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IAC-14.D2.4.5

A Technical Overview of a Based European Launch Service Operator

Mark Hempsell, Reaction Limited, Building D5, Culham Science Centre, Abingdon, Oxon, OX14 3DB , [email protected]

Julio Aprea1, Roger Longstaff2, Giorgio Ferrari3, Steven Hens4, Sebastian Soller5, Roger Dewell6, Dan Ahearn7, Catriona Francis8

Between 2012 and 2014 an industrial consortium led by conducted a feasibility study for the with the objective to explore the feasibility of SKYLON as the basis for a launcher that meets the requirements established by ESA for the Next Generation European Launcher. SKYLON is a fully reusable single stage to launch system under active development. The purpose of the Study, which was called SKYLON Based European Launch Service Operator (S-ELSO), was to support ESA decision making on launch service strategy by exploring the potential implications of this new launch system on future European launch capability and the European industry that supports it. The launch operator requirements centred on geostationary transfer orbit missions and the Study’s main technical focus was on producing concept designs to demonstrate the feasibility of a complete launch infrastructure consisting of SKYLON, a reusable upper stage, and carriers. The requirement was for S-ELSO to operate from Centre Spatiale Guiana. The Study showed that the provision of new facilities, like SKYLON servicing buildings and a 5.9 km runway, and the links with existing CSG services such as payload preparation and propellant supply were feasible. The study showed an operational infrastructure fully able to meet European launch system requirements could be operational by 2024.

I INTRODUCTION and an altitude of 28 km, at which point the can switch to a staged combustion pure mode SKYLON using as the oxidiser.

SKYLON (Figure) a fully reusable single stage to orbit that can take off from a runway reach with a payload of 15 tonnes at 300 km altitude. Once the mission is completed then it returns to earth for a runway landing [1]. It is under active development and is planned to reach operation in the early 2020s. SKYLON is the result of 30 years of technology development and design studies. It is based on an air-breathing engine concept called Figure 1: The SKYLON Spaceplane SABRE, which uses a combination of a pre-cooler heat exchanger to cool incoming air, and a turbo- The technical feasibility of the SABRE compressor to raise the air pressure high enough to engine was primarily centred on the pre-coolers and then be fed as the oxidiser into a whether they could be made for the mass required for combustion chamber to be burnt with liquid a flight system. The demonstration of the technology hydrogen. The air-breathing mode of the SABRE was completed in 2012 with the completion of a test engine can be sustained to a little beyond Mach 5 programme on a test heat exchanger using flight

1 European Space Agency, , [email protected] representative modules (Figure). The impact of the 2 Reaction Engines Ltd, United Kingdom, [email protected] SABRE engine on the overall system is that the 3 , Italy, [email protected] mass fraction needed to achieve orbit in a single 4 QinetiQ Space Nv, Belgium, [email protected] 5 Defence and Space, Germany, Sebastian.Soller@.eads.net stage is raised from around 13% for a pure rocket 6 Grafton Technology Ltd, United Kingdom, [email protected] vehicle to around 23%. 7 42 Technology Ltd, United Kingdom, [email protected] 8 Jacobs Engineering UK Ltd., United Kingdom, [email protected]

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The Study was entitled S-ELSO (SKYLON based European Launch Service Operator) and was split into two separate activities, one exploring the technical and schedule aspects and the other exploring the business aspects both financial and economic. It was started in 2012 and the final report was delivered and final presentation held in June 2014. This paper gives an overview of the technical conclusions of S-ELSO, a separate paper reports the conclusion of the business study [3]. Figure 2: The Technology Demonstration Pre-cooler on the Test Stand II S=ELSO REQUIREMENTS SKYLON Based European Launch Service Operator As part of its on-going work in launcher development ESA generated a set of requirements Given that SKYLON is a venture based in the United for a next generation European launch system. Kingdom, it is natural that the British Government These requirements were published as a through UK Space Agency should take an interest in Specification that was included in the Request for the potential of the concept and its technologies. Proposal for the Feasibility Study for a New Although the majority of the funding has been European Launch Service (NELS) [2]. These through private equity investment a significant requirements were redrafted for the S-ELSO study public contribution has been made through ESA [4] in part to clarify the requirements where they technology development programme. Further the were ambiguous and also to reword some UK Space Agency has also placed direct consultancy requirements which had wording applicable to an contracts with ESTEC to perform technology to wording that was more evaluation of both SKYLON and SABRE. in line with a solution. It was with this background that Reaction Overall these requirements represented a Engines with an industrial team of 5 subcontractors minimum capability to sustain autonomous European and 2 supporting consultancy teams (Table 1) made a operations and that would reduce that capability from proposal to the ESA Launcher Directorate to perform that currently available with 5. This a study which would assess the degree to which a minimalist approach meant large LEO and SKYLON based commercial European launch human capability were not included. The service provider could meet the requirements that S-ELSO Study’s technical goal was to compare that ESA had generated for the Next Generation minimum capability as defined by the specification European Launcher [2]. The Study was designed to with the capability provided by the proposed S- examine the technical schedule and financial ELSO infrastructure. All the requirements were met feasibility against that specification with the purpose and the overall capability was well matched, but it of supporting ESA decision making on launch should be emphasised that the S-ELSO infrastructure service strategy by exploring the potential would inherently have other capabilities that were implications of this new launch system on future not included in the specification. This would mean European launch capability and the European that with S-ELSO Europe would not lose any overall industry that would support it. capability with the introduction of the post- system, but instead gain new capabilities, such as the Table 1: S-ELSO Study Team ability to return payloads. Company Study Role Technical Study The dominant requirement was for a 6.5 tonne Reaction Engines Prime payload capability into Geostationary Transfer Orbit Airbus DS Subcontractor SOMA Engine Grafton Technology Subcontractor (GTO). A payload capability to an 800 km Sun QinetiQ Space Subcontractor Payload carriers Synchronous Orbit (SSO) of 4 tonnes was also Thales Alenia Space Subcontractor Upper Stage Jacobs Consultant Spaceport defined. Operations in Medium Earth (MEO) 42 Technology Consultant Payload Interfaces and Low Earth Orbit (LEO) were required but no performance was specified. The requirements were Business Study Reaction Engines Prime supported by a mission model showing frequency the London Economics Subcontractor of the various missions included in the Specification and the likely mass distribution of GTO payloads.

