ule heater was set to emulate one orbital manoeuvre at perigee at -150ºC (blue line), the external heaters power emulating the from minute 50 to minute 60 of every orbit. Finally, the rest sun power affecting the system during sunlight is depicted in of heaters in the equipment (reaction wheels, gyrometer and green (27.2W/heater). The temperature of the equipment is de- star trackers) were set to target temperature control mode with picted in red. For all equipment Target temperature was set to nominal power as maximum power, it is assumed an average 20ºC, excepting for PROP and XTU (30ºC) and DPM (50ºC). power consumption of 5W each as a worst case.

Stage 3: temperature monitoring in orbit 2.2.5 Experimentation methodology

It consisted on testing 3 consecutive orbits (30 minutes eclipse, The experimentation was carried out at Instituto Nacional de 60 minutes sunlight) starting from operational temperature in Técnica Aeroespacial (INTA) premises in Torrejón de Ardoz,

Fig.5 Stage 3 power profiles.

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Fig.6 Structural model instrumented with heaters and thermocouples inside the TVAC chamber.

Madrid, Spain (http://www.inta.es). to the connection racks.

The thermal vacuum chamber was inside a clean room Once the door of the TVAC chamber was closed, it start- (ISO8). Thus all the structural model components, thermo- ed creating vacuum conditions until 10-5 mbar pressure. After couples, heaters, wiring, auxiliary components and tools were that, the shroud was set at -150ºC. The experiment was carried cleaned. out with the following order: 1. TVAC chamber stabilization at high vacuum and low The structural model was instrumented and integrated in- temperature. side the clean room and placed in the TVAC chamber over a 2. Stage 1 experiment. grid and separated from the bottom shroud with Teflon blocks 3. Internal equipment stabilization at operational (Fig. 6). Then the thermocouples and heaters were connected temperature.

Fig.7 TVAC chamber stabilization.

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4. Stage 2 experiment. TABLE 5 Stabilization temperatures 5. Stage 3 experiment. Channel Stabilization Temperature (ºC) 6. Shut down heaters and TVAC chamber stabilization. SPY -122.5 at ambient conditions. SPZ -118.51 3 RESULTS SPX -123.23 DPM -53.28 After 8 hours of cooling down the thermal vacuum chamber with the temperature of the shroud at -150ºC, the equipment CAM -69.61 was considered to be stabilized: the stabilization criteria was ST1 -79.69 that the temperature decremented lower than 3ºC/hour. Fig. 7 ST2 -84.31 depicts the evolution in the temperature of the different components of the testing specimen. RW -73.83 RST-COLD -113.9 The stabilization temperature depended on the exposure of PROP -59.68 the equipment to the shroud of the thermal vacuum chamber XTU -76.69 and the mass of the equipment. For instance, the higher the exposure to the shroud, the lower the stabilization temperature. This effect is more notorious in the external heaters emulating the solar panels and in the components with lower mass: XTU pation of these components was high enough for reaching ac- or ST1 and ST2. The stabilization temperatures are shown in ceptable operation temperatures without active thermal control Table 5. during operation. The stabilization temperatures of the different components in the cold case are shown in Table 6. 3.1 Results of Stage 1: temperature stabilization Besides, the structure (RST-COLD) was affected by the dis- During the experimentation in Stage 1, the stabilization tem- sipated power of the internal equipment, increasing the surface peratures were recorded for cold and hot case (the external temperature up to -78.03ºC from -113.9ºC. The optical equip- heaters of the solar panels were powered off and powered on ment such as CAM, ST1 and ST2 were still in negative range far respectively according to Fig. 3 in the previous section). Fig. from the operation temperature point. This was an indicator of 8 shows the response of the system, in which the temperature need of support from an active thermal control system in these evolution of the different components can be observed. Notice components to maintain the survival temperature range. that at the start of the Stage 1 experiment there is an increment in temperatures for the solar panels and for the DPM. This was In the hot case with the heaters on the solar panels pow- due to an erroneous activation of the external heaters on the ered on (dark blue, dark red and violet lines), the overall tem- solar panels. This was corrected 6 minutes later, and should not perature of the system increased. Also notice that the DPM be considered in the analysis. (purple line) was affected by the radiation of SPY as its control thermocouple was located behind the SPY heater reaching a Thus, the stabilization temperatures in the cold case were clearly overheated state. That indicated that additional ther- positive for DPM (purple), RW (pink), PROP (dark green) and mal isolation between solar panels and internal equipment XTU (light blue). This indicated that the nominal power dissi- shall be included.

Fig.7 TVAC chamber stabilization.

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TABLE 6 Stabilization temperatures in cold and hot cases (Stage 1) COLD CASE HOT CASE Channel Stabilization Channel Stabilization Temp. (ºC) Temp. (ºC) SPY -104.4 SPY -14.82 SPZ -91.8 SPZ -2.08 SPX -97.27 SPX 23.4 DPM 10.53 DPM 99.73 CAM -16.77 CAM 31.9 ST1 -39.61 ST1 -12.79 ST2 -47.71 ST2 -23.46 RW 14.98 RW 39.24 RST-COLD -78.03 RST-COLD -54.7 PROP 8.13 PROP 46.98

XTU 47.21 XTU 65.16 Fig.8 Stage 1 experiment results.

Furthermore, optic equipment was desired to reach an oper- PROP and XTU (30ºC) and DPM (50ºC). Besides, the external ational temperature close to zero but preferably in the positive heaters on the solar panels where powered on, at 03:00 PM. It range. In this case, CAM reached 31.9ºC, which was consid- can be noticed that the target temperature control maintained ered in range, but in normal operation the power dissipation the temperature of the equipment within operational range, was not continuous. Thus, for CAM, ST1 and ST2 active ther- independently of the external stimulation. The only exception mal control should be required. was the DPM, which was thermally coupled to the SPY.

In the case of the XTU, it reached a moderate overheated The results show that all the equipment stabilized near the condition. However, it had continuous power dissipation, which target temperature. This suggests that the structure has good was not representative of a real condition in which the XTU will thermal isolation properties that help to maintain residual heat dissipate power in an intermittent manner. The temperatures of inside, reducing the required power of an active thermal con- the equipment in the hot case are depicted in Table 6. trol system. This was also checked with the nominal power applied by the thermal control, which applied nominal power to 3.2 Results of Stage 2: power analysis every component with a duty cycle of 50%. This indicates that the active control system needed to maintain survival temper- At the beginning of Stage 2, all the power supplies were ature margins requires half of nominal power to maintain the switched off in order to decrease the temperature of all the system within the operative range. equipment under 60ºC not to overheat any component (Fig. 9). When the DPM and the XTU decreased under 40ºC, the tar- The oscillations in the temperature from 7:00 PM were due get temperature was configured for all the channels as stated in to attempts of adjusting the PID control gains of the heaters for Section 2.2.4: 20ºC for all the equipment with the exception of a better control.

Fig.9 Stage 2 experiment results.

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Fig.10 Stage 3 experiment results.