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The payload accommodations in terms of available envelop and loads were set to be similar to Ariane 5 but with a reduction in shock loads. The original NELS specification adopted a similar size as Ariane 5 approach to reliability, availability and responsiveness. However in the context of a reusable launcher these values are extremely low and also do not cover the ability of a reusable system to abort and return intact. So the S-ELSO specification was revised to offer higher values in all these areas as shown in Table 2

Table 2: SKYLON Operational Characteristics Current ELV SKYLON Statistically proven Mission Abort n/a <1/100 on 400 test flights Due to the impact of Payload Loss 1/50 – 1/70 <1/20,000 full abort capability Driven by the Turnaround months 2 days aeroshell inspection Could be days if Responsiveness > 1 year weeks operator required it Full launch Flexibility Limited Yes campaign in less than a week Reattempts On time Limited Yes opportunities hours apart

The NELS requirements defined a first flight in 2020, but foresaw a 3 to 4 year transition to full operation. This would be the pattern for an expendable system and where launcher was an exclusive partnership between the manufacturer and the operator. In the case of S-ELSO, SKYLON has to undergo an extensive test flight programme before it is certified as operational and, because S-ELSO would be only one of many customers, it is not possible to exploit these for commercial gain. Thus the S-ELSO specification defined the date of operational service as 2024. Other requirements included;  a lifetime of 20 years,  costs to be competitive with and ,  operation from CSG (),  low environmental impact,  European autonomy

III S-ELSO INFRASTRUCTURE

SKYLON was designed to only reach LEO, whereas the key requirement for the next generation European Figure 3 – S-ELSO Flight Elements launch system was to launch payloads into GTO, so in order to create a complete infrastructure able to The Study identified a need for two payload meet the customer’s defined needs, additional carriers, one for small payloads and another for elements. such as upper stages. were required to larger payloads into Low Earth Orbit, and an upper compliment the basic SKYLON. stage that can reach orbits beyond SKYLON’s 600 km maximum altitude. Unlike SKYLON, none

3 of these systems had been taken beyond preliminary the purposes of the Study the second decade was concept work before the S-ELSO Study, so the assumed to be the same as the first. Given the examination of these additional systems was the key capacity of the S-ELSO infrastructure further growth focus of the technical work. The flight systems as outside of the model could be easily accomplished, defined by the study are shown in Figure 3. with the likely constraint being the supply of SUS Another new area of study was the stages. There were three different scenarios in the consideration of Centre Spatiale Guyanais (CSG) at model: pessimistic, nominal and optimistic. Table 3 Kourou, as the operational centre for gives the assumed numbers of flight elements S-ELSO. Previous spaceport work on the SKYLON required to meet this model. programme had considered a generic launch site assuming all the facilities would be new and without Table 3: S-ELSO Flight Inventory regard to practical issue that a real site may have. So this was also the subject of detailed examination by System No. Capability Requirement Notes 2 for redundancy the Study. 79-197 SKYLON 2 400 flights Req. range pess. to Thus it was envisaged that S-ELSO would be Missions opt. a pure transportation enterprise, in the sense that it 60 flights 58 missions SUS (Pess.) 6 would take customer’s payloads, and deliver them (6 expend) (4 expend) into the required orbit using the assets that comprise Expendable SUS 110 flights 108 missions mission are required SUS (Nom.) 11 its infrastructure. The S-ELSO requirement was for (11 expend) (8 expend) for GTO over a complete infrastructure to undertake the missions 6.3 tonnes. 170 flights 165 missions SUS (Opt) 17 specified in the requirement specification. The full (17 expend) (11 expend) list of elements that were identified for the complete this infrastructure is shown as a system breakdown in SUS ASE 2 - - 2 for redundancy Figure 4. The additional element that would be 2 can be flown at once required for support of activities SSPC 4 - - + one set for are labelled as LEO Servicing and are shown faded redundancy as they were not required to meet the specification SLPC 2 - - 2 for redundancy and were not examined by the study.

IV SKYLON D1

Airframe

The S-ELSO Study used the D1 configuration of SKYLON as shown in Figure 5.

Figure 5: SKYLON Layout

The airframe consists of a slender containing the propellant tanks and the payload bay, Figure 4: S-ELSO System Breakdown with a wing located roughly midway along the

fuselage. The SABRE engines are mounted in Although the lifetime was set for 20 years, the flight axisymmetric nacelles on the wingtips. Control traffic model was only thought valid for the first authority whilst in the atmosphere is exerted by decade, but in the absence of any better analysis for foreplanes in pitch, ailerons in roll and an aft