3.3 Results of Stage 3: temperature monitoring in orbit tuned to operate in safe temperature thresholds with an active thermal control system of 50% nominal power for every com- During Stage 3, (Fig. 10), three orbits with operational scenar- ponent in representative Low Earth Orbit environmental con- ios were emulated. During the orbits, all the control channels, ditions. The reached stability confirmed that the structure was except the DPM, were controlled with automatic target tem- suitable for space use. perature mode: 10ºC for optics (ST1, ST2), 50ºC for PROP, 20ºC for CAM, 30ºC for RW and 30ºC for XTU. DPM was set Finally, the heaters were powered off and the chamber was set on 13.5W fixed power, due to the fact that DPM is always on in back to ambient conditions. The evolution is shown in Fig. 11. normal operation. 4 CONCLUSIONS The experiment started by powering on the external heaters during 60 minutes to simulate sunlight. During the last 30 min- In this paper, a novel design of an 8U CubeSat structure, based utes of sunlight the CAM, XTU and PROP were sequentially on previous work of the research group, was presented. The activated for 10 minutes (as shown in Fig. 5). new design incorporates mechanical reinforces to resist to larger loads and reduce large displacements; symmetry to re- As shown in Fig. 10, the temperature in all the equipment duce complexity in manufacturing, integration and to increase remained stable between margins and it was very near to the versatility; and reduction of parts by incorporating the brack- target temperature, with the exception of the DPM, which et functionality in the central tray. As the previous version of slowly increased after each orbit. This indicates the need of the structure mechanically qualified, this paper focuses on its a power dissipation method in the DPM such as radiators or thermal vacuum qualification. thermal strips. During the experimentation no outgassing was detected, and After three orbits, the thermal model of the specimen was after the experiment no plastic deformations, crystallization or

Fig.11 Stage 3 experiment results.

424 Vol 71 No.11 November 2018 JBIS REDESIGN & SPACE QUALIFICATION OF A 3D-PRINTED SATELLITE STRUCTURE with Polietherimide functional changes were detected in the ULTEM™ structure. Acknowledgements

Three experiments were carried out in thermal vacuum: i) The research leading to these results has received funding from stabilization temperatures detection of the components with the Horizon 2020 Program of the European Union’s Frame- nominal power dissipation in cold and hot cases, ii) analysis of work Programme for Research and Innovation (H2020-PRO- the power required to maintain the equipment temperature be- TEC-2015) under REA grant agreement number 687500 – tween acceptable thresholds when the satellite is performing in ReDSHIFT (http://redshift-h2020.eu/). orbit and iii) temperature monitoring in three consecutive orbits. This publication reflects only the author’s views and the -Eu During the experimentation, the thermal model behaved ropean Commission is not liable for any use that may be made well and within operational temperature ranges. The only ex- of the information contained therein. ception was the DPM component, which requires active temperature control to perform without overheating and better The authors want to acknowledge the people who also isolation from the solar panel Y, whose radiated power directly made possible this work with their indirect participation: Es- affected the DPM's thermal behaviour. ther Sarachaga, Antonio Ramírez, Rosa Domínguez, Daniel Hernández, Antonio Támara, Silvia Martínez and Florio Dalla However, the whole system always performed within safe Vedova. temperature thresholds, and the active thermal control only had to provide control with a duty cycle of 50%, which sug- Notes gested that the ULTEM™ structure had good thermal isolation properties that help to maintain residual heat inside, reducing The authors from Elecnor Deimos Satellite System contributed the required power of the active thermal control system. to this paper with the design, analysis, manufacture, testing of the specimen, and with the analysis of the results. In any case, the reached thermal stability during the experimentation confirmed that the structure was suitable for The author from INTA contributed with the testing facilities space use. and instrumentation, tests definition, execution and setup.

REFERENCES

1. SpaceFlight. 2018. “Pricing information”. Retrieved from: http:// on 2018/06/06. spaceflight.com/schedule-pricing/#pricing. Accessed on 2018/09/04. 5. J. Becedas and A. Caparrós. 2018. “Additive Manufacturing Applied 2. W. Wassmer, 2015. “The Materials Used in Artificial Satellites and Space to the Design of Small Satellite Structure for Space Debris Reduction”. Structures”. Azo Materials. Retrieved from: https://www.azom.com/ In Applications of Design for Manufacturing and Assembly. Chapter 9. article.aspx?ArticleID=12034. Accessed on 2018/09/04. IntechOpen. United Kingdom. ISBN 978-953-51-6929-1. Păcurar, A.C. 3. M. Mudarra, J. Sellarès, J.C. Cañadas and J.A. Diego. 2015. “Sublinear editor. dispersive conductivity in polyetherimides by the electric modulus 6. J. Becedas, A. Caparrós, A. Ramírez, P. Morillo, E. Sarachaga and formalism”. IEEE Transactions on Dielectrics and Electrical Insulation, A. Martín-Moreno. In press. “Advanced Space Flight Mechanical Vol. 22, No. 6, pp. 3327-3333. Qualification Test of a 3D Printed Satellite Structure Produced in 4. M. Molitch-Hou, 2017. “Made In Space Begins 3D Printing PEI/PC Polyetherimide ULTEM”. In Advanced Engineering Testing. IntechOpen. on the ISS”. Engineering.com. July 2017. Retrieved from: https://www. United Kingdom. ISBN 978-953-51-6706-8. Ali, A. editor. engineering.com/3DPrinting/3DPrintingArticles/ArticleID/15254/ 7. J. Puig-Suari, and R. Nugent. “6U Cubesat Design Specification” The Made-In-Space-Begins-3D-Printing-PEIPC-on-the-ISS.aspx. Accessed CubeSat Program. Cal Poly SLO, revision 1.0, July 2018.

Received 14 January 2019 Approved 14 January 2019

JBIS Vol 71 No.11 November 2018 425 JBIS VOLUME 71 2018 PAGES 426-430

THE EPSILON LAUNCH VEHICLE: Status and Future

RYOMA YAMASHIRO, IMOTO TAKAYUKI, JAXA Epsilon Rocket Project Team 2-1-1, Sengen, Tsukuba, Japan email yamashiro.ryoma@jaxa.jp

The Epsilon Launch Vehicle (LV) is a next-generation solid propellant rocket developed in a Japanese national program led by JAXA. After the completion of the development, it is positioned as a Japanese flagship LV and plays a key role in securing Japan’s autonomous capability to launch small satellites for observation and scientific missions. It also offers greater launch opportunities to small satellites for commercial missions. The Epsilon is a highly-reliable space transportation system fully reflecting Japanese rocket technology long accumulated through many vehicle programs such as the former M-V and currently operated H-IIA/B LVs. The Epsilon offers user-friendly launch services with newly incorporated technologies, such as next-generation ground support / check-out systems, highly accurate orbit injection system, and advanced techniques improving the payload environment. Our immediate plan is to launch the fourth Epsilon carrying multiple payloads mounted on a newly developed PAF (Payload Attach Fitting). At the same time, aiming to create synergy effect, we have been developing parts and components to be shared with the H3 Launch Vehicle, Japanese large- size next-generation launch vehicle. Looking further ahead, we have started the concept study for the future Epsilon.

Keywords: Solid Propellant, Solid Motor, Payload Environment, Launch Operation, M-V, Epsilon

1 EPSILON LAUNCH VEHICLE 2 shows the payloads of the M-V and the Epsilon.

1.1 Overview 1.2 The First Epsilon

The Epsilon LV, the latest small rocket developed by JAXA, The main purpose of the first Epsilon was to demonstrate the is characterized by superior operability and payload environ- operational technologies (described later), the main technolo- ment. The Epsilon is aimed to facilitate participation in space gies which had to be developed in a short period. Therefore, we development by offering more opportunities for small satellites decided to make full use of existing rocket components for the whose technology is showing remarkable progress in recent first Epsilon. We employed the Solid Rocket Booster-A (SRB-A) years so that more people can challenge space. of the H2A LV for Epsilon’s first-stage motor and improved M-V’s third-stage motor and kick motor for its second-stage The Epsilon basically consists of three stages, the first, sec- motor and third-stage motor, respectively. We also utilized ond and third stages, each of which has a solid motor as main many H2A avionics devices. This approach reduced the devel- propulsion system. It can be equipped with a Post Boost Stage opment period and risk, especially, in the improvement of the (PBS) having liquid propulsion system as an option to boost solid motors which required only the sub-size combustion test its capability for missions that require high orbital accuracy, and analysis because the improvement was based on our proven such as launching an Earth observation satellite to SSO. Fig. technologies which can eliminate the need of time-consuming 1 shows the outline of a variant of the Epsilon, the Enhanced full-scale ground combustion test. We concentrated on develop- Epsilon (described later), and its configuration. Table 1 shows ing the system integration and operation technologies, and fi- the Epsilon’s specifications. nally launched the first Epsilon after three years of development.

The M-V LV, the predecessor of the Epsilon, was primar- The first Epsilon was launched in 2013 in an optional con- ily targeted at satellites and spacecraft for scientific missions. figuration with PBS, delivering a planet observation satellite Taking over the role of responding to the demand of science “Hisaki” into the planned orbit. The flight proved Epsilon’s -ex missions, the Epsilon is now expanding into a new field of cellent function and performance [1]. commercial satellites, such as Earth observation satellites, in which new ideas of venture businesses are incorporated. Table 1.3 The Second and Third Epsilon

After demonstrating the operation technologies by taking ad- This paper was presented at the 16th Reinventing Space Conference, vantage of our existing technologies, we began the development London, 30 October–1 November 2018. of the Enhanced Epsilon. Focused on are its second-stage mo-

426 Vol 71 No.11 November 2018 JBIS THE EPSILON LAUNCH VEHICLE: Status and Future

Fig.1 Epsilon LV.

TABLE 1 Epsilon’s Specifications STAGES ITEMS 1st stage 2nd stage 3rd stage PBS* PLF OVERALL (SRB-A3) (M-35) (KM-V2c)

Length [m] 11.7 4.0 2.2 1.3 9.6 26.0 Diameter [m] 2.6 2.6 1.4 2.0 2.6 Weight [ton] 74.5 17.2 2.8 0.4 0.7 95.6 Propellant [ton] 66.0 15.0 2.5 0.1 - Thrust [kN] 2350 445 99.6 0.2 - Burn time [s] 108 129 88 743 - Propellant Solid (HTPB) Solid (HTPB) Solid (HTPB) Hydrazine - Isp [s] 284 295 299 215 - Control TVC & Solid Thruster TVC & Thruster Spin Thruster -

* PBS is installed as option

TABLE 2 Payloads of the M-V and the Epsilon PAYLOAD LAUNCH VEHICLE Name Objective Payload Orbit Mass (Perigee, Apogee, Inclination)

M-V-1 HALCA (MUSES-B) Radio-Astronomical Satellite (560 km, 21,000 km, 31 deg) 830 kg M-V-3 NOZOMI (PLANET-B) Mars Orbiter Areocentric orbit 540 kg M-V-5 HAYABUSA (MUSES-C) Asteroid Sample-return Spacecraft Heliocentric orbit 510 kg M-V-6 SUZAKU (ASTRO-EII) X-ray Astronomy Satellite (550 km, 550 km, 31 deg) 1,700 kg M-V-7 AKARI (ASTRO-F) Infrared Imaging Satellite (700 km, 700 km, 98 deg) 952 kg M-V-8 HINODE (SOLAR-B) Solar Physics Satellite (680 km, 680 km, 98 deg) 900 kg Epsilon-TF1 HISAKI (SPRINT-A) UV planetary space telescope (950 km, 1150 km, 31 deg) 348 kg Epsilon-F2 ARASE (ERG) Exploration of energization and (440 km, 32,000 km, 31 deg) 355 kg Radiation in Geospace Epsilon-F3 ASNARO-2 Earth Observation Satellite (500 km, 500 km, 97 deg) 570 kg

JBIS Vol 71 No.11 November 2018 427 RYOMA YAMASHIRO, IMOTO TAKAYUKI tor development which required a full-scale ground combustion test, design change of the structure, and weight reduction of the avionics equipment. The Enhanced Epsilon increased its launch capability by 30% (to 600 kg from 450 kg into a 500 km SSO) and its payload envelope area by 20% (in the longitudinal direction) [2][3].

The Enhanced Epsilon, in a basic configuration without PBS, was first launched in 2016 as the second Epsilon and suc- ceeded in sending a geospace exploration satellite "Arase" into the orbit. In 2018, the third Epsilon, in an optional configuration with PBS, successfully launched an Earth observation satellite "ASNARO-2" into SSO (Fig. 2). Both of the launches satisfied the requirements of the satellites, verifying Epsilon’s high function and performance again.

2 APPLIED TECHNOLOGIES

2.1 Techniques for Payload Environment Fig.2 The Third Epsilon Launch. Several techniques for improving the payload environment are applied to the Epsilon LV to facilitate the participation in space development by payloads which adopt various ideas, and to ac- 2.2 Solid Motor commodate a wider range of payloads. The first technique is a vibration damping mechanism. A damper is installed in the With the new second-stage motor having been adopted to the PAF structure to reduce sinusoidal vibration applied to payload second and following Epsilons, we are applying new technol- by rocket. We employed this mechanism for the first and third ogies for improving launch capability and cost reduction. We Epsilons and confirmed its effectiveness to reduce the vibration developed a high sealing material to be used for insulation of acceleration to a maximum of 0.2 G (0 to peak) in the first Epsi- motor cases and simplified their laminated structure. We also lon and a maximum of 0.06 G (0 to peak) in the third Epsilon. improved the propellant injection process and increased the ratio of the amount of propellant against the volume of a mo- The second technique is a new flue on the launch pad -de tor case. The latest 4D-C/C manufacturing method is applied signed based on the analysis using CFD to minimize the effect to nozzle throat inserts for efficient production. Moreover, as on payload caused by the sound emitted by the rocket plume at these technologies have improved the launch capability, con- lift-off. Since the launch of the first Epsilon, it has demonstrat- ventional extension nozzles have been replaced by fixed noz- ed its designed performance with less than 135 dB (overall) of zles for cost reduction [6]. the acoustic level on payload (Fig. 3) [4]. 2.3 Avionics The third technique is a low impact separation mechanism, which is operated by electrical signals and mechanical motion Epsilon’s avionics system incorporated two new technologies. instead of conventional pyrotechnic products (Fig. 4). It was One is RINA (Radio and Inertial NAvigation), a device that first installed in the third Epsilon. Although the impact was measures and calculates the position and velocity of LV by us- not be measured during actual launch, the ground test showed ing navigation satellites and onboard IMU (Inertia Measuring that the mechanism lowered the impact to less than 1000 G at Unit), thus reducing traditional radar stations on the ground. separation [5]. It was demonstrated in the second Epsilon and will come into

new flue

Fig.3 New Flue and Reduced Acoustic Level in Epsilon’s Fairing.

428 Vol 71 No.11 November 2018 JBIS THE EPSILON LAUNCH VEHICLE: Status and Future

Fig.4 Techniques for Payload Environment Improvement (left: vibration damping mechanism, right: low impact separation mechanism). operation in the fourth and following Epsilons. take full advantage of this property, we have applied several new operational technologies (Fig. 5). The other is improved PSDB (Power Sequence Distribution Box). Its conventional mechanical relays have been changed ROSE (Responsive Operation Support Equipment) is a new to semiconductor relays, reducing the mass per unit to 12 kg device to collectively inspect the whole LV for normality. Con- from 20 kg. This reduction seems to be a slight improvement ventional LVs have avionics devices which need to be individu- but cannot be ignored, given the fact that three PSDBs need to ally inspected and require a lot of time to prepare for setup, etc. be mounted in a single Epsilon, a small rocket launching a pay- On the other hand, the Epsilon has ROSE which can collect and load of several hundred kilograms. The new PSDBs have been integrate data from these avionic devices and reduce the time applied to the second and following Epsilons. required for inspection. The output from ROSE is received and assessed for normality by LCS (Launch Control System) on the 2.4 Operation Technology ground. LCS currently assesses data for normality according to the formulated threshold value; however, with more Epsilons The Epsilon LV is excellent in operability. The solid motor, to be launched, it will conduct statistical assessment based on the main propulsion system, eliminates the need of propellant accumulated data. We also developed MOC (Miniature Ord- charging operation usually conducted just before launch. To nance-circuit Checker) to enhance operability of pyrotechnics

Fig.4 Techniques for Payload Environment Improvement (left: vibration damping mechanism, right: low impact separation mechanism).