4 mounted fin in yaw. Yaw control during the rocket SABRE Engine ascent is achieved by differential engine throttling. During the rocket ascent, main engine gimballing The SABRE 4 engine used in SKYLON D1 which takes over pitch control progressively as the dynamic was a revision of an engine concept that could pressure reduces, until finally it hands over to operate in air- breathing mode using LH2 as reaction control thrusters at Main Engine Cut- Off propellant from take-off to a transition point at (MECO). The reaction control thrusters retains the around Mach 5, and then convert to pure rocket control authority until a progressive handover back engine using LH2 and LO2. The rocket transition to the foreplanes, ailerons and tailfin during re-entry was designed to occur at 28km. to the Earth’s atmosphere. In air-breathing mode, a pre-cooler heat The SKYLON vehicle was found to be very exchanger cools the incoming air, so it can then be sensitive to the factors affecting Centre of Pressure compressed to high enough pressures to be fed into a (CP) and Centre of Gravity (CG) and the rocket combustion chamber. The heat is extracted by configuration was strongly driven to ensure that pitch a loop, which uses the energy extracted from control authority was maintained throughout all the air (hundreds of MW) to power the compressors flight phases. This was achieved through a and propellant pumps. The helium loop then uses combination of careful attention to the overall the liquid hydrogen fuel as the heat sink in the aerodynamic shape and mass distribution; sizing of thermodynamic cycle. aerodynamic control surfaces; and differential burn- off from the two hydrogen tanks. Part of the solution Typical Mission Profile to the trim issue was to mount the payload as far forward in the bay as practical. Two mounting SKYLON was designed to be operated in a very provisions were included; one 3m aft of the front of similar manner to an aircraft. Payloads would be the bay, and the other 3m forward of the rear of the integrated into the payload bay from above through bay. However in light of this mass properties the payload bay doors; a process very similar to that constraint, only the forward mounting could carry a used to integrate payloads into the NASA Space full payload, meaning that payloads tend to face Shuttle. Once the payload was installed, the vehicle backwards, and would experience primarily negative would be towed out to a refuelling ramp located at longitudinal accelerations as a result, compared with the end of the runway. The hydrogen, oxygen and conventional launch systems. helium propellants would then be loaded and the SKYLON’s main structure design consisted vehicle moved to the roll start point on the runway of a space frame constructed from struts made from itself. titanium with silicon carbide fibre reinforcement. After all pre-flight checks have been The non-structural aluminium propellant tanks were completed the vehicle would start its engines and suspended within the framework by Kevlar ties. The verify full and nominal operations. Mission frame was covered with sheets of a reinforced glass control would then give the launch command and the ceramic material which acted as the aeroshell and vehicle would release its brakes and accelerate along main thermal protection backed by a multilayer the runway. At the take-off speed (155 m/s) the metallic . vehicle would rotate and would be committed to In addition to the main propulsion system flight. In the event of a malfunction during the take- tanks there were a set of secondary cryogenic tanks off run the vehicle would close down all remaining in the nose and tail areas which fed the orbital propulsion and brake to a halt on the runway. In the manoeuvring engines, called SOMA (SKYLON event of a malfunction after the decision point Orbital Manoeuvring Assembly), the reaction control (which commits to the take-off) the vehicle would thrusters and the fuel cell power supply. dump propellant under powered flight and then The SKYLON high level mass breakdown is return to the runway for a landing under gliding shown in Table 4. flight. Vehicle systems were designed to allow a safe recovery of the vehicle with a complete engine Table 4: SKYLON Mass Breakdown nacelle failure immediately after take- off. Item Mass The vehicle would follow a climbing, lifting Dry Vehicle (inc. margins) 53.4 tonnes Consumables (inc. residuals and aux. propellants) 6.5 tonnes and accelerating trajectory using the engines in air- Usable Ascent Propellant 250.1 tonnes breathing mode up to a speed just above Mach 5 and Nominal Payload 15.0 tonnes an altitude of 28 km. At this point the engines Gross Take Off Mass(GTOM) 325.0 tonnes transition to pure rocket mode and the vehicle continues to climb and accelerate to orbital velocity. At the end of powered flight the engines throttled

5 back to limit the axial acceleration to 3g. After Main and one keel) was retained but the trunnions enlarged Engine Cut-Off (MECO) the vehicle would make a and hollowed, the new design being optimised for small ullage burn using the SOMA orbital the nominal 15 tonnes payload. The new trunnion manoeuvring engines and dumps all residual designs are shown in Figure 6. propellants from its main tanks. At MECO SKYLON would be in a transfer orbit with an apogee equal to the altitude of the required circular orbit. At apogee the SOMA engines burn again to circularize the orbit. Orbital operations would begin with opening the payload bay doors. For missions that involve payload delivery to LEO, the orbital phase of the mission would be completed with the deployment of Figure 6: Sill (left) and Keel (right) Payload the payload. However, for missions involving the Mounting Trunnions use of the reusable SKYLON Upper Stage (SUS),

SKYLON would wait in orbit for around a day while the SUS completes its mission. The SUS and The mounting on the payload side consisted SKYLON would then perform a rendezvous and of two sill trunnions taking loads along the X and Z docking operation, and once this is completed the axes, and one keel trunnion taking loads along the X and Y axes (Figure 37). All the trunnions were SUS would be reinstalled into the payload bay. titanium with a chrome surface finish; 150mm in After orbital operations have been completed diameter; and had a 154 mm diameter sphere at the the vehicle would close its payload bay doors and prepare for re-entry. At a pre-calculated time the end which creates the actual point of contact within vehicle performs a retro-burn with its SOMA engines the sleeve so angular movements to do not induce moment loads.. and begins its descent to the spaceport. Re-entry The keel trunnion design incorporated a interface would be passed at an altitude of 120 km simple electrical connection that supplied 28 V and the vehicle would manoeuvre in bank and angle power and a simple command signal/databus of attack in order to control temperatures and heat loads, and in order to meet the pre-calculated provision. These connections would be made downrange and cross range requirements necessary automatically as the trunnion is inserted into its sleeve. It was intended that there would also be a for the return to the spaceport. The vehicle would high speed data link on a separate connector, but the finally enter a gliding approach and landing schedule study did not address this. that would be almost identical to that of the . All flight hardware would then be The SKYLON design provided two trunnion inspected, serviced and prepared for the subsequent mounting locations in the payload bay, flight. approximately 3 m from each end. These locations were designed to be identical, with the following

exceptions: Payload Interfaces  The forward interface was designed for One area of the SKYLON airframe that was refined 17 tonnes and the rear for 3 tonnes as part of the S-ELSO Study was the payload  Only the forward interface has a payload interfaces. There were two reasons for reconsidering ejection and return capability at the interface design. The first reason was that pre-  Only the forward interface has provision study work by 42 Technology Limited suggested the for LH2, LO2 and LHe supply previous design as defined in Issue 1 of the The payload bay’s forward location has the SKYLON Users’ Manual [5] could be optimised to ability to deploy the payload once in orbit. The both reduce mass and reduce the load coupling. This mechanism to perform this function was examined design goal of the interface only carrying the payload during the S-ELSO study and an electrically driven inertial loads was considered important as it was ball screw linear actuator was selected to push on the hoped that the need for coupled structure analysis of sill trunnion (one on each side), providing a force of payload and launcher could be eliminated. The approximately 2 kN acting over 200 mm, providing a second reason for the design change was that the separation velocity of 0.5 m/s. The mechanism was previous design was known to be seriously under designed so that payloads could be placed back into strength when used for payload integration. the mechanism and locked down in orbit for return to 42 Technology Ltd undertook the redesign Earth. activity, the three trunnion mounting system (two sill

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The S-ELSO Study also looked at methods of customers. The use of USIS also allows operational integrating the payload into the bay, as the earlier and performance advantages; for example, the SUS proposed solution was inadequate. The payload USIS that is used as the payload interface is also the would be inserted into the payload bay from above means by which it is recovered in orbit when its through the open bay doors by a crane and an mission is complete. integration frame. The main connection to the frame Although there are currently many docking is by the sill trunnions but as the frame could not berthing systems and launch systems, none are an reach the Keel trunnion a special provision was accepted standard, and none meet the requirements required to carry the moment loads. Two options that have been identified for a common universal were examined: Option A was an integration interface that can be both a docking berthing system trunnion mounted on top the payload that attaches to and a launch system interface. These requirements a sleeve in the mounting frame, whereas Option B are defined in the “USIS Technical Requirement was to have pairs of integration trunnions in line with Specification” [7]. the sill trunnions which would attach to a moveable The key features of the USIS standard as beam that is part of the integration frame. These two specified for the study were that it should be able to options are illustrated in Figure 7. carry a 10 tonne payload with a centre of mass 2.5m above the interface ring, with inertial accelerations derived from consideration of the Falcon 9, Ariane 5 and SKYLON launch systems. It is also was required to withstand a pressure load of 200kPa while connecting masses of around 100 tonnes in orbit and 30 tonnes on the Martian Surface. The requirements called for the USIS to make a permanent connection (for integration), a breakable mating connection (for berthing), and a capture connection (for docking). The specified alignment and connection loads were in line with the published requirements for the International Docking System Figure 7: Integration Frame Showing Option A - Standard (IDSS) [8]; a public standard defined Top Trunnion (left) and Option B - Side Trunnion jointly by NASA, ESA, , CSA and (right) JAXA. The USIS is used on two S-ELSO elements; on the SKYLON Large Payload Carrier (SLPC) as the means to deliver and recover LEO payloads of up V COMMON ELEMENTS to 10 tonnes and on the SUS as both the payload

attachment and stage recovery system. Earlier work Universal Space Interface System had not produced a fully worked through USIS

design that could support the design activities of Past SKYLON studies have used a concept that these two elements so as part of the Study QinetiQ originated in the infrastructure studies around Space used their experience on the European version HOTOL in the 1980’s (and also early NASA Space of the IDSS to produce a USIS concept design based Station work) which is to adopt a common interface on the IDSS technology and design approach. between space systems [6]. In the form that has The QinetiQ USIS concept (Figure 8) exploits emerged from the SKYLON support studies, it is Stewart Platform mounted capture ring technology, called the Universal Space Interface System (USIS). not only to reduce the loads during the capture The USIS was envisaged as being the process, but also to play a part in meeting the interface that would be used by virtually all of the misalignment requirements. Thus the capture system payloads defined in the S-ELSO requirement operates as an active platform that is steered with the specification. These are all conventional supporting linear actuators. The platform actively types that currently would use a structural ring aligns during capture of the mating vehicle. To do connection, typically utilising a Marmon clamp so, the relative position and orientation of the release mechanism. However, the new operational vehicles would be obtained from the vehicles’ environment created by SKYLON means that the guidance and navigation control system. This “one-shot” mechanism is less appropriate, and a reduces the size of the capture guide vanes, helping mechanism that can allow recapture of payloads in to meet the passageway requirements while keeping orbit for return to Earth is a capability that should be the ring diameter down to 1800 mm. included as part of the S-ELSO package to its

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SOMA Engine

In addition to the main propulsion provided by the SABRE Engine, SKYLON has a secondary propulsion system which is used not only in parallel with SABRE during its rocket ascent but also used to perform , orbit change and de-orbit manoeuvres. This SKYLON Orbital Manoeuvring Assembly (SOMA) uses liquid hydrogen and liquid oxygen propellant from dedicated auxiliary tanks mounted alongside it in the tail of the vehicle. It had always been the intention that the SOMA engine would also be used as the engine for a complimentary reusable upper stage which in view of the high cost of liquid rocket engine development Figure 8: QinetiQ USIS Concept was seen as the best route to make reusable upper stages acquisition costs low enough to make them The QinetiQ USIS concept uses 12 active economically feasible. hook mechanisms to make the structural connection. The previous SKYLON studies only used a The loads on the mechanism were very similar when parametrically derived overall concept to obtain carrying 10 tonnes unpressurised and when carrying performance, mass and other interface estimates for the pressurised loads, confirming the close match the engine and had not considered the SOMA design between these two cases (which was the fundamental in detail. In view of this much higher importance of insight that led to the concept of a common universal the engine in the SUS concept the S-ELSO Study connection standard). produced a more complete engine concept design. A key feature of the USIS concept was that it This work was conducted by Airbus Defence and was intended to have many variations with different Space at Ottobrunn. levels of complexity and functionality. There would Starting from SOMA system requirements, be both pressurised and unpressurised versions with several engine cycle concepts have been assessed different capabilities from permanent or semi- and optimised; these include both expander and permanent connections, through berthing staged combustion cycles. This optimisation study connections to full docking connections. The showed that none of the configurations examined objective was to produce a technical definition for could simultaneously meet all requirements for the USIS that both meets the requirements and , engine mass and maximum allows for a wide range of different technical diameter of the nozzle extension. Together with implementations, including simple and light-weight Reaction Engines it was decided to relax the designs, while retaining universal connectivity requirement on the installation volume of the engine. between them. QinetiQ produced three This decision allowed for the design of a high- unpressurised USIS variations for the S-ELSO study pressure expander cycle engine using two thrust a completely passive docking ring (used on the chambers that met all of the remaining requirements. payload), a light version that utilised a Marmon Clamp one shot deployment which was used on the SUS, and a fully active docking system which was used on the SLPC. While the work to refine the USIS will continue, the requirement generation now needs to focus on the one-shot release requirement, the importance of which has been highlighted by the S- ELSO study. It will also need to re-examine the technical approaches for its implementation with an emphasis on reducing the mass, and decreasing the complexity and cost.