JBIS Vol 71 No.11 November 2018 429 RYOMA YAMASHIRO, IMOTO TAKAYUKI at the launch site. This device can be attached to onboard pyro- lease mechanism, is a modified version of J-SSOD which is now technic equipment and inspect its circuit safely and efficiently. in operation on ISS to release CubeSats from there [8]. It can be removed before launch and reusable [7]. 3.2 Further Research and Development for the future These technologies have reduced Epsilon’s launch operation period after the first stage installation on the launch pad to nine Since the development of the H3 LV began in 2015, we have days, a significant difference compared to 42 days required by been considering "competitiveness improvement develop- the M-V. ment (provisional)" including utilizing H3’s solid rocket motor (SRB-3), now under development, as Epsilon’s first stage and 3 FUTURE PLANS improving attitude control system, the third stage motor, PL fairing, etc. The development has just completed SRR (System Like the M-V, the Epsilon is a rocket evolving as more launches Requirements Review) and entered the next stage. are conducted. The fourth Epsilon will launch multiple PLs using a newly designed satellite mounting structure. In addition, Moreover, we are continuing to conduct research and de- we have undertaken other development aimed at lowering costs velopment on advanced technologies for further promotion by sharing components with the H3 LV. Research and develop- of “competitiveness improvement development (provisional).” ment for further future is also being carried out in parallel. Promising are an autonomous flight safety system that further improves Epsilon’s operability, a new system that enables its up- 3.1 Multiple Payload Launch per stage to return to the earth, etc. [9].

The fourth Epsilon will conduct multiple payload launch as part 4 CONCLUSION of JAXA's innovative satellite program. A total of seven satellites are to be launched: one small satellite (the innovative small The Epsilon LV is a small solid rocket which adopts state-of- satellite No. 1 (RAPIS-1)), three micro satellites (RISESAT, the-art technologies to enhance its operability and payload Microdragon, ALEe), three CubeSat (OrigamiSat-1, Aoba environment. Until now, we have successfully launched up to VELOX-IV, NEXUS). For this launch, we developed a multiple three Epsilons. We will continue conducting research and de- satellite mount structure and E-SSOD (Fig 6). The multiple sat- velopment aiming at adding to the Epsilon new values, such as ellite mount structure employs Planetary Systems Corporation's multiple PL launch and further improved operability, in order LightBand® to its separation mechanism. E-SSOD, CubeSat re- to invite more new users to space in the future.

Fig.6 Mechanism for Separating Multiple Satellites.

REFERENCES

1. Y. Morita: On the First Flight of Japan’s Epsilon Launch Vehicle, 14th Hakodate, Japan, 2016 (Japanese Only) International Space Conference of Pacific-basin Societies (ISCOPS), 6. K. Kitagawa, S. Tokudome, et. al.: DEVELOPMENT OF SOLID School of Astronautics NPU, Xi’an, China, 2014. PROPULSION SYSTEM FOR ENHANCED EPSILON LAUNCH 2. R. Yamashiro, Y. Morita, et. al.: ENHANCED EPSILON’S VEHICLE AND M-35 STATIC FIRING TEST, 67th Congress of the DEVELOPMENT RESULT AND PREPARATION STATUS FOR THE International Astronautical Federation (IAC), Guadalajara, Mexico, 2016 SECOND LAUNCH, 67th Congress of the International Astronautical 7. K. Hirose, T. Yui, et. al.: Launch Operation, Ground Support Equipment Federation (IAC), Guadalajara, Mexico, 2016, Paper No. IAC- and Facilities of the Epsilon Rocket, JSASS 59th Conference, Kagoshima, 16-D2.1.11. Japan, 2015 (Japanese Only) 3. K. Ui, et. al.: The Development Status of the Structure Subsystem for 8. K. Oribe, Y Noguchi, et al.: The Result of Epsilon Launch Vehicle Third Enhanced Epsilon Launch Vehicle, Joint Conference 30th ISTS/34th Flight and Plan for Multi launches, 69th Congress of the International IEPC/6th NSAT, Kobe, Japan, 2015 Astronautical Federation (IAC), Bremen, Germany, 2018, Paper No. 4. Tsutsumi, S., Fukuda, K., Takaki, R., Shima, E., Fujii, F. and Ui, K.: IAC-18-D2.1.8. “Numerical Analysis of Acoustic Environment for Designing Launch- 9. R. Yamashiro, Y. Morita, et. al.: EPSILON’S SECOND LAUNCH Pad of Advanced Solid Rocket,” 26th ISTS, paper No. 2008-g-05. RESULTS AND DEVELOPMENT STATUS FOR THE FUTURE, 68th 5. H. Ikaida, K. UI, et. al.: Development Status of Low-shock Separation Congress of the International Astronautical Federation (IAC), Adelaide, Mechanism for Epsilon Launch Vehicle, JSASS 60th Conference, Australia, 2017, Paper No. IAC-17-D2.1.6.

Received 14 January 2019 Approved 14 January 2019

430 Vol 71 No.11 November 2018 JBIS JBIS VOLUME 71 2018 PAGES 431-436

THE NORMS OF BEHAVIOUR IN SPACE: Our space – Whose rules?

LESLEY JANE SMITH, Faculty of Law – Leuphana University Lüneburg, Wilschenbrucherweg 69, 21335 Lüneburg, Germany email [email protected]

Among the many challenges outer space activities face today, two are worth mentioning, particularly from the perspective of ‘norms of behaviour’; firstly, how to transpose complex, non-binding rules of ‘post-treaty’ soft law onto those involved in outer space activities? Secondly, how to measure the resilience of national authorities in complying with their international space treaty obligations, should their national stakeholders and actors fail to comply with the ‘soft-law’ rules? The increasingly ‘soft law’ requirements for participation in space activities may, paradoxically, drive the call for enforcement of black letter rules one step further. In space, given the urgency posed by debris in LEO, the best practice responses must become evident in the shorter term. Most importantly, the accountability of states for overseeing the enforcement of soft-law rules could assume a central ‘best’ regulator role; these norms could be made legally enforceable at national level, with further scope for recognition as international state obligations which, when breached, may result in state responsibility for wrongful acts. This perspective is a challenge for legal theory and practice alike. If states have to apply soft law rules in managing their risk-benefit analysis, can they also be called to account for failure to ensure their enforcement? With many states currently reviewing their national licensing mechanisms and considering involvement in overarching efforts relating to space traffic management (STM), enforcing soft law at national level could be a useful tool if combined with further economic instruments that foster sustainability.

Keywords: Space Treaty Law, Soft Law, UN Long term Sustainability Guidelines (LTS Guidelines), Outer Space Treaty, Compliance, ISO Post-mission Disposal Norms, National Space Law, Catalogue of standards for National Space Laws, State Responsibility and Liability, Space Traffic Management [STM], Space Situational Awareness [SSA], Governance in Outer Space

1 INTRODUCTION AND SCOPE activities must conform to international law, where consensus-building across the international community is a goal in This paper calls for further reflection on what is now required itself [1]. More recent soft law rules, notably General Assembly from governments and stakeholders in regard to ‘our space Resolutions (GA Res.), repeat and emphasise the principles of – whose rules’. As the discussions on its future sustainabili- the OST, many of which can claim customary law status [2]. ty continue within the UN Committee on the Peaceful use of This credits them with binding force through adherence and Outer Space – COPOUS – this paper cannot provide a finite recognition by the majority of space states. Interestingly, the recommendation, but shares some considerations regarding recent LTS Guidelines debated within COPUOS formulate the architecture of space norms and rules, particularly for the frameworks and procedures for implementing the very same future. rules; nevertheless, final consensus is still outstanding on several core Guidelines [3]. The value of soft law rules for space activities, a field in which consensus on international principles was secured Well over two-thirds of the world’s nations have signed the through treaty law almost sixty years ago, can be reformulated 1967 Outer Space Treaty. Anchoring outer space as the ‘prov- as follows: does adherence to treaty law also require accept- ince of mankind’ means that all actors have to work with and ance and commitment to voluntary rules following in the same by the same rules if activities in outer space are to remain fea- vein? Outer space is open for freedom of exploration and use, sible. Until now, compliance with existing international space with Art. I Outer Space Treaty (OST) clearly stating that these rules has been variable; for example, while the level of regis- activities are to be carried out for the benefit and in the inter- tration of spacecraft (s/c) has increased over the last decade, ests of all countries. Outer space is the province of mankind, its complexities at national and international level as a result is governed by international law and constitutes an environ- of developments such as e.g. on-orbit transfer and outer space ment over which states cannot exert national sovereignty; their launches (landers on comets, launching from ISS), as well as failures to place in orbit remain a challenge. There is an ongoing need for coherency in application of the registration rules, This paper was presented at the 16th Reinventing Space Conference, and general compliance [4]. The duty to register space objects London, 30 October–1 November 2018. falls within the core duties of a state and cannot be left sole-

JBIS Vol 71 No.11 November 2018 431 LESLEY JANE SMITH ly to the growing private sector to implement. Given the in- 2.1 Can soft-rules of the road be enforced nationally as creasingly deteriorating state of the outer space environment, black letter law? this must not be lost sight of. Few efforts other than technical post mission disposal (PMD) rules have positively impacted on The increasingly ‘soft law’ requirements for participation in this environment. With more recent calls for new rules and a space activities and in particular the latest agreement towards a global space authority to manage space craft traffic manage- Draft Resolution on Long term sustainability Guidelines (June ment (STM), the impact of divergent approaches to national 2018) [14] may, paradoxically, drive the call for enforcement of space legislation relating to outer space – and its enforcement black letter rules one step further. The common denominator – requires attention. More states are looking to develop space among these rules cannot be repeated sufficiently. Some states, capabilities, as the increasing membership of COPUOS bears such as France, are already leading the way as regulatory role witness [5]. Recent national space statutes use different defini- models, by guaranteeing the upper tranches of operator liabili- tions to describe their scope. One such question is where outer ty beyond € 60M, and by imposing transparent technical rules space starts? Some national statutes refer to orbital activities implementing the international debris guidelines [15]. States including space data, others such as New Zealand to ‘high alti- are the only actors competent under international law to col- tude activities’ in its Outer Space and High-Altitude Activates laborate in outer space and oversee the enforcement of the Act 2017 [6]. Other states are discussing whether meso space applicable rules. Irrespective of the form of government-com- should be included [7]. Meanwhile, the UK defines space as mercial public-private-partnerships (PPP) or collaboration, sub-orbital, but has, for example, not specifically regulated the private commercial stakeholders can contribute, but cannot export of space data [8]. Such diversifications can influence dictate which rules apply in outer space. As commercial mar- market growth. As already achieved with many fields of inter- kets develop further, states are called upon to assert their su- national commercial law, approximating national space law on pervision and control over national entities, as is called for in the basis of a standard regulatory model remains a considera- Art VI OST. tion. More recent examples of national space statutes have recently given rise to discussion about their compatibility with International cooperation between states and the other ac- the international treaty system in the area of new space activi- tors in the field of outer space remains mandatory. Even if the ties such as resources mining [9]. The discussion about devel- time required for consensus building at COPUOS level, when oping an international mining regime, as outlined in Art. 11 (5) measured against the time taken to agree the Outer Space Trea- Moon Agreement, is still in early stages [10]. ty, is lengthy, more recent examples of international consensus-building, such as on the UN Convention on the Law of the States regulate their economic priorities with the scale of Sea (UNCLOS), have extended to well-nigh a decade. their national activities in mind. National space markets de- pend in part on their geo-locations, and include factors such 2.2 Public budgets and public perception as the interest in access to space, as well as the availability of national ground segment infrastructures. Now, with the po- Governments must ensure there is public understanding that tential offered by Artificial intelligence (AI) and robotics, the sufficient space budgets are essential to the delivery of societies’ feasibility of new forms of operations alters and new market twenty first century needs; in addition, there must be incentives players emerge. to ensure innovation. There is unlikely to be a fully viable economic model for addressing immediate issues, notably Active 2 CONVERGENCE IN ACCESS TO, ACTIVITIES WITHIN AND Debris Removal (ADR). However, ADR, as with security from RETURN FROM OUTER SPACE e.g. cyber interference, are essential investments, and therefore mandatory. Even in times of publicly-driven participation of If outer space activities are to remain safe and secure, whilst the private sector, the economic balance of choice must fall supporting increasingly advanced space operations, now is in favour of ensuring safety and security of outer space oper- the time to achieve a common position relating to access to, ations. The United States has a long legacy, notably through activities in and return from space. Cooperation relating to statutory-based government procurement schemes, of devis- Space Situational Awareness (SSA) is predominantly bilateral. ing economic incentives for its space industry, most recently, The sensitivities about data sharing are well known, and fore- in the form of the NASA Space Acts [16]. Burden sharing be- close any wider global SSA programme. Nevertheless, various tween public and private sector must be secured in proportion stakeholders show a willingness to share data, and commercial to the respective use of outer space, be this between scientific operators have devised their own data sharing platforms, such research, R&D, and commercial activities, including commu- as that of the Space Data Association, SDA [11]. If on-orbit nication and satellite-based location information for essential manufacturing, servicing, debris removal and in-situ resource services. Nevertheless, governments must meet this call first, utilisation are to become reality, with added support of quan- particularly for policy reasons. tum technology methods, then the focus should be on full consensus about the ‘outer space road map’. This applies all the 2.3 Lack of status more so, given the increasing numbers of cubesats with lack of manoeuvrability, and the challenges posed by mega constella- Outer space knows neither owner nor sovereign. It has no legal tions [12]. status and no formal defender, other than the community of societies dependent on the increasing number of space-based The continuous need for convergence between states’ ap- services, and wishing to see it protected. Like the High Seas, proaches to space also requires some reflection on how the var- outer space is res communis omnium. With no legal person- ious General Assembly (GA) Declarations and the Long-Term ality and only a few international principles addressing how Sustainability (LTS) Guidelines [13] can be enforced by states states are to interact in regard to cooperation, and notification against their commercially-driven sectors. This requires struc- of their national activities in outer space, the space commu- tured thought, including economic models, prior to developing nity alone must secure its own sustainability. Core interna- a more complex system of space traffic management (STM). tional treaty provisions, such as Art IX OST, embodying the