Figure 9: Two SOMA Engines as Installed in SKYLON

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The selected cycle was an expander cycle iv. The dry mass of the Reaction Engines SUS with a single turbine powering both the oxygen and was estimated at 1500 kg and it was hydrogen pumps, and a gearbox to account for their designed to carry 7500 kg of propellant. different speed requirements. In order to meet the These four key features are also incorporated SKYLON and SUS packaging requirements the into the SUS concept design produced as part of the pumps fed two thrust chambers. The overall S-ELSO Study by Thales Alenia Space. However, configuration was designed so that two SOMAs changes to the SOMA design resulting from Airbus’ could be mounted to produce a four nozzle cluster as work, and differences in technology assumptions shown in Figure 9. SKYLON D1 employed this two resulted in a very different overall design concept as engine cluster, whereas the SUS used a single shown in Figure 10. SOMA twin chamber engine. The key SOMA engine features are given in Table 5.

Table 5: SOMA Engine Key Parameters (per twin chambered engine) Thrust 40 kN Chamber Pressure 90 bar Mass 102.5kg Throat Diameter 39.1mm Specific Impulse 4562 Nsec/kg Mixture Ratio 5.2: 1 Expansion Ratio 285:1 Total Length 1328mm

VI SKYLON UPPER STAGE (SUS) Figure 10: Thales Alenia Space SUS Concept Design The SUS System Design The SUS is reusable for up to 10 missions SKYLON on its own can only place payloads into which was set through consideration of the market, LEO. To reach the higher orbits and Earth escape which determines the balance of missions that can be orbits specified for S-ELSO requires an upper stage. launched in reusable mode and how many require the The Study assumed development of only one upper extra performance of the SUS in expendable mode. stage that would provide a full launch capability at To recover the SUS it would be launched together entry into service; this is called the SKYLON Upper with a recovery system in the rear payload mount Stage (SUS). The SUS is optimised to provide the that can dock with the SUS and re-install it in the maximum payload into geostationary transfer orbit. forward payload mount, as shown in Figure 11 However, it can also deliver effective payloads to all high Earth and planetary escape orbits. Before the S-ELSO Study, Reaction Engines studied the SUS and produced a concept design which is outlined in its final form (Configuration B2) in Reference 9. Four key features of the SUS embodied in the REL design were: i. It was reusable, thus extending the economics of reusability beyond SKYLON to the entire launch system. ii. It use of the SKYLON SOMA engine with a Specific Impulse of 4,560 N sec / kg to reduce the development cost. The financial analysis of the SUS conducted for the S- ELSO study confirmed this approach in the Figure 11: SUS with Payload and Recovery ASE context of the defined mission model. Installed in SKYLON’s Payload Bay. iii. The payload would be carried on a USIS, which in turn would be engaged by the The SUS has an aluminium primary structure recovery system to enable it to be returned which mounts on SKYLON’s forward payload to the payload bay and brought back down attachment. It consists of four main assemblies: to Earth after completion of the mission. i. Payload support assembly

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ii. Tank support assembly realization of the tanks. The thickness over the iii. Engine thrust frame middle of the shells is ~1 mm, whereas the thickness iv. Truss assembly of the external area of the shells (where the loads are The Payload Support Assembly consists of a concentrated and around the welds) is assumed to be 30 mm thick CFRP skinned sandwich panel which around 2 mm. The resulting mass was found to be has an aluminium machined payload adapter ring 217 kg. which reinforces the panel at the point where the The ullage gas is helium which was stored as lightweight USIS interface is mounted. The USIS liquid in four off-the-shelf tanks, each with a mass of provides the structural connection with the payload 8 kg and mounted at the base. The tanks fed a single and is also the docking connection with the SUS SOMA engine which is identical to the engine used recovery system that is mounted in the SKYLON by SKYLON’s orbit manoeuvring system. payload bay to effect recovery for a return to Earth. is achieved with 24 cold gas The Tank Support Assembly’s main thrusters using the hydrogen boil off. The hydrogen component is a honeycomb sandwich bulkhead and oxygen were also used to supply the fuel cells panel. The three attachment trunnions to the which supply 2 kW of power. The fuel cells are SKYLON are connected to the panel via a titanium supplemented by Li-ion rechargeable cells to handle bracket that is integral with the panel, as is the peak power loads. propellant tank support ring that provides the The system mass budget is shown in Table 6. interface for the H2 and O2 tanks. This shows the level of margins that would normally The Engine Thrust Frame structural assembly be applied to an early feasibility design resulting in a transmits the thrust from the SOMA Engine to the total of 24% between the raw estimate and the final SUS Structure. It is an aluminium conical structure system mass estimate when all the margins were that mounts the engine at the narrow end and is then combined. connected at the wide end to the Engine Thrust Frame. This assembly includes the interfaces for the Table 6: SUS Mass Budget four helium tanks on a small CFRP panel placed Without % of Dry Mass Contributions Margin Margin Total Total Dry between the engine and the Engine Thrust Frame and kg % kg % by four pairs of struts. Structure 313.60 7.07 335.78 33.98 The Payload Support Assembly and the Tank Thermal Control 58.00 5.00 60.90 6.16 Communications 2.00 5.00 2.10 0.21 Support Assembly are connected together by means Data Handling 38.00 5.00 39.90 4.04 of a Truss Assembly which is composed of eight AOCS 16.00 5.00 16.80 1.70 GNC 3.00 5.00 3.15 0.32 machined aluminium I section beams and eight Propulsion 401.50 13.73 456.63 46.21 machined aluminium angle sections. The aim of Power 50.00 10.00 55.00 5.57 Harness 15.00 20.00 18.00 1.82 these beams is to stiffen the overall SUS structure Total Dry 897.10 988.26 kg and to transmit the thrust to the front interface. All (excl. adapter) System margin 20.00 197.65 kg of the beams and angle sections are independent (excl. adapter) elements that attach to the panels by means of Total Dry with margin 1185.91 kg aluminium machined brackets. (excl. adapter) Wet Mass Contributions Enclosed within the truss framework were the Propellant (losses + He) 215.00 0.00 215.00 hydrogen and oxygen tanks which have a common Adapter mass 112.00 10.00 123.20 bulkhead. The tanks were constructed from Launch mass (including adapter and payload) 1524.11 kg Aluminium-Lithium alloy 2195, selected because its strength (especially at low temperature) and low SUS Mission Design density (2685 kg/m³). This material was used for the shells but it was not available for the big machined All the key missions defined in the S-ELSO ring so for this element Aluminium 2219 was 3 Requirement Specification require the combination chosen, which is slightly denser at 2840 kg/m . of SKYLON and the SUS, so the performance of this The preliminary mass estimation was combination is the critical factor in demonstrating performed on the tanks considering an average shell that the proposed S-ELSO infrastructure can meet thickness of 1.5 mm. The reason behind this the future European launcher needs. This required assumption is that the low pressure inside the tanks the study to identify missions to GTO, MEO, Escape and the low loads induced by SKYLON allow the and SSO. utilization of the thinnest manufacturing limit for the