432 Vol 71 No.11 November 2018 JBIS THE NORMS OF BEHAVIOUR IN SPACE: Our space – Whose rules? principles of due regard, avoidance of harmful contamination, starting point would be to require operators to include deorbit- with Art. XI OST calling for international cooperation, can- ing technology, so that at least 98% s/c launched are equipped not alter the legacy of the community’s failure to secure the with de-orbiting capabilities. This model could be tied with environmental viability of outer space. This is why states must insurance-related solutions that allow bonds or guarantees to now deploy careful leverage against their own private sector. be placed and called up on default. Further economic models Treaty rules – with some few exceptions – only bind states, so should be discussed, including those for ADR. that national rules are also required to ensure compliance by the private sector. The res communis status dictates that, in its 2.6 Implications of increasing subscription to space treaty law activities, the private sector too must be required to uphold the public good interest of outer space for the benefit of all, and The UN space treaty rules, including the original precursor notably future generations. UN GA Resolution 1962 (XVIII) 13 December 1963, are the bulwark of outer space activities [20]. They provide the core 2.4 Paradox in legal hierarchy? framework for regulating all outer space activities, irrespective of whether states or their private national entities; this requires If states are also now to apply soft law rules in managing and robust states to manage and finance high-level risk, and to car- supervising their national space activities, can those very same ry potentially major losses. The current ratification status for states also be called to account for failure to enforce the applica- outer space activities shows increasing support, particularly for ble soft law rules against their private sector? The Draft Articles the four main space treaties, leaving aside the Moon Agreement on the International Responsibility of States for Wrongful Acts [21], first and foremost, the 1967 Outer Space Treaty. There is is a compendium of international rules, decisions and practice an increase in the number of states ratifying the Registration developed over time regarding the existence of international ob- Convention, REG [22]. REG is the most relevant treaty in re- ligations and the consequences of their breach [17]. It provides gard to debris issues; the state of registry retains jurisdiction; a sound basis for examining whether consensual rules can be- the lack of control over dysfunctional debris does not dilute come binding international obligations, and if so, when. What the letter of Art VI OST, which requires states to continuously level of state practice is required to recognise the development monitor and control activities of its national entities. Nor does or existence of a new, binding state obligation? Could this trig- a space object lose its status as a space object, merely because of ger claims against States for breach of state supervision under loss of control at end of life. Article VI Outer Space Treaty (OST)? These questions decry a simple assessment, but are a worthy exercise. Answers to such An increase in ratification can be noted for the Liability questions appear the only way to influence current attitudes to Convention [23]. While a lower level of support for the Moon the res communis and common space heritage. Attitudes to de- Agreement is generally given [24], this too is on the increase. bris creation were initially over-complacent. Up until now, causing debris has neither been found to constitute or be defined as The growth in subscription to treaty rules by entrant states illegal or a fault in law. Some ADR now appears essential, subject speaks first and foremost for their interest in complying with to requirements that include identification of object and related the treaty frameworks, alongside an interest in participation in debris. Further formal requirements appear in the LTS Guide- outer space activities. These statistics may also reflect the for- lines (2018) [18]. Generally speaking, there are no insurmount- mal calls by COPUOS for countries to pass national space leg- able legal reasons standing in the way of ADR; however, there is islation and to ensure registration of their space objects. little scope for efficient commercial models, at least for pre-existing debris [19]. Some form of penalty is required for failure 3 Liability Models to comply with post-mission disposal requirements, preferably in conjunction with an economic incentive model for the en- Liability is normally linked to commercial interests and failures tire industry to demonstrate commitment and burden-sharing. to maintain standards, but not so in space. International liabili- Currently, the over-arching responsibilities are carried by the ty for space affects states alone, and exclusively launching states. states themselves, both by virtue of Article VI OST, as well as National laws are required to lay down the risk sharing scheme the provisions of the Liability Convention [see below, 3]. between states and their non-government entities; these operate as incentive models, providing legal certainty. Whether or 2.5. Technological solutions combined with performance not there is a commercial-only interest in the particular opera- bonds or common funds as enhancement? tions depends on the space programme in question, along with the government policy towards incentive-creation. At contrac- Voluntary codes are frequently hailed as successful regulatory tor level, it is generally reflected both in accessing the market models, especially if there are identifiable incentives and trade- through the applicable procurement and contractual rules. Un- offs. While the trade-offs in space are clearly identifiable, see- su der certain government schemes – notably the US – the stand- pra, they are not equally distributed; high-risk operations bene- ard public law government rules grant benefits such as waivers fit society, and the majority of operations provide a high level of or relief from liability for government contractors [25]. The criticality, ensuring core services. A focus on incentive models statutory US commercial space operator liability model has would appear sensible. Downstream services are coupled with been a useful prototype, notably when developing the French an ever-increasing demand for connectivity and services such national space law. Operator liability in Europe is managed by as positioning, navigation, timing, (PNT). This has led to large a compulsory third party liability insurance coverage for out- satellite constellations, designed to meet the need for connectiv- er space operations over and above mandatory launch insur- ity and IOT. Yet these constellations raise their own distinct is- ance. France – and other European governments launching sues in relation to sustainability of the outer space environment, with ESA – underwrite the excess above the capped insurance particularly with the need for constant refreshing. amount of €60M. The model of state guarantee under French law enables to government to assume the third party liabili- States are the gateway to orbits and to frequency allocations. ty at the latest, one year after the licensing requirements are Without states, activities in outer space cannot take place. One met in full, or should have been met. This mixed model allows

JBIS Vol 71 No.11 November 2018 433 LESLEY JANE SMITH the government to share burdens and create incentives for the been allowed to develop over the past fifty years. While owner- private party in a manageable way. Further incentives or statu- ship of a space object can never be relinquished in law (aban- tory exceptions exist to enable industrial development, both in donment, unlike ships on the high seas, does not exist in outer France and the Unites States, for example, the privileged status space), the LTS Guidelines are moving towards creating proce- of experimental space flight, but are not expanded here. dures and clarifying parameters. Until the existence of debris is defined as sub-standard, debris-causation will continue to go The rationale of (state to state) liability for space operations with impunity. No form of liability – whether at international is that (in contrast to civil aviation) only states can bolster the or national level – will be imposed in the absence of clarifying unique – and unlimited – absolute liability for damage caused these rules of the road. If states do not enforce voluntary sus- by space objects, on earth and in air flight [26]. tainability obligations agreed at international level to ensure a future for outer space activities, what is their value, and who is One urgent call to be addressed relates to in-orbit damage. to be called upon for diligence in application? Considerations Are changes in insurance practice conceivable? Until now, very such as these are important in discussions relating to Space few satellite operations are covered by on-orbit insurance. In- Traffic Management (STM). surance practice has, in the event of damage, led to an absence of indemnity claims against the registered owner of the space 4.1 Technical standards and legal presumptions object – or the debris – causing the in orbit damage. Various factors have contributed to this approach; one is that parties Technical standards represent one way of managing com- carry their own risk in such hazardous operations, and if re- pliance with outer space rules, with ISO norms applicable to quired, should insure their own loss. The second relates to prac- debris, notably 24113:2011, going a long way towards setting ticalities, such as the difficulties in proving which debris caused global standards. Other areas of classic industrial law, notably the damage, let alone which is its state of registry. Addressing in manufacturing and products liability, operate with stand- this particular issue from a common ground would encourage ards, which trigger various risk allocation models. Reference workable solutions relating to debris-induced damage. One has been made supra to the market share rule. The applicable li- possibility could be a resort to the so-called market-share lia- ability regime is generally based on either fault or strict liability bility developed in the US courts. That principle was laid down systems, where trade-offs are balanced between manufacturers by judges when assessing risk allocation between market par- and insurers on the basis of compulsory insurance models. This ticipants when the individual perpetrator or tortfeasor could allows the operators to benefit from limitation of liability. The not be traced with certainty, but the exact number, identity and rationale behind fault liability for damage in outer space is that volume of business of companies active on the particular mar- parties accept and remain responsible for their own interests ket segment was well-known [27]. That principle required all in what continues to remain a seriously hazardous environ- market players to share a proportion of the risk. ment. Third parties are only accountable for damage to others, if at fault [30]. Parties are to bear their own risk in hazardous Liability allocation models have been developed to enable circumstances. No further substantive reasons are stipulat- market participants to bear risks equitably. In outer space, the ed in the pre-treaty working documents, namely the travaux common good relating to its environment should dictate that préparatoires, other than that the fault principle applies. all parties participating contribute to shouldering the burden of its future sustainability. Meanwhile, as insurance claims for Nevertheless, in the absence of clear rules of the road, fault space-related damage increase, some induced by space weath- liability for damage resulting from in-orbit activities will re- er, others by debris, further economic modelling is urgently main a paper tiger. This is where technical standards can offer needed [28]. relief. Failure to de-orbit according to the ISO norms dictating after-life de-orbiting capability could be managed with a pre- The unlimited liability of a launching state for damage sumption of fault rule. No alterations to the existing and well caused by a space object should not be confused with liabil- considered rules would be required. This would merely require ity for damage arising from space activities, for which other consensus as to interpretation, and require a shift in legal pre- variables can apply, but for which no practical solution has yet sumption, generally known as a paradigm shift. Transport and been discussed. Jurisdiction can always be claimed by national traffic management regimes operate with presumptions of fault courts; decisions from domestic tribunals are not excluded in (and with strict liability, if not also the lifting of capped liability the Liability Convention. The discussions about on-orbit dam- on wilful cause, where human life is involved). There are many age take on the characteristics of an alarm clock. It is a core valuable international liability regimes from which to extract element in crafting national regulations that include penalties the best risk-management regime. for lack of compliance with sustainability requirements. The unsustainable outer space environment does not necessarily 4.2 Governance fall within the concept of fault liability for damage resulting from a space object or its component parts. The community This paper has neither examined the ongoing discussions must look beyond this treaty rule, alter attitudes to the outer about a possible new authority to govern outer space traffic, space environment, and consider rules that are workable. No nor has it looked in detail at the subject of outer space resource cases have been reported relating to fault liability for damage in mining. Reference has been made above to the parameters for orbit, as opposed to s/c induced damage on earth [29]. Yet the national space laws, notably that these must conform to inter- outer space environment continues to deteriorate. national law, and that international treaty law does not provide any justification for assuming national sovereignty over celes- 4 CAN STM FOSTER DEVELOPMENT OF STRICTER tial bodies or their resources in outer space [31]. As regards ATTITUDES TO COMPLIANCE? the governance of resources mining, analogies are made to the Deep Sea Bed Authority, the institution ensconced within the If the soft law norms are enforced at national level, this might Convention on the Law of the Seas, an important international enable a break-out from the foregoing comfort zones that have document that took at least decade to find agreement. The Ant-

434 Vol 71 No.11 November 2018 JBIS THE NORMS OF BEHAVIOUR IN SPACE: Our space – Whose rules? arctic Convention also provides a model governed by interna- to adapt. This will allow specific monitoring and open up – as tional agreement and subject to a moratorium [32]. required – a further discussion of whether a separate international resources mining agreement is to be mapped out, as stip- In posing the question whether any change in governance ulated by the Moon Agreement. may be required for the future of space governance, the concept of cui bono – in whose interest – should be considered. In- Space needs to adapt its 20th Century models of internation- ternational government cooperation in outer space is manda- al consensus building to include technological and economic tory. Whether or not a new authority is required to undertake input. National governments must retain control of their na- oversight of the space traffic management rules (STM) that are tional space activities, with a view to aligning definitions, and now being called for, requires further assessment. There are if required, to cover excess risk. The basic principles of nation- sufficient authorities already endowed with legitimate compe- al monitoring and authorization, peaceful cooperation, along tence and trust. Additional issues arise with space traffic man- with technical incentives, must not be lost sight of. agement, requiring all aspects of access to, operations in and return from outer space, to be considered along with manage- Space is ours and the rules must reflect this. This could begin ment of data sharing, security and efficiency within. with a system whereby polluters pay, the exact model still to be discussed with the economists. This model could be bolstered As regards STM, those existing known and recognised bod- by extending the rule whereby states have a legal obligation to ies should be called upon first. Whether this should take place maintain the outer space environment as the common – and within an expanded ICAO, or a specific operative division of accessible – province and heritage of mankind. The application a trusted body such as COPUOS, need not be decided now. and interpretation of the states’ accountability for supervising its Governance should first come from national authorities - set- national operations under Art VI OST should be looked at more ting and applying standards for their national space industry, closely. For want of achieving consensus, these rules must be and enforcing these within an internationally acceptable mod- forward-looking, consensus-driven, and avoid pointing fingers el to which insurance or other economic models are required at those states which led the early eras of outer space activities.

REFERENCES 1. Treaty on Principles Governing the Activities of States in the Nov.25, 2015, Luxembourg Law on Exploration and Use of Outer Exploration and Use of Outer Space, including the Moon and Other Space Resources, available at http://legilux.public.lu/eli/etat/leg/ Celestial Bodies (Outer Space Treaty), London/Moscow/Washington, loi/2017/07/20/a674/jo (accessed 13 January 2019). For a detailed entered into force 10 October 1967, 107 ratifications and 23 signatures. discussion, see Hobe S., de Man, P., “National Appropriation of Outer 2. F. Lyall, P. B. Larson, Space Law, A Treatise, Elements of the Outer Space Space and State Jurisdiction to Regulate the Exploitation, Exploration Treaty as Customary Law, 2nd ed., 2016, 70-80, Ashgate. and Utilization of Outer Space Resources, 66 Zeitschrift für Luft-und Raumfahrtrecht, ZLW 3/2017, 460 3. Report of Committee on Peaceful Use of Outer Space, COPUOS, June 2018, A/73/20, available http://www.unoosa.org/oosa/en/ourwork/ 10. Agenda item no 14 of the Legal Subcommittee (LSC) session in 2019 copuos/2018/index.html. The progress achieved in June 2018 on provides for a general exchange of views and information on potential the Draft Guidelines for the long-term sustainability of outer space legal models for mining, see Report of COPUOS, 61st session General activities (LTS Guidelines) is contained in the Conference room paper Assembly, 20-29 June 2018, A/73/20, 5 July 2018, at p. 37. by the Chair of the Working Group on the Long-term Sustainability of 11. Fuller information in membership and activities is available at https:// Outer Space Activities, A/AC.105/2018/CRP.21, available http://www. www.space-data.org/sda / (accessed 13 January 2019) unoosa.org/oosa/oosadoc/data/documents/2018/aac.1052018crp/ 12. Report of Committee on Peaceful Use of Outer Space, COPUOS, June aac.1052018crp.21_0.html. The Guidelines on which consensus is still 2018, A/73/20, id., n. 10, Draft Provisional Agenda, Scientific and outstanding, notably Guidelines 20-22, are contained in AC.105/2017/ Technical Subcommittee, Agenda Item No. 12, Long-term sustainability CRP.23, see http://www.unoosa.org/res/oosadoc/data/documents/2017/ of outer space activities, p. 29. aac_1052017crp/aac_1052017crp_23_0_html/AC105_2017CRP23E.pdf (all accessed 13 January 2019). 13. COPUOS, Guidelines for the long-term sustainability of outer space activities, id. n. [3]. 4. For details of recent failures to register and complications regarding on-orbit transfer, see Ram Jakhu, Bhupendra Jasani, C. McDowell, 14. COPUOS, id. “Critical issues related to registration of space objects and transparency 15. French Law on Outer Space Activities, LOS, 2008, revised 2011, 2018, of space activities”, Acta Astronautica, 143 (2018) 406-420. The UN with technical regulations; for an English version of all rules, see, Clerc, General Assembly Resolution 62/101, A/RES/62/101 17 December Philippe, Space law in the European Context, National Architecture, 2007, “Recommendations on enhancing the practice of States and inter- Policy and Legislation in France, 2018 (Eleven Publishers governmental organisations in registering space objects” calls for higher 16. See NASA’s “Use of Space Act Agreements Audit Report”, National level of compliance. Aeronautics and Space Administration Office of Inspector General, IG-14- 5. The increase in committee membership of COPUOS since its creation, 020, 17 (2014). from 18 states in in 1958 to 92 in 2018, can be followed at http:// 17. International Law Commission, Draft Articles on the Responsibility of www.unoosa.org/oosa/en/ourwork/copuos/members/evolution.html States for Internationally Wrongful Acts, 53 UN GAOR Supp. (No. 10) (accessed 13 January 2019) at 43; UN Doc. A/56/10, 2001, Art 3: “The characterization of an act of a 6. Outer Space and High-altitude Activities Bill. http://legislation.govt. State as internationally wrongful is governed by international law. Such nz/bill/government/2016/0179/latest/DLM6966275.html (accessed 13 characterization is not affected by the characterization of the same act as January 2019). lawful by internal law”. 7. For further discussion, see T. Gangale, “How High the Sky? The 18. COPUOS, Guidelines for the long-term sustainability of outer space Definition and delimitation of outer space and territorial air space in activities, id. n. [3] international law”, vol. 13 Studies in Space law, 2018, Brill, Leiden 19. R. Williamson and L.J. Smith, (eds) 2019 “IAA Study Orbital Debris 8. The UK Space Industry Act 2018 received Royal Assent 15th March removal, Policy, Legal, Political and Economic Considerations” 2018. https://services.parliament.uk/bills/2017-19/spaceindustrybill. International Academy of Astronautics, Paris (not yet released). html (accessed 13 January 2019). See L.J. Smith, R. Leishmann, “Up, up 20. All UN space documents, GA Resolutions and the UN space treaties and away; An update on the UK’s latest plans for space activities”, in: Air are available online via UNOOSA, http://www.unoosa.org/oosa/en/ & Space Law, vol. 44, Issue 1, 2019, WoltersKluwer. ourwork/spacelaw/index.html (accessed 13 January 2019) 9. US Commercial Space Launch Competitiveness Act 2015, PL 114-90, 21. See “Status of International Agreements relating to activities in outer