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Figure 12: Low Orbit Deployment Strategy; GTO Mission (9:1 Resonance)

Of the many possible approaches to reach To meet this shortfall and to take the GTO GTO while recovering the SUS, the Study found the delivered payload beyond the requirement the SUS best approach was to launch the SUS and its payload could be used in an expendable mode, that is without from a very low altitude of 185 km. SKYLON’s the propellant to return to SKYLON, and without the payload to this orbit is 16.5 tonnes, which after an estimated 1050 kg of ASE in the SKYLON payload allowance for recovery equipment installed in the bay that is required for docking operations. In this rear payload position means the SUS/payload stack mode the SUS could deliver 8 tonne satellites can have a mass of up to 15.5 tonnes. It was directly to GTO. calculated that SKYLON can maintain an orbital Although not included in the Requirement altitude of 185 km with minimal extra RCS Specification the current interest in using on board propellant usage – in the order of 10 kg per day - to electric propulsion to reach the final Geostationary compensate for atmospheric drag. Thus SKYLON orbit destination (as opposed to convention chemical can remain in orbit to await recovery of the SUS at propulsion) caused ESA to request the Study to also the end of the mission, a process that would need to look at this approach. The Study selected a strategy be accomplished within 2 days from the initial which involved placing the into a 5900 km launch. altitude circular orbit. This higher starting altitude The mission proceeds as shown in Figure 12. has two significant benefits: it would avoid the lower In order to effect an eventual rendezvous of the SUS Van Allan belt and also it would significantly reduce with SKYLON, a 9:1 resonance SUS mission was the transit time to GEO. The payload of SKYLON to examined (there were no solutions found at 8:1 a 300 km circular orbit would be 15,000 kg, resonance). In this case the SUS plus satellite stack therefore, accounting for the recovery equipment that performed a two perigee burn mission sequence to stays with SKYLON which had an estimated mass of GTO, such that the combined periods of the resulting 1,050 kg, the maximum deployed mass of the transfer orbits equated to exactly nine times that of satellite plus SUS was 13,950 kg. The SUS then the SKYLON in its 185 km circular orbit. performed a four burn mission: The SUS performs a three burn strategy: i. Raise the apogee of the orbit to i. Perigee burn to ITO (185 by 5,900 km 6796 km) ii. Circularise the orbit at 5,900 km ii. Perigee burn to GTO (the satellite is released from the iii. Perigee retro-burn to circularise at SUS in this orbit) 185 km for rendezvous with iii. Depress the perigee of the SUS SKYLON. orbit to 300 km (timed to rendezvous with the SKYLON) The maximum mass deployed by this mission iv. Circularise the orbit at 300 km strategy was found to be a little under 6.4 tonnes just short of the 6.5 tonnes specified in the Requirement The SUS could then rendezvous with the Specification. SKYLON that launched it and return to CSG in the payload bay for refurbishment and subsequent re-

11 flight. The maximum mass of the Electric Propulsion potential for SKYLON/SUS in such mission is given satellite delivered to the circular 5900 km orbit was in Reference 10. calculated to be 6,678 kg, giving a BOL mass in This leaves the polar orbit missions, where GEO of up to 5644 kg, using up to 20.5 kW of the requirement was for 4 tonnes into an 800 km electrical power, with a flight time of 153 days. SSO. SKYLON is not specified to go above 600 km A summary of the three Geostationary and in any case its performance at 800 km altitude in satellite launch options is shown in Table 7. a polar inclination would be a third that required. However the 4 tonnes to 800km altitude could Table 7: SKYLON D1/ SUS Performance in be achieved if the spacecraft were placed in a 300km Geostationary Missions lower altitude initial orbit by SKYLON, and then Maximum Mass of Maximum Mass of reached its operational orbit with secondary Satellite into GEO (With SKYLON/SUS Mission Satellite into GTO 320s SI apogee engine) propulsion. Using the SUS as an upper stage for this (tonnes) (tonnes) role was found to be counterproductive as it was so 185 km LEO deployment oversized for the role as to actually make matters Reusable SUS 6.39 (4.0) 9:1 resonance return worse. However the velocity increment could be transfer orbit achieved with a small upper stage optimised to take 185 km deployment Sun Synchronous payloads to their final orbit. Expendable SUS 8.08 (5.07) destructive re-entry Alternatively, given such spacecraft will require a propulsion system capable of de-orbiting the 300 km LEO deployment Reusable SUS spacecraft after 25 years to meet the ESA 5,900 km circular MEO N/A “Requirements on Debris Mitigation”, that EP satellite (8.96 GTO 5.64 20 kW HE thruster equivalent) propulsion capability could be increased with 153 day transit to GEO additional propellant to enable the satellite to raise the orbit using its on board propulsion. The Requirement Specification did not give any orbital parameters or payload performance figures for MEO. The study made an assumption VII PAYLOAD CARRIERS that the MEO mission would be to a 23222 km circular orbit, with an inclination of 56 degrees being SKYLON Large Payload Carrier (SLPC) the orbit of the Galileo navigation satellite.

In this mission scenario the SUS/Navsat stack The SKYLON’s main three trunnion mounting is not was deployed into a 300km circular, 56 degree necessarily the most convenient arrangement for the inclination LEO by SKYLON, followed by a 4 burn potential payloads, particularly payloads following SUS mission: the current generic form of a cylinder core structure i. Burn 1: The SUS / Navsat is raised leading to a ring interface with the launch system. into a transfer orbit with a 23,222 km Another disadvantage is that the three trunnion apogee mount does not allow for payloads to be captured ii. Burn 2: The stack is circularised at and reinstalled in the bay on orbit, except when 23,222 km (the Navsat is deployed in SKYLON is docked to a facility with a manipulator this orbit) arm. iii. Burn 3: After waiting for phasing the The SKYLON Large Payload Carrier (SLPC) SUS is depressed into a transfer orbit was intended to provide an alternative payload with a perigee of 300 km interface that overcomes these problems for payloads iv. Burn 4: The SUS is circularised at that are delivered to LEO. It was specified to use the 300 km for subsequent rendezvous USIS interface and was designed to carry primary with SKYLON payloads in the SKYLON payload bay, in cases

where the trunnion mounting is not appropriate. It was calculated that this mission sequence Such cases include: can deploy a navigation satellite with a mass of up to 1252 kg using this mission strategy.  where the payload’s structure cannot Escape missions were also examined and reach the trunnions even in reusable mode the SKYLON/SUS offers a  where the payload’s structural concept medium launch system capability for escape is incompatible with the trunnion missions. The capability varies greatly depending on mounting the mission and the additional propulsion that is  where the payload requires recovery assumed in addition to the SUS. A discussion of the  where the payload requires a USIS for its subsequent mission (e.g. servicing).

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As part of the study a design for the SLPC Table 8:SLPC Mass Budget that was produced by QinetiQ Space nv. It ITEM Raw Estimate Estimate with 20% (kg) design margin (kg) responded to requirements that called for a carrier Main Structure 1879.40 2255.28 that was based on the USIS and could carry satellites Deployment mechanism 54.80 65.76 up to 10 tonnes and deploy them once in orbit. Deployment Electronics 10.00 12.00 Power Bolt Mechanism 73.20 87.84 Inherent to the USIS is the ability to also dock with Power Bolt electronics 25.00 30.00 satellites that have a passive USIS docking port and USIS 460.00 552.00 return them to Earth. The requirements specify that SLPC TOTAL MASS 2502.40 3002.88 the SLPC can be mounted in either of the payload Maximum Payload Mass 10000.00 10000.00 attachment points, although in practice it would Maximum Installed Mass 12502.40 13002.88 normally be installed in the forward payload interface as this is the only one strong enough to take payloads of this mass. The requirements also specify SKYLON Small Payload Carrier (SSPC) that the power and data connections that form part of the USIS standard will be available to the payload. The requirements specification for S-ELSO did not include any provisions for small satellites through to Nanosats. Satellites in this class are increasingly important and the study team thought an illustration of how such payloads would be carried on SKYLON was required to demonstrate a complete launch infrastructure for European needs. Thus the purpose of the SKYLON Small Payload Carrier (SSPC) was to carry payloads which would be too small to be realistically carried by the main SKYLON interface Figure 13: SLPC Design Shown Stowed (left) or the SLPC. It is the SKYLON equivalent of the and Deployed (right) Space Shuttle’s Getaway Special carrier or the

Ariane 5 ASAP platform and it gives SKYLON the The resulting design is shown in Figure 13. It capability to fly small satellites for deployment and is a truss frame created from hollow circular tubes fixed payloads for short flights. made of titanium Ti6Al4V, The design for the SSPC was produced by The deployment of payloads will necessitate QinetiQ Space nv in response to a specification that the payload being rotated to vertical from the in turn had been produced in response to feedback horizontal launch position due to the orientation of from small satellite suppliers Reaction Engines SKYLON’s payload bay and the payload bay doors. received on an earlier SSPC concept. This The mechanism to achieve this mounts the USIS on a specification called for the new SSPC to carry four platform which is connected to the main frame by a satellites with a common mechanical interface; two simple hinge giving a 90 degree rotation capability. of the spaces large enough to carry Mini-satellites When stowed, twelve M12 motor driven powered (up to a mass of 500 kg) and two spaces able to carry bolts hold the platform against the frame. Before micro-satellites (up to 100 kg). Another new rotation these bolt connections can be undone, and requirement for the SSPC was the capability to remade after re-stowing. For deployment a single launch CubeSats. mechanism consisting of a linear actuator acting on The QinetiQ design shown in Figure 14 was a the centre hinge of a two bar arm pushes the frame space-frame manufactured from aluminium into position or pulls it back against the frame so that (6060T6), that occupied the space between the the motor driven bolts can secure the frame and Payload Attachment system and the payload bay end payload for re-entry and landing. walls. It was envisaged the SSPC would normally be The estimated mass for the SLPC is 3 tonnes mounted in the rear location as a payload of (Table 8) and it is strong enough to carry a 10 tonne opportunity but it can also be mounted in the forward satellite with its centre of mass 2.5 meters from the location. USIS interface plane.

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least 350 mm, with up to 1000 mm available in the small satellite locations. The SSPC can also carry four NanoSat dispensers mounted on the rear cross member. These satellites were assumed to follow the standard model established by California Polytechnic State University in their CubeSat Design Specification [11]. This specification described a “P-POD” dispenser able to carry 3U of 100 mm CubeSats which was used as the model for the SSPC dispensers. The mass of the SSPC was estimated to be 253 kg including a 10% uncertainty factor. When Figure 14: SKYLON Small Payload Carrier fully loaded this would give an installed mass of with Representative Payloads under 1500 kg, but more typically it was expected that the installed mass will be less than 1000 kg. The space-frame supports four identical interface rings with a hollow rectangular section. Any separation mechanisms that would be required VIII CSG SPACEPORT are assumed to be supplied by the payload itself or from an optional dedicated payload adapter Past work on the ground facilities required for a providing this functionality. The Mini-Satellite SKYLON launch were nonspecific with regard to locations offer an envelope corresponding to a location except exploring the impact of the launch 1500 mm cube, and the micro-satellites an envelope site latitude on performance. It also assumed all the with an 800 mm square base and a height of facilities were new and purpose built. While this 1200 mm. These envelopes include any deployment work did enable a generic set of requirements for a tip-off allowance. The design also allows the SKYLON spaceport to be derived and provide the payload to protrude below the interface plane by at basis for initial cost estimates, it did not demonstrate