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space as of January 1, 2018”, A/AC.105/C.2/2018/CRP.3, 4 April 2018. in the French Alps, Spring 2015, available in English in March 2016, (accessed 13 January 2019) available at https://www.bea.aero/uploads/tx_elydbrapports/BEA2015- 22. Convention on Registration of Objects Launched into Outer Space 0125.en-LR.pdf (accessed 13 January 2019) (Registration Convention), entered into force 15 September 1976, 1023 27. Sindell v. Abbott Laboratories, 26 Cal. 3d 588, 607 P.2d 924, 163 Cal. UNTS 15, 67 ratifications and 3 signatures Rptr. 132, 1980 Cal. LEXIS 151, 2 A.L.R.4th 1061 23. Convention on International Liability for Damage Caused by Space 28. See the latest Swiss Re Commercial Division Report, “New space, new Objects (Liability Convention), entered into force 1 September 1972, dimensions, new challenges; How satellite constellations impact on 961 UNTS 187, 95 ratifications and 19 signatures. Compared with that space risk”, 2018, available at http://www.swissre.com/library/expertise- the Agreement on the Rescue of Astronauts, the Return of Astronauts, publication/how_satellite_constellations_impact_space_risk.html and the Return of Objects Launched Into Outer Space, entered into (accessed 13 January 2019). force Dec. 3, 1968, 672 U.N.T.S. 119 shows 96 ratifications and 23 29. Settlement of Claim between Canada and the Union of Soviet Socialist signatures. Republics for Damage Caused by “Cosmos 954” (Released on 2 April 24. Agreement Governing the Activities of States on the Moon and Other 1981) available at http://www.spacelaw.olemiss.edu/library/space/ Celestial Bodies, entered into force July 11, 1984, 1363 U.N.T.S. 3, 18 International_Agreements/Bilateral/1981%20Canada-%20USSR%20 ratifications and 4 signatures. Cosmos%20954.pdf (accessed 13 January 2019) 25. 28 USC § 2680, Federal Tort Claims Act, as applied in various leading 30. Art. III Liability Convention, id. n. [23] judicial decisions, available at https://www.law.cornell.edu/uscode/ 31. S. Hobe, P. de Man, id. n. [9] text/28/2680 (accessed 13 January 2019) 32. The Antarctic Treaty, signed 1951, entered into force 1961. It enables 26. L.J. Smith and A. Kerrest, A. 2011 “Art. III Liability Convention scientific research to be carried out under the Treaty’s aegis, enabling (LIAB)” Commentary on the Convention on Liability for damage from international cooperation and without operations constituting formal Outer Space Objects, in Cologne Commentary CoCoSL vol II, Cologne, claims by the participating states. It stands as a potential prototype for Heymans. Limited liability is applicable under aviation convention law future models should discussion be taken up on mining regimes for the but is lifted on willful intent to cause the damage. See the final accident Moon. Various provisions reflect principles similar to those contained report by the French Bureau d’Enquete pour la sécurité de l’aviation in the Outer Space Treaty. Further details are available at https://www. civile, BEA relating to the German Wings passenger aircraft downing ats.aq/e/ats.htm (accessed 13 January 2019).

Received 21 November 2018 Approved 19 January 2019

436 Vol 71 No.11 November 2018 JBIS DIARY FORTHCOMING LECTURES & MEETINGS OF THE BIS

APOLLO MISSIONS: THE MECHANICS OF RENDEZVOUS & DOCKING BY DAVID BAKER 20 February 2019, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London, SW8 1SZ Starting with Apollo 9 launched on 3 March 1969, a key feature of the Apollo missions was the ability to rendezvous and dock in orbit – a capability that NASA had evolved over the preceding four years. SpaceFlight Editor David Baker describes the process in detail and casts an expert eye over the different options considered by mission planners in the run-up to the lunar landing missions. APOLLO 9 – RENDEZVOUS IN EARTH ORBIT 6 March 2019, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London, SW8 1SZ Jerry Stone continues his series of talks to celebrate the 50th anniversary of the Apollo missions with a uniquely personal take on the story of Apollo 9 – the first test of the full lunar landing package and only the second outing of the Lunar Module. WEST MIDLANDS BRANCH: A NEW SPACE RACE? & PROJECT CHEVALINE 16 March 2019, 1.45pm VENUE: BIS, 27/29 South Lambeth Road, London SW8 1SZ Gurbir Singh posits the beginning of a new space race between India and China, while John Harlow and Paul Jackman look back to the days of Project Chevaline and the famed Twin Chamber Propulsion Unit. APOLLO 10 – RENDEZVOUS IN EARTH ORBIT 22 May 2019, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London SW8 1SZ Jerry Stone continues his coverage of Apollo with the first flight to carry both the Apollo spacecraft and the Lunar Module on a full dress rehearsal of a landing. Call for Papers RUSSIAN-SINO FORUM 1-2 June 2019, 9.30 am to 5pm (tbc) VENUE: BIS, 27/29 South Lambeth Road, London SW8 1SZ The BIS has now scheduled its 39th annual Russian-Sino Forum – one of the most popular and longest running events in the Society's history. Papers are invited. Watch this space for further details. APOLLO MISSIONS: LANDING ON THE MOON BY DAVID BAKER 12 June 2019, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London, SW8 1SZ SpaceFlight's editor looks at the systems evolved by NASA for calculating optimum lunar landing trajectories, and at the descent procedures needed to achieve the maximum chance of success while preserving emergency abort and safety considerations. Journal of the British Interplanetary Society

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