Figure 15: Preferred Runway Location

14 feasibility for any specific spaceport location. existing facilities and the being The S-ELSO Requirement Specification [4] constructed (Figure 15). As apart of the proposed called for the launch site to be located at the Centre layout the main site road will require rerouting. Spatiale Guyanais (CSG) in Kourou, French Guiana. The SKYLON runway design consists of four So as part of the Study Grafton Technology main elements: supported by Jacobs Engineering UK Ltd considered i. Preparation area: 300m x 50m – The the locating of a Skylon spaceport at CSG with a purpose of this area is to allow for the preliminary scoping investigation of the site which preparation of the SKYLON prior to included a site visit in September 2012 departure, including fuelling The existing payload preparation areas, and ii. Starter strip: 500 m x 50 m – One of the other payload support facilities such as propellant runway sections exposed to highest heating loading were found suitable for the preparation of the during the whole take-off run (point of payloads. Also the liquid hydrogen and liquid rotation being the other, when incidence oxygen propellant production capabilities on-site angle of jet blast is most direct), were found to be sufficient for the launch rate iii. The main runway section: 3600 m x 50 m, envisaged for S-ELSO although new transfer with 7.5 m shoulders – To enable SKYLON facilities would be needed to reach the SKYLON to achieve take-off speed under all loading fuelling apron. Mission planning, control facilities conditions, and tracking systems were also considered and found iv. The stop-way: 1500 m x 50 m to enable to be a suitable basis with some detailed adaption for SKYLON to stop if a fault is detected and SKYLON launch operations. the launch has to be aborted. The new facilities required for SKYLON The length and grading of the SKYLON included a 5.9 km runway (including starter strip and runway has been planned according to the likely preparation area), which would be suitable for mechanical and thermal loads imposed during each SKYLON take off operations. A parallel Code F phase of the launch profile. civil runway, 3.2 km long, was included to Spaceport support facilities were also accommodate supporting air traffic movements and addressed. These consisted of the Payload Facility, also provided an alternative SKYLON landing the Maintenance Facility and an Air Traffic Control runaway. This second runway was separated from facility which included the fire station. Both the the take off runway by 250 m which meets ICAO proposed hangars were located adjacent to the requirements for the separation of dependent runways to minimise the taxi time. Once the runways. SKYLON had been loaded with payload in the A provisional location for the spaceport was dedicated payload integration facility it would be identified within the current CSG boundary taking towed forward across the end of the runways to the into account local geography, the location of all preparation area. There it would be fuelled ready for

Figure 16: Spaceport Concept Planning Layout

15 launch. SKYLON would then be pulled forward to payload loss that are some two orders of magnitude the starter strip for final pre-launch checks before better than any currently available launch system. rocket engines were ignited. The design also The Study found that the operational assumed all offices, meeting and inspection areas infrastructure required for S-ELSO could be would be collocated with the payload processing implemented within a decade assuming the design, facility. development and procurement funding was matched The design work has enabled a preliminary to the technical programme planning [12]. This time concept layout of the ground facility to be produced. scale was found possible due to the decades of This is shown in Figure 16. preparatory work already carried out on SKYLON The technical results of the spaceport study and its SABRE engines. were:  The study confirmed the feasibility of a References Launch Complex at CSG, using both existing site services and a new, purpose- 1. R.Varvill, and A. Bond, “The SKYLON Spaceplane”, built SKYLON runway, a separate civil Journal of the British Interplanetary Society, 57, pp.22, runway, plus support facilities including a 32, 2004 dedicated payload preparation facility 2. “Feasibility Study for a New European Launch Service  A launch facility layout showed a preferred – NELS List of Requirements” 4 LAU-SF/AC/2012- position for the servicing and loading 639, Issue 1, 5th April 2012, - Applicable Document 1 to facilities with respect to taxiing, fuelling, ESA statement of Work “Feasibility Study for a New runway elements, etc. European Launch Service – NELS”. GSP-SOW-12- M.05.01, Issue 1, 5th April 2012  A set of facility requirements was generated to guide the eventual implementation of the 3. M. Hempsell, J. Aprea, B. Gallagher, Greg Sadlier, “A CSG Launch Complex Business Analysis of a SKYLON Based European Launch Service Operator” IAC-14.E6.3.8, Presented at the 65th International Astronautical Congress, Toronto

Oct 2014 IX. PROGRAMME 4. “SKYLON Based European Launch Service Operator The Study produced an acquisition programme plan Requirement Specification” ELS-REL-RP-0002, Rev 2, which included all the main infrastructure element 9th January 201 developments. As shown in Figure 17 it was found 5. “SKYLON Users’ Manual”, SKY-REL-MA-0001, Rev that it would be possible to undertake the first of the 1.1 Sept 2010 four hundred orbital test flights in 2022 and be ready 6. M Hempsell, "Standardisation of Interfaces within the for the first operational flights in 2024. Space Infrastructure", Journal of the British Unsurprisingly it was SKYLON that was found to be Interplanetary Society, Vol 39, No 2, February 1986. the schedule driving element in the S-ELSO 7. “USIS Technical Requirement Specification”, Presently development, which in turn is driven by its SABRE Unnumbered Pending the Establishing of a USIS engine. The SOMA engine and Spaceport were Controlling Body, Draft F (this can be downloaded from found to be the other schedule critical items. www.Hempsellastro.com It should be noted that this programme was 8. “International Docking System Standard (IDSS) produced for the Study, matching the constraints the Interface Definition Document (IDD)” 22 Revision A, requirements placed upon it. It does not represent 13th May 2011 the actual SKYLON and SABRE development 9. M. Hempsell and A. Bond, “Technical and Operational planning Design of the SKYLON Upper Stage”, Journal of the British Interplanetary Society, Vol. 63, Pp.136-144, 2010. X. CONCLUSIONS 10. M Hempsell, R, Longstaff, B. Parkinson, “The Potential of SKYLON to Support ” The S-ELSO Study showed that SKYLON with a IAC-14-A3,P.4, Presented at the 65th International suitable upper stage and payload support systems, Astronautical Congress, Toronto Oct 2014 and operating from the existing CSG site in French 11. “CubeSat Design Specification” Rev 12 8th Jan 2009 Guiana could feasibly meet all the requirements for the next generation European launch system. In 12 “Feasibility Study of a SKYLON based European addition it could offer new capabilities such as Launch Service Operator - Final Report”, ELS-REL-RP- payload return and human spaceflight. It could also 0013, May 2014 offer its customers a response time, and a risk of

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Figure 17: S-ELSO Development Programme

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